Engineering Questions with Answers - Multiple Choice Questions

# Aerodynamics – Supersonic Wind Tunnels

1 - Question

Which nozzle is used in supersonic wind tunnel?
a) Convergent
b) Divergent
c) Convergent – Divergent
d) Conical
Explanation: Supersonic wind tunnel produces flow of Mach number above 1. This is achieved using convergent – divergent nozzle. This selection is based on area – velocity relation, according to which the subsonic flow accelerates in the convergent section to sonic speed after which the speed is further accelerated in the divergent section.

2 - Question

Which of these is not a part of supersonic wind tunnel?
a) Nozzle
b) Test section
c) Diffuser
d) Test model
Explanation: Supersonic wind tunnel mainly comprises of three components – convergent – divergent nozzle, test section which is constant area duct and a diffuser to further slow down the speed to low subsonic speed. Test model is not a part of the wind tunnel and is inserted inside to take measurements such as lift, drag.

3 - Question

How many throat/throats are there in a supersonic wind tunnel?
a) 1
b) 2
c) 3
d) 4
Explanation: In case of supersonic wind tunnels, there are two throats present. One is the nozzle throat where the sonic speed is achieved and the second throat is the diffuser throat where supersonic incoming flow is isentropically compressed to Mach 1.

4 - Question

Normal shock diffuser is less efficient than the oblique shock diffuser.
a) True
b) False
Explanation: The goal of the diffuser is to reduce the flow velocity with small total pressure loss. If we reduce the incoming supersonic flow through a series of oblique shock followed by a weak normal shock wave, leads to lower total pressure loss compared to reducing the incoming supersonic flow to subsonic using one strong normal shock.

5 - Question

In reality what happens due to the interaction of the oblique shock waves with the boundary layer on the walls of the diffuser?
a) Higher total pressure loss
b) Reduction of speed
c) Total pressure increase
d) Reduced skin friction drag
Explanation: In real scenario far from hypothetical nature of the diffuser, the oblique shock waves interact with the viscous boundary layer of the diffuser wall. This leads to the separation of the boundary layer thereby increasing the total pressure losses. One thing to remember is that the aim of the diffuser is to reduce the flow velocity by low total pressure losses, but the boundary layer separation prevents that from happening.

6 - Question

The nozzle throat area is same as the diffuser throat area.
a) True
b) False
Explanation: If we consider the nozzle throat area as section 1, then the mass flow rate through the section is m1˙ = ρ1u1A1t. The mass flow rate through th diffuser throat area is m2˙ = ρ2u2A2t. Since there’s steady flow in side the wind tunnel, thus the mass flow rate remains same. m1˙=m2˙ ρ1u1A1t = ρ2 u2A1t But due to the presence of shock waves inside the diffuser, the density and the flow velocity are not same. (ρ1≠ρ2, u1≠u2). This means that the throat areas are also not same.

7 - Question

What is the relation between the nozzle A1t and diffuser throat area A2t in the supersonic wind tunnel?
a) At2 = At1
b) At2 > At1
c) At2 < At1
d) At2 At1 = 1
Explanation: The relation between the throat areas in terms of total pressure is given by: At2At1=p01p02 Since in the diffuser section, the total pressure drops across oblique shock wave, thus p02 < p01. Hence the throat area of the diffuser is larger than the throat area of the nozzle.

8 - Question

In which scenario is the throat area of both the nozzle and diffuser the same?
a) Ideal isentropic diffuser
c) Ideal isobaric diffuser
d) Ideal Isochoric diffuser
Explanation: In case of ideal isentropic diffuser the total pressure remains constant. Thus from the relation between the throat area and total pressure we see that p01 = p02 = constant. Hence the throat areas At2 = At1. It is important to note that ideal isentropic diffuser is a hypothetical case and is not possible to achieve in real life. At2At1=p01p02

9 - Question

What is the ratio of throat area of diffuser to nozzle in supersonic wind tunnel with flow at Mach 2.7?
a) 0.4236
b) 0.8338
c) 2.36
d) 1.199
Explanation: Given M = 2.7 If we assume that there’s a normal shock present at the entrance of the diffuser, then using the gas table we can find the ratio of total pressures p01p02. For M = 2.7, p02p01 = 0.4236 Using the relation beween the throat areas and total pressure we get At2At1=p01p02=10.4236 = 2.36

10 - Question

In how many categories is the supersonic wind tunnel classified into?
a) 2
b) 3
c) 4
d) 6