Inlets, Combustors & Nozzles

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

In the figure shown below, the flow at the nozzle exit is ________.

Subtopic: Combustor Design & Flame Stabilization

In an ideal turbofan engine shown in the figure below, the compressor is driven by the high pressure turbine, and the fan is driven by the low pressure turbine. The stations 0, 2, 3, 4, 4.5, and 5 refer to free-stream, compressor inlet, compressor outlet, combustor exit, high pressure turbine exit, and low pressure turbine exit, respectively, and the subscript 't' refers to the total condition. Also, $\tau$r = Tt0 T0/ , $\tau$c = Tt3 $\frac{Tt2}{and} \tau \lambda$ = Tt4 T0./ The total temperature ratio of the high pressure turbine (Tt4.5 Tt4/ ) is given by ______.

Subtopic: Subsonic & Supersonic Inlets

A centrifugal compressor has a constant-width radial diffuser. The diameters at the diffuser inlet and outlet are 0.2 m and 0.3 m, respectively. The flow at the diffuser inlet and outlet is assumed to be steady and uniform. The average velocity at the diffuser inlet and outlet are (60 êr + 75 ê𝛉) m/s and (u êr + 50 ê𝛉) m/s, respectively. If the flow through the diffuser is treated as steady and incompressible, the value of u is ________ (rounded off to the nearest integer).

Subtopic: Combustor Design & Flame Stabilization

$\delta Q$ and $\delta W$ are the heat and work interactions of a system with its surroundings, and $dU$ is the change in the internal energy of the system. For an adiabatic process in a closed, constant pressure combustor, which one of the following options is correct?

Subtopic: Subsonic & Supersonic Inlets

A perfect gas flows through a frictionless constant-area duct with heat addition. The inlet conditions are as follows: pressure 100 kPa, density 1 kg/m^3, and velocity 100 m/s. At a particular downstream location, the gas velocity is 200 m/s. The static pressure at the downstream location is ______ kPa (answer in integer).

Subtopic: Combustor Design & Flame Stabilization

Air at temperature 300 K is compressed isentropically from a pressure of 1 bar to 10 bar in a compressor. Eighty percent of the compressed air is supplied to a combustor. In the combustor, 0.88 MJ of heat is added per kg of air. The specific heat at constant pressure is Cp = 1005 J/kg/K and the specific heat ratio is $\gamma$ = 1.4. The temperature of the air leaving the combustor is ___ K (rounded off to one decimal place).

Subtopic: Subsonic & Supersonic Inlets

An ideal ramjet with an optimally expanded exhaust is travelling at Mach 3. The ambient temperature and pressure are 260 K and 60 kPa, respectively. The inlet air mass flow rate is 50 kg/s. Exit temperature of the exhaust gases is 700 K. Fuel mass flow rate is negligible compared to air mass flow rate. Gas constant is R= 287 J/kg/K, and specific heat ratio is $\gamma$ = 1.4. The thrust generated by the engine is ___ kN (rounded off to one decimal place).

Subtopic: Subsonic & Supersonic Inlets

A single-stage axial compressor, with a 50 % degree of reaction, runs at a mean blade speed of 250 m/s. The overall pressure ratio developed is 1.3. Inlet pressure and temperature are 1 bar and 300 K, respectively. Axial velocity is 200 m/s. Specific heat at constant pressure, Cp = 1005 J/kg/K and specific heat ratio, $\gamma$ = 1.4. The rotor blade angle at the outlet is ___ degrees (rounded off to two decimal places).

Subtopic: Combustor Design & Flame Stabilization

A surrogate liquid hydrocarbon fuel, approximated as C 10H12, is being burned in a land-based gas turbine combustor with dry air (79% N 2 and 21% O 2 by volume). How many moles of dry air are required for the stoichiometric combustion of the surrogate fuel with dry air at atmospheric temperature and pressure?

Subtopic: Combustor Design & Flame Stabilization

In an airbreathing gas turbine engine, the combustor inlet temperature is 600 K. The heating value of the fuel is 43.4 × 106 J/kg. Assume Cp to be 1100 J/kg K for air and burned gases, and fuel-air ratio f << 1.0. Neglect kinetic energy at the inlet and exit of the combustor and assume 100% burner efficiency . What is the fuel -air ratio required to achieve 1300 K temperature at the combustor exit?

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

An aircraft with a turbojet engine is flying with 250 m/s speed at an altitude, where the density of air is 1 kg/m^3. The inlet area of the engine is 1 m^2. The average velocity of the exhaust gases at the exit of the nozzle, with respect to aircraft, is 550 m/s. Assume the engine exit pressure is equal to the ambient pressure and the fuel-air ratio is negligible. The uninstalled thrust produced by the engine at these conditions is _________ N (rounded off to the nearest integer).

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Which of the following statements is/are TRUE for an axial turbine?

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

The equations of motion for a two degrees of freedom undamped spring -mass system are: mẍ1 + 2kx1 - kx2 = 0 mẍ2 - kx1 + 2kx2 = 0 where m and k represent mass and stiffness respectively, in corresponding SI units, and x1 and x2 are the degrees of freedom. The larger of the two natural frequencies is given by: $\omega$ = $\alpha$√ k m rad/s. The value of $\alpha$ is ________ (rounded off to 2 decimal places).

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

From the options given, select all that are true for turbofan engines with afterburners.

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

The operating characteristics of a pump were measured to be 2 PCa = , where power coefficient 35P PC D= ,  is the flow coefficient, a is a constant, D is a length scale,  is the rotation rate,  is fluid density, and P is the power required. The flow coefficient is a dimensionless volume flow rate scaled with  and D . Assuming that the flow rate remains the same, if the rotation rate is increased to 1.25 , the power changes to P . The value of  is ________ . (round off to two decimal places)

Subtopic: Combustor Design & Flame Stabilization

A gas turbine combustor is burning methane and air at an equivalence ratio 0.5 = , where   stoich FA FA = and [] stoichFA is the ratio of mass flow rate of fuel to the mass flow rate of air at stoichiometry. If the air flow rate is 20airm = kg/s then the mass flow rate of methane is ______ kg/s. (round off to two decimal places)

Subtopic: Subsonic & Supersonic Inlets

A centrifugal air compressor has inlet root diameter of 0.25 m and the outlet diameter of the impeller is 0.6 m. The pressure ratio is 5.0. The air at the inlet of the rotor is at 1 atm and 25oC. The polytropic efficiency is 0.8 and slip factor is 0.92. Use 1.004pC = kJ/kg-K and 1.4 = . The impeller speed in revolutions per minute (RPM) is _________. (round off to the nearest integer)

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

In a single stage turbine, the hot gases come out of stator/ nozzle at a speed of 500 m/s and at an angle of 70 degrees with the turbine axis as shown. The design speed of the rotor blade is 250 m/s at the mean blade radius. The rotor blade angle, $\beta$, at the leading edge is ________ degrees (round off to one decimal place).

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

A convergent-divergent nozzle with adiabatic walls is designed for an exit Mach number of 2.3. It is discharging air to atmosphere under the conditions indicated in the figure. Flow through the nozzle is inviscid, the characteristic gas constant for air is 287 J/(kg-K), and $\gamma$ =1.4. When the reservoir pressure is 25 bar (absolute), and temperature is 300 K, Prandtl- Meyer expansion waves appear at the nozzle exit as shown. The minimum percentage change in the reservoir pressure required to eliminate the wave system at the nozzle exit under steady state is _______%.

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

An ideal ramjet is to operate with exhaust gases optimally expanded to ambient pressure at an altitude where temperature is 220 K. The exhaust speed at the nozzle exit is 1200 m/s at a temperature of 1100 K. Given: $\gamma$ = 1.4 at 220 K; R = 287 J/(kg-K) for air $\gamma$ = 1.33 at 1100 K; R = 287 J/(kg-K) for exhaust gases. The cruise speed of this ramjet is ________ m/s (rounded off to nearest integer).

Subtopic: Subsonic & Supersonic Inlets

Q.47 A single stage axial turbine has a mean blade speed of 340 m/s at design condition with blade angles at inlet and exit of the rotor being 21° and 55°, respectively. The degree of reaction at the mean radius of the rotor is equal to 0.4. The annulus area at the rotor inlet is 0.08 m? and the density

Subtopic: Combustor Design & Flame Stabilization

The operating conditions of an aircraft engine combustor are as follows. The rate of total enthalpy of air entering the combustor = 28.94 MJ/s. The rate of total enthalpy of air leaving the combustor = 115.42 MJ/s. Mass flow rate of air = 32 kg/s. Air to fuel mass ratio = 15.6. Lower heating value of the fuel = 46 MJ/kg. The efficiency of the combustor is _% (round off to two decimal places). O.No. 49

Subtopic: Subsonic & Supersonic Inlets

In a turbojet engine, the compressor outlet temperature increases with decreasing efficiency of the compressor. If the turbine inlet temperature remains constant, with decreasing efficiency of the compressor, the thrust specific fuel consumption of the engine

Subtopic: Subsonic & Supersonic Inlets

In a turbojet engine, the compressor outlet temperature increases with decreasing efficiency of the compressor. For a fixed turbine inlet temperature (TIT), how does this affect the turbine work output and the cycle thermal efficiency?

Subtopic: Subsonic & Supersonic Inlets

The combustion products of a gas turbine engine can be assumed to be a calorically perfect gas with $\gamma = 1.33$ and $C_p = 1150$ J/(kg K). The total temperature and total pressure at the turbine inlet are 1400 K and 1000 kPa, respectively. If the expansion across the turbine has a total pressure ratio of 4.0, the total temperature (in K) drop across the turbine divided by 10 is _________ (round off to the nearest integer).

Subtopic: Subsonic & Supersonic Inlets

The figure shows the velocity triangles for an axial compressor stage. The axial velocity is constant through the stage at $C_a = 150$ m/s. The rotor blade speed is $U = 200$ m/s. If the rotor blade angles at inlet and exit are $\beta_1 = 45^\circ$ and $\beta_2 = 15^\circ$, the specific work input (in kJ/kg) to the compressor stage is _________ (round off to 2 decimal places).

Subtopic: Combustor Design & Flame Stabilization

In an aircraft gas turbine engine combustor, the primary zone equivalence ratio is typically maintained close to stoichiometric mainly to:

Subtopic: Subsonic & Supersonic Inlets

An axial compressor that generates a stagnation pressure ratio of 4.0, operates with inlet and exit stagnation temperatures of 300 K and 480 K, respectively. If the ratio of specific heats ($\gamma$) is 1.4, the isentropic efficiency of the compressor is

Subtopic: Subsonic & Supersonic Inlets

An axial compressor rotor with 50 % degree of reaction, operates with an axial velocity of 200 m/s. The absolute flow angle at the inlet of the rotor is 22 o with reference to the axial direction. If the axial velocity is assumed to remain constant through the rotor, the magnitude of the relativ e velocity (in m/s) at the rotor exit is __________ ( accurate to one decimal place). 9/9

Subtopic: Subsonic & Supersonic Inlets

An axial compressor rotor with 50 % degree of reaction, operates with an axial velocity of 200 m/s. The absolute flow angle at the inlet of the rotor is $22^\circ$ with reference to the axial direction. If the rotor blade speed is 300 m/s, the work done per unit mass (in J/kg) by the compressor rotor on the air is _________ (accurate to one decimal place).

Subtopic: Subsonic & Supersonic Inlets

In the vane-less space between the impeller and the diffuser vanes in a Centrifugal Compressor, the angular momentum varies in the following manner in the radial direction

Subtopic: Combustor Design & Flame Stabilization

In a combustor, gaseous Octane (CsHs) and air are to be burned in stoichiometric proportions. If the required flowrate of air is 1 kg/s, what should be the corresponding flow rate of Octane?

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

A rule states that in order to drink beer. one must be over 18 years old. In a bar, there are 4 people. P is 16 years old. Q is 25 years old. R is drinking milkshake and S is drinking a beer. What must be checked to ensure that the rule is being followed?

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Water exits from a nozzle of diameter 50 mm into atmospheric pressure at a velocity of 20 m/s. The jet strikes a flat vertical plate held perpendicular to the jet. If the density of water is 1000 kg/m$^3$, the horizontal force (in N) required to hold the plate stationary is _________ (round off to the nearest integer).

Subtopic: Combustor Design & Flame Stabilization

In a combustor, gaseous Octane ($\text{C}_8\text{H}_{18}$) and air are to be burned in a stoichiometric ratio. Assuming air contains 21% $\text{O}_2$ and 79% $\text{N}_2$ by volume, the stoichiometric fuel-to-air mass ratio $(f/a)_{stoich}$ is approximately:

Subtopic: Subsonic & Supersonic Inlets

The given diagram represents the velocity triangle of an axial compressor stage with blade speed $U = 200$ m/s, axial velocity $V_z = 165$ m/s, relative inlet angle $\beta_1 = 45^\circ$, and relative exit angle $\beta_2 = 11^\circ$. The specific work input (in kJ/kg) to the air is _________ (round off to 1 decimal place).

Subtopic: Subsonic & Supersonic Inlets

An aircraft with a turbojet engine, having an inlet area of 1 m$^2$, is flying at a speed of 250 m/s at an altitude where air density is 0.4 kg/m$^3$. The mass flow rate of fuel added is negligible. If the jet exhaust speed is 600 m/s relative to the aircraft, the net thrust (in kN) produced by the engine is _________ (round off to the nearest integer).

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

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.

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

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 o. 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

Subtopic: Combustor Design & Flame Stabilization

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 ____________.

Subtopic: Subsonic & Supersonic Inlets

A gaseous mixture of air and fuel enters a constant area combustion chamber at a velocity of 100 m/s and at a static temperature of 300 K. The heat release due to combustion is 1000 kJ/kg. The specific heat at constant pressure of the calorically perfect gas is 1000 J/kgK. The total temperature of air-fuel mixture after combustion is ______ K.

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Consider 1 -D, steady, in viscid, compressible flow through a convergent nozzle . The total temperature and total pressure are oT , oP respectively. The flow through the nozzle is choked with a mass flow rate of om . If the total temperature is increased to oT 4, with total pressure remaining unchanged, then the mass flow rate through the nozzle

Subtopic: Combustor Design & Flame Stabilization

Air enters an aircraft engine at a velocity of 180 m/s with a flow rate of 94 kg/s. The engine combustor requires 9.2 kg/s of air to burn 1 kg/s of fuel. The velocity of gas exiting from the engine is 640 m/s. The momentum thrust (in N) developed by the engine is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Let $\Gamma$ be the boundary of the closed circular region A given by 12 2≤ +y x . Then ( )∫ $\Gamma$ - = dsxyx I 23 9 3 (where ds means integration along the bounding curve) is

Subtopic: Subsonic & Supersonic Inlets

A cruise missile with an ideal ramjet engine is flying at Mach 4.0 at an altitude where the ambient temperature is 100K. Considerratio of specific heats$\gamma$ = 1.4 and specific gas constant R = 287J/kgK. If the stagnation temperature in the combustion chamber is equal to 2310K, the speed of the exhaust gases (in m/s) is _____________.

Subtopic: Subsonic & Supersonic Inlets

Air at a stagnation temperature of 300K (ratio of specific heats, $\gamma$ = 1.4 and specific gas constant R = 287 J/kgK) enters the impeller of a centrifugal compressor in axial direction. The stagnation pressure ratio between the diffuser outlet and impeller inlet is 4.0. The impeller blade radius is 0.3 m and it is rotating at 15000 rev/min. If the slip factor $\sigma$s (Ratio of tangential component of air velocity at the blade tip to the blade tip speed) is 0.88, the overall efficiency (total-to-total) of the compressor (in %) is ___________.

Subtopic: Subsonic & Supersonic Inlets

An aircraft is flying at Mach 3.0 at an altitude where the ambient pressure and temperature are 50 kPa and 200 K respectively. If the converging-diverging diffuser of the engine (considered isentropic with ratio of specific heats, $\gamma$ = 1.4 and specific gas constant R = 287 J/kgK) has a throat area of 0.05 m^2, the mass flow rate through the engine in kg/s is

Subtopic: Subsonic & Supersonic Inlets

Consider two engines P and Q. In P, the high pressu re turbine blades are cooled with a bleed of 5% from the compressor after the compression process and in Q the turbine blades are not cooled. Comparing engine P with engine Q, which one of the following is NOT TRUE?

Subtopic: Subsonic & Supersonic Inlets

The mass flow rate of air through an aircraft e ngine is 10 kg/s. The compressor outlet temperature is 400 K and the turbine inlet temperature is 1800 K. The heating value of the fuel is 42 MJ/kg and the specific heat at constant pressure is 1 kJ/kg- K. The mass flow rate of the fuel in kg/s is approximately ___________

Subtopic: Combustor Design & Flame Stabilization

For a given inlet condition, if the turbine inlet temperature is fixed, what value of compressor efficiency given below leads to the lowest amount of fuel added in the combustor of a gas turbine engine?

Subtopic: Combustor Design & Flame Stabilization

A gas turbine engine is mounted on an aircraft which can attain a maximum altitude of 11 km from sea level. The combustor volume of this engine is decided based on conditions at (A) sea level (B) 8 km altitude (C ) 5.5 km altitude (D) 11 km altitude

Subtopic: Subsonic & Supersonic Inlets

In a 50 % reaction axial compressor stage, the local blade velocity is 300 m/s and the axial component of velocity is 100 m/s. If the absolute inlet flow angle $\alpha$1 = 45o, the work per unit mass done on the fluid by the stage in kJ/kg is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Distance of center of gravity of the aircraft ( CGX ) from the wing leading edge is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

The Headmaster ___________ to speak to you. Which of the following options is incorrect to complete the above sentence?

Subtopic: Subsonic & Supersonic Inlets

The ideal static pressure coefficient of a diffuser with an area ratio of 2.0 is

Subtopic: Subsonic & Supersonic Inlets

The stagnation temperatures at the inlet and exit of a combustion chamber are 600 K and 1200 K, respectively. If the heating value of the fuel is 44 MJ/kg and specific heat at constant pressure for air and hot gases are 1.005 kJ/kg.K and 1.147 kJ/kg.K respectively, the fuel-to-air ratio is

Subtopic: Subsonic & Supersonic Inlets

What is the tangential component of the fluid's velocity at the same radial location as above?

Subtopic: Subsonic & Supersonic Inlets

The blade angle at the inlet of the rotor is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Choose the most appropriate alternative from the options given below to complete the following sentence: Food prices ___ again this month.

Subtopic: Subsonic & Supersonic Inlets

A main objective of by-pass in a turbo-fan engine is to increase:

Subtopic: Subsonic & Supersonic Inlets

In a centrifugal compressor, the pressure rise occurs in:

Subtopic: Combustor Design & Flame Stabilization

AE 2011 AEROSPACE ENGINEERING - AE General Aptitude (GA) Questions Q. 56 - Q. 60 carry one mark each.

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

Choose the most appropriate word(s) from the opt ions given below to complete the following sentence. We lost confidence in him because he never __________ the grandiose promises he had made.

Subtopic: Subsonic & Supersonic Inlets

How does the specific thrust, al constant turbine inlet temperature, produced by a turbofan engine change with an increase in compressor pressure ratio?

Subtopic: Subsonic & Supersonic Inlets

The maximum operating flow rate through a centrifugal compressor at a given RPM is limited by

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

The lift force acting on the wing is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

If the flow through the nozzle is isentropic, what is che mass flow rate of the gases out of the nozzle?

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

In a rocket engine test, the stagnation pressure at the inlet of the convergent-divergent nozzle is measured. Given the combustion chamber pressure $p_c = 4.0\text{ MPa}$ and chamber velocity $V_c \approx 0$, the stagnation pressure $p_0$ at the nozzle inlet is:

Subtopic: Subsonic & Supersonic Inlets

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 To = 1500 K and specific heat of gas c, = 1 kJ/kg/K, the power drawn by the compressor per unit mass flow rate of air is approximately

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

When the afterburner is turned on, (assuming the total pressure remains the same, the mass of fuel added in the afterburner is negligible i.e., the mass flow rate remains the same, and the specific heat ratio (Y) remains constant), approximately by what factor must the nozzle area be changed ?

Subtopic: Subsonic & Supersonic Inlets

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:

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

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

Subtopic: Subsonic & Supersonic Inlets

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

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

If the nozzle expansion is optimum, then the thrust produced by the engine is

Subtopic: Combustor Design & Flame Stabilization

The pressure loss factor (stagnation pressure loss non-dimensionalized by the inlet dynamic pressure) of the combustor is

Subtopic: Subsonic & Supersonic Inlets

If the diffuser pressure recovery ratio is 0.92, then the isentropic efficiency of the diffuser is

Subtopic: Combustor Design & Flame Stabilization

The purpose of a fuel injection system in the combustor is

Subtopic: Convergent & Convergent-Divergent (C-D) Nozzles

The center of gravity location to have dõe dC, = 0 is