Compressors

Introduction to Compressors A compressor is a mechanical device that increases the pressure of

a air by reducing its volume. Compressors are work absorbing devices which are used for increasing pressure of fluid at the expense or work done on fluid.

The compressors used for compressing air are called air compressors. Work required for increasing pressure of air is available from the prime mover driving the compressor.

Generally, electric motor, internal combustion engine or steam engine, turbine etc. are used as prime movers. Compressors are similar to fans and blowers but differ in terms of pressure ratios. Compressors are similar to pumps: both increase the pressure on a fluid and both can transport the fluid through a pipe.

Compressor

Compressed air is a air which is kept under a certain pressure, usually greater than that of the atmosphere.

Compressed air can be used in or for:

pneumatics, the use of pressurized air to do work.

Air dusters for cleaning electronic components that cannot be cleaned with water.

railway braking systems

road vehicle braking systems.

Types of compressors

RECIPROCATING COMPRESSOR

ROTARY COMPRESSOR

compressor work on two principles

1)Reduce volume of a constant amount of air 2)Adding more gas/air in a constant amount of volume .

positive displacement compressor works on first principle it reduces the volume of air by applying force on it but air amount is constant in every stroke or rotation thus increasing the pressure.

centrifugal & axial flow compressor works on second principle it adds more amount of air in a given constant volume thus the pressure increase.

The basic requirement of compressor for aircraft gas turbine application are well known. 1.High air flow capacity per unit frontal area 2.High pressure ratio per stage. 3.High efficiency. 4.Discharge direction suitable for multistaging. The compressor should be designed in such a way to have 1.Minimum length 2.Weight must be as low as possible. 3.The mechanical design should be simple , so as to reduce manufacturing time and cost. 4.High reliability.

Axial flow Compressor History

The basic concept of multistage axial flow compressor operation have been known for approximately 100 years being presented to French academic des science in 1853.

Efficiencies for this type of unit were quite low. Because the blading was not designed for the condition of a pressure rise in the direction of flow.

Beginning of at the turn of 20th century, a number of axial flow compressors were built , in some cases with the blade design based on propeller theory.

The efficiency of these units was still low (50-60%).Due to lack of sufficient knowledge of fluid mechanics at that time.

The advances in aviation during the period of WW I and rapidly developing background in fluid mechanics and aerodynamics give a new impetus to research on compressors.

Axial flow Compressor History

In 1936 the Royal aircraft establishment in England began the development of axial flow compressors for jet propulsion.

Aerodynamic theory was developed specifically for the case of a cascade airfoils.

By 1945 , compressors of high efficiency could be developed by incorporating aerodynamic principles in design and

development.

Geometry and Working principle

The energy level of air or gas flowing through it , is increased by the action of the rotor blades which exert

a torque on the fluid.

This torque is supplied by an external source an electric motor or gas turbine.

Its applications in the industrial gas turbine units the multistage axial compressor is the principle element

of all gas-turbine power plants for land and

aeronautical application.

Axial Flow Compressor

An axial flow compressors are given more preferred then the

radial flow type in the applications of aircraft and industrial gas

turbines . Because axial flow compressor has high efficiency and

is capable of producing higher pressure ratio on single shaft.

The stage pressure ratios of about 1.15:1 are obtained and by combining the stages , the overall pressure ratios of upto 8:1 or

even higher can be achieved.

The axial flow compressors consists of a number of stages where each stage may be considered as a fan.

The main advantage of axial flow compressors are large air handling abilities with a small frontal area ,a straight through

flow systems and high pressure ratios with relatively high

efficiencies.

The main disadvantages is its complexity and cost.

Axial flow compressors An axial flow compressors is composed of an alternating sequence of

fixed and movable sets of blades.

The set of fixed blades are spaced around the inside periphery of an outer stationary casing, and together constitute stator.

The set of movable blades are fixed to a spindle and the combination constitutes the rotor.

The radius of rotor hub and the length of the rotor blades are designed so that there is only a very small tip clearance at the end of the stator

and rotor blade.

The rotor and stator banks are as close as possible for efficient flow.

One set of stator blades and one set of rotor blades constitute a stage.

There are number of stages in compressors depending upon the pressure ratio required.

The successive set of blades are reduced in length to compensate for the reduction in volume resulting from the increased pressure.

Axial flow compressors

1. The K.E is imparted to the air by means of the rotating blades which is converted into a pressure rise.

2. The air enters axially in to the inlet guide vanes where it is turned through a certain angle to impinge on the first row of rotating blades with proper angle of attack.

3. The rotating guide vanes add K.E. to the air. Here slight pressure rise also takes place. The air then is discharged at the proper angle to the first row of stator blades where the pressure is further increased by diffusion.

4. The air then directed to second row of moving blades and the process is repeated through the remaining stages of the compressors.

5. Usually at entry one more stator is provided to guide the air correctly into the first rotor. This blades are some times referred as the Inlet Guide Vanes(IGV).

6. In many compressors there are one to three rows of diffuser or straightener blades installed after the last stage to straighten and slow down the air before it enters into the combustion chamber.

Axial flow compressors

Axial flow compressors

Selection of Pressure Ratio per Stage

Stage velocity triangle

1. The flow geometry at the entry and exit of the compressor stage is described by the velocity triangles at these stations .

2. The velocity triangles for the compressor stage contain, besides peripheral velocity(u) of the rotor blades both the absolute(c) and relative (w)fluid velocity vectors.

3. Velocity triangles are typically used to relate the flow properties and blade design parameters in the relative frame (rotating with the moving blades), to the properties in the stationary or absolute frame.

Velocity triangle

The air angles of absolute and relative systems are denoted by 1, 2, 3 and 1, 2, 3, respectively.

If the flow is repeated in another stage then

c1 = c3 and 1 = 3

subscripts a and t denote axial and tangential directions respectively.

Thus the absolute swirl or whirl vectors ct1 and ct2 are the tangential components of absolute velocities c1 and c2 respectively .

similarly wt1 & wt2 are the tangential components of the relative velocities w1 & w2 respectively.

The following trigonometrical relations obtained from velocity triangles.

From velocity triangles at the entry:

ca1 = c1 cos1 = w1 cos1 ------------------------------------1 ct1 = c1 sin1 = ca1 tan1 ------------------------------------2 wt1 = w1 sin1 = ca1 tan1 ------------------------------------3

u = ct1 + wt1 -----------------------------------4

u = c1 sin1 + w1 sin1 -----------------------------------5 u = ca1 ( tan1 + tan1 ) ------------------------------------6

From velocity triangles at the exit:

ca2 = c2 cos2 = w2 cos2 ------------------------------------7 ct2 = c2 sin2 = ca2 tan2 ------------------------------------8 wt2 = w2 sin2 = ca2 tan2 ------------------------------------9

u = ct2 + wt2 ------------------------------------10

u = c2 sin2 + w2 sin2 -----------------------------------11 u = ca2 ( tan2 + tan2 ) ------------------------------------12

for constant axial velocity through the stage: ca1 = ca2 = ca3 = ca ------------------------------------13 ca = c1 cos1 = w1 cos1 = c2 cos2 = w2 cos2 ------------------------------------14

Equation 6 &12

u/ ca = 1/ = ( tan1 + tan1 ) = ( tan2 + tan2 ) ------15 This relation can also be presented in another form using eqn 4 & 10

ct1 + wt1 = ct2 + wt2

ct2 - ct1 = wt1 - wt2 -----------------------------------16

ca ( tan2 - tan1 ) = ca ( tan1 - tan2 )----------17

Equations 16 & 17 give the change in the swirl components across the rotor blade row .For steady flow in an axial machine, this is proportional to the torque exerted on the fluid by the rotor.

Work input to the compressor Compressor work input in terms of velocity and blade angles . The compressor work input derived based on the assumption that the axial velocity remains constant throughout the machine.

From eqn 15

u = ca( tan1 + tan1 ) = ca( tan2 + tan2 ) Form Eulers eqn for turbo machinery the power needed by rotor is

P = T = (ct2r2 - ct1r1) where = u1/r1 = u2/r2

Above eqn becomes

P = (ct2u2 - ct1u1)

Dividing above eqn by we will get workdone or specific power

W = u(ct2-ct1)

W= u ca( tan2 - tan1 )

In terms of

W = uca ( tan1 - tan2 )

According to Eulers (turbo machinery) energy equation

W = {(c22 c1

2)+(u22 u1

2)+(w12 - w2

2)}

For axial flow compressors u=u1=u2 the above equation reduced to

W = 1/2(c22-c1

2)+1/2(w12-w2

2)

To obtain higher efficiency the work input should be as minimum as possible . To achieve this , the proper care in the design of blade and flow geometries are essential.

Work done factor()

The reduction in work absorbing capacity of the compressor is measured by work done factor(0.98-0.85)

It is a measure of the ratio of the actual work absorbing capacity of the stage to its ideal value as calculated from equation.

W = uca ( tan1 - tan2 )

This work done factor accounts for the effect of boundary layer and tip clearance.

In terms of temperature difference

hs = h03 - h01

CpTs = uca ( tan1 - tan2 )

Ts = uca ( tan1 - tan2 )

Cp

Compressor stage efficiency

It is the ratio b/w ideal work input to the actual work input.

Wideal = h03 h01

= Cp(T03 T01 )

Wactual = h03 h01

= Cp(T03 T01 )

c = (T03 T01 )

(T03 T01 )

Actual Stage work in terms of velocities and air angles

Wactual = h03 h01 = uca( tan2 - tan1 )

= uca ( tan1 - tan2 )

= 1/2(c22-c1

2)+1/2(w12-w2

2)

Performance coefficients

In order to evaluate the performance of the compressor same dimensionless performance coefficients are found useful in various analyses.

1.Flow coefficient

it is defined as the ratio of axial velocity to peripheral speed of the blades. Flow coefficients sometimes called as compressor velocity ratio.

2.Rotor pressure loss coefficient

it is defined as the ratio of the pressure loss in the rotor due to relative motion of air to the pressure equivalent of relative inlet velocity.

3.Rotor enthalpy loss coefficients

it is defined as the ratio of the difference between the actual and isentropic enthalpy to the enthalpy equivalent of the inlet relative velocity.

4.Stator/Diffuser pressure loss coefficient

it is defined as the ratio of the pressure loss in the diffuser

due to flow velocity to the pressure equivalent of actual inlet velocity

of the diffuser.

5.Stator/Diffuser enthalpy loss coefficient

it is defined as the ratio of the difference between the actual and

isentropic enthalpy to the enthalpy equivalent of absolute velocity of

flow at diffuser inlet

6.Loading coefficient

it is defined as the actual stagnation enthalpy rise in the stage to

enthalpy equivalent of peripheral speed of rotor.

Degree of reaction

The degree of reaction prescribes the

distribution of the stage pressure rise b/w the rotor and the stator blade rows.

for an actual compressor stage the degree of reaction is define as (R)

actual change of enthalpy in rotor

actual change of enthalpy in stage

LOW REACTION STAGE:(R ( P)d

Since the rotor blade rows have relatively higher efficiencies , it is advantageous to have a slightly greater pressure rise in them compared to diffuser.

Flow losses

Aerodynamic losses occurring in the most of the turbo machines arise due to the growth of boundary layer and its

separation on the blade and passage surface .

Types of aerodynamic losses

1.Profile loss 2.Tip clearance loss 3.Stage loss

Performance characteristics

The performance characteristics of axial flow compressors or their stages at various speeds can be presented in terms of the plots of the following parameters.

1.Presssure rise vs flow rate

2.Pressure ratio vs non-dimensional flow rate

0ff-design operation The performance of a compressor is defined according to its

design. But in actual practice, the operating point of the

compressor deviates from the design- point which is known as

off-design operation.

Unstable flow in axial compressors can be due to two reasons.

1.Seperation of flow from the blade surfaces called stalling.

2.Complete breakdown of steady through flow called surging.

Compressor surge It is a form of unstable operation and should be avoided.

Surge has been traditionally defined as the lower limit of stable operation in a compressor, and it involves the reversal of flow.

This reversal of flow occurs because of some kind of aerodynamic instability within the system.

Usually, a part of the compressor is the cause of the aerodynamic instability, although it is possible for the system arrangement to be capable of augmenting this instability.

A decrease in the mass flow rate, an increase in the rotational speed of the blade, or both can cause the compressor to surge.

One should note that operating at higher efficiency implies operation closer to surge.

Surge is a reversal of flow and is a complete breakdown of the continuous steady flow through the whole compressor. It results in mechanical damage to the compressor due to the large fluctuations of flow which results in changes in direction of the thrust forces on the rotor creating damage to the blades.

Compressor Stall

There are three distinct stall phenomena. Rotating stall and individual blade stall are aerodynamic phenomena; stall flutter is

an aero elastic phenomenon.

Individual Blade Stall

This type of stall occurs when all the blades around the

compressor annulus stall simultaneously without the occurrence

of a stall propagation mechanism.

The circumstances under which individual blade stall is

established are unknown at present.

It appears that the stalling of a blade row generally manifests

itself in some type of propagating stall and that individual blade

stall is an exception.

Rotating Stall

Rotating stall (propagating stall) consists of large stall zones covering several blade passages and propagates in the direction of the rotation and at some fraction of rotor speed. The number of stall zones and the propagating rates vary considerably .

This stalled blade does not produce a sufficient pressure rise to maintain the flow around it, and an effective flow blockage or a zone of reduced flow develops.

Stall Flutter

This phenomenon is caused by self-excitation of the blade and is an aero-elastic phenomenon. Stall flutter is a phenomenon that occurs due to the stalling of the flow around a blade.

Blade stall causes Karman vortices in the airfoil wake. Whenever the frequency of these vortices coincides with the natural frequency of the airfoil, flutter will occur. Stall flutter is a major cause of compressor blade failure.

Effects of stall

This reduces efficiency of the compressor

Forced vibrations in the blades due to passage through stall compartment.

These forced vibrations may match with the natural frequency of the blades causing resonance and hence failure of the blade.

Centrifugal compressors 1.It consists of a rotating element called impeller , diffuser and a

volute casing.

2.The air enters into the compressor through the suction eye of

the impeller. Due to the rotation of the impeller at a high speed

produces centripetal force which causes the air to move out of

the impeller at a high velocity.

3.Then the air with high velocity enters into a diffuser ring. The

diffuser blades of the diffuser ring are so shaped that these

provide an increased area of passage to the air which is passing

outwards due to which the velocity of air leaving the impeller is

reduced and its pressure is increased.

4.The high pressure air then flows to the divergent passage of

volute casing. The velocity of air is further reduced due to

increased cross sectional area of volute casing causing very small

rise in pressure.

5.From the casing the compressed air leads to exit pipe and finally

comes out of the compressor.

5.This type of compressor is a continuous flow machine suitable

for large flow rate at moderate pressure. The pressure ratios

between 4 to 6 may be obtained in this type of compressor.

Pressure ratio upto 12 can be obtained by multistage centrifugal

compressors.

Types of diffuser The diffuser consists of any annular space known as a vaneless

diffuser.

The diffuser consists of a set of guide vanes it is known as vanned diffuser . The main aim of this diffuser is to increase

the static pressure by reducing the kinetic energy.

Pressure rise across compressor

1

2 3

Inlet Casing

Impeller Diffuser

P

Channel

0

Ideal energy transfer Let us first considered the case of an ideal compressor with the following assumptions for radial vaned impeller.

1.Losses due to friction are negligible

2.Energy loss or gain due to heat transfer to or from the gas is considered very small.

3.The gas leaves the impeller with a tangential velocity equal to the impeller velocity , no slip condition is assumed.(ct2=u2)

4.The air enters the rotor directly from the atmosphere without tangential component.ct1= 0

Applying these assumptions to the Euler's energy equation under ideal conditions becomes.

E = ct2u2-ct1u1 (or)

E = {(c22 c1

2)+(u22 u1

2)+(w12 - w2

2)}

E = u22

This is the maximum energy transfer that is possible. therefore the work done by the impeller on unit quantity of air is given by

W = E = u22

Energy transfer equation from thermodynamic analysis

W = E = h02 - h01 = Cp(T02 T01 )= Cp T01(rc(-1/) -1)

u22 = Cp T01(rc

(-1/) -1)

Blade shapes and velocity triangles

In order to understand the actual energy transfer and flow through compressor we will use two velocity triangles.

1.Entry velocity triangles

2.Exit velocity triangles

The absolute and relative air angles at entry and exit of the impeller are denoted by 1, 2 and 1, 2.

Based on the value of 2 the blade shapes are given the name as forward curved blades (2>90),Radial blades (2=90),Backward curved blades(2

Types of impeller blade

The blades of the compressor or either forward curved or backward curved or radial. Backward curved blades were

used in the older compressors, whereas the modern centrifugal

compressors use mostly radial blades.

Since the change in radius between the entry and exit of the impeller is large the impeller velocities at these stations are

different.

u1 =21 /60

u2 =22 /60

Slip factor

It is the ratio b/w actual and ideal values of the whirl component at the exit of the impeller.

= ct2 ct2

Slip velocity Cs = ct2 - ct2

if the value of slip factor is 1 then the slip velocity is zero(no slip condition)

Performance parameters

Power input factor

In practice the actual energy transfer to the air from the impeller is lower than the ideal energy transfer ,because some energy is

lost in friction b/w the casing and the air carried round by vanes and in

disc friction.in order to take this into account power input factor is

introduced, so the actual energy transfer becomes.

E =Pif u22 Pif value lies b/w 1.035-1.04

Total head temperature rise across the compressor or temperature rise

across the impeller

Tc = T02 - T01 = Pif u22

Cp

Pressure coefficient :

p = Wactual Wisen

p = Cp T01(rc(-1/) -1)

Cp T01(rcm(-1/) -1)

p = Cp T01(rc(-1/) -1)

u22

Compressor efficiency :

It is the ratio b/w ideal enthalpy difference to the actual enthalpy difference.

c = (T02 T01 ) = Cp T01(rc(-1/) -1)

(T02 T01 ) Cp (T02 T01 )

c = p Pif

Combustion chambers

Combustion The combustion process is of critical importance in a gas turbine

cycle.

It is because in this process the chemical energy of the fuel is converted to heat energy which later converted into work by the turbine.

Therefore losses incurred in the combustion process will have direct effect on the thermal efficiency of the cycle.

Process of establishing self sustained fire using fuel and oxidizer of course in a controlled manner.it is basically a chemical process in which is fuel is burnt in presence of oxidizer.

The overall chemical process must be exothermic in nature , which liberates enough heat to sustain combustion process itself.

Commonly encountered combustion devices in our life are

candle flames , lightening of matchsticks , cigarette burning , wood burning , reciprocating engines, gas turbine engines ,rocket motor.

Combustion Combustion can be defined as a complex sequence of

chemical reactions b/w fuel and oxidizer accompanied by liberation of heat and light.

It is very important that fuel and oxidizer in right proportion within the flammable range must be mixed properly .

Besides this sufficient amount of ignition energy is required to initiate the process of combustion.

Hence the combustion process can be conceived as a triangle involving the fuel, oxidizer , ignition energy .

In order to study combustion phenomenon it is important to considered several disciplines such as thermodynamics , chemical kinetics, fluid mechanics , heat and mass transfer and turbulence .

What is fuel and oxidizer Chemically we can define oxidizer as one which accepts the

electrons .

In contrast the fuel can be defined as one which donate the electrons .

This property of elements ability to accept or donate electrons is known as electronegativity, which dictates whether an

element can be classified as fuel or an oxidizer.

Examples F = 4 , O = 3.5, K = 0.8

TYPES OF FUELS AND OXIDIZER

i)Gaseous Fuels LPG, Natural gas , biogas, Acetylene(Oxidizer-

air/O2)

ii)Liquid fuels Gasoline , HSD, Kerosene , Alcohols(oxidizer

air/liquid O2)

iii)Solid fuels- wood , coal , coke ,biomass , animal dung ,special

fuels Nitrocellulose(oxidizer air/O2)

COMBUSTION THEORY APPLIED TO GAS

TURBINE COMBUSTOR

In any combustion process obtaining complete reaction between fuel and air has a chemical aspect and a physical aspect.

Chemical aspect concerned with rate of reaction. physical aspects are concerned with particle size , injection mixing and evaporation.

The are three recognized postulations as to the combustion mechanism

1.Carbon preferential burning: which states that carbon in the

hydrocarbon fuels burns before the hydrogen.

2.Hydrogen preferential burning: which states that hydrogen

hydrocarbon fuels burns before the carbon.

3.Hydroxylation : which states that there is an initial uniting of oxygen

with the hydrocarbon to form a hydroxylated compound. through chain

reactions of molecules , atoms and radical, hydroxylated compound

burns to CO,CO2 and H20.

The modern theory is based on the statistics of probability as well as kinetics.

It is known from kinetic theory of gases that the individual molecules are in motion at some average velocity but with a wide

difference between the velocities of the slowest and fastest

molecules.

For the combustion reaction to take place the process requires the collision of molecules of fuel and oxygen.

The collision must have a sufficiently high energy level so that the molecules are broken down into atoms and radicals.

Since the temperature is a function of the molecular activity raising the temperature increases the probability and intensity of

collision of high velocity molecules. Therefore will be an increase

in the intensity of combustion.

Factors affecting combustion chamber design

The temperature level of the gases after the combustion must be comparatively low to suit the highly stressed turbine blade

materials

At the exit of The combustion chamber the temperature distribution must be of known form if a high turbine

performance is to be realized and the blades are not to suffer

from local over heating.

Combustion must be maintained in a stream of air moving with a high velocity in the region of 30-60m/s , and stable

operation is required over a wide range of air-fuel ratio from

full load to idling conditions . the air fuel ratio might be vary

from 60:1 to 120:1 for a simple gas turbine engines.

Cont. The formation of carbon deposits must be

avoided . small particles carried into turbines along with the high velocity gas stream can erode the blades.

In aircraft gas turbines combustion must be stable over a wide range of chamber pressure because this parameter changes with altitude and forward speed.

aircraft engine combustion chambers are normally constructed of light gauge heat resisting alloy sheet (approx. 0.8mm thick) but are only expected to have a life of some 10000 hours.

Requirements of the combustion chamber

Complete combustion of the fuel must be achieved .

The total pressure loss must be minimum.

Carbon deposits must not be formed under any expected condition of operation.

Ignition must be reliable and accomplished with easy over wide range of atmospheric conditions .

Temperature and velocity distribution at the turbine inlet must be controlled .

The volume and weight of the combustor must be kept within the reasonable limits.

Process of combustion

The process of combustion in a gas turbine combustion involves the following

1. The mixing of a fine spray of fuel droplets with air.

2. Vaporization of the droplets.

3. The breaking down of heavy hydrocarbons into lighter fractions.

4. The intimate mixing of molecules of these hydrocarbons with oxygen molecules.

5. The chemical reactions themselves.

Three stages of combustion chamber

1.About 15-20 per cent of the air is introduced around the jet of fuel

in the primary zone to provide the necessary high temperature for

rapid combustion.

2.Some 30 per cent of the total air is then introduced through holes

in the flame tube in the secondary zone to complete the

combustion. for high combustion efficiency this air must be injected

carefully at the right points to avoid chilling the flame locally and

drastically reducing the reaction rate .

3.In the tertiary or dilution zone the remaining air is mixed with

the products of combustion to cool them down to the temperature

required at the inlet to the turbine . sufficient turbulence must be

prompted so that the hot and cold streams are thoroughly mixed to

give the desired outlet temperature distribution , with no heat

streaks which would damage the turbine blades.

Cutout view of a can type combustion chamber

Combustion intensity

In aircraft gas turbine engines the air flow through the

engine is at high average speed (100 m/s), which requires

high combustion intensity (heat release rate per unit

volume per unit time).

Cont. .

The combustion intensities of some heat engine combustion processes are compared:

Boiler furnaces-----4x105to106 kJ/m3.hr (1x102 to 103kWatts/m3)

Piston engine -----25125x105 to 106kJ/m3.hr (7 35 x 102 to 103kWatts/m3)

Jet engine-----75-150x105 to 106kJ/m3.hr (2142 x102 to 103kWatts/m3)

Rocket engine -----260 x105 to106kJ/m3.hr(72 x 102 to 103kWatts/m3)

Cont.. The condition at the burner inlet is determined by

the outlet operating conditions of the compressor.

This may keep varying with varying flight regimes.

On the other hand, the outlet condition is governed by turbine design operating limits and is generally required to be uniform and stable.

Hence, combustion chamber is expected to be a stable source of hot gas.

That means even if its inlet conditions are variable it is expected to deliver comparatively steady and uniform flow to the turbine.

Types of combustion chamber

Can type

Annular type

Can annular type

Can type

Cannular Type

Annular type

The flame moves in the direction of the air flow inside the combustion chamber at a characteristic speed known as flame

speed. The flame is sustained in a flame zone at the end of

which most of fuel is burned. Outside the flame zone the

combusted gas moves towards the combustor exit.

The process of evaporation of droplets and mixing of fuel and air can occur partly aided by local turbulent vortices artificially

created around the spray zone, and partly by diffusion of liquid

vapour into air.

At the point of ignition all the droplets may not have been evaporated and mixed -hence some of them may burn as liquid

droplets in a surrounding air.

Injection and Evaporation

Flame Front

Mixing of Secondary air

Delivery of Uniform gas flow

Factors affecting Combustion chamber Performance

1.Pressure loss

a) pressure drop due to friction

b) acceleration due to heat addition

2.Combustion intensity heat release rate per unit volume per unit time .

3.Combustion efficiency it is the ratio between actual total head temperature rise to the theoretical total head temperature rise.

Practical problems

1.Flame tube cooling -

2.Fuel injection-meter the fuel flow , atomize the fuel.

3.Ignition-

4.Use of cheap fuels

5.Pollution-unburnt hydrocarbons,CO,NOx, oxides of sulphar