# Aircraft Performance Questions and Answers – Lift Force

This set of Aircraft Performance Multiple Choice Questions & Answers (MCQs) focuses on “Lift Force”

1. Which of the following is the correct formula for lift of an aircraft?
a) L=$$\frac{1}{2}$$ρV2SCl
b) L=$$\frac{1}{2}$$V2SCl
c) L=$$\frac{1}{2}$$ρV2Cl
d) L=$$\frac{1}{2}$$PV2SCl

Explanation: The correct formula for lift of an aircraft is L=$$\frac{1}{2}$$ρV2SCl where L is lift, ρ is density, V stands for velocity of the aircraft, S is span and Cl is coefficient of lift. Lift is one of the fundamental aerodynamic force that is caused for the upward movement of the aircraft.

2. The zero-lift angle of attack is zero if the aerofoil is symmetric.
a) True
b) False

Explanation: The zero-lift angle of attack is zero if the aerofoil is symmetric. The zero-lift angle of attack is represented by α0. For symmetric airfoil the α0=0. For asymmetric airfoil the α0≠0. In the case of cambered airfoil the value of α0 is negative.

3. In the case of cambered airfoil the value of α0 is positive.
a) True
b) False

Explanation: In the case of cambered airfoil the value of α0 is negative. The zero-lift angle of attack is zero if the aerofoil is symmetric. The zero-lift angle of attack is represented by α0. For symmetric airfoil the α0=0, for asymmetric airfoil the α0≠0.

4. What is the theoretical value of lift curve slope of a flat airfoil?

Explanation: The theoretical value of lift curve slope of a flat airfoil is 2π per radian i.e. $$\frac{dC_l}{d\alpha}$$=0 here $$\frac{dC_l}{d\alpha}$$ is lift curve slope. This factor is depending on thickness of the airfoil and aspect ratio. The lift curve slope is directly proportional to thickness and inversely proportional to aspect ratio.

5. The lift curve slope is inversely proportional to thickness and directly proportional to aspect ratio.
a) True
b) False

Explanation: The lift curve slope is directly proportional to thickness and inversely proportional to aspect ratio. The lift slope curve is given by $$\frac{dC_l}{d\alpha}$$. The theoretical value of lift curve slope of a flat airfoil is 2π per radian i.e. $$\frac{dC_l}{d\alpha}$$=0.
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6. What is meant by stalling angle of attack?
a) The angle at which we get maximum lift
b) The angle at which we get maximum drag
c) The angle at which we get minimum lift
d) The angle at which we get minimum drag

Explanation: Stalling angle of attack is the angle at which we get maximum lift. Stalling angle is the maximum angle at which the aircraft can be maintained in a steady state.

7. Which of the following is a correct condition when the angle of attack is beyond stalling angle?
a) Drag is greater than lift
b) Weight is less than lift
c) Drag is less than lift
d) Weight is less than thrust

Explanation: The drag is greater than lift in the condition where the angle of attack is beyond stalling angle. Stalling angle of attack is the angle at which we get maximum lift. Stalling angle is the maximum angle at which the aircraft can be maintained in a steady state.

8. What is the value of Cl of a symmetric airfoil if the angle of attack is 5°?
a) 0.548
b) 0.0548
c) 3.1416
d) 31.416

Explanation: The answer is 0.548. For a symmetrical aerofoil Cl= 2πα. Given α=5°,
Convert α into radians: α=5*$$\frac{\pi}{180}$$
α=0.0873
Now substitute α in the above formula: Cl= 2π*0.0873
Cl=0.548.

9. What is stall buffet condition?
a) It is an aerodynamic vibration caused due to separation or turbulent airflow
b) it is a condition in which the separation is laminar
c) It is a condition in which the flow separation is not considered
d) It is a condition in which the turbulence is not considered

Explanation: Stall buffet condition is a condition in which the aerodynamic vibration caused due to separation or turbulent airflow. This starts at the back of the wing and moves forward as there is increase in angle of attack.

10. Leading edge flap deflection has the effect of increasing lift curve to a higher stalling angle of attack.
a) True
b) False

Explanation: Leading edge flap deflection has the effect of increasing lift curve to a higher stalling angle of attack. This helps in reducing the take-off and landing speeds of the aircraft.

11. Trailing edge deflection decreases the camber of the aerofoil section.
a) True
b) False

Explanation: Trailing edge deflection decreases the camber of the aerofoil section. This results in shifting the lift characteristics upwards when the zero angle of attack becomes more negative.

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