Europaisches Patentamt J» European Patent Office Publication number: 0 222 421 B1 Office europeen des brevets

EUROPEAN PATENT SPECIFICATION

Date of publication of the patent specification: mtci- B64C 21/06 07.11.90 Application number: 86201320.8 Date of filing: 24.07.86

Laminar flow control airfoil.

Priority: 04.11.85 US 794470 Proprietor: THE BOEING COMPANY, P.O. Box 3707 Mail Stop 7E-25, Seattle Washington 98124(US) Date of publication of application : 20.05.87 Bulletin 87/21 Inventor: Gratzer, Louis B., 2201 Third Avenue, Seattle Washinton 98121 (US) Publication of the grant of the patent: 07.11.90 Bulletin 90/45 Representative: Hoijtink, Reinoud et al, OCTROOIBUREAU ARNOLD & SIEDSMA Sweelinckplein 1, NL-2517 GK Den Haag(NL) Designated Contracting States: DEFRGBITNL

References cited: GB-A-2064709 US-A-4575030 AEROSPACE AMERICA, vol. 22, no. 3, March 1984, pages 72-76, New York, US; R.D- WAGNER etal.: "Fresh attack on laminar flow" JOURNAL OF AIRCRAFT, vol. 21, no. 8, August 1984, pages 612-617, New York, US; R.H. LANGE: "Design integration of laminar flow control for transport CM aircraft"

CM CM CM Q Note: Within nine months from the publication of the mention of the grant of the European patent, any person may give notice to the European Patent Office of opposition to the European patent granted. Notice of opposition shall be filed in a written reasoned A statement. It shall not be deemed to have been filed until the opposition fee has been paid (Art. 99(1) European patent Jfi convention). ACTORUM AG EP 0 222 421 B1

Description figure 10 is an enlarged cross-sectional view of the leading edge region lower surface taken along section line X-X of figure 4; The present invention relates to an airfoil for a figure 11 is an enlarged cross-sectional view of swept wing according to the preamble of claim 1 and 5 the trailing edge region of the wing taken along sec- claim 7, respectively, and as described in the article tion line XI-XI of figure 1 ; of Richard D. Wagner and Michael C. Fischer in figure 12 is a perspective view of the trailing "Aerospace America", Vol. 22, no. 3, March 1984, edge region of the wing with the trailing edge flap in pages 72-76, New York, "Fresh attack on laminar a low-speed position; flow". The described system uses suction in the 10 figure 13 is a cross-sectional view of the joint be- leading and trailing edge region of the wing such as tween two adjacent spoiler sections taken along to maintain laminar flow at high Reynolds numbers section line XIII-XIII of figure 12; and during flight conditions beyond those where figure 14 is a cutaway plan view of an air collec- laminar flow can be expected to occur naturally. By tion system for an aircraft wing; applying suction only to the leading and trailing 15 figure 15 is a plan view of the upper surface of an edges the wing box may be conventionally struc- aircraft wing incorporating slotted suction sur- tured with the intended advantages of proven faces; structural integrity, lightweight and minimum cost to figure 16 is a sectional view of a leading edge re- serve as the basic wing load carrier, and also of gion suction surface incorporating slots; proven leakproof wall integrity to serve safely as 20 figure 17 is a sectional view of a trailing edge suc- the primary fuel container, a container structure in- tion surface incorporating slots; cidentally also designed for maximized volumetric figure 18 is a simplified plan view of the upper sur- fuel capacity. face of one form of wing according to the invention, It is an object of the present invention to improve which view graphically depicts the crossflow in upon the known technic. Therefore measures are 25 each region of the wing; taken to offset the adverse effects of conditions figure 19 is a simplified cross-sectional view of that disturb the boundary layer. the aircraft wing of figure 18, taken along section The airfoil according to the characterizing part of XIX-XIX; claim 1 and claim 7, respectively, tailors the pres- figure 20 is a graph that illustrates the pressure sure distribution and suction characteristics in 30 distribution characteristics of the wing of figure 18; each critical region of the wing such as to prevent figure 21 is a simplified plan view graphically de- as much as possible the distorting effects of cross- picting the spanwise suction distribution character- flow (X-F), Tollmien-Schlichting (T-S), and shock istics of the wing of figure 1 8; waves. figure 22 is a simplified cross-sectional view of a Other objects and advantages of the present in- 35 modified form of wing construction that incorpo- vention will become apparent to one skilled in the art rates a leading edge flap; and after a reading of the following description taken to- figure 23 is a graph that illustrates the pressure gether with the accompanying drawings, in which: distribution characteristics of the wing of figure 22.

Figure 1 is a plan view of the upper surface of a 40 Detailed Description wing attached to the fuselage of an aircraft; figure 2 is a plan view of the lower surface of an In figures 1 and 2, wing 10, attached to a fuselage aircraft wing; 12, is divided into three sections or regions: (1) lead- figure 3 is a cross-sectional view of an aircraft ing edge region 18, (2) main box region 20, and (3) wing taken along section Ill-Ill of figure 1 ; 45 trailing edge region 22 as shown in figure 3. Region figure 4 shows an enlarged cross-sectional view 18 extends chordwise between the leading edge or of the leading edge region of the wing taken along nose 14 and a front main box spar 26 forming part of section line I V-IV of figure 1 ; the main box region 20. The trailing edge region 22 figure 5 is an enlarged cross-sectional view of extends aft from rear main box spar 30 to the trailing the leading edge region upper surface taken along 50 edge 16. In accordance with the present novel hy- section line V-V of figure 4; brid wing concept, flow over leading edge region 18 figure 6 is an enlarged partial cross-sectional and trailing edge region 22 is kept laminar with the view of the leading edge region of the wing with the aid of suction applied through their surfaces while leading edge flap in the low-speed position. flow over the surfaces of the main box region 20 is Figure 7 is an enlarged perspective view of the 55 kept laminar naturally by designing the surface con- inner surface of the leading edge region of the wing tours of those surfaces to promote the required fa- wherein the front edge of the leading edge flap vorable pressure gradients. comes up against a front stop plate; In figure 4 nosepiece 54 is configured in an ap- figure 8 is an enlarged perspective view of the in- proximate parabolic profile extending into an upper ner surface of the leading edge region of the wing 60 edge portion 54a and a lower edge portion 54b. Ex- wherein the rear edge of the leading edge flap tending the length of wing 10, nosepiece 54 is pref- comes up against the flap stops; erably of integral titanium honeycomb core con- figure 9 is an enlargend cross-sectional view of struction. Its rearward edges are joined to upright ' a joint between two leading edge flap surfaces tak- nose spar 56, also extending the spanwise length of en along section line IX— IX of figure 2; 65 the nosepiece. EP 0 222 421 B1

Also joined to nose spar 56 and extending aft as a tions of insects and dirt, and of surface erosion continuation of nosepiece upper surface 54a is a that may have occurred. dual skin panel 34 including porous (multiply aper- Slot 86 is formed by the gap or space in between tured) outer skin 60 and a nonporous inner skin 62 the rear edge of the nosepiece lower portion 54b maintained parallel to skin 60 by chordwise-oriented 5 and the front edge of the leading edge flap 80. A parallel spacers 64 (figure 5). The spaces 66 thus substantially planar front stop plate 90 is mounted formed within panel 34 form suction air collection adjacent the rear edge of nosepiece lower portion chambers and flow passages that conduct air drawn 54b and serves as a stop for the front edge of the through the distributed pores or apertures in skin leading edge flap 80 when the flap is retracted into 60 rearward toward upper surface duct 70. Duct 70 10 cruise position. A series of downwardly oriented is formed forwardly of and above the diagonally ori- projections 92, spaced apart spanwise of the wing, ented partition 76 extending between the lower edge extend from the lower surface of this front stop of an upright auxiliary spar 74, spaced forwardly plate and serve as the actual stop abutments. As from the front spar 26, and upper panel 34. In the shown in figure 7, projections 92, the lower surface space between auxiliary spar 74 and front spar 26, 15 of stop plate 90 and the inner surface of the leading • the inner skin 62 contains multiple holes (not shown) edge flap 80 form entry passages 94 for suction which are sized to meter the suction airflow into up- airflow passing from the leading edge front slot 86 per surface duct 70. into purge duct 100. This duct is formed by the nose Still referring to figure 4, the leading edge region spar 56, skin panel 62, auxiliary spar 74, and the lower surface skin panel assembly 42 includes 20 leading edge flap 80. Slot 88 also opens directly into spanwise extending lines of suction slots 86 and 88 purge duct 100. At cruise speeds, purge duct 100 formed at the forward and rear edges of leading under suction created by compressor 207 (figure edge flap 80. Aft of slot line 88 this lower surface 14), is maintained at a pressure lower than that out- assembly 42 comprises a short bullnose section side the slots 86 and 88. 104a contiguous with a hollow skin panel 104 having 25 The forward edge 104a (figure 8) of the leading a porous or multiply apertured outer skin and an in- edge section outer skin 104 is turned inwardly to its ner skin. Leading edge flap 80 is mounted to be de- terminus, so as to form a forwardly facing convex ployed and retracted by a suitable (whose details surface as one side of an orifice comprising the are or may be conventional) mechanism 81, typical rear slot 88. Mounted on the upper side of the for- of leading edge flap systems. When extended from 30 ward edge 104a are a series of forwardly extending retracted cruise setting (figure 4), into the de- flap stops 106 carried in positions spaced apart ployed setting (figure 6), leading edge flap 80 in- spanwise of the wing by the trailing edge of the lead- creases lift for low speed operation in the usual ing edge flap 80. The leading edge flap 80 abuts up- way. Moreover, as a unique additional feature, the wardly against these flap stops when the flap is fully deployed leading edge flap 80 is positioned for- 35 stowed in its cruise position. The flap actuator wardly of nose 54 and oriented at a rearward in- mechanism aided by differential air pressure acting cline. It thereby serves as a deflector that protects upon the flap 80 operates to hold the flap firmyl in the leading edge wing surfaces from the accumula- its cruise position once it is retracted by such mech- tions of insects and dirt, and, to a substantial de- anism. gree, from surface erosion due to dirt, rain, or hail. 40 As previously mentioned, and as shown in fig- Such accumulations and erosion create surface dis- ures 3 and 4, leading edge flap 80 when stowed is continuities that interfere with the maintenance of essentially horizontal. Moreover, its exposed sur- laminar flow over the leading edge region of the face contour is fairly flat. Hence, the values of the wing 10. As shown in figures 4 and 6, the leading coefficient of pressure Cp over most of its chordal edge flap comprises a main panel which forms part 45 length remain uniformly positive and thereby favo- of the leading edge region lower surface and an rable to laminar flow. At its edges suction slot 86 auxiliary flap 82 hinged to it that is normally folded and 88 are effective to eliminate flow disturbances over to be received within the leading edge region and maintain laminar flow. interior with the leading edge flap retracted. When Each wing carries a series of similar leading edge the leading edge flap 80 is deployed, auxiliary flap 50 flaps 80 mounted end to end as shown. Channel- 82 is pivoted downwardly and forwardly and forms shaped flexible metal seals 107 (figure 9) are fas- an aerodynamic leading edge to complete the con- tened to the end edges of the flaps and cooperate tour of the resultant high-lift wing leading edge sec- to seal the joints between such ends sufficiently to tion. prevent all but minor amounts of suction air to pass Surface jogs or discontinuities created in the low- 55 between them into purge duct 1 00. er surface of the leading and trailing sections of Referring to figures 4 and 10, the remaining panel leading edge flap 80 in the stored cruise position 104 of the leading edge region lower surface 42 is are initiating factors creating turbulence and flow formed by a porous outer skin 108 and a vented in- separation. Suction applied adjacent these disconti- ner skin 110 maintained in spaced parallel relation- nuities through the previously mentioned slots 86 60 ship by intervening spacers 112. Suction air drawn and 88 are effective to maintain laminar flow at through the pores or apertures in the outer skin 108 these locations. The suction applied through these into the passage spaces 114 between these skins slots also minimizes crossflow instability along the flows through vents 116 in the inner skin 110 and leading edge or nose of the wing. It also minimizes from there into the lower surface duct 120. These the turbulence-inducing effect of any accumuia- 65 EP 0 222 421 B1

multiple holes or vents are approximately placed and er and inner skins forms a spoiler suction air collec- sized to meter the suction airflow through the outer tion duct 150. The outer skin 142 is affixed by its for- skin 108. The lower surface duct 120 is formed by ward edge to top edge 30a of the rear box spar 30 the front spar 26, the partition 76 and the leading and extends rearwardly to a point adjacent an upper edge lower skin 104. 5 surface crossmember 154. The forward edge of the The main box region 20 is contoured to provide a inner skin 144 is mounted to lie on the lower side of a pressure distribution, to induce laminar flow natural- horizontal flange of a hanger bracket 158 by means ly, that is, without resort to surface pores or slot of sliding joint 160. The hanger bracket 158 is at- distributed over its surface area. It thereby avoids tached to the top edge 30a of the rear spar and ex- the complications and added weight attending use of 10 tends downwardly and rearwardly to a point adja- subsurface structures and ducting going along with cent the forward edge of the inner skin. The sliding a porous or slotted suction surface approach to joint 160 allows the inner skin to move relative to achieving laminar flow, especially when applied bracket 158 as the spoiler 124 pivots with respect to over the major wing section's surface area and to the wing 10. Leakage of air past the sliding joint is some extent, at least, unavoidably intruding upon 15 minimized by the placement of the inner skin 144 be- what is preferably a proven conventional structural low the flange of bracket 158. Since the air pres- and fuel storage wing section. In order to achieve sure inside of spoiler duct 150 is lower than the natural laminar flow over its chordal extent at top pressure outside, inner skin 144 is forced upwardly and bottom surfaces, the contours of these surfac- against the bracket 158. es are only moderately convex. As shown in simpli- 20 The rigid section 140 of the trailing edge spoiler fied general form in figure 3, the main box region top 124 consists of a porous upper skin 164 and a non- skin panel 36 spans between the top edges 26a and porous lower skin 166. The forward edge of the up- 30a of front and rear main box spars 26 and 30, per skin 164 is spaced away from the forward edge while the main box region lower surface 44 similarly of the lower skin 166 by the upper surface cross- spans between the respective lower edges 26b and 25 member 154, and the rear edges of the upper and 30b of these spars. Details of the wing box struc- lower skins converge with one another at the rear ture omitted from the drawings may be conventional, edge of the spoiler 124. The space between the up- bur are not required to be so. per and lower skins 164 and 166 forms a rear spoiler Referring to figure 1 1 , the wing's trailing edge re- suction air collection chamber 168. The upper sur- gion 22 extending aft from main box region 20 com- 30 face crossmember 154 has vent holes 170 that allow prises an upper surface spoiler 124, a lower sur- air to be drawn through the upper skin 164 into the face trailing edge fairing 126, and an extendable rear spoiler chamber 168 and into the flexible spoiler trailing edge main flap 128. Suction surface panels duct 150. are used selectively in the trailing edge region 22 in As shown in figures 1 1 and 12, flexible spoiler sec- order to help maintain laminar airflow proceeding aft 35 tion 138 is designed to bend downwardly in a curve from the main box surfaces. Trailing edge flap com- 171 so as to maintain a smooth fairing with the flap posite 128, preferably conventional, includes a pri- 128 with the latter in its deployed position. If de- mary flap member 132 and a secondary flap member sired, duct suction can also be applied to the porous 134 mounted forwardly of member 132, in fixed rela- upper surfaces of the spoiler during low-speed tionship thereto, defining a slot 176 between them. 40 flight as well as during cruise flight in order to mini- Since the linkage and drive mans for mounting and mize flow separation and thereby increase low- moving the flap composite 128 between deployed speed capability. This is also true of other porous and stowed positions are or may be conventional, and slotted upper surfaces of the wing disclosed the drawings herein are simplified by omitting these herein. The linkages and actuators for moving the details. 45 spoiler assembly are or may be of simple and Trailing edge spoiler 124 comprises a forward straightforward design using standardized compo- . flexible surface section 138 and a rigid section 140 nents and therefore require no detailing herein. continuing aft from it and having about the same With continued reference to figures 11 and 12, the chordal extent. The flexible section 138 is joined by trailing edge fairing 126 comprises a forward benda- its forward edge to the top edge 30a of the rear 50 bly flexible portion 172 and a rear portion 174 of rigid spar 30. Joined to section 138, rigid section 140 ter- form. The flexible portion includes a porous flexible minates aft thereof at a point approximately at the outer skin panel 178 and a nonporous inner plate mid-chord position of the flap, where it closely over- 180. The forward edge of the outer skin 178 is af- lies and fairs into the top side of flap 132 with the lat- fixed to the bottom edge 30b of the rear spar 30 and ter stowed in the cruise position. The lower surface 55 abuts the rear edge of the main box lower skin panel contour of section 140 is relieved concavely so as 44. Spaced inwardly from and oriented substantial- to accommodate the convexed top side contour of ly parallel to the outer skin panel 178 is the inner flap 132 immediately underlying it with the flap plate 180. The forward edge of the inner plate is sup- stowed. ported through a sliding joint 182 to a support mem- In the present embodiment, the flexible surface 60 ber 184. The member 184 is an inverted version of section 138 is configured of a porous, flexible outer hanger bracket 158, with the forward portion of skin 142, which serves as a suction surface, and a member 184 affixed to the bottom edge 30b of rear nonporous, substantially inflexible inner skin 144 spar 30. Sliding joint 182 is essentially identical to that is approximately parallel to and spaced inwardly sliding joint 160 and allows plate 180 to move relative from the outer skin 142. The space between the out- 65 to support member 184 when fairing 126 bends with EP 0 222 421 B1

respect to wing 10. The rear edges of outer skin responding suction requirements for each section panel 178 and plate 180 are separated by a lower of the wing. All suction air entering the compressor surface crossmember 188. The space between the is discharged into the free stream through exhaust outer skin panel 178 and the inner plate 180 forms a duct 210. If a gas turbine (208) is used to drive the flexible fairing duct 1 90. 5 compressor, combustion air for the turbine enters Contiguous with the rear end of flexible portion through inlet 212 and the resultant exhaust gases 172 is the rigid portion 174 of fairing 126. The rigid from the turbine exit by way of exhaust duct 214. portion 174 consists of a porous, rigid outer skin In the embodiment of figure 15, suction is distrib- panel 192 and a nonporous, rigid inner panel 194. uted over the skin surfaces of the leading edge re- The forward edges of members 192 and 194 are 10 gion 18 and trailing edge region 22 by lines of slots spaced apart by the lower surface crossmember 220 rather than by way of porous or multiply aper- 1 88 that has apertures that allow air to flow from a tured skin panels as in. the previous embodiment rear fairing chamber 200 formed within the rigid por- wherein, in addition to slots placed at locations of tion 174 to the flexible fairing duct 190 formed within discontinuities, the major surface panels of these the flexible portion 172. The rearmost edges of mem- 15 sections are porous or multiply apertured. bers 1 92 and 1 94 converging together form the trail- The approximately parallel lines of slots 220 ex- ing edge of fairing 126. tend spanwise of the wing starting adjacent the fu- The flexible fairing duct 1 90 draws boundary layer selage and in a nearly parallel relationship. For cer- air through pores in the lower outer skin panel 178 tain airplane designs and operating conditions, it is of the flexible portion 172 and through those in outer 20 desirable to limit the spanwise extent of the suction skin 192 of rigid portion 174, thereby promoting lami- areas to avoid structural complexity and excess nar flow over the trailing edge fairing 126. suction airflow. Thus, in figure 15 the slots are grad- For low-speed operation with trailing edge flaps uated in length, the shortest being located adjacent extended, the fairing surface 126 is caused to move the leading edge and the longest spaced aft thereof in concert with the flap through a suitable linkage 25 being located adjacent the front main box spar 26 (not shown) which places it in a faired position as il- (see also figure 4). For rigid panels, a suitable way lustrated in figure 12. In the low-speed faired posi- to form and apply suction to the array of slots 220 is tion, air flowing over fairing 126 is directed toward shown in figure 16. The slotted outer skin panel 22 is slot 176 in trailing edge flap 128. mounted on and fastened to an underlying support As with the leading edge flaps, the spoilers 124 30 panel 224 with corrugations forming flow distribu- and fairings 126 are divided into sections extending tion channels 226 which are in register with the re- end to end along the span of the wing. Flexible spoil- spective slots. Air bleed holes 228 in the bases of er duct 150 and flexible fairing duct 190, formed sec- these channels allow suction air from the channels tionally within the spoiler and fairing sections, re- to enter duct passages 232 formed between the out- spectively, require joint seals between their adja- 35 er skin panel 222 and the inner skin panel 234. For cent section ends. Figure 13 illustrates a represent- flexible panels, a modified LFC slot system may be ative form of seal between the ends of two spoiler applied in a similar manner to the trailing edge sec- sections. In this seal, two flanged flexibly resilient tion surfaces. Figure 17 depicts this. It includes metal or plastic strips 202 bear against each other — slots 221 in the outer surface panel 228, a support across the upper joint gap 203 so as to sufficiently 40 panel 240 with corrugations providing flow chan- eliminate most of the airflow occurring through the nels 242 with bleed holes 244, and an arrangement gap, a perfect seal not being required. A flapper wherein those holes lead into the associated main type seal 204 is used along the lower joint gap 205 suction duct (not shown). to seal the gap. Seal 204 is flexible so that when air Referring again to figure 15, a triangularly shap- is drawn through spoiler duct 150, seals 204 press 45 ed area 250 on the inboard upper surface of wing 1 0 against one another not permitting any air to pass is located rearward of rear spar 30. For a wing hav- through gap 205. ing a planform such as that illustrated in figure 15, In the illustrated suction air collection system 206 the surface of area 250 would be a suction surface. (figure 14) an air compressor 207 driven by a small If a porous skin is used, a construction similar to gas turbine or electric motor 208 pulls air through all 50 that illustrated in figure 5 would be employed. If suc- of the leading edge region ducts 70, 120, 100, by way tion slots are used as shown in figure 15, a skin con- of suction manifolds 209 and 209a. Manifold 209 figuration similar to that illustrated in figure 1 6 would collects airflow from the spanwise upper surface be used. Air drawn through the surface of area 250 duct 70 delivering it to the high-pressure ratio sec- is passed into spoiler duct 150. tion of the compressor. Manifold 209a collects air- 55 As noted above, an important aim of the invention flow from the spanwise lower surface duct 120 and is to achieve laminar flow control in a way that purge duct 100 and delivers it to the low-pressure avoids structural compromise and complexity within ratio section of the compressor. The trailing edge the wing box area. This is accomplished through a region ducts 150 and 190 feed airflow into the combination of airfoil shape and suction in the lead- high-pressure ratio and low-pressure sections of 60 ing and trailing edge regions of the wing such that the compressor respectively. Adjustable control the airfoil suction distribution and pressure distribu- valves near the compressor (not shown) permit flow tion characteristics work effectively and in harmo- matching to compressor operating conditions and al- ny to maintain laminar flow over a wide range of op- so allow changes in airflow distribution into the com- erating conditions. pressor as required by flight conditions and the cor- 65 To implement this approach, it is necessary to EP 0 222 421 B1 10 take measures to offset the adverse effects of centage (X/C) of chord length C. Curve U in figure conditions that disturb the boundary layer. Of prin- 20 shows the pressure distribution over the upper cipal concern are the distorting effects of cross- surface of the wing, while curve L represents the flow (X-F), Tollmien-Schlichting (T-S), and shock distribution over the lower surface of the wing. waves. Since these conditions occur at different 5 Starting from the attachment line AL the pressure regions and in different degrees across the wing, it must fall rapidly in order to make the transition to a is necessary to tailor the pressure distribution and level on the upper surface that is needed to devel- suction characteristics in each critical region of the op the wing lift. This produces a very strong X-F, wing. This can be better understood with reference which is partially offset by contouring the surface to figures 18 and 19, which illustrate another form of 10 to make the gradient as steep as possible. The re- wing according to the invention. The wing illustrated sulting surface curvature also helps to stabilize the in these figures is similar to the wing 1 0 discussed X-F. Nose suction is provided in addition to com- above. The principal differences are the omission plete the stabilization process and also to offset of a leading edge flap 80, the addition of suction in flow contamination that may result from upstream the nose region of the wing, and different overalll 15 sources of turbulence. Those skilled in the art will contours of the upper and lower surfaces of the appreciate both the existence and location of the wing as defined by the pressure distribution curves peak pressure position 1 immediately aft of the lead- of figure 20. Like numerals are used to identify the ing edge 14. common structural components. For clarity of dis- To provide additional offset to the strong cross- cussion, details of construction, such as auxiliary 20 flow XF1 , a slight positive pressure gradient is pro- spar 74 (see figure 4), have been omitted. It will be vided, starting immediately aft of the leading edge understood that the wing structure of figures 18 and peak 1 and continuing to the front spar 26. In addi- 19 would include all components as needed to pro- . tion to offsetting the leading edge crossflow, this vide suction, ducting, and the like in the manner pre- positive pressure gradient also provides a residual viously described. 25 effect beyond the front spar 26 that helps to offset To provide the proper combination of pressure the crossflow that is generated downstream in the gradient and suction, it is necessary to accommo- region between the front and rear spars 26 and 30, date the variations in crossflow found in each respectively. The use of a positive gradient to off- chordwise region of the wing. In figure 18, the ar- set the crossflow mode in the leading edge region (1- rows designated XF1 through XF5 emanating from 30 26) tends to destabilize the Tollmien-Schlichting the typical streamline S graphically illustrate the mode, which further increases the T-S amplification magnitude and direction of the crossflow in a wing normally expected in the main box region (26-30). To having a sweep angle A. In the nose region of the stabilize this T-S mode, a negative pressure gradi- wing, beginning at the edge 14, the sweep of the ent is provided aft of the front spar 26. This nega- wing produces crossflow that is directed inboard 35 tive gradient also produces a crossflow XF3, which and that is very high. In the remainder of the leading is now directed inboard and of relatively small mag- edge region, between a position 1 of peak pressure nitude. Care must be taken to avoid having the and the front spar, the crossflow changes direction pressure gradient be so large as to produce exces- and is relatively weak, as shown by the arrow desig- sive crossflow amplification. It will be recognized nated XF2. In the main box region between the front 40 that there is a proper balance that must be spar 26 and the rear spar 30, the crossflow is rela- achieved, depending upon the Reynolds number and tively mild and is directed inboard as shown by the the wing sweep. It will further be appreciated that, arrow designated XF3. In the region of the shock at the cruise Mach number, a shock wave will inevi- wave (30-4) aft of the rear spar, strong crossflow tably occur aft of the rear spar 30. To avoid exces- is directed outboard as indicated by the arrow des- 45 sive shock losses, the peak pressure at the rear ignated XF4. Finally, in the region from position 4 to spar on the upper surface must be limited. To soften the trailing edge 16, a moderate crossflow is direct- the shock, it is also preferable to reduce the initial ed outboard as indicated by the arrow XF5. gradient in the region (26-30) by about 50 percent The various pressure characteristics illustrated to about 75 percent forward of the rear spar. The in figure 20 are selected to offset the unstable 50 point at which this gradient change occurs is aft of crossflow (X-F) and Tollmien-Schlichting (T-S) the front spar, approximately two-thirds of the dis- modes just discussed. In general, either suction or tance between the front and rear spars. a negative gradient (dCp/dx/c) can be utilized to sta- From a position 4, which lies behind the shock bilize the laminar boundary layer. A negative pres- wave, the pressures recover normally past the sure gradient is desirable, since it is simpler to 55 trailing edge 5 of the rigid section 140 of the trailing achieve and less costly to apply, particularly in re- edge spoiler 124 to the wing trailing edge 16. The po- gions where the wing structure must not be compro- sitioning of the beginning of the recovery must be mised. However, on a swept wing any pressure gra- selected at about the rear spar 30 to avoid separa- dient (positive or negative) will produce crossflow tion or excessive boundary layer buildup at the wing that must be stabilized by suction if the X-F mode 60 trailing edge. It will be observed that the recovery il- becomes too large. The graphic presentation of fig- lustrated in figure 20 (shock followed by subsonic ure 20 shows a plot of the coefficient of pressure recovery to a positive value at the trailing edge 1 6) Cp as a function of displacement (X) from the lead- is typical of conventional high-speed, swept wings. ing edge 14 being graduated in term sof decimal per- As shown by curve L, the pressure distribution 65 on the lower surface of the wing substantially mir- 11 EP 0 222 421 B1 12

rors the pressure distribution on the upper sur- is less in subregion 4-5, but higher than that re- face. A steep negative gradient rises from the at- quired on the leading edge region upper surface 34. tachment line AL to a lower leading edge peak posi- On the wing lower surface, a lower suction rate is tion 1a. The gradient then turns slightly positive required for the surface of the trailing edge fairing over the leading edge region lower surface 42, then 5 126 than is required over the trailing edge spoiler changes to a negative gradient at the front spar. 124, since no shock waves are present and since Since no shock waves. are present on the lower sur- positive pressure gradients are less severe. face, it is not necessary to provide a gradient The shaded areas NS, LES, and TES of figure 21 change across the main box region lower surface illustrate the preferred spanwise areas of the noise 44 between the front and rear spars 26 and 30, re- 10 region, leading edge region, and trailing edge re- spectively. Aft of the rear spar 30, the pressure gion, respectively, in which suction is applied. recovers in a normal, positive manner past the trail- These regions are substantially the same for the up- ing edge 8 of the trailing edge fairing 126 to the trail- per and lower surfaces. In the unshaded areas, no ing edge 16 of the wing. suction is applied and, accordingly, laminar flow To complement the airfoil pressure distribution 15 control is achieved through surface contour alone. just described, suction is applied in the nose region, If desired, of course, suction could be provided the leading edge region, and trailing edge region. As entire spanwise dimension of the wing from root to used herein, the nose region is that included part of tip. Since an aim is to complicate the surface of the the leading edge region 18 (figures 19 and 22) for- wing with suction only to the extent necessary, and ward of the nose spar 56. Within this nose region, 20 since the need for suction, particularly on the upper suction is applied on the upper and lower surfaces surface, diminishes in the outboard region of the on an inboard area that extends from position 1, wing, spanwise limitation of the suction is generally through the leading edge 14, to the position 1a. To preferred. The nose suction is terminated in general accomplish this, the nose region structure of figure at a position I for which the attachment line Rey- 4 is modified by substituting a porous skin panel (of 25 nolds number, ReAL, becomes less than 100. The a construction such as that for skin panel 34) for definition of the attachment line Reynolds number in the nosepiece 54. Suction is applied to this porous terms of wing sweep and airfoil nose shape in ac- nose skin by interconnecting the D-shaped duct cordance with established methods will be readily that exists forward of the nose spar 56 with the suc- discernible to those skilled in the art. Suction in the tion air collection system of figure 14. The suction 30 leading edge and trailing edge regions is generally so applied stabilizes the attachment line boundary terminated at a position II that is a function of wing layer and offsets spanwise contamination effects sweep. In general, position II corresponds to a that are associated with wing sweep. As will be dis- spanwise position such that the chord Reynolds cussed in greater detail below in connection with fig- number (Re) satisfies the following condition: ure 21, the nose suction extends only part span to a 35 Rc(sin A)3/2 = 5 x 1 06 for A 2: 1 5°; and position I. Re - 38 x 106 for A < 15°. As noted earlier, the use of the positive pres- Consideration of the structure of a high-speed sure gradient to offset crossflow on the leading aircraft wing may impose practical constraints that edge region upper surface 34 (between the peak dictate where the trailing edge suction is spanwise position 1 and the front spar 26) destabilizes the 40 terminated. For example, the outboard regions of Tollmien-Schlichting mode. Suction is applied in the the upper surface trailing edge of such a wing typi- leading edge region to control this T-S mode, to sta- cally include a control surface such as an aileron. bilize the crossflow mode, and to thin the boundary As well, the spoilers typically included do not gener- layer ahead of the front spar 26. Preferably, the ally extend to the tip of the wing. In view of this, the suction is included on both the leading edge region 45 suction on the trailing edge will generally terminate upper surface 34 and the leading edge region lower either where the control surface (aileron) starts or surface 42. It is generally desirable to also limit the at the most outboard spoiler. spanwise extent of leading edge suction to a region For the airfoil sections outboard of the termina- that lies inboard of a terminating position II (figure tion of the leading and trailing edge suction regions, 21). 50 contour alone is used to promote natural laminar It is again to be observed that in the main box re- flow. These regions have somewhat different pres- gion, aft of the forward spar, suction is not provid- sure distributions and shapes from the suction air- ed, so that the only control over conditions that dis- foils used inboard. In general, the pressure gradi- turb the boundary layer is through the use of prop- ents on both upper and lower surfaces are continu- erly selected pressure gradients. In the trailing 55 ously negative in the leading edge regions edge region that follows the rear spar 30, suction is (beginning at points 1 and 1a in figure 19) aft to the applied on the trailing edge spoiler 124 and trailing rear spar 30. These negative gradients are also edge fairing 126 to stabilize Tollmien-Schlichting and steeper, usually in the range of about ♦ 6 < dCp/dx/c crossflow modes and avoid to premature separation < • 8, in order to provide laminar flow to the rear the surface. The covered by the 60 on upper region spar. trailing edge spoiler 124 is divided into subregions In the simplified view of figure 21, the inboard 30-4 and 4-5. Suction is concentrated in relatively suction areas are shown extending all of the way in- the subregion 30-4 to stabilize the boundary layer ' board to the root chord where the wing joins the fu- flow through the shock wave that will normally ap- selage 12. It will be appreciated by skilled aerody- pear here in cruise flight. The required suction rate 65 13 EP 0 222 421 B1 14

namicists that the typical boundary layer flow char- the trailing edge 16 as used in the configuration of acteristics of the wing region adjacent the root figures 19 and 20 are used between the front spar chord may make it difficult or undesirable to provide and trailing edge in the configuration of figures 22 laminar flow control in this area. In designing a wing and 23. according to the invention, it is generally an aim to 5 As noted above, the invention aims at striking a provide the pressure distribution shown in figure 20 harmonious balance between the use of suction and over as much of the wing as possible through the contour to maintain laminar flow. Since the use of combined use of surface contour and suction. Prac- suction necessarily complicates the design of a re- tical design constraints may require restricting the gion and requires power, it is desirable to minimize laminar flow control to inboard sections of the wing 10 the suction requirements where possible. Thus, in beginning at the positions discussed above and ter- the configuration of figure 22, it is desirable to minating at an inboard position adjacent the root avoid the necessity of. providing suction on the chord. leading edge flap 80 itself. This is accomplished It should also be pointed out that it may not always through the provision of suction at the leading edge be desirable or cost-effective to design for full- 15 front and rear slots 86 and 88 and through the use chord laminarization of the wing surfaces. In such of stepped gradients between the attachment line cases, certain areas may be eliminated from consid- and the rear spar 26. Nose suction is applied in the eration without compromising the benefits due to nose region between the leading edge and the lower laminar flow in those areas that remain. For exam- surface peak position 1a to stabilize the crossflow ' pie, the elimination of trailing edge suction would still 20 and thin the bondary layer ahead of the leading allow laminar flow to the rear spar, while avoiding edge flap 80. Adjacent the peak position 1a, metered the complication associated with suction in an area suction inflow is induced through the front slot 86 to usually occupied by various mechanical systems, stabilize the crossflow mode and avoid the disrup- flaps, and controls. The compromises made will, of tive effects on the flow that the slot would other- course, depend upon the particular design require- 25 wise produce. A relatively flat (but slightly positive) ments and mission of the airplane. gradient is maintained across the leading edge flap In situations where it is desirable or necessary to and additional suction inflow is induced through the incorporate a leading edge device, such as leading aft slot 88. This suction thins the boundary layer edge flap 80, the airfoil suction and pressure distri- and avoids the disruptive effects on the flow that bution are modified to provide the characteristics il- 30 the aft slot would otherwise produce. In addition, lustrated in figures 22 and 23. For the upper sur- this suction through the aft slot provides favorable face extending from the leading edge 14 to the trail- conditions for flow onset to the fixed surface be- ing edge 16, the characteristics shown by the curve hind the flap. In the remainder of the leading edge U are essentially the same as the characteristics il- region between the aft slot 88 and the front spar 26 lustrated in figure 20 for the nonleading edge de- 35 (i.e., on panel 104 of figure 4) suction is provided to vice configuration of figure 19. For the lower sur- stabilize the crossflow mode and to properly condi- face, the presence of the leading edge flap 80 ne- tion the boundary layer for the natural laminar flow cessitates modification of the suction and pressure over the main box region lower surface 44 between distribution characteristics as shown by curve L' in the front and rear spars 26 and 30. The suction con- order to maintain laminar flow and obtain the desired 40 ditions in the trailing edge region between the rear lift characteristics. spar 30 and the tip 8 of the trailing edge fairing 126 Referring to figures 22 and 23, the lower surface are the same as for the wing configuration of figure of the nose region is contoured to provide a steep 19. negative gradient from the attachment line to a first The spanwise suction characteristics of the up- peak position 1a, which is located immediately for- 45 per surface of the wing of figure 22 would be similar ward of the suction slot 86 formed between the rear to those for the upper surface of the wing of figure edge of the nosepiece lower portion 54b and the 19, as discussed in conjunction with figure 21. Thus, front edge of the leading edge flap 80 (see figure the suction in the nose region, the leading edge re- 4). This steep negative pressure gradient is re- gion, and the trailing edge region can be limited to quired to minimize the intense amplification of the 50 the inboard sections of the wing, with the outboard crossflow mode due to sweep. The flap 80 is con- sections of the wing being contoured to promote nat- toured to provide a slight positive pressure gradi- ural laminar flow. In the configuration of figure 22, ent from the peak position 1a to the leading edge aft however, the presence of the leading edge flap 80 slot 88. This gradient tends to stabilize the leading precludes the maintenance of natural laminar flow edge crossflow mode and provides a residual ef- 55 on the outboard portions of the lower surface. Con- fect that tends to offset the crossflow that is devel- sequently, the suction in the nose region and lead- oped downstream of the leading edge flap 80 in the ing edge region of the lower surface for this config- region between the leading edge aft slot 88 and the uration is not limited to an inboard section, but, rath- front spar 26 and in the main box region beyond the er, extends full-span to the tip the wing. front spar. For the main box region and for the two 60 Although the illustrations used here apply gener- subregions (30-8 and 8-16) in the trailing edge re- ally to an airplane wing, the fundamental teachings gion, the boundary layer conditions requiring con- can be adapted equally well to other wing-like sur- trol are the same as for the wing configuration of faces of an airplane, such as in the empennage, in- figure 19. Accordingly, the same pressure distribu- cluding those areas that incorporate control sur- tion characteristics between the front spar 26 and 65 faces. 15 EP 0 222 421 B1 16

Claims board section of said nose region (54) having a ter- minating position for which the attachment line Rey- and 1 . A laminar flow control airfoil for a swept wing nolds number is less than approximately 100; (10), having an upper surface (34,36,38), a lower said leading edge suction means applies suction in surface (42,44,46), a leading edge (14), a trailing 5 an inboard section of said leading edge region (18) edge (16), a tip, and a root, said airfoil including having a terminating position that satisfies the fol- front and rear spars (26, 30) that chordwise divide lowing conditions: said airfoil into a leading edge region (18), a main box Rc(sin A)3/2 = 5 x 1 06, a > 1 5°; and region (20), and a trailing edge region (22), said Rc= 38x106forA<15°; leading edge region (18) extending between said 10 where Re = chord Reynolds number; and leading edge (14), and said front spar (26), said main A = the sweep angle of the leading edge of said air- box region (20) extending aft of said leading edge re- foil. gion (18) between said front and rear spar (26,30), 7. A laminar flow control airfoil for a swept wing, said trailing edge region (22) extending aft of said having an upper surface (34,36,38), a lower sur- main box region (20) between said rear spar (30) and 15 face (42,44,46) a leading edge (14), a trailing edge said trailing edge (16), said airfoil including suction (16), a tip, and a root, said airfoil including front and means internal thereto for applying suction to por- rear spars (26,30) that chordwise divide said airfoil tions of said leading and trailing edge regions (18, into a leading edge region (18), a main box region 22) to remove boundary layer air flowing thereover, (20), and a trailing edge region (22), said leading characterized in that each of the upper and lower 20 edge region (18) extending between said leading surfaces (34-38,42-46) of said airfoil is contoured edge (14) and said front spar (26), said main box re- such and said suction is applied such as to provide gion (20) extending aft of said leading edge region a chordwise pressure distribution having a nega- . (18) between said front and rear spars (26,30) said tive gradient rising steeply from a positive pres- trailing edge region (22) extending aft of said main sure value at an attachment line position adjacent 25 box region (20) between said rear spar (30) and said the leading edge (14) to a peak negative value at a trailing edge (16), said airfoil including a leading peak position located immediately aft of said leading edge flap (80), having a forward edge and an aft edge (14), a positive gradient from said peak posi- edge, said leading edge flap (80) being swingably at- tion aft to said front spar (26), a negative gradient tached to said leading edge region (18) for move- from said front spar (26) aft to said rear spar (30), 30 ment between a stored position and a deployed posi- and positive pressure recoveries from said rear tion, said leading edge flap (80) forming a portion of spar (30) aft to said trailing edge (16). the lower surface (42-46) of said airfoil when said 2. The airfoil of claim 1, wherein the chordwise flap (80) is in the stored position, said leading edge pressure distribution on the upper surface region (18) including a forward slot (86) and an aft (34,36,38) of said airfoil between the front and rear 35 slot (88), said forward slot (86) being adjacent said spars (26, 30) comprises a first gradient from said forward edge, said aft slot (88) being adjacent said front spar (26) aft to a change point and a second aft edge, said airfoil including suction means inter- gradient from said change point aft to said rear nal thereto for applying suction to portions of said spar (30), said second gradient being about 50 to 75 leading and trailing edge regions (18, 22) to remove percent less than said first gradient. 40 boundary layer air flowing thereover, said suction 3. The airfoil of claim 1, wherein said change point means applying suction to said front and rear slots is located aft of said front spar (26) about two- (86, 88), said upper surface (34-38) being con- thirds of the distance between said front and rear toured such and said suction means being applied spars (26,30). such as to provide a chordwise pressure distribu- 4. The airfoil of claim 1 , wherein said leading edge 45 tion having a negative gradient rising steeply from a region (1 8) has a nose region (54), said nose region positive value at an attachment line position adja- (54) including said leading edge (14), and wherein cent the leading edge (14) to a peak negative value said suction means includes: at a peak position located immediately aft of said leading edge suction means for removing boundary leading edge (14), a positive gradient from said peak layer air flowing over said leading edge region (18), 50 position aft to the front spar (26), a negative gradi- said leading edge suction means including nose suc- ent from said front spar (26) aft to said rear spar tion means for removing boundary layer air flowing (30), and positive pressure recoveries from said over said nose region (54); and rear spar (30) aft to said trailing edge (16); charac- trailing edge suction means for removing boundary terized in that said lower surface (42-46) is con- layer air flowing over said trailing edge region (22). 55 toured such and said suction is applied such as to 5. The airfoil of claim 4, wherein each of said provide a chordwise pressure distribution having a leading edge suction means, said nose suction negative gradient rising steeply from a positive val- means, and said trailing edge suction means applies ue at an attachment line position adjacent the lead- suction in an inboard section of its respective re- ing edge (14) to a peak value at a peak position lo- gion (18,54,22), said inboard section extending 60 cated immediately aft of said leading edge (14), said spanwise from an initial position adjacent the root of leading edge peak position being located adjacent said airfoil to a terminating position spaced inboard said forward slot (86), a positive gradient from said of the tip of said airfoil. peak position aft to said aft slot (88), a steep nega- 6. The airfoil of claim 5 wherein: tive gradient from said aft slot (88), to said front said nose suction means applies suction in an in- 65 17 EP 0 222 421 B1 18

spar (26), a negative gradient from said front spar bereich (18) stromt, wobei die Vorderkantensaugein- (26) aft to said rear spar (30), and positive pres- richtung eine Nasensaugeinrichtung zum Entfernen sure recoveries from said rear spar (30) aft to said von Grenzschichtluft, die uber den Nasenbereich trailing edge (16). (54) stromt, aufweist; und 5 eine Hinterkantensaugeinrichtung zum Entfernen Patentanspriiche von Grenzschichtluft, die uber den Hinterkantenbe- reich (22) stromt. 1. Laminarstromungssteuertragflache fur einen 5. Tragfiache nach Anspruch 4, worin die Vor- gepfeilten Fiugel (10), die eine obere Oberflache derkantensaugeinrichtung, die Nasensaugeinrich- (34, 36, 38), eine untere Oberflache (42, 44, 46), 10 tung und die Hinterkantensaugeinrichtung Sog in ei- eine Vorderkante (14), eine Hinterkante (16), eine nem Innenbordabschnitt von ihrem jeweiligen Be- Spitze und eine Wurzel hat, wobei die Tragfiache ei- reich (18, 54, 22) anwendet, wobei sich der Innen- nen vorderen und ruckwartigen Holm (26, 30) auf- bordabschnitt spannweitenweise von einer weist, welche die Tragfiache sehnenweise in einen Anfangsposition benachbart der Wurzel der Trag- Vorderkantenbereich (18), einen Hauptkastenbe- 15 flache zu einer Endposition, die sich innenbords im reich (20) und einen Hinterkantenbereich (22) unter- Abstand von der Spitze der Tragfiache befindet, teilen, wobei sich der Vorderkantenbereich (18) zwi- erstreckt. schen der Vorderkante (14) und dem vorderen Holm 6. Tragfiache nach Anspruch 5, worin: (26) erstreckt, wobei sich der Hauptkastenbereich die Nasensaugeinrichtung Sog in einem Innenbord- (20) hinter dem Vorderkantenbereich (18) zwischen 20 abschnitt des Nasenbereichs (54) anwendet, der ei- dem vorderen und ruckwartigen Holm (26, 30) er- ne Endposition hat, fur welche die AnschluBlinien- streckt, wobei sich der Hinterkantenbereich (22) Reynoldszahl geringer als etwa 100 ist; und hinter dem Hauptkastenbereich (20) zwischen dem die Vorderkantensaugeinrichtung Sog in einem In- ruckwartigen Holm (30) und der Hinterkante (16) er- nenbordabschnitt des Vorderkantenbereichs (18) streckt, wobei die Tragfiache innerhalb derselben 25 anwendet, der eine Endposition hat, welche die fol- Saugmittel zum Anwenden von Sog auf Teile des genden Bedingungen erfullt: Vorder- und Hinterkantenbereichs (18, 22) zum Ent- Re (sin A) 3/2 = 5 x 1 06, fur A ;> 1 5°; und fernen von Grenzschichtluft, die dariiberstromt, Re = 38 x 1 06 fur A < 15°; aufweist, dadurch gekennzeichnet, daB jede der worin Rc= Sehnen-Reynoldszahl; und oberen und unteren Oberflachen (34 bis 38, 42 bis 30 A = der Pfeilungswinkel der Vorderkante der Trag- 46) der Tragfiache derart konturiert ist und der Sog fiache sind. derart angewandt wird, daB eine sehnenweise 7. Laminarstromungssteuertragflache fur einen Druckverteilung vorgesehen wird, die einen negati- gepfeilten Fiugel, der eine obere Oberflache (34, ven Gradienten hat, der von einem positiven Druck- 36, 38), eine untere Oberflache (42, 44, 46), eine wert an einer AnschluBlinienposition benachbart 35 Vorderkante (14), eine Hinterkante (16), eine Spitze der Vorderkante (14) steil zu einem negativen Spit- und eine Wurzel hat, wobei die Tragfiache einen zenwert an einer Spitzenposition ansteigt, die sich vorderen und ruckwartigen Holm (26, 30) aufweist, unmittelbar hinter der Vorderkante (14) befindet, ei- welche die Tragfiache sehnenweise in einen Vorder- nen positiven Gradienten von der Spitzenposition kantenbereich (18), einen Hauptkastenbereich (20) nach hinten zu dem vorderen Holm (26), einen nega- 40 und einen Hinterkantenbereich (22) unterteilen, wo- tiven Gradienten von dem vorderen Holm (26) nach bei sich der Vorderkantenbereich (1 8) zwischen der hinten zu dem ruckwartigen Holm (30), und positive Vorderkante (14) und dem vorderen Holm (26) er- Druckwiederhersteilungen von dem ruckwartigen streckt, wobei sich der Hauptkastenbereich (20) Holm (30) nach hinten zu der Hinterkante (16): hinter dem Vorderkantenbereich (18) zwischen dem 2. Tragfiache nach Anspruch 1 , worin die sehnen- 45 vorderen und riickwartigen Holm (26, 30) erstreckt, weise Druckverteilung auf der oberen Oberflache wobei sich der Hinterkantenbereich (22) hinter dem (34, 36, 38) der Tragfiache zwischen dem vorderen Hauptkastenbereich (20) zwischen dem ruckwarti- und ruckwartigen Holm (26, 30) einen ersten Gradi- gen Holm (30) und der Hinterkante (16) erstreckt, enten von dem vorderen Holm (26) nach hinten zu wobei die Tragfiache eine Vorderkantenklappe (80) einem Umschlagpunkt und einen zweiten Gradienten 50 aufweist, die eine vorwartige Kante und eine hinter- von dem Umschlagpunkt nach hinten zu dem ruck- wartige Kante hat, wobei die Vorderkantenkiappe wartigen Holm (30) umfaBt, wobei der zweite Gradi- (80) zur Bewegung zwischen einer untergebrachten ent etwa 50 bis 75 % weniger als der erste Gradient Position und einer entfalteten Position verschwenk- ist. bar an dem Vorderkantenbereich (18) angebracht 3. Tragfiache nach Anspruch 1, worin sich der 55 ist, wobei die Vorderkantenklappe (80) einen Teil Umschlagpunkt etwa zwei Drittel der Entfemung der unteren Oberflache (42 bis 46) der Tragfiache zwischen dem vorderen und ruckwartigen Holm (26, bildet, wenn diese Klappe (80) in der untergebrach- 30) hinter dem vorderen Holm (26) befindet. ten Position ist, wobei der Vorderkantenbereich (1 8) 4. Tragfiache nach Anspruch 1, worin der Vor- einen vorderen Schlitz (86) und einen hinteren derkantenbereich (18) einen Nasenbereich (54) hat, 60 Schlitz (88) aufweist, wobei sich der vordere wobei der Nasenbereich (54) die Vorderkante (14) Schlitz (86) benachbart der vorwartigen Kante be- umfaBt, und worin die Saugmittel folgendes umfas- findet, wobei sich der hintere Schlitz (88) benach- sen: bart der hinterwartigen Kante befindet, wobei die eine Vorderkantensaugeinrichtung zum Entfernen Tragfiache innerhalb derselben eine Saugeinrich- von Grenzschichtluft, die uber den Vorderkanten- 65 tung zum Anwenden von Sog auf Teile des Vorder-

10 19 EP 0 222 421 B1 20 und Hinterkantenbereichs (18, 22) zum Entfernen caracterise en ce que chacune des surfaces supe- von Grenzschichtluft, die daruberstromt, aufweist, rieure et inferieure (34-38, 42-46) du profil aerody- wobei die Saugeinrichtung Sog auf den vorderen namique a un profil tel et I'aspiration est appliqude und hinteren Schlitz (86, 88) anwendet, wobei die d'une maniere telle qu'une distribution de pression obere Oberflache (34 bis 38) derart konturiert ist 5 est realisee suivant la corde avec un gradient nega- und die Saugeinrichtung derart angewandt wird, daB tif qui augmente fortement d'une valeur positive de eine sehnenweise Druckverteilung vorgesehen la pression, a I'emplacement d'un axe de fixation ad- wird, die einen negativen Gradienten hat, der von jacent au bord d'attaque (14), a une valeur negative einem positiven Wert an einer AnschluBlinienpositi- de crete a une position de crete piacee juste en ar- on benachbart der Vorderkante (14) steil zu einem 10 riere du bord d'attaque (14), un gradient positif de la negativen Spitzenwert an einer Spitzenposition an- position de crete vers le longeron avant (26) qui se steigt, die sich unmittelbar hinter der Vorderkante trouve vers I'arriere, un gradient negatif a partir du (14) befindet, einen positiven Gradienten von der longeron avant (26) vers le longeron arriere (30) Spitzenposition nach hinten zu dem vorderen Holm vers I'arriere, et un retablissement a une pression (26), einen negativen Gradienten von dem vorderen 15 positive a partir du longeron arriere (30) vers le Holm (26) nach hinten zu dem ruckwartigen Holm bord de fuite (16) vers I'arriere. (30), und positive Druckwiederherstellungen von 2. Profil aerodynamique selon la revendication 1, dem ruckwartigen Holm (30) nach hinten zu der Hin- dans lequel la distribution des pressions suivant la terkante (16); dadurch gekennzeichnet, daB die un- corde, a la surface superieure (34, 36, 38) du profil tere Oberflache (42 bis 46) derart konturiert ist 20 aerodynamique entre les longerons avant et arriere und der Sog derart angewandt wird, daB eine seh- (26, 30), comporte un premier gradient entre le lon- nenweise Druckverteilung vorgesehen wird, die ei- geron avant (26) et, vers I'arriere, un point de chan- nen negativen Gradienten hat, der von einem positi- gement, et un second gradient a partir du point de ven Wert an einer AnschluBlinienposition benach- changement, vers I'arriere, jusqu'au longeron arrie- bart der Vorderkante (14) steil zu einem Spitzen- 25 re (30), le second gradient etant inferieur d'environ wert an einer Spitzenposition ansteigt, die sich 50 a 75 % au premier gradient. unmittelbar hinter der Vorderkante (14) befindet, 3. Profit aerodynamique selon la revendication 1, wobei sich die Vorderkanten-Spitzenposition be- dans lequel le point de changement se trouve en ar- nachbart dem vorderen Schlitz (86) befindet, einen riere du longeron avant (26) aux deux tiers environ positiven Gradienten von der Spitzenposition nach 30 de la distance comprise entre les longerons avant hinten zu dem hinteren Schlitz (88), einen steilen ne- et arriere (26, 30). gativen Gradienten von dem hinteren Schlitz (88) zu 4. Profil aerodynamique selon la revendication 1, dem vorderen Holm (26), einen negativen Gradien- dans lequel la region de bord d'attaque (1 8) a une re- ten von dem vorderen Holm (26) nach hinten zu dem gion de nez (54), la region de nez (54) comportant le ruckwartigen Holm (30), und positive Druckwieder- 35 bord d'attaque (14), et dans lequel le dispositif d'as- herstellungen von dem ruckwartigen Holm (30) nach piration comporte : hinten zu der Hinterkante (1 6). un dispositif d'aspiration de bord d'attaque destine a retirer I'air de la couche limite s'ecoulant sur la re- gion de bord d'attaque (18), le dispositif d'aspiration 40 de bord d'attaque ayant un dispositif d'aspiration de Revendications nez destine a retirer I'air de la couche limite s'ecou- lant sur la region de nez (54), et 1 . Profil aerodynamique a reglage d'ecoulement la- un dispositif d'aspiration de bord de fuite destine a minaire, destine a une aile en fleche (10), ayant une retirer I'air de la couche limite s'ecoulant sur la re- surface superieure (34, 36, 38), une surface infe- 45 gion de bord de fuite (22). rieure (42, 44, 46), un bord d'attaque (14), un bord 5. Profil aerodynamique selon la revendication 4, de fuite (16), une extremite externe et un pied, le pro- dans lequel chacun des dispositifs d'aspiration de fil aerodynamique comprenant des longerons avant bord d'attaque, de nez et de bord de fuite applique et arriere (26, 30) qui divisent le profil aerodynami- une aspiration a un trongon interne de sa region que, dans la direction de la corde, en une region de 50 respective (18, 54, 22), ce trongon interne etant dis- bord d'attaque (18), une region principale en forme pose, suivant I'envergure, d'un emplacement initial de caisson (20) et une region de bord de fuite (22), adjacent au pied du profil aerodynamique jusqu'a la region de bord d'attaque (18) etant disposee entre une position terminate distante de I'extremite externe le bord d'attaque (14) et le longeron avant (26), la re- du profil aerodynamique. gion principale en forme de caisson (20) etant dispo- 55 6. Profil aerodynamique selon la revendication 5, see en arriere de la region de bord d'attaque (18) en- dans lequel : tre les longerons avant et arriere (26, 30), la region le dispositif d'aspiration de nez applique une aspira- de bord de fuite (22) etant piacee en arriere de la re- tion a un trongon interne de la region de nez (54) gion principale en forme de caisson (20) entre le lon- ayant une position terminale pour laquelle le nombre geron arriere (30) et le bord de fuite (16), le profil 60 de Reynolds a I'axe de fixation est inferieur a 1 00 en- aerodynamique comprenant un dispositif d'aspira- viron, et tion place a I'interieur et destine a exercer une aspi- le dispositif d'aspiration de bord d'attaque applique ration dans des parties des regions de bord d'atta- une aspiration au trongon interne de la region de que et de bord de fuite (18, 22) afin que de I'air de la bord d'attaque (18) qui a une position terminale qui couche limite s'ecoulant sur ces regions soit elimine, 65 correspond aux conditions suivantes :

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Rc (sinA)3/2= 5 • 1 06, pour A ;> 1 5°, et riere (88) vers I'arriere, un gradient negatif intense Rc= 38 • 106 pour A< 15°, de la fente arriere (88) au longeron avant (26), un Rc etant le nombre de Reynolds a la corde et A la gradient negatif du longeron avant (26) au iongeron fleche du bord d'attaque du profil aerodynamique. arriere (30) vers I'arriere, et un retablissement de 7. Profil aerodynamique a reglage d'ecoulement pression positive du longeron arriere (30) au bord laminaire, destine a une aile en fleche, ayant une de fuite (16) vers I'arriere. surface superieure (34, 36, 38), une surface infe- rieure (42, 44, 46), un bord d'attaque (14), un bord de fuite (16), une extremite externe et un pied, le pro- fil aerodynamique comprenant des longerons avant 10 et arriere (26, 30) qui divisent le profil aerodynami- que, dans la direction de la corde, en une region de bord d'attaque (18), une region principale en forme de caisson (20) et une region de bord de fuite (22), la region de bord d'attaque (1 8) etant disposee entre 15 le bord d'attaque (14) et le longeron avant (26), la re- gion principaie en forme de caisson (20) etant dispo- see en arriere de la region de bord d'attaque (18) en- tre les longerons avant et arriere (26, 30), la region de bord de fuite (22) etant disposee en arriere de la 20 region principale en forme de caisson (20) entre le longeron arriere (30) et le bord de fuite (1 6), le profil aerodynamique comprenant un volet de bord d'atta- que (80) qui a un bord avant et un bord arriere, le volet de bord d'attaque (80) etant monte afin qu'il 25 puisse pivoter sur la region de bord d'attaque (18) et se deplace entre une position rangee et une position deployee, le volet de bord d'attaque (80) constituant une partie de la surface inferieure (42, 46) du profil aerodynamique lorsque le volet (80) est dans la po- 30 sition rangee, la region (1 8) de bord d'attaque com- prenant une fente avant (86) et une fente arriere (88), la fente avant (86) etant adjacente au bord avant, la fente arriere (88) etant adjacente au bord arriere, le profil aerodynamique comprenant un dis- 35 positif d'aspiration place a I'interieur et destine a as- surer une aspiration dans des parties des regions de bord d'attaque et de fuite (1 8, 22) de maniere que de I'air de la couche limite qui circule sur ces re- gions soit retire, le dispositif d'aspiration exergant 40 une aspiration par les fentes avant et arriere (86, 88), la surface superieure (34-38) ayant un profil tel et le dispositif d'aspiration fonctionnant d'une maniere telle que la distribution des pressions sui- vant la corde a un gradient negatif augmentant for- 45 tement d'une valeur positive, a la position d'un axe de fixation adjacent au bord d'attaque (14), a une va- leur negative de crete a une position de crete pla- cee juste en arriere du bord d'attaque (14), un gra- dient negatif allant de la position de crete vers le 50 longeron arriere (26), vers I'arriere, un gradient ne- gatif du longeron avant (26) au longeron arriere (30) vers I'arriere, et un retablissement de pression positive du longeron arriere (30) au bord de fuite (16) vers I'arriere, caracterise en ce que : 55 la surface inferieure (42-46) a un profil tel et I'aspi- ration est exercee d'une maniere telle que la distri- bution de pression suivant la corde a un gradient negatif qui augmente fortement d'une valeur positi- ve a un emplacement d'un axe de fixation adjacent 60 au bord d'attaque (14), a une valeur de crete a un emplacement de crete qui se trouve juste en arriere du bord d'attaque (14), la position de crete du bord d'attaque etant adjacente a la fente avant (86), un gradient positif de la position de crete a la fente ar- 65

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