TWO MARKS QUESTIONS UNIT I AIRCRAFT GAS TURBINES 1. What are the functions of the gas turbine? 1. 2.

Turbine is the part of the jet engine which is used to increase the kinetic energy of gases. It is used to operate the compressor.

2. What are the primary parts of the turbines? 1. The stator nozzles 2. The turbine rotor blades 3. What are the classifications of the turbines? 1. An impulse stage 2. A reaction stage 4. What is meant by impulse turbine stage? An impulse turbine stage is characterized by the expansion of the gasses which occurs only in the stator nozzles. The rotor blades act as a directional vanes to deflect the direction of the flow. Further they convert the kinetic energy of the gas into work changing the momentum of the gas more or less at constant pressure. 5. What is meant by reaction turbine? A reaction turbine is one which the expansion gas takes place both in the stator and in the rotor. 6. Define blade loading coefficient. The blade loading coefficient is defined as the ratio of work done to square of blade velocity. 7. Define degree of reaction. It is defined as the ratio of isentropic change of enthalpy in the rotor to isentropic change in enthalpy in the stage.

8. Define optimum speed ratio. The optimum speed ratio is the ratio of velocity of the blade to velocity of gas.

9. Write short notes on external cooling for turbine blades. The external surface of the gas turbine blade is cooled by making use of compressed air from the compressor. The quantity of the air required for the purpose is from 1 to 3% of main flow entering the turbine stage by which blade metal temperature can be reduced by about 200 300 degrees. 10. Explain internal cooling method adopted for gas turbine blades? Internal cooling of blades is achieved by passing air or liquid through internal cooling passages from hub towards the blade tip. The internal passages may be circular or elliptical. The cooling of the blades are achieved by conduction and convection. 11. What are the disadvantages of liquid cooling? 1. This system is complex. 2. Water is circulated at high pressure above its vapour pressure. 3. It is impossible to eliminate formation of deposits. 12. Write short notes on air cooling adopted in turbine blades. In this method, the air is bled from the high pressure end of the compressor and delivered to the blades and vanes to be cooled. Quantity of the coolant required to about 1 to 3 % of engine air flow per turbine blade row. 13. What are the assumptions made while eliminating the flow through the stage? 1. Flow conditions evaluated at the mean radius. 2. Blade height / mean radius is small, allowing two dimensional flow theory to be used. 3. Radial velocities are zero. 14. Define blade efficiency. It is defined as the ratio of work done per unit mass flow to work available per unit mass flow.

15. Define total to total efficiency.

Total to total efficiency is the ratio of actual work done by the gas to isentropic work done. 16. Define stage efficiency. The stage efficiency is defined as the ratio of work done in a stage to enthalpy drop in stage. 17. Define reaction ratio. Reaction ratio is defined as the ratio of static enthalpy drop across the rotor to static enthalpy drop across the stage. 18. What is the difference between impulse and reaction turbines? In impulse turbine, the fluid expanded completely in the nozzle and it remains at constant pressure during its passage through the moving blades. In reaction turbine, the fluid is only partially expanded in the nozzle and the remaining expansions take place in the rotor blades. 2. The fluid velocity and blade speed for the reaction turbine are low as compared with those of an impulse turbine. 19. What are the requirements should be satisfied when the gas turbine is to be used as an aircraft power plant? 1. Low weight 2. Small frontal area. 20. Define work ratio. It is the ratio of the actual total head temperature drop to the isentropic total head temperature drop from the total heat inlet to static head outlet. 1.

21. Define total to static efficiency. It is defined as the ration between the actual shaft work to the ideal shaft work between the total conditions at the entry and static conditions at exit.

UNIT II RAMJET PROPULSION

22. Define propulsive efficiency of ramjet engine.

The propulsive efficiency of ramjet engine is defined as the ratio of thrust power to power output. 23. Define combustion efficiency. The combustion efficiency is defined as the ratio of enthalpy rise of air to heat supplied to it. 24. Define diffuser efficiency. Diffuser efficiency is defined as the ratio of actual pressure rise to ideal pressure rise. 25. What are the advantages of ramjet engine? i) ii) iii) iv) v) Ramjet engine is very simple and does not have any moving part. Cost is low. Less maintenance. There is no upper limit for flight speed. Light weight when compared to turbojet engine.

26. What are the disadvantages of ramjet engine? i) Since the take off thrust is zero, it is not possible to start a ramjet engine without an external launching device. ii) The combustor required flame holder to stabilize the combustion due to high speed of air. iii) It has low thermal efficiency. iv) It is very difficult to design a diffuser which will give good pressure recovery over a wide range of speeds. 27. What are the applications of ramjet engine? i) It is widely used in high speed aircrafts and missiles due to its high thrust and high operational speed. ii) Subsonic ramjets are used in target weapons. 28. Explain critical inlet mode operation. When the inlet can accept the mass flow of air required to position the terminal shock just inside the cowl lip. This is called critical inlet operation. 29. What is subcritical operation?

When the inlet is not matched to the engine, the normal shock moves upstream. This is called as subcritical operation. 30. What is super critical operation? When the inlet can not capture the mass flow required by the engine, the terminal shock is sucked into the diffuser. This is called super critical operation.

31. What are the factors to be considered to select the fuel for ramjet engine? i) ii) iii) iv) v) vi) The calorific value of fuel The case with which it can be ignited Its physical properties Its storage ability Toxicity Corrosiveness

32. What are the factors affecting the combustion process? i) ii) iii) iv) The burner geometry Physical and chemical characteristics of fuel The air fuel ratio The velocity of working fluid

33. Briefly explain scramjet engine? A scramjet is a variant of a ramjet air breathing jet engine in which combustion takes place in supersonic airflow. 34. Define thrust. Thrust is a force which propels the engine in to the forward direction. Unit for thrust is Newton. 35. Define specific impulse. It is defined as the ratio of thrust to weight flow rate of air-fuel mixture. Unit for the specific impulse is seconds. 36. What is the function of nozzle? The function of the nozzle is to convert the pressure energy of the fluid into kinetic energy. 37. What is mach number?

Mach number is defined as the ratio of velocity of fluid to velocity of sound.

38. What are the burners used in the ramjet engine? 1. Can type burner 2. Baffle type burner

39. What is meant by ram effect? The function of the diffuser is to convert the kinetic energy of the entering air into pressure energy. This energy transformation is called ram effect. 40. What are the assumptions made for calculate the ideal efficiency of the ramjet engine? 1. Steady flow 2. One dimensional flow 3. Isentropic compression and expansion 4. Gas is perfect 5. Heat added at constant pressure 6. Very low Mach number in the combustion chamber.

41. Why ramjet engine does not require a compressor and a turbine? In ramjet engine, due to subsonic and supersonic diffuser, the static pressure of the air is increased to ignition pressure. So there is no need of compressor and turbine.

UNIT III FUNDAMENTALS OF ROCKET PROPULSION

42. State the comparison between jet engines and rocket engines. In the jet engines, oxygen obtained from the surrounding atmosphere for combustion process. In the rocket engines, the propulsion unit consists of own oxygen supply for combustion purpose. 43. What are the classifications of rocket engine? Rocket engine classified as follows: i) Chemical rocket engines ii) Nuclear rocket engines iii) Electrical rocket engines

iv)Solar rocket engines 44. What is under expanded nozzle? It is a nozzle which discharges fluid at exit pressure greater than external pressure, because the exit area is too small.

45. What is over expanded nozzle? It is a nozzle which discharges fluid at exit pressure lower than external pressure, because the exit area is too large.

46. What are the advantages of conical nozzle? i) ii) It has simple configuration It is relatively easy to fabricate.

47. Define effective speed ratio. It is the ratio of speed of flight to velocity of jet. 48. In rocket engine, how the propulsive efficiency varies with respect to speed ratio? In rocket engine, if the speed ratio is increased, the propulsive efficiency is also increased and reaches maximum value when the speed ratio is unity. After that, the propulsive efficiency is decreased with increase in speed ratio. 49. What is weight flow ratio? It is the ratio of propellant flow rate to the throat force. 50. Define IWR? The ratio of total impulse of the rocket to total weight of the rocket is called as impulse weight ratio.

51. Explain the performance of the rocket engine. In rocket engine, if the speed is increased, the propulsive efficiency is increased and reaches the maximum value of one. Then propulsive efficiency is decreased with increase in speed ratio. 52. What is thrust coefficient? It is the of the thrust to the throat force. 53. Define specific propellant consumption.

The propellant consumption rate per thrust is called as specific propellant consumption. 54. Define altitude. The height of the rocket engine from the sea level is called as altitude.

55. What is the advantage of bell nozzle over conical nozzle? The bell nozzle has 20 % less than the length that would be required for a conical nozzle. 56. What is internal ballistics? The rocket motors operation and design depend on the combustion characteristics of propellant, its burning rate, burning surface, and grain geometry. The branch of science describing these is known as internal ballistics. 57. Define characteristic velocity. It is the ratio of jet velocity to thrust coefficient. 58. Define overall efficiency. Overall efficiency is defined as the ratio of thrust power to heat supplied by the propellant. 59. Define thermal efficiency. Thermal efficiency is the ratio of power developed by the engine to heat supplied by the propellant. 60. Define effective jet velocity. The effective jet velocity is the ratio of thrust to mass flow rate of propellant. 61. Write short notes on aero spike nozzle. The aero spike nozzle has common plug at its centre. The small combustion chambers arranged in a circle around a common plug. The gasses coming out from the chambers flow through the divergent section of the nozzle where they are expanded. UNIT IV CHEMICAL ROCKETS 62. Define heterogenous propellants.

In heterogenous propellants solid propellants plastics, polymers and polyvinylchloride are used as fuels. Nitrates an perchlorates are used as oxidizer. 63. Define homogenous propellants. In homogenous propellants solid propellants nitroglycerine and nitrocellulose are used . It combines the properties of fuels and oxidizer. 64. Define burning rate. The velocity at which a solid propellant is consumed during operation is called the burning rate 65. What is monopropellant? A liquid propellant which contains both the fuel and oxidizer in a single chemical is known as monopropellant. 66. What is bipropellant? If the fuel and oxidizer are different from each other in its chemical nature then the propellant is called bipropellant. 67. Classify the rocket engine based on source of energy employed? Rocket engine can be classified as 1. chemical rocket engine 2. solar rocket engine 3. nuclear rocket engine 4. electrical rocket engine. 68. What are the factors increase the burning rate? 1. Combustion chamber pressure 2. Initial temperature of the solid propellant prior to operation 3. Combustion gas temperature 4. Velocity of gas flow parallel to the burning surface 5. Motor motion 69. What are the components of liquid propellant rocket engine? 1. Tanks for storing liquid fuel and oxygen 2. Preheater 3. Combustion chamber and nozzle 70. What the conditions of maximum propulsive efficiency? The jet velocity must be twice more than the free stream velocity for which the propulsive efficiency is 66.7%. 71. What are the disadvantages of liquid propellant rocket engine? 1. Manufacturing cost is high.

2. High vibration 3. The size and weight of propellant rocket.

the engine is more compared to solid

72. What are the types of propellant feed system? 1. Gas pressure feed system 2. Pump feed system

73. What are the basic combustion processes? 1. 2. 3. 4. 5. Injection Atomization Mixing Ignition Chemical reaction between fuel and oxidizer.

74. What are the advantages of solid propellant rocket engine? 1. Simple in design and construction 2. Less vibration due to absence of moving parts 3. Less maintenance 75. What are the disadvantages of solid propellant rocket engine? 1. It is difficult to stop the engine 2. Low specific impulse 3. Decrease of speed is not possible 76. What is the limitation of hybrid rocket engine? In the hybrid rocket engine, the nozzle erosion can not be avoided. 77. What are the advantages of hybrid rocket engine? 1. 2. 3. 4. Speed regulation is possible by regulating the supply of oxidizer High load capacity High fuel density Lighter compared to liquid propellant rockets

78. What is the use of strand burner? Strand burner is used to measure the burning rate of the solid propellant.

79. What is cold gas propellant? A cold gas propellant is stored at very high pressure gives a low performance allows a simple system and is usually very reliable.It has been used for roll control and altitude control. 80. What is gelled propellant? A gelled propellant is a thixotropic liquid with a gelling additive. It behave like a jelly or thick paint. It will not spill or leak. Readily can burn flow under pressure will burn and is safer in some respects. 81. Define Mixture ratio. The propellant mixture ratio for a bipropellant is the ratio at which the oxidizer and fuel are mixed and react to give hot gases.

UNIT V ADVANCED PROPULSION TECHNIQUES

82. What are the methods for ion generation? There are three methods for ion generation. They are 1. Surface contact 2. Electron bombardment and 3. Electric arc. 83. What are the advantages of electrical propulsion system? 1. Simple device and easy to control 2. Simple power containing 3. Low cost and relatively high thrust and efficiency 4. Can use many propellants including hydrazine augmentation. 84. What are the disadvantages of electrical propulsion system? 1. Lowest Isp, , heat loss 2. Dissociation of gas 3. Indirect of heating of gas and erosion. 85. What are the advantages of arc jet propulsion system? 1. Direct heating of gas. 2. Low voltage 3. Relatively simple device and high thrust. 4. Can use catalytic hydrazine augmentation inert propellant. 86. What are the disadvantages of arc jet propulsion system?

1. 2. 3. 4.

Low efficiency Erosion at high power and low specific impulse High current, heavy wiring, heat loss More complex power conditioning.

87. What are the advantages of ion jet propulsion system? 1. High specific impulse 2. High efficiency 3. Inert propellant. 88. What are the disadvantages of ion jet propulsion system? 1. Complex power conditioning and heavy power supply 2. High voltage, single propellant only 3. Low thrust per unit area. 89. What are the advantages of pulsed plasma electrical propulsion system? 1. Simple device and low power. 2. Because of solid propellants, no need of gas or liquid feed system and there is no zero gravity effects on propellants. 90. What are the disadvantages of pulsed plasma electrical propulsion system? 1. Low thrust 2. Teflon reaction products are toxic 3. Corrosive 91. What are the advantages of steady state plasma electromagnetic propulsion system? 1. Can be relatively simple 2. High specific impulse 3. High thrust per unit area 92. What are the disadvantages of steady state plasma electromagnetic propulsion system? 1. Difficult to stimulate analytically 2. High specific power 3. Heavy power supply. 93. What are the advantages of Hall thruster? 1. Desirable Isp range 2. Compact relatively simple power conditioning 3. Inert propellant 94. What are the disadvantages of Hall thruster? 1. Single propellant 2. High beam divergence

3. Erosion. 95. What are the requirements for solar sail powered spacecraft? 1. Continuous force exerted by sunlight 2. A large ultrathin mirror 3. A separate launch vehicle. 96. What are the types of electrical rocket engines? 1. Arc plasma rocket engine 2. Ion rocket engine 3. Magneto- plasma rocket engine. 97. What are the components in Arc plasma rocket engine? 1. Propellant tank 2. Combustion chamber 3. Cooling system 4. Electric power supply. 98. What are the components in Magneto- plasma rocket engine? 1. Propellant tank 2. Propellant pump 3. Thrust chamber 4. Accelerator. 99. What are the components in Ion rocket engine? 1. Propellant tank 2. Thrust chamber 3. Electric power supply 4. Vapourizing chamber. 100. Write down three fuel oxidizer combination for hybrid propellant rockets? Beryllium hydride Fluorine Lithium hydride chlorine trifluoride Hydrocarbon Nitrogen tetroxide Lithium hydride Nitrogen tetroxide 101. What are the advantages of hybrid propellant rockets engine? 1. Speed regulation is possible by regulating the supply of oxidizer. 2. High load capacity. 3. Hybrid rockets are lighter when compared to the liquid propellant type rocket. 4. Higher fuel density.

SIX MARKS QUESTIONS UNIT I AIRCRAFT GAS TURBINES

1. Describe the working of axial flow turbine stage with neat sketch.

The axial flow turbine stage consists of nozzle and rotor blades. Initially gas enters the row nozzle blades where it is expanded. Then the gas enters the rotor blade passages where it is further expanded. During the expansion process, pressure increased from P1 to P2 and the

temperature is increased from T1 to T2. Here, the pressure energy of the gas is converted into kinetic energy. Finally, the gas is ejected in to the atmosphere.

2. Describe the working of radial flow turbine stage with neat sketch.

The radial flow turbine stage consists of volute, diffuser, nozzle vanes. Here, the gas flow with a high tangential velocity is directed inwards. Then the gas leaves the rotor with as small as whirl velocity as practicable

near the axis of rotation.

3. Discuss the limiting factors in turbine design. Centrifugal stresses in the blades are proportional to the square of the rotational speed and the annulus area. Gas bending stresses inversely proportional to number of blades and blade section modulus. Gas bending stresses directly proportional to blade height and blade work output. Optimizing the design. The velocity triangle upon which the rotor blade section depends are partially determined.

4.What are the cooling methods adopted in turbine blades? Explain. External cooling Internal cooling Liquid cooling Air cooling

5. What are the factors considered to select blade profile, pitch and chord? Optimum pitch / chord ratio Aspect ratio Rotor blade stresses Effect of pitch on the blade root fixing

UNIT II RAMJET PROPULSION 1. Explain the performance of ramjet engine? If the altitude increases the net thrust will be decreased. In the sea level the net thrust will be maximum.

For cruise thrust, net specific fuel consumption decreases with increase in mach number. The ramjet engines have highest thrust per unit weight amongst air breathing engines. Ramjet can operate at subsonic velocities just below the sonic velocity.

2. Derive ideal efficiency for Ramjet engine. Define ideal efficiency. Initially derive ideal efficiency = 1-1/t. Using the relation between static and stagnation temperature derive the ideal efficiency of the ramjet engine.

3. Explain baffle type burner with neat sketch. It consists of fuel injecting system, igniter, flame holder and combustion zone. The air is supplied to the combustion zone with very low velocity. Fuel also supplied to the combustion zone. Igniter initiates the combustion of air-fuel mixture. Flame holder is used to stabilize the combustion process.

4. Explain the process involved in supersonic combustion ramjet engine. A supersonic combustion ramjet is a variant of a ramjet air breathing jet engine in which combustion takes place in supersonic air flow. Here, the air enters the inlet at Mach number 6. To avoid the shock wave problem, partially compress and decelerate the incoming flow. The supersonic combustion ramjet engine (scram jet) consists of converging inlet, combustor and diverging nozzle. In the combustor, the gaseous fuel is burned with atmospheric oxygen to produce heated air.

5. Explain typical modes of inlet operation? critical inlet mode operation.

When the inlet can accept the mass flow of air required to position the terminal shock just inside the cowl lip. This is called critical inlet operation. subcritical operation

When the inlet is not matched to the engine, the normal shock moves upstream. This is called as subcritical operation. super critical operation

When the inlet can not capture the mass flow required by the engine, the terminal shock is sucked into the diffuser. This is called super critical operation.

UNIT III FUNDAMENTALS OF ROCKET PROPULSION 1. A rocket flies at a speed of 10,000 kMph with an effective exhaust velocity of 1350 m/s and the heat produced by the propellant is 6600 KJ / kg. If the propellant flow rate is 4.8 kg/s, determine, (i) Propulsive efficiency (ii) Propulsive power (iii) Engine output (iv) Thermal efficiency (v) Overall efficiency Solution: efficiencies. Find out propulsive efficiency from speed ratio. Find out propulsive power from thrust. Find out Engine output from propulsive efficiency. Find out thermal efficiency from engine output. Find out overall efficiency from propulsive and thermal

2. Derive propulsive efficiency for rocket engine. It is defined as the ratio of thrust power to engine power developed. Find thrust power. It is the product of thrust and velocity of flight. Find engine power developed = Thrust power + Power loss. Find propulsive efficiency from thrust power and engine power developed.

3. A rocket is to be designed to produce 5 MN thrust at sea level. The pressure in the combustion chamber is 7 MPa and the temperature is 2800 K. If the working fluid is assumed to be a perfect gas with the properties of air at room temperature, determine the following: (i) Specific impulse (ii) Mass flow rate (iii) Throat diameter (iv) Exit diameter Solution: Find out exit mach number and exit temperature from gas table. Find out velocity of sound at exit from exit temperature. Find exit velocity from exit mach number. Find thrust from exit velocity. Find mass flow rate and specific impulse from thrust. Find throat diameter from mass flow rate. Find exit diameter from throat diameter.

4. Explain conical nozzle with neat sketch. energy. nozzles. In convergent nozzle, the cross sectional area is gradually decreased Conical nozzle is classified into convergent, divergent and CD Function of the nozzle is to convert pressure energy into kinetic

from entry to exit. In divergent nozzle, the cross sectional area is gradually increased

from entry to exit. In CD nozzle, two truncated cones joined top to top along their axis by a suitable radius to form a nozzle throat. The conical apex angle is 15 O 5. Explain bell nozzle with suitable sketch.

 This nozzle used in Atlas sustainer engine. It consists of throat and divergent section. The length of divergent section of bell nozzle is less than that of conical nozzle. To compensate that, the divergent angle at the throat is provided at 40 degrees. The expansion ratio is same for bell and conical nozzles. The bell shaped nozzle has a length 20 % less than that of 15 degrees conical nozzle.

UNIT IV CHEMICAL ROCKETS

1. With the help of a schematic diagram, explain elaborate scheme of a typical turbo pump feed system. Schematic diagram of turbo pump Fuel and oxidizer tank Parts of turbo pump Working principle 2. Explain the hardware components of solid propellant rocket engine with neat sketch. and explain its working principle. Combustion chamber and nozzle Provision for assembly and disassembly of the unit Mounting pads Provisions for preventing over pressurization of the chamber Provisions for holding the propellant grain in place Seals for preventing moisture to reach the grain Devices for changing the vector direction of thrust. 3. What are the advantages of liquid propellant system over solid propellant system? Combustion process is controllable High specific impulse Speed regulation is possible Smoke produced by the propellant is low.

4. List and very briefly explain the methods of cooling adopted for rocket motors. 1. Steady state cooling i. Regenerative cooling ii. Radiation cooling 2. Unsteady state cooling i. ii. Film cooling Sweat cooling

5. Explain strand burner with diagram Strand burner schematic diagram Parts of strand burner Vent gas Chronometer working UNIT V ADVANCED PROPULSION TECHNIQUES

1. Explain briefly about propellant grain design consideration? Grain design Maximum overall length Burning area program Physical strength characteristics 2. Explain combustion process in rocket engine? Combustion process Types of injectors Parallel stream type Impinging stream type Spray injection type

3. Explain briefly about hybrid propellant rocket engine with diagram? Schematic diagram of hybrid propellant rocket engine Parts of hybrid propellant rocket engine Working principle

 Advantages and disadvantages. 4. Explain briefly about arc plasma rocket engine? Schematic diagram of arc plasma rocket engine Parts of arc plasma rocket engine Working principle Plasma gas Propellants 5. Explain briefly about Ion rocket engine? Schematic diagram of Ion rocket engine Parts of Ion rocket engine Working principle Accelerating grid

TEN MARKS QUESTIONS

UNIT I AIRCRAFT GAS TURBINES

1. In a single stage impulse turbine the nozzle discharge the fluid on to the blades at an angle of 65 degrees to the axial direction and the fluid leaves the blades with an absolute velocity of 300 m/s at an angle of 30 degrees 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 blade efficiency. Solution: Draw velocity triangle for turbine. Find absolute velocity at exit using velocity triangle. Find blade angle from the absolute velocity at exit. Find power produced using tangential velocity, mass flow rate and blade speed. Find effective speed ratio from blade speed and absolute velocity at exit. Find blade efficiency from effective speed ratio.

2. A multi stage turbine is to be designed with impulse stages, and is to operate with an inlet pressure and temperature of 6 bar and 900 K and an outlet pressure of 1 bar. The isentropic efficiency of the turbine is 85 %. All the stages are to have a nozzle outlet angle of 75 degrees and equal outlet and inlet blade angles. Mean

blade speed of 250 m/s and equal inlet and outlet gas velocities. Estimate the maximum number of stages required. Solution: Find actual overall temperature drop from turbine efficiency. Applying energy equation to the nozzle of any stage. Find absolute velocity of gas at exit from optimum speed ratio. Find temperature drop in a stage from energy equation. Find number of stages required from actual overall temperature drop and temperature drop in a stage.

3. Explain procedure for matching the turbine and compressure? Select operational speed Assume turbine inlet temperature Assume compressor pressure ratio Calculate compressor work per unit mass Calculate turbine pressure ratio Check if compressor mass flow plus fuel flow equals the turbine mass flow Calculate pressure ratio across the jet nozzle Calculate area of the jet nozzle outlet. If the calculated area does not equal the actual exit area, assume new value of turbine inlet temperature and repeat the procedure.

4. An axial flow turbine has a blade speed at the mean diameter is 300 mps, and mass flow is 2.5 kg/s. The gas temperature at turbine inlet and outlet are 500 degrees and 300 degrees respectively. The fixed blade outlet angle is 20 degrees measured in the same direction of U. The axial velocity remains constant at 200 mps. Determine power developed. Solution: Draw velocity triangle for turbine From the velocity triangle, find relative velocity at exit and absolute velocity at exit. Find temperature difference from gas temperature at inlet and outlet. Find work done from temperature difference. Find power developed from mass flow rate.

5. The blades of free vortex turbine rotor have inlet and outlet angled of 60 degrees and 65 degrees at a mean diameter of 100 cm. The corresponding nozzle angle is 70 degrees. The hub tip ratio is 0.6 and the turbine runs at 3600 rpm. Calculate for the hub, mean and tip sections (a) Blade angles (b) Degree of reaction (c) Blade to gas speed ratio Solution:

Draw the velocity triangle for turbine Find velocity at tip from radius of tip Find velocity at hub from radius of hub Find blade angles from velocity triangle Find work done from the mean velocity Find degree of reaction from work done Find blade to gas ratio from work done.

UNIT II RAMJET PROPULSION 1. Explain can type of burner with neat sketch.

It consist of liner, injector and combustion zone. Multiple chambers are placed after the subsonic diffuser. Each chambers are supplied with separate system of air and it having own fuel jet from common fuel supply line. For maintenance, the individual chamber can be removed and replaced without disturbing the operation. 2. Discuss the problems associated with the supersonic combustion process? Control of flow is difficult Keep combustion rate of fuel constant Fuel injection and management is complex Choking problem Special cooling is required Testing difficulties. 3. A ramjet engine operates at M= 1.2 at an altitude of 6500 m. The diameter of the inlet diffuser at entry is 50 cm and stagnation temperature at the nozzle entry is 15oo K. The C.V of the fuel used is40 MJ/kg. The properties of the combustion gases are same as that of the air (=1.4, R=287 J/kg K). The velocity of air at the diffuser exit is negligible. Calculate the The efficiency of ideal cycle Flight speed Air flow rate Diffuser pressure ratio Fuel air ratio

i) ii) iii) iv) v)

vi)

Nozzle jet Mach number

The efficiency of diffuser is 0.9, combustor = 0.98, nozzle = 0.96 Given data Entry Mach no, M1 = 1.2 Z = 6500 m Di =50 cm = 0.5 m T03 = 1500 K C.V = 40 MJ/kg = 40 106 J/kg M2 =0 (combustor entry velocity is negligible) = 1.4 R = 287 J/kg K i) Efficiency of ideal cycle = = 0.2236 Efficiency of ideal cycle = 22.36 % ii) flight speed, u M1 = At z= 6500 m from data book a1 = 314.5 m/s 1 = 0.624 kg/m3 u = M1 a1 u = 377.2 m/s iii) mass flow rate m.a = 1 A 1u = 44.7 kg/s diffuser pressure ratio , Rod Diffuser efficiency= Rod = 2.24 v) fuel air ratio f= combustion efficiency= =1+ M12

iv)

T01 = T02 = 316.72 K vi) f = 0.03 exit Mach number . M4

M4 = W k t T04 = T4 + At nozzle expansion occurs at constant stagnation temperature , i.e., T03 = T04 Nozzle efficiency = To find T4 Inn combustor combustion occurs at constant stagnation pressure i.e., P02 = P03 For complete expansion of nozzle, P1 = P4 = 0.44 105 N/m2 P4/ P03 = 0.446 From gas tables, for P4/ P03 = 0.446 and = 1.4 T4/ T03 = 0.794 T4 = 1191 K Substituting in nozzle efficiency T4 = 1203.36 K T04 = T4 + C4 = 772.2 m/s M4 = M4 = 1.11

4. A ramjet engine propels on an aircraft at Mach no = 1.4 and at an altitude of 6ooo m. the diameter of inlet diffuser at the entry is 40 cm and C.V of fuel is 43 MJ/kg. The stagnation temperature at the nozzle entry is 1500 K. The properties of combustion gases are same as those of air(=1.4 , R=287 J/kg K). Find i) Efficiency of ideal cycle ii) Flight speed iii) Air flow rate iv) Diffuser pressure ratio v) Fuel air ratio vi) Nozzle pressure ratio vii) Nozzle jet Mach number viii) Propulsive efficiency ix) Thrust Efficiency of diffuser is 0.92 , combustor is 0.97 , nozzle is 0.95 and combustor pressure loss is 0.22P02. Given data Entry Mach no, M1 = 1.4 Z = 6000 m

Di =40 cm = 0.4 m T03 = 1500 K C.V = 43 MJ/kg = 43 106 J/kg M2 =0 ( combustor entry velocity is negligible) P0cc = 0.22P02. = 1.4 R = 287 J/kg K i) Efficiency of ideal cycle = = 0.2816 Efficiency of ideal cycle = 28.16 % flight speed, u M1 = At z= 6500 m from data book a1 = 314.5 m/s 1 = 0.624 kg/m3 P1 = 0.472 105 N/m2 u = M1 a1 u = 443.1 m/s iii) mass flow rate m.a = 1 A 1u = 36.73 kg/s diffuser pressure ratio , Rod Diffuser efficiency = Rod = 2.94 v) fuel air ratio f= combustion efficiency= =1+ M12

ii)

iv)

T01 = T02 = 346.8 K Mf = 0.992 f = 0.027 vi) nozzle pressure ratio , Ron wkt T03 = T3

=(

)^(

P03 = P3 RON = = For complete expansion of nozzle, P1 = P4 = 0.472 105 N/m2 Given P0cc = 0.22P02. P02 = 1.38 105 N/m2 P03 = 1.36105 N/m2 RON = = 2.88 vii) exit Mach number . M4 M4 = W k t T04 = T4 + At nozzle expansion occurs at constant stagnation temperature , i.e., T03 = T04 Nozzle efficiency = To find T4 Inn combustor combustion occurs at constant stagnation pressure i.e., P02 = P03 For complete expansion of nozzle, P1 = P4 = 0.472 105 N/m2 P4/ P03 = 0.347 From gas tables, for P4/ P03 = 0.34 and = 1.4 T4/ T03 = 0.736 T4 = 1104 K Substituting in nozzle efficiency = T4 = 1123.8 K T04 = T4 + C4 = 869.6 m/s M4 = M4 = 1.29 viii) propulsive efficiency = ix) thrust , F F = m.cj ma.u F = 16430 N = 67.5 %

i) ii) iii) iv) v) vi)

5. A ramjet engine flies at an altitude of 6500 m and the flight Mach number is 4. The data for the engine is given below. Air fuel ratio = 52 C. V. of fuel = 44 MJ/kg Diffuser inlet diameter = 0.48 m Efficiency of diffuser = 0.85 Efficiency of combustor = 0.97 Efficiency of nozzle = 0.96

Calculate 1. Ideal cycle efficiency 2. Flight speed 3. Air fuel consumption 4. Diffuser pressure ratio 5. Maximum engine temperature 6. Nozzle pressure ratio 7. Exit Mach number 8. Thrust 9. Air specific impulse 10. Thrust specific fuel ratio Given data Entry Mach no, M1 = 4 Z = 6500 m Di =48 cm = 0.48 m C.V = 44 MJ/kg = 44 106 J/kg M2 =0 ( combustor entry velocity is negligible) ma / mf = 52 = 1.4 R = 287 J/kg K i) Ideal cycle efficiency = Ideal cycle efficiency = 76.2 % ii) flight speed, u M1 = At z= 6500 m from data book a1 = 314.5 m/s 1 = 0.624 kg/m3 P1 = 0.440105 N/m2 u = M1 a1 u = 1258 m/s = 0.762

iii)

iv)

mass flow rate m.a = 1 A 1u = 142 kg/s ma / mf = 52 mf = 2.73 m = ma+ mf m = 144.73 kg/s diffuser pressure ratio , Rod Diffuser efficiency = Rod = 99.3

v)

maximum temperature T3 wkt T3/ T03 Combustion efficiency = =1+ M12

T01 = T02 = 1032.78 K T03 = 1849.2 K vi) nozzle pressure ratio , Ron RON = =

There is no stagnation pressure loss in combustor, i.e , P02 =P03. RON = = Now , ROd= P02= 43.69 105 N/m2 For complete expansion of nozzle, P1 = P4 = 0.44 105 N/m2 Therefore RON = = 99.39 vii) exit Mach number . M4 M4 = W k t T04 = T4 + At nozzle expansion occurs at constant stagnation temperature , i.e., T03 = T04 Nozzle efficiency= To find T4

Inn combustor combustion occurs at constant stagnation pressure i.e., P02 = P03 For complete expansion of nozzle, P1 = P4 = 0.44 105 N/m2 P4/ P03 = 0.01 From gas tables, for P4/ P03 = 0.01 and = 1.4 T4/ T03 = 0.268 T4 = 495.58 K Substituting in nozzle efficiency= T4 = 549.7 K T04 = T4 + C4 = 1616.2 m/s M4 = M4 = 3.44 viii) thrust , F F = m.cj ma.u F = 55276.6 N ix) air specific impulse, Isp Isp= x) Isp = 39.68 s T.S.F.C T.S.F.C = T.S.F.C = 4.938 10-5 kg/Ns

UNIT III FUNDAMENTALS OF ROCKET PROPULSION

1. Describe the working principle of typical rocket engine with neat sketch. Rocket engine consists of fuel and oxygen tanks, combustion chamber

and exhaust nozzle.

Rocket engine works on the principle of Newtons third law of motion. Initially fuel and oxygen (propellants) are separately pumped in to the

combustion chamber. When the combustion of propellants takes place in the combustion

chamber, very high pressure and high temperature gasses are produced. These gasses expanded in the nozzle section to produce the thrust. This

thrust propels the rocket.

2. Derive thrust equation for rocket engine.

Locate control volume of rocket engine. Find pressure force acting on direction of motion in terms of exit and ambient pressures and exit area.

Find momentum thrust force in terms of mass flow rate of propellant and exit velocity.

Finally the summation of pressure force acting on direction of motion and momentum thrust force give thrust equation for rocket engine.

3. Explain classification of rocket nozzles with suitable sketches. Under expanded nozzles Over expanded nozzles Conical nozzles Bell shaped nozzles Free expansion nozzle.

4. A rocket nozzle has a throat area of 20 cm2, combustion chamber pressure of 24 bar and weight flow rate is 45 N/s. If the specific impulse is 128 seconds. Find (i) Thrust coefficient (ii) Propellant weight flow coefficient (iii)Specific propellant consumption

(iv) Characteristic velocity Solution: Find thrust coefficient from specific impulse Find propellant weight flow coefficient from propellant weight flow rate. Find specific propellant consumption from propellant weight flow rate and thrust. Find characteristic velocity from exit velocity and thrust coefficient.

5. A rocket engine has following data: Propellant flow rate = 5.1 kg / s Nozzle exit diameter = 11 cm Nozzle exit diameter = 1.03 bar Ambient pressure = 1.013 bar Thrust chamber pressure = 20 bar Thrust = 6.8 KN. Find the following: (i) Exit Mach number (ii) Nozzle area ratio (iii)Throat area (iv) Thrust coefficient (v) Propellant weight flow coefficient (vi) Characteristic velocity. Take specific heat ratio as 1.3. Solution: Find exit mach number and nozzle area ratio using gas tables.

Find throat area from exit area of nozzle. Find thrust coefficient from throat area. Find propellant weight flow coefficient from mass flow rate of propellant. Find characteristic velocity from exit velocity and thrust coefficient.

UNIT IV CHEMICAL ROCKETS

1. Describe the working of liquid propellant rocket engine with neat sketch.

Schematic diagram of liquid propellant rocket engine Parts of liquid propellant rocket engine Separate tanks for liquid fuel and oxygen Working principle

2. Describe the working of solid propellant rocket engine with neat sketch?

Schematic diagram of solid propellant rocket engine Parts of solid propellant rocket engine Working principle Advantages and disadvantages.

3. With the help of a schematic diagram, explain elaborate scheme of a gas pressure feed system. Schematic diagram of gas pressure feed system Fuel and oxidizer tank Parts of gas pressure feed system. Working principle. 4. What are the important factors that influence the burning rate of a solid propellant? Explain. Chemical composition Method of propellant preparation Initial grain temperature Burning time Chamber pressure Gas velocity adjacent to grain Geometrical shape of the grain

5. Briefly explain about selection criteria of solid propellants? High release of chemical energy High density

Adequate physical properties Fabrication properties Non toxic exhaust gases

UNIT V ADVANCED PROPULSION TECHNIQUES

1. Briefly explain about selection criteria of liquid propellants? Heat of combustion Reaction rate Stability Corrosiveness Vapor pressure

2. Explain briefly about Thrust vector control? Use of Jet-vane Gimbaled engine Use of small control thrust nozzles Use of movable nozzle plug Side injection method

3. Briefly explain about the nuclear rocket engine with diagram? Schematic diagram of nuclear rocket engine Propellant tank Parts of gas pressure feed system. Working principle. 4. Briefly explain about electromagnetic thrusters?

 Schematic diagram of electromagnetic thruster Propellant tank Power supply Working principle.

5. Explain the concept of nozzle less propulsion with example? Concept of nozzle less propulsion Example: Plasma thruster Schematic diagram of Plasma thruster Working principle