Gas dynamics and jet propulsion important questions for AU Apr May 2020 Exams
ANNA UNIVERSITY APR MAY 2020 THEORY EXAMS
POSSIBLE QUESTIONS SET
ME6604 / ME8096 – GAS DYNAMICS AND JET PROPULSION
(REG 2013 & R2017)
UNIT I
1. Air (cp=1.05kJ/kgk, γ=1.38) at p1=3 x 10*5 N/m*2 and T1= 500K flows with a velocity of 200m/s in a 0.3m diameter duct. Calculate: Mass flow rate, stagnation temperature, mach number and stagnation pressure values assuming the flow as compressible and incompressible respectively.
2. Air (γ=1.4, R=287 J/KgK) at an inlet Mach number of 0.2 enters a straight duct at 400K and expands isentropically if the exit Mach number is 0.8 determine the following. i. Stagnation temperature ii. Critical temperature iii. Static temperature at exit iv. Area ratio.
3. A conical diffuser has entry and exit diameters of 15cm and 30cm respectively. The pressure temperature and velocity of air at entry is 0.69bar, 340K and 180m/s respectively. Determine i. The exit pressure, ii. The exit velocity and iii. The force exerted on the diffuser walls. Assume isentropic flow, γ =1.4, Cp =1.005 KJ/Kg-K.
4. (a) Derive the relation of effect of Mach number on Compressibility.
(b) Derive the Bernoulli equation for isentropic compressible flow.
(c) Derive steady floe energy equation.
UNIT II
5. A circular duct passes 8.25Kg/s of air at an exit Mach number of 0.5. The entry pressure and temperature are 3.45bar and 38°C respectively and the coefficient of friction 0.005. If the Mach number at entry is 0.15, determine: i) The diameter of the duct ii) Length of the duct iii) Pressure and temperature at the exit iv) Stagnation pressure loss and v) Verify the exit Mach number through exit velocity and temperature.
6. Air flows out of a pipe with a diameter of 0.3m at a rate of 1000m*3 / min at a pressure and temperature of 150kPa and 293K respectively. If the pipe is 50m long, and assuming that friction coefficient f = 0.005, find the Mach number at exit, the inlet pressure and the inlet temperature.
7. A convergent – divergent nozzle is provided with a pipe of constant cross-section at its exit the exit diameter of the nozzle and that of the pipe is 40cm. The mean coefficient of friction for the pipe is 0.0025. Stagnation pressure and temperature of air at the nozzle entry are 12bar and 600K. The flow is isentropic in the nozzle and adiabatic in the pipe. The Mach numbers at the entry and exit of the pipe are 1.8 and 1.0 respectively. Determine: a) The length of the pipe b) Diameter of the nozzle throat and c) Pressure and temperature at the pipe exit.
8. 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 the gas at exit is 2.5bar 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, Cp= 1.218 KJ/Kg K.
UNIT III
9. 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.
10. The ratio of the exit to entry area in a subsonic diffuser is 4.0 .The Mach number of a jet of air approaching the diffuser at P0 =1.013bar, T =290K 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. 1. Mach number 2. Temperature 3. Pressure 4. What is the stagnation pressure loss between the initial and final states?
11. 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.
12. Starting from the energy equation for flow through a Normal Shock obtain the following relations (or) Prandtl – Meyer relation Cx x Cy =a*² Mx* x My* =1
13. A gas (γ =1.3) at p1 =345Mbar, T1= 350K and M1=1.5 is to be isentropically expanded to 138Mbar. Evaluate (a) the deflection angle, (b) final Mach number and (c) the temperature of the gas.
UNIT IV
14. Explain the principle of operation of a turbojet engine and state its advantages and disadvantages.
15. Explain the working principle of Ramjet engine with a neat sketch.
16. A turbo propels an aircraft at a speed of 900km/hour, while taking 3000Kg of air per minute. The isentropic enthalpy drop in the nozzle is 200KJ/Kg and nozzle efficiency is 90%. The air-fuel ratio is 85 and the combustion efficiency is 95%. The calorific value of the fuel is 42000kJ/Kg. Calculate (i) The propulsive power (ii) Thrust power (iii) Thermal efficiency and (iv) Propulsive efficiency.
17. A turbojet engine is traveling at 850Km/h at standard sea level conditions (101.32KPa and 15°C) .The compressor ratio is 4:1.The turbine inlet temperature is 1000°C. Calculate (i) Specific Thrust (ii) Thrust S A C (iii) Propulsive efficiency. Assume γ = 1.4, Cp = 1.005 kJ/kg. K.
18. A turbojet aircraft flies at 875 Km/hr. at an attitude of 10,000 m above mean sea level. Calculate (i) air flow rate through the engine (ii) thrust (iii) specific thrust (iv) specific impulse (v) thrust power and (vi) TSFC from the following data: diameter of the air at inlet section = 0.75m diameter of jet pipe at exit = 0.5m velocity of the gases at the exit of the jet pipe = 500m/s pressure at the exit of the jet pipe = 0.30 bar air to fuel ratio = 40.
UNIT V
19. (i) Explain the working of Multi-stage rocket with their merits and demerits.
(ii) Describe the importance of characteristic velocity.
20. Explain with a neat sketch the working of a gas pressure feed system and turbo pump feed system used in liquid propellant rocket engines. What are the advantages and disadvantages of liquid propellants compared to solid propellants?
21. Draw the sketch of a liquid and solid propellant rocket engine. Write down its main advantages and disadvantages.
22. A Rocket has the following data: Propellant flow rate: = 203 Kg/s Thrust Chamber Pressure: = 47bar Thrust Chamber temperature: = 3020K Nozzle exit diameter: = 650mm Ambient pressure: = 1.013bar Thrust produced: = 420KN. Calculate effective jet velocity, actual velocity, specific impulse and specific propellant consumption. Recalculate the values of thrust and specific impulse for an altitude of 20000 m.
Warning: These are the questions expected for university exam. These may or may not be asked in exams. Also the data given in the above problems may vary.
