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INSTRUMENTS

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CONTENTS

PITOT STATIC SYSTEM………………………………………………………………………………..3

AIR TEMPERATURE MEASUREMENT…………………………………………………...................4
THE AIR SPEED INDICATOR (ASI) ………………………………………………………………….5
THE PRESSURE ALTIMETER…………………………………………………………………………8
THE VERTICAL SPEED INDICATOR…………………………………………………………….…14
THE MACHMETER…………………………………………………………………………………….16
AIR DATA COMPUTER……………………………………………………………………………......19
TERRESTRIAL MAGNETISM………………………………………………………………………...20
COMPASS………………………………………………………………………………………………..21
GYROSCOPE…………………………………………………………………………………………....25
DIRECTIONAL GYRO INDICATOR (DGI)………………………………………………………....28
ARTIFICIAL HORIZON…………………………………………………………………………….....31
TURN AND SLIP INDICATOR…………………………………………………………………….….35
TURN COORDINATOR…………………………………………………………………………….….36
REMOTE INDICATING MAGNETIC COMPASS………………………………………………......38
INERTIAL NAVIGATION SYSTEM…………………………………………………………….……41
INERTIAL REFRENCE SYSTEM……………………………………………………………….……45
FLIGHT DARA RECORDER……………………………………………………………………….....47
COCKPIT VOICE RECORDER……………………………………………………………………….50
ELECTRONIC FLIGHT INFORMATION SYSTEM………………………………………………..52

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THE PITOT – STATIC SYSTEM

Pitot Static System

 The air speed indicator, the altimeter, vertical speed indicator, Machmeter
work on the pitot-static system.
 Static pressure is the ambient pressure of air and is present everywhere
 It is measured by the static port
 Dynamic pressure is when moving air is brought to rest (the pressure
exerted due to the movement of air) D= ½ ρV²
 It is measured by the Pitot probe
 The Pitot probe measures the total pressure (Static+ Dynamic)
 If air is disturbed around the static port, it senses an error
 Static pressure can be disturbed due higher TAS and angle of attack (AOA)
 Alternate static source is placed in the cockpit in unpressurised aircraft to
avoid icing whereas in pressurised aircraft it is placed outside.
 Pressure inside the cockpit is slightly lower then outside

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Air Temperature Measurement

Static Air temperature (SAT)
 It is the temperature of the undisturbed air through which the aircraft is
about to fly.

Instruments used to measure

 Direct reading – Works on the differential coefficient of expansion with
temperature. A bimetallic strip made up of INVAR and BRASS is used.
(shielded to avoid solar radiation)
 Remote reading – Temperature information is given in the form of electrical
signals, works on the principle of change of electrical resistance with
temperature
• An air temperature probe may be aspirated to measure air temp on
ground, using engine bleed air.
Total Air Temperature (TAT)
 It is the maximum temperature attainable by air when it is brought to rest
adiabatically
 As the aircraft speed increases the air gets compressed and the temperature
of air increases
 The increase in the air temperature at higher speeds (300 kts and above) is
because of compression and friction, and is known as RAMRISE
 The percentage of ram rise sensed and recovered via TAT probe is termed
as ‘Recovery Factory’ (K)
 RAT = SAT + Recovery factor
 SAT=TAT-(V/100) ² where V is TAS
 TAT =SAT + RAMRISE
 When mach number is given SAT=TAT/1+0.2krM²

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Air Speed Indicator

Principal of ASI
 The Pitot and Static pressures are fed into the ASI, via pitot probe and
static ports, where pitot pressure is fed into the capsule and static pressure
is fed into the airtight casing ,a differential pressure gauge measures the
difference and displays it as the aircraft’s speed on the indicator.
• ASI measures the dynamic pressure (Difference between pitot and
static)
• Dynamic Pressure = Pitot Pressure- Static Pressure.
 Speed of the aircraft is measured in relation to the air.
 In vacuum the air speed indication would read zero.
 The ASI will read the true air speed at ISA MSL pressure of 1013.25 mb,
1225 gm/m³ and temperature of 15º C
 The higher we move away from standard conditions the higher is the error
Different Air Speeds and Errors
 IAS CAS/RAS EAS TAS GS
o POS COMP DENS WIND
 TAS=CAS+(1.75% of CAS per 1000 ft)

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Errors of ASI (BLIPDC)

 Instrument error- Construction defects
 Position or Pressure Error - Due to the changes in static pressure.
Example- Changes in angle of attack, speed, flap setting.
(Maximum error would be when pitch changes are high- Manoeuvre
Induced error)
 Compressibility Error- At TAS more than 300kts (0.4M), it causes the
pressure to increase due compression and causes ASI to over read
 Higher the airspeed and altitude, higher will be the error
 Density Error - Whenever the density of air is not 1225g/m³ there will be
difference between the ASI reading and TAS
 So, at higher altitude ASI will under read the TAS and whenever density is
higher than 1225g/m³ ASI will over read the TAS
 If flying to a warmer airmass, TAS will increase
 Tolerance is + 3% or 5 kts whichever is greater
Blockages (PUDSOD)
 Pitot - If the pitot is blocked (Act as an altimeter):
o No changes during straight and level flight
o Over read during climb
o Under read during descend
• Static (Opp. Of Pitot)
o No change during straighten level
o Over read in descend (In an effort to reduce the
speed, you might stall the A/C)
o Under read in climb
Leaks
 A leak in the pitot will cause the ASI to under read
 Static:
o In unpressurised A/C, ASI will over read
o In pressurised A/C ASI under reads

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V Speeds

 Vso – Is stall speed with gears and flaps extended (Lower end white arc)
 Vs1 – Is stall speed in clean configuration (lower end of green arc)
 Vfe – Is maximum speed with flaps extended (Higher end of white arc)
 Vno – Normal operating speed (Higher end of green arc)
 Vne – Never exceeds speed
 Vyse – Single engine best rate of climb at maximum weight (blue line)
 Vlo – Maximum landing gear lowering speed
 Vle – Maximum speed with landing gear extended
 Va – Manoeuvring speed (Maximum speed at which abrupt controls can be
used)
 Vmo – Maximum operating speed (CAS)
 Vfo – Flap operating speed
 Vx – Best angle of climb
 Vy – Best rate of climb
 Yellow Arc: Caution range Vno to Vne (can only be flown in calm air)
 Red and White in mach meter – indicates CAS of VMO at low altitude and
MMO at high altitude.
 Colour coding is imp for exams
 Static Balancing
 When static ports are placed on both sides of the fuselage to help reduce the
position error when the aircraft is side slipping is called static balancing.
 Aircraft side slipping towards the block static port, ASI will over read.
 Aircraft side slipping towards the open port (Opp. Port is blocked) ASI will
under read

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Pressure Altimeter
 It indicates height above a selected pressure datum
 Works only on static pressure
 Static pressure decreases with height which is sensed by the altimeter to
indicate an increase in height above a selected datum

Construction
 It consists of a metal capsule which is evacuated and sealed(vacuum)
 A leaf spring is attached to prevent it from collapsing
 The capsule is fed with the datum pressure (QNH,QFE,QNE)
 This capsule is mounted in an air tight case and the case is connected to the
static port
 If an aircraft climbs the pressure in the casing will fall, allowing the capsule
to expand and transfer this information to a pointer and vice versa in a
descend
 Pitot tube is not connected to the altimeter so a pitot blockage would not
affect the altimeter reading
 Tolerance - + 100ft at MSL

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 QFE – Pressure at the Aerodrome reference point.

 QNH – It is the station pressure reduced to MSL at ISA conditions.
 QFF – It is the station pressure reduced to MSL under existing
conditions(used in charts only)
 QNE – 1013.25 standard pressure (Pressure Altitude).
 HEIGHT – Vertical distance from ground.
 ALTITUDE – Vertical distance from MSL.
 ELEVATION – Vertical distance of a fixed point above MSL.
 PRESSURE ALTITUDE – When 1013.2 or 29.92 set on subscale, also
known as flight level.

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 Flying from low temperature or pressure to high temp or pressure, the

altimeter would under read (indicated alt. would be lower than true alt.)

L H L (Indicated Alt.)
H L H (Indicated Alt.)

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 TRANSITION ALTITUDE – At or below this altitude vertical distance is

given in feet ,It is the last available altitude before changing to FL.
 TRANSITION LEVEL – At or above this the vertical distance is given in
flight level. This is the lowest flight level available, local QNH must be set
below this.
 TRANSITION LAYER – The air space between the transition altitude and
transition level,
DENSITY ALTITUDE

• It is the pressure altitude corrected for non-standard temperature

• The altitude at which the prevailing density occurs in ISA
 Density decreases with increase in temperature, decrease in pressure and
increase in moisture
 Lower the density, higher is the density altitude
 DA= PA+120 (ISA Deviation)
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ERRORS
 Instrument Error – Due to manufacturing defects
 Position Error – Due to the disturbance of airflow around the static port
 Maneuver Induced Error – Due to the abrupt maneuver there is a
disturbance in the static pressure
 Time Lag – There is a delay in the response of the altimeter pointer after
levelling off from a climb or descend due to which the altimeter will lag
(This error is virtually eliminated in a servo assisted altimeter)
 Barometric Error - Due to the wrong setting in the sub-scale
 Pressure Error – Whenever the outside pressure is other than ISA
• Temperature Error – Whenever the outside air temperature is other than
ISA
 Hysteresis Error – Due to time spent at a particular altitude in cold or hot
temperature the capsule contracts or expands which causes a lag in the
instrument during a climb or descend.
A vibrating device is used at the linkage to reduce friction between gears.
Blockage
 No change during straight and level
 During climb it will under-read
 During descend it will over-read
Leakage
 In a pressurized A/C the altimeter will under read
 In an unpressurised A/C the altimeter will over read
Static Balancing
 Only with two static ports
 Altimeter over-reads when slipping towards blocked static port
 Altimeter under-reads when slipping towards open static port
Formulae
 Temperature Deviation = Actual – ISA
 Density Alt = Pressure Altitude + (118.6 x Temperature Deviation)
 Pressure Altitude = Elevation + {30x(1013-QNH)}
 True Alt = Indicated + (ISA Deviation x 4/1000 x Indicated Altitude)

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Sensitive Altimeter

 It is similar to the normal altimeter ,contains two aneroid capsules to give

over power to drive the linkage and also a temperature shield is installed to
remove hysteresis error
 Accuracy - + 70ft at MSL
Servo Assisted Altimeter

 Instead of mechanical transfer of information, an electromagnetic induction

device is installed which provides greater accuracy
 Lag error is virtually eliminated
 Accuracy = +/- 30ft at sea level

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Vertical Speed Indicator

 It indicates the rate of climb or descend by sensing the rate of change of

static pressure
 Static pressure is fed directly into the capsule and it is also fed into the
instrument casing through a metering unit (Choke)
 This metering unit delays the pressure inside the casing due to which a
pressure differential is created between the capsule and the casing, hence
the capsule expands or contracts, which in turn transfers the information to
the pointer
 Dial representation is logarithmic rather than linear (pointer movement is
larger initially and easy to read)
 A capillary and orifice are installed to correct for any temperature changes
due altitude

ERRORS
 Instrument Error - Due to manufacturing Imperfections
 Pressure Error – Manoeuvre Induced error
 Blockage
 Will read zero if static is blocked
 If choke is blocked VSI shows over-reading during climb and descend
 Lag Error – Since there is a delay in the pressure between capsule and
casing so whenever the A/C levels off after a climb or descend the VSI
takes time to read zero. This error is highest after a prolonged climb or
descend especially at a high rate (Instrument always under read)
 VSI is not affected by temperature error as it only measures the change in
air pressure
 During T/O run, VSI shows a false climb due sudden increase in speed
 Advantage of breaking the glass window of the VSI.
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IVSI (Instantaneous Vertical Speed Indicator)

 Dashpot accelerometer is used to eliminate the lag error of VSI

 Dashpot accelerometer is a piston in a cylinder with two sensitive springs
which reduces the time delay between the capsule and casing
 It causes a temporary false indication of climb on entering a turn in level
flight
 Because of the sensitivity of the dashpot, the instrument tends to overreact
to turbulent flying condition
Serviceability Checks
 + 200 ft per min on ground between – 20 to + 50ºC
 + 300 ft per min on ground outside this range

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MACHMETER
Local Speed of Sound (LSS)
 LSS= 38.95√Temp in Kelvin (0ºC = 273k)
 When an aircraft approaches the local speed of sound, the air flow over
some part of the fuselage or the wing may reach the speed of sound and a
shock wave will form, these shock waves cause more drag, less lift, Mach
tuck, buffeting and reduction in control effectiveness or loss of control.
 The CAS speed of the A/C at which the airflow over some part of the
aircraftreaches the speed of sound is called the critical Mach number
(Mcrit)
 Machmeter is a combination of ASI and Altimeter
 It measures the A/C’s speed in relation to the local speed of sound
 The speed of sound is directly proportional to temperature
 Higher the altitude, lower the temperature and lower will be the speed of
sound
 The TAS should not exceed Mcrit
Construction
 Two capsules are incorporated to indicate the ratio of the TAS and the local
speed of sound
 One capsule is the airspeed capsule which shows the changes in the
dynamic pressure
 The other capsule is the sealed Altimeter capsule which expands or
contracts as the static pressure inside the instrument case changes

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MACH No. = TAS = P-S = Dynamic Pressure

LSS S Static Pressure
 Vmo – Max operating indicated speed
 Mmo – Max operating mach number
 At near MSL, the shock waves form at about 661 kts, but the airplane will
collapse before that because of the dense air over-stressing the a/c
 Thus, at lower altitude (t/o, initial climb & final descent) max speed of
the a/c is limited by Vmo
 At higher altitude shockwave forms early therefore, max speed is
limited by Mmo
 While descending, the max speed of the aircraft is limited by M mo initially
and then by Vmo (roughly around 26,000 ft)
 Vmo is expressed as CAS whereas MMO is expressed in MACH No.
 Red and white needle shows the maximum VMO and MMO (Barber’s
poll) and it is affected by temp and TAS

ERRORS
 Just like the other instruments these suffers from manufacturing errors,
position errors and manoeuvre induced error
 Density, temp and compressibility errors are eliminated by installing
the altitude capsule, because density changes do not change the ratio of
dynamic and static pressure

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BLOCKAGES (PUDSOD)
 These errors are similar to Airspeed indicator
 If a static source is blocked
o Machmeter will under-read in a climb and over-read in a descend
 If a pitot source is blocked
o Machmeter will over-read in climb and under read in descend
 Tolerance is + .01 M
 This is the only instrument which is not calibrated to ISA
 Relation between CAS, TAS, MACH No. under ISA conditions

 Also go through the relation in Isothermal and Inversion

 With CAS constant in straight and level flight, if the temperature reduces
what happen to TAS?
o TAS will decrease with no change in Mach number

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AIR DATA COMPUTER

Advantages
 Reduced instrument and lag error (position/pressure)
 Large no of instruments can be fed simultaneously
 Failure warning
Aeroplane systems receiving info
 Flight Data Recorder (FDR)
 Flight Management System (FMS)
 Automatic flight control system (Autopilot)
 Transponder
 GPWS
 FD
Note
 STBY instruments are not connected to the ADC

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MAGNETISM

 North seeking side of the magnet is RED which is attracted to the blue pole
of earth’s magnetism (magnetic north pole)
 Earth has its own magnetic field
 Dip – Angle that a freely suspended magnet makes with the horizontal
 It is zero at the equator and 90º at the poles
 Higher the Lat, higher is the Magnetic dip
 Lines joining places of equal dip are called isoclinic
 Lines joining places of zero dip are called aclinic (Magnetic equator)

 T is the total magnetic component at a particular place which is subdivided

into the Horizontal component (H) and the Vertical component (Z)
 Dip = Tan-1Z/H
 The horizontal component or the directive force (H) is maximum at the
magnetic equator (Dip Zero degree), while the vertical component (Z) is
maximum at the magnetic poles (Dip 90 degree)

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 Magnetic compass – It is an instrument which incorporates freely

suspended magnet in it to indicate the direction with relation to the
Magnetic north, the horizontal component is used as the direction datum.
 Mandatory to be carried on big aircrafts as a standby instrument.
Properties of an ideal compass

 Horizontality – A magnet should remain horizontal to give accurate

readings but it is only possible at the Magnetic equator
 Moving towards the poles, the magnet tends to dip to the nearer pole due to
the vertical component hence magnet will stand vertical at poles.
 Can be achieved by concentrating the mass of the magnet below the pivot
point (moving the C of G below the pivot point)
 The tilt effect due to dip is opposed by the weight of the magnet
 Sensitivity – The magnet assembly should be sensitive enough to seek the
horizontal component
 It can be increased by using large magnets, which is not practical, instead
several small, circular magnets are used to increase sensitivity
 Aperiodicity – After a displacement due manoeuvres or turbulence, the
compass should come to rest at the earliest
 This is achieved by damping liquids, multiple magnets and damping wires
 The assembly is immersed in a liquid, to reduce the effective weight.
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Properties of the liquid used in the compass (eg- Ethyl Alcohol)

 Transparency
 Low freezing point and high boiling point
 Low viscosity and specific gravity
 non- corrosive
 Low coefficient of expansion
 Liquid should be free from bubbles, discolouration and sediments

Acceleration and Deceleration errors

 These errors are zero on a heading of N & S. And increases towards east
and west with Max being on E & W.
 Acceleration causes the compass to show a turn towards the nearer pole and
opposite while decelerating
 The size of the error depends upon the heading of the aircraft, latitude (dip)
and the magnitude of acceleration.
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 ANDS (Acc North, Dec South on E &W headings)

 Acc on West – Compass overreads – turns anti - clockwise
 Decc on West – Compass under reads – turns clockwise
 Acc on East – Compass under reads – turns clockwise
 Decc on East – Compass overreads – turns anti - clockwise
 Everything is opposite in the southern hemisphere
Turning errors

For Northern Hemisphere

 It is maximum on N&S headings and Zero on E&W headings

 SONU – South overshoot and North Undershoot

Liquid Swirl
 Increases error while turning through North
 Decreases error while turning through south
 Is present at the equator also and under reads the turn

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 Attitude error - While doing a steep climb or descent

 Compass swing- It is to find out deviation on different headings and make
a deviation card
 Hard iron – is difficult to magnetize but once done it tends to retain its
magnetism, it is not usually influenced by earth’s magnetic field, e.g.- iron,
cobalt steel, chromium steel
 Soft iron – is easy to magnetize and easy to demagnetize
 Coefficient B = Deviation on E – Deviation on W
2
 Coefficient C = Deviation on N – Deviation on S
2
 Coefficient A = Deviation on (N+E+S+W+NE+SE+SW+NW)
8
 Deviation due to A is generally due to a misaligned lubber line
 Total deviation on a heading = A+ Bsinϴ + Ccosϴ
Compass swing should be done
 When compass components are installed or replaced
 Whenever the accuracy is in doubt
 If an inspection is due as per the maintenance
 If modification is done involving magnetic material
 After a lightning strike
 After having stood on one heading for a long time
 Large change of magnetic Latitudes
 Should be carried out on the apron

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GYROSCOPE

 Any rotating body with its mass on its periphery (edge)

 It has rigidity and precession
Rigidity (Gyroscopic inertia)
 The axis of the Gyro will maintain a fixed direction in space, unless an
external force has been applied to it.
 Rigidity increases with
o Speed of Gyro
o Diameter of Gyro
o By concentrating the mass of the rotor towards the rim
o Increasing rotor mass
Precession
 If a force is applied to a moving rotor the resultant force would act at a
point 90° ahead in the direction of rotation
 Higher the rigidity lowers the precession
 Higher the force applied higher the precession

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The formula for the rate of precession (Ω) is: -

Ω = T/І ω
Where: T is the applied torque
І is the moment of inertia of the rotor
ω is the angular velocity of the rotor
Gimbal
 A Gyro is mounted on Gimbal but allows it to rotate freely
 Gimbals are always perpendicular to each other
Wander

 Deviation of gyro from its axis to which it is set it can be horizontal or

vertical
o When the spin axis shifts in the horizontal plane it is known as
‘Drift’.
o When the spin axis shifts in the vertical plane it is known as
‘Topple’.
Real Wander
 When the spin axis shifts relative to the space
 It is caused by mechanical imperfections like imbalance rotor mass,
imperfectly balanced gimbals and uneven rotor bearing friction.
Apparent Wander
 Earth Rate - Is due to the effect of earth’s rotation
 As the gyro is aligned with the north and indicates heading of the aircraft in
relation to north pole, but since the earth is rotating at 15° per hour, the
Gyro shifts from its aligned axis and is longer pointing towards the north
pole and the axis has to be set manually
 Maximum drift is 15° per hour at the poles (15 x sin lat)
 Transport Wander - When the Gyro is transported in the aircraft, apparent
wander increases or decreases depending upon the direction of flight
 Transport wander = E/W GS X Tan Lat (NH = E = - /W= +)
60
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Types of Gyros:
 Number of Gimbals = Degrees of freedom
 Plane of freedom = Number of Gimbals + 1
 Free Gyro or Space Gyro: This Gyro has freedom of movement in all
three planes, they are used in INS
 Tied Gyro: A gyro which relates its rigidity to a particular attitude or
direction rather than to space has more application and is called a tied gyro.
 A tied gyro would have rigidity relative to the reference to which it is tied
 One axis is tied to restrict the movement.
 Used in directional Gyro indicator
 Earth’s Gyro:
-It is a tied Gyro in which one plane is tied by gravity
-It is used in an artificial horizon
 Rate Gyro: Used in turn and slip indicator and it has one plane of freedom
Power Source for gyro
 Suction Driven – An engine driven or venturi driven pump is used to
create vacuum and air is sucked in through jets moving the rotor, they are
not affected by electric failures but at high altitudes, dust and moisture can
give varying rotor rpm
 Electrically driven – High rotor speed can be achieved rapidly but are
more expensive

Tied Earth Rate

DGI Artificial Horizon Turn and slip indicator
Horizontal Vertical Horizontal
2-degree freedom 2-degree freedom 1 degree freedom
Rigidity Rigidity Precession

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DIRECTIONAL GYRO INDICATOR (DGI)

 The property of Rigidity is used to maintain accurate heading.

 The property of Precession is used to control the gyro and it is accomplished
by installing a latitude nut.
 It uses a Horizontal Axis Gyro.
 2 Gimbals = 2° of Freedom of precession in three planes but one is tied
 The jets help rotate the rotor at 12,000 RPM and mounted on the inner
Gimbal with its spin axis maintained in the yawing plane and free to rotate
in the vertical plane.
 A compass scale is attached on the outer Gimbal
 The rotor is driven by air drawn in through the aircraft vacuum system and
directed into the nozzle into the buckets, the air jets and a wedge plate are
attached to the outer gimbal
 A caging device is installed to synchronize the heading
Caging Device
 Brings the Gimbals 90° to each other
 Used for compass synchronisation (Every 15mins)
 Prevents topple during violent manoeuvres
 Re-erect after topple
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ERRORS
 It tends to be accurate in level turns because the DGI does not suffer from
turning errors.
 Gimbal Error: When the gimbals are not perpendicular to each other, the
instrument will give incorrect indication (gimbal error). When the aircraft is
pitched and rolled at the same time, the instrument would give a false
indication of direction and it will disappear after manoeuvre is completed
 Real Wander: This is due to manufacturing imperfections
 Apparent wader: Due to the rotation of the earth and it is maximum at the
poles (It is negative in the northern hemisphere and positive in the southern
hemisphere)

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 Varying Rotor RPM: Because of the varying density of air at higher

altitude the rigidity decreases and precession increases resulting in gyro
over-corrections
 Transport Wander: At any latitude other than the equator, meridians
(which define local north) are not parallel. If the gyro is aligned to one
meridian, then flown east to west, the new meridian will be inclined to the
old by transport wander.
 Transport Wander = Gs x Tanθ
60
Latitude Nut
 It compensates for apparent drift by attaching a bias weight, which makes
the gyro precess in the opposite direction of apparent wander
 It can only be done on ground
 It is intentional precession to correct for apparent drift and
provides compensation only at the latitude for which it is set
Limitations of a DGI
 If the aircraft exceeds the pitch or roll limits in modern DGI by 85° (55°
in air driven gyroDIs) the gyro will topple.

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ARTIFICAL HORIZON

 The artificial horizon uses an Earth Gyro in which the spin axis is
maintained in or tied to the vertical by earth’s gravity, this means that
the plane of rotor rotation is horizontal, thus providing the stable, lateral
and longitudinal references required.
 Provides attitude reference in pitch and roll
 It replaces the natural horizon in poor visibility
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 It uses an earth Gyro, with a vertical spin axis

 Properties are rigidity and precession 
 Rigidity for attitude and precession for controlling the gyro axis
 The axis is maintained vertically because of a gravity sensing unit so it can
point towards the centre of earth
 2° of freedom in three planes, one is tied to the vertical by gravity
 2 Gimbals
 15,000 RPM in an air driven, and 22,500 RPM in an electrical driven gyro.
 With older designs, typical limits are ± 60° in pitch and 110° each way in
roll. In modern instruments there is complete freedom in roll and up to
85° (plus or minus) in pitch.
Artificial Horizon Gyro Erection System:
 The control system of air driven artificial horizon consists of four slots and
four pendulous (hanging) vanes at the base of the rotor housing. The vanes
hang down in such a way that when the rotor axis is vertical, each slot is
half covered by its vane and four equal jets of air emerged from the slots,
fore and aft and left and right.
 Now because the four jets are of equal strength but opposite directions, no
force is exerted on Gyro therefore no precession occurs - the Gyro rotor
remaining vertical.

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Limitations and Errors:

 Acceleration Error: It gives a false indication of right bank and climb
(especially during take-off) due to lag of the rotor assembly
 The opposite will happen in case of deceleration
 No acceleration or turning errors in an electrically driven attitude indicator
 Pneumatic attitude indicators take about 5 mins to erect after start up
 A caging device can reduce the erection process
 The rotor assembly is made slightly bottom heavy to reduce the time taken
for initial erection.
 The amount that the case can move relative to the Gyro is controlled by
fixed stops, in modern instruments there is complete freedom in roll and
85 pitch
 Turning Errors in Air driven Artificial Horizon-The magnitude of errors
varies with speed, rate of turn & type of horizon.

TURNING BANK ANGLE PITCH

ERRORS ATTITUDE
Turning through 090 Under reads bank angle Indicates a climb
Turning through 180 Bank angle correct Indicates a climb
Turning through 270 Over reads bank angle Indicates a climb
Turning through 360 Bank angle correct Pitch angle is correct

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Electrical Driven Artificial Horizon

 There are two levelling switches, one to sense pitch and one to sense roll.
 Electric AH uses levelling/Mercury switches in torque motors to keep its
platform stable
 They activate the pitch and roll torque motors respectively which precess
the gyro back to the vertical as soon as it starts to wander
 The torque motor on the side of the inner gimbal corrects wander in the
rolling plane
 The pitch torque motor is on the outer (longitudinal) gimbal so that the
precession is about the lateral axis to correct for pitch
 During rapid acceleration, a pitch cut-out switch activates when an
acceleration of 0.18g or greater is detected. Similarly, another cut-out is
activated at 10 degrees angle of bank.
 Greater rigidity due faster rotor speed
 Acceleration errors are reduced due less pendulous and linear
acceleration cut out switches
 There is an off-flag when power supply is disconnected
 Rotor housing is less bottom heavy and therefore roll error is reduced
 Pitch and roll cut out switches are fitted instead of pendulous vanes,
therefore pitch error is reduced

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TURN AND SLIP INDICATOR (TSI)

 It is a rate Gyro and the axis of the gyro is mounted athwartships

 It incorporates two instruments into one, one being the inclinometer to
measure the slip or skid and the other to measure the aircraft rate of turn (It
generally measures the change in direction)
 It has 1° of freedom with one gimbal in two planes
 Horizontal axis Gyro and the property of precession is used for rate of turn, 
 No indication if the aircraft just banks without moving around the vertical
axis
 The inclinometer uses the combination of gravitation and centrifugal force
• RPM of about 4500, which is less than AH and DG. If rotational speed
is less than rated, less rate of turn will be indicated
• If the Gyro over speeds, then it will over read the rate of turn.
 In straight and level flight there will be no centrifugal force and the gravity
will keep the ball in centre, where as in coordinated turn the centrifugal and
gravitational force will keep the ball in centre
 If the air craft is over banked, the centrifugal force will be less than
gravitational force (A/C slipping into a turn)
 If the A/C is under banked, the centrifugal force would be greater than
gravitational force (A/C skidding out of turn)

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TURN COORDINATOR

 It shows the coordination of bank angle and rate of turn

 The longitudinal axis of Gyro gimbal is inclined at 30° to the horizontal, so
now the Gyro will respond to banking as well as turning
Rate of Turn
 Rate of change of direction per second
 Rate one turn = 3° per sec (360°=2min)
 Rate two turn = 6° per sec and so on
 Angle of bank required for rate one turn = 10 % of TAS + 7
Radius of Turn
 V2/gTanθ
 At constant angle of bank, if speed is increased, radius of turn will increase
and rate of turn will decrease
 At constant speed, if angle of bank is increased, radius of turn will decrease
and rate of turn will increase

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REMOTE INDICATING MAGNETIC COMPASS

Basic System Description

It is a system which combines the properties of a DGI with the Magnetic
compass. Such a system is a gyro-magnetic compass.
The gyro-magnetic compass is known by several names. It can be referred to as
the:
• Gyro-magnetic Compass.
• Remote Indicating Compass.
• Slaved Gyro Compass.
Operation
 The detector unit (flux valve) senses the earth’s magnetic field and sends it to
the error detector, where it is compared with the position of the gyro drive
shaft.
 If the two are aligned, no further action takes place. If, the gyro starts to drift,
the drive shaft will not be in alignment with the flux valve field, and an AC
error signal is generated by the error detector and passed to the precession
amplifier, where it is amplified, phase detected, and rectified to DC.
 The DC signal drives the precession motor, which turns the gyro. This gyro
output is fed via the direct drive shaft to the heading indicator’s error detector
for comparison with the flux valve signal.
 If the two are aligned, the compass is synchronized and no further action takes
place. If not, the error correction continues until the compass is synchronized.
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The Flux valve

 The detector unit is positioned in a part of the aircraft least affected by on-
board electrical fields (usually the wing tip or tail fin, where any aircraft
generated magnetic disturbances are at a minimum).
 It contains a pendulous magnetic detecting element mounted on a Hooke’s
Joint which enables the detector to swing within limits of 25° about the pitch
and roll axes, but allows no rotation in azimuth.

 The operation of the flux valve is in accordance with Faraday’s Law of

Electromagnetic Induction.
 When it operates as a DGI, this is referred to as FREE mode, whilst its normal
magnetically monitored operation is referred to as SLAVED mode.

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 When the Heading Indicator is FREE, the pilot adjusts the indicated
heading in order to correct it to an external datum heading by use of the
CCW/CW (counter-clockwise/clockwise) control switch, which is spring-
loaded to the central position.
The annunciator is useful to the pilot for 2 main reasons:
 It is an indication that magnetic monitoring of the gyro is taking place. It
shows that the compass is ‘synchronized’.
 On systems where it is necessary for the pilot to synchronize manually, it 
indicates which way to turn the compass.

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INERTIAL NAVIGATION SYSTEM

An INS is used to provide the a/c position, and velocity by continuously

measuring and integrating its acceleration, using accelerometers.
 This system is a self-contained navigational aid.
 It contains 2 accelerometers, one on N-S axis and another on E-W axis.
 The accelerometer is a pendulous device, When the aircraft accelerates, the
pendulum, due to inertia, swings off the null position. A signal pick-off device
tells how far the pendulum is off the null position.
 The accelerometer output is fed to 2 integrators, which are connected in series.

 The acceleration signal is sent to the first stage integrator, which integrates it
with respect to time, to give the velocity.
 This velocity is then further integrated by the second stage integrator to give
out the distance in either N-S or E-W direction.
 The new position is calculated by comparing the ch. Lat. & ch. Long. with the
previous known position.
 The final products of an INS are position (lat & long), speed (kt) and distance
(nm).
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MODES
1) STANDBY
 The system is warmed up and the current position is inserted.
2) ALIGN
 The platform is levelled and aligned (gyro com-passed),Ready Nav
annunciator illuminates after the process is complete. The equipment
can now be switched into the Nav mode.
 Alignment is always done on the ground, when the a/c is not moving
(Takes up to 15-20 min).
3) ATT REF
 Selecting ATT REF disconnects computing and loses alignment.
Only used when INS fails to provide Nav info. The system now gives
attitude information (pitch and roll) and a limited form of heading.
 If power is turned off in flight to the INS and turned back on ,It could
only be used for Attitude Reference.

Note – Alignment can take place in Align and Nav mode

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 Gyroscopes which are used to stabilize the platform are also mounted on the
inner-most element of the platform. The gyro and the accelerometer are
mounted on a common gimbal. Three integrating gyros are mounted on the
inertial platform, with their input axis mutually perpendicular.
 Should the aircraft electrical supply to the INS fails for any reason, the INS
will automatically switch to its own battery pack.
 As the power from the battery starts to fail, the BATT warning light on
the Mode Selector Unit will illuminate, indicating that the INS is about to
fail.
 The ALERT annunciator warns that the aircraft is approaching the next way-
point. In AUTO mode the alert light will come on, steady, 2 minutes before
the waypoint, and will extinguish as the track changes overhead the waypoint.
 In MANUAL mode the alert light will come on, steady, 2 minutes before the
waypoint; the light will then flash 30 seconds before the waypoint, and will
continue to flash until the track is changed.

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Schuler Period
Should the platform be displaced from the horizontal it would oscillate with a
period of 84.4 minutes, which is known as the Schuler Period.

Bounded Errors
Errors which build up to a maximum and return to zero within 84.4 minutes
Schuler cycle are termed bounded errors. The main causes of these errors are:
 Platform tilt due to initial misalignment
 Inaccurate measurement of acceleration by accelerometers
 Integrator errors in the first stage of integration

Unbounded Errors
Unbounded errors - are either cumulative track errors or distance errors:
 Initial azimuth misalignment of the platform
 Wander of the azimuth gyro

Errors which give rise to cumulative errors:

 Wander in the levelling gyros. This causes a Schuler oscillation of the plat-
form.
 Integrator errors in the second stage of integration.

Inherent Errors
The irregular shape and composition of the earth, the movement of the earth
through space and other factors provide further possible sources of error.

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INERTIAL REFERENCE SYSTEM

 IRS is a solid-state unit of three Ring Laser Gyros detecting accelerations in

3 dimensions giving a shorter spin up time. (10 min)
 Ring Laser Gyros and accelerometers are attached rigidly, or “strapped
down”,to the frame of the aircraft i.e., as the aircraft moves, so does the IRS
platform.
 Three gyroscopes sense the rate of roll, pitch, and yaw; and three
accelerometers detect accelerations along each aircraft axis.
 It integrates them to get the orientation, then mathematically calculates the
acceleration the north/south, east/west and up/down axes.
 With few moving parts strap-down systems are easier to maintain and more
reliable over time.
 Inertial Reference Unit is a computer that integrates IRS outputs and pro-
vides inertial reference outputs for use by other navigation and flight control
systems, including the Flight Management System (FMS).
 The Inertial Reference Unit (IRU) is the heart of the IRS. It provides all re-
quired inertial reference outputs for the aircraft’s avionics.
Outputs are:
Primary attitude Pitch and roll
Heading True, Magnetic
Accelerations Lateral, Longitude, Normal
Angular rates Pitch, Roll, Yaw

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Inertial velocity N/S, E/W, GS, TA, Vertical rate

Position Latitude, longitude, inertial altitude
Wind data Wind speed, wind angle, drift angle
Calculated data Flight path angle and acceleration
 The primary sources of information for the IRU are its own internal sensors,
three laser gyros, and three inertial accelerometers. The only other inputs
required are initial position, barometric altitude, and TAS.
 The TAS input allows the IRU to calculate wind speed and wind direction.
Ring Laser Gyro (RLG)

 A ring laser gyro (RLG) splits a beam of laser light into two beams in opposite
directions through narrow tunnels in a closed circular optical path around the
perimeter of a triangular block of temperature-stable Cervit glass with
reflecting mirrors placed in each corner.
 When the gyro is rotating at some angular rate, the distance travelled by each
beam will differ—the shorter path being opposite to the rotation.
 The phase shift between the two beams can be measured by an interferometer
and is proportional to the rate of rotation (Sagnac effect).
 The most significant potential problem is lock in, also known as laser lock, at
very low rotation rates the output frequency can drop to zero. This phenome-
non is overcome by the introduction of a vibration device known as a piezo
electric dither motor which breaks the lock in.

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FLIGHT DATA RECORDER

The main function of the flight data recorder (FDR) is to preserve the aircraft
data in order to determine the cause of any aircraft accident.
The FDR records the last 10 or 25 hours of aircraft data on a digital storage
device housed in a shock resistant box that is painted red and located at the rear
of the aircraft, normally under the fin. On the front of the unit is an underwater
locating device (ULD).
FDR Components
The FDR consists of the following components:
 a recording system
 a control unit on the overhead panel
 a control unit on the pedestal
 Data Interface and Acquisition Unit (DIAU)
The control unit on the overhead panel also controls the cockpit voice recorder
(CVR). A spring-loaded switch labelled GND CTL can be selected ON or AUTO
as follows:
 ON The CVR and the DFDR are energized and the ON light is lit
 AUTO The CVR and the DFDR are energized:
 on the ground with one engine running
 in flight (with engine running or stopped)
The control on the pedestal consists simply of a push-button labelled ‘EVENT’
which sets an event mark on the DFDR recording. This acts as a kind of book-
mark to enable the “event” to be found rapidly on the recording at a subsequent
analysis
 When on the ground the FDR is automatically stopped 5 minutes after the
final engine shutdown.

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The mandatory aircraft parameters recorded on the FDR are:

 time or relative time count
 attitude (pitch and roll)
 airspeed
 pressure altitude
 heading
 normal acceleration
 propulsive thrust/power on each engine
 cockpit thrust/power lever position
 flaps/slats configuration or cockpit selection
 ground spoilers and/or speed brake selection
Additional parameters include the following:
 positions of primary flight controls and trim
 radio altitude and navigation information displayed to the flight crew
 cockpit warnings
 landing gear position
Regulations and requirements
1. All aeroplanes less than 5700kg MTOW, but has more than 9 seats must be
capable of recording at least last 10hrs of flight data.
2. All aeroplanes more than 5700kg MTOW must be capable of recording at
least, last 25hrs of flight data.
3. All aeroplanes more than 27000kg MTOW must also record the additional
parameters.
Other requirements
 The FDR must start automatically to record the data prior to the aeroplane
being capable of moving under its own power and must stop automatically
after the aeroplane is incapable of moving under its own power.
 This is achieved by starting when the first engine is started, and automatically
switching off 5 minutes after the last engine is shut down.
 The FDR must have a device to assist in locating that recorder in water.
 Aeroplanes of 5700 kg or less may have the FDR combined with the cockpit
voice recorder.
 Aircraft above 5700 kg must have 2 recorders, either separate FDR/CVR, or 2
combined recorders.

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An aeroplane may be dispatched with an inoperative FDR provided that:

 It is not reasonably practicable to repair or replace the FDR before flight
 The aeroplane does not exceed 8 further consecutive flights
 Not more than 72 hours have elapsed since the unserviceability
 Any cockpit voice recorder required to be carried is operative (unless it
iscombined with the FDR)

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COCKPIT VOICE RECORDER

The tape recorder is located inside a crash-proof metal box that is painted red or
orange and normally placed at the rear of the aircraft, often adjacent to the flight
data recorder. The high impact case should be able to withstand shock, high
temperature and fire.
On the front of the unit is fitted an underwater locating device (ULD), that will
emit a continuous series of ultrasonic pulses to help locate a submerged CVR.
The unit is automatically activated by water and the battery will last several days.
Parameters recorded
 Voice communications transmitted from or received on the flight deck
 The aural environment of the flight deck
 Voice communication of flight crew members using the aeroplanes
inter-phone system
 Voice or audio signals introduced into a headset or speaker
 Voice communication of flight crew members using the PA system.
The Control Unit
AUTO / ON
 When the switch is in the AUTO position the CVR will start to record when
the first engine is started and will stop 5 minutes after the last engine is shut
down.
 Selection of the ON position starts the CVR recording immediately and
latches the switch in the ON position until first engine start, when it will
click back to AUTO.
CVR TEST
 Pressing the TEST button activates an extensive set of functional tests
which determine the integrity of the system using the BITE (built-in test
equipment) facility. A successful self-test results in a visual ‘good’
indication.
ERASE
 Erasure of the tapes is only possible with the aircraft on the ground, all
engines stopped and the parking brake set.
 Additionally, the erase button must be held depressed for at least 2 seconds
before the circuit activates.

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Regulations and requirements

1. All aeroplanes less than 5700kg MTOW, but has more than 9 seats must be
capable of recording at least last 10hrs of flight data.
2. All aeroplanes more than 5700kg MTOW must be capable of recording at
least, last 25hrs of flight data.
3. All aeroplanes more than 27000kg MTOW must also record the additional
parameters.

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ELECTRONIC FLIGHT INFORMATION SYSTEMS

The Electronic Flight Instrument System presents attitude and navigation
information to the pilot on two electronic display units generally referred to as
‘EFIS’.It is fully integrated with digital computer-based navigation systems, and
utilizes colour Cathode Ray Tube (CRT) or Liquid Crystal Display (LCD) types
of Attitude Director Indicator (ADI) and Horizontal Situation Indicator (HSI).

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System Units
A complete EFIS installation is made up of left (Captain), and right (First
Officer), systems. Each system in turn is comprised of:
 Electronic Attitude Director Indicator (EADI) or Primary Flight Display
(PFD)
 Electronic Horizontal Situation Indicator (EHSI) or Navigation Display
(ND)
 Control Panel
 Symbol Generator (SG)
 Remote Light Sensor Unit
A third (centre) symbol generator is also incorporated so that its drive signals
may be switched to either the left or right display units in the event of failure of
their corresponding symbol generators.

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Symbol Generators (SGs)

Symbol generators provide the analogue, discrete, and digital signal interfaces
between an aircraft’s systems, the display units and the control panel, and they
also perform symbol generation monitoring, power control and the main control
functions of the ‘EFIS’ overall.
Display Units
The PFD and ND are usually identical units to facilitate spares commonality and
are often interchangeable with the systems display units (EICAS or ECAM).
The Colour Display System
WHITE Display of present situation information.
GREEN Display of present situation information where contrast
with white symbols is required, or for data having lower priority
than white symbols. Engaged auto flight modes.
MAGENTA All ‘fly to’ information such as flight director commands,
deviation pointers, active flight path lines.
CYAN Sky shading on an EADI and for low-priority information
such as non-active flight plan map data.
YELLOW Ground shading on an EADI, caution information display
such as failure warning flags, limit and alert annunciators and
fault messages.
RED For display of heaviest precipitation levels as detected by
the weather radar (WXR).
AMBER cautionary information, faults, flags
BLACK blank areas, display “off”
The Remote Light Sensor
The Remote Light Sensor is a photodiode device which responds to ambient light
conditions on the flight deck, and automatically adjusts the brightness of the CRT
displays to an acceptable level.

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System Symbols

The ‘EHSI’ Display Presentation

Four principal display modes may be selected on the EFIS control panel;
Full VOR Mode
With a VOR frequency selected, the EHSI displays a full compass rose with the
VOR source in the lower left and the frequency in the lower right.
 Weather radar displays are not available

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Expanded VOR
With a VOR frequency selected, the EHSI displays about 90° of compass rose
with the VOR source in the lower left and the frequency in the lower right.
 Weather radar displays are available; when selected “on”, range
arcs arealso visible.

Full ILS Mode

With an ILS frequency selected, the EHSI displays a full compass rose with the
ILS source in the lower left and the frequency in the lower right.
 Weather radar displays are not available.

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Expanded ILS
With an ILS frequency selected, the EHSI displays about 90° of compass rose
with the ILS source in the lower left and the frequency in the lower right.
 Weather radar displays are available, when selected “on”, range
arcs arealso visible.

Map Mode
The MAP display, displays information against a moving map background with
all elements to a common scale.

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 The aircraft active route as derived from the FMC is shown as a magenta
coloured line joining the waypoints.
 The active waypoint is shown as a magenta coloured star.
 The other waypoints making up the active route are called inactive way-
points and are shown as a white star.
 Distance to next waypoint and time at next waypoint are shown at the top
of the display.
 Weather radar (WXR) return data and range arcs are displayed when the
WXR switch is on.
Plan Mode
In ‘PLAN’ mode a static map background is used with active route data orien-
tated to true north.
 This mode allows the pilot to review the planned route by using the
FMC / CDU LEGS page.
 Weather radar display data is inhibited.
 No wind speed or direction information is shown
 The planning section below this is north orientated. The heading from
TOBIX to LOGAN is approximately 020°, NOT 155°.

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Failure Annunciation
 The failure of data signals from such systems as the ILS and radio altimeter
is displayed on each EADI and EHSI in the form of yellow flags ‘painted’
at specific matrix locations on their CRT screens.
 In addition, fault messages may also be displayed, for example, if the
associated flight management computer and weather radar range disagree
with the control panel range data, the discrepancy message ‘WXR/MAP
RANGE DISAGREE’ appears on the EHSI.

POETIC PILOT ACADEMY

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