DE69309666T2 - Supersonic aircraft and process - Google Patents

Description

Background of the invention Field of the invention

The present invention relates to a supersonic aircraft and a method of operating the same, and more particularly to such an apparatus and method incorporating suction boundary layer control (and in some cases laminar flow control) in the wings of the aircraft.

Background technology

There are a number of challenges in the design of supersonic aircraft which meet performance requirements and are still acceptable to the community in terms of noise generation. To achieve improved supersonic cruise performance, a preferred supersonic aircraft configuration employs highly swept (subsonic) leading edge wings. However, this design creates special problems in the high lift conditions typical of climb and approach where high angles of attack are required. More particularly, these highly swept wings develop two leading edge vortices which, although increasing lift, also increase drag, resulting in a poor lift-to-drag ratio (L/D ratio). A higher (L/D) ratio is obtained when there is fully attached flow over the wings.

In the Concorde supersonic transport aircraft, the wing has a highly swept leading edge, but no leading edge devices are used. During takeoff and During climb, the configuration operates at a high angle of attack and the two strong vortices generated away from the leading edge generate sufficient lift for takeoff and climb. However, because of the high drag, the engines are operated at a relatively high power setting and consequently generate noise considerably above the maximum level permitted near most airports. As a result, there are very few airports at which Concorde can operate. Significant research efforts have been directed towards improving the L/D ratio of supersonic aircraft during takeoff and climb.

One prior art method to obtain good high-lift (L/D) performance is to use leading edge devices such as flaps or auxiliary wings to maintain near-attached flow. However, this method is mechanically complex and can still result in linkage separation. Furthermore, it requires additional hardware, and the systems to operate it, thus creating a penalty in both weight and cost. In addition, the space required to accommodate the leading edge devices reduces the available fuel volume in the leading edge area of the wings.

One concept that has been proposed for generating increased lift during takeoff and approach (desirably in combination with leading edge auxiliary wings or other leading edge devices) is to use vortex generators in the form of tip obstructions or guides located on the more forward inboard leading edge portions of both wings. In this case, these obstructions or guides are raised during takeoff and initial climb to generate two strong vortices sweeping over the upper inboard surfaces of both wings to provide increased lift during takeoff and initial climb. However, the drag generated requires a slightly higher power setting for the engines (and thus greater noise is generated). As the climb continues, these obstacles or guides are moved to a stowed position to reduce drag, allowing the engine to operate at a lower power setting and reduce noise.

The subject of laminar flow control has been studied for a number of decades, and a review of these studies is given in a recent publication entitled "Fifty Years of Laminar Flow Flight Testing" written by R.D. Wagner, D.V. Maddaion and D.W. Bartlett, publication number 881393 at the NASA Langley Research Center. Natural laminar flow (NLF) and also laminar flow control (LFC) using suction at the surface are discussed. Also discussed is HLFC (hybrid laminar flow control), which is said to be a "...flow control concept that integrates LFC and NLF and avoids the objectionable characteristics of each." Suction is applied at a forward location to maintain the LFC and natural laminar flow (NLF) exists immediately behind the LFC section. Proposals have been made to incorporate a suction system on a supersonic aircraft for cruise laminar flow control.

In a paper entitled "Application of Boundary Layer Control to HSCT Low Speed Configuration", AIAA/AHS/ASEE Aircraft Design, Systems and Operation Conference, September 17-19, 1990/ Dayton, Ohio (one of the authors of this paper is PH Parikh, the present inventor), the suitability or feasibility of using boundary layer control (BLC) on or in a high-lift configuration of a high-speed civil aircraft is (HSCT) for low speed performance enhancement. This is shown as installed on a supersonic aircraft having a double delta wing configuration wherein there is a highly swept (subsonic) forward inboard wing section and a less swept (supersonic) outboard aft leading edge wing section. Leading edge flaps (specifically camber nose flaps) are provided on the less swept outboard leading edge wing section. Laminar flow control suction regions are provided along the leading edge of the more swept inboard wing sections, and laminar flow control suction regions are provided on the outboard wing sections in regions aft of the flap hinge lines. These LFC regions are provided to reduce drag during supersonic cruise.

As illustrated in Fig. 2 of the same article, boundary layer control suction is applied at a location immediately abaft the hinge line of the camber nose flap, this being done to avoid separation of the air flowing up and back over the upper surface of the flap and then turning to flow over the upper wing surface. Accordingly, this configuration uses a leading edge flap/BLC combination to avoid (or at least reduce) the separated flow that would otherwise occur behind the flap in certain situations.

Summary of the invention

The supersonic aircraft of the present invention includes a wing having a highly swept subsonic leading edge portion arranged to develop a separated flow at high angles of attack which develops into a vortex, which moves over an upper surface of the wing.

This leading edge portion has on or in its outer surface a boundary layer control suction strip means extending along the leading edge portion. There is also provided a suction means for drawing in outside air through the suction strip means.

The suction means is arranged or set up in a manner and the suction strip means is also positioned, configured and arranged or set up in a manner such that the operation of the suction means to draw in the outside air reduces or mitigates air flow separated by the suction strip means so that the development of the vortex is reduced or mitigated.

In the preferred form, the outer surface of the leading edge portion is substantially fixed. Further, the preferred location of the suction strip means is adjacent a leading edge high point region of the airfoil.

In one embodiment, the wing further has a second less swept supersonic leading edge portion. Specifically, in this embodiment, the aircraft has a planar double delta configuration wherein the highly swept subsonic leading edge portion is at a more forward inboard location and the less swept supersonic leading edge portion is at a more rearward outboard location. Additionally, in this particular embodiment, the less swept supersonic leading edge portion has mechanically operable high lift device means to at least partially mitigate separated flow at high angles of attack.

In another embodiment, the swept leading edge portion extends substantially along an entire Leading edge of the wing in a plan arrow configuration.

In yet another embodiment, a laminar flow control suction strip means is additionally provided, located on or in a surface region of the airfoil. The suction means is arranged to draw in outside air also through the laminar flow control suction strip means. In a particular configuration, the laminar flow control suction strip means is located adjacent to and rearward of the boundary layer control suction strip means.

Furthermore, in this further embodiment, the suction means is arranged to operate in a boundary layer control mode to draw in sufficient outside air through the boundary layer control suction strip means to mitigate the separated flow, and also to operate in a laminar flow control mode and draw in outside air through both the boundary layer control suction strip means and the laminar flow control suction strip means at a flow rate to effect laminar flow control through both the boundary layer control suction strip means and the laminar flow control suction strip means.

Furthermore, in a preferred embodiment, a perspiration anti-icing strip means is positioned adjacent to the boundary layer control suction strip means in such a manner that anti-icing air blown outwardly through the perspiration anti-icing strip means blows over at least the boundary layer control suction strip means for anti-icing thereof. Further, the suction means comprises a means for drawing in outside air through the perspiration anti-icing strip means, and there is also an anti-icing means for discharging of anti-icing air to the transpiration anti-icing strip means. Additionally, in a particular configuration, a second laminar flow control suction strip means is located rearwardly of the transpiration anti-icing strip means. The suction means further includes means for drawing outside air through the second laminar flow control suction strip means.

In the preferred embodiment, the boundary layer control strip means has a width dimension in at least a region of the highly swept subsonic leading edge portion taken along a line parallel to the freestream flow with respect to the aircraft, and this width dimension is no greater than about 5% of a chord length of said airfoil. More particularly, the width dimension is between about 1% to 5% of the chord length, and even more particularly, in a preferred form, it is between 1% to 2% of the chord length.

In the method of the present invention, the arrangement of the aircraft is as described above. During modes of operation where the aircraft is at a high angle of attack and where it is desired to obtain an improved lift to drag ratio so as to enable the engines to be operated at a relatively low power setting (to reduce the generation of noise), the aspirator is operated to mitigate the separated flow along the swept leading edge and to improve the lift to drag ratio.

Furthermore, in the method of the present invention, in a mode of operation during takeoff and initial climb where the generation of noise is less critical, the aircraft is operated at a high angle of attack but without operating the suction means (or at most operating the suction means at a low level) so that a vortex is generated across the swept leading edge. This vortex generates lift, but it also results in increased drag, which requires the higher engine setting. As the aircraft continues through the climb and reaches an altitude where noise generation is a greater problem, the aspirator is operated to reduce or mitigate the formation of the vortex so that the engines are able to operate at a lower power setting, thus generating relatively less noise.

In the embodiment in which laminar flow control suction is applied, during takeoff and at least initial climb, the laminar flow control suction means is not used. However, the laminar flow control suction means is operated during modes of operation where laminar flow control is advantageous (for example, especially during supersonic cruise and possibly in other modes of operation) to reduce drag and improve performance.

The operation of the transpiration anti-icing strip means is such that hot de-icing air is blown outwardly through the anti-icing strip means during an anti-icing mode of operation. In addition, during a mode of operation in which laminar flow control is desired, the suction means may be used to draw air through the anti-icing strip means to effect laminar flow control.

Other features will become apparent from the following detailed description.

Short description of the drawings

Fig. 1 is a top view of a state-of-the-art supersonic aircraft having a reasonably optimized present design;

Fig. 2 is a top plan view of a supersonic aircraft similar to that of Fig. 1 but incorporating the teachings of the present invention;

Fig. 3 is a sectional view taken perpendicular to the front edge in line 3-3 of Fig. 2;

Fig. 4 is an isometric view of the aircraft of the first embodiment of the present invention shown in Fig. 2;

Fig. 5 is a bottom plan view of the aircraft of the first embodiment shown in Figs. 2 to 4;

Fig. 6 is a top plan view of a second embodiment of the present invention;

Fig. 7 is a sectional view of the leading edge taken along line 7-7 of Fig. 6;

Fig. 8 is a semi-schematic view of the air flow system used in the second embodiment;

Fig. 8A is a schematic view of one of the valves of the system of Fig. 8, showing the valve in a different operating position;

Fig. 8B is a view of another of the valves of the system of Fig. 8 shown in a different operating position;

Fig. 9 is a top plan view of a prior art supersonic aircraft having a swept wing configuration; and

Figure 10 is a top plan view of a third embodiment of the present invention embodied in an arrow wing.

Brief description of the preferred embodiment

In Fig. 1, to illustrate various elements of the prior art, there is shown in a top plan view a proposed supersonic transport aircraft embodying various concepts proposed in the prior art for the leading edges of the wings. This aircraft 10 includes a fuselage 12 and right and left wings 14. Trailing edge flaps 16 are provided at the rear end of the wings 14 and four engines 18 are mounted on the underside of the wing near the trailing edge on opposite sides of the fuselage. There is a tail assembly 20 which includes the horizontal tail surface 22 and the vertical control surface 24.

The wings 14 are of the double delta configuration, wherein there is a forward swept wing section 26 having a highly swept (subsonic) leading edge at an inboard location, and also an outboard wing section 30 having a more moderately swept (supersonic) leading edge 32. In this particular aircraft 10, leading edge devices 34 are provided along the rear, less swept leading edge sections 32, and these leading edge devices may be, for example, cambered leading edge flaps. The more highly swept inboard leading edge 28 has across the rear portion thereof leading edge slats 36. Forward of the leading edge slats 36 are provided tip guides 38 (these have been previously described in the "Background of the Invention" section).

In an optimized design, the swept inboard leading edges 28 have a rounded cross-sectional configuration with a minimum radius at the leading edge of, for example, between 1 and 2 inches in a cutting line parallel to the free streamline. Such a swept leading edge is referred to as a "subsonic" leading edge because the flow component taken perpendicular to the leading edge is subsonic during cruise. On the other hand, the leading edges 32 of the two outboard wing sections 30 are supersonic leading edges (that is, the flow component taken perpendicular to the leading edge 32 is supersonic during cruise) and the radius of curvature at the leading edge 32 is made as small as possible (i.e., a very small fraction of an inch). During cruise flight, the obstacles or guides 38 and the leading edge devices 36 and 34 are all fully retracted and in their aerodynamically clean configuration.

A typical mode of operation during takeoff and climb of this aircraft 10 would be as follows. During takeoff, the leading edge auxiliary wings or flaps 34 and 36 are fully retracted for increased lift and the tip baffles or guides 38 are swept upward to create two strong vortices that sweep upward over the upper surface of the inboard portions of the wings 14. As previously discussed herein, in the swept wing configuration, these vortices are of sufficient strength to produce adequate lift for takeoff and climb. However (and also as previously discussed herein), these vortices produce a relatively high amount of drag. Accordingly, the engines 18 must be set to a relatively high power setting to carry the aircraft through takeoff and initial climb. Shortly after takeoff, the tip guides 38 would be moved to their stowed position and the leading edge auxiliary wings 36 and flaps 34 would be deployed to increase the L/D (lift/drag) ratio of the climb configuration. The transition from the takeoff configuration to the climb configuration would be completed by the time the aircraft 10 reaches an altitude of approximately 700 feet, at which point noise pollution becomes an acute problem for the community.

Having presented the foregoing as background information, attention is now directed to a description of the first embodiment of the present invention with reference to Figs. 2 to 5.

Referring to Fig. 2 (which, like Fig. 1, is a top plan view), it can be seen that the first embodiment 110 is in certain respects similar to the aircraft of Fig. 1 in that the aircraft 110 of this first embodiment includes a fuselage 112, double delta wings 114, trailing edge flaps 116, engines 118, and a tail assembly 120 (this tail assembly 120 includes the horizontal tail section 122 and a vertical control surface 124). In addition, as in the prior art aircraft of Figure 1, there is the highly swept (subsonic) inboard wing section 126 having leading edges 128 of a larger radius of curvature (e.g., 1 to 2 inches as measured along a section line parallel to the free streamline), and there are also the less swept (supersonic) outboard wing sections 130 having less swept leading edges 132 of much smaller radius of curvature. In a preferred form of this first embodiment, in which the aircraft is designed for Mach 2.4 cruise, the sweep angle of the highly swept inboard leading edge 128 is about 73°. However, it is apparent that the sweep angle of this subsonic leading edge may vary according to the cruise Mach number and possibly other design factors.

In this first embodiment of the present invention, as shown in Fig. 2, the outboard wing sections 130 are (or may be) substantially the same in overall configuration as the outboard wing members 30 of the aircraft 10 of Fig. 1. Accordingly, as shown in Fig. 2, auxiliary leading edge flaps 134 are present, which are (or may be) camber nose flaps or some other leading edge device, and this less swept leading edge 32 is quite sharp. However, the aircraft 110 of the first embodiment of the present invention differs significantly from the aircraft 10 of Fig. 1 with respect to the configuration and operation of the inboard wing sections 126.

In the first embodiment of Figure 2, the leading edge portions 128 of the inboard wing sections 126 are devoid of mechanical high lift devices and accordingly comprise a stationary outer skin structure. Boundary layer control suction is provided by a perforated suction surface along a narrow strip 136 extending from the forward leading edge location 138 rearward to a junction line 140 where the inboard leading edge portion 128 meets the outboard leading edge portion 132. It can be seen from Figure 3 that the forward longitudinal edge 142 of the boundary layer suction strip 136 is at (or closely adjacent to) the planar "high point" of the aircraft 130. (The term "peak" refers to the circumferential edge line of the aircraft 110 as seen in a top view of the aircraft.) The rear longitudinal edge 144 of the suction strip 136 is positioned a short distance behind the forward line 142 and is located in a forward portion of the upper wing surface 146.

The means for applying suction to the suction strip 136 is shown somewhat schematically in Fig. 3, where a pumping device 147 for generating the suction is shown, and a plenum chamber 148 which is arranged on or on the inside surface of the suction strip 136 into which the outside air is drawn. In the description of the second embodiment of Figs. 6-8, the suction system for the second embodiment will be discussed somewhat more fully. It will be appreciated that the appropriate components from this second embodiment, such as a suction pump, valves and ducts (as well as other components known in the art) can be used in this first embodiment to create the suction through this strip 136 and to control the operation thereof.

To describe the operation of the first embodiment shown in Figures 2 through 5, consider first the operation of the aircraft 110 during takeoff and initial climb. As the aircraft 110 gains speed on the runway and turns to a higher angle of attack to effect takeoff and then proceeds with the initial climb, the aircraft 110 is at a relatively high angle of attack (e.g., typically 10 to 14 degrees). At this high angle of attack, as shown in Figure 5, which is a bottom plan view, the airflow along the lower wing surface 149 adjacent each leading edge 128 follows an attachment line 150 that is inboard of the leading edge 128. The flow 152 outboard of this abutment line 150 flows rearward and outward around the leading edge 128 and then upward. The flow 154 inboard of the abutment line 150 continues to flow rearward along the lower surface 148.

The flow 152 flowing upward and around the leading edge 128 has a tendency to separate and develop into a strong vortex which sweeps over the upper surface of the wing 114. However, if sufficient suction is applied by the strip 136, this flow separation is prevented and this vortex is suppressed so that the drag is significantly reduced.

It is helpful at this point to make at least a distinction between suction laminar flow control and suction boundary layer control, which is designed to prevent or inhibit separated flow and thus promote adjacent flow. The suction coefficient for laminar flow control is typically quite small, on the order of 0.0005. On the other hand, the suction coefficient for boundary layer control to prevent separated flow has a suction coefficient that could be, for example, 20 to 40 times as large as the suction coefficient for laminar flow control. This will be discussed more fully later here in connection with the description of the second embodiment where both suction boundary layer control (which might more properly be called "suction separation flow control") and laminar flow suction control are applied in conjunction with each other.

The suction by the two vortex strips 136 is accomplished as follows. During takeoff and the beginning of the climb, no suction is applied by the strips 136 (or at most a small amount of suction) so that two strong vortices are able to develop on the inboard portions of the wings 114 and pass rearward over the upper wing surfaces. As previously indicated, these two vortices in and on a highly swept wing generate substantial lift, the disadvantage being that greater power is required from the engines 118 because of the large amount of drag developed. However, when the aircraft 110 is taking off or at very low altitude, the increased noise developed by the increased thrust would be at acceptable levels for the immediate airport vicinity. As the aircraft continues to climb, however, it is to be expected that it will fly over areas where the higher engine power settings would be environmentally objectionable. At this time, suction boundary layer control would be activated to draw in enough air flowing over the leading edges 128 to substantially eliminate (or mitigate) the effect of the vortices to substantially reduce drag. At the same time, the power setting of the engines 118 would be reduced so that less noise would be generated. Nevertheless, the aircraft would be able to maintain an angle of attack sufficient to develop enough lift to continue the climb.

It will be understood, of course, that activation of the suction through the boundary layer strips 136 could be programmed to optimize performance. Accordingly, during takeoff, there may be no application of suction through the strips 136, and this could be programmed to be increased as the aircraft approaches the altitude where noise becomes an acute problem to the community. In addition, as the aircraft 110 continues to climb and the angle of attack decreases somewhat, the suction through the boundary layer control strips 136 could be further reduced or even turned off completely. In general, during subsonic climb, the boundary layer control mode of operation would continue until the aircraft 110 is away from geographic areas where excessive noise from the aircraft would be objectionable.

An analysis of the performance of an aircraft 110 embodying the present invention was made with respect to the aircraft design illustrated in Figure 1. This analysis shows that when lift divided by drag (L/D) is plotted against lift coefficient during initial climb conditions in which a relatively high power setting is present, the aircraft 110 of the first embodiment of the present Invention shows an improved L/D over that illustrated in Fig. 1.

The analysis also shows that for an angle of attack typical during the outgoing climb, there is an increase in pitch moment that needs to be compensated for under conditions where the boundary layer control mechanism is disabled as compared to when the boundary layer control mechanism is operating. (Disabling the boundary layer control would occur under circumstances where there is a potential malfunction of the BLC system.) However, the horizontal stabilizer, sized for takeoff turn and stall recovery requirements, has extensive control capacity to counteract the pitch moment increase resulting from an inadvertent loss of boundary layer control suction during the outgoing climb.

In the event of an inadvertent loss of boundary layer control suction at the highly swept inboard wing leading edge, the analysis shows that the net lift on the configuration would increase slightly due to the transition from the attached flow mode to the vortex lift mode. This is in sharp contrast to the behavior of low-swept wings (20º to 35º leading edge sweep typical of subsonic transport aircraft) where the loss of leading edge boundary layer control (if equipped) would result in a dramatic loss of lift and deep stall. For this reason, implementing BLC on low-swept wings raises serious safety concerns that are not relevant in the present highly swept wing application.

Overall, the improvements provided by the first embodiment of the present invention are that there is improved lift-through-drag during the initial climb, as well as elimination of systems required to move leading edge devices, weight and cost savings by eliminating leading edge hardware and systems, and also a gain in fuel volume in the leading edge area. On the other hand, the disadvantages of inboard wing leading edge device removal are that there would be a requirement to increase the horizontal tail area by a small amount for takeoff rotation and stall recovery. In addition, there is a slightly reduced lift coefficient during approach at maximum angle of attack for touchdown, this being a consequence of the tip flap removal.

A second embodiment of the present invention is illustrated in Figures 6 through 8. This second embodiment 210 is similar to the first embodiment, except that suction laminar flow control is employed in addition to suction boundary layer control. Further added is the use of hot air blown outward to flow over the leading edge surface for anti-icing.

In Fig. 6, the aircraft 210 is shown having a fuselage 212, wings 214, trailing edge flaps 216, engines 218, and a tail assembly 220 (which includes the horizontal tail section 222 and the vertical control surface 224). In addition, as in the first embodiment, there is the inboard highly swept inboard wing portion 226 with a leading edge 228 and the less swept outboard wing portion 230 with the leading edge 232. Further, as in the first embodiment, there are the leading edge devices 234 on or in the outboard leading edges 232.

The additional features of this second embodiment will now be described with initial reference to Figure 7. It will be seen that, as in the first embodiment, there is a boundary layer control strip 236 which extends from the forwardmost portion 238 of the leading edge 228 to the junction 240 where the leading edge 228 meets the outboard leading edge 232 (see Figure 6) and which has leading and trailing edges 242 and 244 (see Figure 7). In addition, immediately rearward of the boundary layer control strip 236, on each forward upper wing surface 258, there is a laminar flow control strip 260. Positioned immediately adjacent to the boundary layer control strip 236 and extending downwardly and rearwardly from the leading edge 242 of the boundary layer control strip 236 is a combination transpiration anti-icing and laminar flow control strip 262. Then, adjacent to (and immediately rearwardly of) the strip 262 is a lower laminar flow control strip 264. These strips 236, 260, 262 and 264 extend along the length of the forward leading edges 228 of the wings 214.

Briefly discussing the operation of the various strips described immediately above with reference to FIG. 7, the boundary layer control strip 236 performs substantially the same function as it does in the first embodiment. More specifically, in high lift operation, during that portion of the outgoing climb where it is desired to have a high L/D, suction is applied through the boundary layer control strips 236 to substantially eliminate the vortex generated from each swept leading edge (or substantially mitigate the effect of the vortex). However, in addition, if it is desired to utilize laminar flow control (for example, during subsonic cruise), the boundary layer control strips 236 may be used in conjunction with strips 260, 262 and 264 to accomplish laminar flow control.

Yet another function of the leading edge system shown in Figure 7 is that each strip 262 can also be used to vent hot air for transpiration anti-icing. (This concept of using a perforated airfoil surface for both laminar flow or boundary layer control of air by suction and anti-icing by venting hot air is more fully described in U.S. Patent 5,114,100, issued May 19, 1992, entitled "ANTI-ICING SYSTEM FOR AIRCRAFT" and naming as its inventors K.C. Rudolph and Dezso Georgefalvy, which is incorporated by reference herein.)

Fig. 8 shows schematically the control system for the device of Fig. 7 of this second embodiment. There are shown two concentrically formed channels 266 and 268. The channel 266 has an outer cylindrical channel portion 270 and an inner concentric cylindrical channel portion 272. The inner channel portion 272 is connected by a tube means 274 to a chamber 276 positioned immediately adjacent the inner surface of the boundary layer control strip 236. The annular region 278 defined by the outer and inner tubes 270 and 272 leads by a tube means 280 to a chamber 282 immediately adjacent the inner surface of the aforementioned upper laminar flow control strip 260.

The concentric channel 268 includes an outer channel 284 and an inner channel 286. The inner channel 286 is connected by a tube means 288 to a chamber 290 located immediately adjacent the inner surface of the anti-icing and laminar flow control strip 262. The chamber defined by the inner and outer tubes 286 and 284 annular chamber 292 is connected by a tubing means 294 to a plenum chamber 296 immediately adjacent the inner surface of the lower laminar flow control strip 264.

The annular passage 278 leads through a first flow control valve 298 to a second control valve 300. The inner tube 274 has a bypass portion 302 which bypasses the valve 298 and leads back into the outer tube 270 at a location between the valves 298 and 300. The suction is provided by a first suction pump 304 which is driven by a turbine 306 which in turn is driven by the bleed air from one or more of the engines 218.

As stated above, the strip 262 can function in either the laminar flow control mode, wherein it draws in air, or in the transpiration anti-icing mode, wherein it blows out hot air. Accordingly, the inner conduit 286 of the duct 268 passes through a valve 308 which operates in one of two positions. In a first position, shown in Fig. 8, the valve 308 passes hot bleed air from one or more of the engines 218 through the tube 310, through the valve 308 and into the inner conduit 286 to pass into the chamber 290 to cause hot air to pass out through the strip 262. Then this hot air flows from the strip 262 so as to sweep upward and over the surfaces of the strips 236 and 260.

In the laminar flow control mode, the valve 308 is in the position shown in Fig. 8A, wherein the valve 308 is positioned to block the flow of hot bleed air from the tube 310 and the inner conduit 286 leads directly into the inner region 312 of the outer tube 284. The tube 284 in turn leads to a control valve 314, which in turn leads to a second suction pump 316 driven by a second turbine 318, which che is propelled by air from one or more of the engines 218.

Describe the operation of this second embodiment 210, wherein during takeoff and initial climb, the boundary layer control flow through the boundary layer control strips 236 is operated in substantially the same manner as set forth in the first embodiment. In this mode, the valve 298 would be in its closed position shown in Figure 8 and the control valve 300 would be selectively operated to control the boundary layer control flow as set forth in the description of the first embodiment. Additionally, during takeoff and initial climb, the valve 308 is in the position of Figure 8A and the valve 314 is closed so that there is no flow through any of the surface strips 262 or 264.

Now assume that the aircraft 210 is going through the initial climb from takeoff and icing conditions occur. At this time, the valve 308 is moved to the position of Figure 8 and anti-icing bleed air from the engine or engines 218 flows through the tube 310, through the inner conduit 286 and out of the plenum 290 through the strips 236 as anti-icing air. As stated above, this anti-icing air flows upward over the leading edge of the wing and a short distance over the upper surface to prevent the formation of ice on the leading edge portion of the wing.

Now assume that the aircraft has completed passage through the initial part of its climb and has now reached a supersonic speed where laminar flow control can be beneficial. (It would be expected that this would be the case during the latter part of the climb through the supersonic region and also (occurs during supersonic cruise.) At this time, valve 298 is moved from its position in Fig. 8 to its position in Fig. 8A, thus opening annular passage 278 to control valve 300. At the same time, control valve 300 is placed in the appropriate position to create the proper amount of suction for chambers 276 and 282 to draw laminar flow control air in through surface strips 236 and 260. At the same time, valve 308 is moved from the position of Fig. 8 to the position of Fig. 8A, thus shutting off the flow of hot bleed air from engine 218 and also connecting inner conduit 286 to valve 314. Then, the valve 314 is adjusted to the appropriate position to produce the correct amount of suction to obtain the desired laminar flow control at the surface strips 262 and 264.

It will of course be understood that the suction control system illustrated in Figure 8 is somewhat schematic and of course various changes may be made. For example, valves may be added or rearranged to obtain other control characteristics. In one possible arrangement, an additional valve may be placed in the bypass line 302 so that this flow may be controlled in addition to the control provided by the valve 300, or possibly the valve 300 may be omitted. Similar modifications may be made with respect to the valve equipment at 308 and 314.

The analysis has shown that the flow requirements for boundary layer control operation (more precisely separation flow control operation) compared with laminar flow control operation show that the volume flows for the two applications are comparable. Accordingly, the sizes of the suction channels and of the compressors in the systems shown in Fig. 8, when designed for the laminar flow control mode, will be adequate for the boundary layer control application when only the one boundary layer control strip 236 is used. This analysis is as follows:

Typically, the suction coefficient (Cq) for laminar surface flow is approximately equal to 0.0005. On the other hand, the suction coefficient (Cq) for boundary layer control (more precisely flow separation control) is on the order of 0.01 to 0.02. Accordingly, the suction coefficient is generally 20 to 40 times as large for BLC (i.e. flow separation control) as it is for laminar flow control. (The non-dimensional suction coefficient (Cq) is defined as PoVo/G∞.)

On the other hand, it should be considered that laminar flow control typically operates during supersonic cruise, where the Mach number may be, for example, 2.4 and the aircraft altitude 60,000 feet. Furthermore, during supersonic cruise, a relatively high surface temperature is present immediately adjacent to the aircraft surfaces. On the other hand, during the initial subsonic climb, the speed would be, for example, Mach 0.3 and the altitude might be 1,000 feet (depending on the runway altitude). Under these conditions, the free flow mass flow during the climb, if boundary layer control were used, would be approximately equal to 1.5 of the free flow mass flow during supersonic cruise, where laminar flow control is typically used. Therefore, the surface mass flux ( oVo) when boundary layer control is applied during start-up would be 30 to 60 times the surface mass flux during laminar flow control.

Typically, when comparing the pressure, temperature and air density at the suction surface for supersonic cruise (where laminar flow control would occur) to the conditions during the initial climb (where boundary layer control would occur), the air density at the suction surface during the initial climb when boundary layer control was applied would be approximately 20 times the air density at the suction surface during supersonic cruise where laminar flow control was applied. From this, it can be concluded that the suction air velocity at the surface under laminar flow control conditions would be approximately 1.5 to 3 times the velocity during the boundary layer control mode during takeoff. Since the skin pressure drop is proportional to the suction air velocity, the skin pressure drop in the BLC mode would also be 1.5 to 3 times the skin pressure drop in the LFC mode, which is quite acceptable. Accordingly, it can be reasonably concluded from this analysis that the suction system that would be used for laminar flow control would be adequate for the boundary layer control mode. Other facts to consider are that the temperature of the suction air is slightly higher during the laminar flow control mode, thus resulting in a slightly higher viscosity and a slightly higher pressure drop for a given volume flow. A higher pressure drop is also tolerable in the boundary layer control mode because the external pressure is significantly higher compared to the cruise laminar flow control conditions.

The elimination of the variable geometry leading edge devices on or in the highly swept inboard wing also eliminates potential steps and gaps on the surface that are not essential for maintaining laminar flow. would be acceptable, although the incorporation of the fixed geometry leading edge suction surface on or in the inboard wing makes it much easier to achieve the surface smoothness requirements for maintaining laminar flow.

Before describing a third embodiment of the present invention shown in Fig. 10, reference is first made to Fig. 9 which is a plan view illustrating another configuration of a prior art supersonic aircraft. This aircraft 410 has the swept wing configuration in which the entire leading edge of the wing is highly swept. This aircraft 410 includes a fuselage 412, wings 414, trailing edge flaps 416, a plurality of engines 418, and a tail assembly including a horizontal tail section 422 and a vertical control surface 424.

However, the entire leading edge 428 is a highly swept leading edge extending the entire length of the entire aft portion of the wing, with a relatively short wing tip portion 430 that is parallel to the streamline. The leading edges 428 of the wings 414 are provided with a plurality of leading edge flaps or auxiliary vanes 432.

Figure 10 illustrates a third embodiment 510 of the present invention in which the swept wing configuration is used. This third embodiment 510 of the present invention is substantially similar to the prior art aircraft of Figure 9. Accordingly, there is the fuselage 512, wings 514, trailing edge flaps 516, engines 518 and a tail assembly 520-524. There is a swept leading edge 528 that extends the entire length to the wing tip portion 530. However, instead of having the leading edge flaps 432, two boundary layer control suction strips 536 extending along the length of the swept leading edge in a manner similar to the previous two embodiments. It is believed that the configuration and mode of operation of this third embodiment will be readily understood from reading the description of the first two embodiments. More specifically, the suction boundary layer control is operated during climb to provide a relatively high lift-to-drag ratio (L/D) while still providing sufficient lift for the outgoing climb and mitigation of engine noise.

It will of course be appreciated that this third embodiment 510 could be further modified, for example, to incorporate features of the second embodiment shown in Figures 6 to 8. Accordingly, as in the second embodiment, in addition to using the suction boundary layer control strips 536 (more precisely, the flow separation control strips 536), suction laminar flow regions could also be added.

With respect to various design features of the present invention, the formation of the boundary layer control suction regions 236, 336 and 536 is desirably accomplished by utilizing available technology relating to boundary layer control. Current designs indicate that the size of the suction holes could be, for example, between 0.05 and 0.08 mm (0.002 inches to 0.003 inches), and the spacing of these holes could be between 0.25 and 0.64 mm (0.01 inches to 0.025 inches). Current analysis indicates that the width of the boundary layer control strips (taken along a streamline parallel to the longitudinal axis of the aircraft) would be between about 1% to 2% of the wing chord length. Accordingly, as the chord length of the aircraft decreases in an outboard direction, it would be expected that the Width of the boundary layer strips (136, 236 or 536) would decrease accordingly. However, within the broader scope of the present invention, under certain circumstances, this width dimension may vary by perhaps somewhat greater than 2%, perhaps as high as 4 to 5%. In addition, further analysis may show that this 2% dimension of the suction strips 136, 236 or 536 should be higher where the chord length is much shorter.

With respect to the dimensioning of the laminar flow control strip 260 of the second embodiment, the present analysis shows that the overall width dimension (i.e., the dimension taken parallel to the longitudinal axis of the aircraft) would be approximately 10% of the chord length. Accordingly, if the laminar flow control strip 236 is 2% of the chord length, the width of the laminar flow control strip 260 would be approximately 8% (depending on specific design considerations) to make up a total of 10% during LFC. Normally, if laminar flow control is applied over, for example, 10% of the chord length of the wing, the laminar flow will continue over the wing surface up to perhaps an additional 20% of the chord length. Accordingly, it may be desirable to add an additional laminar flow control strip closer to the midspan line in each wing to obtain even more laminar flow over the wing surface.

Regarding the suction pressure differential from the outside to the inside surface where suction is applied, the present analysis shows that this would normally be between 950 and 2400 Pa (20 to 50 pounds per square foot) pressure differential.

Claims (19)

1. Supersonic aircraft comprising: a. an airfoil having an upper surface and a highly swept subsonic leading edge portion arranged to develop, at high angles of attack, separated flow which develops into a vortex moving over the upper surface of the airfoil; b. the leading edge portion having an outer surface and having on or in the outer surface a boundary layer control suction strip means extending along the leading edge portion; c. Suction means for drawing in outside air through the suction strip means; and d. wherein the suction means is arranged in a manner and the suction strip means is positioned, configured and arranged in a manner such that operation of the suction means to draw in the outside air mitigates the air flow separated by the suction strip means so as to mitigate the development of the vortex. 2. An aircraft according to claim 1, wherein the outer surface of the leading edge portion is substantially fixed. 3. An aircraft according to claim 2, wherein the suction strip means is adjacent a leading edge peak region of the wing. 4. The aircraft of claim 1, wherein the suction strip means is adjacent a leading edge peak region of the wing. 5. The aircraft of claim 1, wherein the wing further has a second, less swept supersonic leading edge portion. 6. The aircraft of claim 5, wherein the aircraft has a planar double delta configuration wherein the highly swept subsonic leading edge portion is at a more forward inboard location and the less swept supersonic leading edge portion is at a rearward more outboard location. 7. An aircraft according to claim 6, wherein the less swept supersonic leading edge portion has mechanically actuatable leading edge high lift device means for at least partially mitigating separated flow at high angles of attack. 8. An aircraft according to claim 1, wherein the highly swept leading edge portion extends substantially along an entire leading edge of the wing. 9. An aircraft according to claim 1, wherein a laminar flow control suction strip means is located on a surface region of the wing and the suction means for drawing in outside air is also arranged through the laminar flow control suction strip means. 10. The aircraft of claim 9, wherein the laminar flow control suction strip means is located adjacent to and backward of the boundary layer control suction strip means. 11. The aircraft of claim 10, wherein the suction means is arranged to operate in a boundary layer control mode to draw in sufficient outside air through the boundary layer control suction strip means to mitigate the separated flow, and further to operate in a laminar flow control mode and draw in outside air through both the boundary layer control suction strip means and the laminar flow control suction strip means at a flow rate to effect laminar flow control through both the boundary layer control suction strip means and the laminar flow control suction strip means. 12. The aircraft of claim 1, further comprising a transpiration anti-icing strip means positioned adjacent to the boundary layer control suction strip means in a manner such that anti-icing air blown outwardly by the transpiration anti-icing strip means blows over at least the boundary layer control suction strip means for anti-icing thereof. 13. An aircraft according to claim 12, wherein the suction means comprises means for drawing in outside air through the transpiration anti-icing strip means, and further comprising anti-icing means for delivering anti-icing air to the transpiration anti-icing strip means. 14. The aircraft of claim 13, wherein a second laminar flow control suction strip means is located rearward of the transpiration anti-icing strip means, and the suction means further comprises means for drawing outside air through the second laminar flow control suction strip means. 15. The aircraft of claim 1, wherein the boundary layer control suction strip means has a width dimension at at least a portion of the highly swept subsonic leading edge portion taken along a line parallel freestream flow relative to the aircraft, and the width dimension is no greater than about 5% of a chord length of the wing. 16. The aircraft of claim 15, wherein the width dimension is between about 1% to 5% of the chord length. 17. The aircraft of claim 16, wherein the width dimension is between about 1% and 2% of the chord length. 18. A supersonic aircraft according to claim 1, further comprising: a. a hull; b. a second wing having a second upper surface and a second highly swept subsonic leading edge portion arranged to develop separated flow at high angles of attack which develops into a second vortex moving over the surface of the second wing; c. the second leading edge portion having a second outer surface and on the second outer surface a second boundary layer control suction strip means extending along the second leading edge portion; d. a suction means arranged or adapted to draw in outside air through both the first and second suction strip means; e. the first and second wings being positioned on opposite sides of the fuselage; and f. wherein the suction means is arranged in a manner and the first and second suction strip means are positioned, configured and arranged in a manner such that operation of the suction means to draw the outside air reduces the air flow separated by the first and second suction strip means so that development of the first and second vortices is reduced. 19. A method of operating a supersonic aircraft, the method comprising: a. Providing the aircraft with a wing having an upper surface and a highly swept subsonic leading edge portion arranged to develop, at high angles of attack, separated flow which develops into a vortex moving over the upper surface of the wing; b. providing a boundary layer control suction strip means on or in an outer surface of the leading edge portion extending along the leading edge portion; c. Operating a suction means during a high angle of attack mode to draw outside air through the suction strip means to reduce or mitigate separated airflow so that the development of the vortex is reduced or mitigated.