List of top Aeronautical and Aerospace Engineering Questions asked in GATE Aerospace Engineering

A general aviation airplane is in steady and level flight. The airplane is prone to adverse yaw. Which one of the following options best describes the deflections of aileron and rudder to achieve a coordinated right turn?

In steady, inviscid, incompressible flow, superpose a uniform stream of speed $U$ along $+x$ (left $\to$ right) with a source of strength $\Lambda$ at the origin. Which statement about the stagnation point's location is NOT true?

Day 1: speed $V_1$ at temperature $T_1$ K. Day 2: speed $\sqrt{1.2}\,V_1$ at temperature $1.2\,T_1$ K. For air treated as an ideal gas with the same $\gamma$ and $R$ both days, compare Mach numbers $M_2$ (Day 2) and $M_1$ (Day 1).

Steady, incompressible, inviscid flow past two airfoils is shown. The pressure coefficient at the trailing edge of (I) with a finite included angle is $C_{pI}$; for (II) with a cusp (zero included angle) it is $C_{pII}$. Which statement is TRUE?

As per the International Standard Atmosphere (ISA), which statement about density variation with altitude in an isothermal layer is correct?

The figure shows a control volume to estimate drag on an airfoil with elliptic cross-section. Surfaces 2 and 3 are streamlines. Velocity is measured at upstream (surface 1) and downstream (surface 4). The drag coefficient is $C_d = \tfrac{D{\tfrac{1}{2} \rho U_\infty^2 c}$, where $D$ is drag force per unit span, $\rho$ is density, $U_\infty$ is free-stream velocity, and $c$ is chord. The static pressure is constant over the control surface. Flow is incompressible, 2D, steady. Find $C_d$ (rounded to 3 decimals).} \begin{center} \includegraphics[width=0.7\textwidth]{18.jpeg} \end{center}

Consider the following Fanno flow problem: Flow enters a constant area duct at a temperature of 273 K and a Mach number 0.2 and eventually reaches sonic condition (Mach number = 1) due to friction. Assume $\gamma = 1.4$. The static temperature at the location where sonic condition is reached is ................. K (rounded off to 2 decimal places).

Using thin airfoil theory, find the lift coefficient of a NACA 0012 airfoil at $\alpha=5^\circ$ in uniform flow (rounded to 2 decimal places).

Consider the International Standard Atmosphere (ISA) with $h$ being the geopotential altitude (in km) and $\dfrac{dT}{dh}$ being the temperature gradient (in K/m). Which of the following combination(s) of $\Big(h, \dfrac{dT}{dh}\Big)$ is/are correct as per ISA?

The operating characteristics of a pump are measured as $C_p = a \Phi^2$, where \[ C_p = \frac{P}{\rho \omega^3 D^5}, \Phi = \text{flow coefficient}, a=\text{constant}. \] If $\omega$ increases by 25% (i.e. $\omega \to 1.25\omega$) and the flow coefficient $\Phi$ is constant, determine $\alpha$ such that $P$ becomes $\alpha P$. (round off to two decimal places)

In potential flow, a uniform stream of strength $U$ flows along x-axis. Line sources of strength $\pi/2, -\pi/3, \pi/4, -\pi/5$ are placed at $x=0,1,2,3$ respectively. Find strength of an additional line source at $x=4$ such that a closed streamline encircles all five sources. (round off to two decimal places)

Enstrophy is defined as square of magnitude of vorticity. For velocity field \[ \vec{V} = (4x - 1.5y + 2.5z)\hat{i} + (1.5x - 1.5y)\hat{j} + (0.7xy)\hat{k}, \] find enstrophy at $(1,1,1)$. (round off to two decimal places)

A thin cambered airfoil has lift coefficient $C_l=0$ at angle of attack $\alpha=-1^\circ$. Estimate $C_l$ at $\alpha=4^\circ$, assuming stall occurs at much higher $\alpha$. (round off to two decimal places)

A general aviation airplane has $W=10 \,\text{kN}$, $S=15 \,\text{m}^2$, $\rho=0.60 \,\text{kg/m}^3$, $C_{D0}=0.025$, $K=0.05$, thrust $T=1 \,\text{kN}$. Find the maximum cruise speed.

Airplane A and Airplane B are cruising at altitudes of $2~\mathrm{km}$ and $4~\mathrm{km}$, respectively. The free–stream density and static pressure at $2~\mathrm{km}$ are $1.01~\mathrm{kg/m^3}$ and $79.50~\mathrm{kPa}$; at $4~\mathrm{km}$ they are $0.82~\mathrm{kg/m^3}$ and $61.70~\mathrm{kPa}$. The differential pressure reading from the pitot–static tubes is $3~\mathrm{kPa}$ for both airplanes. Assuming incompressible flow, the ratio of cruise speeds $V_A/V_B$ is \underline{\hspace{1cm}}. \;(round off to two decimal places)

In 2D potential flow, the doublet is the limit of the superposition of which two singularities?

An ideal glider has drag characteristics given by \[ C_D = C_{D0} + C_{Di}, \] where $C_{Di} = K C_L^2$ is the induced drag coefficient, $C_L$ is the lift coefficient, and $K$ is a constant. For maximum range of the glider, the ratio $\dfrac{C_{D0}}{C_{Di}}$ is:

The figures shown in the options are schematics of airfoil shapes (not to scale). For a civilian transport aircraft designed for a cruise Mach number of $0.8$, which among them is aerodynamically best suited as a wing section?

Consider the Blasius solution for the incompressible laminar flat–plate boundary layer. Among the options, select the correct relation for the development of the momentum thickness $\theta$ with distance $x$ from the leading edge along the plate.