POF – 05 Lift – LMS | INSIGHT AVIATION

📄 Detailed Notes ⚡ Cheat Sheet

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📄 Notes ⚡ Cheat Sheet

Q1. With a decrease in angle of attack:

the stagnation point moves forward

the separation point moves forward

form drag will increase

induced drag will increase

Q2. On a non-swept wing, when the aerofoil is accelerated from subsonic to supersonic speeds, the aerodynamic centre:

shifts from 25% to about 50% of the aerofoil chord

shifts aft by about 10%

remains unchanged

slightly shifts forward

Q3. An aeroplane in straight and level flight is subjected to a strong vertical gust. The point on the wing, where the instantaneous variation in wing lift effectively acts is known as the:

neutral point

centre of thrust

centre of gravity

aerodynamic centre of the wing

Q4. The aerodynamic centre is a fixed point where:

The boundary layer changes from laminar to turbulent

The velocity of the relative airflow is reduced to zero

Assuming no flow separation, the pitching moment coefficient does not change with varying angle of attack

The aeroplane becomes longitudinally unstable when the CG is moved beyond it in an aft direction

Q5. The point about which the wing pitching moment is independent of angle of attack is called:

the pitching centre

the centre of gravity

the centre of pressure

the aerodynamic centre

Q6. The aerodynamic centre is the point on the chord line where:

drag acts

the sum of all aerodynamic force act

the geometric centre of the wing is located

the pitching moment remains constant throughout changes in angle of attack within the normal range

Q7. The aerodynamic centre of the wing is the point, where:

change of lift due to variation of angle of attack is constant

pitching moment coefficient does not vary with angle of attack

aerodynamic forces are constant

the aeroplanes lateral axis intersects with the centre of gravity

Q8. The aerodynamic centre is:

the point where the CG meets the lateral axis

the point where all changes in the magnitude of the lift force effectively take place

situated at about 50% chord

always to the rear of the neutral point

Q9. The chord line of an aerofoil section is the line :

relative to the upstream airflow

drawn between the leading and the trailing edges

of points equidistant from upper and lower surfaces

the tangent at the leading edge of the camber line

Q10. The relative thickness of an aerofoil is expressed in:

degrees cross section tail angle

% chord

camber

meters

Q11. A line connecting the leading and trailing edge midway between the upper and lower surface of an aerofoil. This definition is applicable for:

the mean aerodynamic chord line

the chord line

the camber line

the upper camber line

Q12. Consider an aerofoil with a certain camber and a positive angle of attack. At which location will the highest flow velocities occur?

Upper side

Lower side

In front of the stagnation point

In the stagnation point

Q13. Thickness chord ratio of an aerofoil section is expressed in percentage of :

thickness

blade surface

chord

wingspan

Q14. The airfoil chord line is:

a straight line from the wing leading edge to the trailing edge

a line equidistant from the upper and lower wing surfaces

a line tangential to the wing surface at the point of maximum curvature

a line drawn at 15% chord from the root to the tip

Q15. How is the thickness of an aerofoil section measured?

As the ratio of wing angle

Related to camber

As the percentage of chord

In metres

Q16. In a symmetrical airfoil the mean camber line is?

A line joining points of mean camber along the wing

A line joining points of maximum camber along the wing

A curve co-incident with the top surface of the airfoil

A straight line co-incident with the chord line

Q17. The mean camber line of an aerofoil section is:

a straight line from the leading edge to the trailing edge

a line from the leading to the trailing edge equidistant from the upper and lower surfaces

the profile of the upper surface of an aerofoil section

an arc of circle from the leading edge to the trailing edge

Q18. The chord line of an aerofoil is:

a line from wing tip to wing tip

a line from the leading edge to trailing edge equidistant from the upper and lower surfaces

a straight line joining the centre of curvature of the leading and trailing edges

a horizontal line tangential to the wing surface

Q19. The term angle of attack is defined as:

the angle that determines the magnitude of the lift force

the angle between the wing chord line and the relative wind

the angle between the relative airflow and the horizontal axis

the angle between the wing and tailplane incidence

Q20. The angle of attack of a wing profile is defined as the angle between:

the undisturbed airflow and the chord line

the local airflow and the mean camber line

the local airflow and the chord line

the undisturbed airflow and the mean camber line

Q21. The term angle of attack in a two dimensional flow is defined as:

the angle for maximum lift/drag ratio

the angle between the aeroplane climb path and the horizon

the angle formed by the longitudinal axis of the aeroplane and the chord line of the wing

the angle between the wing chord line and the direction of the relative wind/airflow

Q22. The term angle of attack in a two dimensional flow is defined as:

the angle between the aeroplane climb path and the horizon.

the angle for maximum lift/drag ratio.

the angle between the wing chord line and the direction of the relative wind/airflow.

the angle formed by the longitudinal axis of the aeroplane and the chord line of the wing.

Q23. The angle of attack (aerodynamic angle of incidence) of an aerofoil is the angle between the:

bottom surface and the chord line

chord line and the relative undisturbed airflow

bottom surface and the Horizontal

bottom surface and the relative airflow

Q24. The angle of attack of an aerofoil section is the angle between the:

bottom surface and the chord line

bottom surface and the horizontal

bottom surface and the relative airflow

chord line and the relative undisturbed airflow

Q25. The angle between the airflow (relative wind) and the chord line of an aerofoil is:

climb path angle

glide path angle

angle of attack

same as the angle between chord line and fuselage axis

Q26. The angle of attack of a two dimensional wing section is the angle between:

the chord line of the aerofoil and the fuselage centreline

the chord line of the aerofoil and the free stream direction

the fuselage core line and the free stream direction

the chord line and the camber line of the aerofoil

Q27. The angle of attack of an aerofoil is the angle between the :

bottom surface of the aerofoil and the relative airflow

chord line and the plane of rotation

chord line and the relative airflow

plane of rotation and the direction of flight of the helicopter

Q28. An aeroplane’s angle of attack can be defined as the angle between it’s:

Lateral axis and the horizontal plane

Longitudinal axis and the horizontal plane

Speed vector and longitudinal axis

Speed vector and the horizontal plane

Q29. In level flight an increase in angle of attack will cause:

the stagnation point to move down and aft

the boundary layer to become thinner

a decrease in pitch angle

the centre of pressure to move aft

Q30. The angle between the aeroplane longitudinal axis and the chord line of the wing is the:

angle of incidence

glide path angle

angle of attack

climb path angle

Q31. That force, which acts at right angles to the relative airflow, is:

Lift

Thrust

Total reaction

Drag

Q32. The lift and drag forces, acting on a wing cross section:

vary linearly with the angle of attack

depend on the pressure distribution about the wing cross section

are normal to each other at just one angle of attack

are proportional to each other, independent of angle of attack

Q33. CL varies with:

density

velocity

angle of attack

pressure

Q34. Lift is produced by:

An increase in flow velocity giving an increase in pressure on the lower surface

Airflow velocity increasing over top surface giving decreased pressure and velocity decreasing over lower surface giving increased pressure

An increase in flow velocity on the lower surface and decrease on the upper surface

Airflow velocity increasing downward having been deflected by the blade pitch angle and creating an upward pressure on the blade

Q35. Drag is in the direction of – and lift is perpendicular to the:

chord line

relative wind/airflow

horizon

longitudinal axis

Q36. The forces of lift and drag on an aerofoil are respectively, normal and parallel to the:

Relative wind/airflow

longitudinal axis.

Horizon

Chord line

Q37. Lift is generated when:

an aerofoil is placed in a high velocity air stream

the shape of the aerofoil is slightly cambered

a certain mass of air is accelerated downwards

a certain mass of air is retarded

Q38. Lift and drag on an aerofoil are vertical respectively parallel to the:

horizon

relative wind/airflow

chord line

longitudinal axis

Q39. A flat plate, when positioned in the airflow at a small angle of attack, will produce:

both lift and drag

lift but no drag

drag but no lift

neither lift nor drag

Q40. Lift is the:

Vertical component of the total aerodynamic force, perpendicular to the undisturbed airflow

Component of the total aerodynamic force, perpendicular to the local airflow

Component of the total aerodynamic force, perpendicular to the undisturbed flow

Component of the total aerodynamic force, perpendicular to the mean aerodynamic chord.

Q41. The lift force, acting on an aerofoil:

is mainly caused by suction on the upper side of the aerofoil

increases, proportional to the angle of attack until 40 degrees

is mainly caused by overpressure at the underside of the aerofoil

is maximum at an angle of attack of 2 degrees

Q42. Assuming no flow separation, which of these statements about the flow around an aerofoil as the angle of attack decreases ai e correct or incorrect?

Q43. The forces acting on a symmetrical aerofoil element when the upstream airflow is parallel to the chord line is :

only lift

zero

only drag

lift and drag are at a ratio 2/3

Q44. Which statement is correct about the CI and angle of attack?

For a symmetric aerofoil, if angle of attack = 0, CI = 0

For a symmetric aerofoil, if angle of attack = 0, CI is not equal to 0

For an asymmetric aerofoil, if angle of attack = 0, CI = 0

For an asymmetric aerofoil with positive camber, if angle of attack is greater than 0, CI = 0

Q45. A symmetrical aerofoil at zero degree angle of attack will, in level flight produce:

most of its lift on the lower surface

most of its lift on the upper surface

the same amount of lift on the upper and lower surfaces

zero lift

Q46. At zero angle of attack in flight, a symmetrical wing section will produce:

some lift and drag

zero lift with some induced and profile drag

zero lift and drag

zero lift with some drag

Q47. Wing loading is:

the ratio of lift to wing weight

the ratio of wing area to wing weight

the ratio of lift to aircraft weight

the ratio of aircraft weight to wing area

Q48. Lift is a function of:

velocity, density, wing shape and lift coefficient

velocity, wing area, CL and density

velocity, wing shape, wing area and CL

velocity, density, gross wing area and CD

Q49. The most correct list of factors that affect the lift produced by an aerofoil are:

angle of attack, air density, velocity, wing area

angle of attack, air temperature, velocity, wing area

angle of attack, velocity, wing area, aerofoil section, air density

incidence, TAS, wing plan, leading edge radius and thrust

Q50. Which of the following statements are correct?

Drag acts in the same direction as the relative airflow and lift perpendicular to it

Lift acts at right angles to the top surface of the wing and drag acts at right angles to lift

Drag acts parallel to the chord and opposite to the direction of motion of the aircraft and lift acts perpendicular to the chord

Lift acts perpendicular to the horizontal and drag parallel in a rearwards direction

Q51. Both lift and drag of an aerofoil are:

proportional to the square of the velocity of the relative airflow

proportional to increases and decreases in the velocity of the relative airflow

inversely proportional to the air density

inversely proportional to the area of the wing

Q52. Assuming no flow separation and no compressibility effects, the location of the center of pressure of a symmetrical aerofoil section:

Is at about 50% chord, irrespective of AOA

Moves forward when AOA decreases

Is at about 20% chord irrespective of AOA

Moves backward when the AOA decreases

Q53. The centre of pressure of an aerofoil element :

is the point where the resultant aerodynamic force is applied

is located in a fixed point in an asymmetric aerofoil

depends on the weight of the aerofoil element

depends only on the boundary layer

Q54. The point at which the resultant of lift forces acts is

The centre of gravity

The centre of pressure

The moment arm

The hub

Q55. If a symmetrical aerofoil is accelerated from subsonic to supersonic speed the centre of lift will move:

forward to the leading edge

aft to the trailing edge

aft to the mid chord

forward to the mid chord

Q56. Assuming no flow separation, when speed is decreased in straight and level flight on a positively cambered aerofoil, what happens to the:

Q57. The point, where the single resultant aerodynamic force acts on an aerofoil, is called:

neutral point

centre of gravity

centre of pressure

aerodynamic centre

Q58. The point, where the aerodynamic lift acts on a wing is:

the c.g. location

the centre of pressure

the point of maximum thickness of the wing

the suction point of the wing

Q59. The centre of pressure is the point:

at which the extended rotor axis intersects the plane of rotation

of maximum pressure on the under surface of the aerofoil

on the chord line of an aerofoil section through which the resultant of the lift forces acts

through which the four forces acting on a helicopter act

Q60. The centre of the pressure of a symmetrical aerofoil section is behind the leading edge approximately at the following % of the section chord :

10%

15%

25%

35%

Q61. The location of the centre of pressure of a positively cambered aerofoil section at increasing angle of attack will:

shift forward until approaching the critical angle of attack

not shift

shift aft until approaching the critical angle of attack

shift in spanwise direction

Q62. The centre of pressure of an aerofoil is:

the point where the pressure on the upper surface of the wing is lowest

the centre of gravity of the aerofoil

the point where the pressure on the lower surface of the wing is highest

the point on the chord line where the resultant lift force acts

Q63. The CL/CD ratio is:

Max at an angle of attack of 15°

Min at an angle of attack of 4°

Min at an angle of attack of 16°

Max at an angle of attack of 4°

Q64. Comparing the lift coefficient and drag coefficient at normal angle of attack:

CL is much greater than CD

CL has approximately the same value as CD

CL is lower than CD

CL is much lower than CD

Q65. Which one of the following statements about the lift-to-drag ratio in straight and level flight is correct?

At the highest value of the lift/drag ratio the total drag is lowest

The highest value of the lift/drag ratio is reached when the lift is zero

The lift/drag ratio always increases as the lift decreases

The highest value of the lift/drag ratio is reached when the lift is equal to, the aircraft weight

Q66. If the weight of an aircraft is increased, the maximum lift/drag ratio will:

decrease

increase

not be affected

increase although the aircraft will have to be flown more slowly

Q67. Which statement about an aeroplane leaving ground effect is correct? 1. The downwash angle increases. 2. The induced angle of attack increases

1 Is incorrect; 2 is correct

1 is correct; 2 is correct

1 is correct; 2 is incorrect

1 is incorrect; 2 is incorrect

Q68. Which statement, about an aeroplane entering ground effect at constant angle of attack, is correct?

Q69. What will happen in ground effect?

an increase in strength of the wing tip vortices

The wing downwash on the tail surfaces increases

The induced angle of attack and induced drag decreases

A significant increase in thrust required

Q70. Ground effect has the following influence on the landing distance:

decreases

increases

does not change

increases, only if the landing flaps are fully extended

Q71. If an aeroplane flies in the ground effect:

drag and lift are reduced

the effective angle of attack is decreased

the induced angle of attack is increased

the lift is increased and the drag is decreased

Q72. When an aeroplane enters ground effect:

the effective angle of attack is decreased

the lift is increased and the drag is decreased

the induced angle of attack is increased

drag and lift are reduced

Q73. Assuming constant IAS, when an aeroplane enters ground effect:

The effective angle of attack decreases

Induced drag reduces

The induced angle of attack increases

Downwash does not change

Q74. The Coefficient of Lift of an aerofoil section:

Is at its maximum value at that AOA giving the maximum lift to drag ratio

Increases with an increase in AOA up to the stall

Changes with density

Changes with the velocity of the airflow, squared

Q75. A positively cambered aerofoil will generate zero lift:

At a negative angle of attack

It can never generate zero lift

At a positive angle of attack

A zero angle of attack

Q76. The Cl – alpha curve of a positive cambered aerofoil intersects with the vertical axis of the Cl – alpha graph:

in the origin

below the origin

nowhere

above the origin

Q77. The point in the annex showing zero lift is:

Q78. When considering an angle of attack versus coefficient of lift graph for a cambered aerofoil, where does the lift curve intersect the vertical CL axis?

Above the origin

Below the origin

At the point of origin

To the left of the origin

Q79. Regarding the lift formula, if density doubles, lift will:

Remain the same

Be Doubled

Be Halved

Be four times greater

Q80. If the airspeed is doubled, whilst maintaining the same control surface deflection, the aerodynamic force on this control surface will:

Be doubled

Be four times greater

Increase by the square root of the airspeed

Become four times smaller

Q81. The terms q and S in the lift formula are:

square root of surface and wing loading

dynamic pressure and the area of the wing

static pressure and wing surface area

static pressure and dynamic pressure

Q82. The formula for lift can be written as:

L = CL1/2pV^2S

L = CL1/2pVS^2

L = CL1/2pVS

L= CL1/4pV^2S

Q83. Whilst maintaining straight and level flight with a lift coefficient (CL = 1), what will be the new value of CL after the speed has doubled?

0.5

0.25

Q84. When an aeroplane is flying at an airspeed which is 1.3 times its basic stalling speed, the coefficient of lift as a percentage of the maximum lift coefficient (CLmax) would be:

169%

130%

59%

77%

Q85. Whilst maintaining straight and level flight with a lift coefficient (CL = 1), what will be the approximate new value of CL after the speed is increased by 41%

0.5

0.25

0.6

0.3

Q86. Whilst maintaining straight and level flight with a lift coefficient (CL = 1), what will be the approximate new value of CL after the speed is increased by 30%?

0.3

0.5

0.6

0.25

Q87. The lift formula is:

L = CL ½ RHO V2 S

L = W

L = CL2 RHO V2 S

L = Nw

Q88. Assuming standard atmospheric conditions, in order to generate the same amount of lift as altitude is increased, the aeroplane must be flown at:

A lower TAS for any given angle of attack

A lower TAS and a lower angle of attack

A higher TAS for any given angle of attack

The same TAS regardless of angle of attack

Q89. Regarding the lift formula, if airspeed doubles, lift will:

Not change

Be four times greater

Also double

Halve

Q90. Which of the following variables are required to calculate lift from the lift formula?

Square root of the wing area, density and wing loading

angle of attack, aspect ratio and dynamic pressure

Total pressure and wing area only

Dynamic pressure, lift coefficient and wing area

Q91. If the wing area is increased, lift will:

Remain constant

Increase with the square of the wing area

Increase, because it is directly proportional to wing area

Not change because the lift coefficient is constant

Q92. What do ‘S’ and ‘q’ represent in the lift equation?

static pressure and chord

wing span and dynamic pressure

wing area and dynamic pressure

wing area and static pressure

Q93. To maintain level flight, if the angle of attack is increased the speed must be:

reduced

increased in the same ratio as the lift/drag ratio decreases

kept constant

increased

Q94. The lift coefficient (CL) of an aeroplane in steady horizontal flight is 0.4. Increase of angle of attack of 1 degree will increase CL by 0.09. A vertical up gust instantly changes the angle of attack by 5 degrees. The load factor will be:

3.18

1.09

2.13

Q95. Load factor is the actual lift supported by the wings at any given time:

divided by the surface area of the wing

divided by the total weight of the aircraft

subtracted from the aircraft’s total weight

divided by the aircraft’s empty weight

Q96. The load factor is:

the ratio of lift to drag

the ratio of centripetal force to lift

the ratio of thrust to weight

the ratio of lift to weight

Q97. Which statement is correct?

The centre of pressure is the point on the wings leading edge where the airflow splits up

As the angle of attack increases, the stagnation point on the wing’s profile moves downwards

The stagnation point is another name for centre of pressure

The stagnation point is always situated on the chord line, the centre of pressure is not

Q98. With increasing angle of attack, the stagnation point will move (l) and the point of lowest pressure will move (ll).

(l) up; (ll) aft

(l) down; (ll) forward

(ll) down; (ll) aft

(l) up; (ll) forward

Q99. What is the stagnation point?

The intersection of the total aerodynamic force and the chord line

The point where the velocity of the relative airflow is reduced to zero

The intersection of the thrust vector and the chord line

The point, relative to which the sum total of all moments is independent of angle of attack

Q100. Which of the following wing planforms gives the highest local profile lift coefficient at the wing root?

Rectangular

Elliptical

Tapered

Positive angle of sweet

Q101. Wing sweep angle is the angle between:

The wing plane and the horizontal with the aeroplane in an unbanked, level condition

The quarter-chord line of the wing and the normal axis

The leading edge of the wing and the lateral axis

The quarter-chord line of the wing and the lateral axis

Q102. Taper ratio is the ratio of:

the root thickness to the root chord

the span to the root chord

the root chord to the tip chord

the span to the mean chord

Q103. The Mean Aerodynamic Chord (MAC) for a given wing of any platform is:

the chord of a rectangular wing with same moment and lift

the chord of a large rectangular wing

the average chord of the actual aeroplane

the wing area divided by the wing span

Q104. The aspect ratio of the wing:

is the ratio between the wing span and the root chord

is the ratio between the wing span and the mean geometric chord

is the ratio between the tip chord and the wing span

is the ratio between chord and root chord

Q105. What wing shape or wing characteristic is the least sensitive to turbulence:

swept wings

straight wings

wing dihedral

winglets

Q106. Which of the following wing planforms will be least affected by turbulence?

straight, high aspect ratio

swept, low aspect ratio

straight, moderate aspect ratio

swept, high aspect ratio

Q107. Dihedral of the wing is:

the angle between the 0.25 chord line of the wing and the vertical axis

the angle between the leading edge of the wing and the lateral axis

the angle between the 0.25 chord line of the wing and the lateral axis

the angle between the 0.25 chord line of the wing and the horizon

Q108. Aspect ratio is:

the ratio of the mean chord to the maximum wing thickness

the ratio of the wingspan to the square of the mean chord

the ratio of the wingspan to the mean chord

the ratio of the wingspan to the wing area

Q109. Which of the following is the correct definition of aspect ratio?

span divided by tip chord

chord divided by span

span divided by mean chord

chord divided by span, measured at the 25% chord position

Q110. If IAS is doubled, by which of the following factors should the original CL be multiplied to maintain level flight?

0.25

0.5

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