Aircraft Design - Airfoil and Geometry Selection Quiz(MCQ)
A)
it is not a streamline body
B)
it is easy to make than rectangular
C)
it is a streamline shape which provides better aerodynamics
D)
it is a random shape can be changed to rectangular as well

Correct Answer :   it is a streamline shape which provides better aerodynamics

Explanation : Airfoil is a streamlined body which provides much smoother flow than non-streamlined body. Airfoil is designed in such a way that it can provide better aerodynamics than any other shape. Hence, in aircraft we use airfoils.

A)
B)
Fins
C)
Chord
D)
Camber

Explanation : Fins are extended surfaces which are used to improve heat transfer. Leading edge is foremost part of an airfoil. Camber is curve of an airfoil.

A)
lift*area
B)
drag*area
C)
lift/dynamic moment
D)
section pitching moment/dynamic pitching moment

Correct Answer :   section pitching moment/dynamic pitching moment

Explanation : Lift*area will give moment. Drag*area will also provide moment but will be less. The ratio of sectional pitching moment and dynamic pitching moment is called pitching moment co-efficient.

A)
0m/s
B)
16.882m/s
C)
250m/s
D)
288.16m/s

Explanation : Given, temperature T=288.16K
For idle flow, at leading edge flow is brought to rest isentropically.
Hence, velocity at leading edge = velocity of flow at the impact to leading edge = 0m/s.

A)
circle
B)
airfoil
C)
chord
D)
camber

Explanation : Wings are designed with airfoil cross-section. Engine inlets are typically circular in cross-sections. Camber and chord are parameters of an airfoil.

A)
0 N
B)
10.5 N
C)
50 N
D)
100N

Explanation : Lift is an upward force produced by pressure difference.
Given, velocity V=100m/s, AOA = 0°.
For symmetric airfoil at 0° AOA pressure will be same as upper and lower surface.
Hence, lift = 0N.

A)
28.12m
B)
0.9m
C)
0.12m
D)
0m

Explanation : NACA-4 digit series is used to define certain airfoil characteristics.
Here, velocity = 900 m/s, chord = 1m
From NACA-4 digit,
Maximum thickness = last two digits in percent of chord=12% of chord = 0.12*1 = 0.12m.

A)
0.2 m
B)
2.5 m
C)
5 m
D)
10 m

Explanation : Given, velocity = 0.2M = 0.2*340 = 68m/s
Chord = 10m
Given velocity is less than the sonic speed.
Hence, location of aerodynamic center is = 25% of chord = 0.25*10 = 2.5m from leading edge.

A)
21 Pa
B)
21c N
C)
21/c Pa
D)
21c/2 N

Explanation : Given, symmetric airfoil
Flow velocity V=350m/s, Lift L=21N, AOA = 0.008rad
Chord = c m, Span of airfoil = 1 unit
Pressure difference = Lift/Area of airfoil
= 21/c*1 = 21/c Pa.

A)
0
B)
0 N
C)
100
D)
100 N

Explanation : Given, Lift=200N, AOA = 10°
Here, airfoil is mentioned which is 2D shape.
For an airfoil Induced drag Di = 0N.

A)
transonic range
B)
subsonic range
C)
hypersonic range
D)
supersonic range

Explanation : Based on mach number flow can be subsonic, supersonic, transonic or hypersonic. If flow mach number is <0.8 then it is subsonic. If flow mach number is >0.8 but <1.2 then, it is transonic. If flow mach number is in the range of 1.2-5.0 then, it is supersonic. If flow mach number is <5.0 then, it is hypersonic.

A)
To provide lift
B)
To provide lift co-efficient
C)
To provide streamlined flow
D)
To provide thrust for acceleration

Correct Answer :   To provide thrust for acceleration

Explanation : In an aircraft, engines are used to provide thrust for acceleration. Airfoil is used to provide lift, to improve certain flow characteristics. Streamlined flow is also produced by an airfoil.

13 .
What is represented by ‘?’ in following diagram of an airfoil?
A)
Thickness
B)
Chord
C)
Trailing edge
D)

Explaination : Thickness is distance between upper and lower surface of the airfoil. Leading and trailing edge are foremost and most rear parts of airfoil. Thickness will be different at each section of an airfoil.

A)
National Advisory Committee for Aero models
B)
C)
D)
National Authorized Committee for Aeronautics

Explanation : NACA is a National Advisory Committee for Aeronautics. NACA airfoils are simple to adopt. NACA series is used to identify the airfoil.

A)
0m
B)
0.01m
C)
0.02m
D)
0.021

Explanation : Given, NACA 4 digit airfoil as NACA 1235.
Chord c=1m.
Based on NACA 4 digit, first digit is the maximum camber in hundreds of chord.
Hence, maximum camber for NACA 1235 = 0.01*chord = 0.01*1 = 0.01m.

16 .
A thin airfoil is defined as NACA 0009 is at ? = 5°. What is the value of lift curve slope?
A)
0.4 per degree
B)
C)
2? per degree
D)
4 per degree

Explaination : Given, NACA0009 airfoil,
AOA ? = 5°
As mentioned the airfoil is thin. Hence, thin airfoil theory can be used.
According to thin airfoil theory, for thin airfoil,
Lift curve slope = 2π 1/rad = 0.11 1/degree.

A)
B)
camber
C)
camber line
D)
chord line

Explanation : Chord line is defined as the straight line which connects the leading and trailing edge of an airfoil. Camber line or mean camber line is line passing from locus of mid points. Almost every parameter is defined with respect to chord.

A)
lower than lower surface
B)
always same
C)
higher than lower surface
D)
always 30.25 times lower surface

Correct Answer :   lower than lower surface

Explanation : Pressure on the upper surface of an airfoil is lower than that of the lower surface. This will provide pressure difference which will generate lift force. Pressure will be the same if AOA is 0° in case of symmetric airfoil.

A)
turbulent is produced
B)
lift is reduced drastically
C)
flow remains laminar
D)
drag increases drastically

Correct Answer :   flow remains laminar

Explanation : Laminar bucket is phenomena of laminar airfoils. It is the region where flow remains laminar. This laminar flow is very important as it decreases drag and thus improves overall performance. However, for such airfoils Reynolds number is very important consideration. Following diagram shows laminar bucket for a typical laminar airfoil.

A)
NACA 0009
B)
NACA 6512
C)
NACA 65211
D)
NACA 652132

Explanation : NACA series is used to identify airfoils as per our requirement. NACA -4 and -5 digit series are not used to provide a minimum pressure point. NACA 6 digit series is used for such operations.

A)
at half of chord
B)
at quarter chord point
C)
there is no such place
D)
at quarter chord from leading edge

Correct Answer :   at quarter chord point

Explanation : Aerodynamic center is that point at which pitching moment curve slope becomes zero. Hence, it is the point at where pitching moment is almost independent of AOA. For, low speed airfoils it is located at 25% of chord from leading edge.

22 .
A NACA 22104 airfoil is operating at AOA = 3°. The lift-coefficient at this AOA is 0.15. Determine the value of angle of attack at designed lift co-efficient.
A)
4.36°
B)
C)
0.05°
D)

Explaination : Given NACA 22104 airfoil.
AOA = 3°, Cl = 0.15 at 3°
Based on NACA -5 digit series design Cl = first digit*0.15 = 2*0.15 = 0.3
Now, for thin airfoil lift curve slope = 0.11 per degree
0.11 = (design Cl – Cl at 3°) / (AOA –3°)
0.11 = (0.3-0.15) / (AOA -3°)
0.11*(AOA -3°) = 0.15
Hence, at designed Cl, AOA = (0.15/0.11) + 3° = 4.36°.

A)
0.15
B)
0.15N
C)
0.3N
D)
0.315

Explanation : Given, NACA 13250 airfoil
Based on NACA -5 digit series, designed lift co-efficient = first digit*0.15 = 1*0.15 = 0.15.

A)
B)
0.32 from trailing edge
C)
0.32m from trailing edge
D)

Explanation : Given, NACA 13250 airfoil
Based on NACA -5 digit series, location of maximum camber = (second and third digit)/2 in hundredth of chord from leading edge
= (32/2) * 0.01 * chord = (32/2)*0.01*2 = 0.32m from leading edge.

A)
Bluff shape
B)
Thicker airfoil
C)
Supersonic airfoil
D)
Symmetric airfoil

Explanation : Supersonic aircraft needs to fly at much higher speed. Hence, supersonic airfoils are used to provide required velocity. Thicker airfoils cannot generate required amount of flow characteristics such as critical mach number.

A)
0
B)
0N
C)
0.5
D)
1

Explanation : Given, Lift=200N, velocity = 500m/s.
For cambered airfoil the downwash angle = 0°. Hence, there will be no downwash acting on the flow.
Hence, there will be no induced drag. Hence, for any airfoil induced drag = 0N.

A)
airfoil
B)
wing
C)
engine
D)
lofting

Explanation : Wing is generated by elongating an airfoil. Airfoil is 2D shape. Engine is main source of power. Lofting is skin modelling.

A)
120 N
B)
574.18 N
C)
574.18
D)
620 N

Explanation : Given, a twin turboprop
Velocity V = 120m/s, CL = 0.15, Aspect ratio = 9.2, span b=2m
Aspect ratio = b*b/S
9.2 = 2*2/s
S = 0.434m2
Now, lift is given by,
L = (1/2)*ρ*V2*S*CL = 0.5* 1.225*120*120*0.434*0.15 = 574.18N.

29 .
Which type of wing planform is shown below?
A)
Delta wing
B)
Elliptic wing
C)
Rectangular wing
D)
Tapered wing

Explaination : Tapered wings have finite value of taper ratio. In taper wing root and tip chord will have different values of chord. Elliptic wings are of elliptic shape. Rectangular wing has taper ratio of unity which shows a rectangular profile.

A)
0.4m
B)
0.4
C)
0.8m
D)
0.8

Explanation : Given, jet fighter aircraft which is cruising at Mach 1.4.
For such high speed aircraft, location of the aerodynamic center is given by,
Location of aerodynamic center = 40% of chord = 40% of 1 = 0.4*1 = 0.4m.

A)
same as military aircraft
B)
lower than military aircraft
C)
higher than military aircraft
D)
exactly half of the military aircraft

Correct Answer :   higher than military aircraft

Explanation : Sailplanes are different from military aircrafts. Sailplanes are designed to glide through air. For such requirements, it will require higher lift capability which can be provided by using higher aspect ratio wings than military aircrafts.

A)
0.4761m
B)
0.61m
C)
0.71cm
D)
0.81m

Explanation : Given, span = 2m, aspect ratio = 7 and taper ratio T = 0.2
The root chord is given by,
Croot = 2*S / (b*(1+T))
Here, S is not given but aspect ratio is given.
Aspect ratio AR = Span square / S = 2*2 / S
7 = 4/ S
S = 0.5714 m2
Hence, Croot = 2*S / (b*(1+T))
= 2*0.5714 / (2*(1+0.2)) = 0.4761m.

A)
span square
B)
span square plus reference area
C)
span square divided by reference area
D)
span square multiplied by reference area

Correct Answer :   span square divided by reference area

Explanation : Aspect ratio of wing gives relation between span of wing and wing reference area. It is defined as the ratio of span square to the reference or planform area. Higher aspect ratio wings will be long with less area.

A)
sweep only
B)
taper ratio only
C)
aspect ratio only
D)
aspect ratio, sweep angle, taper ratio

Correct Answer :   aspect ratio, sweep angle, taper ratio

Explanation : Wing planform is shape of the wing when viewed from top. Aspect ratio will affect span and area. Overall planform will be affected by aspect ratio, taper ratio, sweep etc.

A)
True
B)
False
C)
Can Not Say
D)
None of the above

Explanation : Wing tip of high aspect ratio wing is far from root as compared to low aspect ratio wing. The wing affected by higher ratio will be less and the strength of tip vortex will be less as well.

A)
0.89
B)
0.89m
C)
0.89cm
D)
0.89inch

Explanation : Given, wing Aspect ratio AR=8, S=0.1m2
Aspect ratio = span square/ reference area
8 = span square / 0.1
Span square = 8*0.1 = 0.8
Hence, span = 0.89m.

A)
pitch
B)
dutch roll
C)
thrust
D)
deflection at nose

Explanation : Dihedral is upward deflection of wing. It is used to provide roll stability. If excessive dihedral is provided then, it will lead to dutch roll; continuous side to side motion involving yaw and roll.

A)
To provide lofting
B)
To reduce lift by tail
C)
To improve stall characteristics at tip
D)
To increase engine thrust

Correct Answer :   To improve stall characteristics at tip

Explanation : Wing twist is used to delay the stall at tip typically. Wing twist can be used to rearrange the lift distribution of the wing. Lofting is skin modelling.

A)
0.25
B)
0.45
C)
0.55
D)
0.65

Explanation : Taper ratio is defined as,
Taper ratio = tip chord/root chord = 0.9/2 = 0.45.

A)
slope
B)
sweep
C)
aspect ratio
D)
taper ratio

Explanation : Taper ratio is the ratio of tip chord of wing to root chord of the wing. Aspect ratio is square of span divided by area. Sweep is angle between wing and fuselage reference line.

A)
increase critical mach number
B)
increase lofting
C)
increase drafting
D)
decrease critical mach number

Correct Answer :   increase critical mach number

Explanation : If wing is at some finite angle from fuselage reference line then it is called wing is sweep by that much degree. Main function of sweep is to increase critical mach number. Lofting is skin modelling and drafting is drawing phenomena.

42 .
Which of the following is correct in accordance to diagram?
A)
Aspect ratio of 1 is highest
B)
Aspect ratio of 2 is maximum
C)
Aspect ratio of 3 is maximum
D)
Similar aspect ratio

Correct Answer :   Aspect ratio of 1 is highest

Explaination : High aspect ratio will give higher value of maximum lift coefficient. As shown in figure for -1, maximum lift coefficient is highest and hence, aspect ratio of 1 will be highest. Similarly for -3 it is lowest and hence, aspect ratio of 3 will be minimum.

43 .
An aircraft with elliptic wing planform has parasite drag coefficient as 0.6. Lift coefficient of wing is 0.25 and aspect ratio is 7.5. If induced drag co-efficient is 0.0235 then, find total drag coefficient for the wing.
A)
0.0235
B)
0.6
C)
0.5235
D)
0.06235

Explaination : Given, parasite drag co-efficient = 0.6, induced drag coefficient = 0.0235
Total drag coefficient = Parasitic drag coefficient + induced drag coefficient = 0.6+0.0235 = 0.06235.

A)
stability
B)
control
C)
stability and control
D)
accelerating force

Correct Answer :   stability and control

Explanation : Typically tail is used to provide stability and control. Accelerating force will be provided from engines. One of the main function of the tail is to provide enough stability and controllability to make aircraft fly smoother at the time of any disturbance.

45 .
Which tail profile is shown in below diagram?
A)
T – tail
B)
H – tail
C)
Cruciform tail
D)
Conventional tail

Explaination : Cruciform tail viewed from front will look like a ‘plus’ as shown in figure. T – tail will look like letter T and similarly H tail has shape of letter H.

A)
Provides lift to fly
B)
Similar to horizontal tail
C)
To reduce tail-effectiveness at high slide slip angle
D)
To improve tail-effectiveness at high slide slip angle

Correct Answer :   To improve tail-effectiveness at high slide slip angle

Explanation : Dorsal fins are small attachments at the vertical tail. Dorsal fins are used for improving tail effectiveness at high slide slip angle. Dorsal fins are not similar to horizontal tail. Lift is primarily provided by wings.

A)
Flap
B)
Rudder
C)
Aileron
D)
Elevator

Explanation : During spin aircrafts is not stable. Aircraft performs unstable yaw motion. For stabilizing such event adequate amount of rudder control is required. Flaps are used to increase lift and drag.

A)
roll
B)
yawing
C)
yaw-roll
D)
pitching moment

Explanation : Elevator is one of the primary control surfaces of an aircraft. The elevator is located at the horizontal tail. Elevator is responsible for pitching motion to provide required nose up or down moment.

A)
To produce thrust
B)
To increase wetted area
C)
To reduce wetted area
D)

Correct Answer :   To reduce wetted area

Explanation : V tail arrangement is used to combine horizontal and vertical tail. It combines in such way that the wetted area reduces. However, it generates adverse yaw roll motion. To prevent such effect inverted V tail can be used.

A)
B)
to increase lift
C)
to increase thrust
D)

Explanation : Inverted V tail arrangement is used to prevent the adverse yaw roll motion. Lift can be increased by flaps. Thrust is increased by an engine.

A)
Aileron
B)
Rudder
C)
Elevator
D)
Horizontal stabilizer