List of top Propulsion Questions on Jet Propulsion asked in GATE Aerospace Engineering

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).

A monopropellant liquid rocket engine has 800 injectors of diameter 4 mm each, and with a discharge coefficient of 0.65. The liquid propellant of density 1000 kg/m³ flows through the injectors. There is a pressure difference of 10 bar across the injectors. The specific impulse of the rocket is 1500 m/s. The thrust generated by the rocket is __________ kN (rounded off to one decimal place).

A monopropellant liquid rocket engine has 800 injectors of diameter 4 mm each, and with a discharge coefficient of 0.65. The liquid propellant of density 1000 kg/m³ flows through the injectors. There is a pressure difference of 10 bar across the injectors. The specific impulse of the rocket is 1500 m/s. The thrust generated by the rocket is \_\_\_\_\_ kN (rounded off to one decimal place).
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).
A monopropellant liquid rocket engine has 800 injectors of diameter 4 mm each, and with a discharge coefficient of 0.65. The liquid propellant of density 1000 kg/m³ flows through the injectors. There is a pressure difference of 10 bar across the injectors. The specific impulse of the rocket is 1500 m/s. The thrust generated by the rocket is \_\_\_\_\_ kN (rounded off to one decimal place).
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).
A monopropellant liquid rocket engine has 800 injectors of diameter 4 mm each, and with a discharge coefficient of 0.65. The liquid propellant of density 1000 kg/m³ flows through the injectors. There is a pressure difference of 10 bar across the injectors. The specific impulse of the rocket is 1500 m/s. The thrust generated by the rocket is \_\_\_\_\_ kN (rounded off to one decimal place).
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).

Which one of the following figures represents the drag polar of a general aviation aircraft? (Axes: vertical \(C_L\), horizontal \(C_D\)).

In an airbreathing gas turbine engine, the combustor inlet temperature is \(600\ \text{K}\). The heating value of the fuel is \(43.4\times 10^{6}\ \text{J/kg}\). Assume\(c_p = 1100\ \mathrm{J\,kg^{-1}\,K^{-1}}\) for air and burned gases, burner efficiency \(=100\%\), and fuel-air ratio \(f\ll 1\). Neglect kinetic energy at inlet/exit. What fuel-air ratio is required to achieve \(1300\ \text{K}\) at the combustor exit?}

A turbojet–powered aircraft flies at \(V_0=250\ \text{m s}^{-1}\). The (uninstalled) thrust is \(T=60{,}000\ \text{N}\). Fuel heating value is \(44\times 10^{6}\ \text{J kg}^{-1}\); fuel flow is \(\dot m_f=3\ \text{kg s}^{-1}\). The engine thermal efficiency is \(35\%\). Exit pressure equals ambient. What is the propulsion efficiency (in %)?

Which of the following statements is/are TRUE for an axial turbine?
A chemical rocket with an ideally expanded flow through the nozzle produces $5 \times 10^6$ N thrust at sea level. The specific impulse of the rocket is $200$ s and acceleration due to gravity at sea level is $9.8$ m/s$^2$. The propellant mass flow rate out of the rocket nozzle is .............. kg/s (rounded off to the nearest integer).
A centrifugal compressor is designed to operate with air. At the leading edge of the tip of the inducer, the blade angle is $45^\circ$, and the relative Mach number is 1.0. The stagnation temperature of the incoming air is 300 K. Consider $\gamma = 1.4$. Neglect pre-whirl and slip. The inducer tip speed is .............. m/s (rounded off to the nearest integer).

A scramjet engine features an intake, isolator, combustor, and a nozzle, as shown. Station 3 indicates the combustor entry point. Assume stagnation enthalpy is constant between Stations 1 and 3, and air is a calorically perfect gas with specific heat ratio $\gamma$. Select the correct expression for Mach number $M_3$ at the inlet to the combustor from the options given.

A supersonic vehicle powered by a ramjet engine is cruising at \(1000~\mathrm{m/s}\). The ramjet engine burns hydrogen in a subsonic combustor to produce thrust. The heat of combustion of hydrogen is \(120~\mathrm{MJ/kg}\). The overall efficiency of the engine \(\eta_0\), defined as the ratio of propulsive power to the total heat release in the combustor, is 40%. Taking \(g_0=10~\mathrm{m/s^2}\), the specific impulse of the engine is \underline{\hspace{1cm}} seconds. \;(round off to nearest integer)
Two missiles $A$ and $B$ powered by solid rocket motors have identical specific impulse, liftoff mass of $5600 \,\text{kg}$ each, and burn durations of $t_A = 30 \,\text{s}$ and $t_B = 70 \,\text{s}$, respectively. The propellant mass flow rates $\dot{m}_A$ and $\dot{m}_B$ are given as: \[ \dot{m}_A = 120 \,\text{kg/s}, \, 0 \leq t \leq 30, \dot{m}_B = 70 \,\text{kg/s}, \, 0 \leq t \leq 70 \] Neglecting gravity and aerodynamic forces, the relation between final velocities $V_A$ and $V_B$ is:
An aircraft is cruising with a forward speed \(V_a\) and the jet exhaust speed relative to the engine at the exit is \(V_j\). If \(V_j/V_a=2\), what is the propulsive efficiency?
In a single–spool aviation turbojet engine, which of the following is the correct relationship between the total work output \(W_t\) of a 2-stage axial turbine and the total work required \(W_c\) by a 6-stage axial compressor, neglecting losses?