JAR 66 CATEGORY B1 MODULE 11.01
1.2 HIGH SPEED FLIGHT
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1.2 HIGH SPEED FLIGHT
Advancement in modern aircraft and engine design has produced very large airliners capable of cruising at 87% of the speed of sound. Typically at an altitude of 11,000 metres (approximately 36,000feet), this will amount to an airspeed of about 575 miles per hour.
Earlier in the course the effects of subsonic air were considered. As airspeed increases, the aerodynamic effects of airflow passing over an aircraft, go through a series of changes, which will now be considered.
1.2.1 SPEED OF SOUND
One of the most important measurements in high speed aerodynamics is based on the speed of sound and so called mach number.
Mach number is named after the Austrian physicist Ernst Mach (1838-1916) and is the ratio of true airspeed of an aircraft to the local speed of sound at that altitude. (This will be covered in more detail later).
Sound waves, like those produced by a stationary object vibrating at certain frequencies, will cause a continuous series of pulses or pressure waves, to radiate outwards equally in all directions from the point of origin and travel in exactly the same manner as the ripples on a pond.
Pressure Waves – Stationary Object Figure 35
The actual speed at which the waves radiate, depends on the type and density of the material in which they are travelling. Air and Water are both fluids but water is more dense than air, so sound waves will travel faster (about 4 times) in water than in air.
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Additionally, in any one of the fluids, speed will vary with a change in temperature. As temperature increases, the speed of sound will increase and vice-versa, so that in Air on a standard day at sea level (15oC approx), the waves will travel at 761mph (661.7 knots), whereas at 11,000 metres altitude, the speed will fall to 661mph, since the temperature has dropped to -56oC at this altitude.
Note: At altitudes above 11,000 metres and up to about 27,000 metres, the temperature and hence the speed of sound, will remain constant.
1.2.2 SUBSONIC FLIGHT
The propagation of the pressure waves from a stationary object has been discussed above.
When an aircraft begins to move through the air at subsonic speeds, (a speed less than pressure wave propagation speed) the waves still travel forward and it is as if a message is sent ahead of the aircraft to warn of its approach.
On receipt of this message, the air streams begin to divide to make way for the aircraft but there is very little, if any change in the density of the air as it flows over the aircraft. This warning message can be detected perhaps 100metres in front of the aircraft.
Consequently, anyone standing ahead of the aircraft, would hear it coming and be able to detect the change in the nature of the pressure waves as the aircraft passed by. It would be similar to the change in the pitch of the siren of a passing emergency road vehicle.
This is often referred to as Doppler shift or Doppler effect.
Pressure waves – Subsonic Flight Figure 36
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
engineering uk
1.2.3 TRANSONIC FLIGHT
At subsonic speeds, the study of aerodynamics is simplified by the fact that air passing over a wing experiences only very small changes in pressure and density. The airflow is termed incompressible as, when it passes through a venturi, the pressure changes without the density changing
At higher speeds, the change in air pressure and density becomes significant and is called the compressibility effect. When air enters a venturi at supersonic speeds, the airflow slows down and must compress in order to pass through its throat. Once a fluid compresses, its pressure and density will both increase.
Subsonic Airflow Figure 37
Supersonic Airflow Figure 38
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The transonic flight range encompasses sound wave velocity and consequently is the most difficult realm of flight since some of the air flowing over the aircraft, particularly the wings, is subsonic and some is supersonic. As the aircraft approaches the speed of sound, the pressure waves ahead of it will be travelling at the same speed as the aircraft and are therefore relatively stationary. They accumulate to form a continuous pressure wave and consequently will result in the removal of any advance warning of the approach of the aircraft.
Transonic Flight Pressure Waves Figure 39
At these speeds other pressure waves, or shock waves form wherever the airflow reaches the speed of sound. These waves will upset the aerodynamic balance of the wing and this phenomenon will be covered later in the notes.
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
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1.2.4 SUPERSONIC FLIGHT
Once the aircraft is supersonic, all parts of it are considered to be above the speed of sound and therefore travelling faster than the rate of propagation of the pressure waves. An infinite number of pressure waves are produced and form a cone, the inclination of which will change as the aircraft speed changes.
Mach Cone Figure 40
1.2.4.1 Mach Number
As previously mentioned, Mach number is the ratio of the true airspeed of the aircraft and the local speed of sound at that altitude. An aircraft travelling at exactly the speed of sound is said to be travelling at Mach 1.
It follows therefore that an aircraft travelling at twice the speed of sound would be travelling at Mach 2 and at half the speed of sound, Mach 0.5, etc,.
The following definitions regarding airflow and mach number apply:
Subsonic Flow Mach Numbers below Mach 0.75
Transonic Flow Mach Numbers between Mach 0.75 and Mach 1.2 Supersonic Flow Mach Numbers between Mach 1.2 and 5.0 Hypersonic Flow Mach Numbers above Mach 5.0
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1.2.4.2 Critical Mach Number
At any constant aircraft forward speed, the speed of the airflow will vary over the curves and cambers on the different areas of the airframe. The behaviour of the airflow over the wing will be particularly significant, since this is the major lift provider for the aircraft.
As air flows over the camber on the upper surface of the wing, its speed will increase as it flows rearwards from the leading edge, reaching a maximum at the thickest part of the wing chord. This means that although the aircraft itself may be travelling at an airspeed well below Mach 1, the airflow over the thickest part of the wing chord, may have already reached Mach 1
As will be discussed later, many unwanted effects occur when the wing approaches and reaches Mach 1. Therefore, the designers may either incorporate features that will lessen the unwanted effects, or limit the aircraft to a predetermined maximum airspeed, that will ensure the wing speed remains below Mach 1 and thus avoids the unwanted effects altogether.
For each aircraft type therefore, a unique maximum aircraft forward speed will be calculated, corresponding to a wing speed of Mach 1. This aircraft speed (always be less than Mach 1) is called the Critical Mach Number or M.crit and non-supersonic aircraft flying in the transonic flight range, will normally be limited to a maximum speed set below the Critical Mach number.
Critical Mach Number Figure 41
A thick wing will cause the airflow to speed up over the camber and reach Mach 1 more quickly than a thin wing of similar chord length. Consequently, the Critical Mach number for the thinner wing will be a higher value than the thicker wing.
This in turn will mean that the aircraft with a thin wing, will be able to fly faster in the transonic flight range than the one with the thicker wing, before the unwanted effects caused by the wing reaching Mach 1 ensue.
Conversely, less lift will be produced by a thin wing, than a thick wing of similar chord length, but this can be overcome by the so called Supercritical wing chord.
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
engineering uk
In this design, the total amount of lift lost by the shallower camber of the thin wing is restored by making the chord longer. This is perfect for transonic cruise conditions, but at low airspeeds, lift on a clean wing will be insufficient and so extensive use of high lift devices (slots, slats and flaps) is necessary
Supercritical Wing Figure 42
1.2.4.3 Adverse Transonic Effects
Even though the onset of compressibility is gradual, it begins to have a significant effect as the Critical Mach number is approached. Unwanted adverse effects including, buffeting, shock waves, increase in drag, decrease in lift and movement of the centre of pressure occur.
If uncontrolled, these effects could result in the aircraft becoming difficult to fly and to behave in a similar manner to a low speed high incidence stall, even though the aircraft is at high speed and low angle of incidence.
1.2.4.4 Compressibility Buffet
Previously discussed has been the build up of the pressure wave in front of the aircraft as it approaches Mach 1, including the fact that other parts of the airframe, in particular the wing, are likely to reach Mach 1 well before the complete aircraft does.
When this occurs the smoothness of the airflow over the wing is severely affected. This region, as well as those on the flying control aerofoils, experience violent vibration and so-called compressibility buffeting of the airframe. If allowed to continue, control loss or possible structural damage can occur.
1.2.4.5 Shock Wave
Previously in the notes, the build up of pressure waves and the change from incompressible to compressible flow as the aircraft or an aerofoil surface approaches the speed of sound, has been discussed. Transonic flight presents major design problems for the aerofoil in particular, because only a portion of the airflow passing over the wing becomes supersonic.
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When an aerofoil moves through the air at a speed below its critical Mach number, all of the airflow is subsonic and the pressure distribution is predictable.The first indication of a change in the nature of the flow will be a breakaway of the airflow from the aerofoil surface as described previously in boundary layer control. Any turbulence resulting from the separation will cause an increase in drag and a corresponding reduction in the amount of lift. As speed begins to increase, the point of separation moves forward, extending the turbulent wake.
Subsonic Flow Over all the Surface Figure 43
However, as flight speed reaches and exceeds the critical Mach number, the airflow over the top of the wing speeds up to supersonic velocity and a shock wave starts to form.
The First Sonic Flow is encountered Figure 44
A Normal Shock Wave Begins to Form Figure 45
JAR 66 CATEGORY B1 about 3 degrees, even a symmetrical aerofoil section would produce the incipient wave on the top surface first.
The wave extends outwards more or less at right angles to the aerofoil surface and is referred to as a normal (perpendicular) shock wave This normal shock wave forms a boundary between supersonic and subsonic airflow.
As we have seen the high velocity airflow over the top of a wing creates an area of low pressure. The shock wave causes it to decelerate to subsonic speed, resulting in a rapid rise in pressure. The separation point and turbulent wake will now start from this point, resulting in a sudden and considerable increase in drag (about 10 times) and therefore a large loss of lift. Severe buffeting is likely, which could even lead to a shock stall and the centre of pressure will be altered, affecting the pitching moment.
This „extra‟ drag, so called Shock Drag, will be made up of two components, namely Wave Drag, resistance caused by the wave itself and Boundary Layer Drag, due to the increased turbulent region over the surface of the wing.
Furthermore, this shock-induced separation is likely to reduce flying control effectiveness
The velocity of the air leaving the shock wave remains supersonic, so both the static pressure and the density of the air increase adding to the high drag/ low lift condition. Additionally, some of the energy in the airstream will be dissipated in the form of heat.
As the aircraft speed continues to increase, the wave will extend outwards and begin to move aft towards the trailing edge of the wing. A second wave begins to form on the lower surface, as the airflow here also speeds up to supersonic velocity
Shock Induced Separation Occurs Figure 46
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As the airspeed reaches the upper end of the transonic range, both shock waves move aft, become stronger and will eventually attach to the wing's trailing edge.
Almost all Flow is Supersonic, Some Shock Induced Separation Figure 47
Further increases in forward speed will now result in the characteristic normal shock wave forming ahead of the aerofoil. This continuous wave, known as a Bow wave, will move towards and subsequently attach itself, to the leading edge of the wing. Once attached, all airflow over the wing will be supersonic and many of the unwanted transonic effects are eliminated.
The Bow Wave is starting to Form Figure 48
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
engineering uk
As can be seen in figure 49, the transonic region has a great affect on the lift and drag. Both values rise until Mach 0.81, when shock induced separation drastically reduces the coefficient of lift. As speed approaches Mach 0.99, a bow wave is forming and airflow over the wing is slowed to subsonic speeds, resulting in an increase in lift coefficient and a reduction of drag.
Lift / Drag Comparison at 2º Angle of Attack Figure 49
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1.2.5 AERODYNAMIC HEATING
One of the biggest problems of sustained supersonic flight is aerodynamic heating of the aircraft structure. An extreme example of aerodynamic heating might be a „shooting star‟, when its material overheats to the point of destruction, from the heat generated by friction heating with the earth's atmosphere.
In the commercial world, Concorde was probably the only airliner where aerodynamic heating presents a significant problem. When the aircraft was flown at Mach 2, the friction of the air passing around the aircraft heats the skin considerably even at altitudes in excess of 17,000 metres. The point of maximum heating is on the nose where the rise in temperature could reach 1750C.
As a precaution, a probe on the nose of the aircraft monitors the temperature during flight. When a reading of 1270C is reached, the flight deck is directed to reduce the speed to about Mach 1.8, to bring the temperature back within limits.
Concorde used conventional aluminium alloys in its construction. If future aircraft were required to travel within the atmosphere at even higher Mach numbers, other materials such as titanium alloy or stainless steel would need to be considered.
Concord Skin Temperature Figure 50
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
engineering uk
1.2.6 AREA RULE
Area rule is an aerodynamic technique used in the design of high-speed aircraft.
If drag is to be kept to a minimum at transonic speeds, aircraft must be slim, smooth and streamlined. In general terms it means that the wings, fuselage, empennage and other appendages have to be considered together when working out the total streamlining. This is necessary so that the cross-sectional area of successive „slices‟ of the aircraft from nose to tail, conform to those of a simple body of streamline shape.
Area rule is defined as: “For the minimum drag at the connections, (wing/fuselage), the variation of the aircraft‟s total cross-sectional area along its length, should approximate that of an ideal shape having minimum wave drag”.
Without area rule, the greatest frontal cross-sectional area of the fuselage would occur where the wings are attached to the fuselage. Therefore, one method of achieving area rule in this situation is to reduce the cross-sectional area of the fuselage, thereby cancelling out the increase caused by the wings.
Alternatively, the fuselage cross-section could be increased with the use of enlarged sections behind and in front of the wings to eliminate sudden changes in the cross-sectional area and achieve the same result.
Area Rule Figure 51
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1.2.7 FACTORS AFFECTING AIRFLOW IN ENGINE INTAKES OF HIGH SPEED AIRCRAFT
Engine intakes on aircraft that operate in the subsonic flight range only can be of almost any form.
The main criteria are that the airflow reaching the compressor stage of the engine during cruise ideally does not exceed Mach 0.5. This is normally achieved by the careful design of the intake ducts.
Obviously, if the aircraft never exceeds Mach 0.5, a parallel intake duct could be employed, but if the aircraft is to cruise at airspeeds in excess of this, yet below Mach 1, a divergent duct must be utilised to slow the airflow at the compressor down to Mach 0.5.
If the aircraft is designed to cruise above Mach 1, the air entering the intakes will be supersonic and will behave in accordance with the rules of supersonic flow. In this case a convergent duct would be necessary to slow down the airflow to the compressor.
However the aircraft must fly through the transonic range in order to reach supersonic speed so both types of duct will be necessary.
One way to overcome the problem is to have moveable doors that change the intake duct shape from divergent to convergent cross-section as the aircraft passes through Mach 1. See figure 52. This technique can be found on the intakes of Concorde.
Other methods to control airflow reaching the compressor is to make use of the fact that air passing through a shock wave slows down to a lower speed. This type of intake design is usually characterised by the „bullet fairing‟, which on some aircraft can translate in and out of the intake to reposition the shock wave during low or high supersonic flight speeds. See Figure 53
JAR 66 CATEGORY B1 MODULE 11.01 Theory of Flight
engineering uk
Intake Moveable doors Figure 52
„Bullet Fairing‟ Intake Figure 53
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1.2.8 EFFECTS OF SWEEPBACK ON CRITICAL MACH NUMBER
In order to fly at high speed in the transonic range without encountering the problems caused by the production of shock waves, the Critical Mach number needs to be as high as possible. As has already been shown, one way is to have as thin a wing as possible. This of course is an acceptable solution in theory, but in practice there will be structural integrity problems, such as wing loading, strength and flexibility.
Another way of raising the Critical Mach number without the structural limitations is by the use of swept wings. Sweepback not only delays the production of the shock wave, but reduces the severity of the shock stall should it occur. The theory behind this is that it is only the component of velocity over the wing chord that is responsible for the pressure distribution and so for causing the shock wave
Another way of raising the Critical Mach number without the structural limitations is by the use of swept wings. Sweepback not only delays the production of the shock wave, but reduces the severity of the shock stall should it occur. The theory behind this is that it is only the component of velocity over the wing chord that is responsible for the pressure distribution and so for causing the shock wave