TRANSONIC AND SUPERSONIC CONFIGU­RATIONS

Aircraft configurations developed for high speed flight will have significant differences in shape and planform when compared with air­craft designed for low speed flight. One of the outstanding differences will be in the selection of airfoil profiles for transonic or supersonic flight.

AIRFOIL SECTIONS. It should be ob­vious that airfoils for high speed subsonic flight should have high critical Mach num­bers since critical Mach number defines the lower limit for shock wave formation and subsequent force divergence. An additional complication to airfoil selection in this speed range is that the airfoil should have a high maximum lift coefficient and sufficient thickness to allow application of high lift devices. Otherwise an excessive wing area would be required to provide maneuverability and reasonable takeoff and landing speeds.

Revised January 19Л5

F6U MODEL AT VARIOUS MACH NUMBERS в «0® $’0°

Figure 3.11. Schlieren Photographs of Supersonic Flight (sheet 1 of 2)

However, if high speed flight is the primary consideration, the airfoil must be chosen to have the highest practical critical Mach number.

Critical Mach number has been defined as the flight Mach number which produces first evidence of local sonic flow. Thus, the air­foil shape and lift coefficient—which determine the pressure and velocity distribution—will have a profound effect on critical Mach number. Conventional, low speed airfoil shapes have relatively poor compressibility characteristics because of the high local velocities near the leading edge. These high local velocities are inevitable if both the maximum thickness and camber are well forward on the chord. An improvement of the compressibility character­istics can be obtained by moving the points of maximum camber and thickness aft on the chord. This would distribute the pressure and velocity more evenly along the chord and produce a lower peak velocity for the same lift coefficient. Fortunately, the airfoil shape to provide extensive laminar flow and low profile drag in low speed, subsonic flight will provide a pressure distribution which is favor­able for high speed flight. Figure 312 illustrates the pressure distributions and variation of critical Mach number with lift coefficient for a conventional low speed airfoil and a high speed section.

In order to obtain a high critical Mach number from an airfoil at some low lift coefficient the section must have:

(a) Low thickness ratio. The point of maximum thickness should be aft to smooth the pressure distribution.

(b) Low camber. The mean camber line should be shaped to help minimize the local velocity peaks.

In addition, the higher the required lift coefficient the lower the critical Mach number and more camber is required of the airfoil. If supersonic flight is a possibility the thick­ness ratio and leading edge radius must be small to decrease wave drag.

Figure 3.13 shows the flow patterns for two basic supersonic airfoil sections and pro­vides the approximate equations for lift, drag, and lift curve slope. Since the wave drag is the only factor of difference between the two airfoil sections, notice the configuration fac­tors which affect the wave drag. For the same thickness ratio, the circular arc airfoil would have a larger wedge angle formed between the upper and lower surfaces at the leading edge. At the same flight Mach num­ber the larger angle at the leading edge would form the stronger shock wave at the nose and cause a greater pressure change on the circular arc airfoil. This same principle applies when investigating the effect of airfoil thickness. Notice that the wave drag coefficients for both airfoils vary as the SQUARE of the thickness ratio, e. g., if the thickness ratio were doubled, the wave drag coefficient would be four times as great. If the thickness were increased, the airflow at the leading edge will experience a greater change in direction and a stronger shock wave will be formed. This powerful variation of wave drag with thick­ness ratio necessitates the use of very thin air­foils with sharp leading edges for supersonic flight. An additional consideration is that thin airfoil sections favor the use of low aspect ratios and high taper to obtain lightweight structures and preserve stiffness and rigidity.

The parameter — appears in the denominator of each of the equations for the aerodynamic coefficients and indicates a de­crease in each of these coefficients with an increase in Mach number. Essentially, this means that any aerodynamic surface becomes less sensitive to changes in angle of attack at higher Mach numbers. The decrease in lift curve slope with Mach number has tremendous implications in the stability and control of high speed aircraft. The vertical tail becomes less sensitive to angles of sideslip and the directional stability of the aircraft will deteri­orate with Mach number. The horizontal tail of the airplane experiences the same

general effect and contributes less damping to longitudinal pitching oscillations. These ef­fects can become so significant at high Mach numbers that the aircraft might require com­plete synthetic stabilization.

PLANFORM EFFECTS. The development of surfaces for high speed involves considera­tion of many items in addition to the airfoil sections. Taper, aspect ratio, and sweepback can produce major effects on the aerodynamic characteristics of a surface in high speed flight. Sweepback produces an unusual effect on the high speed characteristics of a surface and has basis in a very fundamental concept of aero­dynamics. A grossly simplified method of visualizing the effect of sweepback is shown in figure 3.14. The swept wing shown has the streamwise velocity broken down to a com­ponent of velocity perpendicular to the leading edge and a component parallel to the leading edge. The component of speed perpendicular to the leading edge is less than the free stream speed (by the cosine of the sweep angle) and it is this velocity component which determines the magnitude of the pressure distribution.

The component of speed parallel to the lead­ing edge could be visualized as moving across constant sections and, in doing so, does not contribute to the pressure distribution on the swept wing. Hence, sweep of a surface pro­duces a beneficial effect in high speed flight since higher flight speeds may be obtained be­fore components of speed perpendicular to the leading edge produce critical conditions on the wing. This is one of the most important ad­vantage of sweep since there is an increase in critical Mach number, force divergence Mach number, and the Mach number at which the drag rise will peak. In other words, sweep will delay the onset of compressibility effects.

Generally, the effect of wing sweep will apply to either sweep back or sweep forward. While the swept forward wing has been used I in rare instances, the aeroelastic instability of such a wing creates such a problem that sweep back is more practical for ordinary applica­tions.

In addition to the delay of the onset of com­pressibility effects, sweepback will reduce the magnitude of the changes in force coefficients due to compressibility. Since the component of velocity perpendicular to the leading edge is less than the free stream velocity, the magni­tude of all pressure forces on the wing will be reduced (approximately by the square of the cosine of the sweep angle). Since compressi­bility force divergence occurs due to changes in pressure distribution, the use of sweepback will "soften” the force divergence. This effect is illustrated by the graph of figure 3.14 which shows the typical variation of drag coefficient with Mach number for various sweepback angles. The straight wing shown begins drag rise at M=0.70, reaches a peak near M=1.0, and begins a continual drop past _M= 1.0. Note that the use of sweepback then delays the drag rise to some higher Mach number and reduces the magnitude of the drag rise.

In view of the preceding discussion, sweep – back will have the following principal ad­vantages :

(1) Sweepback will delay the onset of all compressibility effects. Critical Mach num­ber and force divergence Mach number will increase since the velocity component affect­ing the pressure distribution is less than the free stream velocity. Also, the peak of drag rise is delayed to some higher supersonic speed—approximately the speed which pro­duces sonic flow perpendicular to the leading edge. Various sweeps applied to wings of moderate aspect ratio will produce these approximate effects in transonic flight:

Sweep angle (A)	Percent Increase in critical Mach number	Percent increase in drag peak Mach number
0° ………………………………………….	0	0
15°…………………………………………	2	4
30° . ………………………………………	8	15
45° ……….. ……………………………..	20	41
60° ………… ……………………………	41	100

EFFECT OF SWEEPBACK ON LOW SPEED LIFT CURVE

SWEPT WING AT SWEPT WING IN A

ZERO SIDESLIP SIDESLIP TO THE RIGHT

SWEPT WING SWEPT WING IN A

IN LEVEL FLIGHT SIDESLIP TOWARD

THE DOWN WING

Figure 3.15. Aerodynamic Effects Due to Sweepback

(2) Sweepback will reduce the magnitude of change in the aerodynamic force coeffi­cients due to compressibility. Any change in drag, lift, or moment coefficients will be reduced by the use of sweepback. Various sweep angles applied to wings of moderate aspect ratio will produce these approximate effects in transonic flight.

Sweep angle (A)	Percent reduction in drag rise	Percent reduction in loss of
o°…………………………………………..	0	0
15°…………………………………………	5	3
30°…………………………………………	15	13
45°…………………………………………	35	зо
60° ,,. …………………………………….	60	30

These advantages of drag reduction and preser­vation of the transonic maximum lift coefficient are illustrated in figure 314.

Thus, the use of sweepback on a transonic aircraft will reduce and delay the drag rise and preserve the maneuverability of the aircraft in transonic flight. It should be noted that a small amount of sweepback produces very little benefit. If sweepback is to be used at all, at least 30° to 35° must be used to produce any significant benefit. Also note from figure 3-14 that the amount of sweepback required to delay drag rise in supersonic flight is very large, e. g., more than 60° necessary at M = 2.0. By comparison of the drag curves at high Mach numbers it will be appreciated that extremely high (and possibly impractical) sweepback is necessary to delay drag rise and that the lowest drag is obtained with zero sweepback. There­fore, the planform of a wing designed to operate continuously at high Mach numbers will tend to be very thin, low aspect ratio, and unswept. An immediate conclusion is that sweepback is a device of greatest application in the regime of transonic flight.

A few of the less significant advantages of sweepback are as follows:

(1) The wing lift curve slope is reduced for a given aspect ratio. This is illustrated by the lift curve comparison of figure 3-15 for the straight and swept wing. Any reduction of lift curve slope implies the wing is less sensitive to changes in angle of attack. This is a beneficial effect only when the effect of gusts and turbulence is con­sidered. Since the swept wing has the lower lift curve slope it will be less sensitive to gusts and experience less “bump” due to gust for a given aspect ratio and wing loading. This is a consideration particular to the aircraft whose structural design shows

‘ a predominating effect of the gust load spectrum, e. g., transport, cargo, and patrol types.

(2) “Divergence” of a surface is an aero – elastic problem which can occur at high dynamic pressures. Combined bending and twisting deflections interact with aerody­namic forces to produce sudden failure of the surface at high speeds. Sweep forward will aggravate this situation by "leading” the wing into the windstream and tends to lower the divergence speed. On the other hand, sweepback tends to stabilize the surface by “trailing” and tends to raise the divergence speed. By this tendency, sweep- back may be beneficial in preventing di­vergence within the anticipated speed range.

(3) Sweepback contributes slightly to the static directional—or weathercock—stability of an aircraft. This effect may be appre­ciated by inspection of figure 3-15 which shows the swept wing in a yaw or sideslip. The wing into the wind has less sweep and a slight increase in drag; the wing away from the wind has more sweep and less drag. The net effect of these force changes is to produce a yawing moment tending to return the nose into the relative wind. This directional stability contribution is usually small and of importance in tailless aircraft only.

(4) Sweepback contributes to lateral sta­bility in the same sense as dihedral. When the swept wing aircraft is placed in a side­slip, the wing into the wind experiences an increase in lift since the sweep is less and the wing away from the wind produces less lift since the sweep is greater. As shown in figure 3.15, the swept wing aircraft in a sideslip experiences lift changes and a sub­sequent rolling moment which tends to right the aircraft. This lateral stability contribution depends on the sweepback and the lift coefficient of the wing. A highly swept wing operating at high lift coefficient usually experiences such an excess of this lateral stability contribution that adequate controllability may be a significant problem. As shown, the swept wing has certain im­portant advantages. However, the use of sweepback produces certain inevitable disad­vantages which are important from the stand­point of both airplane design and flight oper­ations. The most important of these disad­vantages are as follows:

0) When sweepback is combined with taper there is an extremely powerful tendency for the wing to stall tip first. This pattern of stall is very undesirable since there would be little stall warning, a serious reduction in lateral control effectiveness, and the for­ward shift of the Center of pressure would contribute to a nose up moment (“pitch up” or “stick force lightening”). Taper has its own effect of producing higher local lift coefficients toward the tip and one of the effects of sweepback is very similar. All outboard wing sections are affected by the upwash of the preceding inboard sections and the lift distribution resulting from sweep – back alone is similar to that of high taper.

An additional effect is the tendency to develop a strong spanwise flow of the bound­ary layer toward the tip when the wing is at high lift coefficients. This spanwise flow produces a relatively low energy boundary layer near the tip which can be easily sep­

arated. The combined effect of taper and sweep present a considerable problem of tip stall and this is illustrated by the flow pat­terns of figure 3-16. Design for high speed performance may dictate high sweepback, while structural efficiency may demand a highly tapered planform. When such is the case, the wing may require extensive aero­dynamic tailoring to provide a suitable stall pattern and a lift distribution at cruise condi­tion which reduces drag due to lift. Wash­out of the tip, variation of section camber throughout span, flow fences, slats, leading edge extension, etc., are typical devices used to modify the stall pattern and minimize drag due to lift at cruise condition,

(2) As shown by the lift curve of figure 3.13 the use of sweepback will reduce the lift curve slope and the subsonic maximum lift coefficient. It is important to note this case is definitely subsonic since sweepback may be used to improve the transonic ma­neuvering capability. Various sweep angles applied to wings of moderate aspect ratio produce these approximate effects on the subsonic lift characteristics:

Parent reduction of subsonic maximum lift

coefficient and lift

The reduction of the low speed maximum lift coefficient (which is in addition to that lost due to tip stall) has very important implications in design. If wing loading is not reduced, stall speeds increase and sub­sonic maneuverability decreases. On the other hand, if wing loading is reduced, the increase in wing surface area may reduce the anticipated benefit of sweepback in the transonic flight regime. Since the require­ments of performance predominate, certain increases of stall speeds, takeoff speeds,

STRAIGHT WING OF SAME AREA, ASPECT RATIO, AND

STRUCTURAL TAPER

and landing speeds usually will be accepted. While the reduction of lift curve slope may be an advantage for gust considerations, the reduced sensitivity to changes in angle of attack has certain undesirable effects in subsonic flight. The reduced wing lift curve slope tends to increase maximum lift angles of attack and complicate the problem of landing gear design and cockpit visi­bility. Also, the lower lift curve slope would reduce the contribution to stability of a given tail surface area.

(3) The use of sweepback will reduce the effectiveness of trailing edge control surfaces and high lift devices. A typical example of this effect is the application of a single slotted flap over the inboard 60 percent span to both a straight wing and a wing with 35° sweepback. The flap applied to the straight wing produces an increase in maximum lift coefficient of approxi­mately 50 percent. The same type flap applied to the swept wing produces an increase in maximum lift coefficient of approximately 20 percent. To produce some reasonable maximum lift coefficient on a swept wing may require unsweeping the flap hinge line, application of leading edge high lift devices such as slots or slats, and possibly boundary layer control.

(4) As described previously, sweepback contributes to lateral stability by producing stable rolling moments with sideslip. The lateral stability contribution of sweepback varies with the amount of wing sweepback and wing lift coefficient—large sweepback and high lift coefficients producing large contribution to lateral stability. While sta­bility is desirable, any excess of stability will reduce controllability. For the majority of airplane configurations, high lateral sta­bility is neither necessary nor desirable, but adequate control in roll is absolutely neces­sary for good flying qualities. An excess of lateral stability from sweepback can aggra­vate “Dutch roll’’ problems and produce

marginal control during crosswind takeoff and landing where the aircraft must move in a controlled sideslip. Therefore, it is not unusual to find swept wing aircraft with negative dihedral and lateral control de­vices designed principally to meet cross wind takeoff and landing requirements,

(5) The structural complexity and aero – elastic problems created by sweepback are of great importance. First, there is the effect shown in figure 3-17 that swept wing has a greater structural span than a straight wing of the same area and aspect ratio. This effect increases wing structural weight since greater bending and shear material must be distributed in the wing to produce the same design strength. An additional problem is created near the wing root and “carry – through” structure due to the large twisting loads and the tendency of the bending stress distribution to concentrate toward the trail­ing edge. Also shown in figure 3-17 is the influence of wing deflection on the spanwise lift distribution. Wing bending produces tip rotation which tends to unload the tip and move the center of pressure forward. Thus, the same effect which tends to allay divergence can make an undesirable contri­bution to longitudinal stability.

EFFECT OF ASPECT RATIO AND TIP SHAPE. In addition to wing sweep, plan – form properties such as aspect ratio, and tip shape, can produce significant effects on the aerodynamic characteristics at high speeds. There is no particular effect of aspect ratio on critical Mach number at high or medium aspect ratios. The aspect ratio must be less than four or five to produce any apparent change in critical Mach number. This effect is shown for a typical 9 percent thick sym­metrical airfoil in the graph of figure 3.18. Note that very low aspect ratios are required to cause a significant increase in critical Mach number. Very low aspect ratios create the extremes of three dimensional flow and sub­sequent increase in free stream speed to create

local sonic flow. Actually, the extremely low aspect ratios required to produce high critical Mach number are not too practical. Generally, the advantage of low aspect ratio must be combined with sweepback and high speed airfoil sections.

The thin rectangular wing in supersonic flow illustrates several important facts. As shown in figure 318, Mach cones form at the tips of the rectangular wing and affect the pressure distribution on the area within the cone. The vortex develops within the tip cone due to the pressure differential and the resulting average pressure on the area within. thc-Cone is approximately one-half the pressure between the cones. Three-dimensional flow on the wing is then confined to the area within the tip cones, while the area between the cones experiences pure two-dimensional flow.

It is important to realize that the three­dimensional flow on the rectangular wing in supersonic flight differs greatly from that of subsonic flight. A wing of finite aspect ratio in subsonic flight experiences a three-dimen­sional flow which includes the tip vortices, downwash behind the wing, upwash ahead of the wing, and local induced velocities along the span. Recall that the local induced veloc­ities along the span of the wing would incline the section lift aft relative to the free stream and result in “induced drag.” Such a flow condition cannot be directly correlated with the wing in supersonic flow. ‘ The flow pattern for the rectangular wing of figure 3.18 dem­onstrates that the three-dimensional flow is confined to the tip, and pure two-dimensional flow exists on the wing area between the tip cones. If the wing tips were to be “raked” outside the tip cones, the entire wing flow would correspond to the two-dimensional (or section) conditions.

Therefore, for the wing in supersonic flow, no upwash exists ahead of the wing, three­dimensional effects are confined to the tip cones, and no local induced velocities occur along the span between the tip cones. The

supersonic drag due to lift is a function of the section and angle of attack while the subsonic induced drag is a function of lift coefficient and aspect ratio. This comparison makes it obvious that supersonic flight does not demand the use of high aspect ratio planforms typical of low speed aircraft. In fact, low aspect ratios and high taper are favorable from the standpoint of structural considerations if very thin sections are used to minimize wave drag.

If sweepback is applied to the supersonic wing, the pressure distribution will be affected by the location of the Mach cone with respect to the leading edge. Figure 3-19 illustrates the pressure distribution for the delta wing plan – form in supersonic flight with the leading edge behind or ahead of the Mach cone. When the leading edge is behind the Mach cone the com­ponents of velocity perpendicular to the leading edge are still subsonic even though the free stream flow is supersonic and the resulting pressure distribution will greatly resemble the subsonic pressure distribution for such a plan – form. Tailoring the leading edge shape and camber can minimize the components of the high leading edge suction pressure which are inclined in the drag direction and the drag due to lift can be reduced. If the leading edge is ahead of the Mach cone, the flow over this area will correspond to the two-dimensional supersonic flow and produce constant pressure for that portion of the surface between the leading edge and the Mach cone.

CONTROL SURFACES. The design of con­trol surfaces for transonic and supersonic flight involves many important considerations. This fact is illustrated by the typical transonic and supersonic flow patterns of figure 3.19. Trail­ing edge control surfaces can be affected ad­versely by the shock waves formed in flight above critical Mach number. If the airflow is separated by the shock wave the resulting buffet of the control surface can be very objec­tionable. In addition to the buffet of the sur­face, the change in the pressure distribution due to separation and the shock wave location can

create very large changes in control surface hinge moments. Such large changes in hinge moments create very undesirable control forces and present the need for an "irreversible" con­trol system. An irreversible control system, would employ powerful hydraulic or electric actuators to move the surfaces upon control by the pilot and the airloads developed on the surface could not feed back to the pilot, Of course, suitable control forces would be syn­thesized by bungees, “q” springs, bobweights, etc.

Transonic and supersonic flight can cause a noticeable reduction in the effectiveness of trailing edge control surfaces. The deflection of a trailing edge control surface at low sub­sonic speeds alters the pressure distribution on the fixed portion as well as the movable portion of the surface. This is true to the extent that a 1-degree deflection of a 40 percent chord eleva­tor produces a lift change very nearly the equivalent of a 1-degree change in stabilizer setting. However, if supersonic flow exists on the surface, a deflection of the trailing edge control surface cannot influence the pressure distribution in the supersonic area ahead of the movable control surface. This is especially true in high supersonic flight where supersonic flow exists over the entire chord and the change in pressure distribution is limited to the area of the control surface. The reduction in effective­ness of the trailing edge control surface at tran­sonic and supersonic speeds necessitates the use of an all movable surface. Application of the all movable control surface to the horizontal tail is most usual since the increase in longi­tudinal stability in supersonic flight requires a high degree of control effectiveness to achieve required controllability for supersonic maneu­vering.

SUPERSONIC ENGINE INLETS. Air which enters the compressor section of a jet engine or the combustion chamber of a ramjet usually must be slowed to subsonic velocity. This process must be accomplished with the least possible waste of energy. At flight speeds

just above the speed of sound only slight modi­fications to ordinary subsonic inlet design pro­duce satisfactory performance. However, at supersonic flight speeds, the inlet design must slow the air with the weakest possible series or combination of shock waves to minimize en­ergy losses and temperature rise. Figure 3.20 illustrates some of the various forms of super­sonic inlets or "diffusers."

One of the least complicated types of inlet is the simple normal shock type diffuser. This type of inlet employs a single normal shock wave at the inlet with a subsequent internal subsonic compression. At low supersonic Mach j numbers the strength of the normal shock wave is not too great and this type of inlet is quite practical. At higher supersonic Mach num­bers, the single normal shock wave is very strong and causes a great reduction in the total pressure recovered by the inlet. In addition, it is necessary to consider that the wasted J energy of the airstream will appear as an addi­tional undesirable rise in temperature of the captured inlet airflow.

If the supersonic airstream can be captured, the shock wave formations will be swallowed and a gradual contraction will reduce the speed to just above sonic. Subsequent diverging flow | section can then produce the normal shock wave which slows the airstream to subsonic. Further expansion continues to slow the air to lower subsonic speeds. This is the convergent- divergent type inlet shown in figure 3.20. If the initial contraction is too extreme for the inlet Mach number, the shock wave formation will not be swallowed and will move out in front of the inlet. The external location of the normal shock wave will produce subsonic flow immediately at the inlet. Since the airstream is suddenly slowed to subsonic through the strong normal shock a greater loss of airstream energy will occur.

Another form of diffuser employs an external oblique shock wave which slows the super­sonic airstream before the normal shock occurs. Ideally, the supersonic airstream could be

slowed, gradually through a series of very weak oblique shock waves to a speed just above sonic velocity. Then the subsequent normal shock to subsonic could be quite weak. Such a combination of the weakest possible waves would result in the least waste of energy and the highest pressure recovery. The ef­ficiency of various types of diffusers is shown in figure 3-20 and illustrates this principle.

An obvious complication of the supersonic inlet is that the optimum shape is variable with inlet flow direction and Mach number. In other words, to derive highest efficiency and stability of operation, the geometry of the inlet would be different at each Mach number and angle of attack of flight. A typical super­sonic military aircraft may experience large variations in angle of attack, sideslip angle, and flight Mach number during normal oper­ation. These large variations in inlet flow conditions create certain important design considerations.

(1) The inlet should provide the highest practical efficiency. The ratio of recovered total pressure to airstream total pressure is an appropriate measure of this efficiency.

(2) The inlet should match the demands of the powerplant for airflow. The airflow captured by the inlet should match that necessary for engine operation.

(3) Operation of the inlet at flight condi­tions other than the design condition should not cause a noticeable loss of efficiency or excess drag. The operation of the inlet should be stable and not allow “buzz” conditions (an oscillation of shock location possible during off-design operation).

In order to develop a good, stable inlet design, the performance at the design condition may be compromised. A large variation of inlet flow conditions may require special geometric features for the inlet surfaces or a completely variable geometry inlet design.

SUPERSONIC CONFIGURATIONS. When all the various components of the supersonic airplane are developed, the most likely general configuration properties will be..as follows:

(1) The wing will be of low aspect ratio, have noticeable taper, and have sweepback depending on the design speed range. The wing sections will be of low thickness ratio and require sharp leading edges.

(2) The fuselage and nacelles will be of high fineness ratio (long and slender). The supersonic pressure distribution may create significant lift and drag and require con­sideration of the stability contribution of these surfaces.

(3) The tail surfaces will be similar to the wing—low aspect ratio, tapered, swept and of thin section with sharp leading edge. The controls will be fully powered and ir­reversible with all movable surfaces the most likely configuration.

(4) In order to reduce interference drag in transonic and supersonic flight, the gross cross section of the aircraft may be “area ruled” to approach that of some optimum high speed shape.

One of the most important qualities of high speed configurations will be the low speed flight characteristics. The low aspect ratio swept wing planform has the characteristic of high induced drag at low flight speeds. Steep turns, excessively low airspeeds, and steep, power-off approaches can then produce extremely high rates of descent during landing. Sweepback and low aspect ratio can cause severe deterioration of handling qualities at speeds below those recommended for takeoff and landing. On the other hand, thin, swept wings at high wing loading will have rela­tively high landing speeds. Any excess of this basically high airspeed can create an im­possible requirement of brakes, tires, and arrest ing gear. These characteristics require that the pilot account for the variation of optimum speeds with weight changes and adhere to the procedures and techniques outlined in the flight handbook.