Instruction
- Total number of questions : 20
- Time allotted : 30:00 Minutes:Seconds.
- Each question carry 1 mark, no negative marks.
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Aerodynamics [ Question No: 1 To 20 ] :
Assertion (A): Blockage corrections grow rapidly as the model is made larger relative to the test section.
Reason (R): A larger model leaves less area for the flow, forcing the surrounding air to accelerate.
Select the correct answer from the options below.
Assertion (A): Cryogenic tunnels raise the Reynolds number without altering the Mach number.
Reason (R): Cooling the gas raises its density and lowers its viscosity together.
Select the correct answer from the options below.
Consider the following statements about wake surveys:
1. They determine drag from the momentum deficit behind the body.
2. They require no instrumentation on the body surface.
3. They can distinguish skin friction from pressure drag directly.
Which of the statements given above is or are correct?
Riblets, the fine streamwise grooves trialled on transport aircraft, reduce drag by
Which of the following is NOT a recognised approach to reducing the sonic boom reaching the ground?
Which statement about the critical and drag-divergence Mach numbers is TRUE?
Doubling the incidence of a supersonic wing at fixed Mach number changes its wave drag due to lift by a factor of
The wave drag of a supersonic wing is conventionally divided into contributions due to
The Mach angle in a stream at \( M = 1.5 \) is nearest to
Consider the following statements about a shock in the divergent section of a nozzle:
1. The flow is supersonic immediately upstream of the shock.
2. The flow is subsonic immediately downstream of the shock.
3. The stagnation pressure downstream of the shock equals that in the reservoir.
Which of the statements given above is or are correct?
Match each quantity in List I with its behaviour across a normal shock in List II.
List I
A. Static pressure
B. Mach number
C. Stagnation temperature
D. Stagnation pressure
List II
1. Falls below one
2. Rises
3. Falls
4. Unchanged
Select the correct match.
A normal shock stands in a stream at \( M_1 = 2 \). The downstream Mach number is nearest to
Which of the following is NOT a valid reason for a supersonic nozzle to diverge downstream of its throat?
Assertion (A): The critical Mach number of an aerofoil is lower than its drag-divergence Mach number.
Reason (R): Sonic flow first appears at the point of lowest pressure on the surface.
Select the correct answer from the options below.
A flow is described as transonic when
Consider the following statements about the effect of camber on a thin section:
1. It makes the zero-lift angle of attack more negative.
2. It introduces a nose-down moment about the quarter chord when the camber is convex upward.
3. It increases the lift curve slope above two pi per radian.
Which of the statements given above is or are correct?
Assertion (A): A symmetric thin airfoil has zero pitching moment about its quarter chord at every incidence.
Reason (R): A symmetric section has a straight camber line, so the camber-dependent coefficients vanish.
Select the correct answer from the options below.
An aircraft has \( C_{D_0} = 0.020 \) and flies at \( C_L = 0.7 \) with \( AR = 8 \) and \( e = 0.9 \). Its total drag coefficient is nearest to
A trailing edge stall is characterised by separation that
Flow is described as transonic when
Total number of questions : 20
Number of answered questions : 0
Number of unanswered questions : 20
Aerodynamics [ Question No: 1 To 20 ] :
Assertion (A): Blockage corrections grow rapidly as the model is made larger relative to the test section.
Reason (R): A larger model leaves less area for the flow, forcing the surrounding air to accelerate.
Select the correct answer from the options below.
Your Answer : (Not Answered)
Correct Answer : D
Both are true and the second explains the first. Continuity in a confined section means the measured forces come out too high. Keeping frontal area below about five per cent of the section keeps the correction a modest adjustment rather than a dominant one.
Assertion (A): Cryogenic tunnels raise the Reynolds number without altering the Mach number.
Reason (R): Cooling the gas raises its density and lowers its viscosity together.
Select the correct answer from the options below.
Your Answer : (Not Answered)
Correct Answer : B
Both statements are true and the second is the reason for the first. Both effects push the Reynolds number the same way, giving roughly a sixfold gain at 110 K, while the Mach number depends on speed relative to the local speed of sound and is set independently.
Consider the following statements about wake surveys:
1. They determine drag from the momentum deficit behind the body.
2. They require no instrumentation on the body surface.
3. They can distinguish skin friction from pressure drag directly.
Which of the statements given above is or are correct?
Your Answer : (Not Answered)
Correct Answer : A
The first two are correct and are why the technique is so convenient for section testing. The third is false: the wake carries the total momentum deficit and cannot by itself say how much came from shear at the surface and how much from the pressure distribution.
Riblets, the fine streamwise grooves trialled on transport aircraft, reduce drag by
Your Answer : (Not Answered)
Correct Answer : C
Grooves of the right size impede the lateral movement of the streamwise vortices that dominate the near-wall region, cutting turbulent momentum transport and with it the shear stress. Reductions of a few per cent have been measured, though keeping the surface clean and intact in service is the obstacle.
Which of the following is NOT a recognised approach to reducing the sonic boom reaching the ground?
Your Answer : (Not Answered)
Correct Answer : D
A thicker wing adds volume abruptly and strengthens the signature, which is the opposite of what low-boom design seeks. The other three all spread the disturbance over a greater length or a greater propagation distance and soften the wave that arrives.
Which statement about the critical and drag-divergence Mach numbers is TRUE?
Your Answer : (Not Answered)
Correct Answer : B
Sonic flow appears somewhere on the surface at the critical Mach number, but the drag does not rise sharply until the resulting supersonic pocket has grown enough for a strong shock to close it, which happens a little faster. The band between the two is where supercritical sections are designed to work.
Doubling the incidence of a supersonic wing at fixed Mach number changes its wave drag due to lift by a factor of
Your Answer : (Not Answered)
Correct Answer : C
Wave drag due to lift varies with the square of the lift coefficient, and by linearised theory the lift coefficient is proportional to incidence, so doubling the angle multiplies that drag component by four. The volume-dependent part is unaffected, since it depends on thickness alone.
The wave drag of a supersonic wing is conventionally divided into contributions due to
Your Answer : (Not Answered)
Correct Answer : A
Linearised supersonic theory separates the drag into a part depending on the thickness distribution, the volume-dependent wave drag, and a part depending on the loading, the wave drag due to lift. Both are pressure drags present in inviscid flow, and both are quite distinct from induced drag and skin friction.
The Mach angle in a stream at \( M = 1.5 \) is nearest to
Your Answer : (Not Answered)
Correct Answer : B
Correct answer: 41.8 degrees
\[ \mu = \arcsin\left(\frac{1}{M}\right) = \arcsin\left(\frac{1}{1.5}\right) = \arcsin(0.667) = 41.8^{\circ}. \] The angle falls as the Mach number rises - 90 degrees at \( M=1 \), where the wave is normal to the flow, and approaching zero at high Mach number, where the Mach cone closes down about the flight path. The 48.2 offered is its complement.
Consider the following statements about a shock in the divergent section of a nozzle:
1. The flow is supersonic immediately upstream of the shock.
2. The flow is subsonic immediately downstream of the shock.
3. The stagnation pressure downstream of the shock equals that in the reservoir.
Which of the statements given above is or are correct?
Your Answer : (Not Answered)
Correct Answer : B
The first two are correct and follow from the shock being normal. The third is false: the shock destroys stagnation pressure, so the downstream value is below the reservoir value, and it is that deficit which determines where the shock has to sit for a given back pressure.
Match each quantity in List I with its behaviour across a normal shock in List II.
List I
A. Static pressure
B. Mach number
C. Stagnation temperature
D. Stagnation pressure
List II
1. Falls below one
2. Rises
3. Falls
4. Unchanged
Select the correct match.
Your Answer : (Not Answered)
Correct Answer : B
Static pressure jumps up, the Mach number is always taken below one, stagnation temperature survives because the shock is adiabatic, and stagnation pressure falls because it is irreversible. Keeping these four straight is most of what is needed to answer any normal shock question.
A normal shock stands in a stream at \( M_1 = 2 \). The downstream Mach number is nearest to
Your Answer : (Not Answered)
Correct Answer : B
The relation is \( M_2^2 = (1 + 0.2 M_1^2)/(1.4 M_1^2 - 0.2) \). Substituting, the numerator is \( 1 + 0.8 = 1.8 \) and the denominator \( 5.6 - 0.2 = 5.4 \), so \( M_2^2 = 0.3333 \) and \( M_2 = 0.577 \). Every normal shock leaves the flow subsonic, so any answer above one would be wrong on principle.
Which of the following is NOT a valid reason for a supersonic nozzle to diverge downstream of its throat?
Your Answer : (Not Answered)
Correct Answer : A
Static pressure falls, not rises, as a supersonic flow accelerates through the divergent section; the option inverts the physics. The other three are all correct and together make up the standard derivation of the area-velocity relation.
Assertion (A): The critical Mach number of an aerofoil is lower than its drag-divergence Mach number.
Reason (R): Sonic flow first appears at the point of lowest pressure on the surface.
Select the correct answer from the options below.
Your Answer : (Not Answered)
Correct Answer : A
Both statements are true, but the second is not the reason for the first. The critical value marks where sonic flow first appears anywhere on the surface; drag does not rise sharply until the supersonic pocket has grown large enough to end in a strong shock, which happens a little later. Where the sonic point first appears is a separate matter.
A flow is described as transonic when
Your Answer : (Not Answered)
Correct Answer : B
Transonic flow is the awkward mixed regime, typically between about Mach 0.8 and 1.2 in the free stream, in which the flow accelerates past sonic speed somewhere on the body while remaining subsonic elsewhere. The mixed character is what makes it so hard to compute and why shock-induced separation appears there.
Consider the following statements about the effect of camber on a thin section:
1. It makes the zero-lift angle of attack more negative.
2. It introduces a nose-down moment about the quarter chord when the camber is convex upward.
3. It increases the lift curve slope above two pi per radian.
Which of the statements given above is or are correct?
Your Answer : (Not Answered)
Correct Answer : A
The first two are correct and are the whole of what camber does in this theory. The third is false: the slope stays at two pi per radian whatever the camber, because camber enters only the constant term of the solution and never the term that varies with incidence.
Assertion (A): A symmetric thin airfoil has zero pitching moment about its quarter chord at every incidence.
Reason (R): A symmetric section has a straight camber line, so the camber-dependent coefficients vanish.
Select the correct answer from the options below.
Your Answer : (Not Answered)
Correct Answer : C
Both are true and the second explains the first. The quarter-chord moment depends only on the first two camber coefficients, and a straight camber line makes both of them zero. The centre of pressure of a symmetric section therefore sits at the quarter chord at every angle of attack.
An aircraft has \( C_{D_0} = 0.020 \) and flies at \( C_L = 0.7 \) with \( AR = 8 \) and \( e = 0.9 \). Its total drag coefficient is nearest to
Your Answer : (Not Answered)
Correct Answer : D
The induced part is \( 0.7^2 / (\pi \times 0.9 \times 8) = 0.49 / 22.62 = 0.0217 \). Adding the zero-lift part gives \( 0.020 + 0.0217 = 0.0417 \). Note that at this lift coefficient the induced drag slightly exceeds the zero-lift drag, so the aircraft is flying a little below its minimum-drag speed.
A trailing edge stall is characterised by separation that
Your Answer : (Not Answered)
Correct Answer : A
Correct answer: Begins at the rear and moves forward as incidence increases
On a thick, well-rounded section the adverse pressure gradient first defeats the boundary layer near the trailing edge, and the separation point creeps forward as incidence rises. The lift curve therefore rounds over gradually and the stall is gentle and preceded by buffet - the behaviour a designer wants, because it gives the pilot warning.
Flow is described as transonic when
Your Answer : (Not Answered)
Correct Answer : A
Near a free-stream Mach number of one, local acceleration over a wing can carry part of the flow past the speed of sound while the rest remains subsonic, and shock waves appear where it returns. This mixed condition is the hardest regime to analyse or to design for.
Marks
Total number of questions : 20
Number of answered questions : 0
Number of unanswered questions : 20