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 defined by the wing during at least one transonic condition of the wing. The slot (12) may either extend spanwise along only a portion of the wingspan, or it may extend spanwise along the entire wingspan. 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 split and flow over the upper surface (20) of the trailing airfoil element (38) so as to achieve a performance improvement in the transonic condition.

Description

The aircraft wing of cracking

The cross reference of related application

This application requires U.S. Provisional Patent Application No.60/417,355 preceence.U.S. Provisional Patent Application No.60/417,355 are filed on October 9th, 2002, and its content intact is incorporated herein by reference.

The invention source

Invention part described herein is that the employee by United States Government finishes, and this invention can be under the situation of any patent royalties of nonpayment, by/be purpose manufacturing and the use of United States Government with government.

Technical field

In short, the present invention relates to aircraft, the method that relates in particular to the aircraft wing of cracking and improve the aircraft cruise performance.

Background technology

The wing of many aircraft all uses conventional aerofoil profile to design.The upper and lower surface of conventional aerofoil profile is converged formation leading edge blunt nosed or circle, and sharp-pointed trailing edge.

Conventional aerofoil profile also is used to transonic wing (that is, being the designed wing of transonic flight).When carry-on air-flow velocity was the mixing of subsonic flow (for example, less than the flow velocity degree of the velocity of sound) and supersonic flow (for example, greater than the flow velocity degree of the velocity of sound), aircraft will carry out transonic flight.The upper surface of the airfoil flow air is owing to the upper surface rate of curving is accelerated to produce lift.The speed of the aircraft when as a result, a part of air-flow aboard reaches speed of sound (for example, reaching the velocity of sound) may be less than a Mach number.

In brief, Mach number is exactly the ratio of the flying speed and the current height of the aircraft place speed of sound of aircraft.When flying with the speed of sound, aircraft can reach one Mach.Critical Mach number (M _Crit) be exactly the Mach number of air-flow along aircraft aircraft flight speed when reaching speed of sound somewhere

When the air-flow of arbitrary portion on the aircraft reaches speed of sound really, will produce shock wave herein.If the Mach number of aircraft has surpassed critical Mach number, the upper and lower surface of aerofoil profile all can produce supersonic airstream so, thereby causes producing shock wave on whole aerofoil profile.When transonic flight, usually can there be a plurality of local hyprsonic zone of demarcating by shock wave.

Cross shock wave, the pressure of air and density all increase greatly, thereby cause anisentropic or expendable loss, are classified as wave resistance.When the Mach number of aircraft increased, resistance can increase significantly suddenly, and the near-sonic drag of promptly being known as increases.The shock wave air-flow that can slow down, and therefore pressure boost cause crossing shock wave and adverse pressure gradient occurs.The intensity that depends on shock wave, described adverse pressure gradient can cause air-flow to occur semistall at place, shock wave bottom with aerofoil surface.In the transonic flight process, shock wave and always be the chief component of the whole resistance of aircraft by the separation of boundary layer that shock wave causes.

Mach number when near-sonic drag begins significantly to increase is called as " resistance divergence Mach number " (M _Dd).Will cause the obvious increase of aircraft resistance because the Mach number of aircraft surpasses resistance divergence Mach number slightly, usually be unpractical so operate under this condition economically.

Adopted several different methods, be used for near-sonic drag is brought up to higher Mach number, thus the wave resistance of minimizing on given transonic speed.Some more common methods comprise uses the wing of expensive big swept wing, thin airfoil and rearward projection (aft-camber) of manufacturing cost.Supercritical airfoil produces higher critical Mach number.The high projection rear section that supercritical airfoil generally all has the upper surface of the flattening that can reduce the air-flow acceleration/accel and most of lift can be provided.The wing of rear load moves behind the center with lift, thereby causes big nose-down pitching moment.The increase of nose-down pitching moment finally all need wing and tailplane double the running with the aloft aircraft of balance.The resistance relevant with the balance aircraft is flat (trim) resistance of assignment.Bigger nose-down pitching moment generally can increase the trim resistance.

Except considering the aerodynamics factor, other factors also can limit the thin degree of actual aerofoil profile.For example, the fuel capacity that provides of thin wing is less.And, because thin wing has more shallow structure case, so use thin wing usually can increase the overall weight of wing.

Use big wing also can increase resistance divergence Mach number, thereby reduce the wave resistance under given transonic flight speed.For bigger wing area, need to use the aerofoil profile that has than the low lift-to-drag ratios coefficient, thereby also can cause less wave resistance.But, wetting (wetted) area of the increase of big wing can make the friction drag of aerofoil surface be increased to a certain degree usually, causes the additional surfaces friction drag can offset or surpass the minimizing of any wave resistance.

Name is called the US Patent 6 of " have non-sweepback crack the aircraft of Airfoil of cruising ", 293,497 have disclosed a kind of unswept or unswept basically wing, this kind wing uses the aerofoil profile technology of cruising of cracking, compare with the swept-back aircraft wing of not cracking, have higher cruising speed, when low-speed operations, also can obtain bigger lift simultaneously.US Patent 6,293,497 full content is incorporated in here as a reference, discusses fully.

Summary of the invention

Aircraft wing comprises at least one front end foil element and at least one rear end foil element.During at least one transonic state of described wing, described wing should have at least one seam.Described seam can only extend to spanwise along the part of the span, also can extend to spanwise along the whole span.In either event, described seam can make the part air separation that flows along the lower surface of described front end foil element, and it is flowed on the upper surface of described backend machine wing element, thereby has improved the performance of transonic state.In the exemplary embodiment, described wing comprises part span seam, and the preferably beginning and extend out to wing tip in the middle of about span of described seam can make edge effect weaken greatly at least or surpasses described seam effect.

Another kind of form of the present invention provides the certain methods of the aircraft wing that is used to fly.In one embodiment, a kind of method is included in substantially under at least one transonic state the seam that is set between front end foil element and the rear end foil element is adjusted, thereby the performance of transonic state is improved.

In another embodiment, a kind of method comprises substantially uses a seam, makes at least one transonic state of aircraft wing along the mobile part air diverts of wing lower surface so that its separation and mobile at upper surface of the airfoil.Described air be diverted to that I haven't seen you for ages and postpone the burbling that causes resistance to increase, thereby under the transonic state, improve performance.

In a further embodiment, when being included in cruising condition substantially, a kind of method drives the wing flap device adjusting described wing flap device, thus the performance when improving cruising condition.

By the detailed description of this paper back, can obviously draw the darker field that the present invention uses.It should be understood that detailed explanation and concrete example at least one exemplary embodiment of explanation the present invention, its purpose only is in order to illustrate, rather than limits the scope of the invention.

Description of drawings

Can more completely understand the present invention by the detailed description and the accompanying drawings, wherein:

Fig. 1 is the described according to one embodiment of present invention birds-eye view that comprises the swept wing of part span seam;

Fig. 2 is the described according to another embodiment of the present invention birds-eye view that comprises the swept wing of extreme span seam;

Fig. 3 is the conventional not birds-eye view of slotted aerofoil, there is shown shock-wave spot and supersonic airstream zone when medium cruising speed lift coefficient and Mach number;

Fig. 4 is the birds-eye view of part span slotted aerofoil shown in Figure 1, there is shown shock-wave spot and supersonic airstream zone when medium cruising speed lift coefficient and Mach number;

Fig. 5 is the birds-eye view of extreme span slotted aerofoil shown in Figure 2, there is shown the shock-wave spot when medium cruising speed lift coefficient and Mach number and the zone of supersonic airstream;

Fig. 6 is the side cross section view of wing shown in Figure 1, there is shown when being used for cruising flight according to an embodiment of the present invention front-end and back-end aerofoil profile in the slotted aerofoil zone of planar structure catastrophe point;

Fig. 7 is the side cross section view of wing shown in Figure 1, there is shown the wing section of described according to one embodiment of present invention non-slotted aerofoil zone at root and plane structural mutation point place;

Fig. 8 shows the aerofoil profile that as shown in Figure 6 front-end and back-end aerofoil profile is overlapped in planar structure catastrophe point shown in Figure 7;

Fig. 9 has summed up the wind tunnel test Line Chart of gained as a result, and one is the model that wing, fuselage and the vertical tail of part span seam are housed in the employed model of wind tunnel test, and another is to have the conventional transonic speed model of wing, fuselage and vertical tail;

Figure 10 has summed up the wind tunnel test Line Chart of gained as a result, and one is the dummy vehicle that the wing of part span seam is housed in the employed model of wind tunnel test, and another is the dummy vehicle that conventional transonic wing is housed;

Figure 11 is the described according to another embodiment of the present invention birds-eye view that comprises the slotted aerofoil of wing tip equipment;

Figure 12 is the simple block diagram of the ACTIVE CONTROL system relevant with the operation of the slotted aerofoil of adjusting and adjust described seam;

Figure 13 is the described according to another embodiment of the present invention birds-eye view that has the wing of two part span seams;

Figure 14 is the described according to another embodiment of the present invention birds-eye view that has the wing of two part span seams;

The birds-eye view of Figure 15 slotted aerofoil, wherein said seam comprise the adjustable main plot of a plurality of independences section;

Figure 16 A shows the distribution of pressure of the conventional non-aerofoil profile of cracking;

Figure 16 B shows the distribution of pressure of the aerofoil profile of cracking;

Figure 17 is the crack air-flow of Airfoil Design or the computational fluid dynamics of pressure field (CFD) model sample of two dimension;

Figure 18 A is the transparent view according to the finite element model of the described part span of at least one embodiment of the present invention slotted aerofoil;

Figure 18 B is the more detailed transparent view of the flap carriage shown in Figure 18 A;

To be that at least one embodiment according to the present invention is described have and do not have the lip-deep airflow field of lower wing of part span slotted aerofoil of flap carriage or the three-dimensional CFD model sample of pressure field respectively for Figure 19 A and 19B;

Figure 20 illustrates the lateral plan that the quilt of the aerofoil profile that has the trailing edge flap that singly cracks is packed up;

Figure 21 is the lateral plan of aerofoil profile as shown in figure 20, and the trailing edge flap that its single fluting is shown is partly launched;

Figure 22 is the lateral plan of aerofoil profile as shown in figure 20, but the deviation angle that its trailing edge flap that singly cracks is unfolded than the angle of Figure 21 more greatly.

The specific embodiment

The explanation that different embodiments of the invention are carried out only is the exemplary description to its essence below, is not the present invention, its application or purposes are limited.For example, each embodiment of the present invention (for example is supposed to be widely applied on the various aircraft, especially be not limited to fighter plane, commercial machine, private machine, hyprsonic impact type aircraft etc.) and need not consider the mode that aircraft drives (for example, especially direct-type, distance type, autocontrol or convolution etc.).Therefore, the specific herein aircraft of quoting should not be construed and limits the scope of the invention.And, each embodiment of the present invention also is supposed to be widely applied to aircraft and produces on the surface of lift (for example, especially be not limited to fixed wing, variable-geometry wing, rotor blade, right semispan wing, left semispan wing, extreme span wing, straight wing, swept wing, delta wing, tailplane, conical wing, non-conical wing, tiltedly the wing etc.).Therefore, the specific herein wing of quoting should not be construed and limits the scope of the invention.

Therefore and be not used in the restriction of carrying out scope and employed particular term also only is the purpose in order to refer in the following description.For example, the direction during term " top ", " bottom ", " in the above " and " below " refer to reference to the accompanying drawings.The direction of some part of element has been described at term " preceding ", " back ", " back " and " side ", wherein these quote be self-consistentency but be not limited thereto, can understand with reference to text and the relevant drawings of describing the element of discussing.The word of word, derivatives and the similar implication of mentioning specially above this class term can comprise.Similarly, sequence or order do not represented in " first ", " second " and other numbers that structure refers to, unless in context, clearly be illustrated.

Fig. 1 shows described according to one embodiment of present invention aircraft swept wing 10.As shown in the figure, described swept wing 10 comprises front end foil element 36 and rear end foil element 38.In at least one transonic state of described wing 10, set at least one part span seam 12 between described front end foil element 36 and the described rear end foil element 38.

Fig. 2 shows another embodiment of swept wing 110.As shown in the figure, described swept wing 110 comprises front end foil element 136 and rear end foil element 138.In at least one transonic state of described wing 110, set at least one extreme span seam 112 between described front end foil element 136 and the described rear end foil element 138.

Described part span seam 12 and described extreme span seam 112 make the part air separation that flows along the lower surface of anterior member 36,136 and flow on the upper surface 20,120 of described tail end element 38,138, thereby wing are operated in enter or increase section or improve its performance in the one or more stages near the high speed buffet boundary near the near-sonic drag of wing; Wherein transonic cruising condition and transonic control state are exactly the example in stage in this.At least in certain embodiments, described part span seam 12 and extreme span seam 112 all comprise smooth on the Pneu that is set between the foil element of described front-end and back-end, as not have (blunt form indent) indent passage, and be as described below.

Here employed " part span seam " relates to and comprises one or more seams, and each seam all only extends to spanwise along the part of the wing span.That is to say that described part span slotted aerofoil does not have the single seam that extends to wing tip from wing root fully.In the exemplary embodiment, the seam of described part span slotted aerofoil will make edge effect weaken greatly at least or surpass and stitch effect preferably from probably or slightly beginning to extend out to then wing tip more inwards from span middle part.Fig. 1 shows the exemplary wing 10 that has part span seam 12.

Here employed " extreme span seam " relates to and comprises substantially from the wing root near-end and beginning constantly to extend substantially to the seam of wing tip (will the seam effect be descended to edge effect at least), do not comprise connecting the necessary bracing frame of wing structure element that is positioned at before the described extreme span seam with afterwards.This bracing frame generally can have influence on the inlet of the extreme span seam on the wing lower surface, but can not have influence on the outlet of the extreme span seam of upper surface of the airfoil.Fig. 2 shows the exemplary extreme span seam 112 that extends to wing tip 116 from wing root 114.

Here employed " transonic cruising condition " relates to and comprises the relative high speed stage of wing, and the air-flow of process wing comprises the supersonic airstream regional area as shown in the figure, for example, and Fig. 3,4 and 5.In other words, enter or near the wing near-sonic drag increase section or near the wing of high speed buffet boundary with relative high-performance cruise.And employed here " transonic state " relates to and comprises one or more mission phases, and wing flies therein, but and unnecessary when carrying out cruising flight, enter or increase section or near the high speed buffet boundary near the near-sonic drag of wing.The exemplary transonic state of described wing comprises, but is not limited to transonic cruising condition and transonic control.

Fig. 1 and Fig. 2 are the simplification planar structures that is currently applied to the starboard wing design plan on the business flying device, and the design plan among above-mentioned two figure is separately installed with part span seam and extreme span seam.Described business flying device also comprises a port wing that has basic identical performance in the flight curve.Therefore, when described starboard wing was provided with seam, described port wing (not shown) usually also was provided with equivalent part or corresponding seam.

About the term (for example, starboard wing and port wing) of semispan wing, 0% semispan position generally is known as starboard wing and port wing symmetry or becomes the position of mirror image.Generally speaking, 0% semispan position is the middle part of the accompanying fuselage of wing.When the semispan wing was discussed, term " semispan " referred to the distance from 0% semispan position to the 100% semispan position that is positioned at wing tip.But, it should be noted that embodiments of the invention should not be confined in the semispan wing, equally also can be applicable on the extreme span wing (wing for example, especially flies).And as shown in figure 11, employed here term " span " and " semispan " do not comprise the one or more wing tip equipment that can install or be arranged at wing tip.But, this should not be used to limit the scope of the invention, and embodiments of the invention are supposed to be widely applied on the various wings, including, but not limited to, the wing that has the wing of wing tip equipment and do not have wing tip equipment.In other embodiments, described wing tip equipment can be arranged at least a portion of part span seam or extreme span seam really.

Further with reference to figure 1, described part span seam 12 can extend to spanwise along occurring the part that burbling causes resistance to increase in the semispan of described wing 10 when the transonic state of described wing 10.The position that 12 meetings that can be placed on the hydrokinetics calculation simulation suggestion of the three-dimensional air-flow on the described wing 10 cause the burbling pressure field on the upper surface of the airfoil 20 is stitched in the part span.

In the exemplary embodiment, described part span seam 12 extends from general semispan position 28 to general semispan position 30.Described semispan position 28 and 30 is consistent with Ye Hudi (Yehudi) point or planar structure catastrophe point 32 and wing tip 16 respectively, though actual conditions do not need so.In other embodiments, described part span seam 12 can be from other inner sides, and the inclusion that stitches described in these positions can not hinder the low speed control surface or disturb other elements of for example fuel tank and take-off and landing device and that the planar structure of described wing 10 is carried out is integrated.And described part span seam does not need to extend to fully described wing tip.On the contrary, described part span seam can extend to described wing tip basically, but stops when edge effect weakens to some extent to the performance of being improved by described seam.

The specific chordwise location of described part span seam 12 and extreme span seam 112 (Fig. 2) determined by following consideration to small part probably, for example specific low speed control surface and for example other elements of fuel tank and take-off and landing device and described wingpiston structure carry out integrated.In one exemplary embodiment, the chordwise location of each seam 12 and 112 is positioned at general 90% place of general 70%-of wing chord.

In use, each seam 12 and 112 can make the part air separation that flows along the lower surface 18 of described front end foil element 36,136, and it is flowed on the upper surface 20,120 of described rear end foil element 38,138.In this case, describedly be stitched to few separation of boundary layer and will further shift the wing rear portion onto of having postponed by the shock wave that supersonic airstream produced.The existence of seam can be found out by comparison diagram 3 (not slotted aerofoil), 4 (part span slotted aerofoil) and 5 (extreme span slotted aerofoils) for the supersonic airstream (using area B to represent) on whole top wing surface and the effect (seam effect) of shock-wave spot (using solid line A to represent).As described below, this " seam effect " improved the performance of wing in the transonic state.

" seam effect " prevent or the mode that postpones separation of boundary layer at least and adopted as following description, and in the US Patent 6,293,497 of " have non-sweepback crack the aircraft of Airfoil of cruising " by name, have been described in detail.Here intactly quoted US Patent 6,293,497 content is also here discussed fully.

With further reference to Fig. 1, described part span slotted aerofoil 10 comprises that at least one wing of not setting seam zone 22 and at least one set another zone 24 of at least one part span seam 12.Identification and explanation for convenience, and be not to limit, described wing zone 22 also can be used in reference to generation slotted aerofoil zone 22 not, this is because described not stitched open region 22 unqualified seams, described wing zone 24 also can be used in reference to for described slotted aerofoil zone 24, and this is because described stitched open region 24 defines a part span seam 12 at least.But, it should be noted, in the described wing zone 22 and 24 any one or two seams that can have any amount (for example one or more), wherein some only can be arranged on the position of lift apparatus such as not disposing for example leading edge slat, aileron, wing flap, spoiler and/or stable and control convenience.

Therefore as shown in the figure, along the spanwise setting, described slotted aerofoil zone 24 is provided with between semispan position 28 and 30 between semispan position 26 and 28 in described not slotted aerofoil zone 22.Described semispan position 26,28 and 30 is consistent with wing root 14, planar structure catastrophe point 32 and wing tip 16 respectively, but this is not essential.

Described slotted aerofoil zone 24 reaches on the wing zone of critical Mach number when being arranged at the flight of higher cruising speed only.To illustrate below and determine which part wing can reach the method for critical Mach number when cruising.Other zones that Mach number can not reach critical value on the wing can comprise non-slotted aerofoil zone 22.

In this example, described not slotted aerofoil zone 22 is arranged near the inboard (for example, the fuselage) of planar structure catastrophe point 32 as shown in the figure.For the commercial aircraft of needs withdrawal take-off and landing device, the general chord length of length of using of its wing medial region.For relatively long chord length, the corresponding wave resistance of described inside part usually is minimum, and this is because described aerofoil profile is compared with the wing of whole commercial aircraft and had relatively low local lift coefficient (C ₁).If described inside part does not reach critical Mach number when cruising, so just do not need to use described part span seam 12 to increase the size of Mach number.Therefore, described not slotted aerofoil zone 22 can be arranged at the inboard wing part that reaches critical Mach number at the Shi Buhui that cruises, and does not need to increase the aerofoil resistance increase that the Mach number position uses part span seam to be brought thereby can avoid or eliminate when cruising.And, use slotted aerofoil zone 22 not can make conventional elevation system (for example, conventional wing flap and stripe board) can be applied to the inside part of described wing 10 at described inside part, this also is the additional advantage among each embodiment of the present invention.And, it should be noted that each embodiment of the present invention should not be limited as has the wing that does not reach the inside part of critical Mach number when cruising.Each embodiment of the present invention is supposed to be widely used on the various wings really, comprise, but be not limited to, have the wing of the medial region that when cruising, reaches critical Mach number and have the wing that when cruising, does not reach the medial region of critical Mach number.

Though described part span slotted aerofoil 10 has single not slotted aerofoil zone 22 and single slotted aerofoil zone 24 as shown in the figure, this is not essential.Described part span slotted aerofoil 10 (for example can have any amount, one or more) not slotted aerofoil zone 22 and any amount of slotted aerofoil zone 24, each slotted aerofoil zone can have any amount of seam and not deviate from the spirit and scope of the present invention.Along with wing design-calculated particular demands surface is used the not slotted aerofoil zone that surpasses possibly and/or surpassed one slotted aerofoil zone, crack so and semispan that conversion between the slotted aerofoil zone 22 and 24 will not crossed over wing occurs repeatedly.For example, another embodiment of described part span slotted aerofoil comprises inboard not slotted aerofoil zone, middle part slotted aerofoil zone and another not slotted aerofoil zone between wing tip and described slotted aerofoil zone.

Fig. 6 shows front-end and back- end foil element 36 and 38 aerofoil profiles at planar structure catastrophe point 32 places in the described wing 10.Described front end foil element 36 comprises upper surface 40, lower surface 42, leading edge 44 and trailing edge 46.Similarly, described rear end foil element 38 also comprises upper surface 48, lower surface 50, leading edge 52 and trailing edge 54.Between described part span seam 12 leading edges 46 and the trailing edge 52 in the described rear end foil element 38 that are set in the described front end foil element 36.The section of described part span seam 12 as shown is gap or the space of the leading edge in the described front end foil element 36 from trailing edge 52 separation of described rear end foil element 38.In flight course, described part span seam 12 makes the part air separation that flows along the lower surface in the described front end foil element 36 42 and its upper surface 48 at described rear end foil element 38 is flowed.

Further with reference to figure 6, the meeting in the part front end foil element 36 is overlapping or overhang on the part rear end foil element 38.Therefore, the summation of chord length has surpassed a hundred per cent (100%) (for example, the distance between described leading edge 56 terminals and described trailing edge 34 terminals) in described slotted aerofoil zone 24 in the described front and back ends foil element 36 and 38.In at least one embodiment, but described gap is minimizedly to have enough sizes, can't mix mutually or converge with the boundary 1ayer on the upper surface 48 of described front end foil element 38 thereby make along the boundary 1ayer of the lower surface 42 of described front end foil element 36.

Fig. 7 is the cutaway view in described not slotted aerofoil zone, the airfoil section 64 shown in the figure on semispan position 26 at semispan position 28 or the airfoil section 66 in planar structure catastrophe point 32 places and described non-slotted aerofoil zone 22 coincide.Because described not slotted aerofoil zone 22 is reverse swept-backs and is trapezoidal, so leading edge 68 in the described root airfoil section 64 and trailing edge 70 can be placed at the leading edge 72 of airfoil section 66 of planar structure catastrophe point 32 and the front portion of trailing edge 74.

Fig. 8 is the sectional view in slotted aerofoil zone 24 as shown in Figure 6, it is at the front end airfoil section 36 and the rear end airfoil section 38 at described planar structure catastrophe point 32 places shown in the figure, with airfoil section 66 overlaids in the not slotted aerofoil zone 22 at planar structure catastrophe point 32 places, as shown in Figure 7.At planar structure catastrophe point 32 places, the leading edge 72 in described not slotted aerofoil zone 22 transition relatively smoothly is leading edge 56 terminals in described slotted aerofoil zone 24.At planar structure catastrophe point 32 places, the trailing edge 46 relative transition smoothly of the host wing part 58 in described slotted aerofoil zone 24 are the upper surface in the described not slotted aerofoil zone 22.Equally at described planar structure catastrophe point 32 places, thereby trailing edge 34 terminals in described slotted aerofoil zone 24 offset downward trailing edge 34 terminals that make the air that flows through part span seam 12 can pass through described slotted aerofoil zone 24 from the trailing edge 74 in described not slotted aerofoil zone 22 with appropriate amount.

Described part span seam 12 can quite suddenly begin from described planar structure catastrophe point 32.That is to say that the gap that the trailing edge 46 in the described front end foil element 36 is separated with the leading edge 52 in the described rear end foil element 38 is not bevelled, begins can little by little not increase size from described planar structure catastrophe point 32 yet.Therefore, the part span 12 begins the place in described planar structure catastrophe point 32, and carrying out the transition to described slotted aerofoil zone 24 from described not slotted aerofoil zone 22 is not level and smooth relatively.But, it should be noted, other embodiment can comprise part span seam 12 or the bevelled part span seam 12 that begins gradually so that in planar structure catastrophe point 32, have the part span stitch 12 places state not slotted aerofoil zone 22 relatively smoothly transition be described slotted aerofoil zone 24.

In at least one embodiment, the gap that is positioned at planar structure catastrophe point 32 is passed through, for example, a baffle plate (not shown), sealed.Described baffle plate can be the plane, and strides across described gap and be set up, so that described baffle plate is identical with heading.

Described part span seam 12 can be set between the rising or stable and control convenience of host wing part 58 and for example wing flap 60, aileron, spoiler etc.In the exemplary embodiment, described part span seam 12 is set between the leading edge 52 of the trailing edge 46 of described host wing part 58 and described wing flap 60.Therefore, described part span seam 12 can make the part air separation that flows along the lower surface 42 of described host wing part 58, and it is flowed on the upper surface 48 of described wing flap 60.

Described wing flap 60 can be incorporated into line operate with active control system 61 (Figure 12), the described active control system operation with actuator structure again is relevant, the wing flap actuator structure that the US Patent 5,788,190 of " trailing edge flap that cruises cracks " for example by name is disclosed.Here intactly quoted US Patent 6,293,497 content is also here discussed fully.

Described actuator structure be connected to described wing flap 60 and described host wing part 58 on, be used for moving and described host wing part 58 corresponding wing flaps 60, thereby allow the described wing flap 60 of expansion and/or adjust described seam 12 for state of flight.For example, described wing flap 60 can be used to land or the complete expanded position (not shown) of takeoff condition and the stowed position 62 relevant with cruising condition between move.Perhaps, for example, described wing flap 60 can move to dwindle or to widen described seam 12, described wing flap 60 can raise or reduce changing the relative height between described wing flap 60 and the described host wing part 58, and/or described wing flap 60 can rotate to adjust angle or the pitch angle between described wing flap 60 and the described host wing part 58.

In Fig. 2, described wing 110 comprise by the trailing edge 146 of host wing structure " case " or element 136 and described inboard wing flap and outer aileron 138,138 ' leading edge 152,152 ' between the extreme span seam 112 set.As shown in the figure, the rear edge part of described host wing element 136 overlapping or be suspended from described wing flap and aileron 138,138 ' the front portion on.

Described wing flap 138 and aileron 138 ' in any one or two can be coupled with actuator structure so that described seam 112 is adjusted under the particular flight state of described wing 110.By example as can be known, being used to adjust also, the actuator structure of the described seam 112 of trim can use US Patent 5,788, the 190 wing flap actuator structures that disclosed.

It should be noted,, use other arrangement mode also to be fine for part span seam, extreme span seam and trailing edge system (for example, wing flap, aileron, bumper/spoiler etc.).For example, another embodiment comprises the blade main system, and wherein said seam is set between described blade and the main wing flap, and described blade of while is before described seam, and described main wing flap is after described seam.

In at least some embodiment, a closable extreme span or part span seam is housed, this seam can be when flight condition needs, and (for example, comprising low-speed stages such as taking off, land, climb) closes.Close described seam and can eliminate the surface friction drag consumption that described seam brings.(for example, the transonic cruising condition) can partly or entirely open described seam when the high-speed flight state.

In other embodiments, the described part span or extreme span seam can forever be arranged on the described wing, thereby the different elements that described seam is not relied on form wing (for example, wing flap, aileron, stripe board, bumper/spoiler, other raising devices, other are stable and control convenience etc.) mode of position or shape (for example, be unfolded fully, part be unfolded, packed up).Described seam whether exist the mission phase that do not rely on aircraft (for example, land, take off, climb, aerobatics, cruise, flatly fly, quicken, deceleration etc.).For example, described seam can be as the permanent opening in the movable part of wing flap and aileron, so that described being sewn on launched and still can be opened fully during the withdrawal movable part.

Figure 13 show comprise two part spanes seam 212 and 212 ' the exemplary embodiment of swept wing 210.Described seam 212 is set between the leading edge 252 of the trailing edge 246 of described front end foil element 236 and described rear end foil element 238, the described front end foil element 236 of described seam 212 ' be set in ' trailing edge 246 ' and described rear end foil element 238 ' leading edge 252 ' between.

Figure 14 shows another embodiment of swept wing, comprising have two part spanes seam 312 and 312 '.Described seam 312 is set between the leading edge 352 of the trailing edge 346 of described front end foil element 336 and described rear end foil element 338, the described front end foil element 336 of described seam 312 ' be set in ' trailing edge 346 ' and described rear end foil element 338 ' leading edge 352 ' between.

Figure 15 shows another embodiment of swept wing 410, this swept wing comprise one have a plurality of sections 412,412 ', 412 " seam, and each section all is an independent adjustable.As shown in the figure, each seam section 412,412 ', 412 " are set in the trailing edge 452,452 of described host wing structure case 436 ', 452 " and independent movably raise or stablize and control convenience 438,438 ', between 438 " leading edge 446,446 ', 446 ".Each equipment 438,438 ', 438 " be coupled with actuator structure, for example US Patent 5,788, the 190 wing flap actuator structures of describing.Described actuator structure can move independently the equipment 438,438 relevant with described host wing part 436 ', 438 ", thereby adjust and the described seam section 412,412 of trim ' with 412 " to adapt to the particular flight state of described wing 410.

In another form, the invention provides the method for some design aircraft wing.In one embodiment, in short, a kind of method is adjusting the seam that is set between front end foil element and the rear end foil element under at least one transonic state, to finish the improvement of the performance under the transonic state.Adjust described seam and can use one or more following operations: adjust the slit of separating described front and back ends foil element, described slit defines described seam; Adjust the relative height between the foil element of front and back end; And adjust angle between the described front and back ends foil element.In the exemplary embodiment, described front and back ends foil element comprises host wing partial sum wing flap device respectively, adjusts described seam and comprises the described wing flap device of driving.At least in certain embodiments, if state of flight needs, for example subsonic state (for example take off, land, climb etc.), this method may further include closes, or reduces the width of described seam at least.

In another embodiment, the method of design aircraft wing comprises the seam that uses at least one to be set by described wing in short, and this seam makes part air separation that wing flows along described wing lower surface and flows at the upper surface of described wing at least one transonic state.Make air occur turning to and to prevent or postpone at least to occur and cause the burbling of resistance increase, thereby the performance under the transonic state is made moderate progress at the transonic state.But, it should be noted, in all mission phases, and not all need to occur air diverts.For example, if state of flight guarantee, subsonic state (for example, take off, land, climb etc.) for example, this method may further include the width of closing or reducing described seam at least.In addition, this method can be included in also that wing is in or open described seam during near the subsonic state.And this method also can comprise according to the state of flight of described wing to be adjusted described seam.

In a further embodiment, provide a kind of method that designs aircraft wing, in the method, described aircraft wing comprises that host wing part, wing flap device and at least one are set in the seam between the described host wing partial sum wing flap device when cruising.Thereby this method is generally and drives described wing flap device when cruising and improve performance in the process of cruising so that described wing flap device is adjusted.

The part that will reach critical Mach number on the wing depends in part on the distribution of planar structure, thickness distribution and the airload of described wing (span load) on spanwise at least.In order to determine that reliably which wing section can reach critical Mach number, can use to have high-precision computation model method, this model method has compressibility effect complete, non-linear form and has comprised the influence of the boundary 1ayer and the wake flow of viscosity/turbulent flow.The simplification of different stage is approximate also can be included in the computer mode, for example based on the approximate method (non-sticky/Boundary Layer Method of coupling) of boundary 1ayer and not " whole " still passed through simplification to a certain degree Na Wei-Stokes (Navier-Stokes) source code (for example, " thin layer " is approximate, has wherein ignored some and has had the viscosity condition of less influence).

Can obtain the flow characteristic that CFD analysis code based on " fluid solver " is used for determining given aerodynamic shape.Therefore, when the shape of specific wing is known, just analyzes and can determine, for example, the Mach number critical level of wing different piece or the whole aerodynamic properties of wing.The coupling non-sticky/boundary 1ayer type exemplary CFD anacom software MGAERO by Washington, the Analytical Methods of Redmond, Inc. provides.Comprise FLUENT

The exemplary CFD anacom software of Na Wei-Stokes (Navier-Stokes) type by New Hampshire, the Fluent Inc.Corporation of Lebanon provides; By California, the CFD++ that the MetacompTechnologies of Agoura provides

And by Washington, the Analytical Methods of Redmond, the NSAERO that Inc. provides.

The performance of described part span seam configuration has been carried out theoretic analysis by the research of computational fluid mechanics (CFD), and passes through wind tunnel test and verify, thereby the performance of the improvement that the transonic wing than routine designs is provided.About the CFD model method, on the Airfoil Design of cracking of two dimension, carried out two-dimentional research and analysis for many years, be prior art therefore.In Figure 17, show around the airflow field of the two-dimentional Airfoil Design 80 of cracking or the CFD answer sample of pressure field.

Owing to also CFD is not expanded, uses, be not applied on the three-dimensional slotted aerofoil, so embodiments of the invention also need to develop, optimize and use particular tool and method that slotted aerofoil is carried out detailed three-dimensional CFD design and analysis yet.And as described below, various aspects content of the present invention also needs to use wind tunnel test to check the output of CFD.

Shown in Fig. 3,4,5, when the output of CFD comprised with medium cruising speed lift coefficient and Mach number flight, shock wave model sample on the whole wing and hyprsonic flowed regional.More particularly, when Fig. 3,4,5 shows with medium cruising speed lift coefficient and Mach number flight respectively, along the shock-wave spot and the supersonic flow zone of conventional wing, part span slotted aerofoil and extreme span slotted aerofoil upper surface.

Referring now to Figure 18 A and 18B, there is shown the finite element model of part span slotted aerofoil 82.As shown in the figure, described part span slotted aerofoil 82 comprises the part span seam 84 that has flap carriage 85.In Figure 18 B, illustrate in greater detail described flap carriage 85.

In Figure 19 A, described CFD output comprises the airflow field on the lower wing surface 86 of the part span slotted aerofoil 87 that has flap carriage 88 or the model sample of pressure field.In Figure 19 B, described CFD output comprise the part span slotted aerofoil 87 that do not have flap carriage ' lower wing surface 86 ' on air-flow or the isocontour model sample of pressure.Therefore, comparison diagram 19A and 19B just can draw flap carriage and whether have the influence that brings for the lower wing surface pressure.

Use three-dimensional CFD tool and method described here, can draw described part span slotted aerofoil and compare with conventional transonic wing, it is 0.025 that cruise Mach number increases (Δ M), and pneumatic efficiency (Δ ML/D) has increased-1.0%.It should be noted that it only is for illustrative purposes that these values (for example, 0.025 and-1.0%) propose in this article, should not be construed and limit the scope of the invention.In addition, these values are to obtain by the CFD model that use has the CFD model of part span slotted aerofoil, fuselage and vertical tail and has conventional tranisonic aircraft wing model, fuselage and a vertical tail.These two models do not have tailplane, engine pod or pole.

Described three-dimensional CFD design and analysis tool and method and the result who obtains have thus carried out the transonic wind tunnel test.More particularly, different wind tunnel tests is used to draw the Mach number changing value (Δ M) that cruising condition lower part span slotted aerofoil and conventional transonic wing design are compared and drawn, thereby the relative aerodynamic performance (Δ ML/D) between determining described part span slotted aerofoil and conventional transonic wing designing, and determine to carry out the integrated and trim resistance of gondola to the integrated influence that brings of aircraft, precision and the fiduciary level that three-dimensional CFD is analyzed assessed simultaneously.

Fig. 9 and Figure 10 have summed up the result of some wind tunnel tests.Wind tunnel test model among Fig. 9 includes wing (part span slotted aerofoil or conventional transonic wing), fuselage and vertical tail, but does not comprise tailplane, engine pod or support.But, the wind tunnel test model configuration among Figure 10 is complete, comprises wing (part span slotted aerofoil or conventional transonic wing), fuselage, vertical and tailplane, engine pod and support.

The emphasis of described wind tunnel test and computational fluid dynamics research concentrates on or purpose is to study the performance of Pneu.Improvement in order to ensure aerodynamic properties is directly transferable, need carry out limitation and restriction to wing design, thereby guarantees that the improvement of aerodynamic properties can not reduce the performance of other field or part.For example, the adjustment that can improve aerodynamic properties can not increase the weight of structure.Under these constraint conditions, described part span slotted aerofoil has improved cruising speed significantly, acceptable rising and maneuvering characteristics are provided simultaneously, compare with the transonic wing design of routine on it cruises design speed, guaranteed cooresponding Pneu efficient (ML/D) and scope at least.When removing above-mentioned constraint to initial design, the improvement degree that example of the present invention brings is supposed to get a promotion.After having carried out formal cross-cutting crossing research, part span slotted aerofoil may bring more improvement to aircraft efficient.

About the performance improvement aspect of transonic state lower wing operating conditions, embodiments of the invention can reach any one or their combination in the following effect: improved wing cruising speed or critical Mach number, increased airfoil lift, wing thickness and/or kept Mach number size under less swept wing angle.Be described in more detail below the cruise physical factor of airfoil performance of restriction transonic, illustrated also how the designer weighs the raising of technical merit and the improvement of wing thickness, speed, lift or resistance or these combined factors described below simultaneously.

Described seam can be used to increase the resistance-divergence Mach number (M of the wing of given sweepback angle, lift coefficient and a thickness distribution _Dd) size, and improve or when cruising flight, keep at least the cooresponding Pneu efficient (ML/D) and the scope of wing simultaneously.Pneu efficient is nondimensional performance metric method, multiply by lift by Mach number and calculates divided by resistance, is even more important for the long-distance flight device.Having the wing of at least one seam that can improve cruise performance can be before near-sonic drag increases beginning fly under with higher cruising speed.

The aerofoil profile that described seam prevents or the ability that postpones boundary 1ayer or burbling at least can be used in described wing designed to be able to the distribution that produces pressure when the transonic state, compare with conventional aerofoil profile, the suction level of upper surface (for example descends to some extent, the negative pressure coefficient that upper surface is lower), shock wave and pressure recovery subsequently move backward.The pressure that obtains owing to the existence of described seam scatters higher resistance-divergence Mach number (M is provided _Dd), because the conventional aerofoil profile of not cracking separation of boundary layer can occur under the transonic state, so can't reach this resistance-divergence Mach number.

The present invention can not become on the part wing of critical value at Mach number yet, if any, uses wing conventional or that do not crack.For example, can not reach critical Mach number in the process of cruising if determine the inside part of wing, so for inside part, slotted aerofoil is not regional just can avoid or elimination and the described surface friction drag consumption that is sewed with the pass by using.And for described wing inside part, use conventional and the not zone of slotted aerofoil allows to use the elevation system (for example, conventional wing flap and slat) that is applied to inboard routine too.

Though for the aircraft that uses slotted aerofoil, fuel oil consumption is substantially the same, cruising speed that aircraft increased or Mach number size can increase other efficient really.For example, slotted aerofoil can improve the speed that aircraft cruises or flies before near-sonic drag increases beginning, thereby had reduced the time of journey.Except bring very big benefit to aircraft passengers, flight faster also can make airline company benefit owing to the reduction of running cost.For example, the minimizing of duration flight just needs less aircrew's time, therefore also just can pay less expense for the aircrew.In addition, because required repair and maintenance is usually according to the pilot time number that aircraft had, so flight course also can reduce the frequency and the expense of repair and maintenance faster.

US Patent 6,293,497 described the restriction transonic cruise airfoil performance physical factor and relevantly promote the transonic balance that airfoil performance carried out of cruising to greatest extent.The characteristics of airfoil performance can be measured by following four basic sides during transonic was cruised and used:

1) profile thickness usually is expressed as maximum-thickness than (maximum ga(u)ge is divided by chord length).Thickness is good more greatly, because it provides fuel oil and the required space of mechanical system, also because for identical intensity, the wing structure weight that the degree of depth is bigger is lighter.

2) speed under the preferred operations state or Mach number.The Mach number size of described aerofoil profile after the correction of the factor relevant with described wing setting, helps directly to improve the cruising speed of aircraft.

3) lift coefficient of preferred operations state.The increase of lift coefficient is more favourable, (for example, having the long distance of more fuel flights) or higher cruising altitude because it can be gained in weight.

Drag coefficient under other serviceability that can run into when 4) preferred operations state and aircraft are executed the task.Reducing resistance can reduce fuel oil consumption and increase flying distance.

Other the aspect for example pitching moment characteristic when low mach and lift capability also all is important, but not as four top basic sides important.

Described four key propertys tolerance defines the grade of performance jointly, usually is called as " technical merit " of aerofoil profile.These four key propertys have been brought conflicting problem to the designer, that is to say, tend to cause that for improving the design modification that one of them key property carries out at least one performance descends in other three performances.Therefore, for given application, a good or best design just need be carried out favourable compromise between these four performances, and the overall performance of the aircraft of using this aerofoil profile is estimated.It should be noted, the required compromise of carrying out of aerofoil profile that design has higher technical merit (determined by above-mentioned four performance figure) does not always make overall aircraft reach preferably or optimum technical merit, because higher technical merit brings bad influence can for maximum lift, maneuvering characteristics or less jitter limits aspect.

Sometimes, only on first three individual basis of above-mentioned performance metric, technical merit is made restrictive evaluation.On restrictive meaning, the technical merit of aerofoil profile can determine that it is limited than (tmac/c), lift coefficient (Cl) and Mach number (M) by maxim-thickness according to the target that the is arranged in three dimensional space serviceability that cruises.For the position in the three dimensional space being reduced to single " index ", just need additional prerequisite or rule to use following equation:

ΔM＝[-1(Δtmax/c)]+[-1/7(ΔCl)]

Above-mentioned equation is based on following prerequisite: by what constituted identical technical merit and what is offered the method for operating conditions that relates to two kinds of aerofoil profiles with constructed level by.(for example, compare and draw) that the historical data that is based on constant-1 and-1/7 gets to being considered to the cooresponding aerofoil profile of industrial grade.But, it should be noted that described constant-1 and-1/7 only is exemplary, other suitable constant also can use in above-mentioned equation.

For the technical merit of two kinds of aerofoil profiles relatively, exemplary method uses above-mentioned equation that two kinds of aerofoil profiles are adjusted to the identical point of tmax/c and Cl, and the Mach number to gained compares then.Therefore the difference that can represent two kinds of aerofoil profile technical merits by the difference on the Mach number.

The another kind of relatively exemplary method of aerofoil profile technical merit is surveyed and drawn resistance exactly increases curve (be at lift coefficient under the situation of constant relatively drag coefficient and Mach number).The low resistance opereating specification of the described slotted aerofoil that such curve can be used for showing (being shown in the below of the pressure distribution curve of Figure 16 B) extends to higher Mach number than described unit piece aerofoil profile (being shown in the below of the pressure distribution curve among Figure 16 A), has improved lift when thickness is identical a little.Certainly, can redesign described slotted aerofoil,, for example, when identical, reach even higher lift with unit piece aerofoil profile speed to reach the purpose except fair speed to use this technical advantage.

On any given technical merit, generally all may design the independent aerofoil profile that is suitable for different preferred operations conditions widely, and be embodied in the different half-way houses between four key propertys.For example, a kind of aerofoil profile can have higher operation Mach number than another kind, but this is a cost to reduce lift and to improve resistance.For competent designer, using modern computing fluid dynamics instrument to design the different aerofoil profile of given technical merit generally is the task of being simple and easy to do.On the other hand, improve technical level, for example do not have influence on any one difficulty relatively often in other three key propertys by improving one of them key property, employed technical merit was high more when the designer began, and being finished of task is also just complicated more so.If the wing that beginning is studied is on the technical merit representative in the prior art, so just extremely difficulty is made significant the improvement.

The principal element of limiting performance is relevant with the physical property at the stream of aerofoil profile upper surface.Understand these factors and just need check the cruise distribution of pressure of aerofoil profile of representative type transonic, on negative scale with pressure coefficient (C _p) draw for unit, shown in Figure 16 A, (draw) from US Patent 6,293,497.For ease of reference, the shape of described aerofoil profile 101 is shown in the pressure distribution chart below.At C as shown in the figure _pOn the scale, C _pThe=0th, free fluid is supposed with subsonic speed away from the aerofoil profile place, the static pressure that flows.On each point on surface, C _pValue except limiting pressure, also be equivalent to the specific flow speed value outside the thin viscous boundary layer on the abuts on surface.Negative C _pLow pressure that (above horizontal shaft) representative is compared with free stream and high speed, and positive C _p(below horizontal shaft) is equivalent to high pressure and low speed.The negative C of specified level _pBe equivalent to velocity of sound speed, and use dotted line 89 expressions.

Lower curve 90 on the pressure distribution chart has been represented the pressure on the lower surface 91, or the high pressure side, and upper curve 92 has been represented the pressure on the upper surface 93.Vertical distance between two curves has shown the pressure between upper surface 93 and the lower surface 91, and the zone between two curves is proportional with the resultant lift that is produced by aerofoil profile.It should be noted that near leading edge the C that is referred to as " stationary point " 95 places _pHave a high positive peak in the distribution 94, air-flow on the horizon at first " adheres to " position in described airfoil surface, and the flow velocity outside boundary 1ayer is zero.The C that also it should be noted that in upper and lower surface _pDistribution converges at described trailing edge 96, thereby defines monodrome C _p97, this value almost always one less on the occasion of.The C of this level of described trailing edge place _pPhysical property for stream causes and seriously influences.Because described trailing edge C _pMainly by integral airfoil thickness distribution decision, and thickness also usually limits by a large amount of gentle kinetic factors of structure, so the designer can only relatively less control trailing edge C _pExcept the stationary point and the trailing edge of leading edge, the designer is by changing air foil shape control presssure distribution more.

For given profile thickness and Mach number, the problem that reaches the hightech level also just is summed up as the problem that promotes high coefficient with the low resistance level to greatest extent.Do not reduce profile thickness fully to increase lift usually be impossible by increasing lower surface pressure.Therefore, increase lift as much as possible thereby designer's task is exactly the pressure that reduces upper surface, but do not increase resistance significantly simultaneously.At this on the one hand, the distribution of pressure shown in Figure 16 A has just been represented advanced method of designing.Operating conditions shown in the figure is close with the employed preferred operations condition of the part of cruising in early days in the aircraft task.Resistance under this condition is quite low, but when Mach number and/or lift coefficient increase, resistance can increase fast.

It should be noted that the upper surface C in described aerofoil profile 101 first halfs _p92 are positioned on the dotted line 89, and this shows that stream there is that middle rank is ultrasonic.This hyprsonic zone is just in time stopped by weak shock at middle string rear portion, is expressed as C from the teeth outwards _p98 the value characteristic that is increased to subsonic flow suddenly.C _pDistribution in hyprsonic zone 99 is designed to be almost flat consciously, and pressure rises very lenitively, thereby can prevent that shock wave from becoming stronger and cause resistance to increase under other condition.After the shock wave, pressure 100 increases gradually, is " pressure recovery ", up to reaching the less on the occasion of C of leading edge _p97.Reach balance between the lift that the careful design of process of the position of shock wave and distribution of pressure is adapted at increasing in recovering the zone and the resistance of increase.

The trial that increases lift tends to make wing can't reach favourable balance, and increases resistance simultaneously.For example, a kind of method of increase lift is exactly that shock wave 98 is moved backward.But, this just needs bigger recovery (because direct rear shock C _pWith trailing edge C _pBasically all fix), this can make viscous boundary layer's thickening or even from surface isolation, these 2 all can cause tangible resistance to increase.The another kind of method that increases lift is exactly to reduce shock wave front even more preceding pressure (in the front portion of whole described aerofoil profile with C _pMove on the curve 99 and increase the hyprsonic flow velocity), but this can increase the pressure of crossing over behind the shock wave, and this also can cause so-called drag due to shock wave.For the unit piece transonic airfoil of prior art, the compromise of carrying out between lift and resistance design has reached higher fineness, can not carry out any bigger improvement on technical merit.

The cruise shape of aerofoil profile 523 and resulting pressure of the transonic of cracking distributes and (draws from US Patent 6,293,497) shown in Figure 16 B.Described aerofoil profile 523 comprises two elements (anterior member 560 and posterior member 561), and with its separation, air generally flow to upper surface 564 by described seam from lower surface 584 by curved passage (562, described seam).In this example, described seam flange (565, the trailing edge of anterior member) just in time is positioned at 80% rear portion of the whole string that begins from described leading edge, and the lap of element accounts for general 3% of whole chord length.There is shown the distribution of pressure of two elements, at the overlapping local distribution of pressure of foil element also overlaid.Identical with described conventional aerofoil profile, described upper curve 566,567 is illustrated on the described upper surface 564,583 and produces C _pDistribute, described lower curve 568,569 is illustrated on the described lower surface 584,570 and produces C _pDistribute.It should be noted that to have two stationary points 571,572 and their corresponding high pressure crests 573,574, all corresponding stationary point on each crest, herein, stream on the horizon is attached near the surface each leading edge.

Be the physical property that begins to consider to flow, it should be noted that, the preferred operations condition of the described aerofoil profile 523 of cracking (among Figure 16 B pressure distribution chart under) is faster than unit piece aerofoil profile 101 (among Figure 16 A pressure distribution chart under), lift coefficient is higher slightly, and from the structure purpose, two kinds of aerofoil profiles have identical effective thickness.In the operating conditions of the described aerofoil profile of cracking, the unit piece aerofoil profile of any same thickness all has high resistance.The described substantive advantage of aerofoil profile aspect technical merit of cracking comes from this fact, that is to say that the final recovery 575 of pressure comes across rear portion very far away, starts from general 90% place of whole chord length and with weak shock 576.This distribution of pressure can not appear in the unit piece aerofoil profile, because separation of boundary layer is certain to occur, thereby prevents that shock wave from moving to rear portion far away.This mechanism can be " seam effect " by being called general, has prevented in conjunction with following a plurality of factors that boundary 1ayer from producing by the described seam of this mechanism and has separated:

1) boundary 1ayer on the upper surface 583 of front elements 560 is subjected to being in the influence of the weak shock 577 that stitches flange 565 places, but the pressure recovery after on described front elements, not having shock wave.This point is possible, because described posterior elements 561 can impel " distribution speed " in a raising of the trailing edge of described front elements (the trailing edge CP578 on described front elements is a bigger negative value, and the trailing edge CP on the unit piece aerofoil profile generally all be on the occasion of).

2) forwardly the upper surface of element 560 and lower surface boundary 1ayer converge at trailing edge 565 and are formed with wake flow, and this wake flow flows on the upper surface 564 of described posterior elements, still have in essence different with the boundary 1ayer that is formed at described posterior elements upper surface.At the rear portion of whole described posterior elements 561, described wake flow is subjected to the influence of strong pressure raising 575,576, but the mixing of strong turbulent flow can make described wake flow avoid the influence of adverse current.

3) boundary 1ayer at the upper surface 564 of described posterior elements 561 has very short distance, it is beginning to expand near the stationary point 572 the leading edge of described posterior elements on this distance, so also very thin when running into final weak shock 576 and pressure recovery region 575, can still keep attachment state.Launch about distribution of pressure and boundary 1ayer, described posterior elements 561 has general weak shock and the pressure recovery region that begins from the midpoint of its chord length in fact with the aerofoil profile that himself is a separation, and we expect to occur adhering to stream on it to this.

The upper surface distribution of pressure of Figure 16 B can be realized an extreme relatively example of described seam effect.The scope of a lot of less extreme medium distribution of pressure between the unit piece distribution of pressure in Figure 16 B and among Figure 16 A also can be utilized the described effect of cracking.The shock wave of described front elements 560 can not extend to the seam flange 565 at rear portion always, and can not have the hyprsonic zone on the upper surface 564 of described posterior elements 561 yet.In fact, Figure 16 B shows when a series of such intermediate pressure of described aerofoil profile during with the Mach number of state shown in being lower than and lift coefficient operation and distributes.Described seam effect remains needs, to prevent the appearance burbling under these other conditions.

Distribution of pressure on lower surface helps to improve the technical merit of the aerofoil profile 523 of cracking among Figure 16 B.Related pressure on the lower surface 91 of the distribution of pressure 568 on the lower surface 584 in the posterior member 560 of the described aerofoil profile 523 of cracking and the unit piece aerofoil profile 101 of Figure 16 A distributed 90 compare.On the described aerofoil profile 523 of cracking more flat distribution of pressure can cause described aerofoil profile 523 lower surface than small curve, also cause described aerofoil profile 523 to have the bigger degree of depth (generally be whole chord length 15% and 64%) in the front spar of main structure case and the residing position of rear spar.The more flat lower surface and the darker beam wing all are favourable concerning the construct validity of main structure case.This advantage also can transform and be used to promote Mach number and lift coefficient, and the construct validity (flexural strength) that keeps described wing case simultaneously is identical with the unit piece wing.

Figure 20 shows the lateral plan of conventional aerofoil profile 600, and this Airfoil Design is used for cruising with high subsonic speed and/or transonic speed.Described aerofoil profile 600 comprises the trailing edge flap 602 that singly cracks.In Figure 20, described wing flap 602 is in retracted position 604, for example, and when cruising.In retracted position 604, the front end 606 of described wing flap 602 is inserted in and is hidden in the outline line of described aerofoil profile 600.Like this, described aerofoil profile 600 constitute stream line patterns and on Pneu the smooth exterior surface face, this surface can comprise some little rank shape or slits at the most.

It should be noted that the profile of described aerofoil profile 600 and wing flap 602 only is schematic.It should be noted that also the routine aerofoil profile of cruising generally all is equipped with the leading edge raising device, but this equipment is not shown in Figure 20 to 22.

In Figure 21, described wing flap 602 is illustrated with expanded position 608, for example, and in takeoff condition.Figure 22 shows the wing flap 602 at another expanded position 610, but bigger than the drift angle shown in Figure 21.Expanded position 610 as shown in figure 22 can be used for, for example, and landing state.

For described wing flap 602 is expanded to any one expanded position 608 (Figure 21) or 610 (Figure 22) from retracted position 604 (Figure 20), described wing flap 602 all will move to the rear portion.Described wing flap 602 moved to the rear portion can open a cavity 612 when launching wing flap 602, be commonly referred to as " indent ".Shown in Figure 21 and 22, described empty 612 is non-fleetlines, and is included in steep existing lower edge 614 in the rear end 616 of described master or preceding foil element 618.

Having in the conventional trailing edge flap system that surpasses a seam (for example, two trailing edge flaps etc. that crack), when the system of flaps launches, generally can open blunt form indent above one.

Because the existence of blunt form indent can significantly not influence the rising performance, so there is not what necessity to use the seam of the outstanding method design rising usefulness on the Pneu yet.But, in cruising, having observed the existence that is positioned at seam blunt form indent before and can cause significantly, is unacceptable resistance consumption sometimes.For the shape of given routine rising wing flap and the wing flap-indent zone that is set by the wing flap expansion, the seam of rising usefulness is usually closed in cruising flight, thereby avoids occurring by the caused resistance consumption of wing flap-indent.

Shown in Fig. 6,16B and 17, embodiments of the invention comprise the aerofoil profile that has one or more seams, and these sew to limit has level and smooth fleetline profile and do not have the blunt form indent.These seams comprise having the level and smooth passage of design on good fleetline, the Pneu.Eliminate the blunt form indent and described seam is set at and have passage level and smooth on fine fleetline, the Pneu, this also can make described being sewn on cruise and other transonic state is opened down, thus cruise or other transonic state under improve performance.

Except the seam that cruises as described above is provided for aerofoil profile, the global shape of aerofoil profile or profile also can customized design to utilize described seam effect (description of seam effect see above).Comparison diagram 16A and 16B can draw the exemplary difference of crack aerofoil profile 523 and the conventional air foil shape between the aerofoil profile 101 that do not crack.For example, although also have other delicate difference between two kinds of air foil shapes, the upper surface 583 of the described aerofoil profile 523 of cracking is general more smooth than the upper surface 93 of described conventional aerofoil profile 101.

Cruise distance that wing flap (being set with the wing flap of the seam that cruises at least) needs wing flap to move backward of expansion is less than and launches the distance that the conventional rising wing flap that singly cracks moves backward.For example, shown in Figure 20 to 22, singly crack rising wing flap 602 of described routine needs to move bigger distance backward and described indent 612 could be opened to enough width, thereby does not hinder air-flow through described seam 620.From another aspect, though it is shorter under the described overlapping preferable case of cruising between wing flap and the main wing type element, but embodiments of the invention use the seam that cruises, even be set with the described wing flap that cruises that cruises seam when withdrawing fully, still maintenance is opened basically.Because described cruise wing flap cruise and raised position between backward big displacement tend to open the seam that cruises too much and hinder the raising of rising performance, so described at least in certain embodiments cruise wing flap cruise and raised position between when launching backward mobile preferred condition be minimized.

The trailing edge elevation system can integrate with the aerofoil profile of cracking in every way.

For there not being to set the part of seam of cruising along the wing spanwise, the also unnecessary setting of the described trailing edge elevation system seam that cruises.Therefore, any one in the various conventional rising wing flaps can be as the wing section that does not have the seam that cruises.

Have one or more parts of cruising seam in the wing span and can use a lot of schemes.For example, have among at least one embodiment and set at least one wing flap of seam that cruises, and also can be used as the rising wing flap that singly cracks by the deflection angle that increases this wing flap.Cruise seam or be used as the rising wing flap that singly cracks no matter set, the identical wing and the profile of wing flap all are exposed in the air-flow, and just the deflection angle of wing flap is different.

Aerofoil profile among some embodiment has at least one cruise seam and at least one conventional seam that raises, and preferable case is the upstream that these seams is set in the described seam that cruises.In these embodiments, the described seam that cruises also can be used as the seam that raises.

In the preferred embodiment of part span slotted aerofoil, only the Outboard Sections along described wing is set with the seam that cruises, for example the part of extending to spanwise between described planar structure catastrophe point and described wing tip.Except the described seam that cruises, described Outboard Sections also can have the trailing edge elevation system.The rising that the described seam that cruises only can be used as the wing outside is stitched, and in other words, the described seam that cruises can be used as the seam that raises, and raises stitched together with one or more other routines of being set by described wing Outboard Sections.

In at least one preferred embodiment of part span slotted aerofoil, the seam that cruises is not set by inside part, for example the part of extending along spanwise between described wing root and plane structural mutation point.On the contrary, described inside part comprises conventional trailing edge elevation system, and one or more risings are stitched and one or more blunt form indents of trip are sewed in described rising with just having opened after its expansion.But, in cruising flight, withdraw under the described trailing edge elevation system preferred condition, thereby close described rising seam and eliminate described blunt form indent.

Though different preferred embodiments is described, the those skilled in the art also can make amendment or variant under the situation that does not break away from notion of the present invention.Above-mentioned example just is used to the present invention is described rather than limit it.Therefore, should freely on the basis of related art, understand explanation and claim.

Claims (53)

1. a sweepback aerofoil profile comprises:

At least one has the front end foil element of upper surface and lower surface;

At least one has the rear end foil element of upper surface and lower surface; And

The seam that at least one is set under at least one transonic state of described aerofoil profile by described aerofoil profile, described seam makes the part air separation that flows along the lower surface of described front end foil element, and it is flowed on the upper surface of described rear end foil element, to improve the performance under the transonic state, the trailing edge of the contiguous host wing of a part of part span seam.

2. aircraft wing that comprises the described sweepback aerofoil profile of claim 1.

3. wing according to claim 2, wherein, described seam comprises and being set between the foil element of described front-end and back-end not with the smooth passage on the Pneu of blunt form indent.

4. wing according to claim 2 wherein, is provided with described seam by choose one or more standards from the group of following composition, and to improve the performance of described wing, described group comprises:

Improve cruising speed;

Increase lift;

Increase thickness;

Reduce the sweepback angle;

Reduce resistance; Perhaps

Above-mentioned every combination.

5. wing according to claim 2, wherein, described seam only extends to spanwise along the part of the described wing span.

6. wing according to claim 5, wherein, the about planar structure catastrophe point of described seam along spanwise from described wing begins to extend to about wing tip place of described wing.

7. wing according to claim 2, wherein, described be sewn on the wing burbling occurring and causing increasing on the part of resistance and extend under the transonic state.

8. wing according to claim 2, wherein, described seam extends to wing tip from the root of described wing basically continuously along spanwise.

9. wing according to claim 2, wherein, described sewing is arranged to the shock wave that can be produced by supersonic airstream and is pushed into position, back on described wing along described wing.

10. wing according to claim 2, wherein, the described resistance of being arranged to increase described wing-disperse the Mach number size of sewing can keep the cooresponding Pneu efficient of described wing simultaneously at least.

11. wing according to claim 2, wherein, described sewing is arranged to alleviate shock wave and made described wing that higher cruising speed is provided.

12. wing according to claim 2 also comprises the actuator structure that is coupled with described front-end and back-end foil element, this structure is used to make of described front-end and back-end foil element to move relative to another generation, thereby described seam is adjusted.

13. wing according to claim 12, wherein, described actuator structure is configured to and can adjusts described seam by one or more operations selected from operational group, and described operational group comprises:

Adjustment can separate the slit of described front-end and back-end foil element, and described seam has been set in described slit;

Adjust the relative height between the foil element of described front-end and back-end; And

Adjust the angle between the foil element of described front-end and back-end.

14. wing according to claim 12, wherein, described seam comprises a plurality of sections of vertically arranging along described wing, and each section can be adjusted independently by described actuator structure, thereby can differently adjust described seam at described span diverse location.

15. wing according to claim 2, also comprise the actuator structure that is coupled with described front-end and back-end foil element, be used for making of described front-end and back-end foil element to move relative to another generation, thereby under at least one subsonic state, close described seam, and under described transonic state, open described seam.

16. wing according to claim 2, wherein, at least one seam comprises a plurality of seams of vertically arranging along described wing.

17. wing according to claim 2 comprises that also at least one is arranged at two not slotted aerofoil zones between the slotted aerofoil zone.

18. wing according to claim 2 comprises that also at least one is arranged at the not slotted aerofoil zone between two slotted aerofoil zones.

19. wing according to claim 2, wherein, described transonic state is one or more states of selecting from the group that includes cruising condition and operation state.

20. wing according to claim 2, wherein:

Described front end foil element comprises the host wing part;

Described rear end foil element comprises wing flap; And

Described wing also comprises actuator structure, is used for adjusting described wing flap when cruising, thereby improves the performance of wing when cruising.

21. aircraft that comprises the described aerofoil profile of claim 1.

22. an aerofoil profile comprises:

At least one has the front end foil element of upper surface and lower surface;

At least one has the rear end foil element of upper surface and lower surface; With

The part span seam that at least one is set under at least one transonic state of described aerofoil profile by described aerofoil profile, described seam makes the part air separation that flows along the lower surface of described front end foil element, and it is flowed, thereby under the transonic state, obtain the improvement on the performance on the upper surface of described rear end foil element.

23. aircraft wing that comprises the described sweepback aerofoil profile of claim 21.

24. wing according to claim 23, wherein, described seam comprises and being set between the foil element of described front-end and back-end not with the smooth passage on the Pneu of blunt form indent.

25. wing according to claim 23 wherein, is provided with described seam by choose one or more standards from the group of following composition, to improve the performance of described wing, described group comprises:

Improve cruising speed;

Increase lift;

Increase thickness;

Reduce the sweepback angle;

Reduce resistance; Perhaps

Above-mentioned every combination.

26. wing according to claim 23, wherein, the about planar structure catastrophe point of described seam along spanwise from described wing begins to extend to about wing tip place of described wing.

27. wing according to claim 23, wherein, described be sewn on the wing burbling occurring and causing increasing on the part of resistance and extend under the transonic state.

28. wing according to claim 23, wherein, described sewing is arranged to the shock wave that can be produced by supersonic airstream and is pushed into position, back on described wing along described wing.

29. wing according to claim 23, wherein, the described resistance of being arranged to increase described wing-disperse the Mach number size of sewing can keep the cooresponding Pneu efficient of described wing simultaneously at least.

30. wing according to claim 23, wherein, described sewing is arranged to alleviate shock wave and made described wing that higher cruising speed is provided.

31. wing according to claim 23 also comprises the actuator structure that is coupled with described front-end and back-end foil element, this structure is used to make of described front-end and back-end foil element to move relative to another generation, thereby described seam is adjusted.

32. wing according to claim 31, wherein, described actuator structure is configured to and can adjusts described seam by one or more operations selected from operational group, and described operational group comprises:

Adjustment can separate the slit of described front-end and back-end foil element, and described seam has been set in described slit;

Adjust the relative height between the foil element of described front-end and back-end; And

Adjust the angle between the foil element of described front-end and back-end.

33. wing according to claim 31, wherein, described seam comprises a plurality of sections of vertically arranging along described wing, and each section can be adjusted independently by described actuator structure, thereby can differently adjust described seam at described span diverse location.

34. wing according to claim 23, also comprise the actuator structure that is coupled with described front-end and back-end foil element, be used for making of described front-end and back-end foil element to move relative to another generation, thereby under at least one subsonic state, close described seam, and under described transonic state, open described seam.

35. wing according to claim 23, wherein, at least one part span seam comprises a plurality of seams of vertically arranging along described wing.

36. wing according to claim 23 comprises that also at least one is arranged at two not slotted aerofoil zones between the slotted aerofoil zone.

37. wing according to claim 23 comprises that also at least one is arranged at the not slotted aerofoil zone between two slotted aerofoil zones.

38. wing according to claim 23, wherein, described transonic state is one or more states of selecting from the group that includes cruising condition and operation state.

39. wing according to claim 23, wherein:

Described front end foil element comprises the host wing part;

Described rear end foil element comprises wing flap; And

Described wing also comprises actuator structure, is used for adjusting described wing flap when cruising, thereby improves the performance of wing when cruising.

40. wing according to claim 23, wherein, described wing is the swept-back.

41. one kind comprises the aircraft according to the described wing of claim 40.

42. aircraft that comprises aerofoil profile according to claim 22.

43. the flying method of a sweepback aircraft wing, comprise the seam that uses at least one to set by described wing, the upper surface of described wing appears separating and being flowing in its feasible part air diverts that flows along described wing lower surface under at least one transonic state of described wing, the turning to of described air can postpone the burbling that causes resistance to increase at least under the transonic state, thereby the performance under the transonic state is improved.

44., further be included under the transonic state adjustment to described seam according to the described method of claim 43.

45. according to the described method of claim 43, wherein, the adjustment of described seam is comprised at least one or a plurality of operation selected from operational group, described operational group comprises:

Adjust a slit of separating described front-end and back-end foil element, described seam has been set in described slit;

Adjust the relative height between the foil element of described front-end and back-end; And

Adjust the angle between the foil element of described front-end and back-end.

46. according to the described method of claim 43, wherein:

Described front end foil element comprises the host wing part;

Described rear end foil element comprises the wing flap device; And

Adjustment to described seam comprises the described wing flap device of use.

47. according to the described method of claim 43, wherein, described seam comprises part span seam.

48. according to the described method of claim 43, wherein, described seam comprises that basically the about wing root along the whole length of the span of described wing from described wing begins the single seam that extends to about wing tip place of described wing.

49., also be included in and be in or open described seam during near the transonic state according to the described method of claim 43.

50., also be included under at least one subsonic state of wing and close described seam according to the described method of claim 43.

51. according to the described method of claim 43, wherein, described seam comprises and being set between the foil element of described front-end and back-end not with the smooth passage on the Pneu of blunt form indent.

52. the flying method of an aircraft wing, this wing has host wing part, wing flap device and at least one and be set in seam between the described wing flap device of described host wing partial sum when cruising, described method comprises that driving described wing flap device adjusts described wing flap device when cruising, thus the performance improvement when obtaining to cruise.

53. according to the described method of claim 52, wherein, described seam comprises and being set between the foil element of described front-end and back-end not with the smooth passage on the Pneu of blunt form indent.

Images