S.

NO
1
2
3
4
5

NAME OF THE EXPERIMENT

STUDY OF PISTON ENGINES
STUDY OF JET ENGINES
STUDY OF PERFORMANCE OF PROPELLER
STUDY OF WALL JET
FREE JET

PAGE NO
2
6
12
17
23

PRESSURE DISTRIBUTION OVER A SYMMETRICAL

6

26

AEROFOIL
CASCADE TESTING OF MULTIPLE AEROFOIL SECTION OF
7

29

TURBINE BLADES
8
9
10

NOZZLE PERFORMANCE TEST

NOZZLE PERFORMANCE DISTRIBUTION TEST
BOMB CALORIMETER

31
34
37

EXP NO: 1
STUDY OF PISTON ENGINES
AIM
To study the piston engine including study of assembly of subsystems, various components and their
functions and operating principles.
Introduction
A piston engine is a HEAT ENGINE that uses one or more PISTONS to convert PRESSURE into a
rotating motion. The main types are the internal combustion engine used extensively in motor vehicles, the
STEAM ENGINE which was the mainstay of the industrial revolution and the niche application stirling
engine.
There may be one or more pistons. Each piston is inside a CYLINDER, into which a gas is introduced, either
already hot and under pressure (steam engine), or heated inside the cylinder either by IGNITION of a fuel air
mixture (internal combustion engine) or by contact with a hot heat exchanger in the cylinder (stirling engine).
The hot gases expand, pushing the piston to the bottom of the cylinder. The piston is returned to the cylinder
top (top dead centre) either by a FLYWHEEL or the power from other pistons connected to the same shaft. In
most types the expanded or "EXHAUSTED" gases are removed from the cylinder by this STROKE. The
exception is the stirling engine, which repeatedly heats and cools the same sealed quantity of gas.
In some designs the piston may be powered in both directions in the cylinder in which case it is said to be
double acting.
Components and their functions
The major components seen are connecting road, crank shaft(swash plate), crank case, piston rings,
spark plug, cylinder, flywheel, crank pin and valves or ports.
In all types the linear movement of the piston is converted to a rotating movement via a
CONNECTING ROD and a CRANKSHAFT or by a SWASH PLATE. A FLYWHEEL is often used to ensure
smooth rotation. The more cylinders a reciprocating engine has, the more vibration-free (smoothly) it can run
also the higher the combined piston displacement volume it has the more power it is capable of producing.
A seal needs to be made between the sliding PISTON and the walls of the CYLINDER so that the high
pressure gas above the piston does not leak past it and reduce the efficiency of the engine. This seal is
provided by one or more PISTON RINGS. These are rings made of a hard metal which are sprung into a
circular grove in the piston head. The rings fit tightly in the groove and press against the cylinder wall to form
a seal.
Engine terminology
Stroke: Either the up or down movement of the piston from the top to the bottom or bottom to top of the
cylinder (So the piston going from the bottom of the cylinder to the top would be 1 stroke, from the top back
to the bottom would be another stroke)
Induction: As the piston travels down the cylinder head, it 'sucks' the fuel/air mixture into the cylinder. This
is known as 'Induction'.
Compression: As the piston travels up to the top of the cylinder head, it 'compresses' the fuel/air mixture
from the carburetor in the top of the cylinder head, making the fuel/air mix ready for ignighting by the spark
plug. This is known as 'Compression'.

Ignition: When the spark plug ignites the compressed fuel/air mixture, sometimes referred to as the power
stroke.
Exhaust: As the piston returns back to the top of the cylinder head after the fuel/air mix has been ignited, the
piston pushes the burnt 'exhaust' gases out of the cylinder & through the exhaust system.
The following is an additional parameter for a 2 stroke engine
Transfer Port: The port (or passageway) in a 2 stroke engine that transfers the fuel/air mixture from the
bottom of the engine to the top of the cylinder
Types of piston engines
It is common for such engines to be classified by the number and alignment of cylinders and the total
volume of DISPLACEMENT of gas by the pistons moving in the cylinders usually measured in CUBIC
CENTIMETERS (cc).
In-line engine
This type of engine has cylinders lined up in one row. It typically has an even number of cylinders, but
there are instances of three- and five- cylinder engines. An in-line engine may be either air cooled or liquid
cooled. It is better suited for streamlining. If the engine crankshaft is located above the cylinders, it is called
an inverted engine. Advantages of mounting the crankshaft this way include shorter landing gear and better
pilot visibility. An in-line engine has a higher weight-to-horsepower ratio than other aircraft engines. A
disadvantage of this type of engine is that the larger it is, the harder it is to cool. Due to this, airplanes that use
an inline engine use a low- to medium-horsepower engine, and are typically used by light aircraft.
Opposed engine
An opposed-type engine has two banks of cylinders opposite each other. The crankshaft is located in
the center and is being driven from both sides. The engine is either air cooled or liquid cooled, but air cooled
versions are used mostly in aviation. It can be mounted either vertically or horizontally. The advantage of a
horizontally-opposed engine is that it allows better visibility and eliminates fluid lock typically found on
bottom cylinders. An opposed engine also has a relative advantage in being mostly free of vibration. This is
due to the fact that the pistons are located left and right of the crankshaft and act as balance weights for each
other.
V-type engine
Cylinders in this engine are arranged in two in-line banks, tilted 30-60 degrees apart from each other.
The engine can be either air cooled or liquid cooled.
Radial engine
This type of engine has a row of cylinders arranged in a circle around a crankcase located in the
middle. The combination of cylinders must be an odd number in each row and may contain more than one

row. The odd number of cylinders allows for every other cylinder to be on a power stroke, allowing for
smooth operation. The power output is anywhere from 100 to 3,800 hp.
4 Stroke engine
Engines based on the four-stroke or Otto cycle have one power stroke for every four strokes (up-downup-down) and are used in cars, larger BOATS, and many light AIRCRAFT. They are generally quieter, more
efficient, and larger than their two-stroke counterparts. There are a number of variations of these cycles, most
notably the ATKINSON and MILLER cycles. Most truck and automotive diesel engines use a four-stroke
cycle, but with a compression heating ignition system. This variation is called the DIESEL CYCLE. The four
strokes refer to intake, compression, combustion and exhaust strokes that occur during two crankshaft
rotations per working cycle of Otto Cycle and DIESEL ENGINES. The four steps in this cycle are often
informally referred to as "suck, squeeze (or squash), bang, blow."
2 Stroke engine
The two-stroke internal combustion engine differs from the more common four-stroke engine by
completing the same four processes (intake, compression, combustion, exhaust) in only two strokes of the
piston rather than four. This is accomplished by using the beginning of the compression stroke and the end of
the combustion stroke to perform the intake and exhaust functions. This allows a power stroke for every
revolution of the crank, instead of every second revolution as in a four-stroke engine. For this reason, twostroke engines provide high specific power, so they are valued for use in portable, lightweight applications
such as chainsaws as well as large-scale industrial applications like locomotives. Two-stroke engines are still
commonly used in high-power, handheld applications where light weight is essential, primarily string
trimmers and chainsaws. To a lesser extent, these engines may still be used for certain small, portable, or
specialized machine applications. These include outboard motors, high-performance, small-capacity
motorcycles, mopeds, under bones, scooters, snowmobiles, karts, ultra lights, model airplanes (and other
model vehicles) and lawnmowers. In the past, two-stroke cycles were experimented with for use in diesel
engines, most notably with opposed piston designs, low-speed units such as large marine engines, and V8
engines for trucks and heavy machinery
A Very Basic 2 Stroke Engine Cycle
Stroke Piston Direction

Actions Occurring
Explanation
during This Stroke

Stroke Piston travels up the Induction

1
cylinder barrel
Compression

As the Piston travels up the barrel, fresh fuel/air mix is

& sucked into the crankcase (bottom of the engine) & the
fuel/air mix in the cylinder (top of the engine) is
compressed ready for ignition

The spark plug ignites the fuel/air mix in the cylinder, the
resulting explosion pushes the piston back down to the
bottom of the cylinder, as the piston travels down, the
Stroke Piston travels down
Ignition & Exhaust transfer port openings are exposed & the fresh fuel/air
2
the cylinder barrel
mix is sucked from the crankcase into the cylinder. As the
fresh fuel/air mix is drawn into the cylinder, it forces the
spent exhaust gases out through the exhaust port.

A Very Basic 4 Stroke Engine Cycle

Stroke

Piston
Direction

Actions
Inlet & Exhaust Occurring
Explanation
Valve Positions During
This
Stroke

Piston travels
Inlet
valve
As the Piston travels down the cylinder barrel,
down
the
Stroke 1
open/Exhaust
Induction stroke the inlet valve opens & fresh fuel/air mixture
cylinder
valve colsed
is sucked into the cylinder
barrel
Piston travels
up
the Inlet & exhaust Compression
Stroke 2
cylinder
valve closed
stroke
barrel

As the piston travels back up the cylinder, the

fresh fuel/air mix is compressed ready for
ignition

Piston travels
The spark plug ignites the compressed fuel/air
down
the Inlet & exhaust Ignition (power)
Stroke 3
mix, the resulting explosion pushes the piston
cylinder
valve closed
stroke
back to the bottom of the cylinder
barrel
Piston travels
Inlet
valve
up
the
Stroke 4
closed/Exhaust
Exhaust stroke
cylinder
valve open
barrel

As the piston travels back up the cylinder

barrel, the spent exhaust gases are forced out
of the exhaust valve

RESULT
Thus the study of piston engine including study of assembly of subsystems, various components and
their functions and operating principles is done successfully.

EXP NO: 2

STUDY OF JET ENGINES

AIM
To study about the jet engines and its components.
INTRODUCTION
A jet engine is a reaction engine that discharges a fast moving jet of fluid to generate thrust in
accordance with Newton's third law of motion. This broad definition of jet engines includes turbojets,
turbofans, rockets, ramjets, pulse jets and pump-jets, but in common usage, the term generally refers to a gas
turbine Brayton cycle engine, an engine with a rotary compressor powered by a turbine, with the leftover
power providing thrust. Jet engines are so familiar to the modern world that gas turbines are sometimes
mistakenly referred to as a particular application of a jet engine, rather than the other way around. Most jet
engines are internal combustion engines but non combusting forms exist also.
Jet engines are primarily used by jet aircraft for long distance travel. The early jet aircraft used turbojet
engines which were inefficient. Modern jet aircraft usually use high-bypass turbofan engines which help give
high speeds as well as, over long distances, better fuel efficiency than many other forms of transport. A large
proportion of the worlds oil consumption (about 7.2% in 2004) is burnt in jet engines.

Major Components Of A Jet Engine And Their Functions

The major components of a jet engine are similar across the major different types of engines, although
not all engine types have all components.
Cold Section:
Air intake (Inlet) The standard reference frame for a jet engine is the aircraft itself. For subsonic aircraft,
the air intake to a jet engine presents no special difficulties, and consists essentially of an opening which is
designed to minimize drag, as with any other aircraft component. However, the air reaching the compressor of
a normal jet engine must be traveling below the speed of sound, even for supersonic aircraft, to sustain the
flow mechanics of the compressor and turbine blades. At supersonic flight speeds, shockwaves form in the
intake system and reduce the recovered pressure at inlet to the compressor. So some supersonic intakes use
devices, such as a cone or ramp, to increase pressure recovery, by making more efficient use of the shock
wave system.
Compressor or Fan The compressor is made up of stages. Each stage consists of vanes which rotate, and
stators which remain stationary. As air is drawn deeper through the compressor, its heat and pressure
increases. Energy is derived from the turbine (see below), passed along the shaft.

Common:

Shaft The shaft connects the turbine to the compressor, and runs most of the length of the engine. There
may be as many as three concentric shafts, rotating at independent speeds, with as many sets of turbines and
compressors. Other services, like a bleed of cool air, may also run down the shaft.
Hot section:
Combustor or Can or Flame holders or Combustion Chamber This is a chamber where fuel is
continuously burned in the compressed air.
Turbine The turbine is a series of bladed discs that act like a windmill, gaining energy from the hot gases
leaving the combustor. Some of this energy is used to drive the compressor, and in some turbine engines (i.e.
turboprop, turbo shaft or turbofan engines), energy is extracted by additional turbine discs and used to drive
devices such as propellers, bypass fans or helicopter rotors. One type, a free turbine, is configured such that
the turbine disc driving the compressor rotates independently of the discs that power the external components.
Relatively cool air, bled from the compressor, may be used to cool the turbine blades and vanes, to prevent
them from melting.
Afterburner or reheat (chiefly UK) (mainly military) Produces extra thrust by burning extra fuel, usually
inefficiently, to significantly raise Nozzle Entry Temperature at the exhaust. Owing to a larger volume flow
(i.e. lower density) at exit from the afterburner, an increased nozzle flow area is required, to maintain
satisfactory engine matching, when the afterburner is alight.
Exhaust or Nozzle hot gases leaving the engine exhaust to atmospheric pressure via a nozzle, the objective
being to produce a high velocity jet. In most cases, the nozzle is convergent and of fixed flow area.
Supersonic nozzle if the Nozzle Pressure Ratio (Nozzle Entry Pressure/Ambient Pressure) is very high, to
maximize thrust it may be worthwhile, despite the additional weight, to fit a convergent-divergent (de Laval)
nozzle. As the name suggests, initially this type of nozzle is convergent, but beyond the throat (smallest flow
area), the flow area starts to increase to form the divergent portion. The expansion to atmospheric pressure
and supersonic gas velocity continues downstream of the throat, whereas in a convergent nozzle the expansion
beyond sonic velocity occurs externally, in the exhaust plume. The former process is more efficient than the
latter.
The various components named above have constraints on how they are put together to generate the
most efficiency or performance. The performance and efficiency of an engine can never be taken in isolation;
for example fuel/distance efficiency of a supersonic jet engine maximizes at about mach 2, whereas the drag
for the vehicle carrying it is increasing as a square law and has much extra drag in the transonic region. The
highest fuel efficiency for the overall vehicle is thus typically at Mach ~0.85.
For the engine optimization for its intended use, important here is air intake design, overall size,
number of compressor stages (sets of blades), fuel type, number of exhaust stages, metallurgy of components,
amount of bypass air used, where the bypass air is introduced, and many other factors. For instance, let us
consider design of the air intake.

Types, Description, Advantages And Disadvantages Of Jet Engines

There are a large number of different types of jet engines, all of which achieve propulsion from a high speed
exhaust jet.

Type

Description

Water jet

Can run in shallow water,

Can be less efficient than a
Squirts water out the
powerful, less harmful to wildlife, propeller, more vulnerable to
back through a nozzle
(indeed used by squid)
debris

Motor jet

Most primitive air

breathing jet engine.
Essentially a
supercharged piston
engine with a jet
exhaust.

Turbojet

A basic design, misses many

Generic term for simple Simplicity of design, efficient at improvements in efficiency
turbine engine
supersonic speeds (~M2)
and power for subsonic flight,
relatively noisy.

Turbofan

First stage compressor

greatly enlarged to
provide bypass airflow
around engine core,
and it provides
significant amounts of
thrust. Most common
form of jet engine in
use today- used in
airliners like the
Boeing 747 and
military jets, where an
afterburner is often
added for supersonic
flight.

Rocket

Advantages

Disadvantages

Higher exhaust velocity than a

Heavy, inefficient and
propeller, offering better thrust at
underpowered
high speed

Greater complexity
(additional ducting, usually
Quieter due to greater mass flow multiple shafts), large
and lower total exhaust speed,
diameter engine, need to
more efficient for a useful range contain heavy blades. More
of subsonic airspeeds for same
subject to FOD and ice
reason, cooler exhaust
damage. Top speed is limited
temperature.
due to the potential for
shockwaves to damage
engine.

Carries all propellants Very few moving parts, Mach 0 to Needs lots of propellant- very
and oxidants on-board, Mach 25+, efficient at very high low specific impulse
emits jet for propulsion speed (> Mach 10.0 or so),
typically 100-450 seconds.
thrust/weight ratio over 100, no Extreme thermal stresses of
complex air inlet, high
combustion chamber can
compression ratio, very high
make reuse harder. Typically
speed (hypersonic) exhaust, good requires carrying oxidizer on-

cost/thrust ratio, fairly easy to

test, works in a vacuum-indeed
works best exoatmospheric which
board which increases risks.
is kinder on vehicle structure at
Extraordinarily noisy.
high speed, fairly small surface
area to keep cool, and no turbine
in hot exhaust stream.

Ramjet

Must have a high initial speed

to function, inefficient at slow
Very few moving parts, Mach 0.8
speeds due to poor
to Mach 5+, efficient at high
Intake air is
compression ratio, difficult to
speed (> Mach 2.0 or so), lightest
compressed entirely by
arrange shaft power for
of all air-breathing jets
speed of oncoming air
accessories, usually limited to
(thrust/weight ratio up to 30 at
and duct shape
a small range of speeds,
optimum speed), cooling much
(divergent)
intake flow must be slowed to
easier than turbojets as no turbine
subsonic speeds, noisy, fairly
blades to cool.
difficult to test, finicky to
keep lit.

Strictly not a jet at all

a gas turbine engine
High efficiency at lower subsonic Limited top speed (airplanes),
Turboprop (Turbo is used as power plant
airspeeds (300 knots plus), high somewhat noisy, complex
shaft similar)
to drive propeller shaft
shaft power to weight
transmission
(or rotor in the case of
a helicopter)
Turboprop engine
drives one or more
Propfan/Unducted
propellers. Similar to a
Fan
turbofan without the
fan cowling.

Higher fuel efficiency, potentially

less noisy than turbofans, could
lead to higher-speed commercial
aircraft, popular in the 1980s
during fuel shortages

Development of prop fan

engines has been very
limited, typically more noisy
than turbofans, complexity

Pulsejet

Air is compressed and

combusted
Very simple design, commonly
intermittently instead
used on model aircraft
of continuously. Some
designs use valves.

Noisy, inefficient (low

compression ratio), works
poorly on a large scale, valves
on valved designs wear out
quickly

Pulse detonation
engine

Similar to a pulsejet,
but combustion occurs
Maximum theoretical engine
as a detonation instead
efficiency
of a deflagration, may
or may not need valves

Extremely noisy, parts subject

to extreme mechanical
fatigue, hard to start
detonation, not practical for
current use

Essentially a ramjet
where intake air is
Mach 0 to Mach 4.5+ (can also
compressed and burnt run exoatmospheric), good
with the exhaust from a efficiency at Mach 2 to 4
rocket

Similar efficiency to rockets

at low speed or
exoatmospheric, inlet
difficulties, a relatively
undeveloped and unexplored
type, cooling difficulties, very
noisy, thrust/weight ratio is
similar to ramjets.

Scramjet

Similar to a ramjet
without a diffuser;
airflow through the
entire engine remains
supersonic

Few mechanical parts, can

operate at very high Mach
numbers (Mach 8 to 15) with
good efficiencies[5]

Still in development stages,

must have a very high initial
speed to function (Mach >6),
cooling difficulties, very poor
thrust/weight ratio (~2),
extreme aerodynamic
complexity, airframe
difficulties, testing
difficulties/expense

Turbo rocket

A turbojet where an
additional oxidizer
such as oxygen is
added to the air stream
to increase maximum
altitude

Very close to existing designs,

operates in very high altitude,
wide range of altitude and
airspeed

Airspeed limited to same

range as turbojet engine,
carrying oxidizer like LOX
can be dangerous. Much
heavier than simple rockets.

Air-augmented
rocket

The motion impulse of the engine is equal to the air mass multiplied by the speed at which the engine
emits this mass:
I=mc
where m is the air mass per second and c is the exhaust speed. In other words, the plane will fly faster
if the engine emits the air mass with a higher speed or if it emits more air per second with the same speed.
However, when the plane flies with certain velocity v, the air moves towards it, creating the opposing ram
drag at the air intake:m v
Most types of jet engine have an air intake, which provides the bulk of the gas exiting the exhaust.
Conventional rocket motors, however, do not have an air intake, the oxidizer and fuel both being carried
within the airframe. Therefore, rocket motors do not have ram drag; the gross thrust of the nozzle is the net
thrust of the engine. Consequently, the thrust characteristics of a rocket motor are completely different from
that of an air breathing jet engine.
The air breathing engine is only useful if the velocity of the gas from the engine, c, is greater than the
airplane velocity, v. The net engine thrust is the same as if the gas were emitted with the velocity c-v. So the
thrust is actually equal to

10

S = m (c-v)
Turboprops have a wide rotating fan that takes and accelerates the large mass of air but by a relatively
small amount. This low speed limits the speed of any propeller driven airplane. When the plane speed exceeds
this limit, propellers no longer provide any thrust (c-v < 0).
Turbojets and other similar engines accelerate a much smaller mass of the air and burned fuel, but they
emit it at the much higher speeds possible with a de Laval nozzle. This is why they are suitable for supersonic
and higher speeds.
Low bypass turbofans have the mixed exhaust of the two air flows, running at different speeds (c1 and
c2). The thrust of such engine is
S = m1 (c1 - v) + m2 (c2 - v)
Where m1 and m2 are the air masses, being blown from the both exhausts. Such engines are effective
at lower speeds, than the pure jets, but at higher speeds than the turbo shafts and propellers in general. For
instance, at the 10 km attitude, turbo shafts are most effective at about 0.4 mach, low bypass turbofans
become more effective at about 0.75 mach and true jets become more effective as mixed exhaust engines
when the speed approaches 1 mach - the speed of sound.
Rocket engines are best suited for high speeds and altitudes. At any given throttle, the thrust and
efficiency of a rocket motor improves slightly with increasing altitude (because the back-pressure falls thus
increasing net thrust at the nozzle exit plane), whereas with a turbojet (or turbofan) the falling density of the
air entering the intake (and the hot gases leaving the nozzle) causes the net thrust to decrease with increasing
altitude. Rocket engines are more efficient than even scramjets above roughly Mach 15.
For all jet engines the propulsive efficiency (essentially energy efficiency) is highest when the engine
emits an exhaust jet at a speed that is the same as the airplane velocity.

RESULT
Thus the study of the jet engines and its components is completed.

11

EXP NO: 3
STUDY OF PERFORMANCE OF PROPELLER
AIM
To study the performance of the propeller.
BASIC PROPELLER PRINCIPLE
The aircraft propeller consists of two or more blades and a central hub to which the blades and are
attached. Each blade is essentially of rotating wing. As a result of their construction, propeller blade produce
forces/thrust to pull or push the aeroplane through air.
Power to rotate the propeller blades is furnished by the engines. Low powered engine propeller is
mounted on the propeller shaft and that is geared to the engine crank shaft.
PROPELLER NOMENCLATURE
In order to explain the theory and construction of propellers it is necessary first to define the parts of
various types of propellers and give the nomenclature associated with the propeller.
The cross section of a propeller blade is shown in the figure the leading edge of the blade trailing edge,
the cambered side, or back and the flat side or face. The blade has an aerofoil shape similar to that of an
aeroplane wing; it is through that it is a small wing; which has been reduced in length, width and thickness
(small wing shape). When the blade start rotating, airflows around the blade fast as it flows around the wing
of an aeroplane and blade is lifted forward
The nomenclature of an adjustable propeller is illustrated in the figure. This is metal propeller with
two blades clamped into a steel hub assembly. The hub assembly is supporting unit for the blades, and it
provides mounting structure in which propeller is attached to the engine propeller shaft. The propeller hub is
split on a plane parallel to the plane of rotation of the propeller to allow for the installation of the blades. The
sections of the hubs are held in place by means of clamping rings secured by means of bolts.
NOMENCLATURE FOR A CROUND ADJUSTABLLE PROPELLER
The figure shows two views of various cross sections of propeller blades. The blade shank is that
portion of the blade near the butt of the blade it is usually made thick to give its strength, and it is cylindrical
where it fits the hub barrel, but the cylindrical portion of the shank contributes little or no thrust. In order
designs, the aero foil shape is carried to the hub by means of blade cuffs which are thin sheet metals and it
function like cowling.
BLADE ELEMENT THEORY
The theory for the design of aircraft propeller was known as blade element theory. IT Is some time
referred to as the DRYE WIECKI theory as the polish scientist name is DRYE WIECKI.

12

The theory assumes to the tip of the blade is divided into various mall, rudimentary aerofoil sections.
For example if a propeller blade is 54 inch long and can be divided into 54 one-into aerofoil sections. Figure
shows one of these aerofoil sections located at radius r, the chord c will depend on the plan form or general
shape of the blade.
According to the blade element theory, many aerofoil sections or elements being joined together side
by side, unit to form an aerofoil (the blade) that can create thrust when revolving in a plane around central
axis.
The thrust developed by a propeller is in accordance. With Newtons third law of motion. In the case
of propeller the first action is acceleration of a mass of air to rear of the aeroplane. This means that if propeller
is exerting a force of 200 pounds in accelerating a given mass of air, it is the same time exerting at a force of
2000 pounds in pulling the aeroplane in the direction of opposite that in which the aeroplane is pulled
forward. The quantitative realization slip among mass, acceleration, and force can be determined by the use of
formula Newtons second law.
F=m*a
True pitch propeller is one that makes use of the blade. In elemental theory, each element of the blade
travels at different rates of speed that is tip section travels faster than the section closer to the hub.
Types of propeller:
Fixed pitch: The propeller is made in one piece. Only one pitch setting is possible and is usually two blades
propeller and is often made of wood or metal.
Wooden Propellers: Wooden propellers were used almost exclusively on personal and business aircraft prior
to World War II .A wood propeller is not cut from a solid block but is built up of a number of separate layers
of carefully selected .any types of wood have been used in making propellers, but the most satisfactory are
yellow birch, sugar maple, black cherry, and black walnut. The use of lamination of wood will reduce the
tendency for propeller to warp. For standard one-piece wood propellers, from five to nine separate wood
laminations about 3/4 in. thick are used.

Metal Propellers : During 1940 , solid steel propellers were made for military use. Modern propellers are
fabricated from high-strength , heat-treated,aluminum alloy by forging a single bar of aluminum alloy to the
required shape. Metal propellers is now extensively used in the construction of propellers for all type of
aircraft. The general appearance of the metal propeller is similar to the wood propeller, except that the
sections are generally thinner.

13

Ground adjustable pitch: The pitch setting can be adjusted only with tools on the ground before the engine
is running. This type of propellers usually has a split hub. The blade angle is specified by the aircraft
specifications. The adjustable - pitch feature permits compensation for the location of the flying field at
various altitudes and also for variations in the characteristics of airplanes using the same engine. Setting the
blade angles by loosened the clamps and the blade is rotated to the desired angle and then tightens the clamps.

Full Feathering: A constant speed propeller which has the ability to turn edge to the wind and thereby
eliminate drag and wind milling in the event of engine failure. The term Feathering refers to the operation of
rotating the blades of the propeller to the wind position for the purpose of stopping the rotation of the
propeller to reduce drag. Therefore, a feathered blade is in an approximate in-line-of-flight position,
streamlined with the line of flight (turned the blades to a very high pitch). Feathering is necessary when the
engine fails or when it is desirable to shutoff an engine in flight.

Some of the terminologies used in propeller design :

Two-position: A propeller which can have its pitch changed from one position to one other angle by the pilot
while in flight.
Controllable pitch: The pilot can change the pitch of the propeller in flight or while operating the engine by
mean of a pitch changing mechanism that may be operated by hydraulically.
Constant speed: The constant speed propeller utilizes a hydraulically or electrically operated pitch changing
mechanism which is controlled by governor. The setting of the governor is adjusted by the pilot with the rpm
lever in the cockpit. During operation, the constant speed propeller will automatically change its blade angle

14

to maintain a constant engine speed. If engine power is increase, the blade angle is increased to make the
propeller absorb the additional power while the rpm remain constant. At the other position, if the engine
power is decreased, the blade angle will decrease to make the blades take less bite of air to keep engine rpm
remain constant. The pilot selects the engine speed required for any particular type of operation.
Reversing: A constant speed propeller which has the ability to assume a negative blade angle and produce a
reversing thrust. When propellers are reversed, their blades are rotated below their positive angle, that is,
through flat pitch, until a negative blade angle is obtained in order to produce thrust acting in the opposite
direction to the forward thrust. Reverse propeller thrust is used where a large aircraft is landed, in reducing the
length of landing run.
Beta Control: A propeller which allows the manual repositioning of the propeller blade angle beyond the
normal low pitch stop. Used most often in taxiing, where thrust is manually controlled by adjusting blade
angle with the power lever.
Blade Station
Blade stations are designated distances in inches measured along the blade from the centre of the hub
the figure shows the location of a point on the blade at the 42 inches in each station this division of blade into
station provides a convenient means of discussing the performance of the propeller blade locating blade
marking and damage finding the proper point for measuring the blade angle and locating anti-glare areas
Blade Angle
Blade angle is defined as the angle between the chord particular blade section and the plane of rotation
Blade Pitch
Blade pitch is the distance advanced by the propeller in one revolution
Geometric Pitch
The propeller would have been advanced in one revolution
Experimental Mean Pitch
The distance traveled by the propeller in one revolution without producing thrust
Effective Pitch
Actual distance advanced by the propeller in one revolution
Pitch Distribution
The angle gradually decreases towards the tip and towards the shank
Angle Of Attack
This is the angle formed between the chord of the blade and direction of relative air flow
Propeller Slip
Slip is defined as difference between the geometric pitch and the effective pitch
Forces Acting On A Propeller

15

Thrust force
Centrifugal force
Torsion or twisting force
Aerodynamic twisting force
Aerodynamic twisting movement (ATM)
Centrifugal twisting movement (CTM)
Thrust Force
Thrust force is a thrust load that tends to bend propeller blade forward as the aircraft is pulled through
the air
Centrifugal Force
Centrifugal force is the physical force that tends to throw the rotating propeller blades away from the
hub
Torsion or Twisting Force
Torsion force is the force of air resistance tends to bend the propeller blade in a direction that is
opposite to the direction of rotation
Aerodynamic Twisting Force
It is the force that tends to turn the blade to higher blade angle
Aerodynamic Twisting Moment
It is the force that tends to turn the blade angle towards low blade angle
Propeller Efficiency
Propeller efficiency has been achieved by use of this aerofoil section near the tips of the propeller
blades and very sharp leading and trailing edge
.
Propeller efficiency id calculated = thrust horsepower / torque horse power
It is the ratio of thrust horse power to the torque horse power. Thrust horse power is the actual amount
of horse power that an engine propeller transforms x thrust
Propeller Chart
For a given pitch angle B, the efficiency of the propeller is a function of dimensionless quantity T, the
advance ratio such as a plot for a family of pitch angle that is valuable in a propeller can be plotted. This is
called the propeller chart.
RESULT
Thus the study of performance of the propeller is successfully completed.

16

EXP NO: 4
STUDY OF WALL JET
AIM
The main objective of this experiment is to study the performance of wall jet in a flow field.
THEORY
Turbulent wall jet flows consist of two self-similar layers: a top layer and a wall layer, separated by a
mixing layer where the velocity is close to maximum. The top and wall layers are significantly different from
each other, and both exhibit incomplete similarity, i.e., a strong influence of the width of the slot that had
previously been neglected.

17

EXPERIMENTAL APPARATUS AND PROCEDURE

This wall is usually either a thin lip or an "infinite" vertical wall as in the present experiment. The
latter design is simpler to treat computationally, since it, together with a "no inflow" upper boundary, results in
a single, well-defined inflow boundary with known boundary conditions. It was therefore chosen here, in spite
of the inevitable return flow that this configuration generates, a return flow which far downstream of the
nozzle changes the character of the jet. An important criterion in the experimental design was that the spatial
resolution should be sufficiently high to allow the wall shear stress to be determined directly from mean
velocity measurements. This imposes an upper limit on the ratio of measuring control volume diameter to
viscous length scale, but a high enough inlet Re-number must also be retained to allow comparisons with
earlier studies. Once water was chosen as the working fluid, due to the absence of seeding problems in lowspeed water flows, these considerations led to the present combination of slot width and inlet velocity.

WALL JET TEST FACILITY

18

The test facility is shown in Fig. It consists of a large tank into which a jet discharges. The tank is 7 m long
and its width is 1.45 m. One of the side walls is made of glass, as well as the bottom. (Using a glass bottom
improves the conditions for near-wall measurements, since its smoothness minimizes the diffuse surface
reflections. The slot height was measured with water in the tank, by a diver. The results showed the slot height
to be 9.6 + 0.1 mm over most of the slot width. Given the uncertainties involved, this is consistent with an
indirect determination of the slot height using the volumetric flow rate. Consequently, b = 9.6 mm will be
used in the following analysis, giving a jet width-to-height ratio of 151. This was considered large enough to
obtain good two-dimensionality. A large contraction (Morel 1975) with a turbulence-reducing screen inserted
is used to produce a fairly flat mean velocity profile at the inlet. A weir upstream of the contraction keeps the
upstream water level constant, and the flow velocity through the slot is set by an adjustable weir at the
downstream end at the tank. This reference velocity is determined as

Where h is the difference in height between the upstream and downstream free surfaces.
The inlet velocity, U0, was set as close as possible to 1 m/s, corresponding to a water depth downstream of the
inlet of about 1.4 m. For this water depth, the influence of the re-circulating flow on the growth rate of the jet
was negligible for the first 150 slot heights.
Using water of approximately room temperature, one obtains a nominal inlet Re-number

EXPERIMENTAL PROCEDURE AND FLOW QUALIFICATION

Outline of measurements
Extensive Pitot-tube measurements, spanwise profiles at several heights and numerous vertical profiles
at different spanwise positions, were made at the slot (x = 0) to check for symmetry and spanwise variations.
Part of the inlet velocity profile was also measured using LDV, to better resolve the boundary layer and to get
turbulence data. LDV measurements, streamwise and spanwise profiles, were also taken immediately
downstream of the slot. Extensive spanwise measurements were made at several streamwise positions in order
to check the two-dimensionality of the flow. Based on these spanwise measurements, it was decided to make
the main measurements series approximately halfway between the centreline and the glass wall. The flow
conditions in that spanwise position were identical to those at the centreline within the measurement accuracy.
The main measurement series were taken at the following streamwise positions: x=50, 100, 200, 400,
700,1000, 1500, 2000 mm. For the sake of simplicity, we will refer to these positions as x/b = 5, 10, 20, 40,
70, 100, 200, although the actual dimensionless distance was about 4% larger. In figures showing the
streamwise development of a quantity, the correct x/b will be used. Measurements stopped at x/b = 200
because the flow was losing its wall-jet character. This issue will be discussed later on.

Main measurement series

19

The vertical profiles of the main measurement series were taken in order from x/b = 5 and
downstream. h and the water temperature, T>sub>0>, was checked regularly, in order to detect any drift in
inlet velocity or inlet Re-number. There was essentially no change in U0 or Re0 during the individual profiles.
There were, however, small variations between the different profiles due to a 3% variation in the boundary
conditions, i.e. Re0. Where relevant, all velocities have been normalised to the same inlet velocity by
multiplying with [U0(x=0)/U0(x=X)].
The position of the wall, y = 0, was estimated by observing the output signal from the counter, i.e.
after amplifying and filtering, on an oscilloscope. The "wall signal" is very characteristic. The distance from
this preliminary wall position ; was then measured by a dial gauge. Finally, the wall distance was adjusted
after the measurements by shifting the velocity curve up or down to make it pass through origin. This was
relatively simple due to the linear relation. The necessary adjustments typically were of the order of 0.02 mm.

Inlet conditions
The inlet conditions were determined using Pitot tube and LDV-measurements. Mean velocity profiles
from Pitot tube measurements, taken at several spanwise positions at and around the spanwise position finally
chosen for the main measurements, showed no visible differences in the maximum velocity. There were,
however, small differences in the length of the flat parts of the profiles. These are consistent with the earlier
statement of a +0.1 mm variation in slot height. The variation in the spanwise velocity distribution at y = 4.5
mm was less than 0.25%.
LDV measurements of the lower part of the inlet velocity profile were made in order to resolve the
boundary layer and to obtain information on the turbulence levels. The boundary layer thickness, defined as
U= 0.99 Umax, is 1.4 mm. The turbulence intensity in the flat part of the profile is less than 1%. No
corrections for gradient broadening has been applied to the turbulence measurements, meaning that the peak
in turbulence intensity in the boundary layer is exaggerated. We thus have a fairly flat inlet velocity profile
with a mean velocity which is uniform in the spanwise direction within 0.25%. The flow is laminar and the
laminar boundary layers along the walls have a thickness 1.4 mm.
Persistent spanwise variations of the thickness of the wall jet were noted. These variations are
probably associated with the small variation (1%) in slot height. All subsequent measurements were however
made at a spanwise position where "average properties" of the wall jet were prevailing.
In many technical applications, impinging jets are used for cooling and heating tasks when large heat
transfer coefficients are required. Therefore, many experimental investigations have been performed to study
this subject. Overviews of this topic are given in.
In submerged free and impinging jets, fluid exits from a nozzle into a resting environment. Fluid is
entrained into the jet and is accelerated. The jet becomes broader and the jet velocity decreases as a result of
momentum preservation. The core jet, where the initial conditions are still present, becomes smaller with
increasing distance from the nozzle. Depending on the initial conditions, the core jet has disappeared after 4
nozzle diameters to 6 nozzle diameters for the nonpulsating jet. The resulting velocity profile can be
described with a Gaussian curve. In impinging jets the fluid flows toward a wall and is decelerated and
changes its direction. Depending on the flow regime, the wall is placed in the stream downward flow pattern
and especially the heat transfer between the wall and jet shows different characteristics. In the stagnation
regionthe region where the jet is influenced by the wallthe fluid is decelerated in the axial direction and
accelerated in the radial direction. Directly at the stagnation point, the velocity is zero. With increasing radial

20

distance from the stagnation point, the flow is accelerated in the radial direction. The acceleration is conserved
up to that point, where no more fluid from the free jet is mixed into the wall jet. Especially in cases with low
nozzle-to-plate distances, the boundary layer is laminar in this region and is stabilized by the acceleration.
By mixing of fluid from the environment into the jet and increasing wall jet thickness and by
increasing cross-sectional area in axisymmetric jets, the wall jet velocity is reduced. The flow becomes
instable and turbulent. The maximum velocity parallel to the wall is obtained at a distance of 1-2 nozzle
diameters from the stagnation point. At this point, major changes in heat transfer are observed. At greater
distances from the stagnation point, a turbulent wall jet is present.
In Figure 1 the flow regions in impinging jets are illustrated.

In technical applications, impinging jets are used when large heat transfer coefficients are required. Especially
in the region of maximum radial velocity, heat transfer coefficients are obtained which can hardly be achieved
with other flows without phase change.
The radial evolution of heat transfer coefficients is influenced mainly by the Reynolds number and the
nozzle-to-plate-distance. At small Reynolds numbers, heat transfer decreases monotonically with increasing
radial distance. At large Reynolds numbers, a similar characteristic can only be found for large nozzle-toplate-distances (H/D > 6). At small and medium distances, a slight decrease of heat transfer coefficient in a
radial distance of up to 1 nozzle diameter is followed by a large increase up to a local maximum at 1.5-2
diameters and a monotonic decrease for larger radial distances.
Calculating a simple turbulent channel flow is possible with nearly all turbulence models without
difficulty, but flow patterns are present in impinging jets which are difficult to predict with classical
turbulence models:

Entrainment of fluid from the environment and prediction of the spreading angle, connected with the
increased turbulence level in the jet
Relaminarization near the stagnation point
Large acceleration of the flow, followed by a deceleration
Laminar-turbulent transition in the wall jet

21

Different curve characteristics of radial heat transfer evolution, depending on flow velocities and
nozzle-to-plate-distances

Therefore, jet impingement is often used as benchmark flow for improving turbulence models.The first
comprehensive examinations were performed. In their examination with classical k - models, heat transfer
could be predicted well in the wall jet region, while in the stagnation region the prediction was still poor.
Within the last years, Durbin's
model became popular, which also gives good prediction of heat
transfer in the stagnation region. In Durbin's model, additional transport equations are required for predicting
turbulent flow, which increases computational effort.
A large number of other works can also be found in which existing turbulence models have been used and
in which correction terms or modified parameters are applied to give better prediction of heat transfer in
impinging jets.
A comprehensive overview of earlier turbulence models can be found. A newer overview of common
turbulence models for predicting jet impingement heat transfer can be found.
In the present work, several commercially available turbulence models have been tested for their ability to
predict heat transfer and flow structure in impinging steady jets. It is not the aim of this work to improve the
models, but rather, to give a comparison of 13 widely used turbulence models in terms of their ability to
predict impinging jets.
For one promising model, it was tested how sensitively this model reacts to changes in turbulence
intensity and how the model can predict pulsating impinging jets. The calculations have been compared to our
own experimental data. From these results a recommendation can be made of with which settings heat transfer
in impinging jets can be predicted best. After that, we examine how these results can be transferred to
pulsating impinging jets.

22

RESULT
Thus the wall jet experiment is studied and experimental procedures are discussed successfully
EXP NO: 5
FREE JET
AIM
The main objective of this experiment is to study the performance of free jet
THEORY
Flow Properties of a Rectangular Jet

23

Figure 1. Schematics of a free jet flow and its downstream development

A jet is formed by flow issuing from a nozzle into ambient fluid, which is at a different velocity. If the
ambient fluid is at rest the jet is referred to as a free jet; if the surrounding fluid is moving, the jet is called a
co flowing jet. A jet is one of the basic flow configurations which have many practical applications such as in
jet engines, combustors, chemical lasers, ink-jet printer heads, among others. Figure 1 illustrates some
essential features of a jet. The velocity at the exit of the nozzle of a typical laboratory jet has a smooth profile
and a low turbulence level, about 0.1% - 0.5% of the mean velocity. Due to the velocity difference between
the jet and the ambient fluid, a thin shear layer is created. This shear layer is highly unstable and is subjected
to flow instabilities that eventually lead to the formation of large-scale vortical structures (see Figure 1). The
interaction of these structures produces strong flow fluctuations, entrains ambient fluid into the jet flow and
enhances the mixing. The shear layer and consequently, the jet, spread along the direction perpendicular to the
main jet flow.
The central portion of the jet, a region with almost uniform mean velocity, is called the potential core.
Because of the spreading of the shear layer, the potential core eventually disappears at a distance of about four
to six diameters downstream from the nozzle. The entrainment process continues further beyond the end of
the potential core region such that the velocity distribution of the jet eventually relaxes to an asymptotic bellshaped velocity profile as illustrated in Figure 1. Also shown in Figure 1 is the half-width of the jet, y1/2,
defined as the distance between the axis and the location where the local velocity equals half of the local
maximum or centerline velocity, U0. The increase in the jet half-width with downstream distance provides a

24

measure of the spreading rate of the jet. Due to the spreading, the jet centerline velocity, Vc, decreases
downstream beyond the potential core region.
Apparatus

The following apparatus will be used for this experiment:

1. A rectangular jet, with a nozzle of dimensions 6 cm 1 cm.
2. An air pump to force air stream through the jet nozzle.
3. A pitot-static tube and a digital manometer.
4. A Pentium-based PC with LabVIEW software and an associated ADC card.
Experimental Procedure
Jet Centerline and Cross-Stream Velocity Profile Measurements
1. Set the dynamic pressure of the jet exit velocity at the maximum stable setting (usually between 0.06 and
0.07 psi). Note: the digital pressure gage has an upper limit of 0.1 psi. Do not overload the unit!
2. Beginning at a position approximately at the jet nozzle, move the pitot-static probe along the center axis of
the jet. Measure the jet centerline velocity at 1 cm intervals for 31 data points.

25

3. Move the probe to a downstream location of x/d = 4. (The height d of the jet nozzle is 1 cm) measure the
cross-stream velocity profile by using the traverse to move the probe in the vertical direction and recording
the output using lab view.
4. A total of 8 points with a 1 mm increment should be measured.
5. Move the probe to three other downstream locations at x/d = 10, 20 and 30 and measure the velocity
profiles.
6. Record the ADC output for this location also.
7. Use 8 points and a 3 mm increment.
RESULT
Thus the free jet experiment is studied and experimental procedures are discussed successfully

EXP NO: 6

PRESSURE DISTRIBUTION OVER A SYMMETRICAL AEROFOIL

AIM
To determine the pressure distribution over the given symmetrical aerofoil model and to plot the graph
between x/c and Cp

26

APPARTUS REQUIRED

Wind tunnel setup

symmetrical aerofoil with pressure tapping
Multi-bank manometer

FORMULA USED
1. Gauge Pressure
Pg=gh
Where,
=density of manometer fluid
g=Gravitational acceleration
h=Pressure head
h= hn - ho
n=1,2,3,4..20

2. Pressure Coefficient
Cp

= (Ps -P) / (1/2**V )

= Pg / (v2/2)

PROCEDURE

Check the three phase power supply

Switch on the three phase power supply
Clean the model.
Fix the given model at a given angle of attack using string and lock it.
Connect the pressure tapings of the model to the corresponding point in the multi-bank manometer.
Switch on the wind tunnel set up.

27

Set the force indicators to zero.

Set the air velocity to a given value.
Note down the corresponding forces which are indicated in the display
Repeat the same procedure for the different velocities at the same angle of attack and note down the
corresponding forces for different velocities.

TABULATION: Pressure readings at various flow velocities

Flow

1
0

11

12

13

14

15

16

17

18

19

velocity
in m/s
h
in
m

pg
in
n/
m2
cp
x
x/c

GRAPH
Pressure coefficient versus position of pressure taping on the symmetrical airfoil. i.e Cp vs x/c

28

20

RESULT
Thus the pressure distribution over the symmetrical aerofoil was determined and the graph was
plotted.

EXP NO: 7

CASCADE TESTING OF MULTIPLE AEROFOIL SECTION OF TURBINE BLADES

AIM
To determine the pressure distribution over the given symmetrical aerofoil model and to plot the graph
between x/c and Cp

29

APPARTUS REQUIRED

Wind tunnel setup

symmetrical aerofoil with pressure tapping
Multi-bank manometer

FORMULA USED
1. Gauge Pressure
Pg=gh
Where
=density of manometer fluid
g=Gravitational acceleration
h=Pressure head
2. Pressure Coefficient
Cp

= (Ps -P) / (1/2**V )

= Pg / (v2/2)

PROCEDURE

Check the three phase power supply

Switch on the three phase power supply
Clean the model.
Fix the given model at a given angle of attack using string and lock it.
Connect the pressure tapings of the model to the corresponding point in the multi-bank manometer.
Switch on the wind tunnel set up.
Set the force indicators to zero.
Set the air velocity to a given value.
Note down the corresponding forces which are indicated in the display
Repeat the same procedure for the different velocities at the same angle of attack and note down the
corresponding forces for different velocities.

TABULATION : Pressure readings at various flow velocities

30

Flow

Upper aerofoil

Middle aerofoil

Lower aerofoil

Velocity
In m/s

h
in
m

Pg
in
N/
m2
Cp
X
X/
c

GRAPH
Pressure coefficient versus position of pressure taping on the symmetrical airfoil. i.e Cp vs x/c

RESULT
Thus the pressure distribution over the symmetrical aerofoil was determined and the graph was
plotted.
EXP NO: 8
NOZZLE PERFORMANCE TEST
AIM
To conduct a performance on a nozzle for determining,
Effect of back pressure on mass flow rate
Jet velocity and nozzle efficiency for various operating pressure

31

APPARATUS REQUIRED
1. Nozzle pressure test unit
2. Compressed air
FORMULA USED
1.
2.
3.
4.
5.

Theoretical mass flow rate = {0.0404 At* P1} / {(T1)^ 1/2}

Recorded mass flow rate = (mass flow in m/s) * (area in m2)* (density in kg/m3)
Nozzle efficiency = specific kinetic energy / isentropic enthalpy change
Specific kinetic energy (in J/Kg) = C22 / 2
Isentropic enthalpy change = {(RT)* (1-rp) ^ [(-1)/ ])} / {-1}
Where, pressure ratio, rp = P2/ P1 and = 1.4
Density = 1.129 Kg/m3
D= 0.075 m
A= D2 / 4

PROCEDURE
1. Close the valve V2 and V3
2. Open the valve V4 and V1
3. Closed the pressure regulator valve
4. Closed the Compressor line valve
5. Start the compressor and maintain the pressure in the range of 8- 10 kg/m2
6. Open the outlet valve of compressor tank.
7. Switch on the system.
8. Ensure the functioning of indicators
9. Gradually open compressor line valve and adjust the pressure regulator valve
10. Take the 3 readings of pressure, temperature and force.
11. Open the valve V2.
12. Closed the valve V1.
13. Gradually closed V4, and take readings of pressure temperature
14. Draw the graph,

Pressure ratio Vs mass flow rate

Pressure ratio Vs nozzle efficiency

TABULATION
1. Determination of effect of back pressure on mass flow rate

32

Observation table: 1
S.No
Inlet
Chamber
pressure
pressure
(P1)
(P1)
in kg / m2
in kg / m2
1
2
3

Mass flow in
(m/s)
Inlet Chambe
r

Temperature in c
T1

Force in
N

T2

Table for calculation: 1

Recorded
Chamber mass flow rate
In m/s

In kg/s

Chamber
pressure
(P2)
In kg / m2

Theoretical
Inlet
pressure
(P1)
In kg / m2

Chamber
mass flow
rate
In kg / s

P2/P1

2. Determination of Jet Velocity and Nozzle Efficiency

Observation table: 2
S.No

Inlet
pressure
(P1)
in kg / m2

Chamber
pressure
(P2)
in kg / m2

Temperature
in c

Mass flow in (m/s)

Inlet

Chamber

Force in
N

1
2
3
4
5
Table for calculation: 2
Chamber mass
flow rate in kg/s

C2 in m/s
C2= F/ m

P2/P1

33

Nozzle efficiency

RESULT
The jet velocity and nozzle efficiency is found
The required graph is plotted
Thus the nozzle performance test is conducted successfully

EXP NO: 9

NOZZLE PRESSURE DISTRIBUTION TEST

AIM
To determine the effect of inlet pressure on the mass flow rate with constant back pressure

34

 To determine the effect of back pressure on mass flow rate with constant inlet pressure

APPARATUS REQUIRED

Centrifugal air compressor

Nozzle pressure distribution unit

FORMULA USED
1. Theoretical mass flow rate = {0.0404 At* P1} / {(T1)^ 1/2}
2. Recorded mass flow rate = (mass flow in m/s) * (area in m2)* (density in kg/m3)
Where,
Density = 1.129 Kg/m3
A= D2 / 4 = 0.0007065 m2

PROCEDURE

Before starting the compressor, open V12, V11, and inlet valve fully.

Pressure regulator valve should be kept closed.

Compressor line valve should be kept closed.

Start the compressor and maintain the pressure in the range of 8- 10 kg/m2

Open the outlet valve of compressor tank.

Switch on the system.

Ensure the functioning of indicators

Gradually open compressor line valve and adjust the pressure regulator valve.

Open valve V2 and take the readings of pressure p2, temperature, and mass flow rate.

Take 3 reading by adjusting pressure regulator valve.

Gradually closed V12 and Take reading of pressure P2 by maintaining pressure regulator reading
of 2 kg/cm2 for varying back pressure.

35

TABULATION
1. Determination of effect of inlet pressure on the mass flow rate with constant back pressure.
Inlet pressure

Flow rate

in kg / m2

In m/s

Practical mass
rate

Pressure at P2
in kg /m2

In kg / s

Theoretical mass
flow
In kg / s

2. Determination of effect of back pressure on the mass flow rate with constant inlet pressure
Temperature

P2

in c

In kg / m2

in m/s

Recorded
Mass flow in m/s

Mass flow rate in kg/s

Theoretical Mass flow

rate in kg/s

36

Back pressure
In kg / m2

RESULT

The effect of inlet pressure and back pressure on mass flow rate is determined

Thus the pressure distribution test is conducted successfully

EX.NO:10

BOMB CALORIMETER
INTRODUCTION

37

A bomb calorimeter will measure the amount of heat generated when matter is burnt in a sealed
chamber (bomb) in an atmosphere of oxygen gas.
This isothermal bomb calorimeter provides a simple inexpensive yet accurate method for
determination of heat of combustion (calorific value) of solid and liquid fuels. The out fit is complete for
analysis as per method recommended by ISI (IS 1359-1954).
AIM
To determine the calorific value of the given solid or non-volatile liquid fuel using a bomb calorimeter
OPERATING PRINCIPLE
A known amount of sample is burnt in a sealed chamber (bomb) the air is replaced by pure oxygen.
The sample is ignited electrically. As the sample burns heat is generated. The raise in temperature is noted
since baring loss of heat the amount of heat generated by burning of the sample must be equal to the amount
of heat absorbed by the calorimeter assembly. By knowing the energy equivalent of the calorimeter and the
temperature raise, the calorific value can be found out.
PROCEDURE

Find the weight of the empty crucible using a physical balance.

A small quantity of liquid fuel (diesel) is taken in the crucible and is again weighed with fuel in it.
The crucible with fuel is placed over the support. A fuse wire is connected between the electrodes.
The bomb is closed air tight and is filled with oxygen at a pressure of about 25 bars.
The bomb is placed inside the calorimeter vessel filled with water. Noted the initial temperature of
water using the digital thermometer.
The calorimeter water is stirred using a motor drive. The fuel is ignited electrically by passing a high
voltage through the fuse wire which causes the fuse wire to burn.
Heat liberated by the fuel causes the temperature to rise.
After steady condition is reached the temperature raise is measured using the digital thermometer
provided.

OBSERVATION
Weight of the crucible without fuel (m1)
=
Weight of the crucible with fuel (m2) =
gm
Initial reading of the digital thermometer (t1) =
Final reading of the digital thermometer (t2) =

CALCULATION
Mass of fuel burnt (m)
Temperature rise (t)

= m2 - m1
= t2 - t1

38

gm
.
.

c
c

W = energy equivalent of the calorimeter assembly = 9735 J/.c

Cv = calorific value of fuel in J/gm or KJ/Kg
Then W * t = Cv * m
Cv = W * t / m

RESULT
Thus the calorific value of given solid or non volatile liquid fuel is found using bomb calorimeter.

39