HOME
        TheInfoList






A swept wing is a wing that angles either backward or occasionally forward from its root rather than in a straight sideways direction.

Swept wings have been flown since the pioneer days of aviation. Wing sweep at high speeds was first investigated in Germany as early as 1935 by Albert Betz and Adolph Busemann, finding application just before the end of the Second World War. It has the effect of delaying the shock waves and accompanying aerodynamic drag rise caused by fluid compressibility near the speed of sound, improving performance. Swept wings are therefore almost always used on jet aircraft designed to fly at these speeds. Swept wings are also sometimes used for other reasons, such as low drag, low observability, structural convenience or pilot visibility.

The term "swept wing" is normally used to mean "swept back", but variants include forward sweep, variable sweep wings and oblique wings in which one side sweeps forward and the other back. The delta wing is also aerodynamically a form of swept wing.

Design characteristics

For a wing of given span, sweeping it increases the length of the spars running along it from root to tip. This tends to increase weight and reduce stiffness. If the fore-aft chord of the wing also remains the same, the distance between leading and trailing edges reduces, reducing its ability to resist twisting (torsion) forces. A swept wing of given span and chord must therefore be strengthened and will be heavier than the equivalent unswept wing.

A swept wing typically angles backward from its root rather than forwards. Because wings are made as light as possible, they tend to flex under load. This aeroelasticity under aerodynamic load causes the tips to bend upwards in normal flight. Backwards sweep causes the tips to reduce their angle of attack as they bend, reducing their lift and limiting the effect. Forward sweep causes the tips to increase their angle of attack as they bend. This increases their lift causing further bending and hence yet more lift in a cycle which can cause a runaway structural failure. For this reason forward sweep is rare and the wing must be unusually rigid.

The characteristic "sweep angle" is normally measured by drawing a line from root to tip, typically 25% of the way back from the leading edge, and comparing that with the perpendicular to the longitudinal axis of the aircraft. Typical sweep angles vary from 0 for a straight-wing aircraft, to 45 degrees or more for fighters and other high-speed designs.

Aerodynamics

Subsonic and transonic flight

Albert Betz and Adolph Busemann, finding application just before the end of the Second World War. It has the effect of delaying the shock waves and accompanying aerodynamic drag rise caused by fluid compressibility near the speed of sound, improving performance. Swept wings are therefore almost always used on jet aircraft designed to fly at these speeds. Swept wings are also sometimes used for other reasons, such as low drag, low observability, structural convenience or pilot visibility.

The term "swept wing" is normally used to mean "swept back", but variants include forward sweep, variable sweep wings and oblique wings in which one side sweeps forward and the other back. The delta wing is also aerodynamically a form of swept wing.

For a wing of given span, sweeping it increases the length of the spars running along it from root to tip. This tends to increase weight and reduce stiffness. If the fore-aft chord of the wing also remains the same, the distance between leading and trailing edges reduces, reducing its ability to resist twisting (torsion) forces. A swept wing of given span and chord must therefore be strengthened and will be heavier than the equivalent unswept wing.

A swept wing typically angles backward from its root rather than forwards. Because wings are made as light as possible, they tend to flex under load. This aeroelasticity under aerodynamic load causes the tips to bend upwards in normal flight. Backwards sweep causes the tips to reduce their angle of attack as they bend, reducing their lift and limiting the effect. Forward sweep causes the tips to increase their angle of attack as they bend. This increases their lift causing further bending and hence yet more lift in a cycle which can cause a runaway structural failure. For this reason forward sweep is rare and the wing must be unusually rigid.

The characteristic "sweep angle" is normally measured by drawing a line from root to tip, typically 25% of the way back from the leading edge, and comparing that with the perpendicular to the longitudinal axis of the aircraft. Typical sweep angles vary from 0 for a straight-wing aircraft, to 45 degrees or more for fighters and other high-speed designs.

Aerodynamics

Subsonic and transonic flight

Yakovlev Yak-25 swept wing
aeroelasticity under aerodynamic load causes the tips to bend upwards in normal flight. Backwards sweep causes the tips to reduce their angle of attack as they bend, reducing their lift and limiting the effect. Forward sweep causes the tips to increase their angle of attack as they bend. This increases their lift causing further bending and hence yet more lift in a cycle which can cause a runaway structural failure. For this reason forward sweep is rare and the wing must be unusually rigid.

The characteristic "sweep angle" is normally measured by drawing a line from root to tip, typically 25% of the way back from the leading edge, and comparing that with the perpendicular to the longitudinal axis of the aircraft. Typical sweep angles vary from 0 for a straight-wing aircraft, to 45 degrees or more for fighters and other high-speed designs.

As an aircraft enters the transonic speeds just below the speed of sound, the pressure waves associated with subsonic flight converge and begin to impinge on the aircraft. As the pressure waves converge the air in front of the aircraft begins to compress. This creates a force known as wave drag. This wave drag increases steeply until the whole aircraft is supersonic and then reduces.

However, shock waves can form on some parts of an aircraft moving at less than the speed of sound. Low pressure regions around an aircraft cause the flow to accelerate, and at transonic speeds this local acceleration can exceed Mach 1. Localized supersonic flow must return to the freestream conditions around the rest of the aircraft, and as the flow enters an adverse pressure gradient in the aft section of the wing, a discontinuity emerges in the form of a shock wave as the air is forced to rapidly slow and return to ambient pressure.

With objects where there is a sudden reduction in profile/thickness and the local air expands rapidly to fill the space taken by the solid object or where a rapid angular change is imparted to the airflow causing a momentary increase of volume/decrease in density, an oblique shock wave is generated. This is why shock waves are often associated with the part of a fighter aircraft cockpit canopy with the highest local curvature, appearing immediately behind this point.

At the point where the density drops, the local speed of sound correspondingly drops and a shock wave can form. This is why in conventional wings, shock waves form first after the maximum Thickness/Chord and why all airliners designed for cruising in the transonic range (above M0.8) have supercritical wings that are flatter on top resulting in minimized angular change of flow to upper surface air. The angular change to the air that is normally part of lift generation is decreased and this lift reduction is compensated for by deeper curved lower surfaces accompanied by a reflex curve at the trailing edge. This results in a much weaker standing shock wave towards the rear of the upper wing surface and a corresponding increase in critical mach number.

Shock waves require energy to form. This energy is taken out of the aircraft, which has to supply extra thrust to make up for this energy loss. Thus the shocks are seen as a form of drag. Since the shocks form when the local air velocity reaches supersonic speeds, there is a certain "critical mach" speed where sonic flow first appears on the wing. There is a following point called the However, shock waves can form on some parts of an aircraft moving at less than the speed of sound. Low pressure regions around an aircraft cause the flow to accelerate, and at transonic speeds this local acceleration can exceed Mach 1. Localized supersonic flow must return to the freestream conditions around the rest of the aircraft, and as the flow enters an adverse pressure gradient in the aft section of the wing, a discontinuity emerges in the form of a shock wave as the air is forced to rapidly slow and return to ambient pressure.

With objects where there is a sudden reduction in profile/thickness and the local air expands rapidly to fill the space taken by the solid object or where a rapid angular change is imparted to the airflow causing a momentary increase of volume/decrease in density, an oblique shock wave is generated. This is why shock waves are often associated with the part of a fighter aircraft cockpit canopy with the highest local curvature, appearing immediately behind this point.

At the point where the density drops, the local speed of sound correspondingly drops and a shock wave can form. This is why in conventional wings, shock waves form first after the maximum Thickness/Chord and why all airliners designed for cruising in the transonic range (above M0.8) have supercritical wings that are flatter on top resulting in minimized angular change of flow to upper surface air. The angular change to the air that is normally part of lift generation is decreased and this lift reduction is compensated for by deeper curved lower surfaces accompanied by a reflex curve at the trailing edge. This results in a much weaker standing shock wave towards the rear of the upper wing surface and a corresponding increase in critical mach number.

Shock waves require energy to form. This energy is taken out of the aircraft, which has to supply extra thrust to make up for this energy loss. Thus the shocks are seen as a form of drag. Since the shocks form when the local air velocity reaches supersonic speeds, there is a certain "critical mach" speed where sonic flow first appears on the wing. There is a following point called the drag divergence mach number where the effect of the drag from the shocks becomes noticeable. This is normally when the shocks start generating over the wing, which on most aircraft is the largest continually curved surface, and therefore the largest contributor to this effect.

Sweeping the wing has the effect of reducing the curvature of the body as seen from the airflow, by the cosine of the angle of sweep. For instance, a wing with a 45 degree sweep will see a reduction in effective curvature to about 70% of its straight-wing value. This has the effect of increasing the critical Mach by 30%. When applied to large areas of the aircraft, like the wings and empennage, this allows the aircraft to reach speeds closer to Mach 1.

One of the simplest and best explanations of how the swept wing works was offered by Robert T. Jones: "Suppose a cylindrical wing (constant chord, incidence, etc.) is placed in an airstream at an angle of yaw – i.e., it is swept back. Now, even if the local speed of the air on the upper surface of the wing becomes supersonic, a shock wave cannot form there because it would have to be a sweptback shock – swept at the same angle as the wing – i.e., it would be an oblique shock. Such an oblique shock cannot form until the velocity component normal to it becomes supersonic."[1]

One limiting factor in swept wing design is the so-called "middle effect". If a swept wing is continuous – an oblique swept wing, the pressure iso-bars will be swept at a continuous angle from tip to tip. However, if the left and right halves are swept back equally, as is common practice, the pressure iso-bars on the left wing in theory will meet the pressure iso-bars of the right wing on the centerline at a large angle. As the iso-bars cannot meet in such a fashion, they will tend to curve on each side as they near the centerline, so that the iso-bars cross the centerline at right angles to the centerline. This causes an "unsweeping" of the iso-bars in the wing root region. To combat this unsweeping, German aerodynamicist Dietrich Küchemann proposed and had tested a local indentation of the fuselage above and below the wing root. This proved to not be very effective.[2] During the development of the Douglas DC-8 airliner, uncambered airfoils were used in the wing root area to combat the unsweeping.[3][4] Similarly, a decambered wing root glove was added to the Boeing 707 wing to create the Boeing 720.[5]

Airflow at supersonic speeds generates lift through the formation of shock waves, as opposed to the patterns of airflow over and under the wing. These shock waves, as in the transonic case, generate large amounts of drag. One of these shock waves is created by the leading edge of the wing, but contributes little to the lift. In order to minimize the strength of this shock it needs to remain "attached" to the front of the wing, which demands a very sharp leading edge. To better shape the shocks that will contribute to lift, the rest of an ideal supersonic airfoil is roughly diamond-shaped in cross-section. For low-speed lift these same airfoils are very inefficient, leading to poor handling and very high landing speeds.[6]

One way to avoid the need for a dedicated supersonic wing is to use a highly swept subsonic design. Airflow behind the shock waves of a moving body are reduced to subsonic speeds. This effect is used within the intakes of engines meant to operate in the supersonic, as jet engines are generally incapable of ingesting supersonic air directly. This can also be used to reduce the speed of the air as seen by the wing, using the shocks generated by the nose of the aircraft. As long as the wing lies behind the cone-shaped shock wave, it will "see" subsonic airflow and work as normal. The angle needed to lie behind the cone increases with increasing speed, at Mach 1.3 the angle is about 45 degrees, at Mach 2.0 it is 60 degrees.[7] For instance, at Mach 1.3 the angle of the Mach cone formed off the body of the aircraft will be at about sinμ = 1/M (μ is the sweep angle of the Mach cone)[8]

Generally it is not possible to arrange the wing so it will lie entirely outside the supersonic airflow and still have good subsonic performance. Some aircraft, like the English Electric Lightning or Convair F-106 Delta Dart are tuned almost entirely for high-speed flight and feature highly swept wings with little regard to the low-speed problems this creates. In other cases the use of variable geometry wings, as on the Grumman F-14 Tomcat, allows an aircraft to move the wing to keep it at the most efficient angle regardless of speed, although the cost in complexity and weight makes this a rare feature.

Most high-speed aircraft have a wing that spends at least some of its time in the supersonic airflow. But since the shock cone moves towards the fuselage with increased speed (that is, the cone becomes narrower), the portion of the wing in the supersonic flow also changes with speed. Since these wings are swept, as the shock cone moves inward, the [7] For instance, at Mach 1.3 the angle of the Mach cone formed off the body of the aircraft will be at about sinμ = 1/M (μ is the sweep angle of the Mach cone)[8]

Generally it is not possible to arrange the wing so it will lie entirely outside the supersonic airflow and still have good subsonic performance. Some aircraft, like the English Electric Lightning or Convair F-106 Delta Dart are tuned almost entirely for high-speed flight and feature highly swept wings with little regard to the low-speed problems this creates. In other cases the use of variable geometry wings, as on the Grumman F-14 Tomcat, allows an aircraft to move the wing to keep it at the most efficient angle regardless of speed, although the cost in complexity and weight makes this a rare feature.

Most high-speed aircraft have a wing that spends at least some of its time in the supersonic airflow. But since the shock cone moves towards the fuselage with increased speed (that is, the cone becomes narrower), the portion of the wing in the supersonic flow also changes with speed. Since these wings are swept, as the shock cone moves inward, the lift vector moves forward[citation needed] as the outer, rearward portions of the wing are generating less lift. This results in powerful pitching moments and their associated required trim changes.

When a swept wing travels at high speed, the airflow has little time to react and simply flows over the wing almost straight from front to back. At lower speeds the air does have time to react, and is pushed spanwise by the angled leading edge, towards the wing tip. At the wing root, by the fuselage, this has little noticeable effect, but as one moves towards the wingtip the airflow is pushed spanwise not only by the leading edge, but the spanwise moving air beside it. At the tip the airflow is moving along the wing instead of over it, a problem known as spanwise flow.

The lift from a wing is generated by the airflow over it from front to rear. With increasing span-wise flow the boundary layers on the surface of the wing have longer to travel, and so are thicker and more susceptible to transition to turbulence or flow separation, also the effective aspect ratio of the wing is less and so air "leaks" around the wing tips reducing their effectiveness. The spanwise flow on swept wings produces airflow that moves the stagnation point on the leading edge of any individual wing segment further beneath the leading edge, increasing effective angle of attack of wing segments relative to its neighbouring forward segment. The result is that wing segments farther towards the rear operate at increasingly higher angles of attack promoting early stall of those segments. This promotes tip stall on back swept wings, as the tips are most rearward, while delaying tip stall for forward swept wings, where the tips are forward. With both forward and back swept wings, the rear of the wing will stall first. This creates a nose-up pressure on the aircraft. If this is not corrected by the pilot it causes the plane to pitch up, leading to more of the wing stalling, leading to more pitch up, and so on. This problem came to be known as the Sabre dance in reference to the number of North American F-100 Super Sabres that crashed on landing as a result.

The solution to this problem took on many forms. One was the addition of a fin known as a wing fence on the upper surface of the wing to redirect the flow to the rear (see the MiG-15 as an example.) Another closely related design was addition of a dogtooth notch to the leading edge (Avro Arrow). Other designs took a more radical approach, including the Republic XF-91 Thunderceptor's wing that grew wider towards the tip to provide more lift at the tip. The Handley Page Victor had a crescent wing with substantial sweep-back near the wing root where the wing was thickest, and progressively reducing sweep along the span as the wing thickness reduced towards the tip.

Modern solutions to the problem no longer require "custom" designs such as these. The addition of leading edge slats and large compound angle of attack of wing segments relative to its neighbouring forward segment. The result is that wing segments farther towards the rear operate at increasingly higher angles of attack promoting early stall of those segments. This promotes tip stall on back swept wings, as the tips are most rearward, while delaying tip stall for forward swept wings, where the tips are forward. With both forward and back swept wings, the rear of the wing will stall first. This creates a nose-up pressure on the aircraft. If this is not corrected by the pilot it causes the plane to pitch up, leading to more of the wing stalling, leading to more pitch up, and so on. This problem came to be known as the Sabre dance in reference to the number of North American F-100 Super Sabres that crashed on landing as a result.

The solution to this problem took on many forms. One was the addition of a fin known as a wing fence on the upper surface of the wing to redirect the flow to the rear (see the MiG-15 as an example.) Another closely related design was addition of a dogtooth notch to the leading edge (Avro Arrow). Other designs took a more radical approach, including the Republic XF-91 Thunderceptor's wing that grew wider towards the tip to provide more lift at the tip. The Handley Page Victor had a crescent wing with substantial sweep-back near the wing root where the wing was thickest, and progressively reducing sweep along the span as the wing thickness reduced towards the tip.

Modern solutions to the problem no longer require "custom" designs such as these. The addition of leading edge slats and large compound flaps to the wings has largely resolved the issue. On fighter designs, the addition of leading edge extensions, included for high maneuverability, also serve to add lift during landing and reduce the problem.

The swept wing also has several more problems. One is that for any given length of wing, the actual span from tip-to-tip is shorter than the same wing that is not swept. Low speed drag is strongly correlated with the aspect ratio, the span compared to chord, so a swept wing always has more drag at lower speeds. Another concern is the torque applied by the wing to the fuselage, as much of the wing's lift lies behind the point where the wing root connects to the plane. Finally, while it is fairly easy to run the main spars of the wing right through the fuselage in a straight wing design to use a single continuous piece of metal, this is not possible on the swept wing because the spars will meet at an angle.

Sweep theory is an aeronautical engineering description of the behavior of airflow over a wing when the wing's leading edge encounters the airflow at an oblique angle. The development of sweep theory resulted in the swept wing design used by most modern jet aircraft, as this design performs more effectively at transonic and supersonic speeds. In its advanced form, sweep theory led to the experimental oblique wing concept.

Adolf Busemann introduced the concept of the swept wing and presented this in 1935 at the 5. Volta-Congress in Rome. Sweep theory in general was a subject of development and investigation throughout the 1930s and 1940s, but the breakthrough mathematic

Adolf Busemann introduced the concept of the swept wing and presented this in 1935 at the 5. Volta-Congress in Rome. Sweep theory in general was a subject of development and investigation throughout the 1930s and 1940s, but the breakthrough mathematical definition of sweep theory is generally credited to NACA's Robert T. Jones in 1945. Sweep theory builds on other wing lift theories. Lifting line theory describes lift generated by a straight wing (a wing in which the leading edge is perpendicular to the airflow). Weissinger theory describes the distribution of lift for a swept wing, but does not have the capability to include chordwise pressure distribution. There are other methods that do describe chordwise distributions, but they have other limitations. Jones' sweep theory provides a simple, comprehensive analysis of swept wing performance.

To visualize the basic concept of simple sweep theory, consider a straight, non-swept wing of infinite length, which meets the airflow at a perpendicular angle. The resulting air pressure distribution is equivalent to the length of the wing's chord (the distance from the leading edge to the trailing edge). If we were to begin to slide the wing sideways (spanwise), the sideways motion of the wing relative to the air would be added to the previously perpendicular airflow, resulting in an airflow over the wing at an angle to the leading edge. This angle results in airflow traveling a greater distance from leading edge to trailing edge, and thus the air pressure is distributed over a greater distance (and consequently lessened at any particular point on the surface).

This scenario is identical to the airflow experienced by a swept wing as it travels through the air. The airflow over a swept wing encounters the wing at an angle. That angle can be broken down into two vectors, one perpendicular to the wing, and one parallel to the wing. The flow parallel to the wing has no effect on it, and since the perpendicular vector is shorter (meaning slower) than the actual airflow, it consequently exerts less pressure on the wing. In other words, the wing experiences airflow that is slower - and at lower pressures - than the actual speed of the aircraft.

One of the factors that must be taken into account when designing a high-speed wing is compressibility, which is the effect that acts upon a wing as it approaches and passes through the speed of sound. The significant negative effects of compressibility made it a prime issue with aeronautical engineers. Sweep theory helps mitigate the effects of compressibility in transonic and supersonic aircraft because of the reduced pressures. This allows the mach number of an aircraft to be higher than that actually experienced by the wing.

There is also a negative aspect to sweep theory. The lift produced by a wing is directly related to the speed of the air over the wing. Since the airflow speed experienced by a swept wing is lower than what the actual aircraft speed is, this becomes a problem during slow-flight phases, such as takeoff and landing. There have been various ways of addressing the problem, including the variable-incidence wing design on the Vought F-8 Crusader and swing wings on aircraft such as the F-14, F-111, and the Panavia Tornado.

The term "swept wing" is normally used to mean "swept back", but other swept variants include forward sweep, variable sweep wings and oblique wings in which one side sweeps forward and the other back. The delta wing also incorporates the same advantages as part of its layout.

Forward sweep

Sweeping a wing forward has approximately the same effect as rearward in terms of drag reduction, but has other advantages in terms of low-speed handling where tip stall problems simply go away. In this case the low-speed air flows towards the fuselage, which acts as a very large wing fence. Additionally, wings are generally larger at the root anyway, which allows them to have better low-speed lift.

However, this arrangement also has serious stability problems. The rearmost section of the wing will stall first causing a pitch-up moment pushing the aircraft further into stall similar to a swept back wing design. Thus swept-forward wings are unstable in a fashion similar to the low-speed problems of a conventional swept wing. However unlike swept back wings, the tips on a forward swept design will stall last, maintaining roll control.

Forward-swept wings can also experience dangerous flexing effects compared to aft-swept wings that can negate the tip stall advantage if the wing is not sufficiently stiff. In aft-swept designs, when the airplane maneuvers at high load factor the wing loading and geometry twists the wing in such a way as to create washout (tip twists leading edge down). This reduces the angle of attack at the tip, thus reducing the bending moment on the wing, as well as somewhat reducing the chance of tip stall.[9] However, the same effect on forward-swept wings produces a wash-in effect that increases the angle of attack promoting tip stall.

Small amounts of sweep do not cause serious problems, and had been used on a variety of aircraft to move the spar into a convenient location, as on the Junkers Ju 287 or HFB 320 Hansa Jet. But larger sweep suitable for high-speed aircraft, like fighters, was generally impossible until the introduction of fly by wire systems that could react quickly enough to damp out these instabilities. The Grumman X-29 was an experimental technology demonstration project designed to test the forward swept wing for enhanced maneuverability in 1984. The Su-47 Berkut is another notable example using this technology. However no highly swept-forward design

However, this arrangement also has serious stability problems. The rearmost section of the wing will stall first causing a pitch-up moment pushing the aircraft further into stall similar to a swept back wing design. Thus swept-forward wings are unstable in a fashion similar to the low-speed problems of a conventional swept wing. However unlike swept back wings, the tips on a forward swept design will stall last, maintaining roll control.

Forward-swept wings can also experience dangerous flexing effects compared to aft-swept wings that can negate the tip stall advantage if the wing is not sufficiently stiff. In aft-swept designs, when the airplane maneuvers at high load factor the wing loading and geometry twists the wing in such a way as to create washout (tip twists leading edge down). This reduces the angle of attack at the tip, thus reducing the bending moment on the wing, as well as somewhat reducing the chance of tip stall.[9] However, the same effect on forward-swept wings produces a wash-in effect that increases the angle of attack promoting tip stall.

Small amounts of sweep do not cause serious problems, and had been used on a variety of aircraft to move the spar into a convenient location, as on the Junkers Ju 287 or HFB 320 Hansa Jet. But larger sweep suitable for high-speed aircraft, like fighters, was generally impossible until the introduction of fly by wire systems that could react quickly enough to damp out these instabilities. The Grumman X-29 was an experimental technology demonstration project designed to test the forward swept wing for enhanced maneuverability in 1984. The Su-47 Berkut is another notable example using this technology. However no highly swept-forward design has entered production.

The first successful aeroplanes adhered to the basic design of rectangular wings at right angles to the body of the machine, but there were experimentalists who explored other geometries to achieve better aerodynamic results. The swept wing geometry appeared before World War I, and was conceived as a means of permitting the design of safe and stable aeroplanes. The best of these designs imposed "self-damping" inherent stability upon a tailless swept wing. These inspired several flying wing gliders and some powered aircraft during the interwar years.[10]

J. W. Dunne who was obsessed with achieving inherent stability in flight. He successfully employed swept wings in his tailless aircraft (which, crucially, used washout) as a means of creating positive longitudinal static stability.[11] For a low-speed aircraft, swept wings may be used to resolve problems with the center of gravity, to move the wing spar into a more convenient location, or to improve the sideways view from the pilot's position.[10] By 1905 he had already built a model glider with swept wings and by 1913 he had constructed successful powered variants that were able to cross the English Channel. The Dunne D.5 was exceptionally aerodynamically stable for the time and the D.8 was sold to the Royal Flying Corps and, manufactured under licence by Starling Burgess, to the United States Navy among others.

His work ceased with the onset of war in 1914, but afterwards the idea was taken up by G. T. R. Hill in England who designed a series of gliders and aircraft to Dunne's guidelines, notably the Westland-Hill Pterodactyl series. However, Dunne's theories met with little acceptance from the leading aircraft designers and companies at the time.[12]

Development