GDJP - Unit 3, 4, 5 qb
- 1. Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering UNIT -3 NORMAL AND OBLIQUE SHOCK 1)What is mean by shock wave ? 2)What is mean by Normal shock? 3)What is oblique shock? 4)Define strength of shock wave? 5)What are applications of moving shock wave ? 6)Shock waves cannot develop in subsonic flow? Why? 7)Define compression and rarefaction shock? Is the latter possible? 8)State the necessary conditions for a normal shock to occur in compressible flow? 9)Give the difference between normal and oblique shock? 10) What are the properties changes across a normal shock? Flow with normal shock 1) The state of a gas (γ=1.3,R =0.469 KJ/Kg K) upstream of a normal shock is given by the following data: Mx =2.5, px= 2bar ,Tx =275K calculate the Mach number, pressure, temperature and velocity of the gas downstream of the shock; check the calculated values with those give in the gas tables. (16) 2) The ratio of th exit to entry area in a subsonic diffuser is 4.0 .The Mach number of a jet of air approaching the diffuser at p0=1.013 bar, T =290 K is 2.2 .There is a standing normal shock wave just outside the diffuser entry. The flow in the diffuser is isentropic . Determine at the exit of the diffuser. a) Mach number , (4) b) Temperature, and (4) c) Pressure (4) d) What is the stagnation pressure loss between the initial and final states of the flow? (4) 3) The velocity of a normal shock wave moving into stagnant air (p=1.0 bar, t=17°C ) is 500 m/s . If the area of cross- section of the duct is constant determine (a) pressure (b) temperature (c) velocity of air (d) stagnation temperature and (e) the mach number imparted upstream of the wave front. (16) 4) The following data refers to a supersonic wind tunnel: Nozzle throat area =200cm²
- 2. Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering Test section cross- section =337.5cm² Working fluid ;air (γ =1.4, Cp =0.287 KJ/Kg K) Determine the test section Mach number and the diffuser throat area if a normal shock is located in the test section. (16) 5) A supersonic diffuser for air (γ =1.4) has an area ratio of 0.416 with an inlet Mach number of 2.4 (design value). Determine the exit Mach number and the design value of the pressure ratio across the diffuser for isentropic flow. At an off- design value of the inlet Mach number (2.7) a normal shock occurs inside the diffuser .Determine the upstream Mach number and area ratio at the section where the shock occurs, diffuser efficiency and the pressure ratio across the diffuser. Depict graphically the static pressure distribution at off design. (16) 6) Starting from the energy equation for flow through a normal shock obtain the following relations (or) prandtl – meyer relation Cx Cy =a* ² M*x M*y =1 (16) Flow with oblique shock waves 7) Air approaches a symmetrical wedge (δ =15°) at a Mach number of 2.0.Determine for the strong and weak waves (a) wave angle (b) pressure ratio (c) density ratio,(d) temperature ratio and (e)downstream Mach number Verify these values using Gas tables for normal shocks. (16) 8) A gas (γ =1.3) at p1 =345 mbar, T1= 350 K and M1=1.5 is to be isentropically expanded to 138 mbar. Determine (a) the deflection angle ,(b) final Mach number and (c) the temperature of the gas. (16) 9) A jet of air at Mach number of 2.5 is deflected inwards at the corner of a curved wall..The wave angle at the corner is 60°.Determine the deflection angle of the wall, pressure and temperature ratios and final Mach number. 10) Derive the Rankine –Hugoniot relation for an oblique shock 11) Compare graphically the variation of density ratio with the intial Mach number in isentropic flow and flow with oblique shock. (16) 12) The Mach number at the exit of a combustion chamber is 0.9. The ratio of stagnation temperature at exit and entry is 3.74.If the pressure and temperature of a gas at exit are2.5 bar and 1000°C respectively determine (a) Mach number ,pressure and temperature of the gas at entry,(b) the heat supplied per Kg of the gas and (c) the maximum heat that can be supplied. Take γ =1.3 and Cp =1.218 KJ/Kg K (16) 13) The conditions of a gas in a combustor at entry are: P1=0.343 bar,T1= 310K ,C1=60m/s Determine the Mach number ,pressure, temperature and velocity at the exit if the increase in stagnation enthalpy of the gas between entry and
- 3. Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering exit is 1172.5KJ/Kg.Take Cp= 1.005KJ/kg, γ =1.4. (16) 14) A combustion chamber in a gas turbine plant receives air at 350 K , 0.55 bar and 75m/s .The air –fuel ratio is 29 and the calorific value of the fuel is 41.87 MJ/Kg. Taking γ =1.4 and R =0.287 KJ/Kg K for the gas determine : a) The initial and final Mach number, (4) b) Final pressure, temperature and velocity of the gas, (4) c) Percent stagnation pressure loss in the combustion chamber and (4), d) The maximum stagnation temperature attainable. (4) 15) Obtain an equation representing the rayleigh line. Draw Rayleigh lines on the h-s and p-v planes for two different values of the mass flux. Show that the slope of the Rayleigh line on the p-v plane is {dP/dV} r = þ² c² (16) UNIT -4 JET PROPULSION PART A 1) Differentiate jet propulsion and rocket propulsion (or) differentiate between air breathing and rocket propulsion? 2) What is monopropellant? Give one example for that? 3) What is bipropellant? 4) Classify the rocket engines based on sources of energy employed? 5) What is specify impulse of rocket? 6) Define specific consumption? 7) What is weight flow co-efficient? 8) What is IWR? 9) What is thrust co-efficient? 10) Define propulsive efficiency? PART B 1) Aircraft speed of 525 Kmph. The data for the engine is given below Inlet diffuser efficiency =0.875 Compressor efficiency =0.790 Velocity of air at compressor entry =90m/s Properties of air :γ =1.4, Cp =1.005 KJ/kg K 2) The diameter of the propeller of an aircraft is 2.5m; It flies at a speed of
- 4. 4 Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering 500Kmph at an altitude of 8000m. For a flight to jet speed ratio of 0.75 determine (a) the flow rate of air through the propeller, (b) thrust produced (c) specific thrust, (d) specific impulse and (e) the thrust power. (16) 3) An aircraft flies at 960Kmph. One of its turbojet engines takes in 40 kg/s of air and expands the gases to the ambient pressure .The air –fuel ratio is 50 and the lower calorific value of the fuel is 43 MJ/Kg .For maximum thrust power determine (a)jet velocity (b) thrust (c) specific thrust (d) thrust power (e) propulsive, thermal and overall efficiencies and (f) TSFC (16) 4) A turbo jet engine propels an aircraft at a Mach number of 0.8 in level flight at an altitude of 10 km .The data for the engine is given below: Stagnation temperature at the turbine inlet =1200K Stagnation temperature rise through the compressor =175 K Calorific value of the fuel =43 MJ/Kg Compressor efficiency =0.75 Combustion chamber efficiency =0.975 Turbine efficiency =0.81 Mechanical efficiency of the power transmission between turbine and compressor =0.98 Exhaust nozzle efficiency=0.97 Specific impulse =25 seconds Assuming the same properties for air and combustion gases calculate a) Fuel–airratio,(2) b) Compressorpressureratio,(4) c) Turbine pressure ratio,(4) d) Exhaustnozzlespressureratio,and(4) e) Machnumberofexhaustjet(2) 5) Determine the maximum velocity of a rocket and the altitude attained from the following data: Mass ratio =0.15 Burn out time =75s Effective jet velocity =2500m/s Whatarethevaluesofthevelocityandaltitudelossesduetogravity?Ignoredrag and vertical trajectory . (16) 6) A missile has a maximum flight speed to jet speed ratio of 0.2105 and specific impulse equal to 203.88 seconds .Determine for a burn out time of 8 seconds a) Effectivejetvelocity(4)
- 5. 5 Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering b) Massratioandpropellantmassfunctions(4) c) Maximumflightspeed,and(4) d) Altitudegainduringpoweredandcoastingflights(4)
- 6. 1 Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering UNIT -5 SPACE PROPULSION PART A 1) Differentiate jet propulsion and rocket propulsion (or) differentiate between air breathing and rocket propulsion? 2) What is monopropellant? Give one example for that? 3) What is bipropellant? 4) Classify the rocket engines based on sources of energy employed? 5) What is specify impulse of rocket? 6) Define specific consumption? 7) What is weight flow co-efficient? 8) What is IWR? 9) What is thrust co-efficient? 10) Define propulsive efficiency? PART B 1) Calculate the orbital and escape velocities of a rocket at mean sea level and an altitude of 300km from the following data: Radius of earth at mean sea level =6341.6Km Acceleration due to gravity at mean sea level =9.809 m/s² (16) 2) With a neat sketches the principle of operation of: 1. turbo fan engine and (8) 2. ram jet engine (8) 3) Explain the construction and operation of a ramjet engine and derive an expression for the ideal efficiency. (16) 4) Explain the construction and operation of a solid propellant rocket engine. Also name a n y four solid propellants and state its advantages and disadvantages. (16) 5) What are the advantages and disadvantages of liquid propellants compared to solid propellants? (16) 6) Discuss in detail the various propellants used in solid fuel rockets and liquid fuel system. Also sketch the propellant feed-system for a liquid propellant rocket motor. (16) 7) Briefly explain the construction and working of : A. Rocket engine (6) B. Ramjet engine (6) C. Pulsejet engine (4)
- 7. 1 Mr.S.Gokul/Assistant Professor/Mechanical Engineering/Sri Eshwar College of Engineering 8) With the help of a neat sketch describe the working of a ramjet engine. Depict the various thermodynamic process occurring in it on h-s diagram. What is the effect of flight Mach number on its efficiency? (16) 9) Explain with a neat sketch the working of a turbo-pump feed system used in a liquid propellant rocket? (16) 10) A Rocket has the following data: Propellant flow rate : 203 kg/s Thrust Chamber Pressure : 47 bar Thrust Chamber temperature : 3020 K Nozzle exit diameter : 650 mm Ambient pressure : 1.013 bar Thrust produced : 420 kN Calculate effective jet velocity, actual velocity, specific impulse and specific propellant consumption. Recalculate the values of thrust and specific impulse for an altitude of 20000m 11) A rocket nozzle has an exit area ratio 3:1 with isentropic expansion. What will be the thrust per unit area of exit and specific impulse if the combustion chamber temperature is 2973 K and pressure is 20 bar. Assume atm pressure is 1 bar and R=0.287kJ/kg k and γ =1.3.