propulsion ii

PROPULSION II QUESTION BANK

UNIT – I

PART – A

1. What is need for using shrouds in axial flow turbines?

2. Differentiate between impulse and reaction turbine stage.

3. Draw the pressure and velocity variation of a single stage impulse turbine.

4. Draw the pressure and velocity variation of a single stage reaction turbine.

5. What is multistaging and why it is required in turbines?

6. Define blade loading (ψ) and flow coefficient (φ).

7. Write the relation between ψ and φ .

8. Explain blade and stage efficiency.

9. What is velocity compounding of multistage impulse turbine?

10. What is pressure compounding of multistage impulse turbine?

11. Draw h-s diagram for a flow through a turbine stage.

12. What is blade to gas speed ratio and its significance?

13. Define enthalpy loss coefficient, stagnation pressure loss coefficient in stator

blades.

14. What are blade cooling and its need?

15. What do you understand by enthalpy and pressure losses for moving blades?

16. Define total to total stage efficiency of an axial flow turbine.

17. What are limiting factors in gas turbine design?

18. When a stage does is said to be free vortex stage?

19. State when total to total efficiency is appropriate to use.

20. Define degree of reaction.

Part-B

1. Draw the velocity triangles of a single stage machine and show that

tanα2 + tanα3 = tanβ2 + tanβ3.

2. Derive an expression for work output also reduces it for tangential component at

exit is zero.

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3. Derive the relation for maximum utilization factor for a single impulse stage.

4. Show that work done after third stage in turbines is practically not of use.

5. Derive the relation for degree of reaction in terms of pressure, velocity, enthalpy

and flow geometry.

6. Explain briefly about zero degree of reaction, 50% degree of reaction, 100%

degree of reaction and negative degree of reaction.

7. Discuss about losses and efficiencies of turbine.

8. Explain briefly about blade cooling method.

9. A multi stage gas turbine is to be designed with impulse stages and is to be

operating with an inlet pressure and temperature of 6 bar and 900 K and outlet

pressure of 1 bar. The isentropic efficiency of the turbine is 85 %. All the stages

are to have a nozzle outlet angle of 750 and equal outlet and inlet blade angles.

Mean speed of blade is 250 m/s and equal inlet and outlet gas velocities. Estimate

maximum number of stages required. Assume Cp = 1.15 KJ/Kg-K, γ = 1.333 and

optimum blade speed ratio.

10. A gas turbine having single stage rotates at 10,000 rpm. At entry to the nozzle the

total head temperature and pressure of the gas is 700 0 C and 4.5 bar respectively

and at outlet from the nozzle the static pressure is 2.6 bar. At the turbine outlet

annulus the static pressure is 1. 5 bars. Mach number at outlet is limited to 0.5 and

gas leaves in an axial direction. The outlet nozzle angle is 700 to the axial

direction and the nozzle friction loss may be assumed to be 3 % of the isentropic

temperature drop from total head at entry to static condition at outlet nozzle

pressure. Calculate (i) The gas angle at the entry and outlet from the wheel

showing them on velocity diagram for mean blade section.(ii)Output power

developed by the turbine shaft. Assume the mean blades diameter as 64 cm , gas

mass flow rate as 22.5 Kg/s , turbine mechanical efficiency is 99 % , Cp =1.147

KJ/Kg-K and γ= 1.33.

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11. In a single stage impulse turbine the nozzle discharges the fluid on to the blades

at an angle of 650 to the axial direction and the fluid leaves the blades with an

absolute velocity 300 m/s at an angle of 300 to the axial direction. If the blades

have equal inlet and outlet angles and there is no axial thrust, estimate the blade

angle, power produced per kg / s of the fluid and the blade efficiency.

12. Gas at 7 bar and 3000 C expands to 3 bars in an impulse turbine stage. The nozzle

angle is 700 with reference to exit direction. The rotor blades have equal inlet and

outlet angle and the stage operate with the optimum blade speed ratio. Assuming

the isentropic efficiency of the nozzle is 0.9 and that the velocity at entry to the

stage is negligible. Deduce the blade angle used and the mass flow required for

this stage to produce 75 KW. Take Cp = 1.15 KJ/Kg-K

13. A single stage axial flow turbine has a mean radius of 30 cm and a blade height at

the stator inlet of 6cm. The hot gases enter the turbine stage at 1900 k pa and

1200 k the absolute velocity leaving the stator (C2) is 600 m/s and inclined at an

angle 65º to the axial direction. The relative angles at the inlet and outlet of the

rotor blade are 25º and 60º respectively. The stage efficiency is 0.88. Calculate (i)

β3 (ii) rotor rotational speed in rpm (iii) stage pressure ratio (iv) φ , ψ and degree

of reaction. (v) m (vi) Power delivered by the turbine. Take γ=1.33, R= 290 J/kg

and (Ca/U) is constant through the stage.

14. Discuss how pitch and chord for an axial turbine stage will be selected.

15. A mean diameter design of a turbine stage having equal inlet and outlet velocities

leads to the following data mass flow rate = 20kg/s, inlet temp To1 = 1000K, inlet

pressure Po1 = 4 bar, axial velocity (Constant through stage) Ca = 260 m/s, blade

speed U=360 m/s, nozzle efflux angle 2=65º stage exit swirl 3=10º. Determine

the rotor blade gas angles, degree of reaction, temperature drop co-efficient

(2Cp∆Tos/U²) and power output. Assuming a nozzle loss co-efficient λN = 0.05.

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Calculate the nozzle throat area required. Ignoring the effect of friction on the

critical conditions.

16. Air enters a two stage axial flow turbine at a total temp of 1400 K and total

pressure of 2230 k Pa. The actual work developed by each stage is 185 KJ/kg and

each stage has an adiabatic efficiency of 87% calculate.

(a) The total pressure at the exit from each stage.

(b) The overall adiabatic efficiency.

17. An axial flow turbine has following data:

Turbine stage inlet temperature = 1300 K, turbine nozzle outlet temperature =

940 K, turbine stage outlet temperature = 820 K, blade speed = 318m/s, axial-

velocity=184 m/s (constant). Turbine inlet velocity (absolute) is equal to turbine

outlet velocity (absolute). Calculate following (i) DOR, (ii) Blade loading

coefficient, (iii) Flow coefficient, (v) Work done. Assume suitable data if

required.

18. A single stage gas turbine operates at its design condition with an axial absolute

flow at entry and exit from the stage. The absolute flow angle at nozzle exit is 70º.

At stage entry the total pressure and temperature are 311 kPa and 850ºC

respectively. The exhaust static pressure is 100 kPa, the total to static efficiency is

0.87 and the mean blade speed is 500m/s. assuming constant axial velocity

through the stage. Determine (i) specific work done (ii) Mach No. leaving nozzle,

(iii) axial velocity, (iv)total to total efficiency, (v) stage reaction.

19. The gas turbine stage develops 3.36 MW for a mass flow rate of 27.2 kg/s. The

stagnation pressure and temperature entry of stage are 772 kPa and 1000K. The

axial velocity is constant throughout the stage. The gases entering and leaving the

stage without any absolute swirl. At nozzle exit the static pressure is 482 kPa and

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the flow direction is at 18º to the plane of wheel. Determine the axial velocity and

degree of reaction for the stage given that the entropy increase in the nozzles is

12.9 J/kg-k. Assume that the specific heat at constant pressure of the gas is 1.148

KJ/kg-k and R=0.287 KJ/kg-k. Determine also the total to total efficiency of the

stage given that the increase in entropy of gas across the rotor is 2.7 J/kg-k.

UNIT-II

PART-A

1. What are the applications of integral ram-rocket propulsion system?

2. Compare between subsonic and supersonic combustion chambers.

3. Mention the problems associated with sub critical mode of operation of ramjet.

4. Distinguish between ramjet and scramjet.

5. What are the advantages and disadvantages of integral ram-rocket?

6. Is it possible to use subsonic combustion chamber for hypersonic flight? Justify.

7. Mention advantages and disadvantages of ramjet engine.

8. How is a ramjet different from turbojet?

9. What are the possible applications of ramjet?

10. What is need for supersonic combustion?

11. Write down the function of isolator in scramjet engine.

12. How do you classify hypersonic ramjets based on combustion process?

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13. For a ramjet engine γ = 1.4, what Mach number for which maximum value of

TSFC occurs?

PART-B

1. Explain the reason for a ramjet propulsion system as a suit system for missile

propulsion. Why is it that it is not suitable for aircraft propulsion?

2. Sketch a typical ramjet and explain its operating principle. Sketch density,

velocity, pressure, temperature, and Mach number variation along the length of

the ramjet engine duct.

3. With neat sketches explain critical, sub critical, and super critical operation of

ramjet engine.

4. Compare the design features of ramjet combustor and a turbo jet combustor.

5. An ideal ramjet engine is being designed for a Mach number 3.2 aircraft at an

altitude of 33,000 ft. The fuel has a heating value of 43 MJ/Kg and the burner exit

total temperature is 1889 K. A thrust of 42 KN is needed. What is required

airflow? What is the resulting nozzle exit diameter? What is the resulting TSFC?

6. A supersonic combustion ramjet flies at a mach number of 8 where the ambient

air temperature 707 K. Heat is added in the combustor up to maximum

permissible limit. Find out the performance parameters of the engine.

7. For a ramjet the flight Mach number is 2.5 at a constant combustion chamber

temperature of 1945 K. The heat reaction QR of the fuel is 43 MJ/Kg and the inlet

area is 0.0929 m2. Determine the gross thrust, fuel to air ratio, specific impulse,

TSFC and overall efficiency of the ideal engine when it is operated at sea level.

8. A ramjet is to propel an aircraft at Mach number 1.5 at high altitude where

ambient pressure is 11.6 KPa and ambient temperature is 205 K. The maximum

temperature in the engine is 2500 K. The heating value of fuel used is 45 MJ/Kg.

Assume γ=1.4 and Cp=1.0 KJ/Kg-K. If all components are frictionless, determine

1. TSFC 2. Thermal efficiency 3.Propulsive efficiency 4.Overall efficiency.

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9. With the help of T-S diagram explain the working of a ramjet engine. Show on

the diagram the possible losses.

10. Discuss the merits and demerits of a scramjet.

11. A ramjet travels at M=2.8 at an altitude where the temperature is 256 K and the

pressure is 0.4 atm. Air flows through the engine at 40 kg/s with a burner exit

temperature of 2000 K. The fuel has a heating value of 46500 KJ/Kg. Estimate the

fuel- air ratio, thrust and TSFC, assuming γ=1.4.

12. A ramjet has a constant diameter combustion chamber followed by the nozzle

whose throat diameter is 0.94 times the chamber diameter. Air enters the

combustion chamber with total temperature =1450 K and M=0.3. How high may

the temperature rise in the combustion chamber with out necessarily changing the

chamber inlet conditions? For simplicity neglect frictional losses in the chamber

and nozzle. Assume that the working fluid has specific heat ratio γ=1.333.

13. An ideal ramjet is to fly at 20,000 ft at a yet to be determined Mach number. The

burner exit total temperature is to be 1777 K and the engine will use 65.77 Kg/s of

air. The heating value fuel is 43 MJ/Kg. At what Mach number will the TSFC to

be optimized? What is the optimum TSFC? What is the thrust and dimensionless

thrust at this condition?

14. Write short notes on

(i) Combined ejector ramjet

(ii) Separate ejector ramjet.

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UNIT-III

PART-A

1. Differentiate between air breathing and rocket engines.

2. How do you classify rockets?

3. What is effective jet velocity?

4. Define specific impulse, average specific impulse.

5. Define specific propellant consumption. What is its significance?

6. Derive relationship between CW and CF.

7. Define impulse to weight ratio.

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8. What do you mean by characteristic velocity and characteristic length?

9. Write the assumptions in vertical flight in rockets.

10. What are adapted nozzles?

11. Why some time actual mass flow rate is high than actual mass flow rate?

12. Define divergence factor for conical nozzle.

13. What are TVC and its importance?

14. What are the materials used throat sink?

15. Derive relationship for power of jet in terms of specific impulse for rocket propulsion.

Part – B

1. The inlet conditions for rocket nozzle are given by total temperature = 2800 K.

Total pressure 43 bar and the inlet velocity can be neglected when compared to

nozzle outlet velocity. The nozzle throat diameter is 5.2 cm. Estimate the mass flow

rate through the nozzle. Assuming optimum expansion at sea level, determine nozzle

exit velocity, exit mach no. and the thrust developed by the rocket unit. Take ratio of

specific heat as 1.35.

2. A rocket motor using Liquid H2 and Liquid O2 as the fuel and oxidizer has

combustion chamber temperature and pressure as 2700K and 25 atm respectively.

The rocket motor has throat area of 0.07m². The exit area is designed for a standard

pressure of altitude of 17 km (Pe = 8852 N/m²), γ may be taken as 1.26. Molecular

weight of combustion products to be 9.5. Given R=8314 J/kg-k mole. Determine

(i) Velocity at the exit (ii)Isp, (iii) Thrust developed (iv) Me (v) Ae (vi) mp

3. Determine the value of thrust co-efficient for a gas with γ=1.2 at pressure ratios

2.0 & 5.0.

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4. For a solid propellant rocket with a given data calculate (i) mass flow rate, (ii) exit

velocity, (iii) Characteristic velocity and (iv) Isp. Initial mass 1500 kg. Final mass

150 kg burn time 45 s, Average thrust 90 KN. Chamber pressure 6 MPa. Throat

diameter 50mm and nozzle exit diameter 250 mm.

5. A rocket operates at sea level with a chamber pressure of 150 atm and chamber

temperature of 2400ºK. The propellant consumption rate is 1.5 kg/s. assuming the

working fluid to have the properties of air, calculate the ideal thrust and ideal

specific impulse.

6. The following data are related to a rocket thrust chamber. The chamber pressure

and temperature are Po = 12 MPa and To=2950 K. The molecular weight of the hot

gas in the nozzle is 13.4; average specific heat ratio is 1.2. The ambient pressure is

97.5 kPa. The throat diameter of nozzle =150 mm. Determine (i)mass flow rate,

(ii)exhaust velocity, (iii)rocket thrust, (iv) characteristic velocity.

7. An ideal rocket operates at combustion chamber pressure of 3x10^6 N/m² and

expands to an ambient pressure of 3x10 N/m². The specific heat ratio of the

working substances 1.2 and the gas constant is 287 J/kg-k. If temperature of

working substance at the inlet of the nozzle is 2700K, determine the velocity at the

throat, ideal exhaust velocity and nozzle expansion ratio.

8. A rocket operates at sea level (P=0.1013 MPa) with a chamber pressure Po=2.068

MPa, a chamber temperature To = 2222K and a propellant consumption rate of mp

= 1 kg/s. If γ = 1.3 and R=345.7 J/kg-k. Calculate the ideal thrust and ideal specific

impulse.

Unit – IV

Part – A

1. Mention main components of solid propellant rockets.

2. How do you classify SPRs?

3. What is burning rate?

4. Write influence parameters of burning rate.

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5. How do you improve burning rate?

6. Write an empirical relation for burning rate.

7. What is importance of combustion index ‘n’?

8. Define regressive, neutral and progressive burning rate.

9. Define temp. Sensitivity of burning rate and pressure.

10. Define configuration, sliver and restricted surface.

11. What do you understand by web thickness, web fraction and volumetric loading?

12. What are the uses of additives?

13. What is the function of ignitor?

14. Give the reasons for misfire.

15. What are hypergolic propellants?

16. What is the function of an injector?

17. How combustion instabilities occur in LPRs?

18. Define frozen equilibrium.

19. What is shifting equilibrium?

20. What are the advantages of LPRs over SPRs?

21. What are cryogenic propellants?

22. State limitations of hybrid rockets.

23. When do you prefer turbo pump fed system over other systems?

Part – B

1. Discuss about Erosive burning and also explain about factors effecting erosive

burning.

2. Discuss briefly about performance parameters of SPRs.

3. Explain elaborately about characteristics of grain.

4. Write short note on:

i. End burning or cigarette burning grain.

ii. External burning rod

iii. Internal burning rod

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iv. Dog bone type

v. Wagon wheel type

5. Explain about star burning grain, rod and tube burning grain and Multi perforated

grain.

6. What is need for multi level thrust grain?

7. Write classification of solid propellants and explain their relative merits and

demerits.

8. Give some uses of additives.

9. Mention some oxidizers, fuel and binders of solid propellant with their advantages

and draw backs.

10. Discuss briefly about ignitors used in solid propellants.

11. Write short note on

(i) Inhibitors or restrictors.

(ii) Gas generators.

12. Write down the combustion mechanism of solid propellants.

13. Write properties, characteristics and performance for liquid propellants.

14. What are the common problems associated with liquid propellants and what are

the desired properties of liquid propellants.

15. Discuss about oxidizers and fuels of liquid propellant rocks. Also state what is

monopropellant and bipropellant?

16. Write short note on

(i) Gas pressure feed system

(ii) Turbo pump feed system

17. Explain about injector used in LPRs with their relative merits and demerits.

18. Write down the methods used for cooling of LPRs.

Unit – V

Part – A

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1. What is chemical equilibrium?

2. Name the types of tests performed before launching a rocket.

3. What is nozzle less propulsion system?

4. Why electrical rockets are called essentially power limited?

5. What is the basic concept in using advance propulsion technique?

6. What do you mean by resisto-jet?

7. What is the possible application of a solar sail?

8. What is major disadvantage of nuclear rockets?

9. Mention the advantages of using advanced propulsion technique.

Part – B

1. With a neat sketch explain the operation of nuclear rocket system. What are the

advantages and limitation of such system?

2. What are the limitations of an electrical rockets propulsion system? What are the

various types of electrical propulsion system? Explain one system with a neat

sketch.

3. What is water testing of liquid rocket injectors? How is it done? What is its

application?

4. Discuss the following advanced propulsion techniques with suitable sketches.

(i) Ion propulsion rocket.

5. Explain the working of a hybrid rocket with sketch.

6. Why are static tests conducted? List out parameters measured during static tests

and outline the principles of measurement.

7. With suitable sketches discuss the need and methods for cooling of rockets

engine thrust chamber.

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propulsion ii

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