CN100408428C - slotted aircraft wing - Google Patents

Abstract

An aircraft wing includes a leading airfoil element (36) and a trailing airfoil element (38). At least one slot (12) is provided in the wing in at least one transonic state of the wing. The slot (12) may extend spanwise along only a portion of the wing span, or may extend spanwise along the entire wing span. In either case, the slot (12) allows a portion of the air flowing along the lower surface (18) of the leading airfoil element (36) to be separated and flow over the upper surface of the trailing airfoil portion (38) to achieve improved performance in the transonic regime.

Description

slotted aircraft wing

Cross References to Related Applications

This application claims priority from US Provisional Patent Application No. 60/417,355. US Provisional Patent Application No. 60/417,355 filed October 9, 2002, the contents of which are hereby incorporated by reference in their entirety.

source of invention

The inventions described herein were made in part by employees of the United States Government and may be made and used by and for the United States Government for Government purposes without payment of any royalties.

technical field

In general, the present invention relates to aircraft and, more particularly, to slotted aircraft wings and methods of improving cruise performance of aircraft.

Background technique

The wings of many aircraft are designed using conventional airfoils. The upper and lower surfaces of a conventional airfoil meet to form a blunt or rounded leading edge, and a sharp trailing edge.

Conventional airfoils are also used for transonic wings (ie, wings designed for transonic flight). Transonic flight occurs when the airflow velocity on the vehicle is a mixture of subsonic flow (eg, flow speed less than the speed of sound) and supersonic flow (eg, flow speed greater than the speed of sound). The air flowing on the upper surface of the wing is accelerated due to the curvature of the upper surface to generate lift. As a result, the speed of the aircraft at which a portion of the airflow on the aircraft reaches the speed of sound (eg, reaches the speed of sound) may be less than one Mach number.

Simply put, the Mach number is the ratio of the speed of an aircraft to the speed of sound at the aircraft's current altitude. Mach 1 is achieved when the aircraft travels at the speed of sound. The critical Mach number (M _crit ) is the Mach number at which the airflow along the aircraft reaches the speed of sound somewhere.

When the airflow in any part of the aircraft does reach the speed of sound, a shock wave is created there. If the Mach number of the aircraft exceeds the critical Mach number, supersonic airflow will be generated on the upper and lower surfaces of the airfoil, resulting in shock waves throughout the airfoil. During transonic flight, there are often multiple local supersonic regions bounded by shock waves.

Across the shock, both the pressure and density of the air increase considerably, resulting in non-isentropic or irrecoverable losses, classified as wave drag. When the Mach number of the aircraft increases, the drag will increase significantly, which is called transonic drag growth. The shock wave slows the airflow and thus increases the pressure, causing an adverse pressure gradient across the shock wave. Depending on the strength of the shock, the adverse pressure gradient may cause a localized separation of the airflow from the wing surface at the base of the shock. During transonic flight, the shock wave and the boundary layer separation caused by the shock wave are always the main components of the overall drag of the aircraft.

The Mach number at which transonic drag begins to increase substantially is known as the "drag divergence Mach number" (M _dd ). It is often economically impractical to operate under such conditions because the Mach number of the aircraft slightly exceeding the drag divergence Mach number results in a significant increase in the drag of the aircraft.

Various methods have been employed for increasing transonic drag to higher Mach numbers, thereby reducing wave drag at a given transonic speed. Some of the more common approaches include the use of highly swept wings, thin airfoils, and aft-camber wings, which are relatively expensive to manufacture. A supercritical airfoil produces a higher critical Mach number. Supercritical airfoils generally have a flattened upper surface that reduces airflow acceleration and a highly raised rear section that provides most of the lift. A rear-loaded wing moves the center of lift aft, resulting in a larger pitch-down moment. An increase in pitch down moment eventually requires double movement of the wings and horizontal stabilizer to balance the vehicle in flight. The drag associated with balancing the aircraft is referred to as trim drag. Larger pitch down moments generally increase trim drag.

In addition to considering aerodynamic factors, other factors also limit the thickness of the actual airfoil. For example, thinner wings provide less fuel capacity. Also, the use of a thinner wing often increases the overall weight of the wing, since the thinner wing has a shallower structural box.

Using larger wings also increases the drag divergence Mach number, which reduces wave drag for a given transonic flight speed. For larger wing areas, an airfoil with a lower lift coefficient is required, which also results in less wave drag. However, the increased wetted area of a larger airfoil generally increases the frictional drag of the airfoil surface to such an extent that the additional surface frictional drag can offset or exceed any reduction in wave drag.

US Patent 6,293,497, entitled "Aircraft With Unswept Slotted Cruise Airfoil", discloses a non-swept, or substantially unswept, wing using slotted cruise airfoil technology , compared with the swept-back unslotted aircraft wing, it has a higher cruising speed and can also obtain greater lift when flying at low speeds. The entire contents of US Patent 6,293,497 are hereby incorporated by reference for full discussion.

Contents of the invention

The aircraft wing comprises at least one front airfoil element and at least one rear airfoil element. During at least one transonic state of the wing, the wing should have at least one slot. The slit may extend in the direction of the span along only a part of the span of the wings, or may extend in the direction of the span of the wings along the entire span of the wings. In either case, the slots separate some of the air flowing along the lower surface of the front airfoil element and allow it to flow over the upper surface of the rear wing element, thereby improving span. Sonic state performance. In an exemplary embodiment, the wing includes a partial span slot preferably starting approximately mid-span and extending outward to the wingtip at least such that the wingtip effect is substantially attenuated or surpassed by the slot effect .

Another form of the invention provides methods for flying an aircraft wing. In one embodiment, a method consists essentially of adjusting a slot provided between a forward airfoil element and an aft airfoil element in at least one transonic condition such that performance in the transonic condition is improved.

In another embodiment, a method consists essentially of using a slot to divert a portion of air flowing along a lower surface of the wing during at least one transonic regime of an aircraft wing to separate and flow over an upper surface of the wing. The diversion of said air at least delays separation of the airflows leading to increased drag, thereby improving performance in transonic regimes.

In a further embodiment, a method substantially comprises actuating a flap arrangement during a cruise condition to adjust said flap arrangement to improve cruise condition performance.

Further areas of applicability of the present invention will become apparent from the detailed description that follows herein. It should be understood that the detailed description and specific examples, while indicating at least one exemplary embodiment of the invention, are intended for purposes of illustration only and are not intended to limit the scope of the invention.

Description of drawings

A more complete understanding of the invention can be obtained from the detailed description and accompanying drawings, in which:

Figure 1 is a top view of a swept wing including partial span slots according to one embodiment of the present invention;

Figure 2 is a top view of a swept wing including full span slots according to another embodiment of the present invention;

Figure 3 is a top view of a conventional unslotted wing showing the shock location and supersonic flow region at moderate cruise speeds for lift coefficient and Mach number;

Figure 4 is a top view of the partially spread slotted wing shown in Figure 1, showing the shock location and supersonic airflow region at moderate cruise speeds for lift coefficient and Mach number;

Figure 5 is a top plan view of the full-span slotted wing shown in Figure 2, showing the location of the shock and the region of supersonic airflow at moderate cruise speeds for lift coefficient and Mach number;

Figure 6 is a side sectional view of the wing shown in Figure 1, showing the front and rear of the slotted wing region at the point of discontinuity in plan structure, according to one embodiment of the present invention for cruise flight end airfoil profile;

Fig. 7 is a side sectional view of the wing shown in Fig. 1, showing the wing section at the root and at the point of discontinuity in the plane structure of the non-slotted wing region according to one embodiment of the present invention;

Fig. 8 shows that the airfoil section of the front end and the rear end airfoil section overlapping as shown in Fig. 6 is in the airfoil section of the discontinuity point of the plane structure shown in Fig. 7;

Figure 9 is a line diagram summarizing the results of wind tunnel tests, one of which is a model of wing, fuselage and vertical tail with partial span slots, and the other is a model with conventional transonic Models of wings, fuselage and vertical tail;

Figure 10 is a line graph summarizing the results of wind tunnel tests. One of the models used in the wind tunnel test is an aircraft model equipped with a partially spanned wing, and the other is an aircraft model equipped with a conventional transonic wing ;

Figure 11 is a top view of a slotted wing including a wingtip device according to another embodiment of the invention;

Figure 12 is a simplified block diagram of the active control system associated with adjusting and adjusting the operation of the slotted airfoil;

Figure 13 is a top view of a wing with two partial span slots according to another embodiment of the present invention;

Figure 14 is a top view of a wing with two partial span slots according to another embodiment of the invention;

Figure 15 is a top view of a slotted wing, wherein the slots comprise a plurality of independently adjustable sections;

Figure 16A shows the pressure distribution of a conventional non-slotted airfoil;

Figure 16B shows the pressure distribution of a slotted airfoil;

Figure 17 is a computational fluid dynamics (CFD) model sample of the airflow or pressure field of the two-dimensional slotted airfoil design;

Figure 18A is a perspective view of a finite element model of a partially spread slotted wing in accordance with at least one embodiment of the present invention;

Figure 18B is a more detailed perspective view of the flap support shown in Figure 18A;

19A and 19B are sample 3D CFD models of the airflow or pressure fields on the lower wing surface of partially spread slotted wings with and without flap supports, respectively, in accordance with at least one embodiment of the present invention. ;

Figure 20 is a stowed side view showing an airfoil with a single slotted trailing edge flap;

Figure 21 is a side view of the airfoil shown in Figure 20, showing its single slotted trailing edge flap partially deployed;

FIG. 22 is a side view of the airfoil shown in FIG. 20 but with the single slotted trailing edge flap deployed at a greater offset than that of FIG. 21 .

Detailed ways

The following descriptions of different embodiments of the present invention are merely exemplary descriptions of the essence thereof, and do not limit the present invention, its application, or uses. For example, various embodiments of the present invention are expected to be applicable to a wide variety of aircraft (e.g., but not limited to fighter jets, business jets, private jets, supersonic impact aircraft, etc.) regardless of the manner in which the aircraft is piloted (e.g. , especially direct type, remote control type, automatic control type or combined type, etc.). Therefore, specific references to aircraft herein should not be construed as limiting the scope of the invention. Furthermore, the various embodiments of the present invention are also expected to be broadly applicable to lift-generating surfaces of aircraft (e.g., but not limited to, fixed wings, variable geometry wings, rotating wings, right half-span wings, left half-span wing, full-span wing, straight wing, swept wing, delta wing, horizontal tail, tapered wing, non-tapered wing, oblique wing, etc.). Accordingly, specific reference to airfoils herein should not be construed as limiting the scope of the invention.

Also, specific terms used in the following description are for reference purposes only, and thus are not intended to limit the scope. For example, the terms "upper", "lower", "above" and "below" refer to directions in the drawings with reference to them. The terms "front", "rear", "back" and "side" describe the orientation of certain parts of an element, and where these references are used throughout but are not limited thereto, reference may be made to the text describing the element in question and to the associated appendix. diagram to understand. Such terms may include the words specifically mentioned above, derivatives and words of similar import. Similarly, "first," "second," and other numerals referring to structures do not imply a sequence or order unless clearly indicated by the context.

Fig. 1 shows a swept wing 10 of an aircraft according to an embodiment of the present invention. As shown, the swept wing 10 includes a front airfoil element 36 and an aft airfoil element 38 . In at least one transonic state of the wing 10 at least one partial span gap 12 is defined between the front end profile element 36 and the rear end profile element 38 .

FIG. 2 shows another embodiment of a swept wing 110 . As shown, the swept wing 110 includes a front airfoil element 136 and an aft airfoil element 138 . In at least one transonic state of the wing 110 at least one full-span slot 112 is defined between the front airfoil element 136 and the rear airfoil element 138 .

The partial span slot 12 and the full span slot 112 separate part of the air flowing along the lower surface of the leading end element 36, 136 and flow over the upper surface 20, 120 of the trailing end element 38, 138, thereby Operating a wing improves its performance in one or more phases entering or approaching the wing's region of increased transonic drag or approaching the high-speed buffeting boundary; of which the transonic cruise regime and the transonic maneuver regime are examples of such phases . In at least some embodiments, both the partial-span slot 12 and the full-span slot 112 comprise aerodynamically smooth, non-smooth (bluff indented ) concave channel, as described below.

As used herein, "partial span slot" refers to and includes one or more slots, each slot extending spanwise along only a portion of the span of the wing. That is, the partial-span slotted wing does not have a single slot extending completely from root to tip. In the exemplary embodiment, the slots of the part-span slotted wings preferably start a little inward from roughly or slightly in the middle of the span and extend outward to the wingtips, at least so that the wingtip effect is substantially attenuated or exceeds the slot effect . FIG. 1 shows an exemplary wing 10 with partial span slots 12 .

"Full-span slot" as used herein relates to and includes slots starting substantially from the proximal end of the wing root and extending substantially to the wingtip (at least until the wingtip effect reduces the slot effect), excluding joints located at said full-span The necessary support frame for the wing structural elements before and after the seam. Such braces generally affect the entry of full-span slots on the lower surface of the wing, but not the exit of full-span slots on the upper surface of the wing. FIG. 2 shows an exemplary full-span slot 112 extending from wing root 114 to wing tip 116 .

As used herein, "transonic cruise regime" refers to and includes the relatively high speed phase of the airfoil over which airflow as shown includes localized regions of supersonic airflow, eg, FIGS. 3 , 4 and 5 . In other words, a wing entering or approaching a region of increased transonic drag of the wing or approaching the boundary of high speed buffeting cruises at relatively high speed. Also, as used herein, "transonic regime" refers to and includes one or more phases of flight in which, while the wing is in flight, but not necessarily in cruise flight, a section of increased transonic drag into or near the wing or Approaching the boundary of high-speed chattering. Exemplary transonic states of the airfoil include, but are not limited to, transonic cruise states and transonic maneuvers.

Figures 1 and 2 are simplified planar structures of right wing designs currently applied to commercial aircraft. The designs in the above two figures are respectively equipped with partial span slots and full span slots. The commercial aircraft also includes a left wing having substantially the same performance within the flight curve. Thus, when the right wing is provided with slots, the left wing (not shown) is often also provided with equivalent parts or corresponding slots.

With respect to half-span wing terms (eg, right and left wings), the half-span portion of 0% is generally known as the symmetrical or mirror image location of the right and left wings. Typically, the 0% half-span portion is the middle of the fuselage where the wings are attached. When discussing half-span wings, the term "half-span" refers to the distance from the 0% half-span location to the 100% half-span location at the wingtip. It should be noted, however, that embodiments of the present invention should not be limited to half-span wings, but are equally applicable to full-span wings (eg, flying wings in particular). Also, as shown in Figure 11, the terms "span" and "half-span" as used herein do not include one or more wingtip devices that may be mounted or disposed on the wingtip. However, this should not be used to limit the scope of the present invention, and embodiments of the present invention are expected to be applicable to a wide variety of airfoils, including, but not limited to, wings with wingtip devices and wings without The wing of the pointed device. In other embodiments, the wingtip device may indeed be provided in at least a part of a partial or full span slot.

Referring further to FIG. 1 , the partial span slit 12 may extend in the spanwise direction along the part of the half-span of the wing 10 where airflow separation results in increased drag when the wing 10 is in a transonic state. Partial span slots 12 may be placed at locations on said wing 10 where fluid dynamic calculation simulations of three-dimensional airflow suggest would cause airflow separation pressure fields on wing upper surface 20 .

In the exemplary embodiment, the partial span slot 12 extends from an approximately half span region 28 to an approximately half span region 30 . The half-span locations 28 and 30 coincide with the Yehudi point or planar discontinuity point 32 and the wing tip 16, respectively, although this need not be the case. In other embodiments, the partial span slots 12 may originate from other inboard locations where the inclusion of the slots does not interfere with low velocity control surfaces or interfere with other elements such as fuel tanks and landing gear from the wing. 10 planar structures for integration. Also, the partial span slot need not extend completely to the wing tip. Conversely, the part-span slots may extend substantially to the wingtips, but cease when the wingtip effect diminishes the performance improved by the slots.

The particular chordwise locations of the partial span slots 12 and full span slots 112 (FIG. 2) are likely to be determined, at least in part, by considerations such as specific low speed control surfaces and the relationship of other elements such as fuel tanks and landing gear to the aircraft. Integration with wing planar structures. In an exemplary embodiment, the chordwise location of each slot 12 and 112 is at approximately 70% to approximately 90% of the chord.

During use, each slot 12 and 112 separates a portion of the air flowing along the lower surface 18 of the front end airfoil element 36,136 and over the top of the rear end airfoil element 38,138. Flow over the surface 20,120. In this case, the slot at least delays the boundary layer separation and pushes the shock wave generated by the supersonic flow further to the rear of the wing. The effect (slot effect) of the presence of slots on the supersonic airflow (indicated by region B) and shock position (indicated by solid line A) of the entire upper wing surface (slot effect) can be compared by comparing Figures 3 (unslotted wing), 4 (partial spread slotted wing), and 5 (full spread spread slotted wing) are seen. As described below, this "slit effect" improves the performance of the airfoil in the transonic regime.

The manner in which the "slot effect" prevents or at least delays boundary layer separation is described below and detailed in US Patent 6,293,497 entitled "Aircraft With Non-Swept Slotted Cruise Wing Airfoil". The content of US Patent 6,293,497 is incorporated herein in its entirety and is fully discussed herein.

Referring further to FIG. 1 , the partial-span slotted wing 10 comprises at least one wing region 22 without a slot and at least one other region 24 with at least one partial-span slot 12 . For ease of identification and illustration, and not limitation, the wing region 22 may also be used to refer to an unslotted wing region 22 because the unslotted region 22 does not define a slot, and the wing region 24 may also be used to refer to the slotted wing area 24 since the slotted area 24 defines at least one partial span slot 12 . It should be noted, however, that either or both of the wing regions 22 and 24 may have any number (eg, one or more) of slots, some of which may only be provided in unconfigured areas such as leading edge slots. Elevating and/or stabilizing and control equipment such as wings, ailerons, flaps, spoilers, etc.

As shown, the unslotted wing region 22 is disposed spanwise between the half-span regions 26 and 28, and the slotted wing region 24 is thus disposed between the half-span regions 28 and 30. . The half-span locations 26, 28 and 30 correspond to the wing root 14, the planar discontinuity point 32 and the wing tip 16, respectively, but this is not essential.

The slotted wing area 24 can only be arranged on the wing area where the critical Mach number is reached when flying at a higher cruising speed. The method for determining which part of the wing will reach the critical Mach number at cruise will be described below. Other regions of the wing where the Mach number will not reach a critical value may include the non-slotted wing region 22 .

In this example, the unslotted wing area 22 is disposed inside (eg, near the fuselage) of the discontinuity point 32 in the planar structure as shown. For commercial aircraft, where the landing gear is retracted, a longer chord length is generally used in the inboard area of the wing. The corresponding wave drag of the inboard portion is often minimal for relatively long chord lengths because the airfoil has a relatively low local lift coefficient (C ₁ ). If the inboard section does not reach the critical Mach number at cruise, then there is no need to use the section span slot 12 to increase the magnitude of the Mach number. Therefore, the unslotted wing region 22 can be arranged on the inner wing portion that does not reach the critical Mach number during cruising, thereby avoiding or eliminating the use of partial wingspan slots at positions that do not need to increase the Mach number during cruising. The resulting airfoil drag increases. Furthermore, the use of an unslotted wing region 22 on the inboard portion enables conventional elevation systems (e.g., conventional flaps and strips) to be applied to the inboard portion of the wing 10, which is also an aspect of the present invention. Additional advantages in the examples. Also, it should be noted that the various embodiments of the present invention should not be limited to airfoils having inboard portions that do not reach critical Mach numbers in cruise. Embodiments of the present invention are indeed expected to be applicable to a wide variety of airfoils, including, but not limited to, wings with inboard regions that reach the critical Mach number in cruise and wings that do not reach the critical Mach number in cruise. Mach numbers for the inboard region of the wing.

Although the partially spread slotted wing 10 is shown with a single unslotted wing region 22 and a single slotted wing region 24, this is not required. The partially spread slotted wing 10 may have any number (e.g., one or more) of unslotted wing regions 22 and any number of slotted wing regions 24, each of which may be There may be any number of seams without departing from the spirit and scope of the invention. Switching between slotted and unslotted wing areas 22 and 24 is possible as the specific requirements of the wing design surface may use more than one unslotted wing area and/or more than one slotted wing area. will appear multiple times across the half-span of the wing. For example, another embodiment of the partially spread slotted wing includes an inboard unslotted wing area, a central slotted wing area, and another unslotted wing area between the wingtip and the slotted wing area. wing area.

FIG. 6 shows the airfoil profile of the front and rear airfoil elements 36 and 38 of the wing 10 at the discontinuity point 32 of the planar structure. The front end airfoil element 36 includes an upper surface 40 , a lower surface 42 , a leading edge 44 and a trailing edge 46 . Similarly, the rear end airfoil element 38 also includes an upper surface 48 , a lower surface 50 , a leading edge 52 and a trailing edge 54 . The part-span slot 12 is defined between a leading edge 46 in the front end airfoil element 36 and a trailing edge 52 in the rear end airfoil element 38 . The section of the partial span slot 12 is shown as a gap or space separating the leading edge in the front end airfoil element 36 from the trailing edge 52 of the rear end airfoil element 38 . During flight, the partial span slots 12 separate part of the air flowing along the lower surface 42 in the front end airfoil element 36 and cause it to flow on the upper surface 48 of the rear end airfoil element 38 .

With further reference to FIG. 6 , portions of the front end airfoil elements 36 may overlap or overhang portions of the rear end airfoil elements 38 . Thus, the sum of the chord lengths of the leading and trailing airfoil elements 36 and 38 exceeds one hundred percent (100%) of the slotted wing region 24 (e.g., the leading edge 56 terminal and the trailing edge 34 distance between terminals). In at least one embodiment, the gap is minimized but of sufficient size so that the boundary layer along the lower surface 42 of the nose airfoil element 36 cannot contact the upper surface of the nose airfoil element 38. The boundary layers on 48 mix or converge.

Fig. 7 is the sectional view of described non-slit wing area, shows among the figure that the airfoil portion 64 on half-span position 26 is at half-span position 28 or planar structure discontinuity point 32 places and described non-slit machine The airfoil sections 66 of the wing region 22 coincide. Since the unslotted airfoil region 22 is reverse-swept and trapezoidal, the leading edge 68 and the trailing edge 70 in the root airfoil portion 64 can be placed in the airfoil at the point of discontinuity 32 in plan configuration. Front portion of leading edge 72 and trailing edge 74 of portion 66 .

Figure 8 is a side sectional view of the slotted wing region 24 as shown in Figure 6, showing its front end airfoil portion 36 and rear end airfoil portion 38 at the discontinuity point 32 of the planar structure, and the planar structure The airfoil portion 66 in the unslotted wing region 22 at the discontinuity point 32 overlaps, as shown in FIG. 7 . At the discontinuity point 32 of the planar structure, the leading edge 72 of the unslotted wing region 22 relatively smoothly transitions to the end of the leading edge 56 of the slotted wing region 24 . At the planar discontinuity point 32 , the trailing edge 46 of the main wing portion 58 of the slotted wing region 24 transitions relatively smoothly into the upper surface of the unslotted wing region 22 . Also at the planar discontinuity 32, the trailing edge 34 terminal of the slotted wing region 24 is offset downwardly from the trailing edge 74 of the unslotted wing region 22 by an appropriate amount so that the flow-through portion The air of the span slot 12 can pass through the trailing edge 34 termination of the slotted wing region 24 .

The part-span seam 12 may start rather abruptly at the point of discontinuity 32 in the planar structure. That is to say, the gap separating the trailing edge 46 in the front end airfoil element 36 from the leading edge 52 in the rear end airfoil element 38 is not tapered, nor There is no gradual increase in size. Thus, at the beginning of the portion of the span 12 in the discontinuity of planar structure 32, the transition from the unslotted wing region 22 to the slotted wing region 24 is relatively not smooth. It should be noted, however, that other embodiments may include progressively onset partial span slots 12 or tapered partial span slots 12 such that the unopened section has partial span slots 12 in plan discontinuity 32. The slotted wing region 22 transitions relatively smoothly into the slotted wing region 24 .

In at least one embodiment, the gap at discontinuity 32 in the planar structure is sealed by, for example, a baffle (not shown). The baffle may be planar and positioned across the gap such that the baffle is in the same direction as the flight.

The partial span slots 12 may be provided between the main wing section 58 and raising or stabilizing and control devices such as flaps 60, ailerons, spoilers, and the like. In the exemplary embodiment, the partial span slot 12 is defined between the trailing edge 46 of the main wing section 58 and the leading edge 52 of the flap 60 . Thus, the partial span slots 12 separate part of the air flowing along the lower surface 42 of the main wing section 58 and allow it to flow over the upper surface 48 of the flap 60 .

Said flaps 60 may be operated in conjunction with an active control system 61 ( FIG. 12 ) which in turn is associated with the operation of an actuator structure such as the U.S. Patent 5,788,190 discloses the flap actuator structure. The content of US Patent 6,293,497 is incorporated herein in its entirety and is fully discussed herein.

The actuator structure is connected to the flap 60 and the main wing section 58 for moving the flap 60 corresponding to the main wing section 58 to allow deployment of the flap 60 and/or Or adjust said slit 12 for flight conditions. For example, the flaps 60 are movable between a fully extended position (not shown) for a landing or take-off condition and a stowed position 62 associated with a cruise condition. Or, for example, the flap 60 can be moved to narrow or widen the slot 12, and the flap 60 can be raised or lowered to change the relative height between the flap 60 and the main wing section 58. , and/or the flap 60 can be rotated to adjust the angle or pitch between the flap 60 and the main wing section 58 .

In FIG. 2, the wing 110 comprises a main wing structure "box" or element 136 defined between the trailing edge 146 and the leading edges 152, 152' of the inboard and outboard ailerons 138, 138'. The full-span slots are 112. As shown, the trailing edge of the main wing element 136 partially overlaps or overhangs the forward portions of the flaps and ailerons 138, 138'.

Either or both of the flaps 138 and ailerons 138' may be coupled to an actuator arrangement to allow the slots 112 to be adjusted for a particular flight condition of the wing 110. By way of example, the actuator structure for adjusting and trimming the slit 112 can use the flap actuator structure disclosed in US Patent No. 5,788,190.

It should be noted that other arrangements are possible for partial span slots, full span slots, and trailing edge systems (eg, flaps, ailerons, spoilers, etc.). For example, another embodiment includes a blade master system wherein the slot is provided between the blade and the main flap, with the blade before the slot and the main flap behind the slot.

In at least some embodiments, a closeable full-span or partial-span slot is provided, which can be closed as flight conditions require (eg, including low speed phases such as takeoff, landing, climb, etc.). Closing the seam eliminates the loss of surface frictional resistance caused by the seam. During high-speed flight conditions (for example, transonic cruise conditions), the slots may be partially or fully opened.

In other embodiments, the partial-span or full-span slots may be permanently placed on the wing, making the slots independent of the different elements (e.g., flaps, ailerons, strips) that make up the wing. panels, spoilers, other elevating devices, other stabilizing and control devices, etc.) position or shape (e.g., fully deployed, partially deployed, stowed). The existence of the slot is independent of the flight phase of the aircraft (eg, landing, takeoff, climb, aerobatics, cruise, level flight, acceleration, deceleration, etc.). For example, the slots may be used as fixed openings in the movable parts of flaps and ailerons, so that the slots remain fully open when deploying and retracting the movable parts.

Figure 13 shows an exemplary embodiment of a swept wing 210 comprising two partial span slots 212 and 212'. The slot 212 is set between the trailing edge 246 of the front end airfoil element 236 and the front edge 252 of the rear end airfoil element 238, and the slot 212' is set at the front end airfoil element 236' between the trailing edge 246' and the leading edge 252' of the rear airfoil element 238'.

Figure 14 shows another embodiment of a swept wing which includes two partial span slots 312 and 312'. The slot 312 is set between the trailing edge 346 of the front airfoil element 336 and the front edge 352 of the rear airfoil element 338, and the slot 312' is set at the front end airfoil element 336' Between the trailing edge 346' and the leading edge 352' of the rear airfoil element 338'.

Figure 15 shows another embodiment of a swept wing 410 comprising a slot having a plurality of sections 412, 412', 412", each of which is independently adjustable. As shown, each slot segment 412, 412', 412" is set at the trailing edge 452, 452', 452" of the main wing structural box 436 and independently movable lifting or stabilizing and control equipment 438, 438 438', 438" between leading edges 446, 446', 446". Each device 438, 438', 438" is coupled to an actuator structure, such as the flap actuator structure described in US Patent No. 5,788,190. The actuator structure can independently move the devices 438, 438', 438" associated with the main wing section 436 to adjust and trim the slot segments 412, 412' and 412" to fit the wing 410 specific flight status.

In another form, the invention provides methods of designing aircraft wings. In one embodiment, in general terms, a method is to adjust the slots set between the front end airfoil element and the rear end airfoil element in at least one transonic condition to complete the performance adjustment in the transonic condition Improve. Adjusting the slit may use one or more of the following operations: adjusting the slit separating the leading and aft airfoil elements, the slit defining the slit; adjusting the relative height between the leading and aft airfoil elements; and Adjust the angle between the front and rear airfoil elements. In an exemplary embodiment, said forward and aft airfoil elements respectively comprise a main wing portion and a flap arrangement, and adjusting said slot comprises actuating said flap arrangement. In at least some embodiments, the method may further include closing, or at least reducing the width of, the slot if required by flight conditions, eg, in subsonic conditions (eg, takeoff, landing, climb, etc.).

In another embodiment, a method of designing an aircraft wing generally includes the use of at least one slot defined by said wing that allows the wing to flow along the lower surface of said wing in at least one transonic regime Part of the air separates and flows over the upper surface of the airfoil. Turning the air around prevents, or at least delays, the airflow separation that occurs at transonic speeds and leads to increased drag, resulting in improved transonic performance. It should be noted, however, that air turns do not need to occur during all phases of flight. For example, if the flight condition warrants, for example, a subsonic regime (eg, takeoff, landing, climb, etc.), the method may further comprise closing or at least reducing the width of the slot. Additionally, the method may also include opening the slot when the airfoil is at or near subsonic speeds. Furthermore, the method may also include adjusting the slot according to the flight condition of the wing.

In a further embodiment, a method of designing an aircraft wing is provided, wherein the aircraft wing includes a main wing section, a flap arrangement, and at least one cruise setting between the main wing section and the main wing section. Slit between flap devices. Generally, the method is to drive the flap device during cruising to adjust the flap device so as to improve the performance during cruising.

The portion of the wing that will reach the critical Mach number depends at least in part on the planar configuration of the wing, the thickness distribution and the spanwise distribution of the aerodynamic loads (span loads). In order to reliably determine which wing section will reach the critical Mach number, a highly accurate computational modeling approach with full, nonlinear form of compressibility effects including viscous/turbulent boundary layers and wake flow impact. Different levels of simplified approximations can also be included in computer models, such as methods based on boundary layer approximations (coupled inviscid/boundary layer methods) and Navier-Stowe Navier-Stokes source (e.g. "thin-layer" approximation, where some viscous conditions with minor effects are ignored).

的纳维-斯托克斯(Navier-Stokes)类型的示例性CFD分析计算机软件由New Hampshire，Lebanon的Fluent Inc.Corporation提供；由California，Agoura的MetacompTechnologies提供的CFD++

；以及由Washington，Redmond的Analytical Methods，Inc.提供的NSAERO。A "Flow Solver" based CFD analysis code is available for determining the flow properties of a given aerodynamic shape. Thus, when the shape of a particular wing is known, analysis can be performed to determine, for example, the Mach criticality of different parts of the wing or the overall aerodynamic performance of the wing. An exemplary CFD analysis computer software MGAERO of the coupled inviscid/boundary layer type is provided by Analytical Methods, Inc. of Washington, Redmond. including FLUENT

Exemplary CFD analysis computer software of the Navier-Stokes type is provided by Fluent Inc. Corporation of Lebanon, New Hampshire; CFD++ by Metacomp Technologies of Agoura, California

and NSAERO provided by Analytical Methods, Inc. of Washington, Redmond.

The performance of the partial-span slot configuration was theoretically analyzed through computational fluid dynamics (CFD) studies and validated through wind tunnel testing to provide improved performance over conventional transonic airfoil designs . Regarding the CFD model method, two-dimensional research and analysis have been carried out on two-dimensional slotted airfoil design for many years, so it is already a prior art. In FIG. 17 , a sample CFD solution of the airflow or pressure field around a two-dimensional slotted airfoil design 80 is shown.

Since CFD has not been expanded, applied, nor applied to three-dimensional slotted wings, the embodiments of the present invention still need to develop, optimize and use specific tools and methods to carry out detailed three-dimensional CFD design of slotted wings and analysis. Furthermore, as described below, aspects of the present invention also require the use of wind tunnel testing to verify the CFD output.

As shown in Figures 3, 4, and 5, the CFD output includes a shock model sample and supersonic flow region over the entire wing when flying at moderate cruise speeds with lift coefficient and Mach number. More specifically, Figures 3, 4, and 5 show the shock waves along the upper surface of a conventional, partially-slotted, and fully-slotted wing, respectively, when flying at moderate cruise speeds with lift coefficient and Mach number Location and region of supersonic flow.

Referring now to Figures 18A and 18B, a finite element model of a partially spread slotted wing 82 is shown. As shown, the part-span slotted wing 82 includes a part-span slot 84 with a flap bracket 85 . In Fig. 18B, the flap support 85 is shown in more detail.

In FIG. 19A , the CFD output includes a model sample of the airflow or pressure field on the lower wing surface 86 of a partially spread slotted wing 87 with flap supports 88 . In Figure 19B, the CFD output includes a model sample of the airflow or pressure contours on the lower airfoil surface 86' of a partially spread slotted wing 87' without flap supports. Therefore, comparing Figures 19A and 19B, it can be concluded whether the flap support has an influence on the surface pressure of the lower wing.

Using the 3D CFD tools and methods described here, it can be concluded that the cruise Mach number (ΔM) is increased by 0.025 and the aerodynamic efficiency (ΔML/D) is increased by 0.025 for the partially spread slotted wing compared to the conventional transonic wing. -1.0%. It should be noted that these values (eg, 0.025 and -1.0%) are set forth herein for illustrative purposes only and should not be construed as limiting the scope of the invention. Furthermore, these values were obtained using a CFD model with a partially spread slotted wing, fuselage and vertical stabilizer and a CFD model with a conventional transonic wing model, fuselage and vertical stabilizer. Neither model has horizontal stabilizers, engine pods or struts.

The 3D CFD design and analysis tools and methods and the results obtained therefrom have been tested in a transonic wind tunnel. More specifically, different wind tunnel tests were used to determine the Mach number change (ΔM) of a partially spread slotted wing at cruise compared to a conventional transonic wing design, thereby determining the partial wing spread The relative aerodynamic performance (ΔML/D) between slotted wing and conventional transonic wing design, and determine the impact of pod integration and trim drag on vehicle integration, and the accuracy and accuracy of 3D CFD analysis Reliability is evaluated.

Figures 9 and 10 summarize the results of some wind tunnel tests. The wind tunnel test model in Figure 9 includes the wing (partially spread slotted wing or conventional transonic wing), fuselage, and vertical stabilizer, but does not include the horizontal stabilizer, engine pods, or supports. However, the wind tunnel test model in Figure 10 is fully configured, including wings (partially spread slotted or conventional transonic), fuselage, vertical and horizontal stabilizers, engine pods and supports.

The wind tunnel tests and computational fluid dynamics studies focus on or aim at studying aerodynamic performance. To ensure that improvements in aerodynamic performance are directly transferable, constraints and constraints need to be placed on the wing design so that improvements in aerodynamic performance do not degrade performance in other areas or parts. For example, tweaks that improve aerodynamics don't add weight to the structure. within these constraints, the partial-span slotted wing substantially increases cruise speed while providing acceptable lift and handling characteristics compared to conventional transonic wing designs at their cruise design speeds , at least assuring a comparable aerodynamic efficiency (ML/D) and range. The degree of improvement brought about by the examples of the present invention is expected to increase when the aforementioned constraints on the original design are removed. Partially spread slatted wings may bring additional improvements in aircraft efficiency when formal interdisciplinary research is conducted.

With regard to the performance improvement of the wing operating conditions under the transonic state, the embodiments of the present invention can achieve any one of the following effects or their combination: the wing cruising speed or critical Mach number is improved, the wing lift is increased, the Wing thickness, and/or Mach number is maintained at smaller swept wing angles. The physical factors that limit the performance of airfoils for transonic cruise are described in more detail below, as well as how designers weigh improvements in state of the art against improvements in wing thickness, speed, lift or drag, or a combination of these factors as described below .

The slots can be used to increase the drag-divergence Mach number (M _dd ) magnitude of an airfoil for a given sweep angle, lift coefficient, and thickness distribution, while at the same time improving, or at least maintaining, the airfoil's equivalent in cruising flight. aerodynamic efficiency (ML/D) and range. Aerodynamic efficiency is a dimensionless measure of performance, calculated by multiplying Mach number by lift and dividing by drag, and is especially important for long-distance vehicles. Wings having at least one slot that improves cruise performance can be flown at higher cruise speeds before the increase in transonic drag begins.

The ability of the slits to prevent, or at least delay, boundary layer or flow separation allows airfoils for the airfoils to be designed to produce a spread of pressure at transonic regimes, with suction on the upper surface compared to conventional airfoils. There is a drop in level (e.g., lower negative pressure coefficient at the upper surface), and the shock wave and subsequent pressure recovery move backwards. The resulting pressure distribution due to the presence of the slots provides a high drag-divergence Mach number (M _dd ), which cannot be achieved due to the boundary layer separation of unslotted conventional airfoils in the transonic regime. Drag-divergence Mach number.

The invention can also use conventional or unslotted airfoils, if any, on portions of the airfoil where the Mach number will not be critical. For example, if it has been determined that the inboard portion of the wing will not reach a critical Mach number during cruise, then for the inboard portion, the surface frictional drag associated with the slots can be avoided or eliminated by using an unslotted wing area consume. Also, the use of conventional and unslotted wing areas for the inboard part of the wing allows the use of conventional elevation systems (eg conventional flaps and slats) that are also applied to the inboard.

While fuel consumption is essentially the same for a vehicle with slotted wings, the increased cruising speed or Mach number of the vehicle does add other efficiencies. For example, a slotted wing could increase the speed at which an aircraft can cruise or fly before the increase in transonic drag begins, reducing travel time. In addition to providing great benefits to air passengers, faster flight also benefits airlines through reduced operating costs. For example, a reduction in flight duration requires less aircrew time and therefore pays less to the aircrew. In addition, faster flight procedures also reduce the frequency and cost of maintenance, since the required maintenance is often based on the number of flight hours the aircraft has.

US Patent 6,293,497 describes the physical factors limiting the performance of airfoils for transonic cruising and the trade-offs involved in maximizing the performance of airfoils for transonic cruising. Airfoil performance in transonic cruise applications can be characterized in four basic ways:

1) Airfoil thickness, often expressed as the maximum-to-thickness ratio (maximum thickness divided by chord length). Greater thickness is beneficial because it provides the space needed for fuel and mechanical systems, and because a deeper wing structure weighs less for the same strength.

2) Speed or Mach number at preferred operating conditions. The Mach number of the airfoil, after being corrected by factors related to the wing sweep angle, helps to directly increase the cruising speed of the aircraft.

3) Lift coefficient for preferred operating conditions. An increase in the coefficient of lift is advantageous because it allows for increased weight (for example, to fly longer distances with more fuel) or a higher cruising altitude.

4) The drag coefficient for the preferred operating state and other operating states that the aircraft will encounter when performing its mission. Reducing drag reduces fuel consumption and increases flight distance.

Other aspects such as pitching moment characteristics and lift capability at low Mach numbers are also important, but not as important as the above four basic aspects.

The four basic performance metrics collectively define a level of performance, often referred to as the "technique level" of the airfoil. These four basic properties bring conflicting problems to designers, that is, a design change to improve one of the basic properties will often cause at least one of the other three properties to degrade. Therefore, a good or optimal design for a given application requires a favorable compromise between these four properties, evaluating the overall performance of the aircraft to which this airfoil is applied. It should be noted that the compromises required to design an airfoil with a higher technical level (as determined by the above four performance indicators) will not always result in the best or optimal technical level for the aircraft as a whole, because the higher Skill level can have a bad effect on maximum lift, handling characteristics, or to a lesser extent vibration limitation.

Sometimes a restrictive assessment of the state of the art is made on the basis of only the first three of the performance measures above. In a restrictive sense, the technical level of an airfoil can be determined according to the target cruise operating state located in three-dimensional space, which is determined by the maximum-thickness ratio (tmac/c), lift coefficient (Cl) and Mach number (M ) to limit. In order to reduce a position in three-dimensional space to a single "index", an additional premise or rule is required to use the following equation:

ΔM=[-1(Δtmax/c)]+[-1/7(ΔCl)]

The above equations are based on the premise of what constitutes the same technology level and what provides an approach to the operating conditions involving two airfoils with the same technology level. The constants -1 and -1/7 are based on historical data (for example, comparing airfoils considered to be technically equivalent). However, it should be noted that the constants -1 and -1/7 are merely exemplary, and other suitable constants may also be used in the above equations.

To compare the state of the art of the two airfoils, an exemplary method tunes the two airfoils to the same point of tmax/c and Cl using the above equations, and then compares the resulting Mach numbers. Therefore, the difference in the technical level of the two airfoils can be represented by the difference in the Mach number.

Another example way to compare the state of the art of airfoils is to plot drag increase curves (comparing drag coefficient to Mach number for a constant lift coefficient). This type of curve can be used to demonstrate the lower drag operating range of the slotted airfoil (shown below the pressure distribution curve in Figure 16B) than the single element airfoil (shown below the pressure distribution curve in Figure 16A). Below) extends to a higher Mach number, slightly increasing lift for the same thickness. Of course, the slotted wing could be redesigned to take advantage of this technology for purposes other than higher speed, for example to achieve even higher lift at the same speed as a single element airfoil.

At any given level of technology, it is generally possible to design a wide range of individual airfoils suitable for different preferred operating conditions, and embodying different compromises between the four basic properties. For example, one airfoil may have a higher operating Mach number than another, but at the expense of reduced lift and increased drag. For competent designers, using modern computational fluid dynamics tools to design airfoils with varying degrees of skill is generally a straightforward task. On the other hand, it is often difficult to improve the technical level, for example, by improving one of the basic performances without affecting any of the other three basic performances. The higher the technical level used by the designer at the beginning, the more The tasks to be completed are more complex. If the wing under study is at a level representative of the state of the art, it is extremely difficult to make meaningful improvements.

The main factor limiting performance is related to the physical characteristics of the flow on the upper surface of the airfoil. Understanding these factors requires looking at the pressure distribution of a typical transonic cruising airfoil, plotted in units of the pressure coefficient (C _p ) on a negative scale, as shown in Figure 16A (cited in US Patent 6,293,497). For ease of reference, the shape of the airfoil 101 is shown below the pressure distribution diagram. On the _Cp scale as shown, _Cp = 0 is the static pressure of free fluid flowing away from the airfoil, assuming subsonic velocities. At each point on the surface, the value of _Cp corresponds, in addition to defining the pressure, to a specific flow velocity value outside the thin viscous boundary layer immediately above the surface. Negative _Cp (above the horizontal axis) represents low pressure and high velocity compared to free flow, while positive _Cp (below the horizontal axis) corresponds to high pressure and low velocity. A particular level of negative C _p is equivalent to the velocity of sound and is represented using dotted line 89 .

The lower curve 90 on the pressure profile represents the pressure on the lower surface 91 , or high pressure side, and the upper curve 92 represents the pressure on the upper surface 93 . The vertical distance between the two curves indicates the pressure between the upper surface 93 and the lower surface 91, and the area between the two curves is proportional to the total lift produced by the airfoil. Note that there is a high positive peak in the _Cp distribution 94 near the leading edge called the "stagnation point" 95, the point where incoming airflow first "sticks" to the airfoil surface, and at The velocity outside the boundary layer is zero. Note also that the _Cp distributions on the upper and lower surfaces converge at the trailing edge 96, thereby defining a single value _Cp 97, which is almost always a small positive value. This level of _Cp at the trailing edge has severe effects on the physical properties of the flow. Because the trailing edge _Cp is primarily determined by the overall airfoil thickness distribution, and thickness is often limited by a number of structural and aerodynamic factors, the designer has relatively little control over trailing edge _Cp . In addition to the stagnation point at the leading edge and the trailing edge, designers can have more control over the pressure distribution by changing the shape of the airfoil.

For a given airfoil thickness and Mach number, the problem of achieving a high level of technology boils down to a problem of maximizing a high coefficient at a low level of drag. It is often not possible to increase lift entirely by increasing subsurface pressure without reducing airfoil thickness. Therefore, the designer's task is to reduce the pressure on the upper surface to increase the lift as much as possible without greatly increasing the drag. In this regard, the pressure distribution shown in Figure 16A represents an advanced design approach. The operating conditions shown in the figure approximate the preferred operating conditions used for the early cruise portion of the aircraft mission. Drag under these conditions is fairly low, but increases rapidly as Mach number and/or lift coefficient increase.

Note that the upper surface _Cp 92 in the front half of the airfoil 101 lies above the dotted line 89, indicating that the flow there is mid-sonic. This supersonic region is terminated by a weak shock just aft of the midchord, superficially represented by the sudden increase of C _p 98 to the value characteristic of subsonic flow. The distribution of _Cp in the supersonic region 99 is intentionally designed to be nearly flat, and the pressure rises very gently, thereby preventing the shock wave from becoming stronger and causing drag to increase under other conditions. After the shock, the pressure 100 gradually increases, ie "pressure recovery", until it reaches the small positive value C _p 97 of the leading edge. The location of the shock and pressure distribution in the recovery region has been carefully designed to provide a balance between increased lift and increased drag.

Attempts to increase lift tend to throw the wings out of favorable balance and simultaneously increase drag. For example, one way to increase lift is to move shock wave 98 rearward. However, this would require a large recovery (since both the immediate aftershock C _p and the trailing edge C _p are essentially fixed), which would thicken the viscous boundary layer or even detach it from the surface, both of which would lead to significant resistance increases. Another way to increase lift is to reduce the pressure in front of the shock and even further forward (shifting the _Cp curve 99 up and increasing the supersonic velocity across the front of the airfoil), but this increases the pressure across the shock pressure, which also leads to what is known as shock wave drag. For the single-element transonic airfoil of the prior art, the compromise design between lift and drag has reached a high degree of sophistication, and it is impossible to carry out any major improvement on the technical level.

The shape and resulting pressure distribution of the slotted transonic cruising airfoil 523 is shown in Figure 16B (cited in US Patent 6,293,497). The airfoil 523 comprises two elements (front end element 560 and rear end element 561) separated by a curved channel (562, the slot) through which air generally flows from the lower surface 584 to the upper surface 564. In this example, the slot flange (565, trailing edge of the leading element) is located exactly aft of 80% of the total chord from the leading edge, with elements overlapping approximately 3% of the total chord length. The figure shows the pressure distribution of the two elements, which overlap where the airfoil elements overlap. Same as the conventional airfoil, the upper curves 566, 567 represent the generation of _Cp distribution on the upper surfaces 564, 583, and the lower curves 568, 569 represent the generation of _Cp distributions on the lower surfaces 584, 570 distributed. Note that there are two stagnation points 571, 572 and their corresponding high pressure peaks 573, 574, one for each peak, where the incoming flow attaches to the surface near each leading edge .

To begin to consider the physics of flow, it is important to note that the slotted airfoil 523 (just below the pressure profile in FIG. The one below) is faster and has a slightly higher lift coefficient, and for structural purposes both airfoils have the same effective thickness. In the operating conditions of the slotted airfoil, any single element airfoil of the same thickness has extremely high drag. The substantial technical advantage of the slotted airfoil comes from the fact that the final return of pressure 575 occurs very far aft, starting at about 90% of the total chord length with Weak shock wave 576. This pressure distribution is unlikely to occur in a single-element airfoil because boundary layer separation must occur, preventing the shock wave from traveling far aft. This mechanism can be loosely referred to as the "slit effect" by which the separation of the boundary layer is prevented by a combination of the following factors:

1) The boundary layer on the upper surface 583 of the front element 560 is affected by the weak shock wave 577 at the slot flange 565, but there is no post-shock pressure recovery on the front element. This is possible because the rear element 561 promotes an increased "spread velocity" at the trailing edge of the front element (the trailing edge CP 578 on the front element is a larger negative value, while the trailing edge CP on a single-element airfoil is generally positive).

2) Boundary layers on the upper surface and lower surface of the front element 560 converge at the trailing edge 565 and form a wake that flows on the upper surface 564 of the rear element and forms a wake on the upper surface 564 of the rear element. The boundary layer on the upper surface is still substantially different. Throughout the rear of the rear element 561 the wake is subject to strong pressure rises 575, 576, but strong turbulent mixing keeps the wake free from reverse flow.

3) The boundary layer on the upper surface 564 of the rear element 561 has a very short distance at which it expands from the stagnation point 572 near the leading edge of the rear element, so when encountering the final The weak shock wave 576 and the pressure recovery region 575 are still very thin, and can still maintain the attached state. With regard to pressure distribution and boundary layer development, the aft element 561 is effectively a separate airfoil in its own right, with a weak shock and pressure recovery region starting approximately at the midpoint of its chord length, for which we expect Attached flow may occur on it.

The upper surface pressure distribution of Figure 16B enables a relatively extreme example of the seam effect. The much less extreme range of intermediate pressure distributions between the single-element pressure distributions in Figure 16B and Figure 16A can also take advantage of the slit effect. The shock wave of the front element 560 does not extend all the way to the rear slot flange 565 and there is no supersonic region on the upper surface 564 of the rear element 561 . Indeed, Figure 16B shows a series of such intermediate pressure distributions when the airfoil operates at Mach numbers and lift coefficients lower than those shown. The slit effect is still needed to prevent gas flow separation under these other conditions.

The pressure distribution on the lower surface contributes to the technical level of the slotted airfoil 523 in Figure 16B. The pressure distribution 568 on the lower surface 584 in the rear end element 560 of the slotted airfoil 523 is compared to the associated pressure distribution 90 on the lower surface 91 of the single element airfoil 101 of FIG. 16A . The flatter pressure distribution on the slotted airfoil 523 results in less curvature of the lower surface of the airfoil 523 and also in the airfoil 523 where the front and rear spars of the main structural box are located. The location has a greater depth (typically 15% and 64% of the entire chord length). Both a flatter lower surface and a deeper spar wing are beneficial to the structural efficiency of the main structural box. This advantage can also be translated to increase the Mach number and lift coefficient while maintaining the same structural efficiency (flexural strength) of the wing box as that of a single element wing.

Figure 20 shows a side view of a conventional airfoil 600 designed for cruising at high subsonic and/or transonic speeds. The airfoil 600 includes a single slotted trailing edge flap 602 . In Figure 20, the flaps 602 are in a retracted position 604, eg, while cruising. In the retracted position 604 , the front end 606 of the flap 602 is nested and hidden within the contour of the airfoil 600 . In this way, the airfoil 600 constitutes a streamlined and aerodynamically smooth outer surface which may at most include some small steps or gaps.

It should be noted that the profiles of the airfoil 600 and flap 602 are schematic only. It should also be noted that conventional cruising airfoils are generally fitted with leading edge lifts, but this equipment is not shown in Figures 20-22.

In FIG. 21 , the flaps 602 are shown in a deployed position 608 , eg, in a take-off condition. FIG. 22 shows the flap 602 in another deployed position 610 , but at a greater deflection than that shown in FIG. 21 . The deployed position 610 shown in FIG. 22 may be used, for example, in a landing condition.

To deploy the flap 602 from the retracted position 604 (FIG. 20) to either of the deployed positions 608 (FIG. 21) or 610 (FIG. 22), the flap 602 is moved rearwardly. Moving the flap 602 rearwardly to deploy the flap 602 simultaneously opens a cavity 612, commonly referred to as a "cavity". As shown in FIGS. 21 and 22 , the void 612 is bluff and includes a steeply emerging lower edge 614 in the rear end 616 of the main or forward airfoil element 618 .

In conventional trailing edge flap systems having more than one slot (eg, double slotted trailing edge flaps, etc.), more than one bluff indentation is typically opened when the flap system is deployed.

Since the presence of the bluff indentation does not significantly affect the lifting performance, there is little need to design the lifting slots in an aerodynamically superior way. In cruising, however, it has been observed that the presence of the bluff indentation preceding the seam causes a significant and sometimes unacceptable drag loss. For a given conventional raised flap shape and flap-indented area set by flap deployment, the raising slot is often closed in cruise flight to avoid flap-indented resistance consumption.

As shown in Figures 6, 16B and 17, embodiments of the present invention include airfoils with one or more slots defined to have a smooth streamlined profile and without bluff indentations. These slots include well-flowed, aerodynamically smooth channels. Eliminating the bluff indentation and setting the slot to a well streamlined, aerodynamically smooth channel also allows the slot to open at cruise and other transonic state to improve performance.

In addition to providing the airfoil with cruise slots as described above, the overall shape or profile of the airfoil may also be tailored to take advantage of the slot effect (see above for a description of the slot effect). An exemplary difference in airfoil shape between the slotted airfoil 523 and the conventional unslotted airfoil 101 can be seen by comparing FIGS. 16A and 16B . For example, the upper surface 583 of the slotted airfoil 523 is generally flatter than the upper surface 93 of the conventional airfoil 101 , although there are other subtle differences between the two airfoil shapes.

Deploying cruise flaps (flaps with at least one cruise slot set) requires the flap to move aft a smaller distance than deploying a conventional single-slotted raising flap. For example, as shown in FIGS. 20-22, the conventional single-slotted raised flap 602 needs to be moved rearwardly a relatively large distance to open the indentation 612 to a sufficient width so as not to impede airflow through the slot. 620. On the other hand, although the overlap between the cruise flaps and the main airfoil elements is preferably short, embodiments of the invention use cruise slots, even if the cruise flaps provided with the cruise slots When fully retracted, it remains substantially open. Because large rearward displacements of the cruise flaps between the cruise and raised positions tend to open the cruise slot too far and prevent improved lift performance, at least in some embodiments the cruise flaps are Rearward movement when deployed between the cruise and raised positions is preferably minimized.

Trailing edge raising systems can be integrated with slotted airfoils in various ways.

For the parts along the spanwise direction of the wing that do not have cruise slits, the trailing edge raising system does not need to set cruise slits. Thus, any of the various conventional raised flaps can be used as the wing section without the cruise slot.

A number of schemes can be used along the portion of the wing span with one or more cruise slots. For example, at least one embodiment has a flap with at least one cruise slot defined, and by increasing the angle of deflection of the flap, it can also be used as a single slot elevating flap. Whether setting the cruising slot or using it as a single slot raising flap, the same wing and flap profile is exposed to the airflow, only the angle of deflection of the flap is different.

Airfoils in some embodiments have at least one cruise slot, and at least one conventional rising slot, preferably these slots are located upstream of said cruise slot. In these embodiments, the cruise seam may also act as a raised seam.

In a preferred embodiment of a partially spanned slotted wing, cruise slots are provided only along the outer portion of the wing, for example extending in the spanwise direction between the point of discontinuity in the planar structure and the wingtip. part. In addition to the cruising seam, the outer part may also have a trailing edge raising system. Said cruise slot may only be used as a raised slot on the outside of the wing, or said cruise slot may be used as a raised slot in conjunction with one or more other conventional raised slots defined by said wing outboard portion. Together.

In at least one preferred embodiment of a partially spanned slotted wing, the cruise slot is not defined by an inboard portion, such as a spanwise extending portion between said wing root and a point of discontinuity in the planar structure. Instead, the inboard portion includes a conventional trailing edge raising system that, when deployed, opens up one or more raised slots and one or more bluff indentations upstream of the raised slots. However, in cruise flight, the trailing edge lift system is preferably retracted, thereby closing the lift slot and eliminating the bluff indentation.

Although various preferred embodiments have been described, modifications or variations can be made by those skilled in the art without departing from the concept of the present invention. The above examples are only used to illustrate the invention and not to limit it. Accordingly, the description and claims should be construed freely on the basis of the relevant prior art.

Claims (53)

1. A swept airfoil comprising: at least one front end airfoil element having an upper surface and a lower surface; at least one aft airfoil element having an upper surface and a lower surface; and at least one slot defined by said airfoil in at least one transonic regime of said airfoil, said slot separating a portion of the air flowing along the lower surface of said front end airfoil element and allowing it to flow in said Flow over the upper surface of the aft airfoil element to improve performance in transonic conditions, with a portion of the partial span slot adjacent the trailing edge of the main wing. 2. An aircraft wing comprising the swept airfoil according to claim 1. 3. The airfoil of claim 2, wherein said slots comprise aerodynamically smooth passages without bluff indentations defined between said leading and trailing airfoil elements. 4. The airfoil of claim 2 wherein said slots are provided to improve performance of said airfoil by criteria selected from one or more of the group consisting of: Increased cruising speed; increase lift; Increasing the thickness; Reduced sweep angle; reduce drag; or Combinations of the above. 5. The wing of claim 2, wherein the slot extends spanwise along only a portion of the wing span. 6. The airfoil according to claim 5, wherein the slot extends spanwise from approximately a point of discontinuity in the planar structure of the airfoil to approximately the tip of the airfoil. 7. The airfoil of claim 2, wherein the slots extend over portions of the airfoil where flow separation occurs in transonic conditions and results in increased drag. 8. The airfoil of claim 2, wherein the slot extends substantially continuously spanwise from root to tip of the airfoil. 9. The airfoil of claim 2, wherein the slots are configured to push shock waves generated by supersonic airflow along the airfoil to a rearward position on the airfoil. 10. The airfoil of claim 2, wherein the slots are configured to increase the drag-divergence Mach number magnitude of the airfoil while at least maintaining comparable aerodynamic efficiency of the airfoil. 11. The airfoil of claim 2, wherein the slots are configured to mitigate shock waves and allow the airfoil to provide higher cruising speeds. 12. The airfoil of claim 2, further comprising actuator structure coupled to said front and rear airfoil elements for causing one of said front and rear airfoil elements to move relative to the other. One moves, thereby adjusting the seam. 13. An airfoil according to claim 12, wherein said actuator arrangement is arranged to allow adjustment of said slot by one or more operations selected from a group of operations comprising: adjusting a gap separating said leading and trailing airfoil elements, said gap defining said seam; adjusting the relative height between the front and rear airfoil elements; and Adjust the angle between the front and rear airfoil elements. 14. The airfoil of claim 12, wherein said slot comprises a plurality of segments arranged longitudinally along said airfoil, each segment being independently adjustable by said actuator arrangement, whereby The slots can be adjusted differently at different positions of the span. 15. The airfoil of claim 2 , further comprising actuator structure coupled to said front and rear airfoil elements for moving one of said front and rear airfoil elements relative to the other Movement is generated to close the slot in at least one subsonic regime and to open the slot in the transonic regime. 16. The airfoil of claim 2, wherein at least one slot comprises a plurality of slots arranged longitudinally of the airfoil. 17. The airfoil of claim 2, further comprising at least one slotted wing region disposed between two unslotted wing regions. 18. The airfoil of claim 2, further comprising at least one unslotted airfoil region disposed between two slotted airfoil regions. 19. The airfoil of claim 2, wherein the transonic state is one or more states selected from the group consisting of a cruise state and a steering state. 20. The airfoil of claim 2, wherein: said front end airfoil element comprises a main airfoil portion; the rear airfoil element comprises a flap; and The wing also includes actuator structure for adjusting the flap in cruise, thereby improving the performance of the wing in cruise. 21. An aircraft comprising the airfoil of claim 1. 22. An airfoil comprising: at least one front end airfoil element having an upper surface and a lower surface; at least one aft airfoil element having an upper surface and a lower surface; and at least one partial span slot defined by said airfoil in at least one transonic regime of said airfoil, said slot separating a portion of the air flowing along the lower surface of said front end airfoil element and allowing it flow over the upper surface of the aft airfoil element resulting in improved performance in transonic conditions. 23. An aircraft wing comprising the swept airfoil of claim 21. 24. The airfoil of claim 23, wherein said slots comprise aerodynamically smooth passages without bluff indentations defined between said leading and trailing airfoil elements. 25. The airfoil of claim 23, wherein said slots are provided to improve performance of said airfoil by criteria selected from one or more of the group consisting of: Increased cruising speed; increase lift; Increasing the thickness; Reduced sweep angle; reduce drag; or Combinations of the above. 26. The airfoil of claim 23, wherein the slot extends spanwise from approximately a point of discontinuity in the planar structure of the airfoil to approximately the tip of the airfoil. 27. The airfoil of claim 23, wherein the slots extend over portions of the airfoil where flow separation occurs in transonic conditions and results in increased drag. 28. The airfoil of claim 23, wherein the slots are configured to push shock waves generated by supersonic airflow along the airfoil to a rearward position on the airfoil. 29. The airfoil of claim 23, wherein the slots are configured to increase the drag-divergence Mach number magnitude of the airfoil while at least maintaining comparable aerodynamic efficiency of the airfoil. 30. The airfoil of claim 23, wherein the slots are configured to mitigate shock waves and allow the airfoil to provide higher cruising speeds. 31. The airfoil of claim 23, further comprising actuator structure coupled to said front and rear airfoil elements for causing one of said front and rear airfoil elements to move relative to the other. One moves, thereby adjusting the seam. 32. An airfoil according to claim 31 , wherein said actuator arrangement is arranged to allow adjustment of said slot by one or more operations selected from a group of operations comprising: adjusting a gap separating said leading and trailing airfoil elements, said gap defining said seam; adjusting the relative height between the front and rear airfoil elements; and Adjust the angle between the front and rear airfoil elements. 33. The airfoil of claim 31 , wherein said slot comprises a plurality of segments arranged longitudinally of said airfoil, each segment being independently adjustable by said actuator arrangement, whereby The slots can be adjusted differently at different positions of the span. 34. The airfoil of claim 23 , further comprising actuator structure coupled to the front and rear airfoil elements for causing one of the front and rear airfoil elements to move relative to the other. Movement is generated to close the slot in at least one subsonic regime and to open the slot in the transonic regime. 35. The airfoil of claim 23, wherein at least one part-span slot comprises a plurality of slots arranged longitudinally of the airfoil. 36. The airfoil of claim 23, further comprising at least one slotted wing region disposed between two unslotted wing regions. 37. The airfoil of claim 23, further comprising at least one unslotted wing region disposed between two slotted wing regions. 38. The airfoil of claim 23, wherein the transonic state is one or more states selected from the group consisting of a cruise state and a steering state. 39. The airfoil of claim 23, wherein: said front end airfoil element comprises a main airfoil portion; the rear airfoil element comprises a flap; and The wing also includes actuator structure for adjusting the flap in cruise, thereby improving the performance of the wing in cruise. 40. The airfoil of claim 23, wherein the airfoil is swept back. 41. An aircraft comprising a wing according to claim 40. 42. An aircraft comprising an airfoil according to claim 22. 43. A method of flying a swept aircraft wing comprising the use of at least one slot defined by said wing which in at least one transonic regime of said wing causes a portion of flow along said wing lower surface The diversion of the air, which separates and flows over the upper surface of the airfoil, at least delays separation of the airflow which would lead to increased drag in the transonic regime, thereby improving performance in the transonic regime. 44. The method of claim 43, further comprising adjustment of the slot in transonic conditions. 45. The method of claim 43, wherein adjustments to the seam include at least one or more operations selected from a group of operations comprising: adjusting a gap separating said leading and trailing airfoil elements, said gap defining said slot; adjusting the relative height between the front and rear airfoil elements; and Adjust the angle between the front and rear airfoil elements. 46. The method of claim 43, wherein: said front end airfoil element comprises a main airfoil portion; said rear airfoil element includes a flap arrangement; and Adjustment of the slot includes use of the flap arrangement. 47. The method of claim 43, wherein the slot comprises a partial span slot. 48. The method of claim 43, wherein the slot comprises substantially the entire length of the span of the wing from about the root of the wing to about the tip of the wing A single seam extending at 49. The method of claim 43, further comprising opening the slot while at or near a transonic regime. 50. The method of claim 43, further comprising closing the slot in at least one subsonic regime of the airfoil. 51. The method of claim 43, wherein said slot comprises an aerodynamically smooth channel without bluff indentation defined between said leading and trailing airfoil elements. 52. A method of flying an aircraft wing with a main wing section, flap means and at least one slot set between said main wing section and said flap means when cruising, said The method includes actuating the flap arrangement and adjusting the flap arrangement while cruising, thereby obtaining improved performance while cruising. 53. The method of claim 52, wherein the slot comprises an aerodynamically smooth passage without bluff indentation defined between the leading and trailing airfoil elements.

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