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EASA ATPL 081 Principles of Flight study materials
EASA ATPL 081 Principles of Flight study resources

EASA ATPL 081 Principles of Flight practice questions

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Question 1Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

In the ISA, what is the value of the sea-level pressure and the sea-level temperature that form the reference datum for the standard atmosphere?

  1. 1013.25 hPa and 15 °C Correct answer.
  2. 1000 hPa and 15 °C
  3. 1013.25 hPa and 20 °C
  4. 1013.25 hPa and 0 °C

Explanation

The International Standard Atmosphere is defined from mean sea level with a pressure of 1013.25 hPa, a temperature of 15 °C and a density of 1.225 kg/m³. These datums are used to compute the pressure and density ratios needed for aircraft performance. The lapse rate of 1.98 °C per 1000 ft applies only in the troposphere.

Question 2Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Which statement correctly describes how static pressure and density change as altitude increases within the troposphere of the ISA?

  1. Pressure remains constant while density decreases
  2. Pressure decreases while density remains constant
  3. Both pressure and density decrease with height Correct answer.
  4. Both pressure and density increase with height

Explanation

In the troposphere the ISA temperature falls with height, and because pressure is the weight of the overlying air it always decreases with altitude. Density depends on both pressure and temperature, and the falling pressure dominates, so density also decreases with height. This is why true airspeed exceeds indicated airspeed at altitude.

Question 3Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

An aircraft climbs in the ISA from sea level to 18 000 ft. Using the standard lapse rate, by how much has the outside air temperature fallen?

  1. About 54 °C
  2. About 11 °C
  3. About 18 °C
  4. About 36 °C Correct answer.

Explanation

The ISA temperature lapse rate in the troposphere is 1.98 °C per 1000 ft, approximately 2 °C per 1000 ft. Multiplying by 18 gives roughly 36 °C of cooling, so the temperature would be about −21 °C. This linear lapse applies only below the tropopause at 36 090 ft.

Question 4Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

In steady, incompressible flow through a streamtube of varying cross-section, what happens to the flow speed where the cross-sectional area is smallest?

  1. The speed is lowest there
  2. The speed depends only on the static pressure there
  3. The speed is highest there Correct answer.
  4. The speed is unchanged there

Explanation

The continuity equation states that mass flow, the product of density, area and velocity, is constant along a streamtube. With constant density, a smaller area must be matched by a higher velocity. This is why air accelerates over the upper surface of a wing where the streamtube is constricted.

Question 5Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

For an incompressible, frictionless flow along a streamline, Bernoulli's equation relates static pressure, dynamic pressure and which third quantity?

  1. Total pressure, which stays constant along the streamline Correct answer.
  2. Mass flow, which varies with the cross-section
  3. Static temperature, which stays constant along the streamline
  4. Density, which varies inversely with velocity

Explanation

Bernoulli's equation expresses conservation of energy per unit volume: static pressure plus dynamic pressure equals total pressure, a constant along a streamline in steady incompressible flow. Where velocity rises, static pressure falls by the same amount that dynamic pressure rises. Density and mass flow are not the third term of this sum.

Question 6Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

A light aircraft flies at a true airspeed of 50 m/s at sea level in the ISA, where density is 1.225 kg/m³. What is the dynamic pressure?

  1. 3063 Pa
  2. 1531 Pa Correct answer.
  3. 61.3 Pa
  4. 76.6 Pa

Explanation

Dynamic pressure equals one half of density times velocity squared. Half of 1.225 is 0.6125, and 50 squared is 2500, giving 1531 Pa. This dynamic pressure is what the pitot-static system senses to produce indicated airspeed.

Question 7Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Aircraft A flies at 100 m/s at sea level and aircraft B flies at 100 m/s at 6000 m, where the density is about half the sea-level value. How do their dynamic pressures compare?

  1. Both experience the same dynamic pressure
  2. B experiences about four times the dynamic pressure of A
  3. B experiences about twice the dynamic pressure of A
  4. A experiences about twice the dynamic pressure of B Correct answer.

Explanation

Dynamic pressure is half rho v squared, so at equal true airspeed it is directly proportional to density. At 6000 m the density is roughly half the sea-level value, so A has about twice the dynamic pressure of B. This is why the same indicated airspeed corresponds to a much higher true airspeed at altitude.

Question 8Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Why does a pitot-static system indicate a lower value than true airspeed as altitude increases at constant true airspeed?

  1. The dynamic pressure sensed falls with density while the instrument is calibrated for sea-level density Correct answer.
  2. The static pressure at the static port rises with altitude and reduces the indicated value
  3. The instrument is calibrated for a temperature of 0 °C at all altitudes
  4. The total pressure in the pitot tube falls faster than the static pressure

Explanation

The instrument computes speed from the measured dynamic pressure using sea-level density, so it effectively shows equivalent airspeed. At altitude the actual density is lower, so the same true airspeed produces a smaller dynamic pressure and a lower indication. The ratio between true and indicated airspeed is approximately the square root of the density ratio.

Question 9Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

In two-dimensional potential flow, which combination of quantities describes the flow field completely at every point?

  1. Mass flow and total pressure at each point
  2. Static pressure and density at each point
  3. Velocity magnitude and direction at each point Correct answer.
  4. Temperature and entropy at each point

Explanation

A two-dimensional flow field is fully described by the velocity vector at each point, from which streamlines and pressures follow through Bernoulli's equation. Pressure, density and temperature are derived consequences of the velocity field in incompressible potential flow. Mass flow and total pressure are not independent descriptors of the field.

Question 10Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Which statement about streamlines in steady flow is correct?

  1. Streamlines are always straight lines parallel to the free stream
  2. No flow crosses a streamline, and the velocity is tangent to it at every point Correct answer.
  3. Streamlines represent the path of a single particle over time
  4. Streamlines may cross where the flow accelerates

Explanation

By definition a streamline is everywhere tangent to the instantaneous velocity, so fluid cannot cross it in steady flow. A bundle of streamlines forms a streamtube through which the mass flow is constant. Streamlines may be curved and are not the same as the path traced by a particle over time.

Question 11Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Air flows over a wing upper surface where the local velocity is 1.2 times the free-stream velocity. Assuming incompressible flow with no losses, what is the local static pressure coefficient relative to the free stream?

  1. −0.20
  2. −0.44 Correct answer.
  3. −1.20
  4. +0.44

Explanation

The pressure coefficient is one minus the square of the velocity ratio, so with a ratio of 1.2 the square is 1.44 and the coefficient is −0.44. The negative sign shows suction relative to free-stream static pressure. This suction is what generates most of the lift on a typical aerofoil.

Question 12Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

At the stagnation point on the leading edge of an aerofoil in incompressible flow, what is the relationship between total pressure, static pressure and dynamic pressure?

  1. Total pressure equals static pressure minus dynamic pressure
  2. Total pressure equals static pressure plus free-stream dynamic pressure Correct answer.
  3. Static pressure equals total pressure plus dynamic pressure
  4. Dynamic pressure is zero because the air is stationary

Explanation

At the stagnation point the local velocity is zero, so the local static pressure rises to the total pressure of the free stream. By Bernoulli's equation this total pressure equals free-stream static pressure plus free-stream dynamic pressure. The dynamic pressure of the free stream is not zero, only the local velocity is.

Question 13Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

Comparing the tropopause with sea level in the ISA, which statement is correct?

  1. Temperature is about 0 °C and pressure is about a quarter of the sea-level value
  2. Temperature is about −56.5 °C and pressure remains equal to the sea-level value
  3. Temperature is about −56.5 °C and pressure is about half the sea-level value
  4. Temperature is about −56.5 °C and pressure is about a quarter of the sea-level value Correct answer.

Explanation

In the ISA the tropopause is at 36 090 ft, where the temperature reaches about −56.5 °C and the pressure falls to roughly 226 hPa, about a quarter of the sea-level value. Above the tropopause the temperature remains constant in the lower stratosphere. Pressure never stays at the sea-level value once altitude is gained.

Question 14Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

In a venturi of a carburettor, the throat area is one quarter of the inlet area. If the inlet velocity is 30 m/s and the flow is incompressible, what is the velocity in the throat?

  1. 60 m/s
  2. 120 m/s Correct answer.
  3. 30 m/s
  4. 7.5 m/s

Explanation

Continuity for incompressible flow gives area times velocity as constant, so reducing the area to one quarter raises the velocity by a factor of four. Four times 30 m/s is 120 m/s. The corresponding static pressure in the throat falls, which is what draws fuel into the airflow.

Question 15Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

An aerofoil section is tested at a Reynolds number of 3 million and again at 6 million with the same angle of attack. What is the most likely effect on the boundary layer and maximum lift?

  1. The boundary layer becomes relatively thicker and maximum lift tends to decrease
  2. The flow becomes fully laminar over the whole chord
  3. The boundary layer thickness is unchanged but the stall angle falls
  4. The boundary layer becomes relatively thinner and maximum lift tends to increase Correct answer.

Explanation

A higher Reynolds number means inertial forces are larger relative to viscous forces, so the boundary layer is relatively thinner and more resistant to separation. This typically delays stall to a higher angle of attack and increases maximum lift. The effect is a trend, not an exact law, and depends on the aerofoil and surface condition.

Question 16Fundamentals of aerodynamics: atmosphere, airflow, Bernoulli and continuity, two-dimensional flow

In incompressible flow over an aerofoil, the pressure coefficient at a point is −2.0 and the free-stream dynamic pressure is 800 Pa. What is the local static pressure if free-stream static pressure is 101 325 Pa?

  1. 100 525 Pa
  2. 99 725 Pa Correct answer.
  3. 101 325 Pa
  4. 103 925 Pa

Explanation

The pressure coefficient is the local static pressure minus free-stream static pressure, divided by free-stream dynamic pressure. A coefficient of −2.0 means the local static pressure is 2 × 800 = 1600 Pa below free-stream static pressure. Subtracting 1600 Pa from 101 325 Pa gives 99 725 Pa.

Question 17Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

In the standard lift formula L = ½ρV²SC_L, which quantity does the term ½ρV² represent?

  1. Dynamic pressure of the free stream Correct answer.
  2. Reynolds number per unit length
  3. Static pressure at the stagnation point
  4. Total pressure measured by a pitot probe

Explanation

The combination ½ρV² is the dynamic pressure of the undisturbed free stream, and lift scales directly with it. Multiplying by wing area and lift coefficient converts this pressure into a force. Stagnation and total pressure are different quantities that include the static pressure term.

Question 18Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

For a given aerofoil at a fixed angle of attack, what happens to the lift coefficient as the Reynolds number increases from a low value into the fully turbulent regime?

  1. It falls steadily toward zero
  2. It remains exactly constant with Reynolds number
  3. It becomes negative because viscosity dominates
  4. It generally rises and the maximum lift coefficient increases Correct answer.

Explanation

Higher Reynolds number thickens the turbulent boundary layer and delays separation, so the aerofoil sustains attachment to a higher angle of attack. This raises both the slope region performance and the maximum lift coefficient. Lift coefficient is therefore not independent of Reynolds number.

Question 19Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Which statement correctly describes the pressure distribution around a cambered aerofoil producing lift?

  1. The upper surface has higher pressure, and the net force is upward because of skin friction
  2. Both surfaces carry equal suction, and lift comes only from the leading-edge stagnation point
  3. The upper surface has lower pressure than the lower surface, giving a net upward force Correct answer.
  4. Pressure is uniform over both surfaces and lift arises purely from momentum change far downstream

Explanation

Lift arises because the upper surface static pressure is reduced relative to the lower surface, so integrating pressure around the section gives a net force with an upward component. The contribution of skin friction to lift is small and acts along the surface. Suction is not equal on both surfaces for a lifting section.

Question 20Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

At the stagnation point on an aerofoil in subsonic flow, what is true of the local conditions?

  1. Local velocity is zero and pressure coefficient equals −1
  2. Local velocity is zero and pressure coefficient equals +1 Correct answer.
  3. Local velocity equals free-stream velocity and pressure coefficient equals 0
  4. Local velocity is maximum and pressure coefficient equals −1

Explanation

At the stagnation point the flow is brought to rest, so local dynamic pressure is zero and the static pressure reaches the total pressure. In coefficient form this gives C_p = +1. The value −1 corresponds to a point where local velocity is far above free-stream, not to stagnation.

Question 21Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

As angle of attack increases toward the stall on a typical aerofoil, how does the centre of pressure typically move?

  1. It stays exactly at the quarter-chord point for all angles
  2. It moves aft toward the trailing edge
  3. It moves forward toward the leading edge Correct answer.
  4. It moves to the trailing edge and then back to mid-chord

Explanation

Increasing angle of attack raises the suction peak near the leading edge, so the resultant pressure force acts further forward and the centre of pressure migrates forward. The aerodynamic centre, by contrast, stays near the quarter chord. Confusing these two reference points is a common error.

Question 22Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

What is the boundary layer?

  1. The zone of compressed air ahead of the leading edge
  2. The thin region near the surface where velocity rises from zero to the free-stream value Correct answer.
  3. The layer of separated air behind the trailing edge
  4. The region of inviscid flow outside the wing where streamlines are parallel

Explanation

Viscosity causes the flow to adhere to the surface, creating a thin shear layer in which velocity increases from zero at the wall to the local free-stream value at its outer edge. This is the boundary layer, and its behaviour determines skin friction and separation. The inviscid outer flow is a separate region.

Question 23Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Compared with a laminar boundary layer, a turbulent boundary layer has which characteristic?

  1. Lower skin-friction drag and earlier separation
  2. Higher skin-friction drag but greater resistance to separation Correct answer.
  3. Higher skin-friction drag and earlier separation
  4. Lower skin-friction drag and greater resistance to separation

Explanation

Turbulent mixing transports higher-momentum air toward the wall, so the velocity profile is fuller and the flow can withstand a stronger adverse pressure gradient before separating. The same mixing, however, increases the velocity gradient at the wall and therefore raises skin-friction drag. The two effects are independent in direction.

Question 24Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Which condition promotes transition from laminar to turbulent flow in a boundary layer?

  1. A high Reynolds number and a rough surface Correct answer.
  2. A very low Reynolds number and a polished surface
  3. A strong negative pressure gradient with low free-stream turbulence
  4. A favourable pressure gradient over a long distance

Explanation

Transition is triggered by instabilities that grow when the Reynolds number is high and the surface disturbs the flow. Surface roughness and free-stream turbulence add disturbances that accelerate transition. Favourable gradients and low Reynolds numbers tend to stabilise the laminar layer.

Question 25Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Boundary-layer separation on an aerofoil occurs when the adverse pressure gradient becomes too strong for the air near the wall. What is the immediate consequence for lift?

  1. Lift increases because the separated flow smooths the pressure distribution
  2. Lift becomes negative because the resultant force reverses
  3. Lift decreases because the suction pressure over the affected upper surface is lost Correct answer.
  4. Lift is unaffected because only skin friction changes

Explanation

Separation destroys the low-pressure region over the upper surface, so the pressure difference between upper and lower surfaces collapses and lift drops. Drag rises sharply at the same time. This is the mechanism by which the stall develops.

Question 26Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

What is the critical angle of attack?

  1. The angle at which the wing is set on the fuselage
  2. The angle at which the lift coefficient first becomes positive
  3. The angle at which drag is minimum
  4. The angle at which the boundary layer separates and the stall begins Correct answer.

Explanation

The critical angle of attack is the value at which the flow can no longer follow the upper surface and separates, so lift coefficient reaches its maximum and then decreases. It is a property of the aerofoil section, largely independent of weight and speed. Exceeding it produces the stall.

Question 27Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

An aircraft in straight and level flight at high altitude with a constant indicated airspeed approaches the stall. How does the margin to the critical angle of attack compare with the same indicated airspeed at sea level?

  1. The margin depends only on the aircraft mass, not on altitude
  2. The margin is much smaller at altitude because the air is thinner
  3. The margin is essentially the same because indicated airspeed reflects dynamic pressure Correct answer.
  4. The margin is much larger at altitude because true airspeed is higher

Explanation

Lift depends on dynamic pressure, and indicated airspeed is calibrated to represent dynamic pressure at sea-level density. For the same indicated airspeed the lift coefficient and thus the angle of attack are essentially unchanged, so the stall margin is similar. True airspeed is higher at altitude but that does not reduce the stalling indicated speed.

Question 28Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

A wing is flying at 1.3 times its stalling speed in steady level flight. The aircraft then enters a level turn. What is the load factor at which the wing will stall?

  1. About 1.69 Correct answer.
  2. About 0.59
  3. About 2.60
  4. About 1.30

Explanation

Stall occurs when nV² reaches the value at the 1g stall, so n_stall = (V/V_S)². With V/V_S = 1.3, squaring gives approximately 1.69. Thus the wing stalls when the turn load factor reaches about 1.69, independent of the bank angle needed to produce it.

Question 29Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Why does a swept wing typically stall at a higher angle of attack than an unswept wing of the same section?

  1. The sweep removes the adverse pressure gradient entirely
  2. The sweep makes the boundary layer laminar over the whole chord
  3. The sweep increases the chord Reynolds number considerably
  4. Only the velocity component perpendicular to the leading edge affects the section lift Correct answer.

Explanation

Sweep resolves the free stream into a component normal to the leading edge and a spanwise component. The section behaves largely according to the normal component, which is smaller than the free-stream velocity, so a higher angle of attack is needed for the same lift. This also delays the stall.

Question 30Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

What is meant by the term induced drag?

  1. The drag rise that occurs as the aircraft approaches the speed of sound
  2. The drag component associated with the generation of lift by a finite wing Correct answer.
  3. The drag produced by form separation behind bluff bodies
  4. The drag caused by skin friction on the wetted surface

Explanation

A finite wing sheds a trailing vortex system, and the downwash it creates tilts the local lift vector rearward, producing a drag component called induced drag. It increases with the square of the lift coefficient and decreases with aspect ratio. Skin friction and form drag together make up parasite drag.

Question 31Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

How does increasing the aspect ratio of a wing affect induced drag for the same lift coefficient?

  1. It has no effect on induced drag, only on profile drag
  2. It reduces induced drag because the downwash and tip vortices are weaker Correct answer.
  3. It increases induced drag because the wing has more wetted area
  4. It reduces induced drag only at supersonic speeds

Explanation

Induced drag is inversely proportional to aspect ratio, so a long slender wing spreads its lift over a greater span and generates weaker tip vortices and less downwash. The rearward tilt of the lift vector is therefore smaller. Profile drag may change with wetted area, but induced drag falls.

Question 32Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

In a steady level turn, the load factor is 2. What happens to the stalling speed compared with the 1g stall?

  1. It decreases by a factor of about 1.41
  2. It increases by a factor of about 1.41 Correct answer.
  3. It remains unchanged
  4. It increases by a factor of 2.0

Explanation

Stalling speed varies with the square root of the load factor, so V_S increases by √2, approximately 1.41. Doubling the load factor does not double the stall speed. This is why turn entry requires more lift and raises the stall speed.

Question 33Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Which feature is used to prevent the boundary layer from separating too early at high angles of attack?

  1. A thicker trailing edge that increases the adverse gradient
  2. A reduction in wing span
  3. An increase in the wing loading
  4. A leading-edge slat that re-energises the flow over the upper surface Correct answer.

Explanation

A slat forms a slot that accelerates air over the upper surface, adding energy to the boundary layer and delaying separation to a higher angle of attack. This raises the maximum lift coefficient. Devices that worsen the adverse gradient do the opposite.

Question 34Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

What is the effect of vortex generators on the boundary layer?

  1. They increase the adverse pressure gradient to trip the flow
  2. They reduce skin friction by thickening the boundary layer
  3. They smooth the surface and promote laminar flow
  4. They mix higher-momentum air into the near-wall flow and delay separation Correct answer.

Explanation

Vortex generators create small streamwise vortices that transport high-energy air from the outer flow into the boundary layer near the wall. This makes the layer more resistant to the adverse pressure gradient and delays separation. They are typically fitted ahead of control surfaces.

Question 35Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

A wing at a given angle of attack produces a certain lift coefficient. If the free-stream velocity is doubled while density and angle of attack are unchanged, how does the lift change?

  1. It becomes half as large
  2. It remains the same
  3. It becomes four times as large Correct answer.
  4. It becomes twice as large

Explanation

Lift is proportional to the square of velocity, so doubling velocity multiplies lift by four at constant density, area and lift coefficient. The lift coefficient depends on angle of attack and Reynolds number, not directly on speed. This is why speed has such a strong effect on the lift generated.

Question 36Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

At a point on an aerofoil the local velocity is higher than the free-stream velocity. What is the sign of the local pressure coefficient?

  1. Negative, indicating suction relative to free-stream static pressure Correct answer.
  2. Positive, because the total pressure is constant
  3. Zero, because the dynamic pressure is unchanged
  4. Positive, indicating compression relative to free-stream static pressure

Explanation

Pressure coefficient is defined as (p − p∞) divided by dynamic pressure. Where local velocity exceeds free-stream velocity, Bernoulli's relation implies local static pressure below free-stream static pressure, so C_p is negative. Negative C_p therefore indicates suction.

Question 37Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Which statement about the aerodynamic centre is correct?

  1. It is the point where the lift acts at all angles of attack
  2. It is the point of maximum upper-surface suction
  3. It is the point where the pressure coefficient equals zero
  4. It is the point about which the pitching moment coefficient is essentially constant with angle of attack Correct answer.

Explanation

For a conventional aerofoil the pitching moment about the aerodynamic centre stays nearly constant as angle of attack changes, and it lies near the quarter chord. This makes it a convenient fixed reference for stability analysis. The centre of pressure, where the resultant lift acts, moves with angle of attack.

Question 38Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Why does an aircraft stall at a lower angle of attack when ice has accumulated on the leading edge?

  1. The ice reduces the aircraft mass and changes the trim
  2. The ice increases the wing area and reduces the lift coefficient
  3. The ice disturbs the airflow and causes earlier separation of the boundary layer Correct answer.
  4. The ice increases the Reynolds number and delays transition

Explanation

Leading-edge ice roughens and reshapes the profile, promoting early transition and separation and reducing the maximum lift coefficient. The critical angle of attack is therefore reached at a lower value. This reduces the stall margin and can cause unexpected stalls.

Question 39Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

For a symmetrical aerofoil at zero angle of attack, what is the lift coefficient?

  1. Equal to the maximum lift coefficient
  2. Zero, because the pressure distribution is identical on both surfaces Correct answer.
  3. Negative, because the lower surface has higher velocity
  4. Positive, because the upper surface curvature generates suction

Explanation

A symmetrical section at zero angle of attack presents identical geometry to the upper and lower streams, so the pressure distributions cancel and no lift is produced. Lift appears only when the section is given incidence. Camber is what allows lift at zero angle.

Question 40Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

What is the significance of the laminar separation bubble on a low Reynolds number aerofoil?

  1. It is the stagnation region ahead of the leading edge
  2. The laminar layer separates before transition and reattaches as turbulent flow, altering drag and lift Correct answer.
  3. It is the region behind the trailing edge where the wake forms
  4. It is a region of fully attached laminar flow with no penalty

Explanation

At low Reynolds numbers the laminar boundary layer can separate before it transitions. The separated shear layer then becomes turbulent and reattaches, forming a bubble that increases drag and modifies the effective shape of the aerofoil. Its behaviour can cause abrupt changes in lift.

Question 41Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

Why does increasing the thickness of an aerofoil beyond an optimum tend to increase drag?

  1. The lift coefficient becomes negative at all angles
  2. The boundary layer becomes entirely laminar
  3. The stagnation point moves to the trailing edge
  4. The stronger adverse pressure gradient promotes earlier separation and a larger wake Correct answer.

Explanation

A thick section accelerates the flow strongly and then decelerates it over a longer rearward region, producing a steeper adverse pressure gradient. This encourages separation and a wider wake, raising pressure drag. Thickness therefore involves a trade-off between structural depth and drag.

Question 42Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

On a cambered aerofoil at a small positive angle of attack, how does the centre of pressure position change as the angle of attack is increased towards the stall?

  1. It moves forward towards the leading edge.
  2. It moves rearward towards the trailing edge. Correct answer.
  3. It moves rearward until the stall, then jumps forward.
  4. It remains fixed at the aerodynamic centre.

Explanation

On a cambered aerofoil the centre of pressure shifts rearward as the angle of attack rises, because the increasing lift is distributed further aft along the chord. The aerodynamic centre stays near the quarter-chord, but the centre of pressure does not coincide with it except at one angle of attack. This rearward movement continues until separation disturbs the pressure distribution.

Question 43Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

For an aerofoil generating lift, which statement about the stagnation point on the upper surface is correct?

  1. It is located at the point of maximum thickness on the upper surface.
  2. It lies at the trailing edge where the upper and lower streams meet.
  3. It moves rearward along the upper surface as the angle of attack increases.
  4. It moves around the leading edge onto the lower surface as the angle of attack increases. Correct answer.

Explanation

At the stagnation point the local velocity is zero and the pressure reaches its maximum. As the angle of attack increases, this point migrates around the leading edge onto the lower surface, which is why the suction peak on the upper surface grows. At the zero-lift angle it sits on the lower surface near the leading edge.

Question 44Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

In a two-dimensional incompressible flow over an aerofoil, what does Bernoulli's equation relate along a streamline?

  1. Reynolds number, Mach number and skin-friction coefficient.
  2. Density, temperature and velocity.
  3. Static pressure, dynamic pressure and total pressure. Correct answer.
  4. Lift coefficient, angle of attack and camber.

Explanation

Bernoulli's equation states that along a streamline the sum of static pressure and dynamic pressure remains constant, giving total pressure. Where the local velocity rises, the static pressure falls, which explains the suction over the upper surface. This relationship holds for steady, incompressible, frictionless flow along a streamline.

Question 45Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

At a given angle of attack, an aerofoil section reaches a Reynolds number at which the boundary layer becomes turbulent over most of both surfaces. What is the most likely effect on the skin-friction drag coefficient?

  1. It increases, because turbulent mixing raises the wall shear stress. Correct answer.
  2. It stays unchanged, because skin friction depends only on pressure.
  3. It decreases, because the turbulent layer is thinner.
  4. It becomes zero once transition is complete.

Explanation

A turbulent boundary layer has greater momentum exchange near the wall, so the velocity gradient at the surface is steeper and the local shear stress is higher. For the same angle of attack and lift, this raises the skin-friction drag coefficient compared with a largely laminar layer. The turbulent layer is also thicker, contrary to the idea that it is thinner.

Question 46Airflow and aerofoils: lift, drag, pressure distribution, boundary layer, separation, stalls

An aerofoil at a fixed angle of attack is tested in a wind tunnel. If the free-stream velocity is increased while the air density and temperature remain constant, how does the lift coefficient behave while the flow remains incompressible?

  1. It increases in proportion to the square of the velocity.
  2. It increases in proportion to the velocity.
  3. It decreases in proportion to the velocity.
  4. It remains approximately constant. Correct answer.

Explanation

The lift coefficient is a non-dimensional measure of the pressure distribution over the section and depends mainly on angle of attack, camber and Reynolds number. For incompressible flow at a fixed angle of attack, doubling the velocity doubles the lift but also quadruples the dynamic pressure, so the coefficient stays essentially the same. Only compressibility or Reynolds-number effects would cause a noticeable change.

Question 47Wings: planform, aspect ratio, wingtip devices, high-lift devices, sweep, dihedral and anhedral effects

Which parameter expresses the slenderness of a wing and is defined as span squared divided by reference wing area?

  1. Wing dihedral angle
  2. Wing taper ratio
  3. Wing aspect ratio Correct answer.
  4. Wing sweep angle

Explanation

Aspect ratio compares the wingspan with the mean chord, so a long narrow wing has a high aspect ratio while a short stubby wing has a low one. It strongly influences induced drag, because induced drag falls as aspect ratio increases for a given lift coefficient.

Question 48Wings: planform, aspect ratio, wingtip devices, high-lift devices, sweep, dihedral and anhedral effects

For a wing of the same span and area, increasing the aspect ratio will have which primary effect on the induced drag at a given lift coefficient?

  1. Induced drag is unchanged
  2. Induced drag increases
  3. Induced drag decreases Correct answer.
  4. Induced drag becomes zero

Explanation

Induced drag is proportional to the lift coefficient squared divided by pi times aspect ratio, so a higher aspect ratio produces less induced drag for the same lift. This is why gliders and long-range aircraft favour high aspect ratios.

Question 49Wings: planform, aspect ratio, wingtip devices, high-lift devices, sweep, dihedral and anhedral effects

What is the main aerodynamic benefit of fitting winglets to a transport aircraft wing?

  1. A reduction in the profile drag of the fuselage
  2. A reduction in the tip vortices and hence induced drag Correct answer.
  3. An increase in the critical Mach number of the wing
  4. An increase in the maximum usable angle of attack

Explanation

Winglets diffuse the strong vortex flow at the tip, weakening the tip vortices and reducing the induced drag component. This gives a small improvement in lift-to-drag ratio, mainly beneficial at high lift coefficients.

Question 50Wings: planform, aspect ratio, wingtip devices, high-lift devices, sweep, dihedral and anhedral effects

A wing has a span of 40 m and a reference area of 100 m². What is its aspect ratio?

  1. 16 Correct answer.
  2. 4
  3. 8
  4. 40

Explanation

Aspect ratio equals span squared divided by area, so 40 × 40 = 1600, and 1600 ÷ 100 = 16. A high value like this indicates a slender, efficient wing typical of a glider or long-range airliner.