model exam complete answers gd jp

7/30/2019 Model Exam Complete Answers Gd Jp

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1.Write down the relation between static

temperature, stagnation temperature and mach

number for the case of an isentropic flow.


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2.A plane travels at a velocity of 1600 kmph at an

altitude of 8000 m. Find the mach angle and mach

number.


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3.Give one example each for Fanno flow and

Rayleigh flow.Fanno Flow: Aircraft propulsion system; Air conditioning system;

Chemical Process plants

Rayleigh Flow: Regenerator; Intercooler; Combustion Chamber pipes;

Exhaust gas pipes


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4.Differentiate between Fanno flow and Rayleigh

flow.S.No. Fanno Flow Rayleigh Flow1 Friction is present No Friction

2 No Heat transfer Heat transfer is present


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5.Show a normal shock in h-s diagram with the help

of Rayleigh line and Fanno line.


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6.Write the Rankine Hugoniot equation.


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7.List the different types of Jet engines.Turbo jet; Ram jet; Pulse jet engines.


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8.Define the principle of Ram jet engine.

When aircraft flies at very high velocity, incoming air is compressed to

very high pressure without external work. Kinetic energy is converted to

pressure energy by induction of shock in supersonic diffuser. This

principle is called Ram Effect.


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9.What are the types of propellant feed systems?Gas Pressure Feed System; and

Pump Feed system

Gas Pressure Feed System: An inert gas is separately carried at a pressure

greater than the injection pressure. It exerts the required pressure on the

propellant tank. Inert gas pressure feed the fuel and oxidizer into the

combustion chamber. Nitrogen, Helium, air are used for pressurization.

Pump feed System: Both fuel and oxidizer pumps are used to feed into

the combustion chamber and these are driven by a single turbine.

FuelGas

Oxidiser

to injector to injector

O

FHP inert gas

Shut off valve PRV


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10.What are the classifications of a Rocket?1.Based on energy source: Chemical, Electrical, Nuclear, Solar rockets;

2.Based on application: Spy; Missile; Space Craft; Military; Weather rockets;

3.Based on propellant used: Liquid, Solid, hybrid propellant rocket;

4. Based on number of stages: Single stage, Multiple stage rockets;

5.Based on size and range: Short range, Long range and Medium range rockets.


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11.a. i. Derive the expression for velocity of sound in air.


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11.a. ii. A jet fighter is flying at M=2.5. It is observed directly overheard at

a height of 10 km. How much distance it would cover before the sonic

boom is heard on the ground.


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11.b.i.Derive the expression for compressibity in terms of mach number

variation.

11.b. ii. Determine the mach number of an aircraft at which the velocity temperatureof the air at the entry of engine equals static temperature.


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12.a.i. Show that the maximum entropy point occurs at M = 1 in Rayleigh line.


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12.a.ii. Air enters a constant area duct at M1

=3, P1

= 1 atm, and T1

= 310 K.

Inside the duct heat added per unit mass is 3x10 5 J/kg. Calculate the flow

properties static pressure, temperature, density, stagnation pressure and

temperature, and mach number at the exit.


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12.b.i. Show that the maximum enthalpy point occurs at M = 1 / .


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12.b.ii. A long pipe of 0.3 m diameter has a mean coefficient of 0.002. Air enters the

pipe at a mach number of 3, stagnation temperature 310 K and static pressure 0.507

bar. Determine for a section at which the mach number reaches 1.5: (a) Static pressure

and temperature, (b) stagnation pressure and temperature, (c) velocity of air, (d)

distance of the section from the inlet and (e) mass flow rate of air.


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13.a.i. Derive the Rankine Hugoniot ralation.


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13.a.ii. Air at M = 2.5 enters a convergent duct with an area ratio of A 2/A1 = 0.5.Normal shock

occurs at a test section X where AX/A1 = 0.6. Find the exit mach number and pressure ratio

across the duct.


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13.b.i. Derive the Prandtl Meyer relation.

Substituting (2) and (3) in (1) and solving,


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13.b. ii. A gas ( = 1.3) at 345 mbar pressure, 350 K temperature and mach number 1.5 is to be

isentropically expanded to 138 mbar. Determine the deflection angle, final mach number and

temperature of the gas.


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14.a.i. What are the classifications of an aircraft propulsion engine?ii. Explain with a neat sketch the operation of a turbojet engine.

Classifications of aircraft propulsion engine:Turbo jet; turbo fan; turbp prop; Ram jet; Pulse jet engines.


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Performance of Turbo Jet Components

S.No. Component Process Detail Process reference Efficiency

1 Diffuser Increase pressure

transformation

[ i-1 ]

Adiabatic

[ i-1 ]

Isentropic

2 Compressor Increase pressure

work transfer

[ 1-2 ] [ 1-2 ]

Isentropic

3 Combustor Heat added-

Constant pressure

[ 2-3 ] [ 2-3 ]

Rayleigh

4 Turbine Increase velocity

work transfer

[ 3-4 ] [3-4 ]

Isentropic

5 Nozzle Increase velocity

transformation

[ 4-e ] [ 4-e ]

Isentropic


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14.b. A turbojet propels an aircraft at a speed of 900 kmph while taking 3000 kg of air per

minute. The isentropic enthalpy drop in the nozzle is 200 J/kg and nozzle efficiency is 90%. The

air fuel ratio is 85. The combustion efficiency is 95%; Calorific value of fuel is 42 MJ/kg.

Calculate the thrust, propulsive power, propulsive efficiency, thermal efficiency, specific fuel

consumption.

See next slide for answer


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15.a.i. Compare jet propulsion and rocket propulsion.ii. The effective jet velocity of a rocket is 2700 m/s. the forward flight velocity is 1350 m/s.

Propellant consumption rate is 78.6 kg/s. Calculate the thrust, thrust power and propulsive

efficiency.Jet propulsion Rocket propulsion

Oxygen obtained from ambient

air for combustion

Contains own oxygen supply for for

combustion

Jet consists of cold air and

combustion products. Jet consists of exhaust gases only

Mechanical devices used. No mechanical devices used.


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15.b.i. State the advantages and disadvantages of solid propellant rockets.

Advantages of Solid Propellant Rockets:

Simple in design;

Mass production at short notice is possible;

No feed system;

No moving parts;

Lighter in weight; Less frequent problems.

Disadvantages of Solid Propellant Rockets:

Difficult to control in emergency;

Thrust regulation and combustion regulation are difficult;

Low specific impulse;Nozzle cooling and refuelling not possible;

Nozzle erosion occurs.


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15.b. ii. A rocket nozzle has an exit area to throat area 3:1 with isentropic expansion. What will

be the thrust per unit area of exit and specific impulse. If the combustion chamber

temperature is 2700 o C and pressure is 20 bar. Assume atmospheric pressure is 1 bar; = 1.3,

R = 248 J/kgK.

= 1.3, R = 248 J/kgK


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