Turbomachinery - Propulsion - Aerospace Engineering Previous Year Questions

1

2007 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2007

Which one of the following engines should be used by a subsonic passenger transport airplane for minimum specific fuel consumption?

A

Turbojet engine with afterburner

B

Turbofan engine

C

Ramjet engine

D

Scramjet engine

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2

2007 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2007

For an impulse turbine with identical stages, the hot gas exits from the stator blades at the mean blade height at an absolute angle of 70 degrees with the axis of the turbine. If the absolute inlet blade angle with the axis of the turbine at the mean blade height for the rotor blades is 37 degrees, then the absolute exit blade angle with the axis of the turbine at the mean blade height of the rotor blades is

A

33 degrees

B

37 degrees

C

53 degrees

D

53.5 degrees

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3

2008 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2008

A turbofan engine has a bypass ratio of 5 and a total mass flow rate of 120 kg/s. The mass flow rate through the bypass duct is

A

20 kg/s

B

100 kg/s

C

120 kg/s

D

600 kg/s

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4

2008 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2008

An ideal ramjet engine is flying at a Mach number M. The exhaust gas static temperature at the outlet of the nozzle is T_e. The ambient static temperature is T_a. Gas constant R and specific heat ratio \(\gamma\) do not vary through the ramjet. Assuming that nozzle exhaust static pressure is equal to the ambient pressure and fuel air ratio \(f << 1\), the thrust per unit mass flow rate is

A

\(\sqrt{\gamma R T_a} \left[ \sqrt{\frac{T_e}{T_a}} \right]\)

B

\(\sqrt{\gamma R T_a} \left[ \sqrt{\frac{T_e}{T_a}} - 1 \right]\)

C

\(M \sqrt{\gamma R T_a} \left[ \sqrt{\frac{T_e}{T_a}} - 1 \right]\)

D

\(M \sqrt{\gamma R T_a} \left[ \sqrt{\frac{T_e}{T_a}} \right]\)

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5

2008 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2008

The mass flow rate of fuel is

A

300.3 kg/s

B

269.5 kg/s

C

87.4 kg/s

D

49.9 kg/s

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6

2008 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2008

The total temperature rise across the compressor is

A

310.1 K

B

366.3 K

C

392.1 K

D

405.4 K

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7

2009 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2009

The propulsive efficiency of a turbo-jet engine moving at velocity \(U_{\infty}\) and having exhaust velocity \(U_e\) with respect to the engine is given by

A

\(\frac{2}{U_{\infty}/U_e + 1}\)

B

\(1 - \frac{U_{\infty}}{U_e}\)

C

\(\frac{2U_{\infty}U_e}{U_e^2 + U_{\infty}^2}\)

D

\(\frac{2U_{\infty}}{U_e + U_{\infty}}\)

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8

2009 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2009

Thrust available from a turbojet engine

A

increases as altitude increases

B

increases up to the tropopause and then decreases

C

remains constant at all altitudes

D

decreases as altitude increases

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9

2009 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2009

An ideal axial compressor is driven by an ideal turbine across which the total temperature ratio is 0.667. If the total temperature at turbine inlet is \(T_0 = 1500\) K and specific heat of gas \(c_p = 1\) kJ/kg/K, the power drawn by the compressor per unit mass flow rate of air is approximately

A

300 kW/kg/s

B

1000 kW/kg/s

C

600 kW/kg/s

D

500 kW/kg/s

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10

2009 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2009

If the sensible enthalpy of fuel is neglected, the temperature of combustion products from the reaction of primary air stream with fuel is approximately

A

2100 K

B

3200 K

C

2600 K

D

1800 K

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11

2009 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2009

The approximate ratio of mass flow rates of the primary air stream to the secondary air stream required to achieve the turbine inlet total temperature of 1550 K is

A

2:1

B

1:2

C

1:1.5

D

1:1

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12

2011 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2011

Thermodynamic cycle on which the jet engine operates can be

A

open Rankine cycle only

B

either open or closed Rankine cycle

C

open Brayton cycle only

D

either open or closed Brayton cycle

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13

2011 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2011

Propulsion efficiency of a jet engine is

A

directly proportional to both the thrust power and the air mass flow rate

B

inversely proportional to both the thrust power and the air mass flow rate

C

directly proportional to the thrust power and inversely proportional to the air mass flow rate

D

inversely proportional to the thrust power and directly proportional to the air mass flow rate

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14

2014 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2014

For a given fuel flow rate and thermal efficiency, the take-off thrust for a gas turbine engine burning aviation turbine fuel (considering fuel-air ratio < 1) is

A

Directly proportional to exhaust velocity

B

Inversely proportional to exhaust velocity

C

Independent of exhaust velocity

D

Directly proportional to the square of the exhaust velocity

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15

2014 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2014

The thrust produced by a turbojet engine

A

Increases with increasing compressor pressure ratio

B

Decreases with increasing compressor pressure ratio

C

Remains constant with increasing compressor pressure ratio

D

First increases and then decreases with increasing compressor pressure ratio

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16

2016 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2016

Which of the following aircraft engines has the highest propulsive efficiency at a cruising Mach number of less than 0.5?

A

Turbofan engine

B

Turbojet engine

C

Turboprop engine

D

Ramjet engine

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17

2016 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2016

An un-mixed turbofan engine with a bypass ratio of 6.0, flies with a velocity of 200 m/s. The core and the bypass nozzles of the engine, that are both convergent nozzles, operate under choked condition and have exhaust static temperatures of 580 K and 295 K, respectively. The specific gas constant and the ratio of specific heats for both the streams are 287 J/kgK and 1.4, respectively. If the fuel-air ratio is negligible, the thrust per unit mass flow rate generated by the engine is ______ Ns/kg.

Check answer

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18

2016 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2016

A single-stage gas turbine operates with an axial absolute flow at the entry and exit from the stage. The absolute flow angle at the nozzle exit is \( 70^\circ \). The turbine stage generates a specific work of 228 kJ/kg when operating with a mean blade speed of 440 m/s. The absolute velocity at the rotor entry is

A

275.7 m/s

B

551.5 m/s

C

1103.0 m/s

D

1654.5 m/s

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19

2016 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2016

An axial compressor operates such that it has an inlet and an exit total temperature of 300 K and 430 K, respectively. The isentropic efficiency of the compressor is 85 %. If the ratio of specific heats is 1.4, then the total pressure ratio across the compressor is ______.

Check answer

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20

2016 · Aerospace Engineering · Propulsion · Turbomachinery

Aerospace Engineering (AE) 2016

A substance experiences an entropy change of \(\Delta s > 0\) in a quasi-steady process. The rise in temperature (corresponding to the entropy change \(\Delta s\)) is highest for the following process:

A

isenthalpic

B

isobaric

C

isochoric

D

isothermal

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