WILJAM FLIGHT TRAINING

Instrumentation

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Table of Contents

Section 1 – Air Data Instruments

Chapter 1.1 - Air Temperature Measurement
Introduction 1.1.1
Direct Reading Thermometer 1.1.2
Electrical Thermometer 1.1.2
Static Air Temperature Sensor 1.1.3
Total Air Temperature Probe 1.1.4

Chapter 1.2 - Pitot-Static System

Introduction 1.2.1
Pitot Tube 1.2.1
Static Source 1.2.1
Alternate Static Source 1.2.2
Combined Pitot-Static (Pressure) Head 1.2.4
Operating Problems 1.2.4
Pitot/Static System Errors 1.2.4

Chapter 1.3 - Pressure Altimeter

Introduction 1.3.1
Pressure Altitude 1.3.1
Density Altitude 1.3.1
The Simple Altimeter 1.3.1
Datum Sub-scale Settings 1.3.2
The Sensitive Altimeter 1.3.3
Altimeter Displays 1.3.4
Design Errors 1.3.6
Errors due to Calibration 1.3.6
Blockages and Leakages 1.3.7
Servo Altimeters 1.3.8
Operation of a Servo-Altimeter 1.3.9
Servo-Altimeter Power Failure 1.3.10
Altitude Encoding 1.3.10
Advantages of Servo-Altimeters 1.3.10

Chapter 1.4 - Vertical Speed Indicator

Introduction 1.4.1
Principle of Operation 1.4.1
Operation of the VSI 1.4.1
Errors of the VSI 1.4.2
Faults of the VSI 1.4.2
Instantaneous Vertical Speed Indicator (IVSI) 1.4.2
Operation of the IVSI/ILVSI 1.4.2

Chapter 1.5 - Airspeed Indicator

Introduction 1.5.1
Principle of the Airspeed Indicator (ASI) 1.5.1
Operation of a Simple ASI 1.5.1
Sensitive and Servo Airspeed Indicators 1.5.1
Calibration of the ASI 1.5.2
Colour Coding of the ASI 1.5.2
ASI Errors 1.5.3

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ASI Faults 1.5.4

Calculation of CAS to TAS (up to 300 knots) 1.5.6

Chapter 1.6 - Machmeter

Introduction 1.6.1
Critical Mach Number (MCRIT) 1.6.1
Principle of Operation 1.6.1
Machmeter Construction and Operation 1.6.3
Calibration 1.6.4
Errors 1.6.4
Blockages and leakages 1.6.5
Accuracy 1.6.5
Serviceability Checks 1.6.5

Chapter 1.7 - Mach Airspeed Indicator

Introduction 1.7.1
Display 1.7.1
VMO Pointer 1.7.1
Driven Cursor 1.7.1
Bugs 1.7.1
Linkages 1.7.2
Errors 1.7.2

Chapter 1.8 - Central Air Data Computer

Introduction 1.8.1
The Central Air Data Computer 1.8.1
Conversion of Sensing Pressures 1.8.2
Digital Air Data Computer 1.8.4

Section 2 – Magnetism and Aeroplane Compasses

Chapter 2.1 - Basic Magnetism
Introduction 2.1.1
Magnetic Properties 2.1.1
Fundamental Laws of Magnetism 2.1.2
Characteristics of Lines of Magnetic Flux 2.1.3
Magnetic Materials 2.1.5
Permeability 2.1.5
Electromagnetism 2.1.5
An Electromagnet 2.1.6
Magnetic Moments 2.1.7
Period of Oscillation of a Suspended Magnet 2.1.8

Chapter 2.2 - Terrestrial Magnetism

Introduction 2.2.1
Magnetic Dip 2.2.2
Earth’s Total Magnetic Force 2.2.4
Magnetic Variation 2.2.5

Chapter 2.3 - Aeroplane Magnetism

Introduction 2.3.1
Types of Aeroplane Magnetism 2.3.1
Components of Hard Iron Magnetism 2.3.1
Components of Soft Iron Magnetism 2.3.4

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Determination of Deviation Coefficients 2.3.5

Minimum deviation 2.3.6
Maximum Deviation 2.3.6
Joint Airworthiness Requirements (JAR) Limits 2.3.7
Compass Swinging 2.3.8
The Compass Swing procedure 2.3.9
An Example of A Compass Swing 2.3.10
Deviation Compensation Devices 2.3.11

Chapter 2.4 - Direct Reading Magnetic Compass

Introduction 2.4.1
Properties of a Direct Reading Compass 2.4.1
“E” Type Compass 2.4.2
Pre-flight Checks 2.4.3
Principle of a Pendulum 2.4.3
Acceleration Errors 2.4.4
Turning errors 2.4.7

Chapter 2.5 – Remote Indicating Compass

Introduction 2.5.1
RIC Architecture 2.5.1
Principle of a Flux Detector Element 2.5.1
Flux Detector unit 2.5.4
Operation of the Remote Indicating Compass System 2.5.4
Gyroscope Element 2.5.7
Heading Indicator 2.5.7
Modes of operation 2.5.8
Synchronising Indicators 2.5.8
Manual Synchronisation 2.5.9
Operation of an RIC in a Turn 2.5.9
Advantages of a Remote Indicating Gyro Magnetic Compass 2.5.10
Disadvantages of a Remote Indicating Gyro Magnetic Compass 2.5.11

Section 3 – Gyroscopic Instruments

Chapter 3.1 - Gyroscopic Principles
Introduction 3.1.1
Principle of Construction 3.1.1
Gyroscopic Properties 3.1.2
Types of Gyroscope 3.1.4
Power Sources for Gyroscopes 3.1.5
The Disadvantages and Advantages of Air Driven Gyro’s 3.1.6
The Disadvantages and Advantages of Electrically Driven Gyro’s 3.1.6
Gyro Wander 3.1.7
Examples of Gyro Wander 3.1.10

Chapter 3.2 - Direction Indicator

Introduction 3.2.1
Basic Description of the Direction Indicator 3.2.1
Operation of the Direction Indicator 3.2.2
Errors Associated with the Air Driven Direction Indicator 3.2.3
Use of the Direction Indicator 3.2.4
Advanced use of the Direction Indicator 3.2.4
Sample Calculation 3.2.5

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Chapter 3.3 - Artificial Horizon

Introduction 3.3.1
Construction of an Air Driven (Classic) Artificial Horizon 3.3.2
Operation of an Air Driven Artificial Horizon 3.3.3
The Air Driven Artificial Horizon Erection System 3.3.4
Errors Associated with the Air Driven Artificial Horizon 3.3.5
Construction of an Electrically Driven Artificial Horizon 3.3.6
Torque Motor and Levelling Switch System 3.3.7
Electrical – Fast Erection 3.3.7
Errors Associated with the Electrically Driven Artificial Horizon 3.3.8
Remote Vertical Gyro 3.3.9
Standby Attitude Indicator 3.3.9

Chapter 3.4 - Turn and Balance Indicator

Introduction 3.4.1
Construction and Principle of Operation of a Turn Indicator 3.4.1
Operation of the Turn Indicator 3.4.2
Errors Associated with the Turn Indicator 3.4.3
Pre-flight Check 3.4.3
Construction and Operation of the Balance Indicator 3.4.3
Limitations and Errors Associated with the Balance Indicator 3.4.4
Pre-flight Check 3.4.4
Electrically Driven Turn and Balance Indicators 3.4.4
Typical Indications on a Turn and Balance Indicator 3.4.5

Chapter 3.5 - Turn Co-ordinator

Introduction 3.5.1
Principle of Operation 3.5.1

Section 4 – Flight Navigation Systems

Chapter 4.1 - Inertial Navigation System
Introduction 4.1.1
The Principle and Construction of an Accelerometer 4.1.1
Performance 4.1.3
Operation of a Gyro-Stabilised Platform 4.1.3
Setting-up Procedures 4.1.6
Levelling 4.1.6
Alignment 4.1.7
Levelling and Alignment 4.1.7
Corrections 4.1.7
Acceleration Corrections 4.1.8
Wander Azimuth System 4.1.9
The Schuler Tuned Platform 4.1.9
Errors 4.1.11
The Advantages and Disadvantages of an INS 4.1.12
Mode Selector Panel 4.1.12
Control Display Unit 4.1.14

Chapter 4.2 - Inertial Reference System

Introduction 4.2.1
Description of the Strap-Down System 4.2.1
Solid State Gyros 4.2.2

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Advantages and Disadvantages of RLGs 4.2.3

Alignment of the Inertial Reference System 4.2.3
Performance 4.2.4
The Control, Display and Output from an IRS 4.2.4
Description of a Typical IRS 4.2.5
IRS Transfer Switch 4.2.7
IRS Alignment 4.2.7
Loss of Alignment in Flight 4.2.7

Section 5 – Electronic Flight Systems

Chapter 5.1 - Radio Altimeter
Introduction 5.1.1
The Radio Altimeter System 5.1.1
Principle of Operation of a Radio Altimeter 5.1.2
Performance and Accuracy of a Radio Altimeter 5.1.3
Errors Associated with a Radio Altimeter 5.1.3
The Advantages of a Radio Altimeter 5.1.4

Chapter 5.2 - Electronic Flight Instrument System

Introduction 5.2.1
EFIS Architecture 5.2.1
Attitude Director Indicator (ADI) 5.2.3
The Horizontal Situation Indicator (HSI) 5.2.5
EFIS/IRS Interface 5.2.8
Heading Reference Switch 5.2.9
HSI Symbology 5.2.10

Chapter 5.3 - Flight Management System

Introduction 5.3.1
The Flight Management Computer System 5.3.2
Command Display Unit 5.3.4
Control Panels 5.3.5
CDU and FMC Terminology 5.3.6
The Flight Management Computer Memory 5.3.8
General FMS Operation 5.3.10
Pre-Flight 5.3.11
En-Route 5.3.12
Lateral Navigation (LNAV) 5.3.12
Vertical Navigation (VNAV) 5.3.12
Operational Notes 5.3.14
Fuel Monitoring 5.3.15
Flight Control and Management Summary 5.3.15

Section 6 – Automatic Flight Systems

Chapter 6.1 - Flight Director System
Introduction 6.1.1
Flight Director Architecture 6.1.1
Flight Director Control Inputs 6.1.2
Flight Director Computer 6.1.2
Mode Control Unit 6.1.3
Flight Director Displays 6.1.3

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Flight Director Modes of Operation 6.1.4

Mode Annunciator 6.1.10
Operation of the Attitude Director Indicator 6.1.10
The Horizontal Situation Indicator Flight Director Commands 6.1.11

Chapter 6.2 - Automatic Flight Control Systems

Introduction 6.2.1
Classification of an AFCS 6.2.1
Control Channels 6.2.2
Inner Loop Control (Stabilisation) 6.2.2
Operation of an Inner Loop Pitch Stabilisation System 6.2.3
Outer Loop Control (Stabilisation) 6.2.3
Roll Modes 6.2.4
Pitch Modes 6.2.6
Combined Roll and Pitch Modes 6.2.6
Attitude Sensing 6.2.7
The AFCS Computer (Signal Processor) 6.2.8
Servomotor Actuators 6.2.9
Autopilot Terminology 6.2.10
Cross Coupling 6.2.11
ILS Coupling 6.2.11
VOR Coupling 6.2.13
Stability Problems 6.2.13
Yaw Damper 6.2.13
Mach Trim System 6.2.15
Automatic Pitch Trim Control 6.2.15
Interlocks 6.2.17
Synchronisation 6.2.18
Instinctive Cut-Out 6.2.18

Chapter 6.3 - Automatic Landing system

Introduction 6.3.1
Basic Requirements for an Automatic Landing System 6.3.1
Automatic Landing System Terminology 6.3.1
Automatic Landing System Equipment Requirements 6.3.2
Automatic Approach, Flare and Landing Sequence 6.3.3
Weather minima 6.3.5
ICAO Categorisation for Low Visibility Landing Capabilities 6.3.5
The Fundamental Landing Requirement 6.3.6
System Reliability and Integrity 6.3.7

Chapter 6.4 - Thrust Management Systems

Introduction 6.4.1
Determining the Thrust Required 6.4.1
Calculation of Climb and Cruise Thrust 6.4.2
Cruising Methods 6.4.2
Electronic Engine Control (EEC) 6.4.2
Full Authority Digital Engine Control (FADEC) 6.4.3
Auto-throttle (A/T) 6.4.6
Thrust Lever Operation 6.4.7
Thrust Management via the Auto-throttle 6.4.7
Thrust Management Computer (TMC) 6.4.8
Thrust Mode Select Panel (TMSP) 6.4.8

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Section 7 – Warning and Recording Systems

Chapter 7.1 - Central Warning System
Introduction 7.1.1
Central Warning System Annunciator Panel 7.1.1
Aural Warnings 7.1.2

Chapter 7.2 - Altitude Alerting System

Introduction 7.2.1
Altitude Alerting System Operation 7.2.1

Chapter 7.3 - Ground Proximity Warning System

Introduction 7.3.1
GPWS Architecture 7.3.1
GPWS Modes 7.3.2
Warning System 7.3.3
GPWS Control Panel 7.3.4
Discretionary Response 7.3.5
Warning Inhibition 7.3.5
The Reporting of GPWS Events 7.3.5
Operation of the GPWS 7.3.5
Joint Aviation Requirements 7.3.11

Chapter 7.4 - Traffic Collision Avoidance System

Introduction 7.4.1
Aeroplane Installation 7.4.2
Operation of TCAS II 7.4.3
TCAS Aural Warnings 7.4.5
Information Display 7.4.6
Resolution Advisory / Vertical Speed Indicator (RA / VSI) 7.4.6
TCAS Control Panel 7.4.7
Operating Restrictions 7.4.7

Chapter 7.5 - Mach/Airspeed Warning System

Introduction 7.5.1
System Architecture and Operation 7.5.1
Maximum Operating Airspeed schedule 7.5.1

Chapter 7.6 - Stall Warning

Introduction 7.6.1
Light Aeroplane Stall Warning Device 7.6.1
Transport Category Aeroplane Stall Warning Device 7.6.1

Chapter 7.7 - Recording devices

Introduction 7.7.1
Flight Data Recorder (FDR) Requirements 7.7.1
FDR Design 7.7.2
Cockpit Voice Recorder (CVR) Requirements 7.7.3
CVR Design 7.7.3

Section 8 – Engine and System Instrumentation

Chapter 8.1 - General Engine Instrumentation
Introduction 8.1.1

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Piston Engines 8.1.1

Turbo Propeller Engines 8.1.1
Gas turbine Engines 8.1.2

Chapter 8.2 - Pressure and Temperature Sensors

Introduction 8.2.1
Pressure Measurement 8.2.1
Temperature Measurement 8.2.3

Chapter 8.3 - Pressure and Temperature Indicators

Introduction 8.3.1
Pressure Indicators 8.3.1
Temperature Indicators 8.3.3

Chapter 8.4 - RPM Indicators & Propeller Synchroniser Systems

Introduction 8.4.1
Tachometers 8.4.1
Propeller Auxiliary Systems 8.4.5
Synchronisation System 8.4.5
Synchrophasing System 8.4.6
Operation of a Synchrophasing System 8.4.6

Chapter 8.5 - Engine Torque measurement

Introduction 8.5.1
Torque Meter 8.5.1
Negative Torque Sensing 8.5.2

Chapter 8.6 - Vibration Monitoring

Introduction 8.6.1
Vibration Monitoring System 8.6.1

Chapter 8.7 - Fuel Gauging

Introduction 8.7.1
Measurement of Fuel Quantity 8.7.1
Fuel Totaliser 8.7.4
Fuel Flow 8.7.4

Chapter 8.8 - EICAS

Introduction 8.8.1
EICAS Architecture 8.8.1
Engine Displays 8.8.2
Crew Alerting 8.8.3
Master Warning/Caution Light 8.8.4
Inhibits 8.8.4
Display Messages 8.8.4
Status 8.8.5
Maintenance 8.8.6
EICAS Failure Modes 8.8.7

Chapter 8.9 - ECAM

Introduction 8.9.1
Engine / Warning (E/W) CRT Display 8.9.1
The System / Status (S/S) CRT Display 8.9.2
ECAM System Architecture 8.9.3
ECAM Control Panel (ECP) 8.9.4
Attention Getters 8.9.5

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Display Colour Coding 8.9.5

ECAM System Failure 8.9.5
Failure Categorisation 8.9.5
System Operation 8.9.6

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Chapter 1.1
Air Temperature Measurement

Introduction The temperature of air under pure static conditions at the various flight levels is
one of the basic parameters required to establish data, which is vital to the performance
monitoring of modern aeroplanes. The measurement of Static Air Temperature (SAT) by direct
means is however not possible for all types of aeroplanes, because the measurement can be
affected by adiabatic compression of the air at increased airspeeds. The boundary layer over
the outer surface of an aeroplane flying below Mach 0.2 is very close to SAT, but at higher
Mach numbers the boundary layer may be slowed down or even stopped relative to the
aeroplane. This in turn results in adiabatic compression, which will cause the air temperature to
rise well above SAT. This increase is known as ‘RAM Rise’, and the temperature indicated
under such conditions is known as Ram Air Temperature (RAT).

RAT = SAT + RAM Rise or SAT = RAT - RAM Rise

The majority of RAM Rise is due to adiabatic compression as a result of the airflow coming to
rest, with only a relatively small amount being due to friction between the surface of the
aeroplane and the high-speed airflow. The RAM Rise is always pre-calculated and is tabulated
or graphed as a function of Mach number in the Operations or Flight Manual for each type of
aeroplane, a sample of which is shown below.

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The proportion of RAM rise measured is dependent on the ability of the sensor to sense (or
recover) the temperature rise. The sensitivity is expressed as a percentage and is known as the
recovery factor. For example, if a sensor has a recovery factor of 0.80, it will measure SAT +
80% of the RAM Rise.

A useful approximation of RAM Rise is:-

( TAS kt )2
100

eg. for an estimated TAS of 460 kts deduct ( 460

100
kt )2 from the indicated air temperature, which
is approximately equal to 21°C. If this figure is subtracted from the Indicated Outside Air
Temperature (IOAT) it will give the Corrected Outside Air Temperature (COAT), which is the
best determinable value of the temperature of the air through which the aeroplane is flying, and
is that which is required for navigational purposes, ie:-

COAT = IOAT – RAM Rise

This can alternatively be written as:

TM
TS =
1 + 0.2KM 2

where: TS = SAT in degrees absolute.

TM = Indicated temperature in degrees absolute.
K = Recovery Factor.
M = Mach Number.

Various types of air temperature sensors are fitted to aeroplanes, although the particular type is
dependent on whether SAT or RAT is required, and the normal operating speed of the
aeroplane on which it is fitted.

Direct Reading Thermometer

This is the simplest type of thermometer, and it only indicates SAT. It consists of a bi-metallic
element in the shape of a helix, which expands and contracts when subjected to temperature
changes.

As the temperature changes, the helix will wind or unwind, and will cause a pointer to rotate
against a scale calibrated in degrees Celsius. This type of thermometer is normally only fitted on
low speed light aeroplanes, and passes through the fuselage or canopy with the sensitive
element projecting into the airflow. The element is also shielded so that it is not directly affected
by solar radiation.

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Electrical Thermometer
The resistance of an electrical conductor is dependent on temperature, and its magnitude
changes due to thermal changes, thus altering its ability to oppose current flow when connected
in a circuit. The temperature/resistance technique is commonly used in the Electrical
(Wheatstone Bridge) Thermometer.
TEMPERATURE
SENSITIVE R4
RESISTANCE

R1 R4 R1
40 20 0
60 20
A METER METER -70 AIR 30
C

R2 R3
R2 R3

WHEATSTONE BRIDGE CIRCUIT FOR MK. 1 INDICATOR F ANDOR MK. 1

BASIC CIRCUIT AND MK. 1A THERMOMETERS MK. 2 THERMOMETERS
(1) (2) (3)

The bridge is made up of four resistance arms, which are connected across a low voltage
source as shown above. When the system is switched on, current will flow in the circuit and will
divide at point A before flowing through R1 and R2 at strengths that will vary as the temperature
of R4 (positioned outside the fuselage) alters. At point C the currents will re-unite and will flow
back to the voltage source. If the Bridge is balanced resistance-wise, no current will flow and
the moving coil galvanometer (meter) will read zero. If the temperature (resistance) of R4
changes the bridge will become unbalanced, and the resulting current flowing through the
galvanometer will be registered as a change in temperature.

Static Air Temperature Sensor

The majority of temperature sensors use a platinum wire element which is contained either in a
probe mounted in what is termed a ‘flush bulb configuration’, or in a specially designed probe,
that is shielded from solar radiation.

SPILL
PORTS INLET
VENTURI

OUTLET HOUSING
RING

In the type of sensor shown above the probe protrudes through a hole in the aeroplane skin.

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If the sensing element is mounted flush with the aeroplane skin (flush bulb), it will sense only
the Static Air Temperature (SAT). The recovery factor, or the ratio of the indicated to the actual
temperature, of this type of sensor varies from 0.75 to 0.90, which is dependent on the
aeroplanes geometry, and location of the bulb.

Total Air Temperature Probe

For aeroplanes operating at high Mach numbers, it is usual to sense and measure the
maximum temperature rise possible. This is called the Total Air Temperature (TAT) or Ram Air
Temperature (RAT), and is the temperature of the air when it is brought to rest (or nearly so)
without the addition or removal of heat. A typical TAT probe, which is commonly used on
transport-category aeroplanes is the Rosemount probe, as shown below.

AIR FLOW

TWO CONCENTRIC
AIR SCOOP PLATINUM TUBES

AIRFLOW RADIATION
DE-ICING SHIELD
HEATERS

SENSING
ELEMENT
LOCKING SCREW

CENTRE BODY

AIRCRAFT SKIN
CEMENT

5-POLE
CONNECTOR

The advantages of this type of thermometer over the flush bulb type is that it has a virtually zero
time lag, and also has a recovery factor of approximately one. This type of probe is normally
connected directly to a flight deck indicator, and also to the Mach number module of an Air Data
Computer (ADC).

An air intake, which is mounted on top of a small streamlined strut is secured to the aeroplane
skin at a predetermined location around the nose, where it is free from any boundary layer
activity. In flight, air pressure within the probe is slightly higher than outside, and air flows
through the probe. Separation of water droplets from the air is achieved by causing the air to
turn through 90°, before it passes over the sensing element. Bleed holes in the casing also
allow boundary layer air to be drawn off due to the pressure differential, which exists across the
casing. A pure platinum resistance wire, which is sealed within two concentric platinum tubes is
used to sense the temperature, and a heating element is mounted on the probe to prevent any
ice forming. The heater has a minimal affect on the indicated temperature readings, with typical
values being 0.9°C at Mach 0.1 and 0.15° C at Mach 1.0.

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Chapter 1.2
Pitot-Static System

Introduction
The pitot and static system on an aeroplane is used to measure the total pressures created by
the forward motion of the aeroplane, and the static pressure surrounding the airframe. These
pressures are then fed to instruments, which convert the pressure differentials into speed,
altitude and rate of change of altitude. The system is alternatively referred to as a ‘Manometric’
or ‘Air Data System’.

Pitot Tube
The Pitot tube (pitot probe) is used to sense the total or pitot pressure, which is the combined
static and dynamic pressure of the airflow. The tube is fitted to the airframe with its opening
facing directly into the airflow, and the airflow comes to rest, ie. stagnates inside the tube
entrance.

A baffle plate is fitted inside the tube entrance, and is designed to intercept much of the
moisture in the airflow. A pinhole drain also allows any moisture to leak away to atmosphere
without significantly affecting the sensed pressure. A further drain is additionally provided at the
back of the probe where the pitot pressure feeds into the pitot line. This drain may either be a
pinhole, or a larger capacity moisture trap, which may only be drained on the ground, by
activating a manual release button.

The probe is mounted on a part of the aeroplane where there is minimal disturbance to the
airflow, and is designed to extend well forward into the airflow. These probes are typically
mounted close to the nose, at the wing tips, on a pylon extending well below the wing, or at the
top of the fin. The probe is also fitted with a heater, which is powered from the aeroplane
electrical supply (usually 28 Volt DC or 115 Volt AC), and is switched on as required by the
flight deck crew to prevent the formation of ice. An indicator light gives the operative state of
the system. Some types show an amber light when switched ‘OFF’, or alternatively with the
system switched ‘ON’ and the heating element failed. Most transport category aeroplanes have
at least two pitot tubes.

Static Source
The ambient pressure of the air mass surrounding the aeroplane, or "static pressure", is
obtained via a static source. The static source or ‘static vent’ senses the static pressure of the
atmosphere, which is unaffected by the airflow. To achieve this the source (vent) is located on
a part of the aeroplane where the airflow will be undisturbed by its passage, and is also

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positioned with its entrance perpendicular to the airflow. The vent is manufactured and
attached to the surface of the aeroplane so that it does not create local disturbances in the
airflow. A typical design is shown on the next page.

Vent pipe connections are installed with a slight downward angle to ensure adequate drainage,
and it is also important that the vent plates are not painted, as this would impair their thermal
efficiency. This may be indicated on the aeroplane structure next to the plate by a placard. The
direction of the airflow around the static vent may vary as the airspeed and configuration of the
aeroplane changes, and may also induce errors known as position (or pressure) errors. These
errors can be minimised by carefully positioning the static vent, or by using multiple vents to
average out the errors. This is known as ‘Static Balancing’, and is achieved by fitting vents on
either side of the aeroplane fuselage. The purpose of this is to even out any differences of
pressure that may be caused by the sideways motion of the static vents, such as will occur
during a yaw or sideslip condition. Any residual position (pressure) errors are recorded during
initial flight tests and a correction table is produced, for various airspeeds and configurations.
These readings are incorporated into the Aeroplane Operating Manual (AOM).

If failure of the primary pitot/static pressure source should occur, for example icing up of a pitot
or pressure head due to a failed heater circuit, errors may be introduced in the instrument
readings and other areas dependent on such pressure. As a safeguard against partial failure, a
standby system may be installed in some aeroplanes, whereby static pressure and/or pitot
pressure from alternate sources can be selected and connected into the primary system. A
blockage of the pitot source is not serious, as it will only affect the ASI. A blockage of the static
source will however affect all of the instruments, and it is thus common practice to provide an
alternate static supply.

Alternate Static Source

The changeover to an alternate static source is normally achieved by selector valves located in
the static lines, which are located on the flight deck, within easy reach of the flight crew. A
typical internal alternate static source installation is shown on the next page. Such a system will
only operate satisfactorily if the cabin is unpressurised and the air within the cabin is relatively
undisturbed. When calibrating the pressure/position errors of the alternate system, the
manufacturer will lay down the conditions required in respect of the position of such items as
windows, heating/ventilation and doors, all of which must be observed if the system is to work
correctly.

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Moderate or large aeroplanes will normally have a minimum of two separate static systems, and
each will be fed by a pair of balanced static vents, as shown below.

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The second pair of static vents is normally referred to as the auxiliary static ports, whose precise
location is determined during prototype development.

Combined Pitot-Static (Pressure) Head

In some light aeroplanes the complication and expense of separate pitot head(s) and static
sources is avoided by often incorporating both functions into a combined pressure head. The
combined head is usually mounted on a pylon below the wing towards the tip, thus reducing the
position/pressure error to an acceptable level although accuracy is well below the standards
achieved by more sophisticated systems. A typical combined pressure head is illustrated
below.

Static pressure is admitted to the pressure head through slots or holes cut into the static casing
at 90° to the airflow, whilst pitot pressure enters the head via the pitot input port that faces
directly into the airflow. The different pressures are then fed to the flight instruments via
separate, seamless and corrosion-resistant metal pipelines. An electric heating element is also
connected to the aeroplane electrical supply to prevent ice forming inside the pressure head,
which may obstruct the airflow.

This sensor is particularly prone to errors during manoeuvres, when changes in pitch can result
in the pitot/static sources being presented to the relative airflow in such a way that the full
dynamic pressure is not fully sensed. More significantly, dynamic pressure effects may also
intrude into the static supply. Systems using a combined pressure head thus tend to suffer from
increased pressure errors due to the positioning of the sensor, and in particular during
aeroplane manoeuvres. The combined pressure head is thus only suitable for use on relatively
small, low performance aeroplanes.

Operating Problems
Any ingestion of water is normally cleared by the use of drain holes; although, these are unable
to cope with very heavy precipitation. If the ingested water freezes, ice will block the pressure
head. It is, therefore, fitted with a heater whose function forms part of pre-flight checks.
Blockages may also occur due to the ingestion of dirt and insects etc, although the risk of
blockages may be reduced if the systems are duplicated.

Pitot/Static System Errors

The following errors may occur in pitot/static systems:-

Pressure (Position) error. This is due to the positioning of the pitot/static source and
can vary with speed, angle of attack and configuration. This error is tabulated in the
aeroplane flight manual.

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Changes in configuration (flaps, under-carriage, etc). This may cause turbulence

around the static vents, which may result in random pressure fluctuations.

Manoeuvre induced errors. These result from dynamic changes at the static vents.

Turbulence. This results in random accelerations, which will vary in magnitude, and
will make indications on the pressure-fed instruments extremely unreliable.

Use of alternate static vents. These will almost always result in a different pressure
error profile. If the alternate static vent is within the non-pressurised portion of an
aeroplane, the sensed static pressure will tend to be higher than it should be,
depending on the fan setting and window positioning etc.

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Chapter 1.3
Pressure Altimeter

Introduction
The earth is surrounded by a gaseous envelope, which is divided into several concentric layers
that extend outwards from the earth's surface, up to a height of approximately 900 km
(500 miles). The lowest layer is called the Troposphere, and extends to an altitude of
approximately 36,000 ft in temperate latitudes, whilst the outer layer is called the Exosphere.

The atmosphere is held in contact with the earth by the force of gravity, which decreases
steadily with increasing distance from the earth's surface, as does the atmospheric pressure,
eg. at the Tropopause, the pressure is approximately one quarter of its sea level value. Air
density is directly proportional to atmospheric pressure and similarly reduces with increasing
altitude. Another factor, which affects the atmosphere, is temperature, which also steadily
decreases with increasing altitude up to the Tropopause (36090 feet ISA), and thereafter
remains constant.

Relative to altitude, these variables are difficult to continuously measure and compensate for, so
the International Civil Aviation Organisation (ICAO) formulated a table in which the values of
pressure, temperature and density vary at a prescribed rate (with altitude). This table was
accepted internationally as a reference (datum) and is known as the International Standard
Atmosphere (ISA), against which aeroplane performance can be compared, and air data
instruments calibrated. The following table shows how pressure, temperature and density vary
with increasing altitude in the International Standard Atmosphere. (ISA).

Altitude Pressure Temperature Density

(metres) (feet) (hPa) (°C) (Kg/m3)

0 0 1013.25 15 1.225
2000 6562 795 2 1.007
4000 13124 612 -11 0.819
6000 19686 472 -24 0.660
8000 26248 357 -37 0.526
10000 32810 265 -50 0.414
12000 39372 194 -56.5 0.312
14000 45934 142 -56.5 0.288
16000 52496 104 -56.5 0.166

Pressure altimeters or aneroid barometers make use of the fact that the pressure in the
atmosphere decreases with increasing altitude, and are thus calibrated in terms of ISA to show
altitude instead of barometric pressure.

Pressure Altitude
Pressure altitude is the altitude above the standard datum 1013.25 hPa or mbs (29.92 inch of
mercury).

Density Altitude
Density altitude is pressure altitude corrected for temperature. Pressure and density are the
same when conditions are standard. As the temperature rises above standard, the density of
the air will decrease and the density altitude will increase.

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The Simple Altimeter

The altimeter is an instrument that is designed to measure static pressure and, using the
conditions of the standard atmosphere, convert that pressure into a value of altitude, eg. if the
pressure measured is 472 hPa, then a calibrated altimeter will indicate 19686 feet.

A simple or non-sensitive altimeter comprises of a partially evacuated barometric (Aneroid)

capsule, a leaf spring, a mechanical linkage and a pointer, as shown below.

STATIC
PRESSURE
POINTER
FROM
STATIC VENT

LEAF
SPRING
GRADUATED
CARD

GLASS
FACE

ANEROID
CAPSULE

DIAL
SETTING
KNOB
LINKAGE

These components are all housed in a container, which is supplied with static pressure from the
static vent system. As an aeroplane climbs the static pressure will decrease, which will cause
the capsule to expand and drive the pointer to indicate a higher altitude. As an aeroplane
descends, the capsule will compress due to the increasing static pressure, and the pointer will
indicate a lower altitude. The leaf spring in the instrument is designed to prevent the capsule
from collapsing and acts as a balance between the capsule and the static pressure. The
altimeter is normally calibrated at ISA +15°C and 1013.2mb (hPa) until it reads zero, although
the datum can be adjusted via the sub-scale dial setting knob. This allows the instrument to be
adjusted for different values of mean sea level (msl), and heights above an aerodrome at which
the altimeter will read zero.

Datum Sub-scale Settings

The setting of altimeters to datum barometric pressures forms part of the flight operating
procedures, and is essential for maintaining adequate separation between aeroplanes and
terrain clearance during take-off and landing. These settings have been adopted universally and

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form part of the ICAO "Q" code of communication, which consists of three-letter groups, each
having "Q" as the first letter. The codes normally used in relation to altimeter settings are:-

QFE. The pressure prevailing at an airfield, the setting of which on the altimeter sub-
scale will cause the altimeter to read zero on landing and take-off.

QNE. Setting the standard mean sea level pressure of 1013.25 hPa to make the
altimeter read the airfield elevation, ie. the equivalent height in ISA above the 1013.25
hPa pressure level. When QNE is set, the altimeter will indicate the pressure altitude,
which is the reported flight level. Flight levels occur at 500 ft intervals and are
calculated by dividing the pressure altitude by 100, eg. a pressure altitude of 10,000 ft
will equate to FL 100.

QNH. This is the actual msl pressure. Setting the pressure scale will cause the
altimeter to read the airfield altitude above sea level on landing and take-off. This
setting is also handy for checking the height above terrain or a radio mast.

2610'
CONSTANT
LEVEL 3000'
1020 3210' 1013 1000

HEIGHT ABOVE QFE LEVEL

PRESS ALTITUDE 1000 MBS AIRFIELD QFE

ALTITUDE
ABOVE M.S.L. ABOVE 1013 MBS SAS

1020 MBS QNH

1 MB = 30 FEET APPROXIMATELY

If the aeroplane remains at the same altitude winding on hecto Pascals or millibars will wind on
more height and vice versa

The Sensitive Altimeter

The sensitive altimeter is essentially the same as the simple altimeter but employs a minimum
of two aneroid capsules. This provides for a more accurate measurement of pressure and also
provides more power to drive the mechanical linkage.

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The capsules are stacked together with one face fastened down, which permits movement due
to pressure changes at the other end. The movement of the capsules in response to changes in
altitude (pressure) is transmitted via a suitable mechanical linkage to three pointers that display
(against a graduated instrument scale) the aeroplane altitude in tens, hundreds and thousands
of feet. The whole assembly is encased in a container, which is fed with static pressure, but is
otherwise completely airtight. Within the mechanical linkage a bi-metallic insert is fitted to
compensate for temperature changes that could affect the movement. As the aeroplane climbs
and air pressure falls, the capsules will expand; similarly, as the aeroplane descends, the static
pressure will increase and the capsules will contract. Since it is necessary to allow for different
values of mean sea level pressure and also to allow the altimeter to be used for indicating
altitude above the aerodrome, the altimeter is similarly provided with a means of adjusting the
level at which it will indicate zero feet. This is done via a barometric subscale mechanism, which
adjusts the mechanical linkage and operates a set of digital counters, or calibrated dial. This is
displayed in a window in the face of the altimeter, and is the datum pressure setting above
which the instrument is now displaying altitude. The desired setting is again made using the
knurled knob at the bottom of the instrument.

Altimeter Displays
A number of different types of pressure altimeter are manufactured; however, they differ in
detail depending on the altitude band covered, the accuracy of the instrument, and the method
by which the altitude is displayed. Types of display vary from multi-needle to needle plus digital
counters, with accuracy varying from 100 feet at 0 feet to 1,000 feet at 35,000 feet in early
models, to 35 feet at 0 feet to 600 feet at 60,000 feet in later models.

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A selection of typical displays is shown above; including an instrument face showing a sector
flag, barber pole display or low altitude warning sector. The sector flag, which has stripes in
black and white, appears in a window when the altitude is 16,000 feet or lower, giving the flight
crew a clear warning of approaching low altitudes during rapid descents.

The control of aeroplanes along the many air-routes in the vicinity of modern airports requires
that an aeroplanes lateral and vertical position be constantly and accurately monitored if
potentially hazardous air traffic situations are to be avoided. To provide an automated
transmission of altitude (flight level), two digitisers are normally fitted to modern altimeters. The
digitisers are fitted inside the instrument case, as illustrated in the diagram on the next page,
and are connected by a common gearing to the main shaft.

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The angular position taken up by the rotors of each digitiser relative to its stator determines the
value of a pulsed digital signal, which is produced by the assembly. The digital pulses are fed to
a code converter where any ambiguity is resolved, and the total reply pulse is modified into a
coded response suitable for use by an SSR transponder. The digitised altitude signal is not
affected by changes to the hecto Pascal’s, (millibars) counter setting, as it is always referenced
to a datum pressure of 1013.25 hPa.

Modern altimeters are also fitted with a vibrator assembly, which is designed to reduce the initial
opposition to motion of the moving parts, and also to reduce any frictional lag in the system. The
electrical supply additionally energises a warning flag solenoid in the digitiser circuit code
converter, which in the event of a power supply failure, will be de-energised, and will allow a
power failure warning flag to appear in an aperture on the dial. At the same time the code
converter will also revert to a recognised fail-safe position of "ALL Zeros".

Design Errors
Altimeters suffer from the following errors:-

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Instrument Error. Since capsule movements must be greatly magnified, it is

impossible to ignore the effect of small irregularities in the mechanism. Certain
manufacturers' tolerances thus have to be accepted, and errors generally increase with
altitude.

Pressure Error. Pressure errors arise because the true external static pressure is not
accurately transmitted to the instrument. A false static pressure arises because of
disturbed airflow in the vicinity of the pressure head or static vent. Pressure error is
negligible at low altitudes and speeds, but becomes more significant with increasing
airspeed. Correction for pressure error takes the form of a correction, which has to be
applied to the indicated altitude, and must be determined by calibration. Air Data
Computers are designed to compensate for this type of error.

Time Lag. Because the response of the capsule and linkage is not instantaneous, the
altimeter needle will tend to lag behind whenever the altitude changes. Subsequent
over-indication during descent could be dangerous, but should be allowed for in rapid
descents. Time lag is virtually eliminated in the servo-assisted altimeter.

Hysterisis Error. The capsules suffer from hysterisis, which can cause a lag in the
instrument reading during a climb or descent.

Errors due to Calibration

The calibration of the altimeter, ie. the conversion of ambient (barometric) pressure to readings
in feet, is normally based on ISA conditions. If the real atmosphere however differs, the
altimeter will not indicate the true vertical distance above the sub-scale datum. The most
significant errors are:-

Barometric Error. Barometric error occurs when the actual datum level pressure
differs from that to which the subscale has been set. It is caused by the changing
ambient barometric pressure experienced during transit and with the passage of time. If
the aeroplane flies from an area of high pressure into an area of low pressure it will
descend even though the altimeter reading will remain constant.

0
9 1
8 2

3000FT 7 3 9
0
1

6 4 8 2
0
5
7 3 600FT 8
9 1
2
6 4
5
7 3
6 4
5

2000FT

3000FT
2400FT
1000FT

3000FT

MSL
HIGH 1030MB LOW 1010MB

EFFECTIVE LEVEL OF 1030MB SETTING

The figure above illustrates the effect if the subscale is set to 1030 hPa. A subscale
error of 1 hPa is equivalent to an indicated altitude error of 28 to 30 feet. The QNH has
reduced to 1010 hPa, which will represent a altitude change of approximately 600 feet.
The subscale datum will now be at a point that is effectively 600 feet below sea level,
and this is the level from which the altimeter is measuring.

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Flying into areas of low pressure is therefore potentially dangerous since the altimeter
will over-read, which will result in the flight crew over estimating their clearance from
obstacles. Conversely if the aeroplane flies from a low-pressure area into a high-
pressure area the altimeter will under-read. When flying at low altitude it is thus good
practice to periodically reset the altimeter to minimise the barometric error.

Ornographic Error. Differences from standard may occur when air is forced to
rise/descend over hills or mountains. Low pressure tends to occur in the lee of
mountains with high pressure on the windward side. Additionally, vertical movement of
air can result in change of temperature from ISA, which will tend to induce further errors
into the altimeter readings.

Temperature Error. Temperature error arises whenever the mean atmospheric

conditions below the aeroplane differ from the standard atmosphere. If the actual
temperature lapse rate differs from the assumed one, then the indicated altitude will be
incorrect. In general, if the air below the aeroplane is warmer than standard, the air will
be less dense (low pressure) and the aeroplane will be higher than indicated.
Conversely if the air is colder than standard, it will be more dense (high pressure), and
the aeroplane will be lower than standard.

The diagram above illustrates the effect of flying from a warm atmosphere into a cold
atmosphere, whilst the surface pressure or subscale setting remains constant. The
correct altitude may be obtained from that indicated by using the navigation computer.
For ‘rule of thumb’ work, a temperature difference of 10°C, from standard, will result in
an error of approximately 4% of the indicated altitude.

Blockages and Leakages

If the static tube or vents become blocked, the pressure within the instrument case will remain
constant and the altimeter will continue to indicate the altitude of the aeroplane when the
blockage occurred, as shown on the next page.

Leaks can also take place either inside or outside the pressure cabin. Within the pressure cabin
the cabin pressure altitude will be shown rather than the true altitude. In some aeroplanes, an
emergency source inside the fuselage is available. The static pressure inside an aeroplane is
however normally different from that external to the fuselage, since it is influenced by blowers,
ventilation, etc, so that a different correction for pressure error is necessary. This is normally
given in the Aeroplane Flight Manual.

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NO CHANGE OF
STATIC STATIC PRESSURE
PRESSURE CLIMB
BLOCKAGE

UNDERREAD BLOCKAGE

NO CHANGE OF DESCENT
STATIC PRESSURE

SUB
SCALE BLOCKAGE
SET OVERREAD
1013
QNH
QFE

Servo Altimeters
Servo-assisted altimeters use the same basic principles as sensitive altimeters, whereby
pressure changes are measured using the expansion and contraction of evacuated capsules.
The instrument however uses an electrical servomotor to transmit the movement of
conventional aneroid capsules to the instrument display. A digital counter system and a single
pointer indicate the altitude. The counters are visible through four windows and show (reading
from left to right) the altitude in tens of thousands, thousands, hundreds and units of feet. The
pointer moves around the dial, which is calibrated in 50-foot divisions from 0-1,000 feet, and the
combined system indicates altitudes up to 100,000 feet.

The barometric subscale is a conventional drum counter type and is set by a knob on the front
of the instrument. The presentation can be in millibars or inches of mercury, but in some types
both settings are displayed. In this case the displays are interconnected so that any change in
millibars will produce an equivalent change in inches of mercury and vice versa.

Operation of a Servo-Altimeter
The mechanism of a typical servo-altimeter is shown schematically below.

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In this instrument the pressure sensing capsules are coupled mechanically to an electrical E
and I pick-off assembly, and any movement of the aneroid capsules is transmitted through a
linkage to the "I" bar of the E and I inductive pick-up (transducer). The amplitude of the AC
voltage output from the secondary windings depends on the degree of deflection of the "I" bar,
which is a function of pressure change. When the two air gaps become unequal, the reluctance
of each circuit changes and an electrical output is generated. The actual polarity of the output
signal will depend on whether the capsules expand or contract. The output signal is then
amplified and used to drive a motor whose speed and direction of rotation will depend on the
amplitude and phase of the signal. The motor drives the gear train, which then rotates the
altitude digital counters and the pointer. The motor also drives through a gearing arrangement
and a cam which imparts an angular movement to a cam follower, to which the "E" bar of the
inductive pick-off is attached. Movement of the "E" bar is such that it is driven until it reaches a
position where the air gaps between the "E" and "I" bars once again become equal, thus
completing the servo-loop. The system is very sensitive to small changes in pressure, and
through the motor assembly, provides adequate torque to drive the indicating system.

The datum pressure setting knob is linked to the cam via a gear train and worm shaft, as
shown. Rotation of the knob causes the worm shaft to slide forwards or backwards and rotates
the cam. Angular movement of the cam also alters the relationship between the 'E' and 'I' bars,
resulting in an electrical output which will cause the counters to rotate, and will also drive the
inductive pick-off back to its neutral position. The hPa sub-scale displays the value of the datum
pressure set.

Servo-Altimeter Power Failure

In the event of power failure the servo-altimeter will not function, and warning flags will
immediately be displayed on the dial, indicating that the AC power supply has failed. A standard
sensitive altimeter must thus be provided as a standby instrument.

Altitude Encoding
The servo-altimeter has an altitude encoder (digitiser) incorporated in it to provide a coded
height output, which when transmitted to a remote transponder (SSR) will enable the height
sensed by the capsules to be monitored on the ground, as pressure altitude. The transmitted
pressure altitude is always referenced to 1013.25 mb irrespective of the actual sub-scale setting

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Advantages of Servo-Altimeters
Servo-altimeters have the following advantages over simple and sensitive altimeters:-

¾ At high altitude very little pressure change takes place for a given change of
altitude, with the result that capsule movement is considerably less than for the
same change of altitude at lower levels. This factor reduces the efficiency of
ordinary altimeters at high levels, whereas the servomechanism will pick up a
capsule movement as small as 0.0002 inches per thousand feet.

¾ Power transmission gives greater accuracy.

¾ There is practically no time lag between the arrival of a new pressure in the
instrument, and the positioning of the counters.

¾ Being an electrical system, correction for pressure error can be made, and an
altitude-alerting device may be incorporated in the system.

¾ Although conventional altimeters now employ digital presentation, it is generally

more common with servo-altimeters. The digital presentation reduces the possibility
of misreading.

¾ A pointer is still available on the servo-altimeter for use at low level in assessing the
rate of change of altitude.

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Chapter 1.4
Vertical Speed Indicator

Introduction
A Vertical Speed Indicator (VSI) is a sensitive differential pressure gauge, which records the
rate of change of atmospheric pressure in terms of rate of climb or descent, when an aeroplane
departs from level flight.

Principle of Operation
The principle employed is that of measuring the difference in pressure between two chambers,
one of which is enclosed within the other, as shown below.

The pressure of the atmosphere is communicated directly to the inner chamber (capsule) and
through a calibrated choke or capsule case to the outer chamber. If the pressure changes, as in
a climb or descent, the lag between the capsule and outer chamber will result in a pressure
differential across the metering unit, which is a direct measure of the aeroplane's rate of climb or
descent. Movement of the capsule is transmitted via a mechanical linkage to a pointer, which
moves against a calibrated dial on the face of the instrument, where the indications are
arranged in a logarithmic scale. This allows the scale in the range 0 to 1000 feet/minute to be
more easily interpreted while, at the same time, allowing smaller changes in the vertical speed
to be registered in that range. This allows the flight crew to achieve a given flight profile more
easily.

The metering unit of the VSI provides a pressure differential across the capsule case for any
given rate of climb or descent, whilst compensating for variations in temperature and pressure
of the atmosphere with changes in altitude. The compensation is achieved by incorporating in
the metering unit both an orifice and a capillary, whose sizes are chosen so that indicator
readings remain correct over a wide range of temperature and altitude conditions.

Operation of the VSI

In level flight the pressure inside the capsule and the case will be the same, so the pointer will
remain in its horizontal position, indicating a zero rate of climb. If the aeroplane climbs, the
static pressure in the capsule will decrease at a quicker rate than that in the casing and the
capsule will collapse slightly, and will cause the pointer to indicate a rate of climb. Conversely if
the aeroplane descends the static pressure in the capsule will increase at a higher rate than that

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in the casing and the capsule will expand slightly, causing the pointer to indicate a rate of
descent.

Errors of the VSI

The VSI can suffer from the following errors:-

Lag. When an aeroplane is suddenly made to climb or descend, a delay of a few

seconds occurs before the pointer settles at the appropriate rate of climb or descent,
which is due to the time required for the pressure differential to develop. A similar delay
will occur in the pointer showing a zero rate of climb or descent when the aeroplane
resumes level flight.

Instrument Error. This error is due to the manufacturers' tolerances. However, in

service the instrument pointer can be re-adjusted to the zero position using a screw
adjustment.

Manoeuvre Induced Error. Errors induced by manoeuvres or flight in turbulence can

cause any pressure instrument to misread for up to 3 seconds at low altitudes and up to
10 seconds at high altitudes. The times for the VSI may be even longer. Thus, during
any manoeuvre involving a change of attitude, absolute reliance must not be placed on
the VSI, with pitching resulting in the greatest error.

Faults of the VSI

The following faults will have an adverse affect on the VSI reading:-

Blockages. A blockage in the static line will render the instrument completely
unserviceable, and the pointer will register zero regardless of the aeroplane's vertical
speed.

Breakage or Leakage in the Static line. A breakage or leakage in the static pressure
supply line will cause the static value to change as the breakage occurs, eg. if the
breakage occurs in a pressurised section of the aeroplane the VSI will initially show a
high rate of descent and will then stabilise to give a zero indication. This reading will be
maintained until the aeroplane descends below the cabin altitude pressure.

Instantaneous Vertical Speed Indicator (IVSI)

The Instantaneous Vertical Speed Indicator (IVSI) is also sometimes referred to as the Inertial
Lead Vertical Speed Indicator (ILVSI). The basic construction of this instrument is shown on
the next page. It consists of the same basic elements as the conventional VSI, but it is
additionally fitted with an accelerometer unit that is designed to create a more rapid differential
pressure effect, specifically during the initiation of climb or descent. The accelerometer
comprises of two small cylinders or dashpots, which contain inertial masses in the form of
pistons that are held in balance by springs and their own mass. The cylinders are connected in
the capillary tube system leading to the capsule and are thus open to the static pressure source.

Operation of the IVSI / ILVSI

If a change in vertical motion is initiated, the resultant vertical acceleration, ie. due to the
change in vertical velocity along the aeroplanes vertical axis, will produce a force and the
pistons will be displaced from their neutral position. The pistons are arranged so that one
responds to nose up pitch changes (positive G) by reducing the capsule pressure, whilst the
other will respond to nose down pitch changes (negative G) by increasing the capsule pressure.
This will in turn create an immediate pressure change inside the capsule, and will produce an
instantaneous movement of the indicator pointer, and in the correct sense, to the initiation of a
climb or descent manoeuvre. The errors are generally the same as those affecting the
conventional VSI, although the lag and induced manoeuvre errors will be virtually eliminated,

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with the instrument providing a rapid response to changes in the aeroplanes vertical flight path.
The accelerometer output will decay after only a few seconds, and the pistons will automatically
return to their neutral position, by which time a steady pressure differential will have been
established across the metering unit. The instrument will then continue to behave like a
conventional VSI.

The IVSI is however affected by the acceleration forces, ie. g–forces, which act on the pistons
during steep turns when the angle of bank is in excess of 40°, and may produce a false reading,
known as ‘Turning Error’.

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Chapter 1.5
Airspeed Indicator

Introduction
The airspeed at which an aeroplane is travelling through the air is essential to the pilot, both for
the safe and efficient handling of the aeroplane and as a basic input to the navigation
calculations.

Principle of the Airspeed Indicator (ASI)

When an aeroplane is stationary on the ground it is subject to normal atmospheric or static
pressure, which acts equally on all parts of the aeroplane structure. In flight the aeroplane
experiences an additional pressure due to the aeroplane's motion through the air, which is
known as dynamic pressure, and is dependent upon the forward motion of the aeroplane and
the density of the air, according to the following formula:

PT = 1/2ρV2 + PS

where PT = total or pitot pressure(also known as total head pressure or stagnation pressure)
PS - static pressure
ρ - air density
V = velocity of the aeroplane (TAS)

Re-arranging the formula, the difference between the pitot and static pressures is equal to
2
½ ρV (dynamic pressure). The airspeed indicator thus measures the pressure differential
between the two sources, and provides a display indication graduated in units of speed.

Operation of a Simple ASI

In the simple ASI, a capsule acting as a pressure sensitive element is mounted in an airtight
case, as shown on the next page. Pitot pressure is fed into the capsule and static pressure is
fed to the interior of the case which, when the aeroplane is in motion, will contain the lower
pressure. A pressure difference will cause the capsule to open out with any movement being
proportional to the pressure differential across the capsule skin (pitot - static). A mechanical
linkage is used to transfer the capsule movement to a pointer that moves over a dial, and which
is normally calibrated in knots. A bi-metallic strip is also incorporated in the mechanical linkage
to compensate for any expansion/contraction of the linkage caused by temperature variations.

Sensitive and Servo Airspeed Indicators

Sensitive and Servo airspeed indicators both use the same principle of operation as the simple
ASI. The sensitive ASI uses a stack of two or more interlinked capsules, which are connected
to two pointers via an extended gear train. This enables the instrument to respond to smaller
pressure changes and thus smaller changes in airspeed. The capsule assembly has a linear
pressure/deflection characteristic, which is more closely controlled than the single capsule used
in the simple ASI.

The servo airspeed indicator also uses an electrical linkage rather than a mechanical linkage to
position the indicator needles, which is done by firstly amplifying the error signal

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THE AIRSPEED INDICATOR IS A

DIFFERENTIAL PRESSURE GAUGE

Calibration of the ASI

Standard datum values are used in the calibration of air speed indicators since dynamic
pressure varies with air speed and air density. Density also varies with temperature and
pressure,. The values used are the sea level values of the ICAO International Standard
Atmosphere (ISA).

Colour Coding of the ASI

The scale is calibrated in terms of speed, usually knots or miles per hour (MPH), but in some
cases may be kilometres per hour (KPH). It is thus essential that you know which terms are
being displayed on the ASI. On light aeroplanes the dial is normally colour coded as shown on
the next page, with the coloured segments indicating the following:-

White arc. This arc extends from VSO (stall full flap) to VFE (maximum speed with flaps
extended), and marks the flap operating speed range.

Green arc. This arc extends from VSI (stall clean) to VNO (normal operating speed), and
is the normal operating speed range.

Yellow. This arc extends from VNO to VNE (never exceed speed), and denotes the
cautionary speed range. Operations within this speed range should not be carried out
except in smooth air only.

Red Radial line. This line marks VNE.

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Other lines may also be used, eg. a blue radial line, which is sometimes used to indicate the
maximum rate of climb speed in a twin-engine aeroplane with one engine inoperative (VYSE).
Some ASI’s also have adjustable bugs that can be used to set a ‘target’ speed, eg. the
threshold speed.

RED STALL SPEED

FULL FLAP
240 40 VS0
AIRSPEED V S1 STALL SPEED
220 60 CLEAN

V NE (BOTH WINGS LEVEL

200 80 MAXIMUM WEIGHT)
NEVER
EXCEED KNOTS
SPEED 180 100 WHITE

160 120
140
YELLOW V FE
MAXIMUM FLAPS
EXTENDED SPEED
VNO
GREEN
NORMAL OPERATING
LIMIT SPEED

ASI Errors
The dial of the ASI is calibrated to a formula, which assumes constant air density (standard
mean sea level) and no instrument defects. Any departure from these conditions, or disturbance
in the pitot or static pressures being applied to the instrument, will result in a difference between
the indicated and true air speeds. The following sources of error exist:-

Instrument Error. This error is caused by the manufacturers' permitted tolerances in

the construction of the instrument. This error is determined by calibration and if it is
found to be significant is recorded on a calibration card. This correction is normally
combined with that for pressure error.

Pressure Error. This error arises from the movement of the air around the aeroplane
and causes disturbances in the static and pitot pressure. The causes of this error are:-

Position of the Pitot-Static Sensors. This can alter the pressures being fed
to the instrument, and is particularly so in the case of a combined pitot-static
head where the dynamic pressure component may significantly affect the static
supply. To minimise this source of error separate static vents are positioned
well away from the pitot head, which can result in a 95% reduction in the overall
pressure error. The position/pressure error is normally determined by
calibration, and a pressure error card is tabulated in the Aeroplane Flight
Manual. This card may also incorporate any instrument error calibrations.

Manoeuvre Induced Error. This is caused by changes in the aeroplanes

attitude and/or configuration and is normally only short term. The main sources
of error is normally in the static supply, but since the transient affects of
manoeuvre induced error are not predictable or avoidable, the flight crew must
be aware of this problem.

The pressure error will change if any of the following vary:-

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¾ Airspeed
¾ Angle of attack
¾ Configuration (flap setting, undercarriage etc)
¾ Position of the pitot/static sources and sideslip

Compressibility Error. The calibration formula for most airspeed indicators does not
contain any compensation for the fact that the air is compressible. At low airspeeds this
is insignificant but at airspeeds over 300KTAS this factor becomes significant. This is
especially so at high altitudes where the less dense air is easily compressed.
Compressibility causes an increase in the measured value of dynamic pressure, which
will cause the ASI to over-read. Thus, compressibility varies with airspeed and altitude.
The error and correction can be compensated on some mechanical navigation
computers but is tabulated against altitude, temperature and CAS in the handbooks of
others.

Density Error. Dynamic pressure varies with airspeed and density of the air. In
calibration, standard mean sea level pressure is used; thus, for any other condition of
air density, the ASI will be in error. As altitude increases, the density decreases and the
indicated airspeed (IAS), and thus equivalent air speed (EAS) at speeds in excess of
300 KTAS, will become progressively lower than the true air speed (TAS). For example
at 40,000 ft the density is only ¼ of its msl value. The dynamic pressure, which is
proportional to TAS2, will thus be ½ the msl value for the same TAS, ie. an aeroplane
flying at 400 KTAS will have an IAS of 200 knots. The following formula will help to
establish the relationship:

ρ altitude
EAS = TAS
ρ sea level
For accuracy, the correction of CAS to TAS is done on a navigational computer using
the ambient temperature (outside air temperature), at the required pressure altitude. A
useful formula for estimating TAS is:-

TAS = CAS + (1.75% of CAS per 1000 ft of altitude)

For example for a CAS of 100 knots at 10,000 ft:

TAS = 100 + (1.75/100 x100 x 10) = 117.5 knots

The relationship between the various air speeds is as follows:

¾ Air Speed Indicator Reading (ASIR) + Instrument Error Correction=Indicated Air

Speed(IAS)

¾ IAS + Pressure Error Correction = Calibrated Air Speed (CAS)

¾ CAS + Compressibility Error Correction = Equivalent Air Speed (EAS)

¾ EAS + Density Error Correction = True Air Speed (TAS)

In practice, the corrections are combined to give:

¾ ASIR + Instrument Error Correction + Pressure Error Correction = CAS

¾ CAS + Compressibility Error Correction + Density Error Correction = TAS

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ASI Faults
The following faults may occur in the ASI:

Blockages. A blockage of the pitot tube, as shown on the next page, possibly due to
ice, will cause the ASI not to respond to changes of speed in level flight. The capsule
will however behave as a barometer or altimeter capsule, and will react to changes in
the static pressure. If the aeroplane climbs, the ASI will indicate an increase in airspeed
(over-read) and if it descends, it will indicate a decrease in airspeed (under-read).
CLIMB DESCENT

PITOT BLOCKAGE STATIC PITOT BLOCKAGE STATIC

A. STATIC IN CASING A. STATIC IN CASING

DECREASES INCREASES
OVERREAD UNDERREAD
B. STATIC IN CAPSULE B. STATIC IN CAPSULE
REMAINS HIGH REMAINS LOW

If the static line is blocked, the ASI will over-read at lower altitudes, and under-read at
higher altitudes than that at which the line became blocked.

CLIMB DESCENT

STATIC STATIC
PITOT BLOCKAGE PITOT BLOCKAGE

A. STATIC IN CASING A. STATIC IN CASING

REMAINS HIGH REMAINS LOW
UNDERREAD OVERREAD
B. STATIC IN CAPSULE B. STATIC IN CAPSULE
REDUCES INCREASES

Leaks. A leak in the pitot system will cause the ASI to under-read, whilst a leak in the
static line will cause the ASI to over-read in an unpressurised fuselage (cabin pressure
is usually lower than the atmospheric static pressure), and under-read in a pressurised
aeroplane (cabin pressure higher than static).

Whilst any under-reading of the ASI is undesirable, it is not necessarily dangerous, but over-
reading of the ASI is dangerous, since a stall will occur at a higher indicated airspeed than that
specified for the aeroplane.

Some modern ASI’s also employ coloured flags and needles as attention getters, ie. to indicate
any electrical or transmission failure, and also to draw attention to important altitude indicators.

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Calculation of CAS to TAS (up to 300 knots)

To calculate TAS from CAS using a CRP5 rotate the inner disc of the computer to align the
outer edge of the airspeed window with the pressure altitude (altitude with 1013 mb set on the
altimeter subscale) inside the window. Position the cursor through the CAS on the inner circular
slide rule scale and read off TAS on the outer scale.

Example: If the pressure altitude is 18,000ft, the COAT is –30°C and CAS is 170 knots, then
the TAS will be:-

Solution: In the airspeed window set 18 (pressure altitude x 1000 ft) opposite –30°C
on the COAT scale. Next position the cursor through the CAS of 170 knots on the inner
slide rule scale and read of the TAS of 220 knots on the outer scale.

Note: As a reminder of which way to read CAS to TAS there is a red CAS (RAS) on the inner
scale between 35 and 40, and a red TAS on the outer scale also between 35 and 40.

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Chapter 1.6
Machmeter

Introduction
During flight, aeroplanes emit pressure disturbances (sound waves), which radiate in all
directions at the speed of sound. As the speed of an aeroplane increases, it gets progressively
closer to the waves ahead of it until, at the speed of sound, the pressure waves combine into a
shock wave and attach themselves to the aeroplane. The effect of the shock wave is to greatly
increase the drag forces associated with the aeroplane, and to significantly alter its stability and
control characteristics. It is therefore extremely important for the flight crew to know how close
to the speed of sound the aeroplane is. The instrument used to indicate this is the Machmeter,
which measures the ratio of the aeroplane’s airspeed (TAS) to the local speed of sound, and
displays it as a Mach number.

Mach No. = V/A (Where V = TAS and A = Local Speed of Sound)

Critical Mach Number (Mcrit)

This is the speed of an aeroplane that is defined in terms of the speed of the free stream airflow.
In practice the speed of the local airflow at any point around the structure varies considerably
and the local airflow may reach the speed of sound when the free stream speed is much lower.
The free stream speed at which any element of the local airflow reaches the speed of sound will
result in local shockwaves forming on the structure. The free stream Mach number then
becomes the Critical Mach Number (Mcrit). Unless specifically designed for that purpose, an
aeroplane should not be flown beyond Mcrit and the value of Mcrit is highlighted by an index mark
on the face of the Machmeter.
CRITICAL
MACH
INDEX

.8
.7
MACH .9
.6

.5 1.0

In commercial operations, it is common practice to set a lower limit known as the Maximum
Operating Mach number (Mmo)

Principle of Operation
The Speed of Sound varies only with temperature. As the temperature increases so the local
speed of sound increases and vice versa. The precise relationship is that the speed of sound
(A) varies with the square root of the absolute temperature (° Kelvin - K) of the environment.
This can be written as New A/ Old A = √New K/ √Old K.. Using the known values of the speed of sound
at ISA mean sea level of 661.7knots where the temperature (in degrees Kelvin) is 288°, this can
be re-written as:

A = 38.94√ temperature in Kelvin.

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or [644 + 1.2t] where t = OAT in °C (estimation only).

N.B: ISA temp = 15° - (pressure altitude x 2° per 1000 feet)

Since Mach number is the ratio of TAS to local speed of sound, it may also be rewritten in terms
of pitot and static pressures.

TAS (V) is a function of the dynamic pressure and the local density. Dynamic pressure is, of
course, the difference between pitot (P) and static (S) pressures, ie. P-S (airspeed capsule)

The local speed of sound (A) is also a function of temperature. According to basic gas physics,
temperature is also a function of static pressure (S) and density (altitude capsule).

Since Mach No. = V/A we can rewrite this in a simplified form so that:

V ∝ q/ρ A∝°K
∴ ∝ P-S / ρ ∴ ∝ S/ ρ

ρ is common to both so that:

Mach No. = V/A and is a function of P-S / S

The Machmeter is thus designed to measure the ratio of pitot excess pressure (the difference
between pitot and static) and static pressure.

A relationship also exists between CAS, TAS and Mach Number under ISA conditions. With
increasing altitude the following graphs depict what will occur if one value remains constant:

This is summarised in the following tabular format, including the affect of increasing altitude on
the Local Speed of Sound (LSS).

CAS / IAS TAS MACH LSS

constant increases increases decreases
decreases constant increases decreases
decreases decreases constant decreases

The opposite will occur with decreasing altitude under ISA conditions.

Even if the CAS or MN is unaltered on purpose the MN may change due to a change in the
OAT, eg. high temperature to low temperature, the LSS will decrease, thus to maintain a

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constant MN the TAS must decrease. For example if two aeroplanes are travelling at the same
MN, but at different flight levels the aeroplane at the lower flight level will have a higher TAS.

If an aeroplane climbs or descends through an Isothermal Layer the graphs on the next page
depict what will happen to the other airspeeds, if one of them remains constant. At a constant
Mach number the LSS will remain constant so the TAS will also remain constant, although the
CAS will reduce during a climb, but will increase during a descent, due to the density error.
Conversely at a constant CAS the TAS and Mach number will both increase during a climb, but
will both reduce during a descent.

If an aeroplane climbs or descends through an Inversion Layer the graphs below depict what
will happen to the other airspeeds, if one of them remains constant.

During a climb at a constant Mach number the LSS will increase due to the warmer air so the
TAS will increase, but the CAS will reduce due to the reduction in density. If the aeroplane
climbs at a constant TAS both the Mach number and the CAS will reduce. Conversely if the
aeroplane climbs at a constant CAS the Mach number and the TAS will both increase. During a
descent the reverse will occur, because the LSS will decrease due to the colder air, whilst the
density will increase.

Machmeter Construction and Operation

A typical Machmeter, as shown below, consists of a sealed case containing two capsule
assemblies placed at 90° to each other, and a series of mechanical linkages.

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The first capsule unit is an airspeed capsule, and is connected to the pitot pressure pipeline,
while the interior of the instrument case is fed with static pressure. The second capsule unit is
an aneroid capsule, which responds to changes in static pressure. The airspeed capsule
measures the difference between pitot and static pressure and expands or contracts in
response to air speed changes. The airspeed linkage transfers movement of the capsule to the
main shaft, and causes the shaft to rotate, thus moving a pivoted ratio arm in the direction A-B.

The altitude (aneroid) capsule expands or contracts, and responds to changes in altitude.
Movement of the capsule is transferred to the ratio arm via a spring and pin, thus causing it to
move in the direction C-D.

The position of the ratio arm is therefore dependent on both pitot excess and static pressure.
Movement of the ratio arm controls the ranging arm which, through the linkage and gearing.
This turns the pointer, and displays the Mach Number corresponding to the ratio of pitot excess
pressure and static pressure. Any increase in altitude and/or airspeed will result in a higher
Mach number. The Critical Mach number is indicated by a specially shaped lubber mark, which
is located over the Machmeter dial. It is adjustable so that the critical Mach Number for the
particular type of aeroplane may be displayed.

Calibration
Machmeters are calibrated to a formula relating Mach number to atmospheric conditions of
pressure and density.

Errors
As Mach number is a function of the ratio of pitot excess pressure to static pressure, only those
errors in the measurement of this ratio will affect the Machmeter reading. The errors are:-

Instrument Errors. Like all instruments, Machmeters are subject to manufacturing

tolerances, although these are extremely small.

Pressure Errors. These errors are small at the altitudes and speed ranges where
Machmeters are used.

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Density, temperature and compressibility errors. These errors are eliminated

because density changes do not alter the dynamic pressure / static pressure ratio.

Blockages and Leakages

Blockages and leaks have the same affect as they do on the ASI, but the effects may be
increased due to the 2 capsules; a blockage in the static system will cause the Machmeter to
over-read during a descent and under-read during a climb.

Accuracy
The accuracy of the Machmeter is within +/- 0.01M during its normal operating range, but
increases to +/- 0.02 M at the limits of that range.

Serviceability Checks
The instrument should read zero when the aeroplane is stationary and a rough check of
IAS/TAS against Mach number should be carried out using the CRP5 computer.

The TAS can be calculated using the CRP5 by setting the OAT against the mach index and
reading off the TAS on the outer scale against a set Mach Number (MN). The LSS can also be
determined by reading the speed on the outer scale, which relates to Mach 1.0 on the inner
scale.

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Intentionally Left Blank

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Chapter 1.7
Mach Airspeed Indicator

Introduction
The Mach Airspeed Indicator (MASI) is designed to combine the functions of both the airspeed
indicator and the Machmeter in a single instrument. As the Machmeter contains an airspeed
capsule it is a relatively simple matter to incorporate the IAS output from the capsule into the
display.

Display
To improve clarity, digital readouts are provided for Mach number (MN) and IAS, as well as a
conventional pointer for IAS.

VMO Pointer
A special feature of the M/ASI is the VMO Pointer (Barber’s Pole), which has red and white
stripes. At low altitudes, the VMO Pointer retains a fixed position, which indicates the maximum
operating IAS permitted. As altitude increases, VMO corresponds to an increasing MN until MMO
is reached, at which point the VMO Pointer reading will progressively reduce to reflect the
overriding MN limit. Movement of the VMO Pointer can be mechanical, although the M/ASI
display is normally driven from the Air Data Computer (ADC).

Driven Cursor
The Driven cursor can be adjusted, either by the flight crew or automatically by the aeroplane’s
systems, to indicate target speeds for particular phases of flight.

Bugs
Bugs are indexes, which are set prior to flight to indicate important speeds such as V1, V2, VR
etc.

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Linkages
The M/ASI is often linked to the auto-throttle / flight management systems to provide two-way
feedback. Further links to visual and audio warning devices may also be incorporated.

Errors
The MASI suffers from the same errors as the ASI and Machmeter; namely instrument,
pressure and manoeuvre errors.

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Chapter 1.8
Central Air Data Computer

Introduction
Many of the primary flight instruments on an aeroplane are dependent on pressures that are
transmitted from the pitot/static probes through a system of pipelines before reaching the
sensors in the instruments. Larger aeroplanes require longer pipes, which would result in
increased lag errors and a greater risk of breakage/leakage. To overcome this, and to create
other benefits, most modern transport category aeroplanes use Central Air Data Computers
(CADC).

The Central Air Data Computer

In its basic format, a CADC is an analogue device that produces electrical signals equivalent to
pitot and static pressures, and Total Air Temperature (TAT), as shown below.

These signals are computed within the CADC to produce electrical output signals, which are
equivalent to:

¾ Altitude
¾ CAS
¾ Vertical speed
¾ Mach number
¾ TAS

CADC’s are powered whenever the aeroplane AC busbar’s are powered, and the output signals
are used to operate various flight instruments and aeroplane systems. The flight instruments
supplied by the CADC are the:

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¾ Altimeter
¾ Airspeed Indicator
¾ Vertical Speed Indicator
¾ Machmeter

The aeroplane systems which are typically supplied are the:-

¾ Flight Data recorder (FDR).

¾ Flight Management System (FMS).
¾ Automatic Flight Control System (AFCS).
¾ Transponder.
¾ Ground Proximity Warning System (GPWS).
¾ Power Management Computer (PMC).
¾ Flight Director System (FDS).

Calculated values of True Airspeed (TAS) and Total Air Temperature (TAT) or Static Air
Temperature (SAT) are also normally digitally displayed on the aeroplane’s instrument panel.
The CADC additionally compensates automatically for both temperature and compressibility
effects at high Mach numbers, thus enabling accurate instrument readings to be computed over
a wide range of altitudes and airspeeds. The disadvantage of the CADC is that unlike
conventional pressure instruments it requires power to operate, so a back-up system is
provided, via either an alternate power supply, or by simple pressure instruments. Two CADC’s
are normally fitted, to guard against a single failure, and each computer is supplied from
separate Pitot-static sources.

Conversion of Sensing Pressures

Transducers are used to convert the sensed pressures into electrical signals; one senses the
static pressure, whilst the other senses the dynamic pressure. A typical pressure transducer
utilises the expansion of a diaphragm or capsule to actuate an electrical pick-off, as shown
below, for the CADC Altitude Module.

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As the capsule changes shape, due to a pressure change, the position of the ‘I’ bar relative to
the ‘E’ bar will be altered. When the two air gaps are unequal, the reluctance of each circuit will
change, and an electrical output will be generated, as shown on the next page. The resultant
signal will be amplified and applied to a servomotor, which, via an output shaft, will move the ‘E’
bar until the air gaps equalise. When this occurs no further signals will be fed to the amplifier,
and the ‘I’ bar will be in a null position. The output shaft is also connected to a control (CX)
synchro, which has power applied to its rotor, thus enabling the angular position to be measured
in terms of pressure.

NO OUTPUT AIR ELECTRICAL 'I' BAR MOVED

GAP ABOUT PIVOT
PICK-OFF COIL OUTPUT (UNEQUAL AIR
GAPS)
AC AC
SUPPLY SUPPLY
TO TO
PRIMARY PRIMARY
COIL COIL
PRIMARY COIL
'I' BAR IN
MAGNETIC NULL POSITION
CIRCUIT (EQUAL AIR
GAPS)

PICK-OFF COIL AIR

GAP

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Some systems alternatively use Piezo Electric Transducers, which consist of a stack of quartz
discs, where each disc has a metallic pattern deposited on it. When pressure is applied to the
stack, it will cause it to flex or bend, and an electrical signal will be produced. The polarity of the
resultant signal will be dictated by the direction of flex, and its overall strength will be governed
by the amount the disc flex.

Digital Air Data Computer

Some modern aeroplanes are alternatively fitted with Digital Air Data Computers, which use
digital computing and electronic circuits instead of servomotors. In this system the analogue
inputs are directly converted into digital outputs before computation, but if analogue outputs are
required, they have to be acquired from the digital format. A typical system is shown below.

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Chapter 2.1
Basic Magnetism

Introduction
The operating principles of direct reading compasses are based on the fundamentals of
magnetism, in particular the interaction between the magnetic field of a suitably suspended
magnetic element, and the earth’s magnetic field.

Magnetic Properties
The three principle properties of a simple permanent bar magnet are that:-

¾ It will attract other pieces of iron and steel.

¾ Its power of attraction is concentrated at each end of the bar.

¾ When suspended so as to move horizontally, it will always come to rest in an

approximately north - south direction.

A region of influence, called a magnetic field also extends outside a magnet into the
surrounding space, which is made up of invisible lines of magnetic force, or magnetic flux. This
is best demonstrated by sprinkling iron filings on a piece of paper placed over a magnet.

This experiment illustrates that magnetism is concentrated at a magnet's extremities, which are
called poles. Additionally a freely suspended magnetised rod will always align itself
approximately in a north-south orientation. The end, which seeks north, is called the north
seeking or red pole, and the end, which seeks south, is called the south seeking or blue pole.

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The earliest known form of magnetism was the lodestone, which was a natural mineral found in
Asia. It was discovered that if a piece of this ore was suspended horizontally by a thread, or
floated on wood in water, it would automatically align itself in a north-south direction, as shown
on the next page.

This characteristic led to its use as a compass, and the name lodestone, meaning leading
stone. This occurs because the earth itself is a huge magnet with it's own magnetic field.

The fields interact with each other and the lodestone will align itself according to the
fundamental laws of magnetism. Other than the earth itself, lodestone is the only natural
magnet; and all other magnets are produced artificially. For example magnetism can be
induced in an iron bar by stroking it with a piece of lodestone. Another type of magnet is the
electromagnet, where an electric current produces the magnetic field. Magnets are also often
classified by their shape, and can exist as either horseshoe, bar or ring magnets.

A magnet can be demagnetised by:

¾ heating it to a temperature known as its Curie Point.

¾ hitting it with a hammer.
¾ degaussing it with an alternating magnetic field.

Fundamental Laws of Magnetism

The fundamental laws are as follows:

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¾ Red or blue poles cannot exist separately.

¾ Like poles repel each other and unlike poles attract.

¾ The force of attraction or repulsion between the two magnets varies inversely as the
square of the distance between them.

¾ A line joining the poles is called its magnetic axis.

Characteristics of Lines of Magnetic Flux

The region in which the force exerted by a magnet can be detected is known as the magnetic
field. This field consists of magnetic flux, which is measured in Webers (Wb), and may be
represented, in direction and intensity by lines of flux.

¾ The lines of flux have direction or polarity. They flow from the north pole to the
south pole outside the magnet, but flow from the south pole to the north pole within
the magnet. These lines are continuous and always form complete loops.

¾ The lines of flux will not cross one another, like poles repel, as shown on the next
page

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¾ Lines of flux tend to form the smallest possible loops, ie. unlike poles attract.

¾ Magnetic flux is established more easily in some materials than in others. All
materials, whether magnetic or not, also have a property called "reluctance" which
resists the establishment of magnetic flux, and equates to the resistance found in
an electrical circuit.

¾ Lines of magnetic flux can also be distorted by the interaction with other lines of
flux, as shown below.

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Magnetic Materials
Theoretically all materials are affected to some extent by a magnetic field, and are categorised
as follows:

Ferromagnetic. This is the property of a material, which enables it to become a

permanent magnet, ie. when placed in a magnetic field the material will become
magnetised. The most common materials are iron, cobalt, nickel and alloys of these
materials. Ferromagnetic materials can also be divided into two further categories:-

Hard Iron. This is a material, which is difficult to magnetise, but when it is

removed from the magnetic field, it will retain the magnetism for a considerable
length of time, unless it is subject to a strong demagnetising force, eg. cobalt
and tungsten steel. This is known as a permanent magnet.

Soft Iron. This is a material, which is easily magnetised, but once removed
from the magnetic field easily loses its magnetism, eg. silicon iron. This is
known as a temporary magnet.

The above terms are also used to describe the magnetic effects, which occur in aeroplanes.

Paramagnetic. This is the property of a material, which when placed in a magnetic

field slightly attracts the lines of magnetic force, but once removed loses its magnetism.
The most common materials are platinum, manganese and aluminium.

Diamagnetic. This is the property of a material, which when placed in a magnetic field
slightly repels the lines of magnetic force. The most common materials are copper and
bismuth.

Permeability
Permeability (µ) is the ease by which magnetic flux can be induced into a material, and can be
compared to conductance in an electrical circuit. It is the ratio of B/H, where B is the induced
magnetic flux, and H is the magnetising force.

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Electromagnetism
When current flows through a conductor a magnetic field is produced around it, whose
magnitude is directly proportional to the amount of current flow.

The direction of the field around the conductor is determined by the direction of the conventional
current flow, which can be established by using the Right Hand Grasp Rule, as shown on the
next page. By pointing your thumb in the direction of the current flow, and gripping the
conductor, your fingers will indicate the direction of the magnetic field.

The magnetic field produced by a straight piece of wire is relatively weak and is of no practical
use. It has direction, but no north or south poles and, unless the current is extremely high, the
resulting magnetic field will have little useful strength. Its magnetic characteristics can however
be greatly improved by shaping the wire into a loop.

This will cause:-

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¾ The lines of magnetic flux to move closer together.

¾ The majority of the lines of magnetic flux to be concentrated at the centre of the
loop.

¾ North and south poles to be created at the ends of it, and for it to assume the
magnetic characteristics of a permanent magnet. Lines of magnetic flux will
emerge from the north pole and return via the south pole.

An Electromagnet
A strong magnetic field can be produced if the wire is formed into a coil, which is commonly
known as a Solenoid, as shown on the next page. Also the larger the magnitude of the current,
or the higher the number of turns the greater the magnetic strength.

The strength of the magnetic flux around a coil can be further increased if a soft iron bar is
inserted into the coil. This has the affect of concentrating the lines of magnetic flux, and the
polarity of the coil can be determined if the direction of the current through the coil is known, by
using the ‘Right Hand Grasp Rule’. Your thumb will point in the direction of the north pole if the
fingers of your right hand are wrapped around the coil in the direction of the current flow.

Magnetic Moments
The magnetic moment of a magnet is the tendency for it to turn or be turned by another magnet.
It is a requirement of any aeroplane compass design that the strength of the moment is such
that the magnetic detection system will rapidly respond to the directive force of the magnetic
field.

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The diagram above shows a pivoted magnet of pole strength "S" and length of magnetic axis
"L", which is positioned at right angles to a uniform magnetic field "H". In this situation the field
will be distorted and will "pass through" the magnet. The distortion will be resisted and the field
will try to align the magnet with the magnetic field. The forces being applied to the magnet will
act in opposite directions, and the resulting magnetic moments (m = S x L) will produce a
couple, which will swing the magnet into line with the magnetic field.

The greater the pole strength and the longer the magnetic moment, the greater will be the
magnet's tendency to align itself quickly with the applied field.

Period of Oscillation of a Suspended Magnet

If a suspended magnet is deflected from its position of rest in a magnetic field, the magnet will
immediately be subject to a couple urging the magnet to resume its original position. When the
deflecting influence is removed, the magnet will swing back, and if undamped the system will
continue to oscillate about its equilibrium position before coming to rest. The time taken for the
magnet to swing from one extremity of oscillation to the other and back again is known as the
‘period’ of the magnet. As the magnet approaches its aligned position, the amplitude of the
oscillations will gradually decrease, but the period will remain the same, and will not be altered
by simply adjusting the amplitude. The period of a magnet depends upon its shape and size or
mass (the factors which effect its moment of inertia), the magnetic moment, and the strength of
the field in which it is oscillating. The period of oscillation will be increased if the magnet's mass
is increased, and will become shorter if its field strength is increased.

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Chapter 2.2
Terrestrial Magnetism

Introduction
The Earth is itself a huge magnet and is surrounded by a weak magnetic field that culminates in
two internal magnetic poles situated near the North and South geographic poles. The earth's
magnetic field is similar to that produced at the surface of a short but very powerful bar magnet,
and is why the magnetic poles cover relatively large geographic areas, as the lines of magnetic
force spread out. This is also why the lines of force are horizontal in the vicinity of the equator.
The precise origin of the field is however unknown, but for simplicity the analogy of the bar
magnet at the earth's centre is useful in visualising the general form of the earth's magnetic
field.

The poles are joined together by an imaginary line called a ‘Magnetic Meridian’. If a freely
suspended magnetised needle is positioned at various locations within the earth's magnetic
field, it will line itself up with its red pole pointing towards the Earth’s magnetic North pole,
ie. with respect to magnetic North.

The earth's magnetic field however differs from that of an ordinary magnet in many respects
because the magnetic poles themselves continually alter their position by a small amount. The
magnetic field at any point on the earth's surface is also not constant, because it is subject to
both periodic and irregular changes.

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Magnetic Dip
A freely suspended magnetic needle will settle in a definite direction at any point on the earth's
surface, by aligning itself with the magnetic meridian at that point. It will thus not lie parallel to
the earth's surface at all points, because the earth's lines of magnetic flux (force) are
themselves not horizontal.

The lines of force initially emerge vertically from the South magnetic pole, and then bend over to
become parallel with the earth's surface, before descending vertically at the North magnetic
pole. Thus if a magnetic needle is transported along a meridian from North to South, it will
initially have its red end pointing down towards the earth. Near the magnetic equator the needle
will be horizontal; and at the southern end of its travel the blue end will point towards the earth.

The angle that the lines of force make with the earth's surface at any given place is called the
‘Angle of Dip’’ and varies from 0° at the magnetic equator, to virtually 90° at the magnetic poles.
Lines drawn on the earth’s surface joining places of equal dip are known as ‘Isoclinals’ (BB and
CC), whilst a line joining places having zero dip is known as an ‘Aclinic’ line (AA). The Aclinic
Line is also the magnetic equator, which is close to the geographical equator, but is not the
same line.

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Isoclinals can also be plotted on charts of the world to show how the magnetic dip varies
worldwide, as shown on the next page.

ACLINIC
LINE

Dip is conventionally positive when the red end of a freely suspended magnetic needle is below
the horizontal and negative when the blue end dips below the horizontal, as shown below.

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Earth's Total Magnetic Force

If a freely suspended magnetic needle comes to rest in the earth's field, it will do so under the
influence of the total force of the earth's magnetic field at that point. The value of this total force
at any given place is not easy to measure, but needs to be known. It is usual, therefore, to
resolve the total force into a horizontal component termed "H" and a vertical component termed
"Z". If the value of angle of dip (θ) for the particular location is then known, the total force (T)
can be readily calculated. Knowledge of the horizontal component "H" and vertical component
"Z" is of considerable practical value, as both are responsible for inducing magnetism into the
various ferrous metal parts of the aeroplane (both hard and soft iron) which lie in their
respective planes. Both components may, therefore, be responsible for providing a deflecting or
deviating force around the aeroplane's compass position, a force whose value must be
determined and calibrated for, if the compass is to provide a worthwhile heading reference. The
relationship between dip, horizontal, vertical and total force is shown below.

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H H

This figure shows that "H" is a maximum value at the magnetic equator and decreases in value
towards the poles. Conversely, "Z" is zero at the magnetic equator and, like the value of dip,
increases towards the poles.

Examples

1. If the value of H is 0.22 and the value of Z is 0.44 at a place X, then the angle of dip at this
place will be:-

Z = 0.44 = 2
Tan Dip = H 0.22

Angle of Dip = Tan-1 2 = 63°26’

2. If the Angle of Dip = 60° and Z = 0.27, then the values of H and T will be:-

H = TanZ60° = 1.7321
0.27 = 0.1559

T= 0.27 = 0.27 = 0.3118

Sin 60° 0.8660

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Magnetic Variation
In a similar manner, as meridians and parallels are constructed with reference to the geographic
poles, magnetic meridians and parallels may also be plotted with reference to the magnetic
poles. If a map is prepared showing both true and magnetic meridians, the meridians will
intersect each other at angles varying from 0° to 180° at different points on the earth's surface,
as shown below.

‘Isogonals’ or ‘Isogonic’ lines are used on charts to show the amount of variation, and to join all
places on the earth’s surface having the same angle of variation, whereas a line on the chart
where the variation is nil is called the ‘Agonic’ line. When the direction of the magnetic meridian
inclines to the left of the true meridian, the variation is said to be ‘west’, whilst inclination to the
right of the true meridian is said to be variation ‘east’. Variation can change from 0° in areas
where the magnetic meridians run parallel to the true meridian, to 180° in places located
between the true and magnetic north poles. The angle between the true and magnetic
meridians at any place when looking north is known as magnetic variation, an example of which
is shown below.

At some locations on earth, where the ferrous nature of the rock deposits disturbs the earth's
magnetic field, abnormal magnetic anomalies occur, which may cause large changes in the
value of variation over very short distances.

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While variation differs all over the world, it does not maintain a constant value in any one place,
and the following changes, which are not constant in themselves, may occur:-

¾ Secular changes which occur over long periods, due to the changing position of the
magnetic poles relative to the true poles.

¾ Annual change, which is a small seasonal fluctuation super-imposed on a secular

change.

¾ Diurnal (daily) changes, which appear to be caused by electrical currents flowing in

the atmosphere as a result of solar heating.

¾ Magnetic storms associated with sunspot activity. These may last from a few hours
to several days, with the intensity varying from very small to very great. The effect
on aeroplane compasses will thus vary with intensity, but both variation and local
values of "H" will be modified whilst the "storm" lasts.

These factors will also have a similar affect on the Angle of Dip.

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Intentionally Left Blank

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Chapter 2.3
Aeroplane Magnetism

Introduction
According to the JARs it is a mandatory requirement that a direct reading compass be fitted as
standard on all aeroplanes, and must be fitted where the flight crew can easily read it. The flight
deck is however surrounded by magnetic material and electrical circuits, which will influence the
earth’s magnetic field, and cause the compass needle to be deflected away from the local
magnetic meridian. This is known as compass deviation and can result in either deviation east,
or deviation west of magnetic north by an amount, which depends on the aeroplane’s heading
and latitude. Fortunately any deviation can be analysed and resolved into components acting
along the aeroplane’s major axes, and action taken to minimise these effects. The causes of
the deviation can be analysed and corrected for, by carrying out a compass swing, although
some residual deviation will remain, and this has to be recorded on a deviation card.

Types of Aeroplane Magnetism

Aeroplane magnetism can be classified in a similar manner to that of hard iron and soft iron, ie.
how readily the materials become magnetised.

Hard Iron Magnetism. This form of magnetism is of a permanent nature, and is due to
the presence of iron or steel parts used in the aeroplane structure, in power plants and
other equipments. The earth’s magnetic field will influence the molecular structure of
the ferrous parts of the aeroplane during its construction when it lies on one heading for
a long period of time. Hammering and working of the materials will also play a major
part in the magnetism of the aeroplane components, whilst they are lying in the
magnetic field.

Soft Iron Magnetism. This form of magnetism is of a temporary nature and is caused
by the magnetically soft metallic parts becoming magnetised due to induction by the
earth’s magnetic field. The effect of this type of magnetism is dependent on the
aeroplane’s heading and the local Angle of Dip (magnetic latitude), and its geographical
location

Letters normally indicate the components of aeroplane magnetism, which cause deviation;
capital letters indicate permanent hard iron magnetism, whilst small letters indicate induced soft
iron magnetism. Positive deviations (those deflecting the compass needle to the right) are
termed easterly, whilst negative deviations (deflection of the compass needle to the left) are
termed westerly.

Components of Hard Iron Magnetism

The overall effect of hard iron magnetism on the aeroplane compass can be assimilated to bar
magnets lying longitudinally, laterally and vertically about the compass position.

To analyse the effect of hard iron, the imaginary bar magnets are annotated as ‘component P’,
‘component Q’ and ‘component R’ respectively. Assume that the magnetic strength of these
components will remain constant regardless of the aeroplane heading or latitude, but may vary
with time due to a weakening of the magnetism in the aeroplane. By convention if the blue poles
of the imaginary magnets are forward of, to starboard of’, or beneath the compass position, the
components will be positive, as shown on the next page. If the blue poles however act in the
opposite direction, they will alternatively be negative by convention.

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AEROPLANE
HEADING

This occurs because in the Northern Hemisphere the vertical component of the earth’s magnetic
field will induce magnetism into the airframe, which will cause the blue poles to be attracted to
the top of the aeroplane, and the red poles to be repelled to the bottom.

When the effect of this vertical magnetism is analysed in the horizontal plane through the
compass needle, it reveals a net blue pole to the front of the aeroplane, and net red pole to the
rear. Similar effects occur in the lateral vertical planes, with the overall polarity depending on
the actual aeroplane design. The polarity and strength of the magnetism is however not
affected by the aeroplane heading.

If the aeroplane is heading north, the imaginary magnet due to component P will together with
the compass needle, be in alignment with the aeroplanes fore and aft axis, and the earth’s
directive force (H). P will thus be added to, or subtracted from H, but will not cause any
deviation. If the aeroplane is turned through 360°, then as the turn is commenced (ignoring
compass pivot friction and liquid swirl etc) the magnet system will remain attracted to the earth’s
H component. Component P will however continue to act along the aeroplane’s fore and aft
axis, and will cause the compass needle to align itself in the resultant position between the
directive force (H), and the deflecting force (P), which will cause the needle to point a number of
degrees east or west of north, depending on the polarity of P, as shown in the diagram on the
next page.

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The amount of deviation will increase during the turn, reaching a maximum value when
travelling east and west, and zero when travelling north and south. Deviation resulting from a
positive P can be represented by a sine curve, as shown below.

W- E+

This shows that the deviation due to P is proportional to the sine of the aeroplane’s Compass
Heading, ie:-
Deviation = P sin Heading (C)

Component Q also produces a similar effect, but since it acts along the aeroplane’s lateral axis
(wing tip to wing tip), the deviation resulting from Q will be a maximum value when travelling
north and south, and zero when travelling east and west, ie. when the component is aligned
with the directive force (H). Deviations resulting from a negative Q (blue pole to the left of the
compass position) can be represented as a cosine curve, as shown below.

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This curve shows that the deviation due to Q is proportional to the cosine of the aeroplane’s
Compass Heading, ie:-
Deviation = Q cos Heading (C)

Component R acts in the vertical plane and thus has no effect on the compass system when the
aeroplane is in level flight. If the aeroplane however flies with its lateral or longitudinal axes
away from the horizontal, the component R will be displaced from its vertical position, and the
resulting horizontal vector of this component will have an effect on the compass system. This is
demonstrated in the following diagram, which illustrates how an element of R would affect the
components P and Q.

Notably a similar situation will exist with a tail wheel aeroplane when it is on the ground. The
value of R may however vary when the aeroplane is climbing and descending, but because the
angles involved are normally small, any deviation resulting from component R will also be
correspondingly small. Additionally the turning and acceleration errors associated with a direct
reading compass during turns will make the errors due to R of no practical significance. The
effect of component R will also be negligible in remote indicating compasses, since the turning
errors are virtually eliminated in this instrument, as a result of its associated electronic circuitry.

Components of Soft Iron Magnetism

Soft iron magnetism, which is present at the compass position, may be considered as
originating from soft iron rods placed adjacent to the compass in which magnetism is induced by
the earth’s magnetic field. This field has two components, H and Z, but in order to analyse the
effect of soft iron the H component must be further split into two horizontal components, X and
Y. When these components are put together with the Z component of the earth’s magnetic field
they relate directly to the three principle axes of the aeroplane. The diagram on the next page
shows how the polarities and strengths of X and Y alter with a change in aeroplane heading, as
the aeroplane turns relative to the direction of component H.

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Component Z acts vertically through the compass, and therefore does not affect the directional
properties of the magnet system. If the aeroplane is however moved to a new geographic
location, then because of the change in the earth’s magnetic field strength and direction, all
three components of soft iron magnetism will change. The sign of Z will however only change if
the aeroplane changes the magnetic hemisphere in which it is operating.

The soft iron components, which mainly affect the compass are indicated conventionally by the
small letters, ‘c’ and ‘f’, and are related to the earth’s field components X, Y and Z. Out of the
soft iron components, cZ and fZ are the most important, since their polarity remains the same,
regardless of the aeroplane heading. They also act in the same manner as the hard iron
components P and Q respectively. Pairs of vertical soft iron rods (VSI), which are positioned
respectively fore and aft, and laterally about the compass position represent the components,
cZ and fZ respectively. In the northern hemisphere (magnetic) the lower pole of each rod would
be induced with ‘red’ magnetism.

Determination of Deviation Coefficients

In order to minimise the effects of hard and soft iron magnetism on the aeroplane’s compass, it
is necessary to firstly determine the deviations caused by the components of aeroplane
magnetism on various headings. The value of any deviations can then be analysed into the
‘coefficients of deviation’. Five coefficients exist, namely A, B, C, D and E, of which D and E are
purely due to soft iron, and will not be covered in this manual. The remaining coefficients are
important to aeroplane magnetism and are as follows:-

Coefficient A. This is usually constant on all headings and is caused by the

misalignment of the aeroplane compass. This coefficient is calculated by finding the
average of the algebraic sum of the deviations resulting from a number of equally
spaced compass headings. Readings are typically taken on the four cardinal and four
quadrantal headings, thus:-

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Coefficient A = Deviation on N + NE + E + SE + S + SW + W + NW
8

Coefficient B. This is the result of the resultant deviation caused by the presence of
hard iron P and soft iron cZ, with the deviation being a maximum, when heading east or
west. This coefficient is calculated by dividing the algebraic difference between the
deviations on compass heading east and west by two, thus:-

Coefficient B = Deviation on east - Deviation on west

2
For any given heading, coefficient B may also be expressed as:-

Deviation = B x sin heading

Coefficient C. This is the result of deviations caused by hard iron Q and soft iron fZ,
with the deviation being a maximum when heading north and south. This coefficient is
calculated by dividing the algebraic difference between the deviations on compass
heading north and south by two, thus:-

Coefficient C = Deviation on north - Deviation on south

2

For any given heading coefficient C may also be expressed as:-

Deviation = C x cos heading

The total deviation on an uncorrected compass for any given aeroplane compass
heading
may be expressed as:-

Total deviation = A + B sin heading + C cos heading

Minimum Deviation
If Coefficient A is not present, minimum deviation will occur on the heading where the value of
B sin heading + C cos heading is minimum, thus:-

Tan heading = C
B

This heading is at right angles to that for maximum deviation, but if Coefficient A is present the
minimum deviation parameters have to be determined by a compass swing.

Maximum Deviation
If the deviation due to Coefficient A is constant on all headings, maximum deviation will occur
when the value of B sin heading + C cos heading is a maximum, thus:-

B
Tan heading = C
The value of the tangent derived for maximum deviation corresponds to two reciprocal headings
in opposing quadrants, thus to determine the correct heading it is necessary to construct a
swing circle, as shown on the next page.

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The above diagram illustrates the condition where coefficients A = +1, B = +3 and C = +2.
Maximum deviation thus occurs in the north-east quadrant. The actual value is:-

Tan B = 3 = 1.5 which corresponds to a maximum deviation heading of 056°

C 2

Substituting these values in the formula for maximum deviation gives:-

A + B sin heading + C cos heading = 1 + 3 sin 56° + 2 cos 56° = + 4.6°

Joint Airworthiness Requirements (JAR) Limits

JAR (25) for large aeroplanes requires that a compass residual deviation card (placard),
showing the calibration of the magnetic compass (Direction Indicator) in level flight with the
engines running, must be installed on or near the instrument. The placard must show each
calibration reading in terms of magnetic heading of the aeroplane in not greater than 45° steps.
Furthermore the compass, after compensation, may not have a deviation in normal level flight of
greater than 10° on any heading.

The following Joint Airworthiness Requirements (JAR’s) govern the positioning of the compass:-

¾ The distance between a compass and any other item of equipment containing
magnetic material shall be such that the piece of equipment does not result in a
deviation of greater than 1°, nor shall the combined effect of any such equipment
exceed 2°. The same ruling also applies to installed electrical equipment and any
associated wiring when the systems are powered up.

¾ Any movement of the flight controls or undercarriage should not result in a change
in deviation of greater than 1°.

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¾ The effect of the aeroplanes permanent and induced magnetism, as given by

coefficients B and C, together with any associated soft iron components shall not
exceed:-

Direct Reading Remote Reading

Coefficient Compass Compass
(degrees) (degrees)

B 15 5

C 15 5

Note: a). After correction the greatest deviation on any heading shall be 3° for direct reading
compasses and 1° for remote indicating compasses.

b). Emergency standby compasses and non-mandatory compasses need not fully
comply with JAR, but evidence of satisfactory installation is required.

Compass Swinging
This is a special calibration procedure, which is carried out in order to determine the amount by
which the compass readings are affected by hard and soft iron magnetism. This process
enables the deviations to be determined, the respective coefficients to be calculated, and the
deviations to be compensated for. Compass swinging should be carried out:-

¾ On acceptance of the new aeroplane from the manufacturer.

¾ When a new compass is fitted.

¾ Periodically, normally every three months.

¾ Following a major inspection.

¾ Following a change of magnetic material in the aeroplane.

¾ If the aeroplane is moved permanently or semi-permanently to another airfield

involving a large change of magnetic latitude.

¾ Following a lightning strike or prolonged flying in heavy static.

¾ After standing on one heading for more than four weeks.

¾ When carrying ferrous (magnetic) freight.

¾ Whenever specified in the maintenance schedule.

¾ For the issue of a Certificate of Airworthiness (C of A).

¾ At any time when the compass or residual deviation recorded on the compass card
are in doubt.

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The Compass Swing Procedure

There are a number of methods by which a compass swing can be carried out, but the most
common method is to use an engineer with a landing or datum compass mounted on a tripod.
The specialised compass is positioned either in front, or behind the aeroplane, and is aligned
with the aeroplanes fore and aft axis. This process is normally carried out by an experienced
compass adjuster, and is split into two phases; the correcting and the check swing, as follows:-

¾ Ensure that the compass is serviceable.

¾ Ensure that all equipment not normally carried in flight is removed from the
aeroplane.

¾ Ensure that all equipment normally carried in flight is correctly stowed.

¾ Take the aeroplane to a swing site (at least 50m from other aeroplane and 100m
from a hangar).

¾ Ensure that the flying controls are in their normal flying position, the engines are
running, and the radio and electrical equipments are switched on.

¾ Position the aeroplane on a heading of south (M) and note the deviation, ie. the
difference between the datum compass and the aeroplane compass readings.

¾ Position the aeroplane on a heading of west (M) and note the deviation.

¾ Position the aeroplane on a heading of north (M) and note the deviation. Calculate
coefficient C and apply it directly to the compass reading.

¾ Insert the compass corrector key in the micro adjuster box, and turn the key until
the compass needle shows the corrected reading. Remove the key.

¾ Position the aeroplane on a heading of east (M) and note the deviation. Calculate
coefficient B and correct for B in the same manner as for coefficient C.

¾ The correcting swing is now complete.

¾ Carry out a check swing on eight headings, starting on south-east (M), and note the
deviation on each heading.

¾ Calculate coefficient A on completion of the check swing and apply to the compass
reading. Loosen the compass, or, for remote indicating instruments, the detector
head retaining screws, and rotate the device until the compass needle indicates the
correct heading. Re-tighten the retaining screws.

¾ Having applied coefficient A algebraically to all deviations found during the check
swing, plot the (remaining) deviations, and make out a compass deviation card for
placing in the aeroplane.

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An Example of a Compass Swing

The readings taken during a typical compass swing are as follows:-

1. The Correcting Swing.

Datum Compass Aeroplane Deviation

Heading (M) Compass (degrees)
(degrees) Heading (C)
(degrees)

S 182 180 +2

W 274 270 +4

N 000 354 +6

+ 6 - ( +2) Make the compass

Coefficient C = = +2 read 356°
2

E 090 090 0

0 - ( +4) Make the compass

Coefficient B = = -2 read 088°
2

2. The Check Swing.

Datum Compass (M) Aeroplane Compass(C) Deviation Residual Deviation

(degrees) (degrees) (degrees) Following ‘A’

136 131 +5 +2

183 181 +2 -1

225 221 +4 +1

270 268 +2 -1

313 308 +5 +2

000 358 +2 -1

047 044 +3 0

092 090 +2 -1

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Coefficient A = 25 = +3
8
The resultant readings enable a deviation card to be produced, showing residual deviations
against magnetic headings, which are placed in the aeroplane adjacent to the compass
position.

Deviation Compensation Devices

Following the compass swing procedure, the resultant coefficients C, B and A, are used to
correct or offset the compass needle by an amount in degrees equivalent to the existing
deviation using one of the following methods:-

Mechanical Compensation. The majority of these devices consist of two pairs of

magnets, which are fitted in a bevel gear assembly made of a non-magnetic material,
and are mounted above each other, as shown below. This device is known as a ‘micro-
adjuster’, and it ensures that when the magnets are in their neutral position one pair is
parallel to the aeroplanes fore and aft axis to compensate for any Coefficient C
corrections, whilst the other pair lies athwartships to compensate for any Coefficient B
corrections. Using the compass correction key enables a small pinion to be turned,
which in turn will rotate one pair of bevel gears.

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Each pair of magnets will be displaced from their neutral position, as shown in the diagram
above, and will deflect the compass needle to correct for Coefficient B or C, depending on
which pair of magnets is used.

Electrical Compensation. This type is used in the remote indicating compass and
uses two variable potentiometers, which are connected to the coils of the flux detector
unit. The potentiometers correspond to the Coefficient B and C magnets of a
mechanical compensator and when moved with respect to calibrated dials, insert small
DC signals into the flux detector coils. The resulting magnetic fields produced by these
signals are sufficient to oppose those causing deviations and correspondingly modify
the output from the detector head via the synchronous transmission link. This in turn
drives the gyro and hence the heading indicator to display the corrected readings.

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Chapter 2.4
Direct Reading Magnetic Compass

Introduction
The Direct Reading Magnetic Compass (DRC) is based on a simple magnetic needle, and
points towards the northern end of the earth's magnetic field. It is also installed in an instrument
of dimensions and weight that makes it suitable for use in aeroplanes. Under JAR 25 it is a
mandatory requirement that all modern civil transport aeroplanes carry a direct reading non-
stabilised magnetic compass as a standby direction indicator. The most commonly found direct
reading compass is the “E” type, which is illustrated below.

Properties of a Direct Reading Compass

For a direct reading compass to function efficiently, the magnetic element must possess the
following properties:-

Horizontality. This ensures that the magnet system remains as near horizontal as
possible, thereby sensing only the horizontal or directive component of the earth's
magnetic field. This is achieved by making the magnet system pendulous, by mounting
the magnet, below the needle pivot, as shown in the diagram above. The magnet
system when freely suspended in the earth’s magnetic field will tend to align itself with
the direction of that field, ie. align itself in the direction of the total field (T), where T is
the resultant of the earth’s horizontal (H), and vertical (Z) fields. If the system is tilted
the C of G will move out from beneath the pivot, and will introduce a righting force upon
the magnet system, which will tend to oppose and reduce the overall ‘Z’ component.
The compass will thus take up a position along the resultant of the two forces, ‘H’ and
the reduced effect of ‘Z’, thus minimising the effect of dip. In temperate latitudes the
final inclination of the needle will be approximately 2° to 3° to the horizontal, but this
inclination will increase when flying nearer the poles, such that, by about 70° north or
south, the compass is virtually useless. The displacement of the C of G is purely a
function of the system's pendulosity, and is not a mechanical adjustment, so it will work
in either hemisphere, without further adjustment.

Sensitivity. This ensures that the DRC is capable of operating effectively down to low
‘H’ values, and is achieved by increasing the pole strengths of the magnet being used,
so that it remains firmly aligned with the local magnetic meridian. Sensitivity is also

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aided by keeping pivot friction to a minimum by using an iridium-tipped pivot, which is

free to move in a sapphire jewelled cup. The compass bowl is additionally filled with a
liquid, which reduces the overall effective weight of the magnet system, and also helps
to lubricate the pivot.

Aperiodicity. This ensures that the oscillation of the sensitive element about a new
heading, following a turn, is minimised, ie. a ‘Deadbeat Return’ characteristic. If a
suspended magnet is deflected from its position of rest and then released, it will tend to
oscillate around the correct direction for some time before stabilising. This is obviously
undesirable, as it could, at worst, lead to the pilot chasing the needle. The compass
needle should thus come to rest with minimal oscillation, which is achieved by:-

1. Filling the bowl with methyl alcohol or a silicon fluid, and fitting damping
filaments to the magnet system.

2. Keeping the lever arm of the magnet system as short as possible, but
keeping its strength high. This has the effect of maximising its directional
force, whilst reducing its moment of inertia.

3. Using the fluid to reduce the apparent weight of the system.

4. Concentrating the weight as close to the pivot point as possible, to further

reduce the turning moment.

"E" Type Compass

The majority of standby compasses in use today are of the card type, an example of which is
shown below.

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The compass consists of a single circular cobalt steel magnet, to which a light metal compass
card is attached, and is mounted so that it rotates as close as possible to the inner face of the
bowl, thus keeping parallax errors to a minimum during reading. The card is graduated with
white markings on a black background, every 10°, with any intermediate indications being
estimated by interpolation. Figures are also shown every 30° and the cardinal points are
marked with appropriate letters, N, S, E and W. A white, vertical lubber line is engraved on the
inner face of the bowl, against which the actual heading is observed.

The system is suspended on an iridium-tipped pivot that revolves in a sapphire jewelled cup,
which is in turn secured to a central stem, and is firmly attached by a bracket to the base of the
bowl. This gives the magnet system freedom of movement of + 20° from the horizontal, and
360° in azimuth.

The bowl is moulded from plastic, and is painted on the outside with black enamel, except for a
small area at the front through which the vertical card can be seen. This part of the bowl is also
moulded so that it has a magnifying effect on the compass card.

The bowl is filled with a silicone fluid, which has no detrimental affect on the plastic bowl, and
also because its properties are not significantly affected by its temperature/viscosity. The liquid
used in the compass bowl is also transparent and has a high resistance to corrosion. It must
also not discolour during its use. Furthermore a bellows type expansion chamber is located at
the rear of the bowl and compensates for changes in liquid volume, due to any variation in
temperature, which ensures that the liquid neither bursts a seal, or contracts, leaving vacuum
bubbles in the fluid.

One disadvantage of using a liquid in the compass bowl however is that, in a prolonged turn, it
will tend to turn with the aeroplane, thus taking the magnet system with it, and affecting the
compass readings. This is known as ‘Liquid Swirl’, which is minimised by providing a good
clearance between the damping wires, and the sides of the compass bowl. Liquid swirl also
tends to delay the immediate settling of the system when a new compass heading is selected.

The effects of deviation co-efficient B and C are compensated for by permanent magnet
corrector assemblies, which are secured to the compass mounting plate.

Pre-flight Checks
Prior to flight the flight crew should carry out the following checks:-

¾ Check the security of the compass.

¾ Carry out a visual check for signs of any external damage.

¾ Check that the liquid is free from bubbles, discoloration and sediment.

¾ Check that the compass illumination system is serviceable.

¾ Test for pivot friction by deflecting the magnet system through 10-15 ° each way,
and note the readings on return, which should be within 2° of each other.

Principle of a Pendulum
Consider a plain pendulum that is freely suspended in the aeroplane fuselage. If the aeroplane
maintains a constant direction and speed, the pendulum will remain at rest, but if the aeroplane
turns, accelerates or decelerates the pendulum will be displaced from its true vertical position.
This will occur because the inertia of the pendulum will cause the centre of gravity to lag behind
the pendulum pivot, thus deflecting it away from its normal vertical position, directly beneath its
point of suspension.

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The magnet system (in the compass) is pendulous, so any acceleration or deceleration in flight
will similarly result in a displacement of the C of G away from its normal position. This will result
in a torque being established about the vertical axis of the compass, and unless the compass is
on the magnetic equator, where the earth’s field vertical component ‘Z’ is zero, it will be subject
to dip.

Acceleration Errors
The force applied by an aeroplane when accelerating or decelerating on a fixed heading is
applied to the magnet system at the pivot, which is the magnet's only connection with the rest
of the instrument. The reaction to this force will be equal and opposite, and will act through the
C of G of the magnet system, which is below and offset from the pivot (except at the magnetic
equator), as shown below. The two forces will thus constitute a couple which, dependent on the
aeroplane’s heading, will cause the magnet system to alter its angle of dip, ie. attempt to restore
the magnet to its horizontal position, or to rotate it in azimuth.

The diagram above shows how the forces affect a magnet system when an aeroplane is
accelerating on a northerly heading. The resulting acceleration force is similarly applied to the
magnet system at the pivot, whilst an equal and opposite reaction ‘R’ will act through the C of G,
which is below, but offset from the pivot. The resultant couple will cause the northern end of the
magnet system to dip further, thus increasing the angle of dip without any rotation in azimuth.
This will occur because the pivot ‘P’, and C of G, are both in the plane of the local magnetic
meridian. Conversely, if the aeroplane decelerates when flying in a northerly direction, the
resultant couple will tilt the magnet system down at its southern end. The opposite will be
observed when accelerating/decelerating in a northerly direction along the magnetic meridian in
the Southern Hemisphere. If the aeroplane is flying in either hemisphere, any changes in speed
on headings other than northerly or southerly, will also result in azimuth rotation of the magnet
system, and will produce errors in the heading indication, as shown below.

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Acceleration errors are also caused by the vertical component of the earth’s magnetic field,
which occurs because of the magnet systems pendulous mounting, and causes the compass
card to tilt during changes of speed. This deflection will cause a further error, which will be most
apparent on easterly and westerly headings. When an aeroplane is operated in the Northern
Hemisphere and accelerates on either of these headings, the resulting error will cause the
magnet system to rotate, and the compass to indicate a turn to the north. Conversely if an
aeroplane decelerates on either of these headings, the resulting error will cause the magnet
system to rotate, and the compass to indicate a turn to the south.

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CONSIDER AN AEROPLANE IN
THE SOUTHERN HEMISPHERE
INCREASING SPEED WHILST
HEADING WEST

These indications will however be reversed in the Southern Hemisphere. If the aeroplane
decelerates when flying in a westerly direction, the action and reaction of ‘P’ and ‘R’
respectively, will have the opposite effect, and will cause the assembly to turn in the opposite
direction, with all of the forces again turning in the same direction.

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The errors due to acceleration and deceleration are summarised in the following table:-

Heading Speed Needle Turns Visual effects

East Increase Clockwise Apparent turn to north
West Increase Anti-clockwise Apparent turn to north
East Decrease Anti-clockwise Apparent turn to south

West Decrease Clockwise Apparent turn to south

Note:
1. In the Southern Hemisphere, the errors are in the opposite sense.

2. Similar errors can occur in turbulent flight conditions.

3. No errors occur at the magnetic equator, as the value of “Z” is zero and hence the
pivot point and C of G will be co-incident with each other.

Turning Errors
During a turn, the compass pivot is carried along the same curved path as the aeroplane. The
centre of gravity (of the magnet system), being offset from the pivot, which is used to counter
the effect of ‘Z’, is thus subject to centrifugal acceleration. Furthermore, in a correctly banked
turn the magnet system will tend to maintain a position parallel to the athwartships (wingtip to
wingtip) axis of the aeroplane, and will thus be tilted in relation to the earth's magnetic field. This
will place the pivot and C of G out of alignment with the local magnetic meridian. The magnet
system will thus be subject to a component of ‘Z’, and this will cause it, when turning through
North in the Northern Hemisphere, to rotate in the same direction as the turn. This will further
increase the turning error, and will cause the compass to under-indicate, as shown below.

The magnitude and direction of the turning error is thus dependent on the aeroplane’s heading,
its angle of bank (the degree of tilt of the magnet system), and the local value of ‘Z’ (dip). The
turning error will be a maximum value on northerly/southerly headings, and will be particularly
significant within 35° of these headings.

If an aeroplane turns east, as soon as the turn is commenced, the magnet system’s C of G will
be subject to a centrifugal acceleration, and will cause the system to rotate in the same direction
as the turn. This will in turn tilt the magnet system, and will allow the earth's vertical component
‘Z’ to exert a pull on the northern end, which will cause further rotation of the system. The same
effect will occur if the heading change is from north to west in the Northern Hemisphere.

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The speed of rotation of the system is a function of the aeroplane's bank angle and rate of turn.
As a result of these factors, the following indications may be registered by the compass:-

¾ A turn in the correct sense, but smaller than that carried out when the magnet
system turns at a slower rate than the aeroplane.

¾ No turn when the magnet system turns at the same rate as the aeroplane.

¾ A turn in the opposite sense because the magnet system turns at a faster rate than
the aeroplane.

When turning from a southerly heading in the Northern Hemisphere onto an easterly or westerly
heading, the rotation of the system and indications registered by the compass will be the same
as when turning from north, except that the compass will over-indicate the turn.

The effects of turning through North and South in the Northern Hemisphere are summarised in
the following table:-

Turn Needle Visual Liquid Corrective

Direction Movement Effect Swirl Action
Through Same as Under Adds to Error Turn less
North aeroplane Indication than needle
shows
Through Opposite to Over Reduces Turn more
South aeroplane Indication Error than needle
shows

The liquid in the bowl not only provides damping, but it also tends to turn with, and in the same
direction as the turn. This is referred to as ‘Liquid Swirl’, and its motion will either add to, or
subtract from, the overall needle error, which is dependent on its relative movement.

In the Southern Hemisphere the south magnetic pole will dominate and, in counter-acting its
downward pull on the compass magnet system, the C of G will move to the northern side of the
pivot. The errors will thus be in the opposite sense. If an aeroplane turns from a northerly
heading onto a easterly heading, the centrifugal acceleration acting on the C of G will cause the
needle to rotate more rapidly in the opposite direction to the turn, thus indicating a turn in the
correct sense but of greater magnitude than that actually carried out. The turn will thus be over-
indicated. Turning from a southerly heading onto an easterly or westerly heading in the
Southern Hemisphere will, because of its C of G, which is still north of the compass pivot, result
in the same effect as turning through north in the Northern Hemisphere.

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Note: 1. In the Southern Hemisphere, the errors are opposite to those occurring in the
Northern Hemisphere.

2. The Northerly turning error is greater than southerly, as liquid swirl is additive to
the compass magnet system movement.

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Intentionally Left Blank

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Chapter 2.5
Remote Indicating Compass

Introduction
A remote indicating compass (RIC) or gyro magnetic compass (GMC) combines the use of a
gyroscope with the earth’s lines of magnetic flux to give the best of both a Direction Indicator
(DI) and a Direct Reading Compass (DRC). This type of compass essentially consists of a
magnetic compass whose indications are stabilised via a gyroscopic element, thus minimising
the affects of turning and acceleration errors. The magnetic part of the system simply senses
the earth’s meridian, and does not align itself with it, although it is subject to aeroplane
accelerations. The gyroscopic element is alternatively not affected by changing magnetic fields
or normal acceleration forces, but is subject to precessional forces caused by friction etc. This
type of compass thus integrates the heading indication provided by the magnetic compass with
the directional properties of the gyroscope to minimise the overall errors, in order to provide
accurate indications of an aeroplanes heading.

RIC Architecture
The main components of a typical RIC system are shown in the following diagram.

Principle of a Flux Detector Element

Unlike the detector element of the simple magnetic compass, the element used in a remote-
indicating compass is fixed-in-azimuth, and senses the effect of the earth's magnetic field as an
electro-magnetically induced voltage, and operates as follows.

If a highly permeable magnetic bar or coil is exposed to the earth's field, it will acquire a
magnetic flux, which is solely dependent on the magnetic latitude at which the system is
operating. The amount of flux induced will thus be determined by the strength of the earth's
horizontal ‘H’ component, and the direction of the permeable element relative to the direction of
this component.

The diagram on the next page shows the amount of flux, which would be induced in a single coil
when it is placed at different orientations to the earth’s magnetic (H) field. If the coil is placed
with its longitudinal axis parallel to the H field, the maximum magnetic flux will pass through the
coil. If the coil is alternatively rotated through 90°, so that it is at right angles to the field, it will
produce zero magnetic flux, and if the coil was rotated through a further 90°, it will re-align itself
with the H field, but this time in the reverse direction. In this position it will again produce
maximum flux, but will be in the opposite algebraic sense. The coil will thus show a cosine
relationship (zero flux at 90° and maximum flux at 0°) between the field direction and the coil
alignment.

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For example if an aeroplane is on a heading of 060° (M), the flux intensity will be H Cos 60°.
Similarly, the flux intensity due to the earth's magnetic field on a heading of 120° (M) would
again be H Cos 60°, but the polarity of the flux will have reversed, since Cos 120° is negative.
Conversely on a heading of 300° (M) the induced flux will be of the same sign and value as for a
heading of 060° (M).

A simple system is thus impracticable, because in order to determine the magnetic heading it is
firstly necessary to measure the magnetic flux in the coil, which is difficult to establish, and
secondly, it is subject to an ambiguity in heading, which must also be resolved.

If, according to Faraday, " there is a change of flux linked with a circuit, an EMF will be induced
in that circuit", thus the flux could be easily converted into a measurable electrical current. For
an aeroplane, however, at any given position and direction, if a single coil was used the flux
produced would be of constant value. It is therefore necessary to convert the steady flux into a
changing one, so that a current representing the actual heading would flow. This is achieved in
the flux detector unit via a device called a ‘Flux Valve’, as shown below.

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LAMINATED
COLLECTOR HORNS

A flux valve consists of two identical bars or spokes of highly permeable (easily magnetised and
de-magnetised) material, which is mounted on a common hub. The hub is wound with a coil,
known as the Primary Coil, and is connected in series to an AC single-phase power source of
23.5volts, at a frequency of 400 Hz. A pick-off or secondary coil is wound around both bars and
registers the rate of change of flux in the permeable material. The effect of passing an
alternating current through the primary coil will have the following affect on the amount of flux
being produced in each leg.

The amplitude of the flux produced in each leg is identical, although they are 180° out of phase
with each other, so that the algebraic sum of the fluxes, or the total flux will add up to zero. This
is because at any instant of time the two bars will produce flux of equal and opposite (sign)
intensity. In practice this situation will never occur, since a bar placed horizontally in the earth's
magnetic field will always be subject to an ‘H’ component (unless the aeroplane is near the
north or south magnetic pole). This component will produce a steady flux in both bars, which
when added to their individual fluxes, will bias the system by an amount equal to ‘H’, as shown
below.

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The algebraic sum of the fluxes in each leg will no longer be equal to zero, and the resultant
amplitude will be directly proportional to the aeroplane’s heading.

The changing flux in the bars will result in an EMF or voltage being induced in the pick-up or
secondary coil, which is proportional to the ‘H’ component that acts along the axis of the flux
valve.

The single flux valve, however, has ambiguity over four headings, although two of these have
different algebraic signs, which is resolved in the detector by using three spokes (flux valves),
positioned 120° apart, as shown on the next page. In this arrangement a laminated collector
horn is located at the outer end of each flux valve to concentrate the lines of the earth’s
magnetic force along the parent spoke. This increases the overall sensitivity of the detector
head and increases the magnitude of the induced voltage in the secondary coil

It is however still possible to align the compass with a 180° error, but the instrument will detect
this, and will immediately start to precess the gyro unit to the correct heading.

Flux Detector Unit

The construction of a typical flux detector element is shown below.

1. ELECTRO-MAGNETIC DEVIATION COMPENSATOR 4. PENDULOUS WEIGHT

2. EXCITER COIL 5. PICK-OFF COIL
3. FLUX VALVE 6. UNIVERSAL JOINT

The spokes and coil assemblies are pendulously suspended from a universal ‘Hooke’ joint.
This permits limited freedom in pitch and roll, and ensures that the magnetic element senses

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the maximum value of ‘H’. Unlike a Direct Reading Compass, the magnetic element has no
freedom in azimuth. The unit's case is hermetically sealed and partially filled with fluid to damp
out any element oscillation. The complete unit is normally fitted in the wing or fin tip, so that it is
separated from electronic circuits and aeroplane magnetism in the vicinity of the flight deck,
which may cause compass deviation. The complete unit is secured to the aeroplane by means
of a flange, which contains three screw slots, one of which has calibration marks, and allows
any coefficient A error to be removed. A compensating device is also fitted at the top the
detector unit casing, which enables any coefficient B and C errors to be removed.

Operation of the Remote Indicating Compass System

The directional reference established by the remotely located detector unit has to be
electronically transmitted to another location in the aeroplane via a signal selsyn (synchro
receiver) unit, as shown on the next page. It is used to monitor the action of a gyro, or is simply
displayed on an indicator as a value of aeroplane heading. Any monitoring is then carried out
through a series of selsyn (synchro) units, which in turn will synchronise the movements of the
individual system elements.

If the flux detector is positioned on a steady heading, say 000°, as shown below, then a
maximum voltage signal will be induced in the pick-off coil (secondary winding) 1, whilst coils 2
and 3 will have voltages of half strength and opposing phases induced in them.

These signals will then be fed to the corresponding legs of the stator of a signal selsyn, where
the voltages will be reproduced, and will combine to establish a resultant field across the centre
of the stator. The resultant will thus be in exact alignment with the earth's magnetic field that is
passing through the detector unit. If the rotor of the signal selsyn is positioned at right angles to
the resultant, no voltage will be induced in the windings, and it will be in a "null" position. The
directional gyro being monitored will similarly be aligned with the earth’s resultant field vector,
so the heading indicator will also show 000°.

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If the aeroplane however turns through 90°, the disposition of the flux valve pick-off coils will be
as shown below.

No signal voltage will be induced in coil 1, whilst coils 2 and 3 will have increased voltages, with
the voltage in 3 being opposite in phase to 2. The resultant voltage across the signal selsyn
stator will thus have rotated through 90°, and will be directly related to the aeroplane’s magnetic
heading. Assuming that the selsyn and gyro are still in their original positions, the resultant will
now be in line with the selsyn rotor, and a maximum voltage will be induced in the rotor. The
resulting voltage error signal will then be fed to a precession amplifier, where it will be phase-
detected, and amplified, before it is passed to a torque motor fitted around the Directional
Gyro’s (DG’s) horizontal axis, as shown below.

The motor will apply a torque, which will then precess the gyro around the vertical axis, and will
also rotate the data selsyn rotor, attached to the drive shaft of the Gyro Magnetic Compass.
This will induce a voltage in the stator, which will be duplicated in the data selsyn stator in the
Master Indicator unit. If the rotor is out of alignment with the resultant magnetic field a voltage
error signal will be induced in the rotor, which will be fed to a Follow-up Amplifier, where it will
be amplified before being passed to a servomotor, which is mechanically coupled to the data

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and signal selsyn rotors. Both rotors and the dial of the Master Heading Indicator will thus
rotate, until the correct heading is indicated. The rotors are additionally coupled, such that when
rotation is complete, both rotors will lie in the null position of the fields produced in their
individual stators, ie. at right angles to the resultant of the field induced in the selsyn stators,
and hence no current will flow.

The servomotor will also drive a tachogenerator that will supply feedback signals to the Follow-
up Amplifier, and damp out any oscillations in the system. Provision is also made to transmit
heading information to other locations in the aeroplane through the installation of additional
selsyns (servo-transmitters) in the master gyro unit, and the Master Indicator.

In this example the system will again be in an equilibrium position, and the indicator will show
that the aeroplane has turned through 90°.

This system also uses the signal from the detector unit to stabilise the DG and keep it aligned
with the magnetic field sensed by it. Additionally, since the data selsyn rotor is mounted on the
shaft of the Directional Gyro (DG), if the gyro drifts, or is misaligned, an error signal will also be
created in the rotor.

Gyroscope Element
In addition to the use of efficient synchro transmitter/receiver systems, it is also essential to
employ a gyroscope that will maintain its spin axis in a horizontal position at all times. A gyro
erection mechanism is therefore essential. This consists of a torque motor, which is mounted
horizontally on top of the outer gimbal, with its stators fixed to the gimbal, and its rotor attached
to the gyro casing, as shown below.

The torque motor switch is normally of the liquid level type, and is mounted on the gyro rotor
housing, or inner gimbal, so as to move with it.

When the gyro axis is horizontal, the liquid switch is open and no current will flow to the levelling
torque motor. When the axis is tilted, however, the liquid completes the contact between the
switch centre electrode and an outer electrode, providing power in one direction or another to
the torque motor. The direction of current decides the direction of torque. The direction of the

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torque applied will be determined by the direction of the current, and will precess the gyro axis
back to the horizontal, at which time the liquid switch will be broken.

Depending on the type of compass system, the directional gyroscope element may be
contained in a panel-mounted indicator, or it may be an independent master gyro, which is
located in a remote part of the aeroplane. Systems adopting the master gyro are now the most
commonly used, because, in serving as a centralised heading source, they also provide for
more efficient transmission of the data to flight director and automatic flight control systems, with
which they are now closely linked.

Heading Indicator
An example of a basic type of dial indicator, which is used with modern gyro magnetic
compasses, is shown on the next page. In addition to displaying magnetic headings, this
indicator is also capable of showing the magnetic bearing to the aeroplane with respect to
ground stations on which radio navigation systems are sited; ADF (Automatic Direction Finding)
and VOR (Very high frequency Omni-directional Range). For this reason the indicator is
normally referred to as a Radio Magnetic Indicator (RMI), as shown on the next page.

On this instrument the flight crew are able to set a desired heading using the ‘set heading’ knob,
which is mechanically coupled to a heading bug, so that any rotation of the knob will cause the
bug to move with respect to the compass card. For turning under automatically controlled flight,
rotation of the ‘set heading’ knob will also position the rotor of a data selsyn, which will then
supply twin commands to the autopilot system. On more modern equipment, the main heading
references will be found on the EFIS and/or HSI.

Modes of Operation
All gyro compass systems provide for the selection of two modes of operation: SLAVED, in
which the gyro is monitored by the detector element, and (Free Gyro), in which the gyro is
isolated from the detector unit and functions as a straightforward directional gyroscope. The
latter operating mode is selected when a malfunction in the monitoring mode occurs or the
aeroplane is flying in latitudes where the value of ‘H’ is too small to be used as a reliable
reference.

Synchronising Indicators
The function of the synchronising indicator (annunciator) is to indicate to the user that the gyro
is synchronised with the directional reference sensed by the detector unit. The synchronisation
indicator may be integrated with the heading indicator, or may be a separate unit mounted on
the aeroplane’s instrument panel. The annunciator is activated by monitoring signals from the
detector head to the gyro slaving torque motor, and is therefore connected into the gyro slaving
circuit.

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A typical annunciator is shown on the next page. It consists of a small flag marked with a dot
and a cross, which is visible through a window in one corner of the heading indicator (if so
mounted). A small magnet is located at the other end of the shaft, and is positioned adjacent to
two soft iron cored coils, which are connected in series with the precession circuit. When the
gyro is out of synchronisation with the detector head, a current will flow through the coils, and
will attract the magnet in one direction or the other such, that either a dot or a cross shows in
the annunciator window. If the system is correctly synchronised, the annunciator window
should be clear of an image; but in practice the flag will tend to move slowly from dot to cross,
and back again, which indicates that the system is working correctly.

Manual Synchronisation
When electrical power is initially applied to a compass system operating in the "slaved" mode,
the gyroscope may be out of alignment from the detector head by a large amount. The system
will consequently start to synchronise, but, as the rate of precession is normally low (1° to 2° per
minute), some time may elapse before synchronisation is achieved. In order to speed up this
process, a manual synchronisation system is always incorporated.

A manual synchronisation knob, whose face is clearly marked with a dot and a cross, is sited on
the heading indicator. This knob is mechanically coupled to the stator of the gyro data selsyn
and when it is pushed in and rotated in the direction indicated by the annunciator, the stator will
turn, which will induce an error voltage into its rotor. This signal will then be fed to the Follow-up
amplifier and servomotor, which in turn will drive the signal selsyn rotor and gyro, via the
Precession amplifier and precession torque motor, into synchronisation with the detector head.
At the same time the signal selsyn rotor will be driven into a null position, and no error signals
will be present, when the system is fully synchronised.

Operation of an RIC in a Turn

If an aeroplane enters a turn, the gyroscope will maintain its direction with reference to a fixed
point (rigidity) and the aeroplane will thus turn around the gyro. The rotor of the data selsyn
located in the gyro unit will similarly rotate, and error signals will be generated in the stator,
which will be passed to, and reflected in the stator of the data selsyn located in the Master
Indicator. The rotor of the data selsyn will now no longer lie in the null of the induced field, and a
voltage will thus be generated, which will be passed via the Follow-up amplifier to the
servomotor. The servomotor will then drive the face of the indicator round, so that the compass
card keeps pace with the turn, and at the same time will drive the rotors of the data and signal

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selsyns to keep pace with the turn. During this time the detector unit, which is fixed in azimuth,
will be turning in the earth's magnetic field, so the flux being induced in each spoke of the
detector unit will also be continuously changing. This will result in a rotating field being produced
in the stator of the signal selsyn, which would normally result in a change in flux being detected
by the rotor of the signal selsyn, which in turn would be passed as an error signal to the
precession circuit. The rotor of the signal selsyn is however already rotating under the influence
of the servomotor. The speed and direction of rotation of the rotor therefore matches that of the
stator field, and thus no error signal is present for transmission to the precession circuit, so no
gyro precession occurs.

When the aeroplane resumes straight and level flight, rotation of the data selsyn rotor in the
gyro unit will cease. The field between the stators will thus remain constant, and no current will
flow in the servo-loop. The heading indicator display will also stop rotating and the system will
return to its previous synchronised condition.

In a steep and prolonged turn, a slight de-synchronisation may occur due to the introduction of a
small component of ‘Z’, while the detector head is out of the horizontal for a protracted period of
time. However, on coming out of the turn, the compass card will rapidly resume the correct
heading through the normal precession process. Apart from this small error, the system is
virtually clear of turning and acceleration errors.

Advantages of a Remote Indicating Gyro Magnetic Compass

The advantages of a gyro magnetic compass over a DI or direct reading instrument are:-

¾ The DI suffers from real and apparent drift and has to be reset in flight. Also, when
resetting to the magnetic compass, the aeroplane must be flown straight and level,
whereas the detector unit constantly monitors the gyro magnetic compass.

¾ The detector unit can be installed in a remote part of the aeroplane, well away from
electrical circuits and other influences due to airframe magnetism.

¾ The flux valve technique used in the detector unit senses the earth's magnetic
meridian rather than seeking it, which makes the system more sensitive to small
components of ‘H’.

¾ The unit will provide a heading reference to higher magnetic latitudes than the
direct reading magnetic compass.

¾ Turning and acceleration errors are minimised because:-

1. The detector unit is fixed in azimuth.

2. The precession signals aligning the compass are kept to a low value
(typically 5° Minute)
3. There is a roll cut out switch that isolates the precession system during
turns at bank angles of greater than 10°.

¾ The compass may be detached from the detector unit by a simple switch selection
to work as a DI, so a normal DI is therefore not required.

¾ The system can be readily used to monitor other equipment, eg. autopilot, Doppler,
RMI, etc.

¾ Repeaters can also be made available to as many crew stations or equipment’s as

desired.

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Disadvantages of the Remote Indicating Gyro Magnetic Compass
The disadvantages of a gyro magnetic compass over a DI or direct reading instrument are that it
is:-

¾ much heavier than a direct reading compass.

¾ much more expensive.

¾ electrical in operation and therefore susceptible to electrical failure.

¾ much more complicated than a DI or a direct reading compass.

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Intentionally Left Blank

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Chapter 3.1
Gyroscopic Principles

Introduction
A gyroscopic system is one where a heavy metal wheel or rotor is universally mounted, so that it
has three planes of freedom, as shown in the diagram below.

The axes are aligned to provide:-

Spinning Freedom: rotation about the spin axis (XX1).

Tilting Freedom: rotation about a horizontal axis acting at 90° to the spin axis (YY1).

Veering Freedom: rotation about a vertical axis acting at 90° to the other two axes
(ZZ1).

Freedom of movement within the three planes is obtained by mounting the rotor in two
concentrically pivoted rings, called inner and outer gimbal rings. The gimbal system is
mounted, so that with the gyro in its normal operating position, all of its axes will run mutually at
right angles to each other, and intersect at the centre of the rotor. A gyro with its axis of rotation
in the horizontal plane is known as a ‘vertical’ gyro, and a gyro with its axis of rotation in the
vertical plane is known as a ‘horizontal’ gyro. The gyro’s plane of rotation contains the sensitive
axes.

Principle of Construction
A gyroscope consists of a weighted wheel or rotor, which spins at high speed
(8,000-24,000 rpm) and is mounted in a series of hinged mounting rings, called gimbals, as
shown on the next page. A gyro has 3 axes of freedom, one of which is its spin axis, and is
able to move relative to the mounting base around one or both of the remaining axes. Ignoring
the spin axis, one degree of freedom exists when the gyro can rotate around only one axis, and

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two degrees of freedom exist if it is free to move around both axes. A gyro with two degrees of
freedom is known as a ‘Free’ or ‘Space’ Gyro.

Gyroscopic properties
When the rotor is made to spin at high speed, the device becomes a true gyroscope, and
possesses the following fundamental properties:-

Rigidity in Space. The gyro will try to remain pointing in the same direction or position
in space, even when its mounting base is tilted or rotated, ie. the gyro will have a
tendency to maintain its plane of spin, dependent on its speed, mass, and radius about
which the mass is displaced (Newton’s 1st Law). The greater the rigidity the more
difficult it is to move the rotor away from its plane of spin, unless an external force acts
on it. For example if a spinning bicycle wheel was to fall over its spin axis must be
rotated through 90°, and unless an external force is applied, it will not do so as long as
it has a reasonable speed is maintained.

Rigidity (R) = S x I
F
where: I = Moment of Inertia
F = External Force
S = Speed of Wheel (rpm)

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Rigidity is thus proportional to the rotors rpm, and its moment of inertia, which will be
increased if the rotor has a large radius, with the bulk of its mass being distributed
around its rim. This is opposite to that in the DRC where the mass is placed as close to
the pivot point axis as possible to prevent aperiodicity.

Precession. When an external force is applied to the spinning rotor via the gimbal
assembly the gyro does not move in the direction of the force, but in a direction
perpendicular to that of the applied force, eg. if a force is applied to the spin axis the
gyro will not move in the direction of the applied force, but will instead rotate due to a
force being applied 90° later in the plane of rotation, as shown below.

APPLIED
FORCE

THE REACTION OF THE

THIS POINT WILL MOVE, IE.
APPLIED FORCE IS AT
PRECESS IN THE
THIS POINT ON THE RIM
DIRECTION OF THE ARROW

Similarly if a downward push force is applied to the inner gimbal of a gyro system, the
gyro will precess about its outer gimbal pivot point.

The strength and direction of the applied force, the moment of inertia of the rotor, and the
angular velocity of the rotor will all affect the amount of precession. It follows that, the larger
the force the greater the rate of precession, and the higher the rigidity, the lower the rate of
precession. Thus in order to precess a gyro its rigidity must firstly be overcome.

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Applied Force
Precession =
Rigidity

Precession will continue while the force is applied, until the plane of rotation is in line with
the direction of the applied force. At that point, the force will no longer disturb the plane of
rotation, so that there will be no further resistance to the force, and thus no further
precession. The axis about which a force or torque is applied to a gyro is termed the ‘input
axis’, and the one about which precession takes place is termed the ‘output axis’. Most
gyros are rotated at a constant rpm, particularly where precession plays a part in the
operation of a system or indicator. If the direction of the rotor, or direction of the applied
force is reversed then the direction in which precession occurs will also reverse

Types of Gyroscope
The number of degrees of freedom permitted by each type of gyroscope determines its usage,
but for use in aeroplanes they must exhibit two essential reference datum’s. The first is a
reference against which pitch and roll attitude changes can be detected, and the second is a
directional reference against which changes about the vertical axis can be detected.

The following types of gyroscope exist:-

Space (or free) Gyro. This is a gyro having freedom to move in all three planes. It
consists of two concentrically pivoted rings, called inner and outer gimbal rings. The
three planes relate to the three axes of the aeroplane, ie. fore and aft or roll axis, lateral
or pitch axis, and the normal or yaw axis. Furthermore there is no means of external
control over this type of gyro, a feature, which distinguishes it from a tied or earth gyro.
This type of gyro would thus have no practical use in an aeroplane instrument where
the gyro is required to be set to, and maintain a given direction.

Tied (or Displacement) Gyro. This type of gyro is basically a ‘space gyro’, which has
a means of external control, and has freedom of movement about all three planes. This
type is used as a directional gyro, eg. in the Direction Indicator (DI).

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Earth Gyro. This type of gyro is a ‘tied’ gyro, where the controlling force is the gravity
of the earth. This type is used in gyro horizon, or artificial horizon instruments.

Rate Gyro. This is a gyro having one plane of freedom only, its plane of rotation being
90° removed from its plane of freedom. This type of gyro is used to measure the rate of
turn, and employs restraining springs, eg. in the turn and balance indicator or turn co-
ordinator.

Rate Integrating Gyro. This type of gyro is similar to the rate gyro, having a single
degree of freedom, except it uses the viscosity of a fluid (viscous restraint) to damp the
precessional rotation about its output axis, instead of restraining springs. The main
function of this type of gyro is to detect turning about its input axis by precessing about
its output axis. This type is used on inertial navigation stablised platforms.

Solid State (Ring Laser) Gyro. These are not gyros in the true sense, but they
behave like gyros, and sense the angular rate of motion about a single axis. They
consist of a solid block of temperature stable glass within which there is a cavity or laser
path, filled with a lasing medium such as helium-neon. Some are triangular in shape
(Honeywell), whilst others have four sides (Litton). They both have small tunnels drilled
in them, with reflecting mirrors sited at each corner. Two beams of high-energy laser
light are passed in opposite directions around the sealed cavity and initially travel at the
same speed. Any rotation of the ‘gyro’ in the ‘laser’ plane will result in a change in the
path lengths of each beam, and the resultant frequency shift of the beams is measured,
using a control element. The frequency differential is directly proportional to the angular
turning rate.

Power Sources for Gyroscopes

Conventional gyroscopes in aeroplanes are either air (vacuum) driven, or electrically driven. In
some aeroplane all of the gyros are either vacuum or electrically driven, whilst in others the
vacuum (suction) system provides the power necessary to run the attitude and heading
indicators, whilst an electrical system runs the slip and turn indicator. Alternating or direct
current is used to power the electrically driven gyroscopic instruments.

Air Driven Gyroscopes. This type is widely used on small aeroplanes, and is still
found on some large aeroplanes in order to power stand-by, or emergency instruments.
In the air driven gyro, the gyro system is contained in an airtight case, the air having
been removed from the casing via a suction pump or venturi, thus creating a partial
vacuum inside. The vacuum source is normally an engine driven vacuum pump, which
is controlled to a value between 4.5-5.0 inches of Hg, but in very simple aeroplanes, this
may be achieved using a venturi tube attached to the outside of the fuselage. Air under
atmospheric pressure enters the instrument casing via a filter, and flows through a
shroud, which encases the rotor.

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The shroud acts as the inner gimbal and has a small opening in it that directs the airflow
being fed to it via an inlet port onto buckets cut in the rotor periphery. The shroud also
has an exhaust port to allow air to escape to the outer case. The pipeline from the inner
gimbal is fed through the inner gimbal axis along the outer gimbal and through its axis
to a filter system that covers a hole in the outer case.

When the vacuum is applied to the system the pressure in the outer case will drop.
Replacement air will thus enter the system from outside the case and will pass through
the filter, before being directed onto the rotor buckets. The rotor rotates at
approximately 10,000 rpm under normal operation. The air escapes from the rotor
shroud, although it can be used to provide a controlling force (tie) to the gyro.

Electrical Gyros. The majority of gyroscopes used in aeroplanes today are electrically
driven, and normally use AC current, although a 24-volt DC supply can feed some. The
AC powered gyro is preferable, since it avoids the use of commutators and brush gear,
which require frequent servicing. The gyroscope itself is a 3-phase squirrel induction
motor, which is constructed to obtain the maximum gyroscopic effect. To achieve this,
the rotor does not rotate inside the field coils, as it does in conventional motors, but
instead is positioned around the outside of the field coils. This method of construction
ensures that the rotor mass is concentrated as near its periphery as possible, thus
increasing gyro rigidity. The gyro stator is fed with a 115 volt, 3-phase, 400 hertz AC,
and the rotor rotates at approximately 24,000 rpm. AC gyros are capable of higher
rotational speeds than the DC ones, and are therefore the favoured option in
instruments where high rigidity is required. DC powered gyros can also be powered
from the aeroplane emergency electrical supply.

The Disadvantages and Advantages of Air driven Gyro’s

The disadvantages of an air driven gyro’s are that:-

¾ Full rigidity is not be reached until the rotor speed has built up. Suction driven
gyroscopes normally take 4-5 minutes after starting the aeroplane engine to attain
the correct operating speed. The indications provided by the instrument are
however usable after 2 minutes. Conversely, venturi driven systems will not start to
spin up until the aeroplane begins its take-off run, and cloud penetration must be
delayed for a few minutes to allow the operating rpm to be reached.

¾ The speed of the rotor depends on the mass of air flowing through the system
(mass flow). As the aeroplane climbs, the air density falls and the mass flow
reduces. The rotor speed thus reduces and gyro rigidity deteriorates. The other
thing about mass flow is that it requires a clear unimpeded flow of air. If the filters
on the inlet line are blocked or partially blocked, this also will affect the gyro rigidity.

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¾ A major drawback of the air driven gyro is the need to provide airtight joints where
the inlet pipes pass through the inner and outer gimbal axes, which severely limits
the degree of freedom around these axes.

¾ Ingested dust or moisture will cause corrosion and bearing wear.

The advantages of air driven gyro’s are that:-

¾ They are cheap.

¾ They are easy to maintain.

¾ In the event of an emergency, they can operate without electrical power.

The Disadvantages and Advantages of Electrically Driven Gyro’s

The disadvantages of electrically driven gyro’s are that they:-

¾ depend on their power supply, so standby air driven alternatives are normally fitted.

¾ tend to be more expensive than air driven gyros.

The advantages of electrically driven gyros are that:-

¾ higher rigidity is possible.

¾ their operating rpm is more consistent.

¾ their performance is not affected by altitude.

¾ information can be transmitted easily to other systems.

¾ they have more freedom to rotate around their axes.

¾ the Instrument case is completely sealed, which excludes dirt, and also prevents
heating/cooling effects by allowing the components to be maintained at a constant
temperature, if required.

Gyro Wander
Any deviation of the gyro spin axis from its set direction is known as ‘gyro wander’, and is
classified as follows:-

Real Wander. Any physical deviation of the gyro spin axis is called real wander. A
gyro should not wander away from its preset direction, but various forces act on the
rotating mass of a gyro and cause it to precess. For example bearing friction, which is
always present at the spin axis. If this friction is symmetrical, it will merely slow down
the rotor, but if it is asymmetrical it will cause the gyro to precess. Similarly any friction
in the gimbal bearings will cause the gyro to precess. Wear on the gyro may result in
movement of the C of G, which may also result in a precessing force. Such errors are
not constant or predictable, and can therefore not be calibrated for, or corrections
applied to nullify this error.

Apparent Wander. In this case the gyro spin axis does not physically wander away
from its pre-set direction, but to an observer it will appear to have changed its direction.
This is because the gyro maintains its direction with respect to a fixed point in space,

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whereas the observer rotates with the earth, thus with the passage of time the gyro will
appear to have changed direction, with reference to an earth datum. Apparent wander
is also made up of horizontal components called ‘drift’, and vertical components called
‘topple’. The rate of drift and topple depends upon the latitude and can vary from zero
to a maximum of 15.04° per hour (the rate at which the earth rotates). Depending on
whether a gyro has a vertical or horizontal spin axis, the rotation of the earth will also
have a different effect.

Horizontal Axis Gyro. The diagram below shows a horizontal spin axis gyro
positioned at the North Pole.

It shows an observer initially at position A, where the gyro is set so that its spin axis is
directly in line with him. Six hours later, the earth having rotated through 90°, the
observer will now view the gyro from position B. The observer will not however
appreciate his own motion, and the gyro spin axis will appear to have moved clockwise
in the horizontal plane through 90°. Twelve hours later the gyro spin axis will appear to
have moved through 180°, and finally after twenty-four hours, with the observer back in
his original position, the gyro spin axis will again appear as it was first aligned. The
apparent motion in the horizontal plane is known as ‘Gyro Drift’.

If a horizontal spin axis gyro has its axis aligned in a north/south direction along the
equator, during the earth rotation, the gyro spin axis will continue to remain aligned with
the local meridian. This occurs because at the equator all of the meridians are parallel
to one another, and a gyro aligned with a meridian will thus remain with that meridian
over a 24 hour period. This means that the gyros will neither drift nor topple when it is
aligned in this manner.

If the horizontal spin axis gyro is positioned at the poles it will drift through 360° in 24
hours (maximum drift), ie. the rate of drift at the poles will be the same as the angular
velocity of the earth, at 15.04° per hour, whilst at the equator the same gyro with its spin
axis aligned with the local meridian will have zero drift due to earth rotation.

Drift at intermediate latitudes = 15.04° x Sin Latitude° per hour.

The diagram below shows a horizontal spin axis gyro with its spin axis aligned in an
east/west direction along the equator, when observed at point A.

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EQUATOR

DIRECTION
OF EARTH’S
ROTATION

In this case, after six hours duration, and the earth having rotated through 90°, the
observer will again view the gyro from position B, where it will appear to have turned
into a vertical axis gyro. This apparent change in its vertical plane is known as
‘Topple’, and will be maximum value at the equator, but zero at the poles, due to the
earth’s rotation.

Topple at intermediate latitudes = 15.04° x Cos Latitude° per hour.

Vertical Axis Gyro. Diagram (a) below shows a vertical spin axis gyro, which is
positioned at the North Pole, where the rotation of the earth beneath the gyro spin axis
has no effect on the gyro, ie. the gyro axis will appear to neither drift nor topple.
DIRECTION OF
EARTH’S ROTATION

EQUATOR

(a) (b)

Diagram (b) above shows a gyro at the equator, with its spin axis vertical to the
observer, when viewed at position A. After six hours, and the earth having rotated
through 90°, the observer at point B will view the gyro as a horizontal axis gyro. After a
further six hours the spin axis will again appear vertical. This apparent change in the
gyros vertical axis is known as ‘gyro topple’. At the equator gyro topple is 360° in
24 hours, whilst at the poles it is zero.

From the above you’ll notice that the vertical gyro suffers from topple but does not suffer
from drift:-

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Drift = Vertical gyro unaffected

Topple = 15.04° Cos latitude° per hour

Gyro drift and topple may be summarised as:-

Horizontal axis gyro: at the poles: maximum drift and no topple

at the equator: no drift and maximum topple

Vertical axis gyro: at the poles; no drift and no topple

at the equator: no drift and maximum topple

Transport Wander. This is an additional form of apparent topple/drift, which principally

occurs when the gyro is placed on a platform, such as an aeroplane, that is flying in an east
or west direction. The gyro is now being carried in space in the same way as the earth
does, and results in ‘Transport Wander’.

Transport Drift = Rate of change of longitude° per hour x Sin latitude° per hour

Transport Topple = Rate of change of longitude° per hour x Cos latitude° per hour

Examples of Gyro wander

EG 1: If a horizontal spin axis gyro is set with its axis aligned in an east/west direction
at latitude 45°N, the attitude of its spin axis after 3 hours will be:-

Since the gyro axis is aligned in an east/west direction at an intermediate latitude, the
gyro will both drift and topple.

Drift = 15.04° Sin latitude° per hour = 3 x 15.04° x Sin 45° = 31.9°

Topple = 15.04° Cos latitude° per hour = 3 x 15.04° x Cos 45° = 31.9°

Note: In the Northern Hemisphere the gyro axis will drift clockwise, and anti-clockwise
in the Southern Hemisphere

The spin axis will thus be aligned at 090° + 31.9° = 121.9° / 301.9°

The eastern end of the spin axis will appear to have risen by 31.9° from the horizontal,
and the western end will be similarly depressed.

EG. 2: If the rate of change of longitude during a flight is 25° in one hour, at latitude
50°N, the amount of transport drift present will be:-

Transport Drift = Rate of change of longitude° per hour x Sin latitude° per hour

Transport drift = 25° x Sin 50° = 19.15°

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Chapter 3.2
Direction Indicator

Introduction
The Direction Gyro Indicator (DI) uses a horizontal axis tied gyro, which possesses freedom in
three planes, and uses the gyroscope’s property of rigidity, to stabilise an azimuth scale. It is
manually aligned with the direct reading magnetic compass and, in light aeroplane, provides a
stabilised directional reference for maintaining and turning accurately on to a heading. The DI is
non-magnetic and is thus not subject to turning and acceleration errors, dip or magnetic
disturbances. The DI provides an accurate dead-beat indication of heading, and shows any
deviation from the set heading instantaneously. The DI is also not north seeking, so it must be
provided with a directional datum from an outside source, which is normally taken from the
direct reading magnetic compass. It is thus essential that the DI indications be checked, at
regular intervals, against the direct reading magnetic compass, because after the initial
synchronisation the gyro may wander, particularly after aerobatics. The DI is thus designed to
compliment the DRC, and not replace it.

Basic Description of the Direction Indicator

The instrument can be either air driven or electrically driven. In the air driven version the
instrument consists of an air-driven horizontal axis gyro, rotating at approximately 10,000 rpm.

The azimuth scale is graduated from 0° to 360° in 5° divisions, with main graduations every 10°,
and figures every 30°. The scale is then read of against a vertical lubber line.

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Below the window is sited a knob, which is used to cage the gyro, and also to rotate the gyro
assembly when, setting a given heading.

Operation of the Direction Indicator

The diagram below shows the internal mechanism of a typical air driven version of the
instrument.

The rotor spins about its horizontal axis, and is supported in bearings in the inner gimbal ring,
which is free to rotate about the horizontal axis through 110°, ie. 55° either side of its central
position, when it is uncaged. The inner gimbal ring is supported in bearings in the vertical outer
gimbal ring and is free to rotate in azimuth through 360° about the vertical axis. A nozzle is
sited in the outer gimbal, and directs a jet of air onto buckets cut in the rotor periphery. The
action of the air ensures that the rotor reaches its operating rpm after about five minutes, when
full suction is developed by the vacuum pump. The air jet also maintains the rotor spin axis in
the horizontal plane, as shown below. If the gyro topples, a component of the jet force will act at
right angles to the rotor, and will produce a precessing force, which will erect the gyro.

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In this version the indicator scale is attached to the outer gimbal, but on newer models, the
synchronising gear ring also drives a sequence of gears, which connect the movements of the
gyro around its vertical axis, onto a vertical scale, as shown below.

The gyro is initially erected using a caging mechanism, which manually places the gyro spin
axis in its horizontal plane. The mechanism consists of a bevel pinion and a caging arm, which
are both directly controlled by a caging or setting knob sited on the front of the instrument.
When the knob is pushed in the bevel pinion engages with the synchroniser gear ring, and
allows the scale to be adjusted in azimuth by rotating the caging knob.

At the same time the caging arm is raised, which locks the inner gimbal ring in its horizontal
plane, and prevents the gimbal ring and rotor from toppling during resetting. The caging knob is
pulled out to uncage the gyroscope, by disengaging the gears and allowing the caging arm to

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drop, thus releasing the inner gimbal ring. The inner gimbal should also be locked in its
horizontal position during aerobatics, to prevent severe loads being transmitted to the rotor
bearings.

In the electrical version, the rotor is part of an AC ‘squirrel cage’ induction motor, and rotates at
approximately 24,000 rpm. Initial erection is again by use of the caging device, but thereafter
the gyro is tied so that it maintains its spin axis horizontal to the earth’s surface.

Errors Associated with the Air Driven Direction Indicator

The air driven Direction Indicator is subject to the following errors:-

Real Drift and Topple. The gyro is subject to both of these errors, which are caused by
mechanical imperfections such as a slight imbalance of the rotor/gimbal system,
bearing friction and mechanical or electrical latitude correction.

Apparent Drift. Earth rate (due to earth rotation) and transport wander (due to the gyro
being transported in an east/west direction) cause this error. A fixed rate of
compensation is calculated for the latitude of operation of the aeroplane for apparent
wander, and may be applied by means of balancing nuts, which are attached to the
gimbals. The rate of compensation, but no compensation for transport wander is made
apart from periodic resetting of the instrument to the magnetic compass heading. The
drift rate can be as much as 15.04°/hour; and for this reason the instrument should be
realigned with the magnetic heading shown on the direct reading compass
approximately every 15 minutes, whilst the aeroplane is in level and unaccelerated
flight. Any topple experienced by the gyro will be corrected for by the air jet erection
system.

Gimballing Error. Any misalignment between the aeroplane axes and the navigation
system axes will cause this error. This error only exists during manoeuvres, ie. during
climbing, descending or banking, but once level flight is resumed, will disappear.

Use of the Direction Indicator (DI)

On small basic aeroplanes and some older aeroplanes, the direction indicator is the primary
heading reference used to maintain the required heading, although great care must be
exercised when using this instrument, due to the various drift errors that exist.

Prior to departure the DI should be aligned with the magnetic compass using the heading set
knob. Also during the pre-flight process, and during taxi for take-of, the two readings should be
periodically compared to make sure that the gyro is not showing large drift rates. Once in flight
the DI should be again checked and reset against the magnetic compass approximately every
15 or 20 minutes, with the aeroplane in a steady level flight condition.

Advanced Use of the Direction Indicator

In areas where the Earth’s magnetic field does not provide a stable heading reference, it is
necessary to base any headings on an unmonitored DI. This is a highly sophisticated
instrument, where any friction effects are reduced to a minimum, although the instrument still
suffers from drift errors. These errors therefore need to be calculated before they can be
compensated for.

Sample Calculation
A DI has its latitude compensatory device set to correct for operations at latitude 45° N, and is
fitted to an aeroplane flying eastwards along the parallel 60°N, at a ground speed of 240 knots.
The expected drift rate (°/hour) if the gyro is free from random drift errors will be:-

¾ Real drift (gyro drift) from the fixed latitude correction = 15.04° x Sin latitude°/hour.

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¾ Apparent drift (earth’s rotation) at the latitude of operation = 15.04° x Sin

latitude°/hour (- for a decrease in the Northern Hemisphere, or + for an increase in
the Southern Hemisphere).

¾ Transport drift in an Easterly direction = departure x Tan mid latitude°/hour.

Ground Speed
= x Tan mid latitude°/hour
60

Note:(- if easterly tracking, or + if westerly tracking in the Northern Hemisphere)

Solution:

Apparent Drift = 15.04° x sin 60° = -13.03°/hour (decrease)

Transport Drift (convergence) = 240 x tan 60° = - 6.93°/hour (east)

60

Total Drift = Gyro Drift + Transport Drift = 13.03 + 6.93 = -19.96°/hour (decrease)

Latitude correction (real drift) = 15.04° x sin 45° = +10.63°/hour (increase)

The Total Anticipated Drift will be = - 9.33 °/hour (decrease)

By comparing the actual DI heading with the expected one (derived from these calculations),
the pilot can then use this information to establish the real (random) drift error that exists.

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Intentionally Left Blank

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Chapter 3.3
Artificial Horizon

Introduction
The artificial horizon (gyro horizon) comprises of a vertical spin axis earth gyro having freedom
of movement in all three planes, and indicates the aeroplane attitude relative to its pitch (lateral)
and roll (longitudinal) axes, which is essential when a natural horizon is unavailable, eg. when
flying in cloud.

The instrument is either air or electrically driven, although the principal of operation is identical.
The gyro spin axis is maintained vertical with reference to the centre of the earth, and a bar
positioned at 90° to the spin axis represents the local horizon. A symbol representing a
miniature model aeroplane is fixed to the instrument case, and represents the rear view of the
true aeroplane, which on some instruments is adjustable to suit the pilots own eye level, and the
particular aeroplane pitch trim setting. A typical artificial horizon display is shown on the next
page.

In flight, the aeroplane’s movement about its pitch or roll axis is indicated instantaneously by
movement of the case relative to a horizon (natural horizon) bar, which is held in the local
horizontal by gyro rigidity. The position of the model aeroplane relative to the bar represents
the attitude of the aeroplane to the natural horizon, whilst the position of a pointer relative to a
fixed scale represents the aeroplane’s angle of bank.

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Dive and climb are indicated by the model aeroplane moving up and down with respect to the
horizon bar, whilst the angle of bank is indicated by the model aeroplane appearing to bank in
relation to the horizon bar. The indications expected during various flight attitudes are shown
below.

The exact angle of bank is indicated by a pointer at the bottom of the instrument, and provides a
direct indication of any change of attitude, without any lag being involved.

Construction of an Air driven (Classic) Artificial Horizon

A schematic view of an air driven artificial horizon is shown on the next page. This type is
commonly used in light aeroplane, and as a standby instrument in commercial aeroplanes. It is
operated by a vacuum pump, which evacuates the air from the instrument case and gyro
housing (inner gimbal). This creates a depression within the instrument, and the surrounding
atmosphere enters the instrument through a filtered inlet. The air then passes through channels
to jets mounted within the inner gimbal, which direct air onto buckets cut into the periphery of
the rotor, and cause the rotor to rotate at approximately 13,000 rpm, in an anti-clockwise
direction when viewed from above. The air is then evacuated through a pendulous unit,
mounted below the rotor casing, via four ports that are controlled by two pairs of linked

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pendulous vanes, and provide a mechanism for maintaining the gyro spin axis in its vertical
plane.

The rotor spins about a vertical axis (Z-Z1) and is mounted in bearings within a sealed case,
which forms the inner gimbal.

X1

The inner gimbal is mounted in bearings within a rectangular shaped outer gimbal and is free to
rotate 55° either side of its horizontal position, about the lateral axis (Y-Y1). This enables the
aeroplane’s pitch attitude to be determined, and is directly indicated by movement of the horizon
bar. The horizon bar arm is actuated by a guide pin, which protrudes from the gyro stabilised
rotor housing (inner gimbal), and moves in a curved guide slot in the outer gimbal.

The outer gimbal is mounted in an air tight instrument case, with its pivots along the fore and aft
axis (X-X1), and is free to rotate through 110° either side of its central position, in order to
determine the roll attitude of the aeroplane. A background plate representing the sky is fixed to
the front end of the outer gimbal and carries a bank pointer, which registers against a bank-
angle scale. Movement in both cases is limited by resilient stops, which prevent any internal
damage to the instrument.

The instrument is gyro stabilised, and arranged so that when the gyro is erect, the horizon bar is
horizontal with reference to the earth’s surface, and the angle of bank pointer is in its centre
position, showing the gyro to be vertically erect with reference to the earth’s surface.

Bank indication is given by an index on the sky plate, which reads against a scale printed on the
glass face of the instrument. When the aeroplane banks, the rotor, inner gimbal and outer
gimbal remain rigid in their level position, whilst the instrument case, and hence printed scale,
moves with the aeroplane; thus the position of the sky plate index indicates the aeroplane’s
bank angle against the scale.

Operation of an Air driven Artificial Horizon

During level flight the aeroplane's vertical axis is parallel to the rotor spin axis, with the guide pin
in the centre of the slot in the outer gimbal, and the horizon bar centralised. During a climb or

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descent, the rotor, and hence inner gimbal will remains rigid with reference to the local vertical,
whilst the outer gimbal and instrument case will move with the aeroplane, and turn about the
Y-Y1 axis. When the aeroplane starts to climb, the rear of the instrument case and outer gimbal
will follow the nose of the aeroplane and will rise up. This will cause the guide pin to move,
relative to the inner gimbal, thus displacing in the slot in the outer gimbal, and placing the
horizon bar below the model aeroplane, giving a relative indication of a climb. Conversely when
the aeroplane starts to descend the rear of the instrument case and outer gimbal will be
depressed with the nose of the aeroplane. The movement of the guide pin will cause the
horizon bar to move above the model aeroplane, thus providing a relative indication of a dive.

During a roll manoeuvre the instrument case and model aeroplane will rotate about the fore and
aft axis (X-X1), but the gyro assembly, including the inner gimbal, outer gimbal and horizon bar
will remain level. The model aeroplane will thus turn in relation to the horizon bar, and will
provide an indication of bank.

The Air Driven Artificial Horizon Erection System

The air driven instrument incorporates a mechanical pendulous vane unit, which erects the
gyroscope into its vertical position, and also maintains its spin axis in that position during its
operation. A typical unit is shown below.

VANE

The unit is fastened to the underside of the rotor housing and consists of four knife-edged,
pendulously suspended vanes, which are fixed in diametrically opposed pairs, on two shafts
supported in the unit body. One shaft is parallel to the pitch axis (Y-Y,), whilst the other is
parallel to the roll axis (X-X,) of the gyroscope. In the sides of the unit body are four small,
elongated ports, one located under each vane. Suction air, after spinning the gyro rotor, is
exhausted through the ports, and the reaction of these diametrically opposed streams of air

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applies a force to the unit body. The vanes, under the influence of gravity, always hang in the
vertical position, and govern the amount of airflow from the ports. They also control the forces
applied to the gyroscope by the exhaust air reaction forces. When the gyroscope is in its
vertical position the knife-edge of each vanes will equally bisect each port, thus making all four
port openings of equal dimension, as shown on the next page.

The air reactions will similarly be equal, and the resultant forces about each axis will be in
balance. If the spin axis is however displaced from its vertical position, as shown below, the
pair of vanes positioned on the Y-Y, axis will remain vertical, thus fully opening one port whilst
the diametrically opposing port will be fully closed.

Y1

The increased reaction force produced by the air being expelled from the fully open port will
result in a torque being applied to the gyro body in the direction of the arrow, and thus according
to the law of precession, the unit will rotate about the pitch axis (Y-Y1),. The spin axis will
therefore be returned to its local vertical or erect position, when the vanes will again equally
bisect the ports, and will result in equal reaction forces again.

Errors Associated with the Air Driven Artificial Horizon

The air driven artificial horizon suffers from both acceleration and turning errors, and for the
purpose of explanation it is assumed that the gyro rotor rotates in an anti-clockwise direction
when viewed from above.

Acceleration Error. This error is also known as the ‘Take-off error’, since is most
noticeable during the take-off phase of flight, and is caused by the pendulous unit and
its associated vanes. The pendulous unit makes the rotor housing (inner gimbal)
bottom-heavy, so that when the aeroplane accelerates, a force due to the unit's inertia,
which is effective at the bottom of the rotor system, will act in the direction of the flight
crew. The resulting force will be precessed through 90° in an anti-clockwise direction,
and will lift up the right-hand side of the outer gimbal. This will cause the sky-plate,

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which is attached to the outer gimbal to rotate anti-clockwise, and will indicate a false
turn to the right against the bank angle index. Additionally during the acceleration both
of the laterally (left and right) mounted side vanes will additionally be thrown aft, with the
result that the right-hand side port will fully open and the left-hand side port will fully
closed. This will in turn produce a reaction force on the right-hand side, which when
precessed through 90°, will lift the inner gimbal, and will indicate a false climb. A
classic artificial horizon will thus indicate a false climbing turn to right during the
take-off phase of flight.

Turning Error. During a turn the longitudinally (fore and aft) mounted vanes on the air-
driven artificial horizon will be displaced due to the centrifugal force acting on the
pendulous unit. This will cause one port to open, whilst the opposing port will close, and
a reaction force will be set up along the fore and aft axis of the aeroplane (X-X1). After
precessing the force through 90°, it will tend to lift the outer gimbal on the left or right
hand side depending on the direction of the turn. This will result in a false bank
indication, or ’Erection Error’. It follows that during a left turn the instrument will
indicate a reduced left bank indication, whilst during a right turn the instrument will
indicate a reduced right bank indication.

The centrifugal force will additionally cause the pendulous unit to swing outwards in the
opposite direction to that of the turn, which will cause the inner gimbal to give a false
indication of climb or descent. This is alternatively known as a ‘Pendulosity Error’.
During a left turn the classic artificial horizon will indicate a false climb, and during a
right turn will indicate a false descent. These two forces will act together, and during a
360° turn will reach a maximum value at 180°, and return to zero when the turn is
complete. In modern gyroscopes however the axis of rotation is slightly offset from its
true vertical to counter these errors, although this is only valid for one particular rate of
turn, and airspeed. The scales are similarly offset so that the indications are not
affected during straight and level flight.

Construction of an Electrically Driven Artificial Horizon

An example of an electrically driven artificial horizon is shown below.

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It is made up with the same basic components as the vacuum-driven type, except that the
vertical spin axis gyroscope is a squirrel-cage induction motor. Unlike conventional induction
motors, where the rotor normally revolves inside the stator, in order to make the motor small
enough to be accommodated within the space available in a modern miniaturised instrument,
the rotor is designed so that it rotates in bearings outside the stator. This ensures that the mass
of the rotor is concentrated as near to the periphery as possible, thus ensuring maximum inertia,
and adequate rigidity.

The "squirrel-cage motor" design is not only used in the artificial horizon, but is also used in
other instruments that employ electrical gyroscopes. The motor assembly is carried in a housing
that forms the inner gimbal, and is supported in bearings in the outer gimbal, which in turn is
supported in bearings in the front, and rear casing of the instrument. The horizon bar assembly
is in two parts, and like the air driven version is similarly pivoted at the rear of the outer gimbal.
The instrument is fitted with a torque motor erection system, which maintains the gyro in its
vertical axis. The electrical motor rotates the rotor at approximately 22,500 rpm, and if the power
supply fails it is indicated by a solenoid-actuated ‘OFF’ flag, which appears in the face of the
indicator.

Torque Motor and Levelling Switch System

The torque motor electrical system consists of two torque control motors, which are
independently operated by mercury levelling switches; one is mounted parallel to the
longitudinal axis, and one is mounted parallel to the lateral axis. The lateral switch detects
displacement of the gyroscope in roll, and is connected to the torque motor mounted across the
pitch axis, whilst the longitudinal switch detects displacements in pitch, and is coupled to a
torque motor mounted across the roll axis. Each levelling switch consists of a sealed glass tube
containing three electrodes and a small quantity of mercury.

When the gyro is running and in its normal operating position, the mercury in the levelling
switches will lie in the centre of the tubes, and will only be in contact with the centre electrode,
whilst the two outer electrodes, which are connected to their respective torque motors, will
remain open as shown above.

The autotransformer in the system reduces the voltage to a nominal value (20V), which is then
fed to the centre electrode of the switches, so that no current flows to the torque motors when
the system is level. If the gyro is displaced, eg. about the pitch axis (Y-Y1), the pitch-levelling
switch will be displaced, and the mercury will roll in the direction of pitch to make contact with
one of the outer electrodes. This will result in the electrical circuit being completed to the
laterally mounted pitch torque motor, thus energising the motor, and causing it to apply a torque
force. According to the law of precession, a subsequent force will act on the gyro about the

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pitch axis, and will return the gyro spin axis back to its vertical position. The pitch levelling
mercury switch will now be returned to its normal operating position, with the mercury element
now only in contact with the central electrode, thereby removing the electrical supply from the
pitch torque motor.

Electrical - Fast Erection

On many electrically operated gyro horizons that employ the torque motor method of erection,
there is also a roll cut out switch, which is designed to prevent false erection signals being sent
to the erection torque motors during a prolonged turn. In this arrangement, if the gyro rotor spin
axis becomes more than 10° misaligned from the vertical, a commutator switch, fitted to the
outer gimbal ring will interrupt the current flow between the mercury switches and the torque
motors. If the gyro had been switched off, and had subsequently toppled, this device will also
prevent automatic erection taking place.

In order to overcome the problem, and to bring the gyro to its operational state as quickly as
possible, a fast erection system is provided. This system is designed to bypass the roll cut out
switch, and to apply a higher than normal voltage to the erection motors. A typical system is
shown below, where only one mercury switch and one torque motor are shown to emphasise
how the system operates

If the gyroscope exceeds the appropriate limits of movement from its vertical position, it is
important that the gyro is brought back to normal as quickly as possible, and this is achieved by
pushing in the fast-erection switch. In this position the torque motors are supplied directly with
115volts, which increases the torque motor output and hence produces greater torque. This will
result in the erection rate increasing from the normal 5° per minute, to between l20° and 180°
per minute, depending on the particular design. In order to prevent the torque motor
overheating it is important that the fast-erection switch is not be used continuously for longer
than 15 seconds, nor is it used when the gyro is in its vertical position.

Errors Associated with the Electrically Driven Artificial Horizon

The electrically driven artificial horizon like the air driven derivative similarly suffers from both
acceleration and turning errors.

Acceleration Error. In the case of the electrically driven gyro horizon, the inner gimbal
does not have a pendulous erection unit hanging below it as in the case of the air driven
or classic version, and is therefore not subject to the apparent turn component of
acceleration error. However, the mercury in the longitudinally mounted switch will hang
back and complete the circuit to the pitch torque motor, and will cause the instrument to

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indicate only a false climb, and not an apparent climbing turn to the right, as in the
case of the air driven variant.

Turning Error. The sole effect on an electrically driven gyro is to displace the mercury
in the lateral mercury switch, to complete the circuit via one or other of the outer
electrodes to the roll torque motor, causing the instrument to indicate only a false bank,
and not a false indication of turn and climb or descent, as in the case of the air driven
variant

Remote Vertical Gyro

On many modern aeroplanes, the attitude indicator is fed from a remote vertical gyro unit, which
is normally sited in the avionics bay. This gyro works in the same way as the electrical gyro just
described except that it is not linked directly to a presentation. Pitch and bank data is fed to the
remote (panel mounted) indicator by means of an electrical synchro transmission system. The
same attitude information can also be fed to the autopilot so that it can use the same data as
the flight crew are viewing. The biggest advantage of using the remote vertical gyro is that it can
provide greater degrees of freedom, and also the indicator can be constructed to present all
attitudes with virtually unlimited freedom.

Standby Attitude indicator

Many modern aeroplane employ integrated flight systems, which include indicators that can
display not only pitch and roll attitude data from a remotely located vertical axis gyroscope, but
also associated guidance data from radio navigation systems. In these systems there is no
longer a need for a separate artificial horizon to be fitted, but in order to satisfy the airworthiness
requirements one has to be fitted, as a standby attitude indicator. This provides the necessary
indication should the circuits controlling the aeroplane attitude display fail.

An example of the face of a typical standby indicator is shown above. This instrument uses an
internal gyroscope, which is electrically operated and is powered during normal operation by the
aeroplane’s II5V 3-phase supply. If the normal power supplies fail a static inverter, will provide
28V DC from the battery busbar, and will automatically supply the standby artificial horizon.
Power from this source is always available, so attitude indications are continually displayed.

In place of the conventional stabilised horizon bar method of displaying pitch and roll, a
stabilised spherical element is adopted as the reference against an aeroplane symbol. The
upper half of the element is coloured blue (sky) to display climb attitudes, while the lower half is

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black, to display descending attitudes. Each half is graduated in l0° increments up to 80° climb
and 60° descent. A pointer and scale indicates the bank angle in the normal manner. The
indicator also has a pitch-trim adjustment and a fast-erection facility. When the knob is rotated
in its "IN" position, the aeroplane symbol may be positioned through ±5° variable pitch trim.
Pulling the knob out, and holding it, will alternatively energise a fast-erection circuit.

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Chapter 3.4
Turn and Balance Indicator

Introduction
The turn and balance indicator, previously known as the turn and slip indicator, is essentially
two instruments in one casing, which provide separate indications on a common dial. A turn
indicator displays the rate of, and direction of turn, using gyroscopic principles; and a balance
indicator to show whether the aeroplane is performing a balanced or unbalanced turn (skidding
or side-slipping). The dial presentation of a typical turn and balance indicator is shown below.

Construction and Principle of Operation of a Turn Indicator

The turn indicator comprises of a horizontal spin axis gyro, which is supported in a gimbal ring,
and is mounted with its plane of rotation acting along the fore and aft or roll axis (X – X1) of the
aeroplane.

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It uses a rate gyro, and has freedom of movement in the rolling plane only. The rotor is either
electrically driven, and includes a power failure warning flag, or is air driven. Both types of drive
are structured to produce a low rotor speed of approximately 9,000 rpm, because in level flight,
the gyro axis is maintained in its horizontal position by an adjustable spiral spring.

The spring is attached between the gimbal and the instrument case. A pointer is also attached
to the gimbal, and moves over a scale showing the aeroplane’s rate of turn, which is positioned
adjacent to the zero datum mark, when the gyro is in its horizontal position, ie. when the
aeroplane is in level flight. A damping device, usually a piston cushioned by air in a cylinder, is
additionally fitted to the gimbal to ensure that the instrument reacts smoothly to changes in the
rate of turn, and at the same time reacts to a definite turn rate without pointer oscillation.

When the aeroplane turns the gyro will precess, thus tilting the rotor and gimbal ring until the
precessing force is matched by the tension of the spring. At this point the precession will cease,
and the gyro will remain inclined for the duration of the turn, giving an indication of the actual
rate of turn, which is shown by the pointers position on the scale. When the aeroplane stops
turning the gyro will return to its original horizontal position under the action of the spring.

Operation of the Turn Indicator

For example when an aeroplane enters a left turn the gyro axis, which is rigid, will oppose the
turn and a force will be experienced about the vertical input axis.

The gimbal ring will also turn with the aeroplane, but the resultant turning moment will be
resisted due to the rigidity of the gyroscope, and will precess about the longitudinal (X – X1)
axis. During a left turn a force will be applied at the front pivot of the gimbal ring, which is the
same as applying a force at point F on the rotor rim. Due to primary precession, a subsequent
force will act 90° later in the plane of rotation, ie. at point P, and will cause the gimbal ring to tilt
about the fore and aft axis. The pointer, which is connected to the gimbal ring will also move,
and in doing so will indicate the direction of turn via reverse gearing. The rate of turn can also
be established, since the force exerted by the spring is directly proportional to the amount of
gimbal deflection.

During a left turn, the gyroscope, in precessing, will stretch the spring until the force it exerts
prevents further deflection of the gyro. As the gimbal ring is deflected under the influence of

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force P, the stretched spring will exert a downward force where it is attached to the gimbal. This
equates to a force pressing on the left-hand lower part of the gyro rotor, ie. opposite to force P,
and when precessed through 90°, will produce a rotational force about the input axis, acting at
point K on the rim. Force K acts in the same direction as the original turning force, F. This is
known as Secondary Precession. When the rate of turn is established, force F will reach a
constant value, and when force K reaches the same value, ie. the forces applied are equal and
opposite, the gyro will be unable to tilt any further. Force F is due to the rigidity of the gyro, and
force K is a precessing force. The angle of tilt is therefore entirely dependent on the magnitude
of force F, whilst the rate of turn is a function of gyro tilt.

The scale showing the rate of turn is calibrated in what are termed standard rates and, although
seldom marked on the instrument, are classified by the numbers 1 to 4, corresponding to turn
rates of 180°, 360°, 540° and 720° per minute respectively. On commercial aeroplanes the
scale is normally only graduated to indicate rate one turns, since turns in excess of this rate are
not normally performed in these types of aeroplanes. This is because the majority of
passengers do not like to experience the acceleration forces imposed during tighter turns, and it
would also subject the airframe to unnecessary high load factors.

Errors Associated with the Turn Indicator

The turn indicator does not suffer from apparent wander because the spring prevents topple in
the vertical plane, and drift in the horizontal plane is impossible due to the instruments
construction. Mechanical or real wander is also normally negligible, providing that the spring
tension has been correctly adjusted.

Erroneous indications may however be caused if the rotor speed fluctuates too far from its
normal operating rpm. If the instrument case of an air-driven gyro is not airtight, air will be
drawn into the case via the leaks, resulting in a loss of efficiency. This will result in a reduction
in the rotor speed and the pointer will indicate a lesser rate of turn; similarly, if the speed is too
high, the pointer will indicate a higher rate of turn than that being flown. The most likely fault is
a rotor speed falls below the design RPM, which will result in both the gyro rigidity and the
precessional forces being reduced. Of these, the reduction in the precessional forces is the
most important as they will no longer be able to overcome the spring tension to the same
degree. The Turn indicator will therefore under read. In effect the following rule is easy to
remember and summarises this:

Under speed of the rotor under indicates the rate of turn.

Pre-flight Check
If the indicator is air driven approximately five minutes should be allowed for the rotor to reach
its operating rpm prior to taxiing. With the aeroplane still stationary on the ground the turn
pointer should be aligned with the zero datum, but during taxiing for take-off the pointer should
respond accordingly to left and right turns. In most light aeroplanes applying hand pressure to
one corner of the flight instrument panel will also enable the turn indicator to be checked. This
is because the panel is normally fitted on shockproof mountings, and any movement results in
the turn pointer indicating a momentary rate of turn.

Construction and Operation of the Balance Indicator

This part of the instrument uses a mechanical method to indicate that an aeroplane is correctly
banked for a given rate of turn. It uses the force of gravity, which acts upon a black ball in a
liquid filled glass tube, and maintains it in its true vertical position whilst the aeroplane is in
straight and level flight, as shown below.

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The liquid acts as a damping medium for the ball, and two expansion chambers are concealed
behind the dial, to cater for temperature changes. The back of the tube is painted on the
outside with fluorescent paint to provide a contrasting background for the black ball, and the
whole assembly is firmly secured to the back of the dial by a bracket. The ball itself has weight,
and is thus affected by aeroplane manoeuvres.

If the ball remains in the centre the turn is balanced, and no slip or skid is present, as shown in
diagram (A) below.

Diagram (B) shows the aeroplane making a left turn at a certain angle of bank. During this
manoeuvre the indicator case and scale will both move with the aeroplane. The ball is
additionally subject to a centrifugal reaction, since the aeroplane is in a turn, which will force the
ball away from the centre of the turn. If the turn is however carried out with the correct angle of
bank the two forces will be in balance, and the ball will remain in the zero position. Any increase
in airspeed during the turn will increase both the bank angle and centrifugal force. The ball will
continue to remain in line with the resultant of the two forces, as long as the bank angle is
correctly maintained.

If the angle of bank for a particular rate of turn is incorrect, for example the aeroplane is under
banking, as shown in diagram (C), the aeroplane will tend to skid out of the turn. This will occur
because the centrifugal force predominates, and the ball is displaced away from the zero
towards the outside of the turn. By comparison if the aeroplane is alternatively over-banked, ie.
the angle of bank is excessive for the rate of turn, as shown in diagram (D), the aeroplane will
tend to slip into the turn, since the force of gravity will now predominate, and the ball will move
away from its zero position towards the inside of the turn. If the aeroplane skids or sideslips,
the turn is said to be unbalanced, and if the ball remains in the centre, the turn is said to be
balanced.

Limitations and Errors Associated with the Balance Indicator

The balance indicator has no operational limitations, and is also not subject to any errors.

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Pre-Flight Check
With the aeroplane on level ground the ball should be in its central (zero) position, but during
any turns when taxiing, the ball will register a skid.

Electrically Driven Turn and Balance Indicators

The internal mechanism of a typical electrical driven variant is similar to that of an air driven
variant, as shown on the next page.

In this type it is important prior to flight to ensure that the ‘OFF’ flag has disappeared from view,
and during taxiing, the needle should indicate a turn in the correct direction, and the ball should
indicate a skid. The flag will come into view if the rotor is not at its operating RPM, ie. due to a
power failure, and that the instrument is unreliable.

Typical Indications on a Turn and Balance Indicator

The diagrams below show the indications that a typical turn and balance indicator would be
likely to show during different types of turn.

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Chapter 3.5
Turn Co-ordinator

Introduction
A turn co-ordinator is a development of the turn and balance indicator and is used in place of
such instruments in a number of small, general aviation aeroplanes. The primary difference is in
the position of the precession axis of the rate gyroscope. In this instrument the gimbal is spring-
restrained and mounted with its axis is at approximately 30° with respect to the aeroplane's fore-
and aft axis, as shown below.

This has the affect of making the gyroscope sensitive to movements about the aeroplanes roll
and yaw axes, ie.to banking, as well as turning

Principle of Operation
The turn co-ordinator integrates both the rate of roll and the rate of turn together. It shows the
pilot what the aeroplane is actually doing, and not what it has done by indicating the two rates
on a display like the one shown below.

The aeroplane symbol on the turn co-ordinator moves in the direction of turn or roll, which is
unlike the artificial horizon, where the symbol is fixed to the instrument case and the horizon bar
moves. "NO PITCH INFORMATION" is normally printed across the indicator scale, in order to
avoid confusion in pitch control, due to the instrument's similarity to an artificial horizon.

REMEMBER - THIS INSTRUMENT IS NOT AN ATTITUDE INDICATOR

If the wing of the aeroplane is lowered, even very slightly, the turn co-ordinator will immediately
show a deviation from straight and level flight, whereas the turn and balance indicator will show

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nothing until yaw is present. The turn co-ordinator will therefore anticipate a change from
straight and level flight, whilst a turn and balance indicator will measure a deviation, thus
enabling the pilot to anticipate the resulting turn. The pilot can then simply control the turn at the
required rate, which is indicated by the alignment of the aeroplane with the outer scale.

The gyroscope is usually powered by a DC motor, and rotates at 6,000 rpm, although some turn
co-ordinators are powered by an AC motor, which is supplied from a solid state inverter that is
housed within the instrument. The inclusion of Silicon fluid or a graphite plunger in a glass tube
is also often used in the instrument design, to assist in the damping out of any gyroscopic
movements.

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Chapter 4.1
Inertial Navigation System

Introduction
An Inertial Navigation System (INS) provides the aeroplanes velocity and position by
continuously measuring and integrating its acceleration. This system relies on no external
references, is unaffected by weather, and can operate during the day or night. All corrections
associated with the movement of the earth, and transportation over the earth's surface are also
applied automatically. The products of an INS are position (latitude/longitude), speed (knots),
distance (nautical miles), and other navigational information. The quality of information is
however dependent on the accuracy of the initial (input) data, and the precision with which the
system is aligned to True North.

An Inertial Navigation System is based on the measurement of acceleration in a known

direction, or along a sensitive (input axis), which is detected and measured by an
accelerometer. The output from the accelerometer is then integrated, firstly to provide the
velocity along the sensitive axis, and a secondly to obtain the distance along the same axis. The
process of integration is used because the acceleration is rarely a constant value.

For navigation in a horizontal plane, two accelerometers are needed, which are normally
aligned with True North and True East. The accelerometers are placed with their sensitive axes
at 90° to each other and these axes must be maintained in the local horizontal, in order to avoid
any contamination due to gravity. To keep this reference valid, the accelerometers are mounted
on a gyro-stabilised platform, which maintains the correct orientation during all aeroplane
manoeuvres.

The Principle and Construction of an Accelerometer

The principle of an accelerometer is the measurement of the inertial force that displaces a mass
when acted on by an external force (acceleration). In its simplest form an accelerometer
consists of a mass, which is suspended in a cylindrical casing in such a way that it can move
relative to the case when the case (aeroplane) is accelerated, as shown below.

In this system the final position of the mass is dictated by retaining springs, so that when the
aeroplane is travelling at a constant velocity the mass will be positioned in the centre of the
cylinder, but when the aeroplane accelerates the mass will be displaced according to the
magnitude of the force being exerted on the aeroplane. The final position of the mass is
controlled by the pull of the springs, and the displacement of the mass is proportional to the
acceleration.

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The most common type of accelerometer used in the INS, is the ‘Force Balance Accelerometer’,
that is based on the angular displacement of a pendulum, which will rotate about a pivot point,
as shown on the next page, when the aeroplane accelerates.

This type of accelerometer consists of a weighted pendulum, which is prevented from moving
from its central position by two force balance (restoring) magnetic coils that act on a permanent
magnet. With the outer case (aeroplane) at rest and horizontal, or moving at a constant velocity,
the pendulum will be in its central position, and no pick-off current will flow. If the aeroplane
accelerates or decelerates the inertia of the pendulum will cause it to lag behind, either to the
left or right of its datum, which will in turn be detected by the pick-off coils. The resulting current
will then be fed to the restorer coils, and will be of sufficient magnitude to return the pendulum to
its central position. The strength and direction of the current in the coils required to resist this
tendency will give a direct indication of the acceleration/deceleration being applied along the
accelerometer sensitive axis. This thus provides a measure of the rate of change of speed with
time, and is known as its output. Conversely any acceleration sensed perpendicular to this axis
will have no effect on the accelerometer.

The output signal from the restoring coils is relayed to an integrator, where it is multiplied by
time to give a measure of the rate of change of distance with time, ie. Velocity. The resulting
output signal is then relayed to a second integrator where it is again multiplied by time to give
the distance travelled by the aeroplane. A single accelerometer can thus be used to determine
how far an aeroplane has travelled in a definite direction for a given time. By using two
accelerometers, one aligned along the North/South axis, and the other along the East/West
axis, it is thus possible to establish the new position of an aeroplane. The INS computer
processes the outputs from the accelerometers, and based on a known starting position of
latitude and longitude is able to calculate the distances travelled along each axis. For display
purposes, the distances are converted into changes of latitude/ longitude, and are applied to
give the aeroplane’s new latitude/longitude position. Changes of latitude (North/South direction)
in minutes of arc equate directly to the distance covered, but due to convergence, the distance
covered in the East/West direction has to be multiplied by the secant of the mean latitude to
derive the change of longitude.

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Performance
Accelerometers that are used in INS applications should meet the following requirements:

Sensitivity threshold: detect accelerations in order of 1 x 10-6 g

Sensitivity Range: be accurate over the range –10g to +10g
Input/ Output: tolerance of 0.01%
Scaling Factor: amplification of restorer current of about 5ma/g
Zero Stability (Null uncertainty) The perfect accelerometer will give a zero
output when the input is zero, but instrument
error may result in an output when the input is
zero. The null position should be within +/-1 x
10-4 g
Small and Light
Shock Loading: They must be able to withstand a shock loading of 60g,
and also have a low response to vibration

Operation of a Gyro-Stabilised Platform

For navigation with respect to the Earth, the accelerometers in an INS have to be kept in their
local horizontal position, and must also be aligned with True North. This is achieved by
mounting the accelerometers in the North and East directions on a North aligned gyro-stabilised
platform, which isolates the accelerometers irrespective of the aeroplane’s manoeuvres, or
changes of aeroplane direction. Rate integrating gyros are used as sensors to detect any
departure of the platform from its level and desired alignment. These gyros are extremely
sensitive and operate on the same principle as the gyro in a turn indicator.

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Three gyros are normally used; one to detect rotation about the North datum, the second to
detect rotation about the East datum, and the third to detect rotation about the vertical axis,
ie. they sense manoeuvres of the aeroplane in pitch, roll and change of heading (yaw). Any
rotational movement of the platform is detected by the gyros, which precess, and in doing so
generate a correction signal, that causes an electric current to flow to the appropriate torque
motor, and thus returns the platform to its original orientation, with respect to the earth. This
effect is virtually instantaneous, and keeps the platform in a stable condition.

To enable the rate integrating gyros to achieve the requisite high degree of accuracy, the rotor
and gimbal assemblies are immersed in a fluid, which tends to reduce any gimbal friction. A
typical gyro is shown below.

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Any torque (rotation) about the gyro’s input (sensitive) axis will cause the inner can to precess
about its output axis, ie. relative motion between the inner and outer cans takes place. Pick-off
coils sense this movement, and the resultant output signal is proportional to the input turning
rate. To avoid any temperature errors, the whole unit is closely temperature controlled.

If the platform is arranged as shown on the previous page, when the aeroplane is heading in a
northerly direction, the North gyro will be sensitive to roll and the east gyro will be sensitive to
movements in the pitch axis. Any yaw will additionally be detected by the azimuth gyro, and all
three of the rate integration gyros will act to turn their respective torque motors, to maintain the
platform’s alignment.

If the platform is alternatively arranged with the aeroplane heading in an easterly direction, as
shown on the next page, the East gyro will sense roll and the North gyro will sense pitch. For all
intermediate headings, the simultaneous action of the rate gyros/torque motors will be
computed, and the appropriate corrections applied. In summary, the platform isolates the
accelerometers from the angular rotations of the aeroplane, and maintains the platform in a
fixed orientation relative to the earth. The assembly of accelerometers, rate integration gyros,
torque motors, and gimbal system, are all mounted together on a platform, which is known as
the ‘stable element’.

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Setting-up Procedures
The accuracy of an INS is totally dependent on the alignment of the platform in azimuth and the
attitude of the stable element. The platform must be horizontal (level) and aligned to the
selected heading datum, normally true north, which must be maintained throughout the flight. It
is important that the levelling and alignment processes are carried out on the ground with the
aeroplane stationary.

Since the gyros and accelerometers used in an INS are normally fluid filled, it is necessary to
firstly bring the containing fluid up to its correct operating temperature before the platform is
aligned. The first stage in the sequence is thus a warm-up period, where the gyros are run up to
their operating speeds, and the fluid is temperature controlled. This phase normally takes 3-4
minutes, and whilst this is taking place the current position can be entered. The alignment
sequence can only begin when this sequence has been completed.

Levelling
Coarse levelling is achieved by driving the pitch and roll gimbals until they are at 90° to each
other, and aligned within 1°-2° in only a few seconds. The platform is roughly levelled using
either the airframe as a reference, or by using the outputs from gravity switches to operate
torque motors. ‘Fine Levelling’, follows this process where the earth’s component due to
gravity, is sensed by the accelerometers, and the resultant outputs are used to drive the

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appropriate torque motors until zero acceleration is achieved, which takes approximately
1-2 minutes to complete.
Alignment
Course azimuth alignment is achieved by turning the platform until the heading output agrees
with the aeroplane’s best known true heading, and is carried out during the coarse levelling
process. ‘Gyro Compassing’, or ‘Fine Alignment’, is automatically initiated once the platform
has been levelled. If the platform is not accurately aligned with True North (in azimuth), the
East/West accelerometer will sense an acceleration force due to the rotation of the earth. If the
accelerometer is alternatively lying with its sensitive axis exactly in an East/West direction, the
earth's rotation will have no effect on the accelerometers. If the alignment is not precise when
the INS is first switched on, the East gyro will detect a component of earth rotation, which is
used to torque the azimuth gyro until the platform is aligned, and the East/West output is
reduced to zero. Depending on the degree of mis-alignment, this process can take up to
10 minutes, and the aeroplane should not be moved until it is completed.

NOTE: The magnitude of the earth rate, which is affecting the East/West accelerometer is
dependent on 15.04°cos.lat°/hour, thus if the aeroplane is operating at high latitudes, this
component will verge towards zero, and will make any alignment to True North virtually
impossible. The effect of latitude on the fine alignment process will thus limit the initial alignment
of the platform to mid-latitudes and equatorial regions, so the usefulness of the North aligned
system is limited.

Levelling and Alignment

For a conventional gyro system the process of levelling and alignment will take approximately
15 to 20 minutes, although this time will vary from equipment to equipment. It is important that
the aeroplane is not moved during this process, because the resultant accelerations will upset
the system, and will prevent the platform from aligning.

This process also requires corrections to be applied, which are dependent on the system being
‘told’ the accurate value of the present latitude and longitude. The alignment process will be
unsuccessful if the wrong latitude is initially entered into the system, and if the error is small will
lead to poor alignment and degraded accuracy. If the latitude error is however large, the
platform will be unable to be aligned.

If the longitude is however wrongly inserted during the initialisation process it will not adversely
affect the alignment process but will affect the accuracy of any subsequent position. This is
because the INS is a direct reading system, and if you start with a wrong position all subsequent
positions will be wrong.

The inter-relationship between levelling and alignment is complex, and any slight discrepancy in
one will directly affect the other. It is therefore important that from the moment fine levelling is
complete, the necessary correction to keep the platform horizontal with respect to the earth is
applied. This is because the earth is continuously rotating, and with a gyro-stabilised device the
gyros will try to maintain spatial, rather than terrestrial rigidity. The platform is thus ‘tilted’ as the
earth moves round to maintain its position horizontal to the earth’s surface.

Corrections
Accelerometers and gyros have sensitive axes, which extend infinitely in straight lines; ie. they
operate with respect to inertial space. The earth is however not like that, because the local
vertical axes are not constant, since the earth is a curved surface, and it also rotates.

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Corrections for earth rate and transport wander therefore have to be made, as do the
accelerations caused by the earth's rotation. Any control gyro is rigid in space and, in order to
maintain an earth reference, it must thus be corrected for both earth rate and transport wander.
Further corrections for Coriolis (sideways movement caused by earth rotation except at the
equator) and the centripetal acceleration must also be made. The latter is caused by rotating
the platform to maintain its alignment with the local vertical reference frame.

The stabilised platform possesses the same characteristics as a single rigid gyro, and therefore
suffers from drift and topple, due to earth rate and transport wander as follows:-

Apparent Wander:

1) Apparent (Earth Rate) Drift: The azimuth gyro must be torque compensated for earth
drift, which allows for the familiar 15.04 x sin.lat°/hr.
2) Apparent (Earth Rate) Topple: The North gyro must be torque compensated at a rate
of 15.04 x cos lat°/hr.

Transport Wander:

1) Transport Wander Drift: Transport wander causes misalignment of the gyro input
(sensitive) axis at a rate, which varies directly with speed (along the sensitive axis) and
latitude. For a correctly aligned platform, the speed in an East/West direction is the first
integral of easterly acceleration, ie. the output of the east accelerometer. Latitude is
also calculated by the platform and, with these two values, the INS computer can
calculate and apply the correction for transport wander drift.
2) Transport Wander Topple: A stabilised platform, which is transported across the
surface of the earth, will appear to topple in both the East/West and North/South planes.
To keep the platform locally horizontal, transport wander corrections must thus be
applied to the pitch/roll torque motors by the appropriate amount.

Acceleration Corrections
To maintain earth orientation, the platform is slewed to maintain north alignment, and toppled to
maintain in the horizontal to the earth’s surface. During this process the accelerometers are
subject to centripetal and Coriolis effects, which will produce false outputs, and the INS
computer thus has to correct for these as follows:

Coriolis. This is caused by the sideways force, which affects the output of both the
North/South and East/West accelerometers, and is caused by the rotation of the earth
about its axis. An aeroplane following an earth-referenced track will follow a curved path
in space and the resultant tiny error will be computed. The necessary corrections will
then be applied to the outputs of the accelerometers.

Centripetal Acceleration. A body moving at a constant speed in a circle (such as an

aeroplane flying over the surface of the earth where the centre of the earth is the centre
of the circle) will have a constant acceleration towards the centre of the earth. This
acceleration will affect the accelerometers on an inertial platform and corrections to
compensate for this movement must be made, and applied to the outputs of the
accelerometers. The corrections made to the gyros and the accelerometers in an INS,
are summarised in the following table.

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AXIS GYROS ACCELEROMETERS

EARTH TRANSPORT CENTRIPTAL CORIOLIS
RATE WANDER FORCE EFFECT
NORTH Ω cos λ U -2 Ω U sin λ
- U tan λ
2
R R
EAST Nil -V U V tanλ 2 Ω V sin λ
R R
AZIMUTH/ Ω sin λ U tanλ U2 +V2 2 Ω U cos λ
R R
VERTICAL
V = Velocity North
U = Velocity East
λ = Latitude
R = Radius of Earth
Ω = Rotation of Earth (15.04° / hour)

Wander Azimuth System

The effect of slew on the accelerometer’s increases in magnitude towards the poles, to the
extent that a North aligned INS will not function properly at high latitudes. This is overcome by
not physically maintaining the system aligned to True North following the initial alignment
process. The alignment of the horizontal accelerometers is determined by calculation from the
sensed tilt outputs of the two gyros having horizontal sensitive axes. Using this technique, the
computer keeps a record of the platform azimuth alignment corrections and adds them to the
initial sensed alignment mathematically. The platform does not have to be physically rotated
and so the system becomes usable at high latitudes. In this system the INS computer
calculates the extent by which the platform has wandered from its North aligned position, and
accordingly corrects the outputs from the accelerometers.

The Schuler Tuned Platform

When a pendulous gyroscopic mass, such as an INS platform, is accelerated over the earth, it
will tend to oscillate, and as a result the outputs from the sensors will become inaccurate. Dr
Maximillian Schuler solved the problem by demonstrating that no oscillation will occur if the
platform exhibits the characteristics of an ‘Earth Pendulum’.

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Assume that the length of the pendulum is equal to the radius of the earth, and its bob (plumb
weight) is positioned at the earth's centre with its point of suspension at the earth's surface. If
the suspension point of such a pendulum were to be accelerated over the earth's surface, its
inertia coupled with the force of gravity would combine to hold the bob stationary at the centre of
the earth, and the shaft of the pendulum would remain vertical throughout. If the bob of an earth
pendulum were disturbed, as would be the case if the aeroplane was the suspension point, it
would oscillate with a period of 84.4 minutes, as shown on the next page.

Any errors experienced, as a direct result of Schuler’s oscillation will thus average out to zero
over the 84.4 minute cycle. For example if a ground speed error of 5 knots was generated, the
error would build up to 5 knots during the first 21.1 minutes, reduce to zero by 42.2 minutes,
increase to 5 knots in the opposite sense by 63.3 minutes, and return to zero again by 84.4
minutes.

In practical terms, this means that the output from the INS will be correct three times every
Schuler Period; once when the period starts and then again at the end. It will also be correct in
the middle, at 42.2 minutes. At 21.1 minutes the error will be a maximum high value and at
63.3, it will be a maximum low value. For example if an aeroplane is travelling at a real ground
speed of 350 knots and is subject to a bounded error of 5 knots when the platform is slightly
disturbed. The following table illustrates how the ground speed will vary during a typical Schuler
period.

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Period (min) 0 21.1 42.2 63.3 84.4

INS G/S (kt) 350 355 350 345 350

Errors
The INS has the following errors associated with it:-

Bounded Errors. These errors build up to a maximum positive and negative values,
and return to zero within each 84.4 minutes Schuler cycle. The main causes of these
errors are:-

a. platform tilt, due to initial misalignment

b. inaccurate measurement of acceleration by accelerometers
c. integrator errors in the first integration stage

This error thus averages out over the time period.

Unbounded Errors. The following errors exist:

Cumulative Track Errors. These errors arise from the misalignment of the
accelerometers in the horizontal plane, and the main causes of these errors
are:

1. Initial azimuth misalignment of the platform.

2. Wander of the azimuth gyro.

Cumulative Distance Error. These errors give rise to cumulative errors in the
recording of the total distance travelled. The main causes of this error are:

1. Wander in the levelling gyros. NOTE: wander causes a Schuler

oscillation of the platform, but the mean recorded value of the
distance travelled is increasingly divergent from the actual
distance travelled.
2. Integrator errors in the second stage of integration.

The sensitivity of any INS system is due to inaccuracies in the manufacture of

the rate integrating gyros and despite tight tolerances, less than 0.01°/hr is
normal, thus real wander is the main culprit of unbounded errors.

Inherent Errors. The irregular shape and composition of the earth, together with the
movement of the earth through space, and other factors provide further possible
sources of error. Such errors vary from system to system, depending upon the balance
achieved between accuracy on one hand and the simplicity of design, reliability, ease of
construction and cost of production on the other. The main cause of this error is:-

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Radial Error. This error provides an indication of the continuing serviceability

of the INS, and is derived from the following:-

Distance of final ramp position from INS position (NM/hour)

Radial Error =
Elapsed time in hours
When calculating the distance between two positions; the latitude must be
considered. It is also good practice to monitor the performance of the INS at the
end of each flight. Comparing the INS display with the ramp position does this,
where the radial error is established using Pythagoras, which when divided by
the recorded time in the navigation mode, gives the average error.

Overall minimal inbuilt errors exist in the INS, and according to the
manufacturers 1 nautical mile per hour is the maximum error that will exist.
This will however slightly increase during prolonged flights

The Advantages and Disadvantages of an INS

The following advantages exist:

¾ Indications of position and velocity are instantaneous and continuous.

¾ Self contained; independent of ground stations.
¾ Navigation information is obtainable at all latitudes and in all weathers.
¾ Operation is independent of aeroplane manoeuvres.
¾ Given TAS, the Waypoint/Velocity can be calculated and displayed on a continuous
basis.
¾ If correctly levelled and aligned, any inaccuracies may be considered minor as far
as civil air transport aeroplane are concerned.
¾ Apart from the over-riding necessity for accuracy during pre-flight data insertions
there is no possibility of human error.

The following disadvantages exist:

¾ Position and velocity information degrades with time.

¾ The system is not cheap, and is difficult to maintain and service.

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¾ Initial alignment is simple enough in moderate latitudes when stationary, but is

difficult above 75° latitude, and should not be carried out in flight.

Mode Selector Panel

The functions and information displayed in the INS are controlled through a mode select panel
(MSP), as shown on the next page.

The functions provided are:-

SELECTION MEANING

OFF Power Off.

STBY (Standby) Power ON

Power is supplied to the system, the gyros spin up and
the temperature of the oil is stabilised. The start
position is normally inserted as a latitude /longitude,
which is accurate to the nearest 1/10th of minute of arc.

ALIGN Automatic alignment begins, and the platform is

levelled and aligned. When the Gyro-compassing is
complete, the READY NAV light (green) will illuminate,
and will indicate that the NAV mode can be selected.
The aeroplane must not be moved during this process,
although the aeroplane can continue to be loaded as
the INS is not affected by vibration, gusts etc.

NAV When this mode is selected, the aeroplane is ready to

move. The NAV annunciator light will extinguish. This
mode may be heavily indented to prevent accidental
movement. This mode also requires the TAS to be fed
from the ADC.

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SELECTION MEANING

ATT REF This selects the attitude reference mode. In this mode
the INS provides pitch, roll and platform heading
outputs only. The Control Display Unit (CDU) does not
display this information and selecting this mode
disables the navigational capability of the INS for the
remainder of the flight. This is selected if the
Navigation mode fails, thus providing continual data to
the EFIS and AFCS.

BATT The illumination of this light (red) indicates that the AC

power supply has failed and that the INS is running
directly off the aeroplane battery (DC). This light will
remain on until the battery power is insufficient for the
INS unit to function. A further annunciator light will
also illuminate on the Control Display Unit (CDU).

Control Display Unit

The diagram below shows the principle controls and displays on the Control Display Unit (CDU).

The functions are as follows:

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SELECTION FUNCTION

TK/GS Track and Ground speed. Track is given to the nearest

tenth of a degree in the left-hand window, and the ground
speed to the nearest knot is displayed in the right-hand
window.
HDG/DA Heading and Drift Angle. True heading to the nearest
tenth of a degree is given in the left-hand window. The
drift angle is displayed in the right-hand window and is
preceded by L (left drift) or R (right drift).

XTK/TKE Cross Track Distance and Track Angle Error. The

cross track distance is the perpendicular displacement of
the aeroplane from the direct Great Circle track between
two selected waypoints, and is shown to the nearest tenth
of a nautical mile in the left hand window. The track angle
error is the angle between the current track made good,
and the desired great circle track between the selected
waypoints. The L or R preceding this value indicates that
the actual track is to the left or right of the desired track.

POS Latitude and longitude of present position. Latitude is

displayed in the left-hand window to the nearest tenth of a
minute of arc, and Longitude is displayed in the right-hand
window to the nearest tenth of a minute of arc.

ALERT Annunciator Comes on steady to warn the operator that the aeroplane
has 2 minutes to run to the next waypoint, and then goes
out when changing to the next track. If manual track
changing is selected, instead of the alert light going out, it
will start to flash for a new track to be inserted.

WARNING Annunciator Illuminates when there is a system malfunction.

BATTERY Annunciator Illuminates when operating on internal power.

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Intentionally Left Blank

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Chapter 4.2
Inertial Reference System

Introduction
Gyro-stabilised platforms are generally costly, heavy and require a lengthy alignment
procedure. With the introduction of relatively low cost, high performance digital computers,
these systems have been steadily replaced by mathematical computer software models, which
modify the output signals from accelerometers that are strapped directly to the airframe. This is
referred to as an Inertial Reference System (IRS), which works on the same fundamental
principles as the INS, and has the following functions:-

¾ It measures vector accelerations.

¾ it determines the horizontal components of these accelerations.
¾ it integrates the above to obtain vector velocities and distances.
¾ it adds the above results to a start position, to obtain the present position.

The fundamental difference between the INS and the IRS is that the latter is a ‘Strap-down
System’. The IRS senses the aeroplane’s displacement about 3 axes to provide:

¾ primary attitude.
¾ true and magnetic headings.
¾ vertical speed.
¾ aeroplane position relative to the earth.
¾ accelerations and angular rates.
¾ wind velocity and direction.
¾ ground speed.

Each IRS consists of three laser gyros, three accelerometers, power supplies, a
microprocessor, built in test equipment (BITE), and output circuitry. Three totally independent
IRS’s are normally installed on an aeroplane, and each receives barometric altitude, altitude
rate, and TAS data from the Central Air Data Computer (CADC). Coupled with the gyro and
accelerometer data the aeroplane’s vertical speed can be determined, and the wind parameters
calculated.

Description of the Strap-Down System

The strap-down system dispenses with the gimbal mounted stable element and instead uses
solid-state ring laser gyros (RLG). These gyros are not required to stabilise the accelerometers,
as in the case of an INS, but provide aeroplane orientation, by determining the rate of rotation
around each of the aeroplane axes. The orientation data is used to process (modify) the
accelerometer outputs to represent those, which, under the same conditions, would be the
expected outputs from the accelerometers, if they were positioned along the North, East and
vertical axes. The transform matrix (a Quaternion) is generated by digital computation, and
gives the analytical equivalent of a gimballed system.

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The sensitive axes of both the accelerometers and the RLG’s are strapped or fixed at 90° to
each other on the airframe, along the aeroplane’s principle axes. There is thus no isolation from
the aeroplane’s movements, and the outputs represent linear accelerations (accelerometers)
and angular rates (RLG’s), with respect to the aeroplane’s axes.

Solid State Gyros

Solid-state gyros have no moving parts, but can achieve a significantly higher level of accuracy
and serviceability than a conventional gyro. Two types of solid state gyro are currently being
used in commercial aviation applications; the Ring Laser Gyro (RLG), and the Fibre Optic Gyro
(FOG), which both operate on similar principles.

Ring Laser Gyro. Unlike conventional gyros that are maintained in a level attitude by a
series of gimbals, the RLG is fixed in orientation to the aeroplane axes. Any change in
orientation, as a result of an aeroplane manoeuvre, is sensed by measuring the
frequencies of two contra-rotating beams of light within the gyro.

In the IRS a triad of RLGs (orthogonal axes), with their sensitive axes positioned
mutually perpendicular is utilised. A block diagram of one of these is shown above. The
example shown has a triangular path of laser light, whose path length is normally 24, 32
or 45 cm. Other models alternatively use a square path, ie. one more mirror. The RLG
is produced from a block of a very stable glass ceramic compound, which has an
extremely low coefficient of expansion. The triangular cavity contains a mixture of
helium and neon gases at low pressure through which a current is passed. The gas (or
plasma) is ionised by the voltage, which causes helium atoms to collide with, and
transfer energy to, the neon atoms. This raises the neon to an inversion state, and the
spontaneous return of neon to a lower energy level produces photons, which then react
with other excited neon atoms. This action is repeated at speed and creates a cascade

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of photons, throughout the cavity, ie. a sustained oscillation, and the laser beam is
pulsed around the cavity by the mirrors at each corner.

The laser beam is forced to travel in both directions around the cavity. Thus, for a
stationary block, the travelled paths will be identical, and the frequencies of the two
beams will be the same at any sampling point. If the block is however rotated, the
effective path lengths will differ; one will increase, and the other will decrease. Sampling
at any point will thus give different frequencies, and if the frequency change is
processed, it will give both an angular change, and a rate of angular change.

By processing the difference in frequency between the two-pulsed light paths, the RLG
can be used as both a displacement, and as a rate gyro. There is however a limit of
rotation rate below which the RLG will not function: because of minute imperfections
(instrument error) in the mirrors. Consequently one laser beam can ‘Lock-in’ to the
other, and therefore no frequency change will be detected. If this occurs the RLG will
cease to be a gyro, which is equivalent to gimbal-lock in a conventional gyro system.
Using an AC piezo-electric motor, which operates at a frequency of approximately
350Hz, and gently vibrates or ‘dithers’ the complete block, prevents ‘Lock-in’ of the two
laser beams. The outputs of the RLG are digital, not mechanical, and the reliability and
accuracy should exceed those of a conventional gyro by a factor of several times.

Fibre Optic Gyros. Like the RLG, the FOG comprises of a triad of gyros, which are
positioned mutually perpendicular to each other, and similarly three accelerometers.
The FOG senses the phase shift proportional to angular rate in counter-directional light
beams travelling through an optical fibre. FOGs are dimensionally similar to RLGs,
although the FOG benefits from less weight and is overall cheaper. The FOG is
however not quite as rugged, nor as efficient as the RLG.

Advantages and Disadvantages of RLGs

RLGs suffer from the following advantages and disadvantages:-

The advantages of RLGs are:

¾ High reliability.
¾ Very low ‘g' sensitivity.
¾ No run-up (warm-up) time.
¾ Digital output.
¾ High accuracy.
¾ Low power requirement.
¾ Low life-cycle cost.

The disadvantage of RLGs is that they are initially expensive to buy.

Alignment of the Inertial Reference System

Although the assembly is ‘bolted’ to the aeroplane frame, a Inertial Reference System (IRS) or
RLG INS, still needs to be aligned to an earth reference. Instead of levelling and aligning as in
the case of a stable platform, the speed and flexibility of a digital computer allows a transform to

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be calculated and compiled. 'The transform is a mathematical solution as to where the

horizontal and True North lie with respect to the triad of RLGs and accelerometers. Full
alignment takes less than 10 minutes, at the end of which an offset to each output of the RLGs,
and accelerometers is established. This consequently allows the local horizontal and True North
references to be determined. The initial calculated values are applicable at that place, on that
heading, at that time. The earth will certainly move on, and if the aeroplane moves as well, the
vital references must be safeguarded. This is achieved by making sure that the NAV mode is
engaged. The complexities of 3-D motion, ie. the interactions of pitch, roll and yaw, require an
extensive mathematical and trigonometrical juggle to be quickly conducted.

This process is the reverse of the techniques used in a conventional INS, where instead of
creating a reference from a gimballed system, a reference is created from data taken from a
completely different set of values. If the aeroplane heading has not been altered since the IRS
was last used, then a rapid alignment taking approximately 30 seconds is possible.

Performance
The performance of an IRS (RLG INS) is generally slightly better than that of a conventional
INS, the principal advantage being reliability. The system has the following performance
criteria:

a. Position accuracy - 2nm/hr *

b. Pitch/roll - 0.050
c. Heading (T) - 0.40°
d. Groundspeed - ±8 kts
e. Vertical velocity - 30'/second
f. Angular rates - 0.1°/second
g. Acceleration - O.0lg
h. 95% probability, assuming no update with other navigation sources

The Control, Display and Output from an IRS

Control and Display of an IRS is very similar to that of the conventional INS, and a typical
master switch unit (MSU) is shown below.

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The outputs are displayed on a display panel and give the information selected by the crew,
with any changes being made using a menu driven keypad, as shown below.

Description of a Typical IRS

In a typical system magnetic variation between latitudes 73°N and 60°S is stored in the IRS
memory, and data corresponding to the present position is combined with the aeroplanes true
heading to determine its magnetic heading. The IRS is the aeroplanes normal sole source of
attitude and heading information, although a standby attitude indicator and standby magnetic
compass are still fitted.

The IRS outputs are independent of any external navigation aids, and in the normal NAV mode
provides the following data to the various aeroplane systems:-

¾ Attitude.
¾ True and magnetic headings.
¾ Acceleration.
¾ Vertical speed.
¾ Ground speed track.
¾ Present position.
¾ Wind data.

A block schematic of the overall input/output functions for a typical IRS is shown below.

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The IRS also integrates with navigation aids and other equipment on the aeroplane. A
multitude of data is additionally passed to the Horizontal Situation Indicator (HSI), which may be
either an electro-mechanical instrument on its own, or alternatively may form part of an
Electronic Flight Information System (EFIS) display.

The system normally consists of two independent IRS’s, which can operate on either AC or DC
power. If AC is not normal the systems will automatically switch to backup DC power from the
battery busbar. Backup power to the right IRS is also automatically terminated if AC power is
not restored within 5 minutes.

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IRS Transfer Switch

If either IRS fails, the IRS transfer switch is used to switch all associated systems to the
functioning IRS.

IRS Alignment
The alignment of the IRS is carried out on the ground as follows:-

Normal Alignment. Like the gyro stabilised INS the alignment of the IRS must be done
on the ground and the system initialised with the current aeroplane position, before it
can enter the NAV mode. The position is normally entered on the POS INIT page of the
FMC CDU during the alignment period, which like the INS must be carried out whilst the
aeroplane is stationary. The IRS display unit may be alternatively used to enter the
necessary data.

Alignment between latitudes 70°12’N and 70°12’S is normally initiated by rotating the
IRS mode selector switch from the ‘OFF’ position, directly to the ‘NAV’ position. The
IRS performs a short DC power test, during which the ‘ON DC’ light will illuminate, and
when it extinguishes the ‘ALIGN’ light will illuminate, indicating that the alignment
process has begun. The aeroplane’s present position should be entered at this time via
the FMC CDU, and after approximately 10 minutes the IRS will automatically enter the
‘NAV’ mode, at which time the ‘ALIGN’ light will go out.

At latitudes between 70°12’ and 78°15’ the mode selector switch must be left in the
‘ALIGN’ position for 12 minutes, and then be manually rotated to the ‘NAV’ position, at
which time the IRS will immediately enter the ‘NAV’ mode.

Fast Alignment. During transit or through flight stop-overs, with only short ground hold
over times, a 30 second realignment and zeroing of the ground speed error may be
selected by selecting ‘ALIGN’ from ‘NAV’, whilst the aeroplane is parked. The present
position should then be simultaneously updated by manually entering the current
latitude and longitude prior to reselecting the ‘NAV’ mode.

Loss of Alignment in Flight

If the alignment of the IRS is lost in flight due to the loss of DC or AC, or the mode selector
switch is moved out of the NAV mode detent, the position and ground speed outputs will be
inoperative for the remainder of the flight. If the selector switch is however rotated to the ‘ATT’
position it will allow the attitude mode to be used to relevel the system, and provide attitude
indications on the Attitude Director Indicator (ADI). In steady level flight (SLF) the levelling
process will take approximately 30 seconds to complete. The attitude mode may also provide
heading information, but in order to establish compass synchronisation the crew must firstly
manually enter the initial magnetic heading. All heading information will be invalid, and heading
flags will come into view until the actual magnetic heading is entered into the system.
Thereafter the IRS will be subject to drift (15.04° sin latitude), so when operating in this mode it
is important to periodically cross check and update the magnetic heading in the IRS against the
operating compass system as required.

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Chapter 5.1
Radio Altimeter

Introduction
The Radio Altimeter (RA) is a device, which accurately measures the height above the surface
immediately below an aeroplane up to 2,500 feet and is particularly suited to low altitude terrain
clearance measurement. It provides an instantaneous and continuous readout on the flight
deck of the height above water, mountains, buildings, or other objects on the surface of the
earth, but gives no information regarding high ground immediately ahead of the aeroplane. This
information is also supplied to the:-

¾ Automatic Flight Control System (AFCS) to facilitate automatic landings using the
Instrument Landing System (ILS).

¾ Ground Proximity Warning System (GPWS) to provide height, and rate of change of
height information.

The outputs from the Radio Altimeter can additionally be fed directly, or via a data bus, to the
Electronic Flight Instrument System (EFIS) and the Flight Management Computer (FMC).

Importantly the height being measured by the Radio Altimeter is absolute, so flight over
undulating terrain will result in sympathetic variations in the indications of the height of the
aeroplane on the display.

The Radio Altimeter System

A Radio Altimeter determines the time taken for a radio wave to travel from the aeroplane to the
ground directly beneath the aeroplane and back again. The system consists of a
transmitter/receiver, a modulator, an integral timing or beat frequency counter, a transmitter
aerial, a receiver aerial and a display as shown below.

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Some Radio Altimeter systems alternatively use a mechanical circular display, as shown below,
where the altitude is displayed linearly up to 500 feet and logarithmically from 500 - 2,500 feet,
which makes the lower range of altitudes easier to read more accurately.

In this type of display the maximum altitude (2,500 feet) is obvious, but it is not so apparent
when using a moving vertical scale presentation, as shown on the previous page.

Notably all radio altimeter displays have a method of setting the decision height, which is
normally set at 100 feet, and a flashing DH indicator light will be given when this point is
reached. The required height is set using a decision height (DH) setting knob, and a bug or
index marker indicates the set height. The setting control knob on some systems also normally
doubles up as a press-to-test (PTT) facility, which when engaged, will drive the display to a
predetermined value, which is typically 100 feet.

With reference to the upper display, an ‘OFF’ or ‘FAIL’ flag will be visible if:-

¾ There is a power failure.

¾ The returning signal is too weak.
¾ Local reflections are received from the airframe itself.

A mask will also cover the height pointer if:-

¾ The equipment is switched off.

¾ There is a fault in the transmitted signal.
¾ The altitude exceeds 2500 feet.

The Decision Height (DH) light will flash continuously if the aeroplane goes below the set height
and will remain so until the aeroplane climbs, or until the DH is set at a lower value. At
approximately 50 feet above the set decision height an audible alert will sound with increasing
loudness until the actual decision height is reached.

Principle of Operation of a Radio Altimeter

A Radio Altimeter measures the time taken for a radio wave to travel from the aeroplane to the
surface directly beneath and back again, and provided that the path followed by the wave is
vertical, the total elapsed time will be a function of the aeroplane’s height. During this time the
transmitted frequency will change and the equipment will measure the difference between the
transmitted and received signals. The frequency change is a measure of the time taken for the
radio wave to travel to and from the surface and thus, the greater the frequency change the
greater the height. To achieve this the Radio Altimeter system makes use of primary radar
principles and transmits a Frequency Modulated Continuous Wave (FMCW), at a frequency of
4250 MHz to 4350 MHz, which is in the Super High Frequency (SHF), or Centimetric
wavelength band.

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A complete modulation cycle or frequency sweep is illustrated on the next page. In this system
the total sweep of the modulated or carrier frequency is automatically varied by + 50 MHz
approximately 300 times per second, from an initial datum of 4300 MHz.

Throughout the cycle there will be two very short periods when the modulation changes from
positive to negative and vice-versa. The frequency difference, which occurs when the
transmitter alters the direction of its frequency sweep is overcome by relating the aeroplane
height directly to the average beat frequency, ie. the difference between the transmitted and
received frequency, observed over a short sampling period. The frequency changeover points
can essentially be ignored in the height calculation, so that the difference in frequency will be
directly proportional to the aeroplane’s height.

At low altitudes the reflected radio wave will return almost instantaneously, which will give an
erroneous height, so a wider sweep is necessary to provide a measurable frequency difference.
In order to overcome this ambiguity, the sweep rate is lowered, ie. the time for a complete
frequency sweep is made longer, so that all normal heights within the normal operating range of
the radio altimeter are covered.

Performance and Accuracy of a Radio Altimeter

The accuracy of the radio altimeter is normally:-

¾ 0 - 500 feet: ±3 feet or 3% of the height, whichever is the greater.

¾ Above 500 feet: 5 % of the height.

When the aeroplane is on the ground, the Radio Altimeter may show a small negative value,
since the equipment is normally calibrated to indicate zero when the main wheels first come into
contact with the runway surface on landing. This effect is particularly noticeable on aeroplanes
with multi-wheel undercarriage assemblies, which are inclined at an upward angle when
deployed in flight.

Errors Associated with a Radio Altimeter

A Radio Altimeter may be susceptible to the following errors:-

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Leakage Error. This may occur if the Transmitter (Tx) and Receiver (Rx) antennae on
the underside of the aeroplane are fitted too close together, ie. the spilling through of
the side-lobes directly into the Rx antenna. The antennae are thus placed far enough
apart to avoid any interference, which also provides adequate screening.

Mushing Error. This may occur if the antennae are placed too far apart. As the
aeroplane comes close to the ground, the Tx antenna, reflection point and Rx antenna
will form a triangle, so that the actual distance travelled by the wave can become
greater than twice the vertical height between the surface and the aeroplane, thus
giving a false height indication, as illustrated below.

The Advantages of a Radio Altimeter

Radio altimeters have the following advantages:-

¾ They indicate the actual (absolute) height of an aeroplane.

¾ They provide an easy crosscheck with the barometric altimeter for terrain
clearance.
¾ They provide a aural warning signal prior to reaching the preset DH, and a visual
warning when the DH is reached.

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Chapter 5.2
Electronic Flight Instrument System

Introduction
The Electronic Flight Instrument System (EFIS) is a highly sophisticated Flight Director System
(FDS), which uses cathode ray tube (CRT) or light emitting diode (LED) technology to provide
attitude and navigation information. This system replaces the electro-mechanical type of
instruments and provides the necessary high reliability for safe operations. In this system all of
the information is integrated into a single presentation and is placed in the flight crews preferred
line of sight. The information is clearly presented using colour symbols and is also easily
understood. Any relevant information can also be easily selected without having to scan a large
instrument panel.

EFIS Architecture
A basic EFIS layout is shown below.

The system comprises of:-

¾ 2 x Attitude Director Indicators (ADI).

¾ 2 x Horizontal Situation Indicators (HSI).
¾ 3 x Symbol Generators (left, right and centre).
¾ 2 x Mode Control panels.
¾ 2 x Light Sensors.

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The Attitude Director Indicator is alternatively known as the Primary Flight Display (PFD), and
the Horizontal Situation Indicator as the Navigation Display (ND). The EFIS is capable of
interfacing with multiple avionic, radio and navigation (long and short range) equipments, the
Thrust Management Control System, and also the Automatic Flight Control System (AFCS).

This system is also inter-linked with the Flight Management System (FMS), and provides flight
progress, and MAP displays. It is additionally linked to the Inertial Reference System (IRS) to
provide attitude and heading data, and the symbol generator to process, and display data.

1. Symbol Generator (SG). These components are a central part of the EFIS, and receive
inputs from various aeroplane sensors and avionic equipments. The data is then processed
and converted into suitable data for presentation on the ADI/ HSI. When the symbol
generators are powered they provide appropriate displays on these instruments, but the
displays can be interrupted if the following faults occur:-

¾ the screen will go blank if a power failure, over temperature, or failure of the
relevant symbol generator occurs.

¾ a partial loss of colour capability may cause an odd colour presentation, which may
be due to an over-temperature.

¾ when information is unreliable or the received radio signals are not received, the
display will disappear.

¾ If the aeroplane equipment fails a failure flag will be displayed.

2. Instrument Comparator Unit (ICU). This component detects data faults associated with
the ADI/ HSI and constantly monitors the following cross-cockpit displays for disagreement:-

¾ Attitude (pitch and roll)

¾ Heading
¾ Track

If conditions are detected outside set parameters, the appropriate Master Warning light will
illuminate, and an aural signal will sound. Instrument comparison monitoring is inhibited
when both pilots are using the centre symbol generator.

3. Compression Mode. On most EFIS’s, if either ADI or HSI screen fails, modified
information from the failed CRT will be automatically presented on the serviceable CRT in a
compressed format.

4. Temperature Sensing Units. These units are fitted to each ADI and HSI and are set to
‘low’ and ‘high’ values, which are spaced approximately 20°C apart. If the low value is
exceeded on the ADI the sky and ground shading will be switched off, thus alerting the pilot,
but will be automatically restored if the temperature drops. If the temperature however
exceeds the high value the whole display will switch off, but will similarly reset itself, if it
subsequently cools down.

5. Mode Control Panels (MCP). The ADI and HSI mode control panels administer the
symbology options, mode ranges, and the brightness of the respective ADI and HSI
displays. This panel also allows the radio altimeter decision height (DH) to be selected, as
shown on the next page.

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6. Light Sensors. These are fitted in close proximity to the displays and ensure that the light
intensity of the displays being selected by the flight crew is compatible with the ambient light
conditions on the flight deck.

Attitude Director Indicator (ADI)

The ADI is normally the upper of a pair of cathode ray tubes, which are sited in the instrument
panel in front of both pilots’. The information is displayed on approximately 16cm square colour
screens, and provides indications on the current:-

¾ Attitude (pitch and roll).

¾ Flight Director commands.
¾ Localiser and Glide slope deviations.
¾ Airspeed deviations.

The ADI additionally displays:-

¾ Auto-flight annunciations, eg auto-throttle.

¾ Ground speed.
¾ Angle of attack.
¾ Radio altitude.
¾ Decision Height (DH).

The primary part of the ADI display is the aeroplane’s attitude, which is supplied by the Inertial
Reference System (IRS), when it is aligned in the navigation, or attitude reference mode.
Attitude data is however unavailable during the pre-flight checks, and whenever the ‘ATT’ flag
appears.

Individual generators contained within the symbol generators produce the individual displays via
a scanning process. The coloured background is produced by ‘raster scanning’, which is the

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method by which electron beams travel back and forth across the screen to form a solid image.
The sky shading is cyan (light blue) in colour, whilst the ground shading is brown, and all other
symbols/characters are produced by a digital stroking technique known as ‘stroke scanning’.
A typical display is shown below.

Flight Director (FD) commands are presented on the ADI by conventional command bars, or in
some systems by V-bars. A failure in either axis will cause the respective command bar to
disappear, but if both axes are unreliable, both bars will disappear, and a FD flag will appear.
The FD bars will also not be displayed until the IRS has been properly aligned. The auto-flight
modes are usually presented in the select mode in white, and in the acquisition mode in green.
Additionally, if the normal data sources for the ADI are not available, alternate data can be
accessed via an Instrument Source Selector Panel. The following data is also displayed on the
ADI:-

Radio Altitude. Above 2500 feet the radio altitude display will remain blank, but
between 2500 feet and 1000 feet, the radio altitude will be digitally displayed on the
ADI. A white analogue ring scale and digital readout will replace the sole digital display
when the aeroplane is below 1000 feet. This ring is calibrated in 100 feet segments,
and steadily disappears with reducing altitude.

Decision Height (DH). This is normally displayed digitally, but from 1000 feet to
touchdown the selected DH is displayed as a magenta (pink) triangle on the circular
radio altitude display. During the descent, when the aeroplane is 50 feet above the
selected DH, an aural alert chime will sound at an increasing rate until the DH is
reached. At this point the ring and scale will change from white to amber, and the DH
marker will change from magenta to amber. The ring, scale, and DH marker will also
flash for a couple of seconds. The flashing can be cancelled by pushing a reset button
on the mode control panel, which will cause the scale and marker to return to their
original colours.

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Localiser and Glide Slope Indication. The ADI localiser deviation pointer sense is
reversed whenever the aeroplane's track is greater than 90° from the ILS forward
course. This ensures that the ADI deviation pointer is usable on a back-course
approach, and it also retains the compatibility between the HSI and the ADI localiser
deviation directions at all times. ILS Deviation Monitoring will also alert the flight crew of
any ILS deviations during an auto-pilot or flight director approach, when flying below
500 feet AGL. With the APP (Approach) mode selected on the MCP, if the glide slope
deviation is greater than one dot per second, or the localiser deviation is greater than
one fifth of a dot per second, the respective localiser or glide slope scale will
automatically change colour, from white to amber, and the pointer will flash. This alert
condition will cease when the localiser and/or glide slope parameters return to within
their normal limits.

The Horizontal Situation Indicator (HSI)

The HSI is a cathode ray tube (CRT), which is usually fitted beneath the ADI, and presents
plan view orientation navigational information to the flight crew. It also presents a selectable,
dynamic colour display of the flight progress. The display modes include MAP, PLAN, ILS and
VOR. Heading data is forwarded to the HSIs by respective IRSs (CAPT HSI - L IRS, F/O HSI -
R IRS), whilst the centre IRS is available as an alternate source. A typical normal HSI display is
shown below.

The HSI compass rose is automatically referenced to magnetic north when operating between
73°N and 60°S latitudes, with the NORM/TRUE switch selected to NORM, and to true north
when operating outside these latitudes. The compass rose may be also be referenced to true
north by manually selecting the NORM/TRUE switch to TRUE, regardless of the latitude.

TRU will be displayed at the top of the HSI, and will be enclosed by a white box if the HSI is
referenced to true north. When the HSI is referenced to true north and the aeroplane descends

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2000 ft at more than 800 fpm the box will change from white to amber. The box will then
continue to flash for 10 seconds, and thereafter remains amber. The box will return to white if
the aeroplane climbs 2000 ft at more than 500 fpm. A green box will be displayed around the M
for 10 seconds when the HSI is returned to magnetic referencing.

In the MAP mode, shown above, the HSI presents the following information against a moving
map background: -

¾ Heading.
¾ Routes.
¾ Curved trend vector.
¾ Range to altitude.
¾ Wind.
¾ Distance.
¾ Estimated time of arrival.

The respective Flight Management Computer (FMC) supplies track information, whilst the
opposite FMC is available as an alternate source. If the information from the FMC is unreliable,
an IRS will automatically provide the necessary information. Selected navigation data points
are also programmed into the FMC, so it is important to be familiar with the colours and
Symbology used. Purely for reference a comprehensive listing is provided at the end of this
chapter, although these symbols may vary from manufacturer to manufacturer.

The recommended display mode for most phases of flight is the MAP mode, as shown above.
Other available modes are as follows:-

PLAN mode. This display is presented, on the bottom 2/3 of the HSI against a static
map background, and shows the active route data oriented to true North.

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The top part of the HSI maintains a display of track and heading information as in the
MAP mode. Sequencing through the Flight Management Computer Control Display Unit
(FMC CDU) allows the active route to be viewed by the pilot.

VOR and ILS modes. When these modes, as shown below, are selected the HSI
presents expanded track scale and heading orientation. Wind information, and system
source annunciation is also provided with conventional VOR/ILS navigation information.
Conventional full compass rose VOR and ILS modes are also available.

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The HSI displays weather radar data in cyan, green, amber and red. Red depicts areas of:

¾ Greatest return intensity, and thus the highest risk area for intense turbulence

¾ Reductions in the return intensity are indicated by a change of colour from red to
amber, and then to green.

¾ Cyan and amber are additionally used for message displays.

¾ Weather radar data can only be displayed when the system is switched on and the
respective HSI is in the expanded VOR, expanded ILS or MAP mode.

Note. When Weather radar (Wx) is displayed in the VOR or ILS mode, the scale shown
applies only to the Wx display, and not to the deviation display.

Like attitude, heading/track data is unavailable until the associated IRS has completed its
alignment and has entered the navigation mode. HDG (heading) flags do not appear in this
case. In addition to previously mentioned EFIS failure indications, other discrepancy
messages can be displayed on the HSI. For example WXR/MAP RANGE DISAGREE,
indicates both Flight Management Computers, and the Weather Radar range, disagree with
the current control panel range data.

EFIS/IRS Interface
A block schematic of a typical EFIS/IRS interface is shown below.

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Electronic Flight Instrument (EFI) switching determines the Centre Symbol Generator (C-SG)
inputs and outputs. If the left EFIS switch is selected to Alternate (ALTN), the left system
instrument sources will supply the C-SG. When the First Officer (F/O) EFI switch is in the ‘ALTN’
position, the right system instrument sources will supply the C-SG. If both flight crew select
‘ALTN’ with their respective EFI switches, the left system instrument sources will supply data to
the C-SG. The C-SG will always use the centre ILS and centre radio altimeter. Each ADI and
HSI Control Panel is connected to the symbol generator via the EFI switch. Additionally each
IRS switch will permit flight crew selection of the alternate data source for heading/track, attitude
and vertical reference data.

Heading Reference Switch

This switch is normally sited on the centre instrument panel and permits the selection of a
magnetic or true heading reference for each Horizontal Situation Indicator (HSI), the Radio
Magnetic Indicator (RMI), the Flight Management Computer (FMC), the Auto-pilot Flight Director
System (AFDS) and the Flight Control Computer (FCC).

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The AFDS uses true heading only when the switch is in TRUE, but if TRUE is selected when the
AFDS is in the HDG SEL mode, it will automatically change to ‘HDG HOLD’. HDG SEL may
then be re-selected.

HSI Symbology
The following symbols can be displayed on each HSI/ ND depending on the switch selection on
the EFIS control panel. Symbols can be displayed with different colours but the general colour
presentations are as follows: -

GREEN (G). engaged flight mode displays, dynamic conditions.

WHITE (W). present status situation, scales, armed flight mode
displays.
MAGENTA (M) (pink). command information, pointers, symbols, ‘fly-to’
condition.
CYAN (C) (blue) non-active and background information.
RED (R) WARNING.
AMBER (A) or YELLOW (Y). cautionary information, faults, flags.
BLACK (B) blank areas, ‘off’ condition.

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Chapter 5.3
Flight Management System

Introduction
The Flight Management System (FMS), as shown below, is made up of the four primary
systems; a Flight Management Computer System (FMCS), an Autopilot/Flight Director System
(AFDS), an Autothrottle (A/T) and Inertial Reference Systems (IRS’s).

These components are all independent systems, and can either be operated individually or in
various combinations. Together these components provide continuous automatic navigation,
flight guidance, and performance management. The FMS is capable of four-dimensional
navigation (latitude, longitude, altitude and time), whilst optimising the aeroplane’s performance,

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in order to achieve the most economical flight possible. It also provides centralised flight deck
control of the aeroplanes flight path and performance parameters.

The primary function (flight management role) of the FMS is to provide performance
management, navigation guidance and automatic flight control. In this role, the FMS is
additionally interfaced with the engine ‘Power Management Control’ (PMC) and the Automatic
Flight Control System (AFCS). This isolates the flight crew from the control loop, and allows the
FMS to act in a totally integrated fashion. The FMS thus provides optimum control over the
aeroplane’s engine power settings, and total control over its flight path.

The secondary function (advisory role) of the FMS is to provide inputs to the various flight deck
displays in order to assist the flight crew in manually flying the aeroplane. For example it
provides the HSI map display (for orientation) and positions the bugs on the airspeed and EPR
(N1) indicators (to assist in manually flying precise flight profiles).

The FMS thus relieves the flight crew, so that they can attend more closely to the tasks of
monitoring and decision-making.

The Flight Management Computer System

The Flight Management Computer System (FMCS) is the heart of a modern aeroplane’s
electronic systems, and gathers information from other subsystems.

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The FMCS also allows the flight crew to manage the aeroplane’s lateral (L-NAV) and vertical
(V-NAV) flight path through a keyboard on the Command Display Unit (CDU). Computations
relating to L-NAV include items such as courses to be flown, ETA’s, and distances to go.
Conversely computations relating to V-NAV include items such as fuel-burn data, optimum
speeds and recommended altitudes. When operating in the Required Time of Arrival (RTA)
mode, the computations include required speeds, take-off times and enroute progress
information. The FMC additionally integrates the information that has been entered by the flight
crew on the CDU and the information stored in its memory with information fed from external
sources, eg. navigation data, air data and engine data. This information is used to calculate the
present position of the aeroplane from which pitch, roll and thrust commands can be derived,
which will in turn be used to optimise the aeroplane’s flight profile. From this information control
and guidance commands are also fed to the AFCS and Autothrottle. This allows for an
integrated FMS operation using both automatic lateral and vertical navigation, from the initial
climb to the final approach. Advisory information is additionally forwarded to various flight deck
displays, eg. the EFIS and the Mach/Airspeed Indicators. The FMC thus reduces the need for

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the flight crew to make calculations, refer to maps, and read manuals. The flight crew must
however still continue to monitor the FMC to ensure that it is following the planned route.

Command Display Unit

The Command Display Unit (CDU), as shown below, is the means by which the flight crew can
communicate with the computer.

An alpha/numeric CRT screen dominates the upper part of the unit, while the lower part has a
keyboard. Using the keyboard, the flight crew can enter the desired vertical and lateral flight
plan data into the Flight Management Computer. The parts of the CDU and their individual
functions are as follows:-

A – Line Select Keys. (Six Each side of screen) Push to select or enter data on the adjacent
line.

B – CDU Display. Displays the page of data selected by function, mode or Line Select key.

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C – Page Title. Indicates the type of data displayed. When in lateral or vertical navigation
modes displays ACT or MOD indicate the page status.

D – Page Number. The first digit represents the page number and the second digit indicates
the total number of related pages.

E – Boxes. Indicates the data required by the FMC for full navigation capability. Data will be
entered from scratch pad using Line Select key.

F – Dashes. Indicates that a data entry is required by the system. If known, the data should be
keyed in and transferred from the scratch pad using the appropriate Line Select key.

G – Line Title. Indicates the type of data on the line. This will be blank if data is not recognised
by the FMC.

H – Scratch Pad Line. This is the bottom line of the display and displays: -

¾ System generated messages to the flight crew.

¾ All keyboard entries (before they are transferred to the required line).
¾ Data being moved from one line to the other.

Control Panels
In a totally integrated system two further control panels exist, which are the:-

¾ AFCS Mode Control Panel (MCP)

¾ EFIS Control Panel (ECP)

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The diagram below illustrates how information flows to, within, and from the various FMS
components.

Most modern aeroplanes are fitted with two FMS units, which are usually arranged as illustrated
on the next page.

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Both FMC’s are cross-coupled, or matched as a dual system, so loading data into one CDU will
automatically feed both FMC’s. It is thus advisable to nominate one CDU for data loading while
using the other to monitor or cross check the data being loaded. In this arrangement each CDU
and EFIS map display will relate to its own computer, ie. the left FMC feeds the left CDU and
the left EFIS.

If the FMC switch is placed in the ALTN position, the related CDU and EFIS displays will be
connected to the other computer. This level of interconnection enables all displays to be
maintained, provided at least one CDU and one FMC is operating.

CDU and FMC Terminology

Active. This refers to flight plan information that is currently being used to calculate
lateral (LNAV) or vertical (VNAV) guidance commands. For example; the active
waypoint is the point the system is currently navigating toward; the active performance
(VNAV) mode is the climb, cruise or descent speed schedule currently being used for
pitch and thrust commands. In this mode ACT is displayed in the associated page titles.

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Activate. This refers to the process of designating one or two routes as active, and is a
two-step process. Firstly, push the ACTIVATE line select key for the desired routes, and
secondly, push the illuminated EXEC key.

Altitude Constraint. This refers to a crossing restriction at a given waypoint.

Cost Index. This is a figure that is determined by the operator and is used to bias the
computations for the speed schedule. It is based on a trade off to get the optimum
balance between minimum fuel and least time. The index is determined by dividing the
operating cost of the aeroplane by the cost of fuel, eg. If the fuel costs are high, the
number will be low. A cost index of zero will result in an economic speed, ie. the
aeroplane will be flown at its maximum range speed, and the higher the Cost Index the
closer the aeroplane will fly to VMO/MMO. Additionally, unless the Cost Index value has
been arrived at in a scientific manner, it will also only provide a means of choosing
appropriate climb, cruise and descent speeds

Econ. This refers to a speed schedule that is calculated to minimise the operating cost
of the aeroplane, and is based on the ratio of time costs to fuel costs. The economy
speed is based on a cost index that is entered into the FMC/CDU during pre-flight.

Enter. This refers to the process of typing or line selecting characters into the CDU
scratch pad line, and then line selecting the desired location for the data.

Erase. This refers to removing modified data from the system by pushing the line
select key adjacent to the word ERASE.

Execute. This refers to making entered data part of the active flight plan by pushing the
illuminated EXEC key.

Inactive. This refers to route, climb cruise or descent information that is not currently
being used to calculate LNAV or VNAV commands.

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Initialise. This refers to the process of entering data into the CDU that is required to
make the system operative.

Message. This refers to information that the system automatically writes in the scratch
pad to inform the flight crew of some condition.

Modified. This refers to active data that has been changed. When a modification is
made to the active route or performance mode, MOD is displayed in the associated
page title. ERASE also appears next to one of the line select keys and the EXEC key
illuminates. Pushing the ERASE line select key removes the modification, and pushing
the EXEC key will change the modified data to the active status.

Prompt. This refers to something displayed on the CDU to aid the flight crew in
accomplishing a task. Boxes or dashes remind the flight crew to enter information on
the associated line, or alternatively a word indicates what action is required next.

Resynchronisation. This is the automatic process of one FMC loading data into the
other, when a significant difference between the two FMC’s is detected.

Select. This refers to pushing a key to obtain the desired data or action.

Speed Restriction. This refers to a flight crew entered airspeed limit below a specified
altitude.

Speed Transition. This refers to an automatically entered airspeed limit below a

specified altitude.

Waypoint. This refers to a point in the route. It may be a fixed point such as a latitude
and longitude, VOR or NDB station, intersection on an airway, etc., or a conditional
point. An example of a conditional point is when reaching 1000 feet.

The Flight Management Computer Memory

The FMC storage is made up of three types of memory:

Erasable Programmable Read-Only Memory (EPROM) or Bubble Memory. This

holds the bulk of the navigation and performance data bank.

Non-Volatile Random Access Memory (NVRAM). This holds specific navigation and
performance data, which has been down loaded from the EPROM. Power for this
memory is provided from the computer power supply whenever power is applied to the
FMC. When power is removed, the memory elements are automatically switched to a
low-power standby state, which is specially designed for data retention.

High Speed Volatile Random Access Memory (HS RAM). This holds the operating
programme, which can be altered by CDU inputs.

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The FMC contains storage space for two data bases:

1. The Performance Database, which contains an average model of the aeroplane and its
engines. This model provides the FMC with the data required to calculate pitch and thrust
commands. Additionally, the relevant data can be displayed and this reduces the need for
the flight crew to refer to a performance manual during flight. The data stored includes:

¾ aeroplane drag characteristics

¾ engine performance characteristics
¾ maximum, optimum, minimum altitudes
¾ speeds and speed limits
¾ speed and altitude capability with one engine inoperative

The maintenance group for individual aeroplanes can refine this data, by entering correction
factors for drag and fuel flow, which are retained in the NV RAM for continued use.

2. The Navigation Database contains numerous elements of data, which relate to the normal
operational area of the operator and type of aeroplane. Each data package is originated in
the Flight Operations department and is loaded onto a magnetic tape. The data is then
loaded into the NV RAM, using a portable data transfer unit connected to the FMC.

There are normally two Navigation Databases – one active and one inactive. The active
part is effective until a specified expiry date, and the inactive part holds a set of data
revisions for the next period of affectivity.

To cover changes in navigational data and procedures, each navigation database is

renewed at intervals not exceeding 28 days, although the database holds up to 56 days of
data.

The navigation databases contain items such as:-

ITEM DATA HELD

Radio Aids Identifier; position; frequency; type of aid; DME elevation. VOR,
magnetic variation, ILS category and centre line bearing. The
maximum distance an aid can be tuned at normal cruise altitudes).

Waypoints ICAO identifier; type (en-route/terminal).

Position (latitude/longitude).

En-route Designator; outboard magnetic course.

Airways
Airports ICAO four letter identifier; position; elevation; alternates.

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ITEM DATA HELD

Runways ICAO identifier; number; length; heading; threshold position; final

approach fix (FAF) Ident. Any threshold displacement.

Airport ICAO code; type (SID, STAR, ILS, RNAV etc) runway number/
Procedures transmission; path and termination code.

Company Route Origin airport; destination airport; route number; details of SID,
route, STAR, approach.

NB. It is the flight crew’s responsibility, during the pre-flight checks, to ensure that the data
package in use is valid. CHECK THE DATE!!

At least one large aeroplane has crashed with a probable cause of an inaccurate data package
being loaded

General FMS Operation

In order to make full and effective use of the FMS the following rules apply:-

¾ To avoid errors, work in a slow deliberate manner while operating the CDU. Avoid
pushing more than one key at a time. Avoid entering data into both CDU’s at the
same time. Do not push CDU keys when the system is going through a
resynchronisation. Resynchronisations take about 30 seconds to complete. During
this time one map and CDU will show a failed condition, whilst the other CDU will
display the ‘RESYNCING OTHER FMC’ message.

¾ When selecting a CDU page read the page title to ensure the correct page appears.

¾ Check that the scratch pad line is blank before trying to enter data on the line. Use
the CLR key as required to blank the scratch pad.

¾ When entering data on the scratch pad line ensure that it is correct before
continuing with the procedure.

¾ Use care when pushing line select keys to ensure the correct key is being pushed.

¾ Confirm that data displayed on the CDU is correct before pushing the EXEC key. If
an error has been made, correct the erroneous data or push the ERASE line select
key, and then restart the procedure. Data cannot be entered on a blank line.

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¾ Messages that commonly indicate an error has been made are: - ‘NOT IN
DATABASE’, ‘INVALID ENTRY’ and ‘INVALID DELETE’.
Pre-Flight
During pre-flight, information from the flight plan and load sheet is entered into the CDU. This
information will define the starting point of the flight for initialisation of the inertial reference
systems, the desired route to the destination to initialise LNAV, and performance information to
initialise VNAV. If necessary, the CDU may also be used to modify the flight plan while in flight.

When electrical power is initially applied to the aeroplane, the CDU will display the appropriate
page for starting the pre-flight checks, as shown below.

After checking and entering the necessary data on each pre-flight page, the lower right line
select key should be pushed to select the next page. When ‘ACTIVATE’ is selected on the
route page the EXEC key will illuminate. The EXEC key should then be pushed to complete the
task of making the route active before continuing the sequence.

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If a Standard Instrument Departure (SID) has to be entered into the route the DEP/ARR key is
pushed. After selecting the desired SID the lower right line select key is again used to proceed
with the pre-flight sequence.

When the TAKEOFF REF page is reached it displays ‘PRE-FLT COMPLETE’, which confirms
that all of the required pre-flight entries have been made.

If the IDENT page is not displayed at the beginning of the pre-flight, the IDENT page can be
alternatively selected by pressing the INIT REF (initialisation) key.

En-Route
In-flight the CDU is used to modify the flight plan and display navigation and performance
information.

The first step is to select the appropriate page of data by pushing the Function / Mode key that
says what you want to do. For example, DIR = go direct; CLB = change climb conditions;
HOLD = enter or exit holding pattern, etc. Then, if required, enter the desired modification.

For example, to fly from the aeroplane’s present position direct to a waypoint, push the
DIR/INTC key and enter the waypoint in the box prompts that appear. Alternatively if it is to
answer a what if' type question, the page that displays the desired information can be selected
and the modified conditions entered via the keypad. The CDU will then display predictions of
what will happen if the modification is executed. The flight crew will then have the option of
either erasing or executing the modification.

Lateral Navigation (LNAV)

LNAV guidance outputs from the FMC are normally great circle courses between the waypoints
that make up the active route. If a procedure stored in the FMC database is however entered
into the active route the FMC will be able to supply commands in order to maintain a constant
heading / track, or follow a DME arc, as required to comply with the procedure.

The FMC determines the present position of the aeroplane using inputs from the IRS, DME,
VOR and localizer receivers. The FMS is certified to navigate accurately within a VOR/DME
environment, although its tolerance is tighter if it uses GPS updating. The FMC System position
may be solely based on IRS data, but if available, it will normally use DME or VOR/DME inputs
to refine and update the FMC Position. The FMC then uses its calculated present position to
generate lateral steering commands along the active leg to the active waypoint.

Whilst the aeroplane is on the ground the FMC will calculate its present position based only on
data received from the IRS systems, for which the FMC requires a present position input from at
least one IRS. Inertial systems will however accumulate position errors as a function of time, so
the position information being used by the FMC will also slowly accumulate errors. These
position errors can be detected by observing the position of the aeroplane on the HSI map. If an
extended ground delay occurs and a significant map error is noticed, the IRS should be
realigned, and the aeroplane’s present position re-entered.

While the aeroplane is in flight the FMC refines its position calculations based on inputs from
GPS, the three IRS, DME, VOR, and ILS. The refinement of position calculations is made by

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use of two DME stations, if available, or one DME and one co-located VOR. During an ILS
approach the position can also be refined by using the localiser signals.

The FMC additionally obtains basic position and velocity data from the IRS, which is then
checked against data received from the ground and/or space based sensors to determine the
IRS drift rates and offset corrections. Using all of this data the FMC is able to generate the Best
position every 5 seconds.

The FMC normally automatically tunes the VOR and DME to provide the best available signals
for updating the FMC calculated present position. The flight crew can alternatively select
frequencies manually and the FMC can continue to use the signals for position updating.

Vertical Navigation (VNAV)

Climb and Cruise. After takeoff, VNAV mode will engage after a thrust reference other
than take off is selected and the MCP altitude window is set to an altitude above the
aeroplane. Once VNAV is engaged, the MCP may be set to any altitude, even below
the aeroplane, without causing a level off. The VNAV mode will however disengage, if
the MCP altitude is intercepted before the aeroplane reaches its FMC cruise altitude.

The VNAV profile that the FMC commands, if not modified by the flight crew, is a climb
with the climb thrust set to achieve the airspeed limit associated with the origin airport
until above the limit altitude, and then a climb at the economy speed until the entered
cruise altitude is reached. During the climb, the aeroplane will remain within all altitude
constraints that form part of a SID that has been entered into the active route. The
aeroplane will then cruise at its economy speed until the top of descent point has been
reached, during which time the thrust will be limited to maximum cruise thrust.

If flying the climb speed profile would cause a violation of any altitude constraint the
‘UNABLE NEXT ALT’ message will appear on the CDU. The flight crew must then
manually select a different speed that provides a steeper climb angle.

Descent. When a (E/D) point is entered the FMC will calculate a descent path. (An E/D
is a waypoint altitude constraint that requires a descent from cruise altitude. The E/D is
normally entered on the legs page as a result of selecting a STAR or APPROACH).

Target speeds may also be changed by entries on the legs or descent page. Wind and
thrust assumptions may additionally be changed on the descent forecast page.

If the MCP is set to an altitude below the aeroplane, when the TOD point is reached the
FMC will command idle thrust and pitch to track the descent path. VNAV will
automatically disengage if the MCP altitude is reached before the lowest altitude
constraint. During the descent the MCP may be set to an altitude above the aeroplane
without VNAV disengaging or stopping the descent.

If an unexpected (not entered on descent forecast page) headwind is encountered, that

a significantly decreases the airspeed, the engine thrust will be automatically increased
to regain the target speed. If the autothrottle is not engaged, a ‘THRUST REQUIRED’
message will be displayed on the CDU. Conversely, if an unexpected tailwind is

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encountered, that significantly increases the airspeed, a ‘DRAG REQUIRED’ message

will alternatively be displayed. If the airspeed is limited the aeroplane will fly at the limit
speed, even if it must leave the computed flight path.

For VFR and non-precision approaches, the FMC computed path is designed so that
the aeroplane flies down to a point, which is 50 feet over the approach end of the
runway. It is then the flight crew’s responsibility not to descend below the Minimum
Descent Altitude (MDA) until adequate visual contact has been achieved. At the missed
approach point the vertical profile will initiate a climb to the missed approach altitude,
using climb thrust.

Operational Notes

¾ When operating in LNAV and VNAV modes, continue to monitor the system
operation for undesired pitch, roll or thrust commands. If undesired operation is
noticed switch over to the heading select and flight level change modes.

¾ The system should be carefully monitored for errors following the activation of a
new database, resynchronisation, power interruption or IRS failure.

¾ During twin IRS operation each FMC uses a different IRS for position calculations.
The IRS positions are not averaged as during normal operation, which can result in
a difference between the two HSI maps and the descent paths when radio updating
is not available.

¾ When operating significantly off the active route, the active waypoint may not
change as it is passed. When the LNAV mode is armed it can only capture the
active leg, and will not capture an inactive leg in the active route. The ‘DIRECT TO’
or ‘INTERCEPT LEG/COURSE TO’ procedures may by used to make the desired
leg active.

¾ When the same waypoint is used more than once in the route, certain route
modifications (such as ‘DIRECT TO’ and ‘HOLD’) will use the first waypoint.

¾ Some standard instrument departures contain a heading vector leg. These show on
the CDU LEGS page as a VECTORS waypoint, and on the map as a magenta line
leading away form the aeroplane symbol or waypoint. If the VNAV mode is
engaged, the ‘DIRECT TO’ or ‘INTERCEPT LEG/COURSE TO’ procedure may be
used to restore the waypoint sequencing.

¾ When entering airways into a route page the beginning and ending waypoint must
be in the database, otherwise the route segment must be entered as a ‘DIRECT’
leg.

¾ Occasionally a procedure in the database contains a hidden discontinuity that

appears on the LEGS page as ---- for the inbound course.

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¾ If an ILS procedure is entered into the active route, and it contains a leg to intercept
the inbound course, the LNAV mode will not sequence past the (INTERCEPT)
waypoint until the LOC. mode engages.

¾ If the engines are not shutdown on landing, a cruise altitude entry must be made
prior to the next flight to ensure that the vertical profile is rebuilt. If in descent and a
diversion to another airport is entered, a cruise altitude entry must be made to
rebuild the vertical profile.

¾ When operating outside the FMC navigation database area the following operating
characteristics will be noticed:

1. Origin, destination and runways cannot be entered into the route.

However, any origin that is in the database may be entered. An origin
entry is required for VNAV operation.

2. All waypoints must be entered as latitudes and longitudes.

3. The FMC will not use radio signals to update its calculated position and
will not tune the VOR or DME.

4. The HSI cannot display airports, navigation aids or waypoints that are
not in the route.

Fuel Monitoring
The FMC receives fuel data from the fuel quantity (totaliser) system and EICAS. The FMC
additionally calculates a separate fuel quantity, which prior to engine start is set to agree and
track the totaliser value unless the flight crew make a manual fuel quantity entry. When the FMC
receives a positive fuel flow signal (engine start) the calculated value is disconnected from the
fuel quantity system until the engines are shutdown after the flight. After start the calculated
value decreases at the rate the fuel flow signals indicate. The calculated value is displayed on
Progress Page 2. The calculated value is also displayed on the Performance Initialisation page
where it is labelled ‘CALC’, unless a manual entry of fuel quantity is made. In that case it is
labelled ‘MANUAL’.

If fuel is loaded after the FMC receives a positive fuel flow signal, the calculated value will not
include the new fuel loaded. This could occur if the engines are shutdown at one location, then
restarted to taxi to the fuelling location. Making a manual Fuel quantity entry on the
Performance Initialisation Page, followed by deletion of the manual entry can restore normal
operation.

The fuel flow signals are also used for calculating the fuel used by the engines. FUEL USED is
displayed on Progress page 2 and is reset to zero following a flight, before both engines are
shut down.

If the FMC determines a significant difference between the totaliser and calculated values the
‘FUEL QTY ERROR-PROG 2/2’ message will be displayed on the CDU scratch pad. The flight
crew may then select which value the FMC should use for fuel calculations for the remainder of

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the flight. The FMC will also continually estimate the amount of fuel that will remain on board the
aeroplane when it reaches the destination airport if the active route is flown. If the estimate is
less than the fuel reserve value entered on the Performance Initialisation Page the
‘INSUFFICIENT FUEL’ message will be displayed.

Flight Control and Management Summary

The FMS is capable of commanding the aeroplane along a pre-selected lateral (navigation) and
vertical path (performance). This takes place shortly after takeoff, and continues until the
system captures the localiser and glideslope. A typical FMC receives twenty-four digital inputs
and three discrete inputs, and outputs to nine different digital customers.

The FMC performs seven major functions (typical):

1. The input/output function of the FMC receives and transmits digital data to and from the
various systems on board the aeroplane, and checks that all received data is valid.

2. The CDU function of the FMC, formats updates and sends data to the CDU fix display, and
provides alerting and advisory messages to the CDU for display on the scratch pad.

3. The bit and monitoring function of the FMC performs a self-test of the FMC during power up
and upon request. It continuously monitors the FMS during normal operation, and any
failures would are recorded (on the memory disk) for retrieval at a later date.

4. The navigation function of the FMC houses the navigation data base, and is responsible for
computing the aeroplane’s current position, velocity and altitude. It also selects and
automatically tunes the VOR receivers and DME interrogators. The navigation function
computes the aeroplane’s present position by determining the distance to two auto-tuned
DME stations and GNSS, if installed. Positional information from the three IRUs is used to
solve any ambiguity that may occur, or as a prime source when the aeroplane is on the
ground. The aeroplane’s velocity is computed using IRU inputs, and altitude is computed
using both IRU and ADC inputs.

5. The performance function of the FMC computes performance parameters (limits) and
predictions for the vertical path of the flight profile, by utilising the performance database,
and the CDU input data.

6. The guidance function stores the active vertical and lateral flight plan input from the CDU.
The present aeroplane velocity and position information is then calculated by the navigation
function. The guidance function the compares the actual and desired position, and
generates steering commands, which are forwarded to the appropriate flight control
computer (FCC). Using the current computed vertical profile data from the performance
function, the guidance function also compares the actual and desired altitude, and also the
altitude rate. It then generates pitch and thrust commands, which are inputed to the
appropriate FCC, and the thrust management computer (TMC).

7. The EFIS function of the FMC provides dynamic and background data to the EFIS symbol
generator, and also provides the navigation function with a list of the closest NAV aid array
for auto tuning.

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Intentionally Left Blank

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Chapter 6.1
Flight Director System

Introduction
In a Flight Director system (FDS), Pitch and Roll commands are computed to achieve or
maintain a particular condition of flight, which is then applied manually by the flight crew. The
Flight Director is able to combine information from various sensor input sources, and display
them as a series of Pitch and Roll commands. This reduces the scan, simplifies the
interpretation and removes the need for the flight crew to analyse the inputs before making the
necessary control inputs. The commands from the Flight Director computer are also modelled
so that any command given will not unduly over stress the airframe.

Flight Director Architecture

The Flight Director System is made up of the components and displays illustrated in the
diagram below. It also shows the range of input sources and flows within the system.

Key: Flight Director Computer (FDC) Signal amplifier

Mode Selector Unit (MSU) Annunciator Panel

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Attitude Director Indicator (ADI) Horizontal Situation Indicator (HSI)

Flight Director Control Inputs

In its operation the Flight Director display receives Pitch and Roll Commands from the Flight
Director System Computer, which may form part of the Automatic Flight Control System, or may
be completely independent. The system typically receives inputs from the sources, which are
shown in the diagram below.

These sources are:-

Vertical Gyro. Pitch and Roll attitude are prime to the system and essential to all
computer modes. Any manoeuvre will require to be referenced to a pitch and/or roll
attitude. Any failure of the attitude reference will thus result in a total failure of all Flight
Director functions. The pitch and roll inputs may be derived from a Vertical Gyro unit, or
more commonly from the gyro stabilised Inertial Navigation System (INS), or from the
Inertial Reference System (IRS).

Compass. Magnetic Heading inputs are required for the heading hold and select
functions of the system. Compass information is also an essential input during the radio
modes of operation, where any heading deviation signals will be obtained from the
Remote Indicating (Gyro Magnetic) Compass (RIC) System.

Localiser/VOR. The VHF Navigation Unit provides inputs for the tracking of VOR
radials, and for the capture and tracking of the Localiser. The localiser inputs may be
used as a pure lateral navigation function or may be displayed in combination with glide
slope derived commands during the approach phase.

Glide Path. The Glide slope receiver provides inputs to the computer, equating to the
direction and degree of displacement from the glide path during the approach phase.
As a condition of the selected modes, when the pitch channels of the computer are
processing the glide slope deviations, the roll channels will also be generating localiser
commands.

Air Data Computer. The ADC provides signal inputs related to IAS, Mach, Vertical
Speed and Altitude. A Flight Director Computer may also generate commands based
on any, or all of these inputs.

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The Flight Director Computer (FDC)

The information derived from the various sources is processed, in a solid-state digital computer,
and depending on the setting on the Mode Control Unit (MCU), will provide the necessary
attitude and steering commands. These commands will then be fed to the displays, or
alternatively to the Automatic Flight Control System (AFCS).

Mode Control Unit

This unit allows the flight crew to communicate with the FDC, regarding the phase of flight, and
a typical example is shown below.

Flight Director Displays

In its simplest form, the display of the Flight Director consists of an indicator having Pitch and
Roll Command Bars or ‘V’ bars, as shown below, which are often combined with those of the
attitude indicator.

Movement of the Pitch Command bar up will require the aeroplane to be positioned by pulling
the control column back until the command bar is centred, as shown below.

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When the command bar is centred, it indicates that the required pitch attitude has been
achieved, and that the control column should similarly be centred at this point.

Conversely, movement of the roll command bar to left or right will require the aeroplane to be
rolled to the left or right, in order to centre the command bar, and to satisfy the computed
command, as shown on the next page.

By virtue of the fact that the command bar deflections are the resultant of a combination of
system inputs, they should not be regarded as positional to acquire a required state. They are
simply a command for change in Pitch and/or Roll attitude, eg:-

¾ Pitch Command Bar UP - Pitch Up

¾ Roll Command Bar RIGHT - Roll Right

Alternatively if the delta and WING type of display is used the following commands will be
given:-

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Flight Director Modes of Operation

The following modes of operation are available:-

Attitude Mode. With no other modes selected, this is the basic function of the system.
The command bars will be positioned to hold the pitch and roll attitude at the time the
Flight Director was switched ON. Any deviation from this datum will be sensed by the
Attitude Reference, VG, and will result in corrective Pitch and Roll Commands being
displayed.

In many systems if the aeroplane is in a near wings level state, typically 3° - 5°, the
Flight Director Command bar will call for wings level and from this point, datum to wings
level and Magnetic heading hold.

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Some Flight Directors have the facility where the command bars can be re-datumed by
the flight crew. This is achieved via a switch that enables the command bars to be re-
centred at a particular pitch and roll, or by a thumbwheel that that enables the pitch
command bar to be accordingly repositioned, as shown below.

The following explanations assume that the Attitude and Flight Directors are separate
instruments, but these are more usually combined together on the same instrument, as
an Attitude Director Indicator (ADI).

Heading. With Heading selected the Flight Director roll bar will command the
aeroplane to fly to the Pre-Selected Heading, and subsequently to hold that heading.
To achieve this the computer will receive inputs from the attitude reference source, the
VG, and the course from the compass system.

Either Pitch Attitude or Altitude Hold are compatible and available in the Heading mode,
and when the aeroplane is flying on the selected heading the command bars will be
centred, as shown on the next page.

If a new heading is subsequently selected, a heading error, ie. the difference between
the selected and actual heading, will position the roll command bar accordingly.

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As the aeroplane is banked to satisfy the command the attitude reference will input to
the computer a roll attitude signal, which will oppose the signal error. When the
established bank angle is appropriate to the heading change required, the heading error
and attitude signals will cancel. The roll command bar will then centre, and the control
wheel will be returned to its centre position.

With the aeroplane banked and turning the heading error will steadily be reduced, which
will result in the computer commanding opposite roll, in response to the attitude input, to
roll the aeroplane out onto the new heading.

With the aeroplane on its new heading, the command bars and control wheel will again
be centred, as shown on the next page.

The heading/bank angle commands, which are determined by the computer are limited,
primarily for passenger comfort, and typical maximum bank angle limits will be between
30° - 35°.

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Altitude Hold. If this mode is selected the attitude hold mode will automatically be
cancelled, and will provide appropriate pitch commands to hold the existing altitude.
This mode is compatible with all lateral command modes, and also with the Glide slope
mode, until the glide slope is captured.

Having selected this mode, any deviation from the datum altitude will result in pitch
commands being generated. In response to the pitch command, a pitch attitude signal
will be applied to the computer, and will oppose the altitude error signal.

Signal summation and the resultant command generated will be similar to that in the
heading mode.

A sophisticated Flight Director may have a number of Pitch modes, eg. IAS Hold, Mach
Hold, Vertical Speed, and some may have the facility to provide commands relative to
Vertical Navigation profiles generated by the Flight Management Computers. In
common however, they all allow only one pitch mode at any one time to be engaged.

Localiser/VOR (LOC/VOR). This selected mode is identified through the frequency of

the particular navigation aid that has been tuned on the VHF Navigation Control Unit.
VOR provides the facility for tracking a selected VOR Radial, although the word Course
is more appropriate since the Flight Director facilitates passage over the VOR station
and continued guidance relative to what would now be an Outbound radial.

In the VOR mode the aeroplane equipment will automatically sense any over-station
condition or ‘Cone of Confusion’, and will provide Flight Director Commands in a
memory capacity during the relatively short over-station condition.

The Localiser mode will provide for Capture and tracking of the Localiser signal. Apart
from VOR over-station consideration the principles of Localiser mode operation will
apply equally to VOR. In this mode the computer will construct its roll commands
around the Radio deviation signals that are generated as a result of the aeroplane being
to the left or right of the localiser. Importantly the computer will also sum the radio
deviation signals with the heading inputs that correspond to the Runway Inbound
Course or QDM. In the case of a VOR this will become the QDM, or the reciprocal
dependent upon whether the aeroplane is flying ‘To’ or ‘From’ the station. The
Heading input is necessary since there is NO content in the radio deviation signal that
tells the aeroplane how far it should fly to the left or right to reduce the error. In an
extreme case if an aeroplane were to be well to the left of the localiser it would fly to the
right in circles unless the radio deviation signal was co-ordinated with the heading.
Importantly all manoeuvres require to be co-ordinated against pitch and roll attitude.

The following diagram shows how the Flight Director roll commands are computed
relative to Radio deviation, Attitude and Heading. The arrows identify the sense of the
input signals to the computer.

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1. The aeroplane is wings level and is positioned to the left of the localiser,
which will result in the ‘Fly Right’ Localiser deviation signal giving a Flight
Director ‘Roll Right’ command.

2. As the aeroplane is banked to the right, the ‘attitude error’ and ‘radio
deviation’ signals will oppose each other. The roll command bar will thus
be centred, and the aeroplane will turn towards the runway centre-line.

3. As the aeroplane turns away from the runway ‘Magnetic Heading’ a ‘Roll
Left’ signal will be commanded, and the aeroplane will roll out until the
wings are level.

4. The aeroplane is wings level and approaching the runway at an optimum

angle. The ‘heading error’ and ‘radio deviation’ signals now act in
opposition to each other, and the roll command bar is centred.

5. The ‘radio deviation’ signal is slowly diminishing, and the predominant

‘heading error’ will result in the Flight director Computer commanding
‘Roll Left’ in order to reduce the relative angle to the runway centre-line.

6. The aeroplane will continue to roll out onto the centreline using co-
ordinated radio deviation, roll attitude and heading signals.

7. The aeroplane is now on course and lined up with the runway centre-line.

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Glideslope (GS). This mode provides for the capture and tracking of the
Localiser/Glide slope, when the UHF Glide slope frequency is paired with the Localiser.
With the mode selector switch in the Glide slope detent, and the ILS frequency tuned,
Localiser and Glide slope deviation signals will be directed to the computer.

The heading mode may be used to set up an initial approach to the Localiser beam, or
Altitude Hold may be used as a pre-capture mode for the glide slope. When the
aeroplane is within the capture area for the localiser the mode selector will be
positioned to Glide slope. The Flight Director will now generate the necessary roll
commands and once the Glide slope is captured the Altitude Hold mode will be
automatically disconnected. The diagram below shows the typical command indications
when the Glide slope mode is selected.

1. The start of the Capture phase. ‘Altitude Hold’ disconnects and the
computer disregards the initial ‘Fly Up’ signal, thus preventing commands,
which would call for the aeroplane to ‘Pitch Up’ to acquire the Glide path.

2. The aeroplane is on the Glide path, with zero Glide slope deviation, and no
pitch attitude error present.

3. The aeroplane passes through the Glide path and the computer responds
to the Glide slope ‘Fly Down’ by commanding ‘Pitch Down’. As the
aeroplane assumes a nose down pitch attitude, it results in an increased
‘Pitch Attitude Error’ signal, in opposition to the Glide slope signal. An
Integrator circuit within the computer then builds up a signal in opposition to
the VG ‘Pitch Attitude’ signal, and pitch commands are displayed in order
to enable the aeroplane to be positioned on the Glide path.

4. The aeroplane is now on the Glide path with a slightly nose down pitch
attitude.

Go-Around (GA). This mode is selected for a ‘go around’ after a missed approach,
and the computer will signal the ‘command bars’ to instigate a wings level pitch up

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attitude. This mode is also automatically selected if the GA button on the throttle is
activated.

Mode Annunciator
This unit provides a visual reminder of the mode selected by the flight crew, and is usually in the
form of a panel showing a number of coloured lights, with one light for each mode. When a
mode is selected, the appropriate light illuminates.

The Mode annunciator’s are sometimes solely dedicated to the Flight Director, whilst others
combine both the FD and AFCS modes, with the FD annunciator’s normally sited on the left
hand side, and the AFCS annunciator’s on the right hand side of the EADI.

Operation of the Attitude Director Indicator

The flight crew will select the pitch attitude for initial climb prior to take-off, by manually
positioning the command bars on the instrument. On take off, the flight crew will rotate the
aeroplane until the delta ‘fits’ into the command wings, at which point the climb attitude will
have been achieved.

As the aeroplane approaches the preset target, the FD command bars will indicate the
corrective action necessary in order to achieve the required transition to the new flight path. If,
for example, the aeroplane is to climb to a specified cruise altitude, when the altitude acquisition
signal is sent to the flight director’s computer it will generate a ‘Fly Up’ command. The flight
crew will then fly the attitude that will place the aeroplane symbol into the command ‘wings’. As

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the aeroplane approaches the target altitude the computer will generate a ‘Fly Down’
command. The flight crew will then reduce the pitch attitude of the aeroplane progressively until
the attitude for a level cruise is reached, as the designated altitude is reached.

If an essential input fails in flight it will be displayed to the flight crew by red flags, which are
easily identifiable by appropriate text as follows:-

¾ ‘GS’. This flag will drop over the glide slope scale if the glide slope signal fails or is
alternatively sufficiently weakened.

¾ ‘GYRO’. This flag will appear if the attitude information input fails.

¾ ‘COMPUTER’. This flag will appear if the system detects a failure in the computer output’s
or command signals.

The Horizontal Situation Indicator (HSI) Flight Director Commands

On a typical instrument a compass will dominate the instrument display, as shown below, which
is driven from the heading reference system (RIC or IRS), and rotates against a lubber line.

At the centre of the display is a fixed aeroplane symbol, and inside the compass ring is the
symbology, which represents lateral guidance signals, similar to the ADI display. The
symbology typically consists of:-

¾ a course arrow adjusted/set by the course set knob.

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¾ a course deviation bar driven by signals from the VOR/LOC receiver (as
appropriate)a course deviation scale on which each dot represents a displacement
of 1° from the selected course or localiser.

¾ a TO/FROM flag that indicates whether the selected course is extending TO or

FROM the VOR, which will disappear when a localiser frequency is selected.

¾ In the RNAV mode, each dot will normally represent a 1NM displacement, but in the
RNAV APR mode the scale will be altered so that each dot represents ¼ NM.

To the left of the indicator is a glide path scale, which replicates the glide path indications on the
ADI, and digital read outs are also provided. The range figure is determined from a DME while
the flight crew select the course. For intercept purposes the flight crew may also manually
select a heading, which will directly adjust the position of the heading bug on the compass card.

The warning flags on the HSI, like the ADI are red in colour, and are identified by appropriate
text, as follows:-

‘GS’. This flag will appear if the glide slope signal fails, and will cover any Glide slope
indications

‘COMPASS’. This flag will appear if the heading reference fails.

‘VOR/LOC’. This flag will appear if the VOR or localiser (as appropriate) signal fails.

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Intentionally Left Blank

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Chapter 6.2
Automatic Flight Control System

Introduction
An Automatic Flight Control System (AFCS) is a complex system, which is designed to relieve
the workload of the flight crew, and allow them to fully concentrate on the
management/monitoring of the flight. An AFCS thus provides the following functions:-

Stability and Control Augmentation. These two functions are linked and apply equally
to both manual and automatic flight control. Stability and control are closely allied to
each other, and one cannot exist without the other, eg. if the aeroplane stability is too
low, its controllability will be too high. The function of the AFCS is thus to ensure that
the correct levels of each exist.

Attitude Hold. This function ensures that the attitude of the aeroplane around selected
axes is maintained.

Flight Control. In this function the AFCS responds to externally sourced inputs such
as altitude, airspeed, heading, and navigational information. The autopilot when
operating in this mode will maintain a set condition, but will also respond to changes of
flight profile, eg. track guidance, automatic approach, flare and landing, which are
dictated by inputs from external sources. The external sources are not a part of the
AFCS, but must be fully integrated with it, if the system is to work effectively. The
diagram below shows the typical flow of information to the AFCS.

Classification of an AFCS
The degree of complexity of the AFCS fitted to an aeroplane is dictated by the:-

¾ size of the aeroplane.

¾ age of the aeroplane.
¾ length of the flights.
¾ intended route structure.
¾ complexity of the aeroplane.

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¾ number of flight crew.

¾ cost benefits achievable.
¾ safety benefits

An AFCS is also classified according to the number of axes around which it exercises control,
which are as follows: -

Single Axis. This type of autopilot normally only provides control around the roll axis,
using the ailerons for lateral control.

Two axes. This type of autopilot provides control around the roll and pitch axes, using
the ailerons and elevators respectively. Stability about these axes are the most basic
functions of an autopilot. It can also be used to supply information directly to the Flight
Director, to allow a pre-selected flight path to be followed manually, or automatically by
the autopilot system.

Three Axes. This type of autopilot provides control around the pitch, roll and yaw
axes, using the elevators, ailerons and rudder respectively. It is normally integrated
with a Flight Director and also provides sufficient control to carry out automatic landings.

Control Channels
The movement about each axis is controlled by a control channel, and these are classified as
the:-

¾ Pitch channel
¾ Roll channel
¾ Yaw channel

Inner Loop Control (Stabilisation)

Each axis control channel works on the same basic principle, so it is appropriate to look in depth
at one channel only, eg. the pitch channel. The components of a basic single axis pitch control
channel are shown below.

Together these components form a ‘Closed Loop Control System’, or an ‘Inner Loop
Stabilisation System’. The functions of the individual elements are:-

Attitude Sensing. This is achieved by gyroscopes or accelerometers, which

sense attitude changes about the relevant axis.

Error Sensing. This is achieved by synchro-transmitters or E-I bars, which change the
attitude signal into an electrical error signal.

Signal processing. This is achieved by a discriminator and amplifier circuit, which

together process the electrical error signals, and provide an output to the servomotor
actuator.

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Servomotor Actuator. This device moves the control surface.

Position Feedback loop. This element advises the signal processor of any
control surface movement.

Operation of an Inner Loop Pitch Stabilisation System

During normal operation the autopilot system will maintain the aeroplane pitch attitude prior to
its engagement. If the aeroplane pitch however changes, as a result of some aerodynamic
disturbance or out of trim condition, an electrical pick off will carry the displacement error signal
to the signal processor. The error signal will be processed, and the demand will determine the
degree of corrective control input required to be forwarded to the servomotor actuator. The
servomotor will then move the control surface in the appropriate direction, and by the
appropriate amount. The signal processing is carried out in the AFCS computer, which is
electrically advised of any control surface movement via a feedback loop.

The position feedback signals are amplified, and the resulting signal will oppose the error signal.
As the aeroplane is returned to its original attitude an aerodynamic feedback signal will steadily
reduce the attitude error signal, and the position feedback signal will begin to dominate, thus
moving the elevator back into its former flared position.

Outer Loop Control

The primary function of the autopilot is stabilisation, which is achieved by ‘inner loop control’,
but by inputting signals from raw data {heading, airspeed, altitude, radio links, Lateral
Navigation (LNAV), Vertical Navigation (VNAV)} into the inner loop, the system can perform
other tasks. These data inputs are known as ‘outer loop control’, and are fed into the system
by the flight crew through an appropriate mode select panel. For example in a light twin-engine
aeroplane commands to roll or pitch the aeroplane are manually inserted via a Mode Control
Unit (MCU), as shown on the next page. When these controls are activated, a synchro
transmitter rotor, or potentiometer applies the appropriate command to the channel servomotor.

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In a modern transport category jet-engine aeroplane the outer loop signals are applied to the
system via a more complex Mode Control Panel (MCP), as shown below.

These outer loop signals are categorised into either roll or pitch modes depending on their
function:-

Roll Modes

Heading Hold. When this mode is engaged the autopilot on receiving a heading hold
error signal from a Remote Indicating Compass (RIC) or Inertial Reference System
(IRS), will control the bank angle to maintain the heading at the time of autopilot
engagement.

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Heading Select. When this mode is engaged the aeroplane will turn onto a heading,
which has been digitally inputted, or manually inputted by the movement of a heading
bug. The diagram on the next page illustrates how an aeroplane responds to a change
of heading from 090° to a new heading of 150°.

HEADING 90°

MAG
MAG
MILES No 1 MILES No 2
MILES No 1 MILES No 2
9 12 9 12
6

NE
6
15
3

W
15
3
18 21

HE
0

18 21
0
33

AD
33

IN
24
27 30 27 30
24

G
15
0°
MAG
MILES No 1 MILES No 2

15 18
12
21
3 6 9

24 2
7

30
0 33

DIAGRAM 19.21 HEADING SELECT MODE

VOR/LOC Capture. The VHF Navigation Unit provides inputs for the capture and
tracking of VOR Radials and also for the capture and tracking of the localiser. The
aeroplane is controlled towards the radio beam in the Heading Select mode until a
device called a ‘Lateral Beam Sensor’ checks the strength of the signal and initiates
the capture mode at a predetermined signal level.

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SEQUENCE OF OPERATION (AP IN CMD)

AP-FD VOR-MODE PRESENT A/C HDG : 090
(ARM, CAPT) PRESENT MODE : HDG SEL
3 1) DIAL TO DESIRED VOR FREQUENCY
2) DIAL TO DESIRED VOR COURSE (070)
3) PRESS V/L MODE PUSH BUTTON
8 9 0 6 7 8
2 4) V/L AMBER APPEARS
5) V/L AMBER CHANGES TO GREEN AND HDG SEL
DISAPPEARS
6) ROLL CMD BAR MOVES TO LEFT AND CENTERS
AS AIRCRAFT ENTERS TURN
CMD CMD
HDGSEL DURING TURN :
V/L V/L
BANK ATT 25°
BANK RATE 2.5°/SEC VOR

CAPTURE
070
POINT
5

090
20° F
O
ER
E NT M
C A
BE CWS IS AVAILABLE
ADI
DIAGRAM 19.22 VOR CAPTURE MODE

Lateral Navigation (LNAV). In this mode computed flight path information from the
Flight Management system (FMS) is fed into the autopilot and steers the aeroplane
along the designated route.

Pitch Modes

Altitude Hold. When this mode is engaged the aeroplane will remain at the altitude at
the time of selection. If it deviates from this altitude by more than a preset amount
(typically ± 200 to 300 ft) an altitude alert warning (aural and visual) will be sounded.

Altitude Capture (Acquire). When this mode is activated a pre-selected altitude when
climbing or descending will be captured, and the aeroplane will automatically level out
and maintain that altitude. A warning of altitude capture is given approximately 1,000 -
700 ft. prior to the preset altitude being reached, and stays annunciated until
approximately 300 - 200 ft. from this altitude.

40,000 FT ALTITUDE

CAPTURE MANEUVER IF ALTITUDE IS PRE-SELECTED

CAPTURE POINT

DIAGRAM 19.23 ALTITITUDE CAPTURE

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Vertical Speed (VS). When this mode is selected the nose of the aeroplane will be
pitched up or down in order to maintain a selected rate of climb or descent, and this
mode should be used with caution. It is essential to ensure that the engine thrust
setting is correct for the required rate of climb or descent, and that the aeroplanes
speed is constantly monitored, which is normally done in conjunction with the
Autothrottle system.

Glideslope (G/S). When the approach (APP) mode is selected, and the glide slope is
captured, the autopilot will respond to the glide slope beam, and will automatically steer
the aeroplane down to the runway threshold.

Vertical Navigation (VNAV). When this mode is selected, like in LNAV, the aeroplane
will fly a computed flight plan in the vertical sense. The VNAV flight plan can be
connected to the autopilot and/or Autothrottle.

Flight Level Change (FLCH) (Speed Mode). This mode is a ‘pitch speed’ or
‘manometric’ lock’, which enables the autopilot to climb or descend the aeroplane at a
selected speed, to a selected altitude, and may be either IAS or MACH defined. This
mode requires pressure (manometric) data inputs, which are normally derived from the
Air Data Computer (ADC).

Combined Roll and Pitch Modes

Go-Around (GA). During an automatic approach the flight crew can elect to abandon
the approach, and climb out with the wings level. Depending on the aeroplane type the
GA mode may be Flight Director and/or autopilot operated. This mode can only be
initiated if the auto-approach mode is operative, by either operating the Take-Off Go-
Around (TOGA) switch on the throttle lever, and/or by moving the throttle levers to the
fully forward position. When the TOGA switch is pushed the throttles will automatically
advance, and the aeroplane will climb out at maximum safe rate.

Control Wheel Steering (CWS). If this mode is selected the flight crew will be able to
control the aeroplane in pitch and roll using normal control forces on the control column,
through the autopilot. When the control wheel is moved the aeroplane will take up the
new attitude, and on releasing the wheel, the new attitude will be maintained. In this
mode the pitch and roll forces applied are detected by force transducers, which are built
into the control column. The transducers may be in the form of piezo crystal elements,
which vary their electrical resistance when they are put under pressure. The generated
signal outputs, which are proportional to the input forces, are then amplified and fed as
output signals to the appropriate control channel where the controls are activated in
proportion to the applied signal.

Touch Wheel Steering (TWS). If this mode is selected via the TWS switch, the
appropriate control channels and servos will be temporarily disengaged. The flight crew
will then be able to manually manoeuvre the aeroplane to a new attitude and heading.
On releasing the TWS switch, the autopilot will be automatically re-engaged, and the
new attitude will be maintained.

Turbulence Penetration. Flight in turbulent conditions can impose heavy structural

loads on an aeroplane. In these conditions it is thus normal to disengage the autopilot,
and fly the aeroplane manually. In some aeroplanes turbulence penetration can be
alternatively selected, which reduces the gain on the pitch and roll channels, and
softens the AFCS’s response to turbulence.

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Attitude Sensing
This is achieved by using either a rate sensing gyro, or a vertical axis (two degrees of freedom)
displacement gyro, which are positioned as shown below, along each aeroplane axis.

Rate Sensing Gyro. When a gyro is aligned with its sensitive axis parallel to an axis, it
will detect any rotation of the aeroplane around that axis, and will also be able to
determine the rate of rotation being experienced.

Displacement Gyro. In this system the attitude of the aeroplane is monitored against a
vertical gyro (VG) reference unit. When a displacement is sensed the signal processor
will determine the magnitude of the displacement, and the corrective control input
required as compensation for the displacement. The vertical reference can be taken
from an integral gyro source, or may be alternatively taken from an independent source,
such as a remote Vertical Reference Unit (VRU), INS or IRS.

In many systems, the detector signal is derived from a combination of both rate and
displacement signals. This reduces the time delay incurred by deriving a rate from a
displacement signal, and also has the advantage of ‘damping’ the tendency to overshoot the
correction, which is a common problem with the displacement type of system.

The pitch and roll channels are normally operated from the VRU, but the yaw channel requires
an input from a horizontal axis tied gyro.

The AFCS Computer (Signal Processor)

The function of this component is to process the displacement signal and determine the amount
of control movement required to counteract the displacement. It must also monitor the feedback
to ensure that the required control has been activated, and that the desired effect is being
achieved.

The computations are very simple in an attitude hold system, but are extremely complex in a full
multi-mode AFCS. The computations and processes carried out in a computer are:-

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Amplification. This boosts the power of the detected signal to a sufficiently

high level to act as an output.

Integration/Differentiation. These are mathematical processes, which are used to

derive information, such as attitude change from rate of
rotation, or vice versa.

Limiting. This restricts the range of any parameter change, eg. Pitch rate, to a
specific limit.

Shaping. This modifies the computer output so that the required flight path or
handling characteristics are achieved.

Programming. This is the individual processes, which are designed to instruct

the aeroplane to follow defined manoeuvres.

The degree of computer power varies according to the role and complexity of the AFCS. In all
of these computations however, ‘Control Laws’, are used to determine exactly how the control
demands are translated into control movements. The ‘C’ control law for example is commonly
used in large transport aeroplanes in the pitch axis. This law provides stability of the aeroplane
at a selected flight path angle, and also compensates for problems associated with flight at low
airspeeds.

These computations also consider the airframe load factors, which are likely to be imposed on
the aeroplane during a manoeuvre. These are a function of the airspeed (dynamic pressure)
and thus require an input from the ADC, or other dynamic pressure source. On older or simple
autopilots there is no allowance made for ‘gust loads’, which can exist during flight in turbulent
conditions. If these conditions prevail with this type of autopilot the system should be
immediately disconnected, and the aeroplane flown manually. In more modern and complex
systems a function switch is available, which, when operated, ‘softens’ the control demands,
thus enabling the aeroplane to ride with the gusts. This is achieved by increasing the ‘limits’
argument in the computations. The switch should however only be activated in turbulent
conditions, as it will desensitise the autopilot.

Servomotor Actuators
The following types of actuators are available, but in their design, consideration must be given
to the balance between the range of control surface movement, and the rate of movement (in
the event of a failure occurring), the normal rate of movement, and the magnitude and accuracy
of any movement for control and/or stability:-

Electro Mechanical Actuator. These may be either DC or AC powered. In the DC

system a motor is coupled to the flight control via an electro magnetic clutch and a
mechanical linkage. A feedback is provided from a potentiometer driven by the motor. In
the AC system, the motors used may be either of the hysterisis type, or of the two-
phase induction type. A synchro transmission system provides information on the
position of the control surface whilst a tachogenerator provides the necessary feedback
to the signal processor.

Electro Pneumatic Actuator. A typical actuator of this type is shown below, where the
valve assembly is operated electro mechanically by signals from the AFCS computer.

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The necessary power is supplied from a pneumatic feed taken from a source such as an
engine driven pump, or a compressor bleed from a jet engine. When a command input is
received the opening of one valve is increased, whilst the other is reduced. This produces a
differential pressure in the two cylinders and results in a differential motion of the two rods,
which in turn will cause the output linkage to rotate, and a resultant control movement to
occur.

Electro Hydraulic Actuator. In most modern transport aeroplanes the primary flight
controls are operated through Power Control Units (PCU), as shown on the next page, which
utilise the muscle of hydraulic power to activate the control surfaces. These units respond
directly to signals from the AFCS computer and do not require independent servo actuators.
The signal from the AFCS computer is then fed to a solenoid, which operates a ‘Transfer
Valve’ within the hydraulic system. This activates the control surface and a position
transducer provides a feed back to the computer.

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Autopilot Terminology
The following terminology is associated with modern complex autopilot systems:-

Rate Damping Systems. These systems having more to do with stability than control.
They will not return an aeroplane to a specific attitude, but will prevent unwanted
divergence rates from developing, and will also help to smooth rate demands
commanded by the flight crew.

System Protection. This system prevents control surface runaway, or any other
undesirable malfunction, by limiting the actuator authority.

Comparators. These devices compare the outputs from both the sensors and the
actuators. If the attitude change being sensed is in the same direction as the actuator is
moving the control surface, the comparator will automatically disconnect the circuit.

Rate Trigger Systems. The characteristics exhibited by an aeroplane during a system

runaway are very marked, and are significantly different from those expected in normal
flight. This system will automatically disconnect the autopilot if the rate exceeds a set
threshold.

Simplex System. This is a single automatic control system, which is allied to a number
of sub-channels. Some of the components may be duplicated, although a single failure
elsewhere will render the system unserviceable. This type of system is alternatively
known as a ‘Single Non-Redundant System’.

Multiplex Systems. These are systems, which comprise of two or more independent
simplex systems, and sub-channels, such that in the event of a failure of a system or
sub-channel, the remaining system will be capable of performing the controlling
function. The number of systems and sub-channels are qualified as, ‘Duplex, Triplex
and Quadruplex’.

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Duplex Systems. These systems incorporate two complete control systems (Lanes)
for each channel. They use independent sensors, computers, actuators and actuator
power supplies. The two control systems are connected to the same control surface,
and in the event of a runaway in one lane, the other lane will sense a disturbance and
apply a corrective actuator input. The speed of response of this system (to a runaway)
can be improved by using a comparator system, which continually monitors the
positions of the individual actuators, and if one is detected to be diverging at a
significant (pre-set) rate, that channel will be automatically disengaged.

Triplex Systems. These systems are an extension of the duplex system, and utilise
three lanes on a control channel rather two. Monitoring of all three lanes is conducted
continuously and a comparator circuit is fitted to detect a divergence of any lane. Voting
then takes place within the system to decide which lane is divergent, and that lane is
then closed down, thus downgrading the operational channel to a duplex system. This
system is therefore termed a ‘Single Failure Survivable system’.

Quadruplex Systems. This system uses four lanes per channel, and therefore
provides additional redundancy.

Model Following. This system is not used alone, but is used in association with
another systems, such as a Duplex system. It is basically a software fix in which the
flight characteristics of the aeroplane are programmed into the AFCS computer. The
computer then uses this data to determine the anticipated response for a given control
demand and compares this with the response detected by the sensors. This effectively
gives a Duplex system the properties of a Triplex system.

Automatic Change of Gain. When a lane is shut down the remaining lanes must carry
the additional load. This is achieved by automatically adjusting the gain of the system
so that a given disturbance will demand a greater movement, or rate of movement from
the remaining actuator.

Monitoring. This term applies to multiplex systems, where comparisons are made
either between two or more outputs (or inputs) or between an output (or input) and a
selected datum. If the values exceed a preset limit the system will automatically
disconnect.

Duplicate-monitored. This term refers to a system comprising of two systems, which

have separate power supplies, and operate in parallel. The systems are either self-
monitoring or have their outputs checked by parallel comparator circuits. Only one
system is active at any one time, whilst the other system is switched on, but acts only
as an active standby system. If a fault is detected the standby system will
automatically switch over, and become the active system.

Cross Coupling
This is the process by which a yawing motion is set up when the ailerons are used to turn an
aeroplane, which subsequently has to be countered by the application of some rudder.
Additionally, the banked aeroplane will also have its lift vector tilted towards the lower wing, so
that the vertical component is reduced, which has to be recovered by increasing the aeroplanes
angle of attack (pitch attitude). These factors are all taken into account and, in order to obtain a
balanced turn, although the primary action comes from the roll channel, secondary inputs from
both the pitch and yaw channels are also added. In two axes systems the roll and pitch
channels normally work together to achieve the necessary balance.

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ILS Coupling
Whenever an aeroplane navigation receiver is tuned to a specific Localiser it will also
automatically tune the glide slope. When the localiser and glide slope signals become strong
enough they are captured and fed into the autopilot, thus enabling it to fly the aeroplane on an
approach to the runway, which is normally done in two stages.

¾ Firstly the aeroplane intercepts and captures the localiser signal. The pre-selected heading
and localiser signals are then mixed, to enable the autopilot to smoothly align the aeroplane
with the extended runway centre line.

¾ Secondly the aeroplane captures the glide slope beam, and the autopilot modifies the pitch
attitude so that the aeroplane flies down the beam.

During an automatic approach an allowance must additionally be made for the effects of
crosswinds, otherwise the aeroplane would take up a position downwind of the localiser beam
centre, and would cause the aeroplane to fly a stand off track parallel to the beam. An extra
signal equivalent to the magnitude of the aeroplane drift must therefore be fed into the autopilot
to correct for the crosswind, thus when the aeroplane is correctly established on the localiser
beam, the glide path will be intercepted. The mode prior to interception of glide slope will be
either ‘Attitude Hold, Altitude Hold, Indicated Airspeed Hold, or Vertical Speed Hold’, but
the selected mode will be automatically cancelled when the glide slope is captured.

When the vertical beam sensor detects that the aeroplane is approaching the glide slope beam
centre, the capture mode is automatically triggered, and the autopilot will be instructed to
achieve a vertical descent of approximately 700 ft. per minute. After a short period
(approximately 10 seconds) during which the glide slope error signal will reduce to
approximately zero, the glide slope track mode will be engaged. As the aeroplane approaches
the runway both the glide slope and localiser beams will converge, and to reduce the possibility
of the aeroplane entering an oscillatory motion the error signals will be attenuated as an inverse
function of radio altitude. During an automatic approach some form of beam deviation warning
must be fitted, and this is indicated either via a ‘Beam Deviation Light’ in association with a

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Land status light, or as part of the EADI. Deviation warning is normally given as a flashing
localiser or glide slope indication.

VOR Coupling
VOR coupling is similar to that in the Localiser mode, except in this instance the VOR is
captured when a lateral beam sensor selects it. As the aeroplane approaches the selected
radial the autopilot will automatically capture it when the deviation is approximately one dot (5°).

Prior to capturing the beam the aeroplane will normally be on a pre-selected heading or
alternatively on another intercept mode, and this mode will be automatically discontinued
whenever the VOR capture mode is initiated. Tracking of the VOR radial will occur when the
deviation signal, heading error signal and roll attitude signal approaches zero.

As the aeroplane nears the VOR transmitter its error signals will tend to become erratic due
‘Beam Convergence’, and the ‘Cone of Confusion’. To prevent this it is usual to cut off the
VOR signal as the aeroplane nears the ‘Cone of Confusion’, and this is achieved by an ‘Over
Station Sensor (OSS)’, which automatically deselects the VOR signal for a preset time period,
until the aeroplane has over flown the ‘Cone of Confusion’.

Stability Problems
The two most common stability problems that are likely to be encountered by an aeroplane are:-

¾ Dutch Roll. This is an oscillation in yaw and roll, where roll predominates yaw. This
motion is most common on aeroplanes with swept wings, and if it is not dampened
out by the natural stability of the aeroplane, will lead to a divergent Phugoid. To
correct for this a ‘Yaw Damper System’ is fitted, which is designed to pick up
extremely small deviations in yaw away from the flight path, and automatically apply
inputs to the rudder.

¾ Tuck Under. This is the phenomenon encountered in aeroplanes flying at high

subsonic or transonic speeds. At these speeds there may be a significant rearward
movement of the centre of pressure, and a subsequent uncontrollable ‘nose down’
motion. To compensate for this a ‘Mach Trim System’ is fitted.

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Yaw Damper
This unit is designed to counteract the yawing tendency at an early stage of onset, and before
the roll motion associated with the onset of ‘Dutch Roll’ can develop. This system is
completely independent of the AFCS and is operational whenever it is switched on. It is
however an operational requirement that all ‘Yaw Damper systems’ are switched on, and are
also operational prior to take-off, although only one system is active at any one time. A typical
system is shown below.

The sensor unit is a rate gyro that has its sensitive axis vertical, and is contained within a unit
known as the ‘Yaw Damper Coupler’. This unit also contains the yaw damper circuit computer,
which provides the functions of:-

¾ Filtering the detected error signals and comparing them to a reference signal before
passing a command to the next stage of processing. This additionally removes any
error that may be caused by fuselage flexing.
¾ Integration of the filtered signals to form the servo signal input.
¾ Amplification of the servo signal.

The amplified servo signal is fed via a ‘Transfer Valve’, which is independent of the normal
AFCS actuator, and directs hydraulic fluid under pressure to the yaw damper rudder actuator.
The inputs from the yaw damper and from the rudder controls are normally arranged in ‘Series’
or ‘Parallel’, which relates to the method by which the system affects the rudder control. In a
‘Series’ system there is no resultant rudder pedal movement, but in a ‘Parallel’ system operation
of the yaw damper system will result in rudder pedal movement. Some systems alternatively
operate in ‘Series or Parallel’ dependent on the phase of flight. During the ‘Cruise’ the system
will operate in ‘Series’, when limited rudder is required, but during the ‘Autoland’ stage, when
increased rudder authority is required, the system will switch over to a ‘Parallel’ system. A
feedback loop is also fitted so that, so that when the yawing motion has been stopped, the
rudder will be automatically returned to its normal position.

The yaw damper signal, for a given rate of oscillation, is inversely proportional to airspeed, so a
signal from the Air Data Computer is required. The yaw damper signal may also have to be
modified for different configurations. If this is necessary, a signal from the flap position indicator
circuit will be applied to a gain circuit on the yaw damper output, which will increase the rate of
response when the flaps are extended.

Operation of the Yaw Damper system is monitored on a flight deck indicator, which also shows
the position of the rudder. A test circuit is additionally provided, which simulates a yaw
oscillation by applying torques to the rate gyro sensor. The displaced gyro in turn generates an

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error signal, which displaces the rudder, and its respective movement is relayed to the flight
deck indicator.

Mach Trim System

This system is attached to the pitch channel, and like the yaw damper does not depend on the
AFCS for its operation. A typical system is shown on the next page.

The heart of the system is the coupler unit, which receives signals, corresponding to Mach
number, from the ADC. If these signals exceed a predetermined value (depending on the
aeroplane type) the trim coupler unit will release the brake, and the speed signal from the ADC
will be fed to the motor. This will cause the stabiliser to move in such a way that the elevators
are driven upward, and will thus counteract any tendency for the aeroplane to ‘Tuck-Under’.
This system is also internally monitored, and if the system fails a fail indicator light will illuminate
on the flight deck.

Automatic Pitch Trim Control

Automatic trim control in an AFCS is usually only provided about the pitch axis, and must be
active whenever the AFCS is engaged.

The design of the trim control varies significantly from type to type, but for the purpose of this
manual only two types will be compared.

Small General Aviation Twin-Engine Aeroplane. In this type of aeroplane the elevators will
be operated by a system of cables that are powered either manually or electrically. A typical
automatic trim system is shown below.

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In this system a sensor compares the tension of the ‘Up’ elevator control cable with that of
the ‘Down’ control cable, and if the elevator is subject to an ‘Up’ command, the sensor will
detect the imbalance in the cable tension. This will in turn produce an error signal, which will
be amplified and forwarded to a trim motor. The trim tab will then be driven downwards until
the control cable tensions are reduced to a minimum. A ‘Flap Compensation’ circuit may
also be added to the command signal detection circuit, to compensate for any pitch change,
which may occur when the flaps are extended, and will be operated via a relay attached to
the flap position indicator.

If the aeroplane is alternatively being flown manually, trimming is achieved via a control
switch, which controls the application of power to the servo system.

Large Transport Category Jet Engine Aeroplane. In this type of aeroplane, the elevators
and variable incidence tailplane are normally hydraulically powered, so the associated
operating system is much more complex. The control of the tailplane is affected by an
automatic trim signal, which controls a hydraulic motor. The trim motor is connected via a
clutch assembly, and screw jack to the tailplane. .A block schematic of such a system is
shown below.

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In this system automatic trimming is achieved by moving the variable incidence tailplane
only, whilst the elevators are left with full movement authority, to control any pitch changes
commanded by the AFCS. If the detectors detect that an elevator is too far out of the faired
position for several seconds, the tailplane will move up or down as required until the need for
the autopilot to hold the elevators out of the faired position is reduced. When the elevators
get close to the faired position the detectors will stop the operation of the tailplane. For
example if a nose down demand from the AFCS is ordered, the pitch control channel will
activate the Power Control Unit (PCU), and this will drive the elevator downward. At the
same time a signal will be supplied to the trim circuit, and when the relative position between
the stabiliser and elevator reaches a threshold value, the trim circuit will be activated. A
signal will then be passed to the solenoid, which in turn will activate the hydraulic motor, and
will thus reposition the stabiliser. A new elevator ‘neutral’ position will be established, such
that a signal will be fed back to the elevator positioning circuit, and the elevator will be
repositioned accordingly. The reason for elevator displacement or neutral shift is to augment
the control authority of the tailplane, which is achieved via a neutral shift rod mechanism.

Limit switches prevent the tailplane from being driven beyond pre-set limits, and a stabiliser
trim indicator on the flight deck shows the position of the trimming surfaces. The system is
monitored by a fault detection circuit and will provide the following warnings:-

1. A warning light will illuminate on the flight deck if the automatic trim system fails.
2. An aural alert will be sounded if an excessive trim input is detected.

Interlocks
Before an autopilot can be safely engaged with the flight controls, certain conditions must be
fulfilled. These conditions vary slightly between autopilots, but basically take the form of a
number of switches called interlocks, which are connected in series. A list of some of the
conditions that must be operative before the interlocks will allow the autopilot to be engaged are
as follows. The:-

¾ AC and DC Power supplies must be in tolerance.

¾ gyroscopes must be trimmed and erected.

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¾ Yaw damper system must be engaged.

¾ automatic pitch trim system must be available.
¾ Mach trim system must be available.
¾ Instinctive Cut-Out (Disengage buttons) must not be not pressed.
¾ CADC must be operative.
¾ Manual controls must be in their neutral position, ie. Control wheel and column, turn
knob, pitch switch).

Only once all the interlocks have been enabled the system can be engaged.

Synchronisation
It is important that there is a smooth transition from manual to automatic flight, ie. if an
aeroplane is in a climb, it would not be desirable to have a zero pitch datum when the autopilot
is engaged. Prior to engagement, any signal present on a particular channel is passed to a
‘Synchroniser’ circuit, where an opposing signal equivalent to the attitude error builds up,
which continues right up until the autopilot is engaged.

For example, if the aeroplane is pitched up prior to engagement, the resulting attitude error
signal will be passed through the channel, and will develop an opposing signal in the
synchroniser circuit. When the autopilot is engaged the input to the synchroniser will be
disconnected, and the synchroniser level will be frozen at its current value, which will provide a
datum for the system.

Instinctive Cut Out

It must be possible for the pilot to disengage the autopilot, and this is done via ‘Instinctive Cut-
Out Buttons’, which are located on the control column. These buttons are positioned on the
outside of the control wheel, so that the hand, which is not used for controlling the throttles, can
operate them. It should also be possible to disengage the autopilot by manual operation of the
control column. If the autopilot is disconnected via the ‘instinctive cut out’, the flight crew will
receive a short-term audio and visual warning.

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Chapter 6.3
Automatic Landing System

Introduction
If the weather is poor at a particular destination, it is still possible to land there, and not divert to
another destination, saving the expense of transporting passengers to their original destination
by other means, by using an ‘Automatic Landing System’. Any diversion also means that the
aeroplane will be out of position, and most probably would have to be flown to its original
destination with no fare paying passengers.

For an aeroplane to be certified with an automatic landing capability it must comply with the
minimum requirements laid down in ‘JAR-AWO (ALL WEATHER OPERATIONS)’.
Basic Requirements for an Automatic Landing System
The basic requirements for an Automatic Landing System are that:-

¾ The safety achieved by an automatic landing must not be less than a manual
landing, and the risk of a fatal accident should be better than 1 x 10-7.

¾ The flight crew must be able to adequately monitor the landing phase, so that if a
critical malfunction occurs, the autopilot can be manually disengaged, and manual
control of the aeroplane taken at any time, with the minimum skill required to keep
the aeroplane under control.

¾ The aerodrome must have the required and suitably calibrated radio aids.

¾ It is improbable that the landing performance will be outside the following limits, as
shown in the following diagram.

a. Touchdown must occur between 60m – 900m after the threshold.

b. The outboard landing gear must be no more than 21 m from the runway
centre line assuming a runway width of 45 m.

Automatic Landing System Terminology

In order for an aeroplane to carry out a fully automatic approach and landing it is necessary for
the aeroplane to be fitted with two or more autopilots, and an Autothrottle system. The
utilisation of multiple systems ensures that it is unlikely that in the event of a single major
component failure in the autopilot system, it will not cause the aeroplane to deviate away from

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its approach path. The following terminology is associated with flight failures of an automatic
landing system:-

Fail-Operational. An automatic landing system is ‘Fail-Operational’ if, in the event of a

failure below the alert height, during the approach, flare and landing flight phase, the
landing can still be completed by the remaining part of the automatic system. In the
event of a failure, the automatic landing system will be downgraded to a ‘Fail-Passive’
system. A ‘Fail-Operational’ system is alternatively known as a ‘Fail-Active’ system.

Fail-Passive. An automatic landing system is fail-passive if, in event of a failure, there

is no significant out-of-trim condition or deviation of flight path or attitude but the landing
cannot be completed automatically. A Fail-Passive system is alternatively known as a
‘Fail Soft’ system.

Dual-Dual. This term is used by some manufacturers to define a twin ‘Fail

Operational’ control system, which have two passive monitoring systems. This type of
system is not the same as a Duplex system, since the control systems may or may not
be active simultaneously. In the event of a monitor detecting a failure in its associated
system, the second system with its respective monitor will be automatically switched on.

Alert Height. The alert height is a specified radio height, based on the characteristics
of the aeroplane and its ‘Fail Operational Landing System’. In operational use, if a
failure occurs above the alert height in one of the required redundant operational
systems in the aeroplane, (including, where appropriate, ground roll guidance and the
reversionary mode in a ‘Hybrid System’) the approach would be discontinued, and a
‘Go-Around (GA)’ executed, unless reversion to a higher decision height is possible. If
a failure in one of the required redundant operational systems occurred below the alert
height, it would be ignored and the approach continued.

Decision Height. This is the wheel height above the runway elevation by which a GA
must be initiated unless adequate visual reference has been established, and the
aeroplane position and approach path have been assessed as satisfactory to continue
the approach and landing safely.

Automatic Landing System Equipment Requirements

The following list of systems and equipment is required in aeroplanes with automatic landing
systems, in order to achieve the following decision heights:-

Decision Height 199-100ft

1. Autopilot with an ILS coupling mode. (Note: A flight director system with an
ILS coupling mode may be approved for use following failure or disconnect
of the autopilot)
2. Autothrottle (unless it can be shown speed control does not add
excessively to the crew workload).
3. Radio altimeter.
4. Excess ILS deviation warnings.

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Decision Height 99-50ft

1. As for decision height 199-100ft.

2. Autopilot with autoland mode.

Decision Height Below 50ft

1. As for decision height 199-100ft

2. (a) Autopilot with a fail-operational autoland mode and an automatic missed
approach mode.
OR
(b) Autopilot with automatic landing and missed approach modes and a
landing guidance display.

No Decision Height

1. As for decision height below 50ft.

2. (a) Autopilot with a fail-operational ground roll mode.
OR
(b) Fail operational head-up ground roll guidance display.
OR
(c) Autopilot with ground roll mode and a head up ground roll guidance
display.
3. Anti-skid braking system.

Automatic Approach, Flare and Landing Sequence

The following diagram shows the stages involved during the automatic approach, flare and
landing phases of flight of an aeroplane fitted with a triple channel autopilot, such as that fitted
to a Boeing 757.

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Depending on the number of channels engaged the system will perform an automatic landing
with a Land 2 or Land 3 status, which is displayed on the EADI.

Land 2 signifies 2 autopilot channels engaged giving the system a ‘Fail Passive’
capability.

Land 3 signifies 3 autopilot channels engaged giving the system a ‘Fail Operational’
capability.

During the cruise and approach to an airport a single autopilot is normally engaged and
controlling the aeroplane along the designated route. As the aeroplane approaches the airport
the flight crew manually select the other autopilots, and when the ‘Approach Mode (APP)’ is
selected on the AFCS control panel, the localiser/glide slope, together with the two remaining
autopilots will be armed. As the aeroplane passes through 1500 ft radio altitude, with the
localiser and glide slope already captured, the two off line autopilots will be automatically
engaged. This will be indicated to the flight crew by a ‘LAND 3’ status, and the aeroplane will
continue to fly down the glide slope.

Note: If a single failure occurs between 1500-200ft radio altitude, the system will
downgrade itself to a ‘Fail-Passive’ system, and a ‘LAND 2’ status will be annunciated.

At a radio altitude of 330ft the aeroplane will be trimmed nose up by an automatic adjustment of
the variable incidence tailplane. As the aeroplane passes through the alert height (normally
200ft radio altitude) the reversion to ‘LAND 2’ due to a fault will be inhibited until the aeroplane
is below 40 knots, during the roll out.

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When the aeroplane is 45ft above the ground the flare mode will be automatically engaged and
the aeroplane rate of descent will be progressively reduced to achieve a rate of 2 ft. per second
on touchdown. At the same time the Autothrottle will also reduce the amount of available
engine thrust, in order to maintain the flare path. At 5ft gear altitude the flare will be disengaged
and there will be a transition to touchdown, and the subsequent roll out mode. The Land status
will remain engaged until de-selected by the flight crew.

Importantly these systems are designed to carry out automatic landings under all visibility
conditions, and must have a reliability factor better than 1x10-7.

Weather Minima
In low visibility operations, the weather limits for landing are given in terms of the following:

Runway visual range (RVR). This is an instrumentally derived value that represents
the range at which high-intensity lights can be seen in the direction of landing along the
runway. These readings are transmitted to the aeroplane by ATC, and provide the flight
crew with the visibility conditions, which currently exist at the airport.

Decision height. This is the wheel height above the runway threshold by which a
‘Go-Around (GA)’ must be initiated, unless adequate visual reference has been
established, and the position and approach path of the aeroplane have been visually
assessed as satisfactory to safely continue the approach or landing.

Minimum values of these two quantities, which are referred to as `Weather Minima', are
specified by The National Licensing Authorities, for various types of aeroplanes, and also for
various airports.

ICAO Categorisation for Low Visibility Landing Capabilities

This system is based on the principle that the probability of having an adequate short visual
reference, for the range of permitted decision heights, should be as high as possible.

Category 1: Operation down to a minima of 200 ft decision height, and a RVR of 800
m, with a high probability of approach success.

Category 2: Operation down to a minima below 200 ft decision height and RVR of 800
m, and to as low as 100 ft decision height and RVR of 400 m, with a high probability of
approach success.

Category 3A: Operation down to, and along the surface of the runway, with external
visual reference during the final phase of the landing down to a RVR minima of 200 m.

Category 3B: Operation to, and along the surface of the runway and taxiways with
visibility sufficient only for visual taxiing comparable to a RVR value in the order of 50
m.

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Category 3C: Operation to, and along the surface of the runway and taxiways without
external visual reference.

These categories also serve as an indication of the stages through which automatic approach
and automatic landing development progresses, and thus designates the capabilities of an
individual AFCS.

The capabilities of the ground guidance equipment available at a particular airport are also
designated, and are as listed below.

¾ ILS localiser and glide path.

¾ Approach lighting.
¾ Runway and taxiway lighting.

In connection with automatic landing systems, the term `All Weather Operations' is frequently
used. This term, which is frequently taken to mean that there are no weather conditions that can
prevent an aeroplane from taking-off and landing successfully, is not strictly correct, eg. an
aeroplane cannot perform a landing task in wind conditions in excess of those for which it has
been certificated, or on a runway which, because of contamination by water, slush or ice, is not
fit for such an operation.

The Fundamental Landing Requirement

During an automatic landing an aeroplane should be controlled in such a way that:-

¾ Its wheels make contact with the ground comfortably within the paved surface of the
runway and that the landing point is not too far down the runway.
¾ It lands at a very low vertical velocity in order to avoid collapse of the landing gear.
¾ The speed at touchdown should be sufficiently low to allow the aeroplane to be
brought to a halt within the remaining length of the runway.

To facilitate the above the aeroplane’s:

¾ Final rate of descent should be no greater than 1 to 2 feet per second.

¾ Airspeed should be reduced from 1.3Vs during the approach to 1.15Vs, by a
progressive reduction of engine thrust during the landing flare.
¾ Wings should be levelled prior to the actual landing, and any drift `Kicked-Off'
before touchdown.

To achieve all of the above, requires that the aeroplane be controlled about all three axes
simultaneously.

Any such system, as well as being capable of achieving at least the targets listed above, must
also be designed with the following aims:

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¾ Achieving the highest integrity and reliability of systems, bearing in mind that they
need to be entrusted with very considerable authority over the controls of an
aeroplane, including the throttles, and in the presence of the ground.
¾ The provision of adequate monitoring information on the progress of the approach
and landing manoeuvre, and which will enable the flight crew to take over, under
the most critical conditions of a system malfunction in the presence of the ground.
¾ The substitution of the flight crew’s direct vision with an automatic externally
referenced guidance system, having an integrity and reliability of the same high
order as that demanded of the `on board' system.

System Reliability and Integrity

Devices designed to limit the authority of automatic control systems in the event of `runaway'
conditions that may result from malfunctions, may be incorporated in more conventional control
systems. These are however normally only effective for the intended purpose down to any
specified `break-off' height, ie. the approach height at which a control system is disengaged,
but these systems do not satisfy the requirements for systems designed for autoland.

The setting of safety devices is dictated by the conflicting requirements listed below:-

¾ They must limit the effect of a `runaway' such that a safe recovery can be effected
by the flight crew;
¾ They must allow sufficient authority to the control system so that the required flight
path can be followed accurately in the presence of disturbances.

A further factor limiting the application of safety devices (in the manner of conventional control
systems) is their inability to protect against passive failures. While not producing flight path
changes directly, these failures would nevertheless mean that the predetermined and accurate
flight manoeuvre of automatic landing, could not be maintained, and so could set up an equally
dangerous situation.

It therefore follows, that to achieve the objective of automatic landing, the operation of an AFCS
must be of such a nature that it will:-

¾ Not disturb the flight path as a result of an active malfunction.

¾ Have adequate authority for sufficiently accurate control along the required flight
path.
¾ Warn of a passive failure.
¾ Not fail to complete the intended flight manoeuvre following an active or a passive
failure.

In order to resolve these problems the concept of `system redundancy' is applied. ie. the use
of multiple systems, operating in such a manner that a single failure within a system will have an
insignificant effect on the aeroplane's performance during the approach and landing operation.

¾ The autothrottle will remain in control of the engines until reverse thrust is
demanded.

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¾ The AFCS will remain in control of the aeroplane until the crew disengages it.

At any time from 2000 feet down to the decision height, the flight crew can elect to abort the
approach, by pressing the TOGA switch on the throttle, which will cause the throttles to advance
to a pre-set reduced thrust ‘Go-Around’ value. A second press of the TOGA switch will
command the throttles to advance the engines to full power. The ‘Go-Around’ phase will then
interact with the AFCS, and will cause a ‘GA’ annunciation on the ADI. The pitch channel will
generate a ‘pitch-up’ command, and the aeroplane will be placed in the correct climb attitude.

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Chapter 6.4
Thrust Management Systems

Introduction
Most modern aeroplanes are equipped with systems to control and compute engine thrust.
Knowledge of how engine thrust is computed, how it is controlled, and how engine inputs are
collected is therefore necessary.

Determining the Thrust Required

Thrust curves and charts are published in the aeroplane manual, and also in the performance
manuals. These are used to determine the required ‘Engine Pressure Ratio (EPR)’ and/or
‘Fan Speed (N1)’ for any desired engine rating, which depends solely on the prevailing ambient
temperature and barometric pressure. Time, engine speed, and the ‘Exhaust Gas Temperature
(EGT)’ limit the take-off rated thrust. Maximum continuous, maximum climb and maximum
cruise thrust ratings are alternatively limited by EGT for a given length of time, or continuously,
as defined below.

Engine Pressure Ratio (EPR). This is the amount of useful thrust being developed by
an engine. It is the product of the mass of air passing through the engine, and its
velocity at the exhaust nozzle, minus the drag due to the air passing through the
engine. By comparing the air pressure across the engine, ie. the ratio of the exhaust
pressure to the compressor inlet pressure, it gives the Engine Pressure Ratio (EPR),
which is an indication of the thrust output from the engine. EPR is usually given as a %
thrust value.

RPM, N1, N2 or N3. These are normally given as a % of the maximum value.

EGT. The exhaust gas temperature, which must be monitored in order to prevent
excessive heat damaging the turbine.

Rated Maximum Continuous Thrust. This is the thrust that is approved for
unrestricted periods of use, and according to JAR, is defined in the aeroplane flight
manual.

Maximum Continuous Thrust (MCT). This is the amount of thrust, which is authorised
for emergency use at the discretion of the pilot only. This is also used, for aeroplane
certification requirements, and for climb operations as determined by the airframe
manufacturer.

Maximum Climb Thrust. This is the maximum thrust approved for the climb phase,
which on some engines is identical to the rated maximum continuous thrust level. This
rating is selected by positioning the throttle to give the required EPR or N1, for the
prevailing climb profile and engine inlet temperature. The climb thrust curves or charts
are contained in the aeroplane performance manual.

Maximum Cruise Thrust. This is the maximum approved thrust for cruise operation.
These thrust limitations must be closely adhered to, because are allied to specific

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warranty limitations, which have been established by the engine manufacturer, to

ensure optimum operation, and engine life.

Calculation of Climb and Cruise Thrust

To determine the necessary thrust to obtain the desired climb and cruise performance, the flight
crew must consult the charts in the aeroplane performance manual. A simplified method to
calculate the amount of reduced thrust is the assumed temperature method. This is based on
using the take-off thrust and aeroplane speeds for an assumed temperature higher than the
actual ambient temperature (OAT). The assumed temperature procedure enables the flight crew
to make thrust reductions according to the prevailing conditions, and within practical limits.

Climb. This is the thrust, which is established by adjusting the throttle to obtain the
appropriate parameter indication (N1 or EPR) in accordance with the published climb
thrust setting charts or curves. On some engines, the maximum continuous and
maximum climb thrust levels may be the same.

During the climb, at a fixed throttle setting, as ‘Total Air Temperature (TAT)’ falls with
increasing altitude, the N1 and EPR values will progressively increase in accordance
with the thrust curves. Normally only one or two throttle adjustments should be
necessary throughout the climb, depending on whether a high speed or long range
climb is being performed. Significant temperature and speed deviations may however
require slight throttle adjustments in order to avoid exceeding the thrust curve values.
By comparison typical changes in the ambient temperature and Mach number will not
require any such adjustments.

Importantly, the EGT must be continually monitored throughout the climb so that the
established maximum climb EGT is not exceeded.

Cruise. Upon reaching the selected flight level, the climb thrust may only be
maintained long enough to allow the aeroplane to accelerate to its designated cruising
speed. To maintain the desired speed, the required N1 or EPR is set in accordance with
the cruise charts, or curves, which are applicable to the prevailing cruise conditions.

Cruise thrust values are based on cruising speeds, in KIAS, and Mach number,
aeroplane gross weights, TAT, and pressure altitudes. The charts and curves cover all
realistic combinations of these factors, within the operational scope of the engine and
aeroplane up to, and including the maximum cruise thrust rating.

The Maximum Cruise Thrust is an engine warranty limitation, and should thus not be
exceeded whilst attempting to maintain a given altitude and speed.

Cruising Methods
Procedures used for cruising depend primarily on the length of time spent in the cruise during a
flight. For short flights, within given periods of time, depending on the aeroplane-engine type,
and cruising environments, the fixed throttle cruise will provide a favourable or acceptable
balance of fuel consumed, versus time saved. Once the thrust level has been set to obtain the
desired cruising speed, the throttle position may remain fixed throughout the cruising portion of
the flight.

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Electronic Engine Control (EEC)

The first Electronic Engine Control System (EEC) was purely a supervisory control, which
utilised proven hydro-mechanical controls. The major components in this system included the
control itself, the fuel control of the engine, and the bleed air and variable stator vane control.

In this system, the flight crew simply moved the thrust lever to a desired thrust or maximum
climb position, and the control system automatically adjusted the EPR to maintain the thrust
rating irrespective of changes in flight, and ambient conditions. The control also limited the
engine speed and temperature, which ensured that the engine/engines were operated safely
throughout the entire flight envelope.

If a fault were to occur in this system, the control would automatically revert to a hydro-
mechanical system, whilst maintaining the required thrust level. Full reversion to the hydro-
mechanical system can also be instigated at any time.

Some electronic control systems alternatively function as a limiter only. For example, if the
engine shaft speed, or EGT approaches the limits of safe operation, then an input is
automatically made to a fuel flow regulator, to reduce the fuel flow, and thus maintain the
desired shaft speed, or EGT at a safe level.

Modern Full Authority Electronic Engine Control (EEC) are fully redundant systems, which
control all engine functions, and also eliminate the need for a back-up hydro-mechanical control
system. This control system is more commonly referred to as a ‘Full Authority Digital Engine
Control (FADEC)’ system.

Full Authority Digital Engine Control (FADEC)

FADEC is a digital electronic fuel control system, which is specifically used on gas turbine
engines and functions during all engine operations. It includes total electronic engine control,
and operates with the Flight Management Computer, to schedule the fuel to the engine. A
typical FADEC system is shown on the next page. One of the basic purposes of FADEC is thus
to reduce the flight crew workload, particularly during the critical phases of flight. This is
achieved by the FADEC's control logic, which simplifies the power settings for all engine-
operating conditions. The thrust levers thus achieve engine thrust values for a set lever
position, regardless of the flight or ambient conditions. For example, assume a given EPR at a
particular OAT; if the OAT consequently changes, the system is designed to automatically
adjust the amount of fuel being supplied to the engine, in order to maintain a set EPR.

The FADEC system establishes the amount of engine power through direct closed-loop control
of the EPR, which is the thrust rating parameter. The selection of EPR is normally calculated as
a function of thrust lever angle, altitude, Mach number, and TAT.

The Air Data Computer supplies altitude, Mach number and TAT to the control system, whilst
sensors provide measurements of engine temperatures, pressures, and speeds. This data is
used to provide automatic thrust rating control, engine limit protection, transient control and
engine starting. The control system also implements EPR schedules to obtain the EPR rating at
various throttle lever angle positions, and provides the correct rating at a constant throttle lever
angle during changing flight or ambient conditions.

The FADEC has the following advantages over a mechanical system:-

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¾ The system requires no engine adjustment, and therefore no engine running, which
saves fuel.
¾ The system reduces fuel consumption through improved engine bleed air control.
¾ The system fully modulates the active clearance control systems, producing a
substantial benefit in performance by reducing the engine blade tip clearances.
¾ The higher precision of the digital computer ensures more repeatable engine
transients (ie. acceleration/deceleration) than that possible with a hydro-
mechanical system. The latter is subject to manufacturing tolerances, deterioration
and wear, which will affect its ability to consistently provide the same acceleration
and deceleration times.
¾ The system ensures improved engine starts by means of digital schedules and logic
that adjusts for measured conditions.
¾ The system provides engine limit protection by automatic limiting of the critical
engine pressures and speeds. Direct control of the rating parameter also prevents
inadvertent overboost of the selected rating when the power is being set.
¾ The engine idle speed remains constant regardless of changes in ambient
conditions and bleeds requirements, whereas with the mechanical system, the
engine speed changes with ambient conditions.

FADEC takes over virtually all of the steady state and transient control intelligence and replaces
most of the hydro mechanical and pneumatic elements of the fuel system. The fuel system
solely reduced to a fuel pump and control valve, an independent shut-off cock and a minimum
of other additional features, which are necessary to keep the engine safe in the event of
extensive electronic failure. FADEC also furnishes information to the engine instrument and
crew alerting system.

The FADEC is mounted on the engine compressor casing on anti-vibration mounts and is air-
cooled. The figure below indicates the signals, which are transmitted between the engine,
mounted components and the engine/aeroplane interface.

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The control has dual electronic channels, each having its own processor, power supply,
programme memory, selected input sensors and output actuators. A dedicated engine gearbox
driven alternator also provides power to each electronic control channel. If computational
capability is lost in the primary channel, the FADEC will automatically switch to the secondary

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channel. If a sensor is lost in the primary channel, cross-talk with the secondary channel will
automatically supply the necessary information.

In the unlikely event of the loss of both channels of the electronic control, the torque motors are
spring loaded to failsafe positions. The fuel flow goes to minimum flow, the stator vanes are set
to fully open (to protect take-off), the air/oil cooler goes to wide-open, and the active clearance
control will be shut off.

Autothrottle (A/T)
The Autothrottle (A/T) system is a computer controlled electromechanical system that controls
engine thrust within engine design parameters. The system:-

¾ Computes and displays EPR and/or N1 and speed (IAS or MACH) information.
¾ Provides automatic control from start of the take off, through the climb, cruise,
descent, approach and Go-Around or landing.

The throttle position for each engine is controlled in order to maintain a specific engine thrust
(N1 or EPR) or target airspeed, for all flight regimes as directed by the Thrust Management
Computer. Autothrottles are fitted in modern jet engine aeroplanes, primarily to conserve fuel,
and a typical system is shown below.

It has inputs from a Thrust Management Computer (TMC), which integrates signals from the
engines, a Thrust Mode Select Panel (TMSP), a Flight Management Computer (FMC), and also
receives signals from the Air Data Computer (ADC). With full forward throttle the TMC provides
maximum engine power without exceeding the specified operating limits during the take-off and
go-around phases of flight. The system can however be manually over ridden at any time.

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The A/T and AFCS operate together to maintain a specific airspeed and vertical path profile.
With the A/T ‘ON’, and either the autopilot or the flight director ‘ON’, one system or the other will
maintain a set airspeed. When the AFCS mode is controlling the airspeed, the A/T will
automatically control a specific engine thrust value, but when the AFCS mode is controlling the
aeroplanes vertical flight path, the A/T will automatically maintain the desired airspeed through
thrust control.

The engine’s on Airbus aeroplanes are driven by FADEC, but alternatively use electrical signals
for thrust control in order to eliminate the weak points of the conventional Autothrottle system,
which is mechanically operated. On these aeroplanes, the actual throttle (thrust lever) position
does not move automatically, unlike the early auto-throttles, thus making them much more
reliable.

The Airbus system can be operated in either manual or auto-thrust modes. In the manual thrust
mode the flight crew will move the thrust levers between idle and full thrust as usual, but in auto-
thrust, the thrust levers are set in a fixed position, which is defined by the maximum amount of
available thrust. Whether in manual or auto-thrust, speed and power changes are monitored via
N1, the IAS, and speed trends, as on any aeroplane.

Thrust Lever Operation

The Autothrottle system when engaged moves the thrust levers together, in response to error
signals generated, and compares the aeroplane’s actual flight conditions, against selected
datum’s. In the majority of aeroplanes the throttles can be manually repositioned at any time
without disengagement of the Autothrottle.

NOTE: In the Airbus Fly by Wire type of aeroplane the throttle levers are in a
fixed position on the throttle box, and the throttle levers do not move.

Thrust Management via the Autothrottle

The Autothrottle operates in response to flight crew mode control panel inputs from the AFCS,
or alternatively via automatic FMC commands. The Autothrottle system:-

¾ Uses reference thrust limits calculated by the FMC

¾ Commands the thrust levers
¾ Commands thrust equalisation through the electronic engine controls

The FMC calculates a reference thrust for the following modes (typical):-

¾ Takeoff
¾ Reduced takeoff (also called derated takeoff)
¾ Assumed temperature takeoff
¾ Climb
¾ Reduced climb
¾ Cruise
¾ Go-around

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The mode used for thrust reference, automatically changes during the respective flight phase,
as selected by the FMCS, and the selected thrust reference mode is displayed on the thrust
mode display.

Thrust Management Computer (TMC)

The TMC operates the throttles in response to manual inputs from the Thrust Mode Selector
Panel (TMSP), and automatic commands from the Flight Management Computer (FMC) when
operating in the VNAV mode. The basic functions of the TMC are to:-

¾ Calculate thrust limits (using outside pressure and temperature), and thrust
settings, or follow any FMC thrust commands.
¾ Provide automatic control of the three primary modes (EPR, MACH HOLD and
speed) for various flight phases, as shown on the next page.
¾ Detect and transmit A/T failures.

Thrust Mode Select Panel (TMSP)

The TMSP, as shown below allows:

¾ the TO/GA, CLB, CON, CRZ reference thrust modes to be selected.

¾ temperature de-rating to be selected, in order to prolong the engine life.

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Chapter 7.1
Central Warning System

Introduction
In modern aeroplanes many systems require constant monitoring, which in turn require a
corresponding multitude of warning devices; both visual and aural, to be fitted. These warnings
consist of flashing lights, horns and bells, which if operating in various parts of the flight deck
could pose an unnecessary distraction. In order to reduce this probability, aeroplanes are
equipped with a Central Warning System (CWS).

Central Warning System Annunciator Panel

In its basic form, the system comprises of a group of warning and indicator lights, which are
connected to signal circuits allied to the various aeroplane systems. Each light displays a
legend, which denotes the system, and a malfunction or advisory message. The lights are blue,
red or amber, and are contained on an annunciator panel, which is normally installed on the
centre control panel, as shown below

SYSTEM LIGHTS

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In aeroplanes carrying a flight engineer, a functionally integrated duplicate warning panel is

additionally installed. The character of the signals also varies according to the degree of
urgency or hazards involved. Aural, visual and tactile signals are given singularly or in
combinations to provide both warning and information simultaneously, regarding the nature of
the problem. If the condition requires immediate corrective action by the flight crew a red
warning light will indicate. These lights indicate engine, wheel well, or Auxiliary Power Unit
(APU) fires, autopilot disconnect, and landing gear unsafe conditions. If the condition requires
timely corrective action by the flight crew a cautionary amber warning light will indicate it. A
series of blue lights will alternatively indicate whether any associated system valves are in
transit, or merely in disagreement with the appropriate control switch.

If a fault occurs in a system, a fault-sensing device will additionally transmit a signal to an

electronic device known as a ‘Logic Controller’, which will determine whether the fault is of a
hazardous nature, or is one that simply requires caution. If the fault is considered to be
hazardous, the controller output signal will illuminate the red ‘master warning’ light, but if
caution is simply required, the signal will illuminate the amber ‘master caution’ light. Each
master warning light incorporates a switch unit, so that if the caps are pressed in, the active
signal circuits will be disconnected, and the lights will be extinguished. This action will also
reset the master warning system, so that it can accept signals from other faults, which might
subsequently occur in other aeroplane systems. The system lights are not of the resetting type
and remain illuminated until the system fault is corrected. Dimming of the lights, and testing of
the bulb filaments is normally carried out by means of switches mounted adjacent to the
annunciator panel.

Aural Warnings
Various aural signals on an aeroplane alert the flight crew if warnings or cautions exist, and
some typical ones are listed below:-

¾ a clacker will sound if the airspeed limits are exceeded.

¾ a warning tone will sound if the autopilot is disconnected.
¾ an intermittent horn will sound if the cabin altitude limits are exceeded.
¾ a steady horn will sound if a landing gear disagreement exists.
¾ an intermittent horn will sound if the take-off configuration is incorrect.
¾ a fire warning bell will sound if a fire exists.
¾ Ground Proximity warnings and alerts will be indicated by voice warnings.

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Chapter 7.2
Altitude Alerting System

Introduction
The Altitude Alerting System (AAS) provides both aural and visual warnings of an aeroplane
reaching, or deviating from a pre-selected altitude, by utilising an output from a pressure
altimeter (or ADC).

Altitude Alerting System Operation

If a target altitude is selected on the mode control panel (MCP), a signal from an altitude sensor
will be mixed with the signal created by the control panel pre-set altitude, and as long as the
actual altitude and pre-set altitude are different, a signal difference will exist.

For example the sequence of events which takes place during a descent from 31000 feet to
15000 feet are as follows:-

¾ Prior to leaving 31000 feet the control unit will be set to 15000 feet, and no warning
will be given at this stage, as the signals are too different.
¾ At approximately 1000 feet above the target altitude an aural warning will sound (‘C’
chord) for approximately one second, and an altitude alert light will illuminate
adjacent to each primary altimeter.
¾ The light will remain illuminated until approximately 250 feet above the target.

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¾ If the aeroplane moves greater than 250 feet above the target altitude or continues
through the target altitude to 250 feet below it, an aural warning will sound and the
light will illuminate again.

The same sequence of events will also occur if the target altitude is alternatively approached
from below. Altitude alerting is additionally inhibited whenever the flaps are in the landing
configuration, or whenever the glide slope (G/S) is captured.

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Chapter 7.3
Ground Proximity Warning System

Introduction
Most transport aeroplane crashes occur due to ‘controlled’ flights into the ground, for reasons
such as inattention, confusion, vertigo, distraction, instrument reading error, poor visibility and
navigation error. The Ground Proximity System (GPWS) is thus designed to prevent this sort of
accident, by giving the flight deck crew some advanced warning, both aurally and visually, if an
unsafe flight condition close to the ground exists.

The radio altimeter and the Air Data Computer (ADC) are continuously monitored for the
aeroplane height above the ground, and the barometric rate of change of height; which enables
the rate of closure with the terrain immediately beneath the aeroplane to be continuously
assessed. In a typical GPWS, a red ‘PULL-UP’ light, together with a ‘WHOOP-WHOOP PULL-
UP’ audible command, which will give a warning of an unsafe proximity to the ground. When the
dangerous condition has been corrected, the warnings will cease and the system will
automatically reset itself.

Importantly, no warning is provided of steeply rising ground directly ahead of the aeroplane.
Additionally, the system will not prevent a properly configured aeroplane from landing short of
the runway in the absence of an ILS glide path. The GPWS is normally activated between
50 feet and 2,450 feet above the surface, which is determined by the radio altimeter. The
GPWS must never be de-activated (ie. by pulling the circuit breakers) except when using
approved procedures at airports where GPWS inhibition is specifically required.

GPWS System Architecture

The GPWS comprises of a Central Processing Unit (CPU), which accepts inputs from various
sources. The CPU continually examines these inputs, and if a collision risk with the terrain
exists, appropriate visual and aural warnings will be generated. Computer failure, and any
failure in the input signals will also be displayed on a warning panel

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The GPWS system additionally incorporates a fully integrated self-test function, which is
capable of checking the signal path from all of the inputs described above. If the system checks
are carried out satisfactorily when the test switch is depressed, the normal indication to the flight
crew will be the simultaneous activation of both visual and aural warnings. Notably system
testing by this means is normally prohibited when the aeroplane is airborne.

GPWS Modes
Four main conditions give rise to audible and visual alerts. These four conditions are described
as MODES, and are numbered from 1 to 4. The condition associated with each mode is as
follows:-

MODE 1 Excessive descent rate below 2450 feet AGL.

MODE 2 Excessive Terrain Closure:

¾ In the cruise configuration.
¾ In the landing configuration.

MODE 3 Height loss after T/0, or during a missed approach.

MODE 4 Unsafe terrain clearance when not in the landing configuration below:-
¾ 500 feet AGL with the landing gear not locked down.
¾ 200 feet AGL with the flaps not set for landing.

Other modes which exist are:

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MODE 5 Excessive downward departure from the ILS glide path.

MODE 6 Aural altitude call-outs on approach, including decision height.

MODE 7 Windshear detection, warning and guidance.

In addition to the inputs from the ADC and the radio altimeter, the GPWS CPU also obtains
information from the following sources:-

Main Landing Gear Selector Assembly - The position of the landing gear governs
whether or not MODE 3 is activated, and will also determine the height / barometric rate
of descent conditions that will activate a MODE 4 warning.

Flap Selector Assembly - The position of the flap governs whether or not MODE 3 is
activated, and will also determine the height / terrain closure rate, which would activate
a MODE 2 warning, and the height / barometric rate conditions, which would activate a
MODE 4 warning.

ILS Receiver - The degree of deviation from the glide path, together with glide path
validity signals, are used in MODE 5.

Stall Prevention - The stall prevention devices, which are fitted to the aeroplane, eg.
stick shakers and stick pushers, will also feed a signal to the GPWS computer to inhibit
the GPWS warnings during the incipient stall and / or stalled condition.

Warning System
To avoid violent manoeuvres, current equipments generate alerts and warnings. The alert is
regarded as a preliminary to a warning, and the flight crew must respond immediately by
correction of the flight path, or configuration, so that the alert ceases. This is a cautionary
indication, and if ignored will be followed by an imperative command or warning. If a warning is
initiated the flight crew must immediately level the wings and initiate a maximum angle of climb
until a minimum safe altitude is reached. Where an advanced GPWS is fitted, the cause of the
warning should be verified after the climb has been initiated. The aural alerts and warnings for
both the basic and advanced systems are summarised in a tabular format on the next page.

GPWS Mode Vertical Basic Equipment Advanced Equipment

Limits
(feet) Alert Warning Alert Warning
1 Excessive 50 - 2450 - Whoop Sink Rate, Whoop
descent rate Whoop Sink Rate Whoop
Pull-up Pull-up
2 Excessive 50 - 2450 - Whoop Terrain, Whoop
terrain closure Whoop Terrain Whoop
2A Cruise Pull-up Pull-up

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GPWS Mode Vertical Basic Equipment Advanced Equipment

Limits
(feet) Alert Warning Alert Warning
Excessive 200 - 790 Whoop Terrain, Whoop
terrain closure Whoop Terrain Whoop
2B Configured Pull-up Pull-up
for landing
3 Altitude loss 50 - 700 - Whoop Don't Sink
after take-off Whoop
or go-around Pull-up
4 4A Proximity to 50 - 500 - Whoop Too Low Too Low
terrain gear Whoop Gear Terrain
not down & Pull-up
locked
4B Proximity to 50 - 200 - Whoop Too Low Too Low
terrain & flaps Whoop Flaps Terrain
not set for Pull-up
landing
5 Below glide 50 - 1000 Glide slope - Glide slope -
slope
6 Below 50 - 1000 - - Minimums -
mimimums
7 Windshear - - - Windshear -
encounter

The alert for Mode 5 consists of a steady amber glide slope light, and an aural glide slope
warning that becomes more frequent, the greater the deviation below the glide slope, and also
becomes louder if the deviation exceeds 2 dots below 300 feet.

GPWS Control Panel

The GPWS warning and alert lights are typically sited on a control panel, as shown below.

PULL-UP AMBER BELOW

RED PULL-UP BELOW G/S
WARNING LIGHT PUSH TO GLIDESLOPE
(PRESS TO TEST) PRESS TO TEST INHIBIT ALERT LIGHT

FLAP OVERRIDE GEAR OVERRIDE INOPERATIVE

SELECTOR SELECTOR
& &
INDICATOR INDICATOR WARNING
LIGHT
(AMBER)

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Pressing the ‘Pull-up’ light will test the system integrity. This will cause the ‘Pull-up’, ‘Below
Glide Slope’ and ‘INOP’ lights to illuminate, and aural warnings to be activated. The INOP light
will also illuminate if the system or power supply fails, or any input to the computer is lost.

A flap-override switch is also available that should be operated whenever a non-standard flap
setting is used during the approach to land. With the switch in this position the associated
warning will be suppressed and the GPWS flap-override indicator will illuminate. Similarly a
gear-override selection will disable the warning if the aeroplane has to land with one or more of
its wheel assemblies retracted.

Discretionary Response
Regardless of the type of GPWS, basic or advanced, on receipt of an alert or a warning, a
response must be made. In order to avoid an excessive manoeuvre, a warning may sometimes
be considered as if it were an alert, but only when the following conditions exist:-

¾ The aeroplane is being operated during the day in meteorological conditions, which
enable it to remain 1 nm horizontally and 1000 feet vertically away from cloud, with
a visibility of at least 5 nm.

and

¾ If it is obvious to the Commander that the aeroplane is NOT in a dangerous

situation with regard to the terrain, configuration or present manoeuvre.

Note: Although some manufacturers of GPWS equipment may show in their literature
‘TOO LOW TERRAIN’ to be an alert, the view of the JAA is that the response should be
as it is for a warning.

Warning Inhibition
The Mode 5 indication alone may be inhibited, so that localizer-only, or back-beam approaches
may be flown without provoking an alert. Inhibition is achieved by pressing the lamp housing of
the amber indicator light, which is only effective in the soft alerting height band above 300 feet.
In all other cases climbing out of the envelope that initially triggered the event will only silence
the warning or alert. There is intentionally no on/off switch, and under no circumstances should
the power supply circuit breaker be tripped to silence a warning.

The Reporting of GPWS Events

All GPWS events should be reported to the regulatory authority on purpose-designed forms.
Reports should additionally be submitted where no warning is received, but when one would
normally have been expected.

Operation of the GPWS

A typical GPWS operates as follows, although some differences do exist between
manufacturers, eg for the Boeing 737-400 the lower limit is 30 feet, but for JAA purposes the
range is always between 50 and 2450 feet.

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MODE 1. This mode is activated whenever the barometric descent rate is excessive
with respect to the aeroplane’s height above the terrain, as determined by the radio
altimeter, and the barometric rate signal, which is obtained from the ADC. The mode 1
envelope is divided into two areas: the ‘initial penetration area’ or ‘sink rate area’, and
the inner warning area, or pull-up area. This mode is also completely independent of the
landing gear and flap positions. The warning envelope for MODE 1 has an upper limit
of 2,450 feet above the ground, and at this height a warning will be given if the
barometric rate of descent exceeds 7,125 ft/min. At the lower limit of the envelope,
which is 50 ft above the ground, a barometric descent rate of 1,500 ft/min or more will
cause MODE 1 activation. The full operating parameters for MODE 1 are shown on the
next page.

Figure 3.3

Penetration of the first boundary will activate the ‘PULL-UP’ light and a repetitive aural
alert of “SINK RATE”. Penetration of the second boundary will result in a repetitive
aural warning of “WHOOP WHOOP PULL-UP”

MODE 2. This mode will be activated whenever the aeroplane has an excessive
closure rate with respect to rising terrain, and is achieved by measuring the terrain
closure rate as determined by the radio altimeter. This mode also consists of two sub
modes; Mode 2A if the flaps are NOT in the landing configuration, and Mode 2B if the
flaps are in the landing configuration. The full operating parameters for Mode 2 are
shown below.

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Like Mode 1 this mode also possesses two boundaries; the initial penetration area and
the inner warning area. Penetrating the first boundary will result in an aural alert of
“TERRAIN” being repeated twice and will be followed by a repetitive aural alert of
“WHOOP WHOOP PULL-UP”

A Mode 2A “TERRAIN” alert will also be given if the landing gear and flaps are NOT in
the landing configuration when the aeroplane is exiting the ‘PULL-UP’ area, and will
continue until at least 300 feet of barometric altitude has been gained, or the landing
gear has been lowered. As the airspeed increases from 220 knots to 310 knots with the
landing gear retracted, the radio altitude at which Mode 2A alerting or warning varies
according to an airspeed expansion function up to a maximum of 2450 ft.

At airspeeds less than 220 knots, the upper boundary will be 1,650 ft radio altitude and
a Mode 2A warning will be activated if the terrain closure rate is equal to, or in excess of
6,000 ft per minute. At the lower limit of MODE 2A operation, a warning will be given if
the terrain closure rate exceeds approximately 2,000 ft per minute, and the flaps are
NOT in the landing configuration.

A Mode 2B warning will be activated if an excessive terrain closure rate occurs with the
flaps in the landing configuration. It will be activated at the upper parameter of 789 ft
radio altitude if the terrain closure rate is equal to, or exceeds, 3,000 ft per minute, or at
the lower parameter of 200 ft radio altitude if the terrain closure rate is equal to, or
exceeds, 2,250 ft per minute. The lower boundary cut-off varies between 200 ft and 600
ft and is dependent solely on the barometric rate. In this mode no altitude gain is also
required to silence the “TERRAIN” alert after the “PULL-UP” warning has been
activated. Additionally during the approach, with the landing gear and flaps fully
extended, the altitude gain function will be inhibited, and the “WHOOP WHOOP PULL-
UP” annunciation will be replaced by “TERRAIN -TERRAIN”.

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MODE 3. This mode will be activated if an excessive height loss is experienced during
the initial take-off climb or during a go-around procedure. The full operating parameters
for Mode 3 are shown below.

If the aeroplane enters the envelope a repetitive aural alert of “DON’T SINK” will sound
until a positive rate of climb is established between 30 ft and 700 ft radio altitude. The
alert will be repeated based on the original descent altitude if the aeroplane descends
again before climbing. This mode is only active during take-off or when either the flaps
or undercarriage are raised during a missed approach. When the aeroplane passes
through 700 ft radio altitude with the undercarriage retracted, Mode 4 will normally be
armed if an accumulated barometric height loss of 70 ft or more is sensed by the ADC.
An expanded upper altitude limit of 1333 ft radio altitude is however used at airspeeds
in excess of 190 knots in order to prevent premature switching from Mode 3 to Mode 4
warning during the climb-out, when the accumulated barometric height loss is in excess
of 128 ft. Mode 3 is also inactive when the landing gear and flaps are both in the
landing configuration.

MODE 4. This mode will be activated whenever an unsafe terrain clearance situation is
experienced, and the aeroplane is NOT in the landing configuration. Similar to Mode 2
this mode also has two sub modes; Mode 4A with the undercarriage retracted, and
Mode 4B with the undercarriage extended and the flaps not in the landing position.

Mode 4A will be activated whenever the terrain clearance reduces to 500 ft

radio altitude, Regardless of the barometric rate UNLESS the landing gear is
fully extended. The standard upper boundary for this sub mode is 500 ft radio
altitude, and at airspeeds less than 190 knots a repetitive aural alert of “TOO
LOW GEAR” will be sounded. At airspeeds between 190 knots and 250 knots
a repetitive aural alert of “TOO LOW TERRAIN” will alternatively be sounded,
and the upper boundary of the envelope will be extended to 1000 ft. The full
operating parameters associated with Mode 4A are shown below.

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Mode 4B will alternatively be activated UNLESS the flaps are also in the
landing position. The standard upper boundary for this sub mode is 245 ft radio
altitude and at airspeeds less than 159 knots a repetitive aural alert of “TOO
LOW FLAPS” will be sounded. The full operating parameters associated with
Mode 4B are shown below.

At airspeeds between 159 knots and 250 knots the upper boundary of the
envelope will be similarly increased to 1000 ft radio altitude and the aural alert
“TOO LOW TERRAIN” will be repeated.

MODE 5. This mode will be activated whenever the aeroplane falls significantly below
the ILS glide path by more than 1.3 dots. This mode has two specific alerting areas; soft
and loud, although the repetitive aural alert of “GLIDE SLOPE” will be sounded if either
area is entered. The ‘BELOW G/S’ lights will also be illuminated at the same time. The

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amplitude of the repetitive aural alert will additionally increase in volume on entering the
‘Loud Area’ and the repetition rate will increase as the glide slope deviation increases,
whilst the radio altitude decreases. The full operating parameters associated with Mode
5 are shown below.

This mode is armed whenever a valid signal is received by the Commander’s Glide
Slope receiver, and whenever the radio altitude is 1000 ft or below. This mode can be
inhibited or cancelled when flying below 1000 ft radio altitude by pressing either flight
crew ‘BELOW G/S’ light and will be rearmed whenever the aeroplane climbs above
1000 ft radio altitude or below the lower limit. Additionally the glide slope warnings may
occur at the same time as pull-up warnings when the ‘PULL-UP’ alert is due to an active
MODE l, 2, or 4, but NOT Mode 3. If Mode 3 is activated, Mode 5 will automatically be
inhibited.

Mode 7. This mode provides both aural and visual warnings, if a windshear condition
exists during the take-off and approach phases of flight, below 1500 ft radio altitude.
The following diagram shows the typical sequence of events during a windshear
encounter during the take-off phase of flight.

The aural warning consists of a two-tone siren, which is followed by an aural warning of
“WINDSHEAR, WINDSHEAR, WINDSHEAR”. A visual warning message is also

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displayed on each Electronic Attitude Director Indicator (EADI), and remains there until
the windshear condition subsides. The windshear warnings also take priority over all
other GPWS modes, and only become active on take-off after rotation.

Joint Aviation Requirements

The requirements state that:

¾ All turbine-engined aeroplanes of a maximum certified take-off mass in excess of

15000 kg or authorised to carry more than 30 passengers, for whom the individual
Certificate of Airworthiness is first issued on or after 1 July 1979, shall be equipped
with a GPWS.
¾ All turbine-engined aeroplanes of maximum certified take-off mass in excess of
5700 kg or authorised to carry more than 9 passengers shall be equipped with a
GPWS from 1 January 1999.
¾ A GPWS shall provide automatically a timely and distinctive warning to the flight
crew when the aeroplane is in a potentially hazardous proximity to the earth’s
surface.
¾ From 1 January 1999 a GPWS shall provide, as a minimum, warnings of the
following circumstances:-

¾ Excessive descent rate.

¾ Excessive terrain closure rate.
¾ Excessive altitude loss after take-off or go-around.
¾ Unsafe terrain clearance while in the landing configuration.
¾ Gear not locked down and the flaps not in the landing position.
¾ Excessive descent below the instrument glide path.

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Intentionally Left Blank

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Chapter 7.4
Traffic Collision Avoidance System

Introduction
With the ever-increasing traffic flow, the risk of an airborne collision has dramatically increased
and, in order to preserve the safety element, modern aeroplanes are fitted with equipment that
provides collision avoidance assistance. The current equipment used to achieve this function is
known as ‘TCAS’, which is an abbreviation for a ‘Traffic Alert and Collision Avoidance
System’, and provides traffic information within approximately 30 miles of the aeroplane. The
purpose and function of the TCAS, is to alert the flight crew to the presence of other aeroplanes
in their vicinity, and where possible provide an escape manoeuvre, should a collision risk arise.
This is achieved by equipment on board the aeroplane only, and without any reference to the
ground installations used by air traffic control. It is thus designed to complement, and not
replace conventional air traffic management methods.

The TCAS has the following levels of capability:-

TCAS I. This system provides a Traffic Advisory (TA) only, ie. information that would
advise the flight crew of a potential traffic hazard as an aid to visually acquiring the
target and avoiding it. The TA display shows the range and bearing of the aeroplane
posing a potential threat. This capability serves only as a warning, and simply provides
an aid to visual acquisition and avoidance, but does not recommend an escape
manoeuvre.

TCAS II. This system has the same capability as TCAS I, but is additionally capable of
providing an escape manoeuvre recommendation, called a Resolution Advisory (RA), in
the vertical plane only. Bearing information is also displayed, but only as an aid to
visual acquisition. The recommended escape manoeuvre is based on Mode C reports
from any conflicting traffic.

TCAS III. This system has the same capability as TCAS II, but is also able to provide
RA's and manoeuvre guidance in the horizontal plane.

TCAS II is the most common system and provides the necessary vertical manoeuvre advice. In
accordance with the Joint Airworthiness Requirements, as of the year 2000, all commercial
aeroplanes over 15000 kg, and with a seating capacity of 30 or more passengers, when
operating in European airspace, will be required to carry TCAS II. This system provides:-

¾ The generation of Traffic Advisories (TA)

¾ Threat detection
¾ The generation of Resolution Advisories (RA)
¾ Co-ordination
¾ Surveillance
¾ Communication with ground stations

The first four of these items is processed during each complete cycle of operation, and takes
approximately 1.2 seconds to complete.

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Aeroplane Installation
The equipment required to support a TCAS II is shown below.

The above equipment includes:-

¾ A TCAS Computer (transmit/receive unit)

¾ Two TCAS antennas; an upper directional antenna, and a lower antenna, which
may be directional or omni directional.
¾ A combined ATC Transponder/TCAS control panel
¾ A flight-deck display.
¾ A Mode S transponder(s) with top and bottom omni-directional antennas.

TCAS II receives inputs from:-

¾ The altimeter and /or the Air Data Computer (ADC) with regard to pressure altitude
¾ The Radio Altimeter in order to establish heights for the various operating
restrictions described later
¾ The gear and flap circuits which provide aircraft configuration status
¾ The FMC which provides operational performance data such as operational ceiling

Note: TCAS II is not connected to the autopilot, nor the FMS (this includes INS/IRS), and
remains independent of them, therefore it will continue to function in the event of the failure of
either of these systems.

TCAS relies on other aeroplane transponders to indicate their presence, so aeroplanes not
equipped with this system will appear transparent. If targets respond with Mode A only, it will be
assumed that they are at the same flight level as the interrogator, and only a traffic advisory will
be generated. TCAS-protected airspace is defined horizontally by time-to-convergence, and by
ATC vertical separation minima (400 to 740 feet depending upon the altitude).

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Traffic Advisories (TA) will appear on targets detected 30-45 seconds away from convergence
within the protected vertical envelope, and Resolution Advisories (RA) will appear on targets
20-30 seconds from convergence.

Operation of TCAS II
With reference to the diagram below the TCAS 11 system operates as follows:-

¾ Aeroplane A is TCAS equipped, and uses a Mode S transponder to transmit its

unique aeroplane identifier once per second in the form of an omni-directional
broadcast known as ‘Squitter’.
¾ Other TCAS equipped aeroplanes in the area will monitor the 1090 MHz
transmissions, and when a valid ‘Squitter’ signal is detected, the identity of the
transmitting aeroplane will be added to a ‘List’ or ‘Roll Call’ for subsequent
interrogation.

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¾ Aeroplane A will also compile a ‘Roll Call’ of its own, which will be derived from
‘Squitter’ from other TCAS equipped aeroplanes.
¾ The TCAS II transmitters of aeroplanes A and B will additionally transmit
interrogating signals to detect other aeroplanes in the area that are equipped only
with standard transponders. In this mode the TCAS transmitter will radiate a
general interrogation signal to all standard transponders. To separate replies when
a number of aeroplanes are in the vicinity, the transmitter has the following
features:

¾ The antenna design permits individual bearing transmissions to each of

the four quadrants in sequence.
¾ The transmissions consist of a series of pulses, starting at low power,
and progressively increasing to a higher power. This is known as
‘Whisper/Shout’, and causes close-by aeroplanes to respond before
more distant ones. TCAS is therefore able to track individual replies,
and also limit unnecessary overall energy radiation.

¾ Aeroplane C will reply to such interrogations, but the Mode S equipment on

aeroplane B will only respond to Mode S interrogations.
¾ Having compiled the roll-call aeroplane A will interrogate each aeroplane on the list.
The TCAS antennas then receive the replies, and this enables the relative bearings
of intruders to be determined. Should the intruder be able to respond with altitude,
this information will also be used by the TCAS. The range of intruders is
determined from the time interval between the interrogation signal and the reply.
¾ The altitude, altitude rate, range, and range rate are also determined by tracking the
replies from each interrogation. By computer analysis of the replies, the TCAS will
determine which aeroplane represents a potential collision threat, and will then
provide the appropriate advisory to the flight crew if any aeroplane is predicted to
pass within 1200 feet. Where multiple threats exist, each threat is processed
individually to produce the optimum avoidance solution.
¾ Where a Mode S signal is received and processed, if a collision risk is established,
the computer will establish an air-to-air Mode S data link with the TCAS II computer
on the other aeroplane. The computers will then ‘agree’ and co-ordinate
‘Resolution Advisories’ as necessary.

TCAS Aural Warnings

If the computer determines the likelihood of a conflict, an aural announcement “TRAFFIC,
TRAFFIC” will be generated. Should evasive action later be required, a Resolution Advisory
(RA) will follow a Traffic Advisory (TA), and will provide a verbal statement of the action needed,
eg. “CLIMB, CLIMB”. The RA’s may be:-

Corrective. This will instruct the flight crew to, “CLIMB” or “DESCEND” or
“DON’T CLIMB/DESCEND”.

Preventative. This will advise the flight crew to avoid certain manoeuvres in order to
maintain the separation between other aeroplanes. A selection of possible aural
resolution advisories is given in the following table.

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Aural Resolution Advisories

No. Advisory

1 Climb, Climb, Climb

2 Descend, Descend, Descend
3 Reduce Descent, Reduce Descent
4 Reduce Climb, Reduce Climb
5 Monitor Vertical Speed, Monitor Vertical Speed
6 Clear of Conflict
7 Climb, Crossing Climb, Climb, Crossing Climb
8 Descend, Crossing Descend, Descend, Crossing Descend
9 Increase Climb, Increase Climb
10 Increase Descent, Increase Descent
11 Climb-Climb Now, Climb-Climb Now
12 Descend-Descend Now, Descend-Descend Now
.
When TCAS determines that a threat has passed, a statement ‘CLEAR OF CONFLICT’ will be
generated.

Information Display
The action required to comply with the Resolution Advisory is typically displayed on a
specialised Resolution Advisory/Vertical Speed Indicator (RA /VSI) instrument using a colour
liquid crystal screen, which shows the rate of climb, and intruder information as shown on the
next page. On other aeroplanes the intruder information will be displayed on the EFIS
Navigation Display screen.

In either case, the symbology used will be the same:-

Hollow White or blue diamond. Other traffic not offering a threat

Solid White of blue diamond. Traffic within 6NM and 1200 feet vertically.
Solid Yellow Circle. Traffic advisory.
Solid Red Square, Resolution advisory.

Each symbol will also have a data tag attached to it, which will appear in the same colour as the
associated symbols.

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This data will provide the altitude of the target in reference to your own aeroplanes altitude,
eg.+ 06 will indicate 600 feet above. An arrow will also indicate whether the ‘target’ is climbing
or descending. There is also often provision to momentarily display absolute pressure altitude in
hundreds of feet of pressure altitude.

Where the target is a mode A transponder with no height transmissions no data tag will appear
and no RA will be given. If the target does not have a functioning transponder the TCAS
system will be unable to detect its presence.

The nominal maximum tracking range of TCAS is 14NM. However, in areas of high-density
traffic, the system range can be reduced to 5NM.

A RA will be generated between 15 to 35 seconds before the point of closest approach of the
intruder, and a TA will be generated 5 to 20 seconds in advance of the RA’s.

The TCAS equipment will be capable of handling a maximum surveillance capacity of

30 aeroplanes, but is nominally capable of surveillance of approximately 27 ‘high closing speed’
targets, within 14NM of the aeroplane.

Resolution Advisory / Vertical Speed Indicator (RA / VSI)

The required action to comply with the resolution advisory will usually be commanded by means
of a specialized vertical speed indicator, as showing on the next page. The TCAS VSI uses a
colour liquid crystal display to show conventional rates of climb, plus intruder information, which
is displayed as symbols in the centre. The two lower buttons will change the displayed range
between 4, 8, and 16 nautical miles. The aeroplane symbol will also appear 2/3 from the top. A
range ring is always present at 2 miles, and a second ring appears at 6 miles when the 16-mile
scale is selected.

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A Resolution Advisory is commanded by part of the VSI scale appearing in red, and the target
climb or descent rate as a green sector to show the preferred rate.

A panel dimmer also controls the display brightness, and an automatic sensor adjusts the
display for differing flight deck conditions.

TCAS Control Panel

A Mode S Control panel as shown below controls the TCAS.

F/L
ATC1

8888 T
XPDR
C
STBY
A FAIL A
XPDR
N TST S
TA
B TCAS
RA/TA

DIAGRAM 26.6 TCAS CONTROL

The control panel switch labelled A/N/B permits the selection of various altitude bands in which
intruders are displayed, which are as follows:-

A 7000 feet above to 2700 feet below

N 2700 feet above to 2700 feet below
B 2700 feet above to 7000 feet below

Note: This switch also has no effect on the generation of advisories.

Operating Restrictions
TCAS may be inhibited either totally or partially, to avoid conflicts with other handling
requirements, as follows:-

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¾ When either the ground proximity warning or windshear system is active to ensure
that these alerts take priority.
¾ Below certain altitudes:

¾ TCAS aurals are suppressed below 400 feet AGL

¾ All RA's are inhibited below 500 feet AGL
¾ Descend RA's are inhibited below 1000 feet AGL and sometimes 700 feet.
¾ Increase Descent RA's are suppressed below 1800 feet AGL.

¾ Climb or Increase Climb RA's may be inhibited above a defined barometric altitude.
¾ Climb and/or Increase Climb RA's are inhibited in those circumstances where the
manoeuvre cannot be safely executed due to performance limitations.
¾ The flight crew also has the capability through the control panel to select TA's only
in order to eliminate unnecessary RA's during some operations, eg. during parallel
approaches.

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Chapter 7.5
Mach/Airspeed Warning System

Introduction
The Mach/Airspeed Warning System provides a distinct aural warning if the maximum operating
speeds of VMO or MMO are exceeded. These airspeeds exist due to structural limitations at low
altitudes, and the handling characteristics of the aeroplane at higher altitudes.

System Architecture and Operation

The system consists of two independent Mach/Airspeed units, which are normally driven by
outputs from the Central Air Data Computer (CADC), via an internal mechanism comprising of
interconnected altitude, and airspeed capsules. A typical system layout is shown below.

If the limiting speed, as indicated by the bugs on the Mach/airspeed indicator is exceeded, an
audio warning (Clacker) will be activated, which can only be silenced if the airspeed is reduced
below VMO / MMO. The system can also be tested at any time by pushing a ‘Test’ function switch,
which will in turn activate the ‘clacker’. Notably this system provides the only audible warning
of ‘over-speed’.

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Maximum Operating Airspeed Schedule

The diagram on the next page shows a typical airspeed envelope, where VMO at sea level is
selected to be 340 knots (IAS). This airspeed will remain constant until the aeroplane reaches
approximately 26,000 ft, when the MMO limit of 0.82 is reached.

VMO will then continue to decrease with increasing altitude, if the selected Mach number is
maintained.

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Chapter 7.6
Stall Warning

Introduction
In order to warn the flight crew that the aeroplane is approaching a stall condition, a warning
device is fitted, which varies depending on the type of aeroplane.

Light Aeroplane Stall Warning Device

In addition to pre-stall buffet light aeroplanes are normally fitted with an audible stall-warning
device, which operates just before the stall. This device is activated by way of a stall warning
sensor in the form of a moving vane, which is positioned approximately midway along the wing,
just below the leading edge, as shown below.

They are held in their non-active position by static pressure, due to the forward motion of the
aeroplane, and the pressure distribution around the wing, with the stagnation point positioned
above the device. As the air speed reduces/angle of attack increases, the stagnation point will
moves below the device, and the spring-loaded flap is pushed upwards, as shown below.

When a pre-set deflection is reached, an electrical circuit will be completed and a warning horn
will be activated. The warning is usually a continuous sound that stops only when the angle of
attack has been reduced, ie. when the airspeed is increased.

Transport Category Aeroplane Stall Warning Device

On most modern transport aeroplanes the stall-warning device is activated via an ‘Angle of
Attack’ or ‘Alpha sensor’, as shown on the next page. One of these sensors is located either
side of the fuselage near the nose, and detects the change / rate of change of angle of attack of
the aeroplane as the airspeed varies. The sensor is shaped like an aerofoil surface and varies

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its position relative to the airflow. This results in an electrical signal being generated and when
a predetermined limit is reached, nominally 12 – 14°, a stall warning will be activated. These
devices are also heated, so that they remain operational throughout the flight.

The sensor is connected via a pivot to a digital stall warning system as illustrated below, which
is deactivated on the ground via a Squat switch on the nose wheel gear (weight on
undercarriage or air/ground safety sensor).
.

Two independent digital computers are normally installed, which compute the proper stall
warning based on the angle of attack, flap configuration, and thrust. The computers receive
inputs from the angle of attack sensor (Alpha probe), the flap position transmitter, the N1 and N2
indicators, the air/ground relay, the air data computers, and the leading edge module.

At a pre-set angle of attack the circuit will be completed and an eccentric weight motor attached
to the base of either, or both control columns will operate. The motor will in turn vibrate the
control column, thus imitating the effect of aerodynamic buffet, and alerting the flight crew
before a stall develops. This will also be visually displayed on the Primary Flight Display (PFD).

The stall warning system additionally incorporates a ‘test’ function facility, which when operated
will activate the stick shaker, by completing the circuit from the sensor to the motor.

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Chapter 7.7
Recording Devices

Introduction
Recording devices are installed in aeroplanes as part of the never-ending quest to prevent
accidents, since no transport system has a 100% safety record. The two root causes of all
incidents and accidents are:

¾ mechanical failure
¾ human error

If an aeroplane incident or accident occurs it is important to establish whether the cause was
the result of mechanical or human failure, or even a combination of both.

The investigation of a mechanical failure is made easier if a complete record of the behaviour of
every mechanical/ structural component is available right up to the instant of the occurrence.
To facilitate this extensive pre-flight and maintenance documentation for aeroplane is retained
on the ground. A complete record of the flight crew actions prior to flight is also maintained,
together with any relevant certification and training certificates, since this may also provide
invaluable information into the cause of the event.

To enhance the recording system certain aeroplanes are required to carry a ‘Flight Data
Recorder (FDR)’ and a ‘Cockpit Voice Recorder (CVR)’. The FDR is designed to record
mechanical features; whilst the CVR is designed to record all voice communications with, and
on the flight deck. It is also a requirement for the flight crew and ATC to keep in-flight records
although, in the event of an accident, the on-board documentary records could be destroyed.

Flight Data Recorder (FDR) Requirements

In accordance with JAR-OPS commercial transport category aeroplanes with a Certificate of
Airworthiness (C of A) first issued on or after 1st April 1998, with more than 9 passenger seats,
and a maximum take-off mass over 5700kg, should not be flown in a JAA Member State or
elsewhere unless it is equipped with a FDR. This device must also use a digital method of
recording and storing data, and be equipped with a method of readily retrieving that data from
the storage medium.

The parameters to be recorded may also vary according to the maximum certificated take-off
mass and age of the aeroplane, as specified in JAR-OPS 1 Section K.

For aeroplanes required to carry an FDR the following data must be recorded against a
common time scale:-

¾ altitude
¾ airspeed
¾ heading
¾ acceleration
¾ pitch and roll attitude

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¾ radio transmission keying, unless an alternative means is provided to enable the

FDR and CVR recordings to be synchronised
¾ thrust or power on each engine
¾ configuration of lift and drag devices
¾ air temperature
¾ use of Automatic Flight Control Systems (AFCS)
¾ angle of attack

For those aeroplanes with a maximum certified take-off mass over 27000Kg the following
additional parameters must also be recorded:-

¾ positions of the primary flight controls and pitch trim

¾ radio altitude
¾ primary navigation information being displayed to the flight crew
¾ cockpit warnings
¾ landing gear position

The FDR should furthermore:

¾ be capable of retaining the data recorded during at least the last 25 hours (10 hours
for aeroplanes of 5700 kg. or less) of its operation.
¾ be able to obtain accurate data correlation (matching) with information displayed, or
presented to the flight crew from aeroplane sources.
¾ must automatically start to record the data prior to the aeroplane being capable of
moving under its own power, and must also automatically stop after the aeroplane
is incapable of moving under its own power.
¾ be fitted with a device to assist in locating the recorder in water.

On some modern aeroplanes a large number of the parameters are taken from the aeroplane’s
integrated data source, and on some FDR models certain parameters can also be transmitted at
regular intervals through a data link with a ground station.

If the FDR is unserviceable the flight may still be conducted, as listed in JAR-OPS 1, if:-

¾ It is not reasonably practicable to repair or replace the unit before the

commencement of the flight.
¾ the aeroplane does not exceed 8 further consecutive flights with the FDR
unserviceable.
¾ not more than 72 hours have elapsed since the FDR was first reported
unserviceable.
¾ any CVR required to be carried is operative, unless it is combined with the FDR.

FDR Design
The FDR is powered from the aeroplane’s 24V DC or vital DC busbar and is contained in a
shockproof yellow or orange box, as shown below.

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The box is fitted at the rear of the aeroplane, normally behind the rear pressure bulkhead, and
must be capable of sustaining extremely high impact forces. The box must also be fireproof
and completely waterproof. The box is additionally fitted with an underwater locator (pinger) or
small beacon transmitter, which will enable it to be located in deep water, and it is also designed
to survive crash conditions. The transmitter has a self-contained power supply and is designed
to commence operation as soon as it enters water. The transmitter can operate continuously
for 30 days, and has a range of 2–3 miles.

Cockpit Voice Recorder (CVR) Requirements

In accordance with JAR-OPS commercial transport category aeroplanes with a Certificate of
Airworthiness (C of A) first issued on or after 1st April 1998, with more than 9 passenger seats,
and a maximum take-off mass over 5700kg should not be flown in a JAA Member State or
elsewhere unless it is equipped with a CVR.

The CVR must be capable of recording, with reference to a common time scale:

¾ voice communications transmitted from or received on the flight deck by radio.

¾ the aural environment of the flight deck, including without interruption, the audio
signals received from each boom or mask microphone in use.
¾ voice communications of flight crew members on the flight deck using the
aeroplane’s interphone system and public address system.
¾ voice or audio signals identifying navigation or approach aids introduced into a
headset or speaker.

The CVR should furthermore:

¾ be capable of retaining the data recorded during at least the last 2 hours (30
minutes for aeroplanes of 5700 kg. or less) of its operation.
¾ must automatically start to record the data prior to the aeroplane being capable of
moving under its own power, and must also automatically stop after the aeroplane
is incapable of moving under its own power. In addition, depending on the
availability of electrical power the recorder must start to record as early as possible
during the cockpit checks prior to engine start at the beginning of the flight, until
cockpit checks immediately following engine shutdown at the end of the flight.
¾ be fitted with a device to assist in locating the recorder in deep water.

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If the CVR is unserviceable the flight may still be conducted, as listed in JAR-OPS 1, if:

¾ it is not reasonably practicable to repair or replace the unit before the

commencement of the flight.
¾ the aeroplane does not exceed 8 further consecutive flights with the CVR
unserviceable.
¾ not more than 72 hours have elapsed since the CVR was first reported
unserviceable.
¾ any FDR required to be carried is operative, unless it is combined with a CVR.

CVR Design
The CVR is powered from the 24V DC or vital DC busbar, and like the FDR is located in a box
at the rear of the aeroplane. The box is orange or yellow in colour, and must also be impact
resistant, shockproof, fireproof and waterproof. Additionally in aeroplanes over 5700Kg, the
CVR must be a separate unit from the FDR.

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Chapter 8.1
General Engine Instrumentation

Introduction
Instrumentation is vital to maintain the safe and efficient operation of the aeroplane engine /
engines, and its associated systems. It varies according to the type of engine fitted, eg. piston,
turboprop or gas turbine, and in many cases the individual instruments are identical

Piston Engines
The level of instrumentation required for an aeroplane fitted with a piston engine is largely
governed by the complexity of the engine, for example the pilot of an aeroplane with a
supercharged piston engine will need to monitor the:-

Engine Speed. This is a measure of how much work is being done by the engine, and
is measured in revolutions per minute (RPM).

Induction Manifold Pressure, or Boost Pressure (MAP). This is a measure of the

engine power of a supercharged engine (absolute pressure). MAP is usually given in
inches of mercury (Inches of HG), and boost is normally measured in pounds per
square inch (PSI).

Torque or Turning Moment. This acts on the output shaft of the engine and is
proportional to the horsepower being developed. It is also sometimes used to provide
information for power control.

Cylinder Head Temperature. This temperature is important because excessive

temperatures can cause engine damage. The temperature is measured in degrees
Celsius (°C)

Lubricating Oil Pressure and Temperature. This ensures that the engine is
adequately lubricated. In pressure terms it maybe ‘HIGH’ or ‘LOW’, whereas the
temperature is usually given in Degrees Celsius (°C).

Fuel Flow. This provides a measure of the economy of the engine, and is measured in
pounds, kilogrammes or gallons / hour.

Fuel Quantity. This ensures that there is sufficient fuel to complete the flight, and is
measured as Fuel Mass or Volume.

Fuel Pressure. This is measured using a pressure gauge, and any drop in fuel
pressure may indicate a partially blocked fuel filter.

Turbo Propeller Engines

For an aeroplane with a turbo propeller engine the parameters which need to be monitored are
the:

¾ RPM
¾ Torque
¾ Engine Exhaust Gas Temperatures (EGT), which is measured in °C.
¾ Lubricating oil pressure
¾ Lubricating oil temperature
¾ Fuel flow
¾ Fuel quantity

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¾ Fuel Pressure.

Gas Turbine Engines

For an aeroplane with gas turbine engines the pilot must monitor the: -

Engine Pressure Ratio (EPR). This is the amount of useful thrust being developed by
the engine. It is the product of the mass of air passing through the engine and its
velocity at the exhaust nozzle minus the drag due to the air passing through the engine.
By comparing the air pressure across the engine, ie. the exhaust pressure to the
compressor inlet pressure (EPR), it provides an indication of the thrust output from the
engine. EPR is usually given as a % thrust value.

RPM, N1, N2 or N3. This is normally given as a % of its maximum value.

EGT. The exhaust gas temperature, which must be monitored in order to prevent
excessive heat damaging the turbine.

Oil Temperature and Pressure. These are monitored to ensure the safe operation of
the engine.

Fuel Pressure and Temperature. These are both monitored to ensure that a supply of
non-cavitated fuel is supplied at an acceptable pressure and temperature. A ‘low fuel
pressure’ warning light may back this up.

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Chapter 8.2
Pressure and Temperature Sensors

Introduction
Devices are fitted to aeroplanes that sense, or measure pressure and temperature, and then
create a signal, which is proportional to that measurement.

Pressure Measurement
In aviation, pressure is measured in pounds per square inch (PSI), inches of mercury (in. Hg),
millibars (mbar) or kilopascals (kPa). Pressure is also compared to some reference value, and
the three most common types are:-

Absolute Pressure. This is the pressure compared to a perfect vacuum, which is

either equal to, or greater than this value. It is given as PSIA or in. Hg absolute.

Gauge Pressure. This pressure is compared to ambient pressure, and is given as

PSIG.

Differential Pressure. This is the difference between two different pressures in an

aeroplane, and is given as PSID.

Aeroplane instruments which are used to register these pressures are typically the:-

Manifold pressure gauge. This gauge measures absolute pressure.

Oil pressure gauge. This gauge measures gauge pressure

Cabin differential pressure gauge. This gauge measures the difference in pressure
between the inside and the outside of the aeroplane, and is calibrated in PSID

Pressure measurements are required for various applications such as:-

¾ Static air pressure

¾ Fluid pressure
¾ Manifold pressure
¾ Differential pressure
¾ Pressure ratios

Pressures are usually measured by using a flexible metal chamber (aneroid capsule or
bellows), which is spring loaded against the effect of changes in pressure, or a Bourdon tube.

Aneroid Capsule. To measure static pressure, the capsule is partially evacuated

and then sealed, is prevented from collapsing by the action of a spring. The spring may
be fitted externally or, for some applications, may be fitted internally, as shown on the
next page.

If the pressure acting on the external face of the capsule is reduced, the spring will
cause the capsule to open, but if the external pressure is increased the effectiveness of
the spring will be reduced, and the capsule will collapse. This type of device is used to
measure relatively low pressures.

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Bellows. In some cases it is necessary to measure the difference between two sensed
pressures. One common use is to measure the differential pressure, where the bellows
are divided into two separate chambers, and a different pressure source is connected to
each side.

Expansion and contraction of the bellows will thus be dependent on the algebraic some
of these pressures. This type of device is used to measure medium pressures.

Bourdon Tube. This device is manufactured from a metal such as phosphor bronze or
beryllium-copper. It is in the form of a coil, as shown on the next page, and when
affected by a change in pressure will extend or contract. The Bourdon tube may be
used to measure oxygen pressure, hydraulic pressure, and engine oil pressure. This
type of device is used to measure relatively high pressures.

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In all of these types of pressure sensor, the change in pressure acting on them is converted into
a mechanical motion by the change in shape of the sensor. The sensors have an initial
resistance to any change in shape that results in a time lag between the time the pressure
changes, and the resultant change in shape.

Temperature Measurement
The temperatures that require to be measured on an aeroplane are:-

¾ Air temperatures
¾ Gas temperatures
¾ Component temperatures
¾ Fluid temperatures

The variation in the physical properties of a substance is used to measure temperature, and
any devices used on aeroplanes to measure temperature, are called ‘Temperature-Measuring
Systems’. Aeroplane temperature indicators give readings in degrees Celsius (°C) or in
degrees Fahrenheit (°F). To make a comparison the temperatures in the following paragraphs
will be given in °F.

Bi-metallic Temperature System. This system is used to measure temperatures

up to 140°F, and uses the property of expansion. Different materials expand and
contract at different rates when subjected to the same change in temperature. If two
thermally dissimilar metals, eg. iron and brass, are strapped together, and heat is
applied, one will expand more than the other, and the bi-metallic strip will distort.

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The bi-metallic strip can be formed into a coil so that any change in temperature will
cause the strip to wind or unwind, and produce a rotational movement. This motion will
then rotate a needle around a scale, and will display the temperature on a temperature
gauge. This system is commonly used on small aeroplanes, where it is normally
installed through the side window.

Mechanical Bulbs Temperature System. This system consists of a Bourdon

tube gauge, which measures pressure, a thin walled bulb, which is at the point of
measurement, and a thin capillary tube to connect them together, as shown below.

It uses the principle of the increase in vapour pressure within a confined space to
measure temperature. The system is filled with a chemical, eg. Methyl Chloride, which
in its normal state is part liquid and part gas. The system is sealed and as the
temperature increases the pressure changes within the tube, and gives an accurate
reading of temperature on the Bourdon tube gauge. This method is used on small
aeroplanes to measure engine oil pressure, and on some jet aeroplanes to measure the
compressor inlet temperature of the engine.

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Wheatstone Bridge System. This method of measuring temperature requires

electrical power, and is useful for measuring temperatures up to 300°F. The bridge
circuit consists of three fixed resistors and one variable resistor.

The variable resistor is the temperature probe, and contains a coil of fine nickel wire.
As the temperature of the coil increases, its resistance will also increase, and current
will flow in the bridge. This will in turn move the needle on the gauge, and will give a
temperature reading. A disadvantage of the Wheatstone bridge is that any bad
connections can drastically affect the resistance value, and produce an inaccurate
temperature reading.

Thermocouple Temperature System. This system is used to measure temperatures

up to about 500°F or more, and is the most commonly used device. Unlike the
Wheatstone bridge, this system does not require an electrical power source.

When one junction of two ‘dissimilar’ metals is heated a voltage proportional to the
temperature between the ‘hot’ and ‘cold’ junctions will occur, and current will flow in
the circuit, with a reading being taken at the cold junction. Two commonly used metals
are Copper-Constantan and Iron-Constantan, which are both able to withstand high
temperatures and produce a useable voltage. The actual voltage produced is very low,
so this device is not usually used to measure temperatures below 400°F. To measure
high temperatures up to 1000°F and above a combination of Chromel-Alumel is used.

Thermocouples are used on piston engines to measure the cylinder head temperature.
If only one probe is used it is fitted in the hottest running cylinder, eg. the rear cylinder
on a horizontally opposed engine. The probe is often in the form of a spark plug gasket,
which fits under the spark plug, or alternatively a bayonet type probe, which fits into a
special recess in the cylinder head. Thermocouples are also used to operate exhaust
gas temperature (EGT) gauges on jet engines, because high temperatures can severely
damage the turbine sections. They are fitted around the engine jet pipe and are

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connected in parallel, so that the failure of one thermocouple, will not adversely affect
the overall reading.

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Chapter 8.3
Pressure and Temperature Indicators

Introduction
Various pressure and temperature indicators are used on board an aeroplane, so that the
engine parameters can be effectively controlled

Pressure Indicators
A number of pressures on an engine have to be sensed. These include gas pressures and fluid
pressures, although all indicators use the same basic principle.

Manifold Pressure (MAP). These indicators, colloquially termed ‘boost gauges’, are
of the direct-reading type, and are calibrated to measure absolute pressure in inches of
mercury. This pressure is representative of that produced at the induction manifold of a
supercharged piston engine. In order to measure the pressure delivered by the
supercharger and obtain an indication of engine power, it is necessary to have an
instrument that indicates absolute pressure. This is measured between the throttle valve
and inlet valve. The mechanism of a typical indicator is illustrated below.

In this system the measuring element is made up of two bellows, one open to the
induction manifold and the other evacuated and sealed. A controlling spring is fitted
inside the sealed bellows and the distension of both bellows is transmitted to the pointer
via a lever, quadrant and pinion mechanism. A filter is located at the inlet to open the
bellows; where there is also a restriction to smooth out any pressure surges.

When pressure is applied to the open bellows the latter expands causing the pointer to
move over the scale (calibrated in inches of mercury) and indicate a change in pressure
from the standard sea level value of 29.92 in. Hg (‘zero `boost'). With increasing

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altitude, there is a tendency for the bellows to expand a little too far because the
decrease in atmospheric pressure acting on the outside of the bellows offers less
opposition. This tendency is counteracted by the sealed bellows, which also sense the
change in atmospheric pressure, but expands in the opposite direction. Thus a
condition is reached at which the forces acting on each bellow are equal. This cancels
out the effects of atmospheric pressure, and the manifold pressure is measured directly
against the spring. A typical manifold pressure gauge is shown below.

Engine Pressure Ratio (EPR). This measuring system consists of an engine inlet
pressure probe, a number of pressure-sensing probes projected into the exhaust unit of
the engine, a pressure ratio transmitter, and an indicator. The interconnection of these
components based on a typical system is schematically shown below.

The inlet pressure (P1) sensing probe is similar to a pitot probe, and is mounted so that
it faces into the airstream in the engine intake or, as in some power plant installations,
on the pylon, and in the vicinity of the air intake. The probe is also protected against
icing by a supply of warm air from the engine anti-ice system.

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Pipelines terminating at a manifold interconnect the exhaust pressure-sensing probes

(PEX) in order to average out the pressures. In some engine systems, pressure sensing
is done from chambers contained within the EGT sensing probes. A pipeline from the
manifold, and another from the inlet pressure probe, is each connected to the pressure
ratio transmitter. The indicators may be of the servo-operated type, but in electronic
display systems, the transmitter output signals are supplied directly to the appropriate
system computer. If a circuit malfunction occurs, an integrity monitoring circuit within
the indicator will activate a warning flag circuit, which will obscure the digital counter
display.

In some aeroplanes, a maximum allowable EPR limit indicator is also provided. It is

integrated with a TAT indicator and also with the CADC; its purpose being to indicate
limits related to air density and altitude values from which thrust settings have been
predetermined for specific operating conditions. These conditions are climb, cruise,
continuous and go-around. They are selected as appropriate by means of a mode
selector switch, which is connected to a computing and switching circuit, and generates
a datum signal corresponding to each selected condition. The signal is then supplied to
a comparator, which also receives temperature signals from the TAT sensor and
altitude signals from the CADC. These signals are compared with the datum signal and
the lower value of the two is automatically selected as the signal representing the
maximum EPR limit for the selected operating condition. The Comparator transmits this
signal to an amplifier, and then to a servomotor, which drives a digital counter to display
the limiting values.

Fuel and Oil Pressures. The oil pressure gauge is the most important gauge for
satisfactory engine operation. If the oil pressure fails, bearing failure will occur quickly.
The face of the gauge has a green arc showing the normal pressure range, a yellow arc
for the caution range and a red line for maximum oil pressure. It is important that the oil
pressure registers on the gauge within 30 seconds of the engine starting. Metal
capsules (or diaphragms) are normally used as sensors for measuring oil and fuel
pressures. These are connected by an electrical transmission system to the indicators
on the flight deck.

Temperature Indicators
The type of sensors, which are used to indicate engine associated temperatures are:-

Cylinder Head Temperatures. High temperatures are sensed in this area of the
engine so a probe capable of withstanding these values is used. This would typically be
a thermocouple, which can be in the form of a gasket beneath the spark plug, or a
bayonet type probe placed in the hottest part of the engine within a recess. The gauge
is positioned at the ‘cold’ junction

Exhaust Gas Temperatures. In a reciprocating engine aeroplane the EGT gauge is

used to manually lean the fuel-air mixture for better economy. Rich mixtures reduce the
engine exhaust gas temperature and weak mixtures increase it. Any adjustment of the
mixture using the EGT must be carried in the cruise and below 75% power. To facilitate
this a thermo-couple probe is installed in the exhaust pipe and connected to a simple
gauge on the flight deck. If the EGT gauge is used for leaning the mixture there is no
redline on the gauge, but in a turbo-supercharged reciprocating engine aeroplane the
gauge has a redline because the turbo-supercharger can be damaged by high
temperatures. In gas turbine engines a thermo-couple is also used to measure the
‘Turbine Gas Temperature (TGT)’, which is sometimes referred to as ‘exhaust gas
temperature’, or ‘Jet Pipe Temperature (JPT)’ This is because this type of engine can
be severely damaged by high temperatures in the turbine sections, which in some
regions can reach up to 1100°C. Thermocouples are used in parallel with multiple pick-
up points, so that the failure of one will not affect the reading

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Fuel and Oil Temperature. These are of the variable resistance type with the sensitive
element contained in a bulb and immersed in the fluid. This data is transmitted by a
Wheatstone bridge arrangement to a gauge on the flight deck.

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Chapter 8.4
RPM Indicators & Propeller Synchroniser Systems

Introduction
The measurement of engine speed is an extremely important parameter, which together with
manifold pressure, torque and exhaust gas temperature, allows the performance of an engine to
be accurately controlled.

The speed of a reciprocating engine is measured at the crankshaft, whilst with turboprop and
turbojet engines the rotational speed of the compressor shaft is measured, which gives a useful
indication of the amount of thrust being produced. These instruments are normally referred to
as ‘Tachometers’, and operate either mechanically or electrically.

In the case of aeroplanes fitted with multi-propeller installations the RPM’s are carefully
matched to reduce flight crew workload by automatically reducing the noise and vibration during
the cruise.

Tachometers
The main types of tachometer are: -

Magnetic Drag Tachometer. This type of tachometer is like a car speedometer and
is used on a light aeroplane. It uses a series of small permanent magnets, which are
rotated via a flexible shaft at half the engine speed, from a spur gear on the engine.

DIAL

INPUT SHAFT POINTER

DRAG DISC

HAIRSPRINGS
SERIES OF SMALL
MAGNETS

A highly conductive metal cup or disc (copper or aluminium alloy) is mounted on a shaft,
which is free to rotate in very low friction bearings, within the rotating magnetic field.
The shaft also carries a pointer, which is positioned by a calibrated hairspring so that it
registers zero when the magnets are at rest.

As the magnets rotate the resultant magnetic field will induce eddy currents in the disc,
which then interact with the magnetic field, and drag it along with it, hence the name
‘Drag Cup’. The eddy currents are such that the amount of drag increases
proportionally with speed, whilst hairsprings apply torque to the system. The torque
produced is proportional to the rotation of the drag cup shaft, and the pointer
correspondingly rotates over a linearly spaced dial, as shown on the next page. The

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cautionary operating range is clearly marked with a yellow arc, whilst a green arc
indicates the normal operating range.

Tacho-Generator and Indicator System. This system uses a remotely driven

tachometer. The detector (or pick-up) is an AC type generator, which consists of a
permanent magnet rotor that rotates within a slotted stator. The AC generator is bolted
directly to a mounting pad at the appropriate accessories drive gear outlet from an
engine, and the rotor is driven by a splined shaft coupling. In order to limit the
mechanical loads on generator, ratio gears are used in the engine drive system to
reduce the operating speed of the rotor.

The signal from the detector unit is passed through a synchro system to the indicator
unit. A typical indicator consists of two interconnected elements, a ‘driving element’
and a ‘speed-indicating element’.

Three wires connect the AC generator to the indicator unit, and as the permanent
magnet rotor is rotated within the stator, a three-phase AC supply, whose frequency
and voltage is proportional to the engine speed, is generated. The output from the
generator is then fed to directly to stator of an ‘AC three-phase Squirrel Cage

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Induction motor’, which in turn drives a cylindrical permanent magnet on a shaft. As

the magnet rotates it induces eddy currents into a drag cup, whose rotation is controlled
by a calibrated hairspring, which is attached to one end of the shaft. At the front end of
the shaft, a gear train is coupled to two concentrically mounted pointers; a large one
indicating hundreds of RPM, and a small one indicating thousands of RPM, as shown
below. The indicator may also be designed to read percentage of RPM.

Tachometer Probe and Indicator System. This system has the advantage of
providing a number of separate electrical outputs in addition to those required for speed
indication, eg. automatic engine control and flight data recording. It also has the
advantage that there are no moving parts, and thus will not be subject to high rotational
loads.

The stainless steel, hermetically sealed probe comprises of a permanent magnet, a

pole piece, and a number of ‘Cupro-Nickel’ or ‘Nickel-Chromium’ coils that are wound
around a ferromagnetic core, as shown below. Separate windings (from five to seven
depending on the type and application of a probe) provide outputs to the indicator and
other processing units requiring engine speed data.

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1. Body 5. Soft Iron Core

2. Potting 6. Spool (insulation)
3. Coil 7. Permanent magnet
4. Gear

This type of probe is mounted on the engine at a station in the high-pressure compressor
section so that it extends into this section. In some turbofan engines, a probe may also be
mounted at the fan section for measuring fan speed. The fan speed indicating system is, in
effect, a fan blade counting device. The sensor heads are mounted flush in the fan shroud
panel, and contain permanent magnets.

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The passage of each fan blade disrupts the magnetic field set up by the sensor magnets, and
causes an electrical signal pulse. The frequency of the pulses is equal to the number of blades
times the RPM, thus giving a signal frequency, which is proportional to the fan speed. The
signal is then amplified, conditioned and transmitted to the cockpit indicator in order to provide a
N1 readout in % RPM.

To ensure that the probe is correctly orientated, a locating plug is provided in the mounting
flange, as shown on the next page. The permanent magnet produces a magnetic field around
the sensing coils, and as the fan blades pass the pole pieces, the intensity of flux through each
pole varies inversely with the width of the air gap between the poles and the blades. As the
blades move, the air gap varies, and an EMF is induced in the sensing coils, the amplitude of
which varies with the rate of flux density change

The output signals (from the probe) then pass through a signal-processing module, and then
through a servo amplifier to a torque motor. This in turn rotates a pointer and indicates the
change in probe signals, in terms of speed. The servo potentiometer is supplied with a
reference voltage that controls the summation of signals to the servo amplifier and ensures that
signal balancing occurs at the various constant speed conditions. In the event of a power supply
or signal failure, the pointer of the indicator is designed to return to an `off-scale' position under
the action of a pre-loaded helical spring.

In this type of indicator, the indication of a power failure differs in that a flag is also energised to
obscure the counter display.

Propeller Auxiliary Systems

Propeller auxiliary systems include systems, which increase the efficiency of the propeller
operation, and provide automatic operation of the Constant Speed Unit (CSU) or Propeller
Control Unit (PCU), and feathering mechanisms. This increases safety, and reduces the
workload of the flight crew. The following systems may be found on either light twin engine
aeroplanes, or large turbo-propeller aeroplanes.

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Synchronisation System
The synchronisation system is used to set all of the engine CSU's/PCU's at exactly the same
RPM, thus eliminating excessive noise and vibration. It also avoids the need for the flight crew
to continually adjust the engine controls. A ‘Tacho-generator’ or a ‘Frequency Generator’ is
fitted to each engine, and these generate a signal proportional to engine speed. One of the
engines acts as the ‘Master’ engine, whilst the other engine(s) act as the ‘Slave’ engine(s).
The slave engines always maintain the same RPM as the master. On four-engine aeroplanes
any engine can be selected to act as the master, but on a light twin-engine aeroplane it is
always the left engine. The diagram below shows a typical two-engine synchronisation system.

A synchroniser control box is used to compare the RPM signal of the slave engine(s) to the
RPM signal of the master engine. The engine, which generates the higher voltage or frequency
will determine the direction in which the actuator will rotate, and will adjust the CSU/PCU spring
setting, which in turn will adjust the RPM. Generally the RPM of the slave engines must be
within approximately 100 RPM of the master engine for synchronisation to occur. This system is
used during all phases of flight, except for take-off and landing, when failure of the master
engine would result in all the engines attempting to follow the master engine.

Synchrophasing System
Synchrophasing is a refinement of Synchronisation, and allows the pilot to set the blades of the
slave engines a number of degrees in rotation behind the blades of the master engine.

This system is used to further reduce noise and vibration. The Synchrophasing angle can be
varied by the pilot to adjust for different flight conditions to achieve a minimum noise level.

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Operation of a Synchrophasing System

A pulse generator is fitted to the same blade root of each propeller, eg. No. 1 blade, and the
signals generated are used to ensure that all of the No. 1 blades are in the same relative
position at the same instant.

The pulse generator serves the same function as the tacho-generator does in the
synchronisation system. By comparison, when the signals from the slave pulse generators
occur in relation to the master engine pulse, the mechanism synchronises the phase
relationship of the slaves to the master engine. A propeller phase control in the cockpit then
allows the flight crew to select the phase angle, as shown below, which will give the minimum
amount of vibration.

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Intentionally Left Blank

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Chapter 8.5
Engine Torque Measurement

Introduction
The power produced by a propeller is proportional to the torque, which can best be described as
the turning moment produced by the propeller about the axis of the output shaft.

The power being produced by a turbo-propeller engine is indicated on the flight deck by a
torque meter. The torque meter system forms part of the engine itself, and is usually built in
with the reduction gear assembly between the output shaft and the propeller shaft.

Torque Meter
A typical torque meter system is illustrated below:-

In this system the helical gears, driven by the engine shaft, create an axial thrust, and oil
pressure, acting on a number of pistons, resists this axial thrust, which is in turn directly
proportional to the torque. This value is then transmitted to a suitably calibrated indicator dial.
One advantage of this type of torque measuring device, as fitted on a turboprop aeroplane, is
that the system can also be used to operate the propeller-feathering device if the torque meter
oil pressure suddenly reduces due to a power failure. On some aeroplanes it is additionally
used to automatically operate the water injection system that boosts the take-off power when
operating at high altitude/ high temperature aerodromes.

The system on the next page is alternatively based on the principle of the tendency for some
part of the reduction gear to rotate, and is resisted by pistons working in hydraulic cylinders,
which are secured to the gear casing. The pressure being created by the pistons is then
transmitted to a pressure gauge on the flight deck.

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Torque may also be sensed by an electrical strain gauge system. A strain gauge is a fine
insulated conductor wire bonded to a component. When strain is applied, ie. deflection under
load, the resistance of the wire changes, and this can be measured in a ‘Wheatstone Bridge’
circuit.

Negative Torque Sensing

The Negative Torque Sensing (NTS) system, as shown on the next page, is designed to
prevent the propeller driving the engine, eg. in the in-flight idle power setting, during a lean fuel
schedule, or temporary fuel interruption, or due to air gusts momentarily acting on the propeller.

The NTS device takes its signal from the torque indicating system, and at a specified torque the
NTS will activate the propeller control system, which will cause the propeller blades to coarsen
until the torque value rises above a specified value. This system is therefore a blade
coarsening system, and not necessarily a feathering system.

During negative torque the gear ring will close the NTS valve, and oil pressure will be fed to the
feather valve, which in turn will force it to the right. Oil will then be released from the propeller
hub, and will allow the feathering spring to coarsen the blades. When positive torque is
reapplied to the gear wheel, the NTS valve will open, thereby releasing the oil, which will allow
the feather valve to move to the left, and will allow normal propeller operation to be resumed.

When the NTS is activated the torque meter will fluctuate, which will tend to indicate engine
failure, and manual feathering should be carried out.

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Intentionally Left Blank

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Chapter 8.6
Vibration Monitoring

Introduction
A turbo jet engine has a very low vibration level, and any change in vibration, could be the first
indication of an impending problem, which could easily go unnoticed. Any problems may be
caused by: -

¾ A damaged compressor blade

¾ A turbine blade that has a crack or is subject to `creep',
¾ An uneven temperature distribution around the turbine blades and rotor discs may
be set up.

Any of these problems could result in an unbalanced condition of the main rotating assemblies,
and could lead to possible disintegration. In order to provide the flight crew with a timely warning
of increased vibration, jet engines are fitted with vibration monitors, which continually monitor
their vibration levels. On a multi jet aeroplane a monitor is fitted to each engine, and these are
designed to indicate if the maximum amplitude of vibration of the engine exceeds a pre-set
level. These indicators are located within the control group of instrumentation and are usually
milli-ammeters that receive signals through an amplifier from engine-mounted transmitters, and
are displayed as relative amplitude.

Vibration Monitoring System

The monitor is mounted on the engine casing and consists of a vibration pick-off (sensor), which
is mounted at right angles to the engine axis, an amplifier-monitoring unit, and an indicator
calibrated to show vibration amplitude in thousandths of an inch (mils) as shown below.

The sensor (transducer) is a spring-supported permanent magnet, which is suspended in a coil

attached to the interior of the case. As the engine vibrates, the sensor unit and core move with
it, but the magnet, tends to remain fixed in space because of its inertia. The motion of the coil
causes the turns to cut the field of the magnet, thus inducing a voltage in the coil and providing
a signal to the amplifier unit. The signal, after amplification and integration by an electrical
transmission system, is then fed to the indicator via a rectifying section.

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An amber indicator light also forms part of the system, together with a test switch. The light is
supplied with dc from the amplifier rectifying section and comes on if the maximum amplitude of
vibration exceeds a pre-set value. A test switch also permits functional checking of the system's
electrical circuit. In some engine installations, two sensors may be fitted to an engine: for
example, in a typical turbofan engine, one monitors vibration levels around the fan section, and
the other around the engine core section. A system of filters in the electrical circuit to the gauge
makes it possible to compare the vibration is recorded against a known frequency range, and so
enables the source of vibration to be located. A multi-selector switch enables the flight crew to
select a specific area, and to obtain a reading of the level of vibration.

In systems developed for use in conjunction with LCD and CRT display indicators, the vibration
sensors are of the type whereby vibration will cause signals to be induced in a piezoelectric
sensor stack.

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Chapter 8.7
Fuel Gauging

Introduction
Fuel quantities and flows are essential measurements for the safe and efficient conduct of a
flight. Fuel quantity is of great interest to the flight crew, since there must be sufficient fuel on
board to complete the flight, and also maintain adequate reserves. A check on the amount of
fuel remaining against the planned quantity at significant route points will reveal any excessive
fuel burn. Fuel quantity may also be given in volume or mass.

Measurement of Fuel Quantity

The following types of fuel contents gauges exist:-

Float Type. This type of gauge may be either electrical or mechanical. In a direct
reading mechanical fuel gauge a float drives a pointer directly, via gearing and a
magnet. The pointer rotates over a dial, which is calibrated in units of volume, eg.
gallons, and is separated from the fuel.

The electrical gauge is similar to the mechanical type, except that in this system the
float is arranged so that it drives an electrical design type transmitter, which enables the
gauge to be positioned remotely from the fuel tanks.

Ratiometer Type Fuel Gauge. This type of system is most commonly used on light
aeroplanes and indicates the quantity of fuel by volume in a fuel tank.

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As the level of the float changes, it moves a wiper arm, which alters the resistance of a
variable resistor, and changes the overall resistance in the circuit. This in turn will alter
the current flow in a DC circuit and will give an indication on a gauge similar to a
voltmeter or ammeter. The gauge uses two opposing magnetic fields and a pointer,
which react to the ratio of the current flow in the two sections.

The float and resistance systems will only be accurate when the aeroplane is in steady
straight and level flight. Additionally if the aeroplane accelerates or the temperature
changes the volume of fuel indicated on the gauge will also vary. This type of gauge
will also be subject to inaccuracies if the voltage fluctuates, which may be caused by
voltage regulator settings or a weak battery.

Capacitance Type of Fuel Gauge. This system is the most common type used on
modern jet engine aeroplanes, and uses an electronic fuel-measuring device, which
indicates fuel quantity. The quantity is measured in mass or weight in pounds or
kilograms, and not volume. The principle used in this type of gauge is that the
capacitance of a capacitor of fixed dimensions is dependent on the dielectric between
the plates. The tank units that form the plates of a capacitor consist of two concentric
tubes, as shown below.

The space between the plates is filled with fuel, air, or a mixture of both, forming the
dielectric of the capacitor; thus the capacitance of the tank units will be directly proportional
to the amount of fuel in the tanks. The dielectric constant of jet engine fuel compared to air
is approximately 2:1. A simplified capacitance bridge circuit is shown below.

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If the two capacitances are equal, the voltage drop across them will also be equal, and the
voltage between the centre tap and point P will be zero. If the fuel quantity increases the
tank unit’s capacitance will similarly increase, and the bridge circuit will become unbalanced.
A voltmeter will therefore indicate a voltage, which is proportional to the change in
capacitance, and hence fuel mass. This is useful with jet engines because the amount of
thrust being produced by them is more a factor of the amount of mass consumed, rather
than the volume of fuel consumed. This system also enables accurate readings to be
produced in large or irregular shaped fuel tanks. A number of probes are normally fitted in
each tank, depending on its size, and are connected in parallel, which ensures that the
indications remain the same regardless of the attitude or wing flex of the aeroplane. The
capacitors must also be matched or characterised to their specific locations, and the sum of
their capacitance gives a measure of the actual quantity of fuel in the tanks. The tank units
are connected to an amplifier in place of a voltmeter, and the output drives a pointer, which
shows the total mass of fuel in the tank.

Any changes in temperature will also affect the density of fuel, so the volume occupied by a
given quantity of fuel will increase if the temperature increases, and its density will fall. To
compensate for this error, a balancing (short) capacitance unit is installed at the lowest part
of the tank where it will be totally immersed whenever there is a useable quantity of fuel in
the tank, as shown below.

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By totally immersing this unit, the only variation in the capacitance it will measure will be due
to changes in the fluid density. This variation is thus used to adjust the calibration of the fuel
quantity probes.

If there is a significant amount of water in the tank, it will lie at the bottom of the fuel tank
and will also affect the compensating or balancing probe, which will result in an incorrect
adjustment of the calibration. Foe example if the tanks were drained and then refilled to the
same level with water, the pointer will tend to move up the scale to a new point of balance.
This is because water has a greater density and dielectric constant than aeroplane fuel

This method of quantity measurement may also be used for other fluids, eg. Hydraulics.

Fuel Totaliser
This device digitally displays the amount of fuel in each tank on a flight deck indicator and
provides an indication of the total amount of fuel on board the aeroplane, as shown below.

On some aeroplanes the Totaliser is combined with a gross weight (mass) indicator and uses
two digital scale windows as indicated.

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The upper scale shows the sum of the individual tank readings and the lower scale shows the
aeroplane’s gross weight, which is normally manually set after it has been loaded. As the fuel is
used, the fuel remaining, and the weight scales will both reduce.

On some indicators the flight crew may also be able to enter the zero fuel weight (ZFW) at the
commencement of the flight, so that the gross weight can be displayed throughout the flight.

Fuel Flow
Fuel flow metering systems are designed to provide the crew with a continuous indicationthe of
the instantaneous rate of fuel flow to each engine and in some instances the amount of fuel
consumed. Fuel flow is measured at the fuel intake of each engine (LP fuel supply line on a gas
turbine). If the fuel flows increases, for a particular engine power, it will indicate a reduction in
efficiency and probably an impending mechanical problem.

The basic method of measuring fuel flow is by a rotating vane flow meter, as shown on the next
page. In this system, the shaft carrying the vane will form the transmitter of a synchro
transmission system. The receiver rotor is attached to, and drives a needle moving against a
scale calibrated in fuel used per unit time, which can be either volume or mass related.

A more complex system, which is in common use in large aeroplanes, is the integrated flow
meter system, as shown below.

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In this system the fuel flow to the engine is measured, and is not only presented on a suitable
scale, but is also passed to an integrator where it is processed (integrated) with respect to time
to obtain the amount of fuel used. The transmitter/sensor unit consists of a tube, narrowed at
the ends and fitted into the appropriate engine fuel supply line. Within the tube there is a motor
driven impeller through which the fuel passes, which makes the fuel swirl at a rotation rate that
varies with the flow rate. On leaving the impeller, the swirling fuel impacts a receiver turbine
and induces a rate of rotation of that turbine that is directly proportional to the swirl rate. This
rotation is electronically detected and transmitted to the fuel flow indicator via a synchro system.
Signals are also sent via the integrator, to the fuel-consumed indicator.

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Chapter 8.8
EICAS

Introduction
The Engine Indication and Crew Alerting System (EICAS) displays primary engine indications
and also provides a centrally located crew alerting system for non-normal situations. The
system also shows the status of systems not otherwise displayed on the flight deck.

On the ground, EICAS additionally provides maintenance personnel with a variety of system
data.

EICAS Architecture
Two EICAS computers receive inputs from engine and system sensors. The information from
the sensors is displayed on two Cathode Ray Tubes (CRTs) as dials and digital readouts of
warnings, cautions and advisory messages.

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The parameters that need to be set and monitored by the flight crew are permanently displayed
on the screen, but the system also monitors the remaining parameters and displays them only if
one or more exceed safe limitations.

A Computer Selector panel determines which computer controls the EICAS. When the selector
is in the AUTO position the left computer is used, and if the left computer fails, control will
automatically switch to the right computer. When the L position is selected only the left
computer can control the EICAS and when the selector is in the R position only the right
computer can control the system. The EICAS computers monitor over 400 inputs through a
comprehensive warning and caution system, to provide a quick and unambiguous identification
of problems as they arise.

The Brightness and Balance Controls are used to adjust the brightness level of both CRTs.

The Event Record Switch is used to store systems data in an EICAS memory for later use by
maintenance personnel. When the switch is pushed, current data from the engine and system
sensors is recorded, and any previously recorded data is erased from the memory.

System lights and a Standby Engine Indicator (SEI) provide backup indications for the CRT
displays.

Engine Displays
The CRT screens are located in the centre of the instrument panel above the throttle pedestal
and are typically displayed as follows:-

¾ Primary engine indications appear on the upper screen.

¾ Secondary Engine Indications appear on the lower screen.

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In the event of an upper or lower CRT failure all engine indications will be automatically
transferred to the other CRT, and the display will appear in a compacted format.

This allows the flight crew to still retain a full capability in the event of a CRT failure, or if the
lower CRT is being used for status displays (or maintenance displays on the ground). In the
compacted mode, the primary engine indications are shown in their normal format, whilst the
secondary information, and oil system indications are shown in a digital format.

Crew Alerting
The crew-alerting portion of EICAS continually monitors all of the aeroplane systems, and if a
fault occurs, or any system fault light illuminates in the cockpit, the EICAS will display a crew-
alerting message on the upper CRT. In addition to the display messages, some crew alerts are
also indicated by aural tones and Master Warning/Caution lights.

All ‘crew alert’ messages are divided into one of three categories:-

Warnings (Level A). These are indicated, in red, and reflect an operational or
aeroplane system condition that requires immediate crew awareness and prompt
corrective action. These are the most urgent types of crew alert, of which an engine fire
is a typical warning.

Cautions (Level B). These are indicated in amber, and reflect an operational or
aeroplane system condition that requires immediate crew awareness and future
compensatory action. These are less urgent than warnings; of which an engine
overheat is a typical caution.

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Advisories (Level C). These are indicated in amber, and reflect an operational or
aeroplane system condition only for crew awareness that requires corrective action on a
time available basis. Advisories are the least urgent type of crew alert. A yaw damper
fault is a typical advisory.

If a parameter goes out of tolerance, an alert, status or maintenance message will automatically
be generated, depending on the urgency of the malfunction.

Master Warning/Caution Light

Two Master Warning Lights illuminate when a warning occurs, and they remain on as long as
the warning exists, or until either Master Warning/Caution reset switch is pressed.

Pushing the reset switch will silence the fire bell and cabin altitude siren, and may also silence
the landing configuration siren, depending on the reason for its activation.

Inhibits
Parts of the crew alerting system are inhibited or deactivated during certain phases of the flight
to prevent distractions, eg. the gear light will illuminate as soon as the landing gear begins to
retract, but the ‘GEAR DISAGREE’ message is inhibited for 25 seconds to allow their normal
stowage.

Display Messages
Crew alerting messages appear on the upper CRT to indicate all non-normal conditions
detected by EICAS, where up to 11 messages can be displayed. If more than 11 messages are
generated the last message is removed and replaced by a page number. Page 2 can be
displayed by pressing the CANL (CANCEL) switch, and page 1 can be recalled by pressing the
RCL (RECALL) switch.

Some typical examples are shown below.

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Warnings are indicated by red messages at the top of the message list, and are also arranged
according to their urgency, and order of occurrence.

Cautions appear as amber messages below the lowest warning.

Advisories appear below the lowest caution and are also indicated by amber messages. They
are also indented by one space so that they can be distinguished from cautions.

The most recent warning, caution and advisory message appear at the top of its respective
group of messages.

A message is automatically removed from the display when the associated condition no longer
exists, and then all messages that appeared below the deleted message will move up one line.

If a new fault occurs, its associated message is inserted on the appropriate line of the display,
which may cause older messages to move down one line. For example, a new caution message
would cause all existing caution and advisory messages to move down a line.

If there are more messages than can be displayed at one time, the lowest message will be
removed, and a white page number will appear on the lower right side of the message list.
Messages ‘bumped’ from the bottom of one page will automatically appear on the next page.

The Cancel and Recall switches are used to manipulate the message lists. Pushing the Cancel
Switch will remove the caution and advisory messages from the display.

Warning messages cannot be cancelled.

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If there is an additional page of messages, pushing the Cancel Switch will display the next
page. Warning messages are carried over from the pervious page, and when the last page of
messages is displayed, pushing the switch once more, will remove the last caution and advisory
messages, together with the page number.

Pushing the Recall Switch will display the caution and advisory messages that were removed
with the Cancel Switch if the associated faults still exist. If there is more than one page of
messages, page one will be displayed. A white RECALL message will also appear for about
one second on the lower left side of the message list to indicate that the Recall Switch has been
pushed.

New display messages appear on the page being viewed. For example, if page three is
selected and a new caution occurs, the caution message will appear on page three below any
warning messages. If the Recall Switch is subsequently pushed, the new caution message will
appear as the top caution message on page one.

Multiple display messages of a similar nature are sometimes replaced by a single, more general
display message. For example, if only the forward or aft entry door is open on the left side, a
‘L FWD ENT DOOR’ or ‘L AFT ENT DOOR’ message will appear. If both doors are open, only
a ‘L ENTRY DOORS’ message will appear.

Status
The status portion of EICAS is used to determine the aeroplane’s readiness for dispatch, and if
the STS (Status) Switch is pushed, the status display will appear on the lower CRT, as shown
on the next page.

The status display includes system indicators, flight control position indicators, status messages
and brake temperature indicators.

System indicators will appear in the top left corner of the display, and will show hydraulic
quantity and pressure, APU EGT, RPM and oil quantity; and oxygen pressure.

The flight control positions of the rudder, ailerons and elevators will also appear in the bottom
left corner of the status display.

White, status messages will additionally appear on the right side of the status display. These
messages will indicate any equipment faults that require awareness at dispatch, particularly
those that are not otherwise shown on the flight deck.

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Status messages are arranged by order of occurrence, and the most recent status message will
appear at the top of the list.

A message will be automatically removed from the display when the associated condition no
longer exists, and messages, which appear below the deleted message, will each move up one
line.

If a new status fault occurs, its associated message will be inserted at the top of the list, and all
other messages will move down one line.

If there are more messages than can be displayed at one time, the lowest message will be
removed and a white page number will appear on the lower right side of the message list.
Messages bumped from the bottom on one page will automatically appear on the next page.

If there is an additional page of status messages, pushing the STS (Status) Switch will display
the next page. When the last page of messages is displayed, pushing the switch will remove the
status display from the lower CRT, and will blank the screen.

New status messages will appear at the top of the page being viewed. If the status display is
deselected and subsequently reselected the message list will be reordered, with the newest
status message now appearing at the top of the first page.

A Status Cue will appear in the left upper corner of the lower CRT if a new status message
occurs, whilst the status display is not currently selected, and the aeroplane is in the air. The
cue will disappear if the status page is displayed. Status messages do not need to be checked
in flight; however, they can be useful in anticipating possible ground maintenance actions.

Brake temperature indications will additionally appear on the lower right side of the status
display.

Maintenance
The maintenance portion of the EICAS provides a flight deck display of system data for use by
maintenance personnel. Maintenance displays can only be used on the ground and are
designed to provide flight deck display of maintenance information for the use of flight deck
crew post flight logbook entries and ground crew. For convenience all status messages are
repeated on the Maintenance page, and also any significant information not covered by the alert
messages. Maintenance messages are displayed on the right hand side of the lower CRT by
pressing the ‘ECS/MSG’ button on the EICAS maintenance panel when the aeroplane is on the
ground.

EICAS Failure Modes

If a CRT fails, status can only be displayed on the ground. If the EICAS Control Panel fails an
EICAS CONT PNL advisory message will be displayed, and the EICAS full up engine mode will
also be automatically displayed. The cancel and Recall Switches will not operate when the
EICAS Control Panel fails. In the event of the failure of both EICAS computers, or both CRTs, a
Standby Engine Indicator (SEI) will be automatically activated. The SEI system lights and
system indicators are used to monitor the engines and system operation if a total EICAS failure
occurs.

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Chapter 8.9
ECAM

Introduction
The Electronic Centralised Aircraft Monitoring (ECAM) System is part of the electronic
instrument system (EIS) consisting of six screens, and provides the flight crew with assistance
in system management in both normal / abnormal conditions. This operational assistance is
given by the EFIS system plus two centre mounted CRT's identified as the:-

¾ Engine / Warning Display (E/W).

¾ System / Status Display (S/S).

The physical layout of both CRT displays is dependent upon the flight deck layout. The layouts
on the A320 and MD11 are respectively shown below.

Engine / Warning (E/W) CRT Display

The E/W Display as shown on the next page shows the:-

¾ Engine parameters.
¾ Fuel on board.
¾ Slats and flap position.
¾ Warning and caution messages.
¾ Memos when no failures exist.

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The System / Status (S/S) CRT Display

The S/S Display as shown above shows the:-

¾ System synoptic diagrams.

¾ Status messages.

ECAM System Architecture

A typical ECAM system is shown below.

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In this system information is derived from the various aeroplane systems, and after processing
in the following units, is displayed on the ECAM Display Units.

Flight Warning Computers (FWC’s). These units generate all of the alert messages,
aural alerts, and vocal messages (Radio Height), which are derived:-
¾ Directly from aircraft sensors or systems to generate red warnings.

¾ Through the system data acquisition concentrators (SDACs) to generate

amber cautions.

Display Management Computers (DMC’s). These units are common to EFIS/ECAM

and generate the images displayed on the PFD, ND, E/W and S/S display units. They
thus provide a similar function to that of the symbol generators in an EFIS system.
System Data Acquisition Concentrators (SDAC’s). These units acquire data,
process and then distribute it to the:-
¾ DMCs, for the display of engine parameters and system page data.

¾ FWCs, for the generation of alert and procedure messages.

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ECAM Control Panel (ECP)

The ECP shown below allows the system pages or the status (STS) page to be selected prior to
them being displayed on the S/S Display.

The ECP has the following functions:-

System Page Push Switches. If these switches are depressed the S/S display
changes to the corresponding system page, and also illuminates following a manual
selection, or when an advisory is detected.

STS Push Switch. If this switch is depressed the status page will be displayed on the
S/S display, but if there is no status message present, ‘NORMAL’ will be displayed on
the CRT for 5 seconds.

CLR Push Switch. This switch will illuminate if a warning/caution message on the
ECAM display unit requires flight crew action or acknowledgement. Whilst the CLR
switch is illuminated, pressing it will change the ECAM display until all actions required
by the ECAM system have been carried out.

ALL Push Switch. If this switch is depressed, all of the system pages will be
successively displayed at one second intervals. This switch will also allow the system
pages to be successively presented in the case of ECAM control failure, and will allow
the flight crew to stop at the desired page.

EMERG Cancel Push Switch. If this switch is depressed any present:-

Aural Warning (including GPWS) will be cancelled for as long as the failure
condition is present.

Caution (CRT message + master caution + single chime) will be cancelled for
the remainder of the flight. The status page will then be automatically displayed
and the ‘CANCELLED CAUTION’ message displayed, together with the title of
the failure, which was inhibited.

TO CONFIG Switch. If this switch is depressed a take-off power application will be

simulated, and a warning will be triggered if the aeroplane is not correctly configured for
take-off.

OFF / BRT Knobs. These knobs are used to control the on/off and brightness of each
ECAM display unit.

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Attention Getters
In abnormal operation, the ECAM system will direct the flight crew’s attention by visual and
audio attention getters. The visual attention getters are the:-

Master Warning Light (red). This gives an indication of any system faults, which
require immediate crew awareness and immediate corrective action.

Master Caution Light (amber). This gives an indication of any system faults, which
require immediate crew awareness, but not immediate corrective action.

Display Colour Coding

The displays are colour coded as follows:

Red Warnings
Amber Cautions
Green Normal long term operation
White For functions not in normal operation (switch in off position)
Blue For action to be carried out

ECAM System Failure

If a system failure occurs, the flight crew can manually switch the display functions to other units
within the electronic instrument system, and still maintain full operational capability as follows.

Failure Categorisation
The following categories of failure can occur on an aeroplane:-

Independent Failure. This type of failure only affects an isolated item of equipment or
a system, and will not affect any other systems in the aeroplane.

Primary Failure. This type of failure will affect an item of equipment or a system, and
will result in the loss of other systems in the aeroplane.

Secondary Failure. This type of failure will result due to the primary failure of an item,
or system.

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System Operation
The ECAM system provides warnings and cautions whenever the aeroplane flies out of its flight
envelope, or the failures affect the aeroplanes integrity. These warnings and caution messages
are written in plain English, and:-

¾ Provide improved failure analysis.

¾ Provide guidance for any corrective action required.
¾ Minimise the need for the flight crew to refer to paper checklists.
¾ Improve the understanding of the aeroplane and system configuration after failure

In normal operation the ECAM system eases the flight crew workload by eliminating the need
for frequent scanning of the various aeroplane system panels. For example, if the ‘SEAT
BELT’ or ‘NO SMOKING’ signs are on, or the ‘APU’ is running, they will be displayed on the
E/W display unit as a memo message. Checklists will be additionally displayed on the memo
page during the take-off and approach phases of flight. Synoptic diagrams are routinely
presented on the S/S display unit, and are automatically adapted to the current flight phase, or
can be manually called up by the flight crew, using the ECP push buttons. The synoptic pages
are used to amplify system failures and switch selections that have been made in conjunction
with warnings displayed on the E/W display. Some system parameters are also continually
monitored throughout the whole flight, and are automatically displayed on the relevant system
page when their values drift out of the normal range, but well before the warning level is
reached.

In the case of a failure being detected by the ECAM system the following actions will occur:-

¾ The E/W display will present the warning and caution messages. The warning
message will include a failure title and any associated procedures, as well as the
title of the systems affected by the failure.
¾ The master warning or master caution lights will illuminate.
¾ An audio alert will be triggered.
¾ The S/S display will present the affected system page.
¾ The CLR push switch will illuminate on the ECP.
¾ When all of the messages and synoptic diagrams associated with the failure have
been actioned by the flight crew, the CLR push button must be depressed until the
normal configuration returns, and the CLR push button is extinguished.

The sequence of events, which occur during a typical failure are detailed on pages 8.9.7–
8.9.13 inclusive.

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Step 1

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Step 2

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Step 3

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Step 4

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Step 5

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Step 6

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