1、r -_ ., - I i/ ,- ,: _ , / .- t _ . .- _-, _ -/ _ , x _ . . _ - :, .I / _. -/ . , , : : :-_ _ ). . . . /. , _ _ - _. Government I?tin!hg Offibe. varks nwording to size ,- - - , ,- . ., D. C. Yearl$ snbscription; $10; foreign. 45 cents - Provided by IHSNot for ResaleNo reproduction or networking perm
2、itted without license from IHS-,-,-REPORT 1218 EFFECT OF GROUND INTERFERENCE ON THE AERODYNAMIC AND FLOW CHARACTERISTICS OF A 42” SWEPTBACK WING AT REYNOLDS NUMBERS BJP TO 6.8 x 10” By G. CHESTER FURLONG and THOMAS V. BOLLECH Langley Aeronautical Laboratory Lmgley Field, Va. I I Ill II III Provided
3、by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-National Advisory Committee for Aeronautics Headquarters, 1512 H Street NW., Washington 2ti, D. C. Created by act of Congress approved March 3, 1915, for the supervision and direction of t,he scientific study of
4、 the problems of flight (U. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2, 1929, and to 17 by act approved May 25, 1948. The members are appointed by the President, and serve as such without compensation. JEROME C. HUNSARER, SC. D., Massachusetts In
5、stitute of Technology, Chairman LEOX., General, United States Air Force, Chief of Staff. HUGH L. DRYDEN, PH. D., Director JOIIN TV. CROMLIY, JR., B. S., Associate Director for Research JOHN F. VICTORY, LL. D., Execzctil,e Secretary EDWARD H. CHAMBERLIN, Emxutie 0,ficer HENRY J. E. REID, D. Eng., Dir
6、ector, Langley Aeronautical Laboratory, Langley Field, Va. SMITH J. DEFRAKCE, D. Eng., Director, Ames Aeronautical Labora.torr, 3Ioffett Field, Calif. EDWARD R. SHARP, SC. D., Director, Lewis Flight Propulsion Laboratory, Cleveland, Ohio WALTER C. WILLIAMS, B. S., Chief, High-Speed Flight Station, E
7、dwards, Calif. II ._. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORT 1218 EFFECTOFGROUNDINTERFERENCEONTHEAERODYNAMICAND FLOW CHARACTERISTICS -OF A 420 SWEPTBACK WING AT REYNOLDS NUMBERS up TO 6.8xiot Rp G. CHESTER FURLOS: and THOMM V. BOLLE
8、CH SUMMARY The eflects of ground interference on the aerodynamic charac- teristics of a 42” sweptback u namely, the plain wing and the wing with inboard trailing- edge split flaps and outboard leading-edge flaps deflected. Certain aspects of the effects of the ground interference on the aerodynamic
9、characteristics of unswept wings have been thoroughly investigated both theoretically and experi- mentallv (refs. 1 to 6). The experimental results of these The present report contains force and moment data obtained throughout the angle-of-attack range at several values of Reynolds number and contou
10、r charts of downwash, sidewash, and dynamic-pressure ratio at two longitudinal tCombination of the recently deolsssitlod NACA RM LSQ22, “Downwash, Sidewash, and Wake Surveys Behind a 42 Sweptback Wing at B Reynolds Number of 6.8X10 0 With and Without a Simulated Ground” by 0. Chester Furlong and Tho
11、mns V. Bolleeh, 1948 and NACA TN 2487, “Effect of Ground Interference on the Aerodynamic Characteristics of a 42 Swept- back Wing” by ct. Chester Furlong and Thomss V. Bollech, 1951. 1 I -. _.: . . . +a* . -. .- -_ - -. Provided by IHSNot for ResaleNo reproduction or networking permitted without lic
12、ense from IHS-,-,- - 2 REPORT 12 1 g-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS stations behind the wing (2.0 and 2.8 mean aerodynamic chords). The locations of the tip vortices have been shown on the contour charts of dynamic-pressure ratio for the plain wing without the ground pressnt. Integratio
13、ns have been made to obtain variations of average downwash and dynamic pressure with angle of attack. Values of downwash have been calculated by extending the method present.ed in ref- erences 7 ancl 8 to account for the sweep of the 0.25-chord line. The ground was sim.ulated in the tunnel by m.eans
14、 of a ground board. Although this method of ground represen- tation is not ideal, the results of the present tests are believed to be indicative of t,hc ground-interference effects on a swcp tbnck wing. SYMBOLS lift coefficient l* PS Drag drag coefficient, qs pitching-moment coeficicnt about O.25c,
15、Pitching moment qsc angle of attack of wiiig root chord, deg free-stream dynamic pressure, q, lb/sq ft pvc Reynolds number, P wing area, sq ft wing span, ft local chord, ft mean scrod)-namic chortl, z s 612 so 2 dy, ft mass tlclnsity of air, slugs/cW ft stream velocity, ft/scc local stream d,vnamic
16、pressure, lb/sq ft local downwash angle, deg smcep angle of 0.25-chord line, deg siclrwash angle, inflow positive, deg coefficient of visc0sit.v of air, slugs/ft-set ratio of local-stream dynamic prcssurc to free- stream dynamic pressure vertical distance from chord plane cstcntlcd, ft, longitutlina
17、l distance from 0.25-chord point of root chord vortex scmispan (always positive), ft lateral distance from plane of s-mnictrj-, ft downwash factor total induced downward velocity, ft/sec section lift coefficient vortex strength calculated downwash a.ngle, cleg downward displacement, measured normal
18、to the relative wind, of the center line of t,he wake and the trailing vortex sheet from its origin at the trailing edge, ft Integrated air-stream surveys: !a*/dao average qJq, obtained by Go average e, obtained by where chord of fictitious tail span of fict.itious tail area of fictitious tail spanw
19、ise distance rate of change of cat, with angle of attack GROUND, MODEL, AND APPARATUS GROUND REPRESENTATION AND GROUND DISTANCE Several methods such as the reflection method, the partial plate and reflection method, and the plate method are avail- able for ground simulation in a wind tunnel (refs. 4
20、 to 6). The most feasible arrangement for ground tests in the Langley 19-foot pressure tunnel is the plate method (com- monly referred to as the ground-board method). The vertical distance from the 0.25C to the ground boarcl (regardIess of boundary-layer thickness on the ground board) is referred to
21、 as the ground distance. Inasmuch as no standard point of reference exists, the 0.25i has been used because it. was the most convenient point of reference from considerations of test procedure. The model was supported in the tunnel at the 0.25Z, and to maintain a constant ground distance for any oth
22、er point of reference would have necek tntecl moving the ground board as the angle of attack of t,ht! wing was changed. Based on tbc preceding tlcfinition of ground distance, tbc ground distances used in the present tests were 0.68Z and 0.92;. MODEL The model mounted on the Ilormal wing-support syst
23、em of the Langley 19-foot pressure tunnel is shown in figure 1. The wing had 42 sweepback of the leading edge, a tape1 ratio of 0.625, an aspect ratio of 4.01, and NACA G41-112 airfoi1 sections normal to the 0.27khord line. The principal tlimcnsions of the motlcl and flaps arc given in figure 2. It,
24、 was found that a slight discontinuity existed along the 0.20- chord line of the wing. Tl lc results obtained in the present tests, therefore, do not necessarily represent. exactly those which would bc obtainccl on a wing with true NACA 641-1 12 airfoil sections. The model was maintained in a, smoot
25、h condition during the tests. For tests with flaps deflected, the 0.20 t,railing-edge split flaps were deflected 60” from t.he lower surface and extended from the root to 0.50%. The leading-edge flaps extended spanwise from 0.400: to 0.975%. Provided by IHSNot for ResaleNo reproduction or networking
26、 permitted without license from IHS-,-,-GROUND INTERFERENCE EFFECTS ON SWEPTBACK WINGS 3 (b) Rear view. FIGURE l.-Concluded. (a) Front view. FIGUU 1.-A 42O sweptback wing mounted in the Langley 19-foot pressure tunnel. Flaps deflected; ground board in. Ground distance, 0.92?. 64,-I I2 sections -34.1
27、25 : -Flap loins upper surface : approximately l/2 inch behind L.E -:50 diameter FIGJRE 2.-Layout of 42O sweptback wing. (All dimensions are in inches.) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 - REPORT 12 1 S-NATIONAL ADVISORY COMMITTEE FOR
28、 AERONAVTICS APPARATUS The aerodynamic forces were measured by a simultane- ously recording, six-component balance system. Survey apparatus.-The Langley 19-foot-pressure-tunnel survey apparatus and multiple-tube survey rake (fig. 3) were used to obtain downwash and dynamic pressure behind the wing.
29、The multiple-tube survey rake consists of six pitot- static tubes with pitch and yaw orifices in the hemispherical tips. The survey apparatus maintained the rake in a verti- cal position as it was moved laterally along the span. This survey rake had been previously calibrated through known pitch and
30、 yaw angles. All pressure leads were conducted to a multiple-tube manometer and during the tests the data were photographically recorded. A probe containing three tufts spaced 1.5 inches was used to locate the tip vortex. The probe was attached to the survey strut. Ground board.-The ground board con
31、sisted of a steel framework covered with plywood on both the upper ancl lower surfaces, which resulted in an overall thickness of 4 inches. (See fig. 4.) A 1 t s o extending the full width of the grouncl board and. located 1 foot in front of the 0.25C of the wing was provided as a means of boundary-
32、layer control. The ground board was supported in the tunnel test section by means of wall brackets and center posts. (See figs. 1 and 4.) The support system allowed a ground-board travel from 16.0 to 31.9 inches below the center line of the tunnel (center of rotation of the model). - r ldary- slot (
33、Center of rotation) Static orifice I “Yaw orifice /Y I (a) Photograph of survey rake. (b) Sketch of survey-rake tube. 1 F IGURE 3.-Langley 14foot pressure tunnel airstream survey rake. 0-l Direction -of oir flow B -2 View B-B “.3lot flap Section A-A FIGYRE 4.-Sk b.ence the usual moclel blockage corr
34、ection has been applied to the dynamic-pressure measurements. The ground board reduced the tunnel-clear stream angle approximately 0.15. Visual tuft studies of the flow on the ground board with the boundary-layer slot closed and open were macle through the angle-of-attack range of the model. When th
35、e slot was closrtl but not completely sealed, an unsteady flow condition existed along the nose of the slot. The flow condition at the nose of the slot was improved when the slot was open. An unsteady flow condition existed in an area near the center of the board between 2.OC and 2.8C with either th
36、e slot open or clostd. This unsteady flow conclition can be attributed to the diffusion of the flap wake. There was no indication of actual flow separation on the board throughout the angle-of- attack range of the moclel. By use of the boundary-layer- control slot the maximum thickness of the bounda
37、ry layer was reduced from approximately 1 .O inch to 0.4 inch beneath tbe wing and from 1.6 inches to 1.0 inch at a distance 2.8; rearward of the 0.257. The flow through the slot was not, matcrin.lly affected by the presence of the model. The dis- continuity in bounclary-layer thickness due to the f
38、low through the slot corresponcls to an effective discontinuity in ground distance, which, however, is believed to have a negli- gible effect on the test results. Presence of a boundary layer on the ground boarcl may be less troublesome under a swept- back wing than under an unswept wing, mainly bec
39、ause the maximum lift is considerably lower for the sweptback wing. Force and moment tests.-Force ancl moment data were obtained for the two model configurations through an angle- of-attack rnnge from -4 t,hrough the stall. The tests were made with the grouncl boarcl out and with the ground board lo
40、cated at ground distances of 0.6% and 0.92-E for several values of Reynolds number. The Reynolds numbers of the test,s based on C were 3.0, 4.3, 5.2, ancl 6.8 X 106. Airstream surveys.-Downwash, sidewash, and dynamic- pressure surveys were macle for each model and ground- board configuration at two
41、longitudinal stations. The posi- tions for t.he survey apparatus were selected so that they npprosim.ated, through the angle-of-attack range of the tests, stations 2.OC and 2.G behind the 0.25C of the wing measurecl along the chord plane extended. The maximum variation of the stations 2.OC and 2% fr
42、om the locations of the survey apparatus was only 0.5 inch through the angle-of-attack range of the test. Due to the fact that the trailing edge of the wing was swept back, the distance between the survey rake and tb.e trailing edge of the wing decreased as the rake was moved from the plane of symme
43、try. Data were ob- tained at three angles of attack for the wing with flaps neu- tral and at four angles of attack for the flapped wing. The angles of attack for the tests in the presence of the grouncl were selected-so that the values of lift-_coef -_- 1.36Cr. 1.53cr, Ae 7x - m With the ground boar
44、d in the tunnel test section, it was not possible to obtain corrections for support-tare and st.rut interference. The ground-board-out corrections for support- tare and strut interference, however, have been applied to the ground-board-in clata in the belief that they would be of the same nature, al
45、though not necessarily of the same magnitude, as would be obtained with the ground board in. Calculations made for other ground investigations (such as ref. 4) have shown that at small ground heights, jet-boundary corrections are negligible; hence, they have been neglect,ecl in the present tests. EF
46、FECTS OF GROUND INTERFERENCE A discussion of the concepts of ground interference appears pertinent before the results of the present tests of a swept- back wing are presented. Although the concepts have been derived largely to explain the effects of grouncl interference on an unswept wing, they shou
47、ld, in general, apply to a swept- back wing as well. The ground effect on a wing may be considered as the interference due to the reflected image of the wing in the ground. Computations of the effects of the image wing on the real wing can be made by replacing it with a bound vortex and a system of
48、trailing vortices. Inasmuch as these computations are based on thin-wing theory, the effect of the thickness of the image wing must also be determined. The separate effects of the bound vortex, trailing vortices, Provided by IHSNot for ResaleNo reproduction or networking permitted without license fr
49、om IHS-,-,-6 REPORT 12 1 S-NATIONAL ADVISORY COM3IIITEE FOR AERONAUTICS and wing thickness can then be added. In reference 1 the interference from the trailing vortices of the image wing was considered in detail; whereas in reference 6 the interferences from the bound vortex and wing thickness of the image wing were also c
