This set of Aircraft Design Multiple Choice Questions & Answers (MCQs) focuses on “Aerodynamic Coefficients”.

1. How do you define the lift coefficient?

a) Wing lift to weight of aircraft

b) Lift to drag

c) Ratio of aerodynamic lift to the dynamic lift

d) Thrust to weight

View Answer

Explanation: Lift coefficient is defined as aerodynamic lift divided by Dynamic lift. Dynamic lift is defined as product of dynamic pressure and reference area. Lift coefficient of airfoil and wing will be different.

2. Which of the following is correct?

a) D = q*S*CD*ρ

b) D = q*S*CD

c) D = q*CD

d) D = q

View Answer

Explanation: Above equation is a general equation of drag force. Drag on an aircraft or any other object depends on number of factors such as local dynamic pressure q, area S etc. A typical value of drag can be given as, Drag D = q*S*CD where, CD = drag coefficient.

3. If an aircraft as pitching moment of 10 Nm and dynamic pitching moment is about 8.25Nm. Find the moment coefficient Cm.

a) 1.21

b) 3

c) 5.62

d) 0.0921

View Answer

Explanation: Moment coefficient Cm = pitching moment / dynamic pitching moment

Cm = 10/8.25 = 1.21.

4. If an aircraft is operating with dynamic pressure of the free stream q=20Pa and has area of wing is 10m^{2} then evaluate drag experience by the aircraft. Given drag coefficient is 0.9.

a) 567 N

b) 345 N

c) 234 N

d) 180 N

View Answer

Explanation: Given, free stream q=20Pa, area of wing S = 10m

^{2}, drag coefficient CD = 0.9

Drag D = q*S*CD = 20*10*0.9 = 180N.

5. If Lift produced by wing is 350N then, determine lift coefficient. Given q = 35Pa and S=8.5 m^{2}.

a) 5.6

b) 2.8

c) 1.174

d) 4.37

View Answer

Explanation: Lift coefficient = lift / q*S

= 350/35*8.5 = 1.174.

6. Following diagram represents _____________

a) drag polar for non-symmetric wing

b) typical drag polar

c) wing lift curve

d) thrust required for wing

View Answer

Explanation: The above diagram is illustrating a typical schematic diagram of airfoil drag polar. Drag polar is nothing but a graph which shows variation of drag coefficient with respect to lift coefficient. Wing lift curve is used to show lift variation.

7. For a symmetrical airfoil drag coefficient at zero lift is 0.05 and induced drag coefficient is 0.0025. Find the total drag coefficient.

a) 5.25

b) 0.45

c) 0.0525

d) 52.5

View Answer

Explanation: Total drag coefficient = drag coefficient at zero lift + induced drag coefficient

= 0.05+0.0025 = 0.0525.

8. Cambered wing has minimum drag coefficient of 0.05 and constant K of 0.023. If CL is 0.8 then find the value of CD. Given minimum drag occurs at CL of 0.1.

a) 0.6721

b) 6.1

c) 61.2

d) 0.06127

View Answer

Explanation: Given, minimum drag coefficient CD

_{min}= 0.05, constant K of 0.023, CL is 0.8 and minimum drag occurs at CL of 0.1. Hence, CL

_{mindrag}= 0.1.

Now, CD is given by,

CD = CD

_{min}+ K*(CL – CL

_{mindrag})

^{2}

= 0.05+0.023*(0.8-0.1)

^{2}

= 0.06127.

9. Following diagram represents ______________

a) cambered airfoil drag polar

b) cambered wing drag polar

c) symmetric wing drag polar

d) drag polar of an airfoil

View Answer

Explanation: Above diagram is showing typical drag polar for cambered wing. Drag polar will be different for different types of wing. Drag polar is graphical representation of drag characteristics. It shows relationship between drag coefficient and lift coefficient typically.

10. A wing is designed to operate with free stream velocity of 20m/s and air density of 1.225 kg/m^{3}. Find aerodynamic efficiency of given wing. Consider S as 8 m^{2}, CL as 0.9 and CD as 1.25.

a) 0.72

b) 2

c) 3

d) 5.23

View Answer

Explanation: Given, CL = 0.9, CD = 1.25

Aerodynamic efficiency is defined as the ratio of CL and CD of the aircraft.

Hence, Aerodynamic efficiency = CL/CD = 0.9/1.25 = 0.72.

**Sanfoundry Global Education & Learning Series – Aircraft Design.**

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