Aircraft Design MCQ – Aerodynamic Coefficients
1 - Question
How do you define the lift coefficient?
a) Ratio of aerodynamic lift to the dynamic lift
b) Lift to drag
c) Wing lift to weight of aircraft
d) Thrust to weight
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 - Question
Which of the following is correct?
a) D = q*S*CD
b) D = q*S*CD*ρ
c) D = q*CD
d) D = q
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 - Question
If an aircraft as pitching moment of 10 Nm and dynamic pitching moment is about 8.25Nm. Find the moment coefficient Cm.
Explanation: Moment coefficient Cm = pitching moment / dynamic pitching moment Cm = 10/8.25 = 1.21.
4 - Question
If an aircraft is operating with dynamic pressure of the free stream q=20Pa and has area of wing is 10m2 then evaluate drag experience by the aircraft. Given drag coefficient is 0.9.
a) 180 N
b) 345 N
c) 234 N
d) 567 N
Explanation: Given, free stream q=20Pa, area of wing S = 10m2, drag coefficient CD = 0.9 Drag D = q*S*CD = 20*10*0.9 = 180N.
5 - Question
If Lift produced by wing is 350N then, determine lift coefficient. Given q = 35Pa and S=8.5 m2.
Explanation: Lift coefficient = lift / q*S = 350/35*8.5 = 1.174.
6 - Question
Following diagram represents _____________
a) typical drag polar
b) drag polar for non-symmetric wing
c) wing lift curve
d) thrust required for wing
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 - Question
For a symmetrical airfoil drag coefficient at zero lift is 0.05 and induced drag coefficient is 0.0025. Find the total drag coefficient.
Explanation: Total drag coefficient = drag coefficient at zero lift + induced drag coefficient = 0.05+0.0025 = 0.0525.
8 - Question
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.
Explanation: Given, minimum drag coefficient CDmin = 0.05, constant K of 0.023, CL is 0.8 and minimum drag occurs at CL of 0.1. Hence, CLmindrag = 0.1. Now, CD is given by, CD = CDmin + K*(CL – CLmindrag)2 = 0.05+0.023*(0.8-0.1)2 = 0.06127.
9 - Question
Following diagram represents ______________
a) cambered wing drag polar
b) cambered airfoil drag polar
c) symmetric wing drag polar
d) drag polar of an airfoil
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 - Question
A wing is designed to operate with free stream velocity of 20m/s and air density of 1.225 kg/m3. Find aerodynamic efficiency of given wing. Consider S as 8 m2, CL as 0.9 and CD as 1.25.
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.