PROPULSION-I Questions Bank 2014

Anna University, Chennai

 

16 MARKS QUESTION UNIT - I

1. An advanced fighter engine operating at Mach 0.8 and 10Km altitude where, Ta=223.297K & Pa=0.2649 bar has the following uninstalled performance data and uses a fuel with C.V=

42,800KJ/Kg:

Thrust = 50 KN Mass flow of air = 45Kg/s Mass flow of fuel = 2.65 Kg/s

Determine the specific thrust, thrust specific fuel consumption; exit velocity, thermal efficiency, propulsion efficiency, and overall efficiency (assume exit pressure equal to ambient pressure). 16

2. Find specific thrust and SFC of a simple turbojet engine, having the following component performance at which the cruise speed and altitude are M 0.8 and 10000m. Select ambient condition from the gas table.

Compressor pressure ratio Turbine inlet temperature Isentropic efficiency:	8.0 1200K
Of compressor ηc	0.87
Of turbine ηt	0.90
Of intake ηi	0.93
Of propelling nozzle ηj	0.95
Mechanical transmission efficiency ηm	0.99
Combustion efficiency ηb Combustion chamber pressure loss ∆Pb	0.98 4% of compressor outlet pressure.

C.V of fuel is 43,000 KJ/Kg, assume data if necessary, Cpa ≠ Cpg 16

3. (a) Explain with neat sketch operating principles of turbofan engine 8 (b) What is thrust augmentation? Explain any two methods of thrust augmentation with

sketches. 8

4. Compare the characteristics, advantages & disadvantages of turbojet, turbofan and turboprop engine.

5. (i)Discuss the different methods of thrust augmentation. Draw T-S diagram for turbojet engine with thrust augmentation. 8

(ii) Discuss the typical turbojet cycle performance with suitable sketches. 8

6. A turbojet engine is traveling at 270 m/s at an altitude of 5000m. The compressor pressure ratio is 8:1 and maximum cycle temperature is 1200K. By assuming the following data,

Ram efficiency 93% Isentropic efficiency of compressor 87%

Pressure loss in combustion chamber 4%of compressor delivery pressure

Calorific value of fuel 43,100 kj/kg Combustion efficiency 98% Mechanical transmission efficiency 99% Isentropic efficiency of turbine 90% Propelling nozzle efficiency 95% Ambient conditions at 5000 m are 0.5405 bar and 255.7 K.

Calculate the

(i) Specific thrust and

(ii) TSFC 16

7. (i) Define thrust of an engine and derive the thrust equation for a general propulsion system. 8

(ii) Discuss the typical turbojet cycle performance with suitable sketches. 8

8. An ideal turbojet flies at sea level at a Mach number of 0.75. It ingests 74.83 kg/s of air, and the compressor operates with a total pressure ratio of 15. The fuel has a heating value of 41,000 kj/kg, and the burner exit total temperature is 1389 K. Find the thrust developed and the TSFC. Assume that the specific heat ratio is 1.4. 16

9. Air enters a turbojet engine at a rate of 12*104 kg/h at 150C &1.03 bar and is compressed adiabatically to 1820C & four times the pressure. Products of combustion enter the turbine at

8150C & leave it at 6500C to enter the nozzle. Calculate the isentropic efficiency of the compressor, the power required to drive the compressor, the exit speed of gasses & thrust developed when flying at 800 km/h. Assume the isentropic efficiency of the turbine is same as that of the compressor and the nozzle efficiency is 90%.Assume the data required suitably. 16

10. A jet propelled plane consuming air at the rate of 18.2 kg/s is to fly at Mach number of 0.6 at an altitude of 4500m (Pa = 0.55 bar, Ta = 255K ). The diffuser which has a pressure coefficient of 0.9, decreases the flow to a negligible velocity. The compressor pressure ratio is 5 & maximum temperature in the combustion chamber is 1273 K. After expanding in the turbine, the gases continue to expand in the nozzle to a pressure of 0.69 bar. The isentropic efficiency of compressor, turbine and nozzle are 0.81, 0.85 & 0.915 respectively. The heating value of the fuel is 45870 kj/kg. Assume Cp = 1.005 kj/kg-K, Cpg = 1.147 kj/kg-K. Calculate

(i) Power input to the compressor

(ii) Power output of the turbine

(iii) The fuel air ratio

(iv) The thrust provided by the engine

(v) The thrust power developed. 16

1. (i) Explain successive steps in the acceleration and over speeding of a one- dimensional supersonic inlet with sketches. 8

(ii) Derive the relation between area ratio Amax/Ai and external deceleration ratio ui/ua. 8

2. A supersonic inlet is designed with a two-dimensional conical spike (with two half-cone angles 100 and 200 relative to the axial centerline, respectively). The inlet is to operate at a flight Mach number of 1.9.The two standing oblique shocks are attached to the spike and cowl, and a converging inlet section with a throat of area A* is used to decelerate the flow through internal compression. Assume γ = 1.4 and internal diffuser pressure recover factor Πr = 0.97. Estimate the overall recovery factor Πd on the assumption that the inlet starts (i.e., the normal shock is swallowed). Also, find the required A*/A1.

3. What are the different modes of inlet operation? Explain with suitable sketches. 16

4. Air enters a two-dimensional supersonic diffuser at a pressure of 14.102 kPa, a temperature of

217 K, and with a Mach number of 3.0. The two-dimensional oblique shock diffuser has an oblique shock angle of 27.80, which is followed by a normal shock. Determine, assuming constant specific heats.

(i) The velocity, total temperature and pressure of the air entering the oblique shock. (ii) The Mach number, total pressure after the oblique shock.

(iii) The flow deflection angle.

(iv) The Mach number, total and static pressure and static temperature after the normal shock.

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UNIT-III

1. (a) What are the important factors affecting combustor design? 8 (b)Write down the methods of flame stabilization and explain with sketch. 8

2. (a)What are the three types of combustion chamber? Compare its advantages and disadvantages. 8

(b) Name the material used for combustion chamber and discuss the special qualities of the material used for combustion chamber? 8

3. (a)What are the factors affecting combustion chamber? Explain briefly? 8

(b) With the aid of a simplified picture explain the operation of a flame holder. 8

4. (i) With a neat sketch explain the working of a combustion chamber. 8

(ii) Consider n-decane fuel, balance the chemical equation for the stoichiometric combustion of this fuel in air and find the stoichiometric fuel-to-air ratio. 8

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UNIT-IV

1. (a) Plot Mach number, static temperature, static pressure and static density variations along the longitudinal axis of a convergent-divergent nozzle, when it flows full. Explain the variations. 8

(b)A De Laval nozzle has to be designed for an exit Mach number of 1.5 with exit diameter of

200 mm. Find the ratio of throat area/exit area necessary. The reservoir conditions are given as Po = 106 Pa, To = 200C. Find also the maximum mass flow rate through the nozzle. What will be the exit pressure and temperature? 8

2. A converging-diverging is designed to operate with an exit Mach number of 1.75. The nozzle is supplied from an air reservoir at 68bar (abs.). Assuming 1-d flow, calculate:

(i) Maximum backpressure to choke the nozzle. 4 (ii) Range of backpressure over which a normal shock will appear in the nozzle. 4 (iii) Back pressure for the nozzle to be perfectly expanded to design M. 4 (iv) Range of back pressure for supersonic flow at the nozzle exit plane. 4

3. (i) What are the types of nozzle? Explain various operating conditions of a C-D nozzle with suitable sketch. 8

(ii) Write short notes on the following:

(a) Ejector and variable area nozzles 4 (b) Thrust reversing 4

1. An axial compressor stage has a mean diameter of 60cm and runs at 15000rpm. If the actual temperature rise and pressure ratio developed are 300C and 1.4 respectively.

(i) The power required to drive the compressor while delivering 57 Kg/s of air; assume mechanical efficiency of 86 % and an initial temperature of 350C.

(ii) The stage loading coefficient. (iii) The stage efficiency and

(iv) The degree of reaction if the temperature at the rotor exits is 550C.

2. (i) Explain the working of a centrifugal compressor and draw the velocity

triangles. 8

(ii) A centrifugal compressor has an impeller tip speed of 366 m/s. Determine the absolute Mach number of the flow leaving the radial vanes of the impeller when the radial component of velocity at impeller exit is 30.5 m/s and the slip factor is 0.9. Given that the flow area at impeller exit is 0.1m2 and the total-to-total efficiency of the impeller is 90%, determine the mass flow rate. 8

3. (i) A sixteen-stage axial flow compressor is to have a pressure ratio of 6.3. Tests have shown that a stage total-to-total efficiency of 0.9 can be obtained for each of the first six stages and 0.89 for each of the remaining ten stages. Assuming constant work done in each stage and similar stages fine the compressor overall total-to –total efficiency. For a mass flow rate of 40 kg/s determine the power required by the compressor. Assume an inlet total temperature of 288

K. 8

(ii) Discuss the factors affecting stage pressure rise of an axial flow compressor with suitable sketches. 8

4. A stage of a radial compressor is to be analyzed. It rotates at 12,300 rpm and compresses

31.75 kg/s of air. The inlet pressure and temperature are 241.325 kPa and 306K respectively. The hub and tip radii of the blades at the inlet are 7.62 and 13.97cm respectively. The exit radius is

27.94cm and the exit blade height is 2.54cm. The slip factor is unity. Flow enters the inducer with no prewhirl and the impeller has straight radial blades. The efficiency of the stage is 88%. The value of Cp and γ are 1.005 kj/kg-K and 1.397 respectively.

Find the following:

(i) Mean relative flow angle at the inlet.

(ii) The static pressure at the impeller exit. (iii) The total pressure ratio for the stage,

(iv) The Mach numbers at the impeller inlet and exit.

(v) The required power for the stage. 16

5. An axial flow compressor stage is designed to give free-vortex tangential velocity distributions for all radii before and after the rotor blade row. The tip diameter is constant and 1.0m; the hub

diameter is 0.9m and constant for the stage. at the rotor tip the flow angles are as follows: 16

Absolute inlet angle, α1 = 300

Relative inlet angle, β1 = 600

Absolute outlet angle, α2 = 600

Relative outlet angle, β2 = 300

.Determine,

(i) the axial velocity

(ii) the mass flow rate

(iii) the power absorbed by the stage

(iv) the flow angles at the hub

(v) the reaction ratio of the state at the hub

Given that the rotational speed of the rotor is 6000 rpm and the gas density is 1.5 kg/m3 which can be assumed constant for the stage. It can be further assumed that stagnation enthalpy and entropy are constant and after the rotor row.

6. The mass flow rate of flow at 288 K and 101.3 KPa at the inlet to the impeller of the centrifugal-flow compressor is 1.814 kg/s. The inlet flow is in the axial direction. The impeller eye has the minimum diameter of 3.81cm and a maximum diameter of 12.7cm and rotates at

35,000rpm. Assuming no blockage due to the blade, calculate the ideal angle at the hub and tip at the inlet to the impeller. Draw velocity diagram at the hub and at the tip

PROPULSION-I Two Marks Questions With Answers 2014

2 MARKS QUESTION WITH ANSWERS

UNIT-1

1. Define Froude efficiency, what is its effect on thrust?

From the propulsion power and the thrust power, we can derive the propulsive

efficiency _prop, also known as the Froude efficiency. It is given by

Propulsive efficiency = Pthrust/Pprop

2. Compare air breathing engine and rocket engine.

air breathing engine are combustion engines that use the oxygen in atmospheric air to oxidize ('burn') the fuel carried, rather than carrying an oxidizer, as in a rocket. Theoretically, this should result in a better specific impulse than for rocket engines.

A continuous stream of air flows through the Air-breathing engine. This air is compressed, mixed with fuel, ignited and expelled as the exhaust gas. Thrust produced by a typical air- breathing engine is about eight times greater than its weight. The maximum velocity of Air- breathing engines is limited to 1-3 km/s due to extreme temperature and dissociation of the exhaust gas; however, the maximum velocity of a hydrogen-breathing engine of the same design is about 4 times higher.

3. Define SFC.Write down its significance.

SFC, is an engineering term that is used to describe the fuelefficiency of an engine design

with respect to thrust output. It allows the efficiency of different sized engines to be directly compared.

SFC is dependent on engine design, but differences in the SFC between different engines using the same underlying technology tend to be quite small. Increasing overall pressure ratio on jet engines tends to decrease SFC.

4. What are the factors affecting thrust?

The Jet engine is much more sensitive to operating variables . Those are:

1.) Engine rpm.

2.) Size of nozzle area.

3.) Weight of fuel flow.

4.) Amount of air bled from the compressor

6. Define by pass ratio.

It is defined as the ratio between the mass flow rate of air drawn through a fan disk which

bypasses the engine core (un-combusted air), to the mass flow rate passing through the engine core which is involved in combustion to produce mechanical energy

8. Differentiate between Scramjet & Ramjet engine.

In a ramjet, the combustion chamber - where the air is mixed with fuel and ignited - only works

at subsonic speeds. So the intake slows the air down, releasing some of its energy as a shock wave, but this reduces fuel efficiency. The scramjet is an innovation on the ramjet in which the combustion chamber is specially designed to operate with supersonic airflow. Scramjet is simply an acronym for 'supersonic combustion' ramjet.

10. What are the advantages and disadvantages of cooling gas turbine blades?

There are three main types of cooling used in gas turbine blades; convection, film, and

transpiration cooling. While all three methods have their differences, they all work by using cooler air (often bleed from the compressor) to remove heat from the turbine blades.

11. Mention relative merits of jet engines over piston engines.

• Simple design with few moving parts, giving a compact engine with low maintenance costs and reduced frictional losses.

• The operational flexibility through the variable compression ratio allows operation optimisation for all operating conditions and multi-fuel operation. The free-piston engine is further well suited for homogeneouschargecompressionignition (HCCI) operation

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UNIT-II

1. What are the requirements of an aircraft intake?

The air intake requires enormous effort properly to control airflow to theengine.

The intake must be designed to provide the appropriate amount of airflow required by the engine

.Furthermore this flow when leaving the intake section to enter thecompressor should be uniform stable and of high quality.

Good air intake design is therefore a prerequisite if installed engineperformance is to come close to performance figures obtained at thestatic test bench

2. Write notes on pressure recovery factor of the intake?

The ratio between pressure of air intakes the engine to the pressure level of atmosphere

P01/Pa = P01/PaxP0a/Pa

3. What are the starting problems in supersonic inlets?

For air-breathing engines on supersonic vehicles,

usually want to slow flow down to subsonic speeds inside engine

– need diffuser (M>1→M<1) for engine inlet

– exception: supersonic combustion (e.g., SCRAM jets)

• Goal

– lowestpo loss (highest thrust)

• given flight M

• mass flow rate requirement (thrust)

• stable operation (nothing drastic for small changes in flight conditions)

4.What are the factors to be considered while designing a subsonic inlet?

1. High total pressure ratio

2. Good uniform of flow

3. Low installation drag

4. Good starting and stability.

5.What are the factors to be considered while designing a supersonic inlet?

1. Provide adequate subsonic performance

2. High pressure recovery factors

3. Good pressure distribution at the compressor inlet

6. What is meant by sub critical mode of inlet operation? State its advantages and disadvantages.

Some inlets are designed to operate in the subcritical mode with a margin of stable operation.

Stagnation

pressure recovery is normally equal to or less than the value for critical operation. The capture area ratio is less than one, since air is expelled from the inlet opening, with a resulting increase in drag due to the subsonic air-stream passing over the cowl lip.

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UNIT-III

1. Define equivalence ratio and stochiometric fuel air ratio.

A stoichiometric amount or stoichiometric ratio of a reagent is the optimum amount or ratio

where, assuming that the reaction proceeds to completion:

1. all reagent is consumed

2. there is no shortfall of reagent

3. no residues remain.

2. Define efficiency of the combustion.

Combustion efficiency is a calculation of how well your equipment is burning a specific fuel, shown in percent. Complete combustion efficiency would extract all the energy available in the fuel. However 100% combustion efficiency is not realistically achievable. Common combustion processes produce efficiencies from 10% to 95%. Combustion efficiency calculations assume complete fuel combustion and are based on three factors:

The chemistry of the fuel.

The net temperature of the stack gases.

The percentage of oxygen or CO2 by volume after combustion.

3.What is the purpose of primary air in combustion chamber?

The purpose of the primary air is as follows,

1. 30% of the air intakes the engine used in combustion process

2. 20% of air used in compression process

3. Remaining 50% of used primary used in cooling process

4.What is the purpose of secondary air in combustion chamber?

The purpose of the secondary air is as follows,

1. Used in external turbine cooling

2. Thrust reversal operation

3. Afterburner operation

5.What is the purpose of dilution air in combustion chamber?

Dilution air is also called as tertiary air, It is used in cooling the burning

surfaces.

6.Define combustion intensity?

The ratio between heat released by the combustor and the addition of combustion volume and

pressure

7.State the advantages and disadvantages of annular combustor.

i)The burner at the left is an annular combustor with the liner sitting inside the outer casing which has been peeled open in the drawing. Many modern burners have an annular design.

ii)The burner in the middle is an older can or tubular design. The photo at the top left shows some actual burner cans. Each can has both a liner and a casing, and the cans are arranged around the central shaft.

iii)A compromise design is shown at the right. This is a can-annular design, in which the casing is annular and the liner is can-shaped. The advantage to the can-annular design is that the individual cans are more easily designed, tested, and serviced.

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UNIT-IV

1. What is choked nozzle?

An expansion nozzle in which the gas velocity has reached sonic level at the

throat

2. What are the types of nozzle in gas turbine engine?

1. Convergent nozzle

2. CD nozzle

3. Give any four functions of an exhaust nozzle.

1. Allow for cooling of walls,

2. Allow for thrust reversing

3. Thrust vector should be controllable

4. Define under expanded nozzle.

An under expansion nozzle in which the gas exit pressure at the nozzle is greater than the

back pressure.

5. Define over expanded nozzle.

An over expansion nozzle in which the gas exit pressure at the nozzle is lesser than

the back pressure.

6. Define optimum expansion nozzle.

An optimum expansion nozzle in which the gas exit pressure at the nozzle is equal to

the back pressure

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UNIT-V

1. Write down the difference between centrifugal and axial flow compressor.

*Axial compressors are rotating, airfoil based compressors in which the working fluid

principally flows parallel to the axis of rotation. This is in contrast with centrifugal, axi- centrifugal and mixed-flow compressors where the air may enter axially but will have a significant radial component on exit.

*Centrifugal fan/blowers are more suited to continuous-duty applications such as ventilation fans, air movers, cooling units, and other uses that require high volume with little or no pressure increase. In contrast, multi-stage reciprocating compressors often achieve discharge

pressures of 8,000 to 10,000 psi (59 MPa to 69MPa). One example of an application of centrifugal compressors is their use in re-injecting natural gas back into oil fields to increase oil production

2. Define Elementary Airfoil Theory

When a single airfoil is parallel to the velocity of a flowing gas, the air flows over the

airfoil as shown in figure 7a. The airdivides around the body, separates at the leading edge, and joins again at the trailing edge of the body. The main stream itself suffers no permanent deflection from the presence of the airfoil. Forces are applied to the foil by the local distribution of the steam and the frictionof the fluid on the surface. If the airfoil is well designed, the flow is streamlined with little or no turbulence.

3. Define laminar flow airfoils.

Just before and during World War II, much attention was given to laminar-flow airfoils.

These airfoils are designed so that the lowest pressure on the surface occurs as far back as possible. The reason for this design is that the stability of the laminar boundary layer increases when the external flow is accelerated (in the flow with a pressure drop), and the stability decreases when the flow is directed against increasing pressure. A considerable reduction in skin friction is obtained by extending the laminar region in this way, provided that the surface is sufficiently smooth.

5. Define degree of reaction of axial compressor.

The degree of reaction in an axial-flow compressor is defined as the ratio of the change of static

head in the rotor to the head generating the stage

6. Define radial equilibirium.

The flow in an axial-flow compressor is defined by the continuity, momentum, and

energy equations. A complete solution to these equations is not possible because of the complexity of the flow in an axial-flow compressor. Considerable work has been done on the effects of radial flow in an axial-flow compressor. The first simplification used considers the flow axisymmetric. This simplification implies that the flow at each radial and axial station within the blade row can be represented by an average circumferential condition. Another simplification considers the radial component of the velocity as much smaller than the axial component velocity, so it can be neglected.

7. Define Diffusion Factor

The diffusion factor should be less than 0.4 for the rotor tip and less than 0.6 for the rotor

hub and the stator. The distribution of the diffusion factor throughout the compressor is not properly defined. However, the efficiency is less in the later stages due to distortions of the radial velocity distributions in the blade rows. Experimental results indicate that even though efficiency is less in the later stages, as long as the diffusion loading limits are not exceeded, the stage efficiencies remain relatively high.

Anna University, Chennai