Aeroscience

Requires the pilot to stabilise the aircraft manually

Uses a system of wires connected to pulleys linked to the pilot’s controls

Converts the movements of flight controls to electronic signals that are transmitted by wires, and uses flight control computers to determine how to move the actuators at each control surface to provide the ordered response

Uses a system consisting of three gyroscopes and three accelerometers

1. Also known as a high-speed stall

2. The breakdown in lift as the result of a shock wave

3. When an aerofoil exceeds the critical angle and there is insufficient lift to support the aircraft's weight, causing the aircraft to stall

4. Both 1 and 2 are correct

Lower

Higher

decreases, increases, remain constant

decreases, increases, decrease

increases, decreases, remain constant

increases, decreases, decrease

decreases, increases, increase

increases, decreases, remain constant

increases, decreases, increase

decreases, increases, remain constant

decreases, increases, increase

increases, decreases, decrease

increases, decreases, increase

decreases, increases, decrease

decreases, increases, decrease

decreases, increases, increase

increases, decreases, increase

increases, decreases, decrease

Beyond MFS 2.0

MCRIT to 1.0

MFS 1.1–1.5

MCRIT

When the TAS and MCRIT/MMO arrive at the same value

When the stall speed and MCRIT/MMO arrive at the same value

When the stall speed and TAS arrive at the same value

When the TAS and LSS arrive at the same value

The Mach number relative to the local speed of sound measured at a specific point on the aircraft. Depending on where this measurement is taken, this speed may be greater or less than the free-stream Mach number.

The Mach number of the airflow at a point unaffected by the pressure of the aircraft but measured relative to it

The Mach number of the airflow at a point where the drag coefficient rapidly rises. This is due to the flow separation and adverse pressure gradient behind the shock wave.

The lowest free-stream Mach number at which a local Mach number of 1.0 will occur at any point on the aircraft

It must have the ability to return to normal following a stall.

It must have the ability to return to normal following a disturbance about the vertical axis.

It must have the ability to return to normal following a disturbance about the longitudinal axis.

It must have the ability to return to normal following a disturbance about the lateral axis.

Winglets

Wing fences

Leading-edge sawtooth

All of these

increase; decrease

decrease; remain constant

remain constant; decrease

decrease; decrease

The decreased turbulent flow behind the shock wave. It strikes the tailplane section with enough force to buffet the controls.

An aerofoil exceeding the critical angle. The airflow breaks away from the aerofoil and strikes the tailplane.

A low-speed buffet

The increased turbulent flow behind the shock wave. It strikes the tailplane section with enough force to buffet the controls.

The ratio of the indicated airspeed (IAS) to the local speed of sound (LSS)

The ratio of the calibrated airspeed (CAS) to the local speed of sound (LSS)

The ratio of the local speed of sound (LSS) to the true airspeed (TAS)

The ratio of the true airspeed (TAS) to the local speed of sound (LSS)

As an aircraft accelerates through the transonic range, the centre of pressure moves rearwards in conjunction with the aft-moving spread of supersonic flow. From a starting position of approximately 15% of the chord, the C of P may end up at 20% chord, well behind the point of maximum camber (depending on the wing). This rearward shift of the C of P increases the moment arm between the centre of pressure and the centre of gravity, causing a nose-down pitching moment.

As an aircraft accelerates through the transonic range, the centre of pressure moves forwards in conjunction with the aft-moving spread of supersonic flow. From a starting position of approximately 25% of the chord, the C of P may end up at 50% chord, well behind the point of maximum camber (depending on the wing). This forward shift of the C of P increases the moment arm between the centre of pressure and the centre of gravity, causing a nose-down pitching moment.

As an aircraft accelerates through the transonic range, the centre of pressure moves rearwards in conjunction with the aft-moving spread of supersonic flow. From a starting position of approximately 25% of the chord, the C of P may end up at 50% chord, well behind the point of maximum camber (depending on the wing). This rearward shift of the C of P increases the moment arm between the centre of pressure and the centre of gravity, causing a nose-down pitching moment.

As an aircraft accelerates through the transonic range, the centre of pressure moves rearwards in conjunction with the aft-moving spread of supersonic flow. From a starting position of approximately 5% of the chord, the C of P may end up at 70% chord, well behind the point of maximum camber (depending on the wing). This rearward shift of the C of P increases the moment arm between the centre of pressure and the centre of gravity, causing a nose-down pitching moment.

38.94 × √T°K

39.84 × √T°K

38.94 × √T°C

39.84 × √T°C

The aircraft should be as large as possible.

The aircraft should be as small as possible.

The aerodynamic shape of the surface in question should change in cross-sectional area as sharply as possible, increasing rapidly to a maximum, then decreasing just as rapidly.

The aerodynamic shape of the surface in question should change in cross-sectional area as smoothly as possible, increasing gradually to a maximum, then decreasing just as gradually.

A higher critical angle

A lower critical angle

The decreased critical Mach number

The increased critical Mach number

Air density

Air temperature

Air pressure

Air pollution

The distance travelled by a sound wave propagating through the air

The distance travelled by a light wave propagating through a given medium

The distance travelled by a sound wave propagating through a given medium

The distance travelled by a sound wave propagating through the water

Mach 0.8 and below (approx. 530 kt)

Mach 0.8 and 1.2 (approx. 530–790 kt)

Mach 1.2 and 5.0 (approx. 790–300 kt)

Mach 5.0 and 10.0 (approx. 3,300–6,600 kt)

Critical drag rise Mach number (also known as drag divergence Mach number, or MCDR) is reached. This is the Mach number where the amount of drag experienced rapidly rises. The large increase in drag is caused by the formation of a shock wave on the upper surface of the airfoil.

The MFS reaches Mach 1.0.

MDET is reached. This is the speed where all airflow over the aircraft is supersonic. The amount of drag is slightly higher than the subsonic range due to the presence of the bow wave.

MCRIT is reached.

A form of induced drag

The drag caused by the turbulent boundary layer generated by the shock wave.

Typically seen on aircraft flying in the subsonic speed range

Typically seen on aircraft flying in the supersonic speed range

MDET is reached, where all airflow over the wing is supersonic (>M1.0). The detached bow wave has formed and interferes with the airflow over the wing. Although a stable flow, lift is slightly reduced.

Both shock waves, above and below the wing, have reached the trailing edge. The airflow over the wing is now supersonic, and a stable laminar flow is aiding in the production of lift.

CL increases with an increase in speed. The airflow accelerates over the wing, increasing lift. MCRIT is reached where the airflow over the aircraft is now supersonic in some places

Shock waves formed above and below the wing interfere with the airflow behind the shock wave. This causes separation of the air from the wing behind the shock wave, creating drag and the reduction in lift.

Reduces the critical Mach number. When you increase angle of attack, you are effectively increasing the velocity of the air over the upper section of the aerofoil. As a result, the first shock wave (normal shock wave) develops at a lower free-stream Mach number, and it is for this reason that the MCRIT is reached at a lower Mach number.

Increases the critical Mach number. When you decrease angle of attack, you are effectively increasing the velocity of the air over the upper section of the aerofoil. As a result, the first shock wave (normal shock wave) develops at a lower free-stream Mach number, and it is for this reason that the MCRIT is reached at a higher Mach number.

Reduces critical Mach number. When you increase angle of attack, you are effectively decreasing the velocity of the air over the upper section of the aerofoil. As a result, the first shock wave (normal shock wave) develops at a lower free-stream Mach number, and it is for this reason that the MCRIT is reached at a lower Mach number.

Increases the critical Mach number. When you increase angle of attack, you are effectively increasing the velocity of the air over the upper section of the aerofoil. As a result, the first shock wave (normal shock wave) develops at a lower free-stream Mach number, and it is for this reason that the MCRIT is reached at a higher Mach number.

The point where the aircraft will change from climbing at constant Mach to constant IAS

Where a descending aircraft would change from constant IAS to constant Mach

The point at which a pilot would change the subscale on the altimeters from area QNH to standard 1013 hPa

Where a climbing aircraft would change from climbing at constant IAS to constant Mach

Mach would decrease; TAS would decrease

Mach would increase; TAS would increase

Mach would increase; TAS would decrease

Mach would decrease; TAS would increase

The Mach number relative to the local speed of sound measured at a specific point on the aircraft. Depending on where this measurement is taken, this speed may be greater or less than the free-stream Mach number.

The Mach number of the airflow at a point unaffected by the pressure of the aircraft but measured relative to it

The Mach number of the airflow at a point where the drag coefficient rapidly rises. This is due to the flow separation and adverse pressure gradient behind the shock wave.

The lowest free-stream Mach number at which a local Mach number of 1.0 will occur at any point on the aircraft

The lowest free-stream Mach number at which a local Mach number of 1.0 will occur at any point on the aircraft

The Mach number of the airflow at a point where the drag coefficient rapidly rises. This is due to the flow separation and adverse pressure gradient behind the shock wave.

The Mach number of the airflow at a point unaffected by the pressure of the aircraft but measured relative to it

The Mach number relative to the local speed of sound measured at a specific point on the aircraft. Depending on where this measurement is taken, this speed may be greater or less than the free-stream Mach number.

The Mach number relative to the local speed of sound measured at a specific point on the aircraft. Depending on where this measurement is taken, this speed may be greater or less than the free-stream Mach number.

The Mach number of the airflow at a point unaffected by the pressure of the aircraft but measured relative to it

The Mach number of the airflow at a point where the drag coefficient rapidly rises. This is due to the flow separation and adverse pressure gradient behind the shock wave.

The lowest free-stream Mach number at which a local Mach number of 1.0 will occur at any point on the aircraft

The Mach number where all of the airflow over the wing is supersonic. This usually occurs around MFS 1.2 to 1.3.


The lowest free-stream Mach number at which a local Mach number of 1.0 will occur at any point on the aircraft

The Mach number of the airflow at a point where the drag coefficient rapidly rises. This is due to the flow separation and adverse pressure gradient behind the shock wave.

The Mach number relative to the local speed of sound measured at a specific point on the aircraft. Depending on where this measurement is taken, this speed may be greater or less than the free-stream Mach number.

Mach 0.8 and below (approx. 530 kt)

Mach 0.8 and 1.2 (approx. 530–790 kt)

Mach 1.0 to Mach 5.0 (approx. 790–3,300 kt)

Mach 5.0 and 10.0 (approx. 3,300–6,600 kt)

MDET is reached, where all airflow over the wing is supersonic (>M1.0). The detached bow wave has formed and interferes with the airflow over the wing. Although a stable flow, lift is slightly reduced.

Both shock waves, above and below the wing, have reached the trailing edge. The airflow over the wing is now supersonic, and a stable laminar flow is aiding in the production of lift.

CL increases with an increase in speed. The airflow accelerates over the wing, increasing lift. MCRIT is reached where the airflow over the aircraft is now supersonic in some places

Shock waves formed above and below the wing interfere with the airflow behind the shock wave. This causes separation of the air from the wing behind the shock wave, creating drag and the reduction in lift.

MDET is reached, where all airflow over the wing is supersonic (>M1.0). The detached bow wave has formed and interferes with the airflow over the wing. Although a stable flow, lift is slightly reduced.

Both shock waves, above and below the wing, have reached the trailing edge. The airflow over the wing is now supersonic, and a stable laminar flow is aiding in the production of lift.

CL increases with an increase in speed. The airflow accelerates over the wing, increasing lift. MCRIT is reached where the airflow over the aircraft is now supersonic in some places

Shock waves formed above and below the wing interfere with the airflow behind the shock wave. This causes separation of the air from the wing behind the shock wave, creating drag and the reduction in lift.

CL increases with an increase in speed. The airflow accelerates over the wing, increasing lift. MCRIT is reached where the airflow over the aircraft is now supersonic in some places

Shock waves formed above and below the wing interfere with the airflow behind the shock wave. This causes separation of the air from the wing behind the shock wave, creating drag and the reduction in lift.

Both shock waves, above and below the wing, have reached the trailing edge. The airflow over the wing is now supersonic, and a stable laminar flow is aiding in the production of lift.

MDET is reached, where all airflow over the wing is supersonic (>M1.0). The detached bow wave has formed and interferes with the airflow over the wing. Although a stable flow, lift is slightly reduced.

The MFS reaches Mach 1.0.

MDET is reached. This is the speed where all airflow over the aircraft is supersonic. The amount of drag is slightly higher than the subsonic range due to the presence of the bow wave.

Critical drag rise Mach number (also known as drag divergence Mach number, or MCDR) is reached. This is the Mach number where the amount of drag experienced rapidly rises. The large increase in drag is caused by the formation of a shock wave on the upper surface of the airfoil.

MCRIT is reached.

MCRIT is reached

Critical drag rise Mach number (also known as drag divergence Mach number, or MCDR) is reached. This is the Mach number where the amount of drag experienced rapidly rises. The large increase in drag is caused by the formation of a shock wave on the upper surface of the airfoil.

The MFS reaches Mach 1.0.

MDET is reached. This is the speed where all airflow over the aircraft is supersonic. The amount of drag is slightly higher than the subsonic range due to the presence of the bow wave.

Increase, Increase

Increase, Decrease

Decrease, Decrease

Decrease, Increase

stronger lateral stability with weaker directional stability

stronger lateral stability and strong directional stability

weaker lateral stability with stronger directional stability

weak lateral stability with weaker directional stability