List of top Propulsion Questions on Jet Propulsion

In the figure shown below, the flow at the nozzle exit is .
Which of the following process(es) is/are involved in the compression of air in an ideal ramjet engine?
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).

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?
A rocket operates at an absolute chamber pressure of 20 bar to produce thrust \(F_1\). The hot exhaust is optimally expanded to 1 bar (absolute pressure) using a convergent-divergent nozzle with exit to throat area ratio \(\left(\frac{A_e}{A_t}\right)\) of 3.5 and thrust coefficient, \(C_{F,1} = 1.42\). The same rocket when operated at an absolute chamber pressure of 50 bar produces thrust \(F_2\) and the thrust coefficient is \(C_{F,2}\). Which of the following statement(s) is/are correct?