NASA NACA-RM-L51A23-1951 Flight determination of the effects of wing vortex generators of the aerodynamic characteristics of the Douglas D-558-I airplane《机翼旋涡发生器对道格拉斯D-558-I飞机空气动力特.pdf

《NASA NACA-RM-L51A23-1951 Flight determination of the effects of wing vortex generators of the aerodynamic characteristics of the Douglas D-558-I airplane《机翼旋涡发生器对道格拉斯D-558-I飞机空气动力特.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-RM-L51A23-1951 Flight determination of the effects of wing vortex generators of the aerodynamic characteristics of the Douglas D-558-I airplane《机翼旋涡发生器对道格拉斯D-558-I飞机空气动力特.pdf（24页珍藏版）》请在麦多课文档分享上搜索。

1、h .I n RESEARCH MEMORANDUM FLIGHT DETERMIXATION OF THE EFFECTS OF WING VORTEX GENERATORS OM THIYAERODYNAMIC CHARACTERISTTCS OF THE DOUGLAS D-558-1 AIRPLANE By De E. Beeler, Donald R. Bellmm, and John R. Griffith Langley Aeronautical Laboratory Langley Field, Va. “ “ “_“ “ “ -“- NATIONAL ADVISORY COM

2、MITTEE FOR AERONAUTICS WASHINGTON . August 14, 1951 I .1 . “ - . I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 -9 NACA FN L5lA23 NATIONAL ADVISORY comm FOR AERONAUTICS FLIGHT DETERMINATION OF THE EFFECTS OF WING VORTEX GENERATORS ON THE AERODYN

3、AMIC CHARACTERISTICS OF TBE DOUGLAS D-558-1 AlRPLANE By De E. Beeler, Donald R. Bellman, and John H. Griffith Tests were made of the Douglas D-558-1 airplane to determine the effect of wing vortex generators on some of the typical undesirable handling characteristics such as buffeting, lateral unste

4、adiness, change in trim, and loss in control effectiveness which have occurred on present-day aircraft flying at supercritical speeds. An arbitrary size, shape, and location of the generators were selected and the investiga- tion was initiated to determine what benefits might be derived from the ins

5、tallation and whether further investigations would be warranted. The use of the vortex generators resulted in a reduction of sepa- rated regions over the wing at Mach numbers greater than 0.85 for level flight. At higher normrtl-form coefficients for Mach numbers greater than 0.85, regions of separa

6、tion and forward movement of the shock were reduced. The buffet boundary and wing-dropping tendency were delayed approximately 0.05 in Mach number for level flight; however, no change in the small-amplitude oscillation was detected. The pilot reported the buffeting to be appreciably reduced during f

7、lights penetrating the buffet region for lift coefficients below the stall. No detrimental effects on the 1ongitudTnal and lateral control characteristics were encountered for the flight conditions investigated. The drag of the airplane was increased by the use ofthe generators. t I ! Present-day ai

8、rplanes flying at supercritical speeds have demon- strated the effects of compressibility in the form of buffeting, lateral unsteadiness, change in trim, and loss. in contrpl effectiveness. In the z course of flight testing the research.airplanes, modifications to the CQp-Ws. . Provided by IHSNot fo

9、r ResaleNo reproduction or networking permitted without license from IHS-,-,-2 - MACA RM I311123 specific configurations of these airplanes have been considered for the purpose of reducing the magnitudes and/or delaying the effects of com- pressibility which are objectionable from the handling and t

10、he structural standpoint. Based on results of reference 1, an investigation was made at the mACA High-speed Flight Research Station at Edwards Air Force Base, Calif., to determine the effect of wing vortex generators on the handling and buffeting characteristics of the Douglas D-558-1 research airpl

11、ane. The D-558-1 is capable of penetrating the- buffet region to some degree and experiences a lateral unsteadiness in level flight, an abrupt lateral trim change (reference 2), and a loss in control effectiveness. This paper presents the results of flight t-ests to determine the effect of the vorte

12、x generators on some of the forementioned undesirable ckacteristics of the airplane. cNA CD +/4 F/4 D Fa M M.A.C. P mT S V W airplane nom-force coefficient (nW/qS) drag coefficient (D/qS ) pitching-moment coefficient about- wing quarter chord? (stall moment is positive) one-quarter of M.A.C., feet t

13、otal airplane drag, pounds aileron force, pounds Mach number mean aerodynamic chord pressure coefficient r1 ; loss in total head, free-stream impact preesure minus local hpac t pre S sure wing area, square feet true velocity, feet per second weight of airplane, pounds - Provided by IHSNot for Resale

14、No reproduction or networking permitted without license from IHS-,-,-. NACA RM L5LA23 b wing spari, feet CW wing chord, feet Y/C ratio of height above mean chord line to section chord it stabilizer incidence, degrees n normal acceleration P rolling velocity, radians per second PO free-stream static

15、pressure, pounds per scpaze foot P1 9 free-strem aynamic pressure, pounds per square foot local static pressure, pounds per square foot a Ee elevator deflection, degrees aileron deflection, degrees 3 The Douglas D-558-1 airplane used for these tests is a single-place low- monoplane powered by a Gene

16、ral Electric TG-180 turbojet engine. The center of gravity was located at 23.34 percent of the mean aero- aynamic chord and the gross weight at take-off was 10,610 pounds. Table I lists the more pertinent physical characterbtics of the air- plane and figures 1 and 2 aze a three-vlew sketch ad a phot

17、ograph of the airplane, respectively. I Vortex Generators The vortex generators were made from small airfoils having a sec- tion of NACA 0012 and a chord of 0.5 inch. The enerators were mounted perpendicular to the wing surface and projected 0.5 inch above the sur- face with centers at 2-inch eterva

18、ls. The mean chord line of each generator was inclined at 13 to the normal to the 30-percent chord toward the fuselage and the remafning ones were inclined away from the fuselage. The wing surface was smooth and the same section profile as existed with the generators removed was retained; the only p

19、rotuberances, line of the wing. (See fig. 3.) fiternate generators were inc1-d i Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 - NACA RM L5U3 as may be seen from the photographs of figures 4 and 5, were the individual generators. Instrumentation

20、Standard mACA recording instruments were used to measure altitude, airspeed, accelerations, rolling velocity, control-surface positions, and control forces. Chordwise surface pressures. over the upper and lower surfaces of one span station of the right wing (fig. 3) were measured with an NACA record

21、ing manometer. Procedure The firstflighMas made with the vortex generators extending spmwise from approximately the.midflap to themidaileron station. A lift-off and landing was made followed by a flight to altitude to deter- mine, primarily, the low-speed handling characteristics of the airplane wit

22、h the generators installed. The second flight was made with gener- ators extending from the midflap span station to the tip. The third flight was made with generators extendiw across the complete span. The flight8 at altitude consisted, in general, of low-speed stalls, pull-ups through the buffet re

23、gion at Mach numbers up to approxi- mately 0.89, and abrupt aileron rolls at Mach numbers above 0.70. The flight conditions were repeated with generators removed. The data presented herein are for the full-span configuration except that the lateral-control effectiveness data are for the con- figurat

24、ion of the second flight and the drag data are for all configurations. IIESUGTS AND DISCUSSION The effect of wing vortex generators of one specific size, type, and location on the pressure distribution is presented and discussed. This information is followed by a presentation and discussion of the e

25、ffect of wing vortex generators on the over-all airplane measurements of buffeting, lateral unsteadiness, trim Changes, later-sp region I1 shows flight conditions where buffeting occurred for the basic configuration, but where no occurrence of buffeting with generators installed is evident; region I

26、11 shows flight conditions where buffeting occurred with generators installed. No measurements were made to determine the effect of generators on the buffeting intensities. It is expected, however, that reductions in separated regions by use of the generators would result in a reduction in buffeting

27、 intensities. The pilot reported that, during flights made into the buffet region below maxirmun normal-force coefficients, the buffeting intensities were appreciably reduced with the generators installed. No detrimental effects of the generators on the control characteristics of the airplane were r

28、eported by the pilot. Lateral unsteadiness and trim change.- It is difficult in flying the D-558-1 airplane to select the proper lateral trim to maintain precisely a “wings-level“ flight condition and pilots have described the occurrence as attempting to balance the airplane on a pivot point. As the

29、 Mach number approaches a value of- about 0.84 however, the abrupt left roll has been delayed to a Mach number of about 0.89 when the airplane normal-force coeffi- cient is equal to 0.35. It is indicated that the difficulty of main- tawing wings-level flight after the abrupt roll atill date, as dld

30、the difficulty fn the case with the basic configuration. Lateral-control effectiveness.- A loss in lateral-control effective- ness for the D-558-1 airplane without generators has been Fndlcated at Mach numbers greater than approximately 0.86 as shown in figure 11. Data were obtained with the vortex

31、generators installed fram the midflap span station to the tig station of the wing, but only to a Mach number of 0.84 and, as might be elrpected from an inspectfon of the pressure- distribution measurements, no beneficial effects from the generators result at these Mach numbers. If it is assuIlled th

32、at the loss in lateral- control effectiveness results from the development of a region of sepa- rated flow over the aileron, it is indicated from the pressure distribu- tions (fig. 6) that for Mach numbers greater than 0.88 the generators might reduce the loss in effectiveness. Longitudinal trim cha

33、nges.- The variation of the wing-section pitching moment with Mach llllILiber for the basic am3 full-span generator configurations, 8s determined by integration of the chordwfse pressure distributions, is shown in figure 12. Also fncluded in figire 12 is the variation of the elevator angle required

34、for trim for a given sta- bilizer incidence. The data are for a normal-force coefficient corre- sponding to 1 g flight. It is indicated that no large changes in me wing-section pitching moment occurred because of the installation of the generators, and the longitudinal trim data indicate that no app

35、reciable changes in the downwash resulted from the fnstallation. Drag.- The comparison of the measured airplane total drag variation with.Mach rimer far all configurations at a normal-force coefflcient of approximately 9.25 is presented in figure 13. It is evident from fig- ure 13 that the use of th

36、e generators resulted in an increase in drag. Because of the scatter of the data, however, little can be concluded regarding the effect of spanwise location of the generators on the drag. I i . I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 CONC

37、LUDING REMARKS NACA RM L5lA23 The results indicated that vortex generators of a specific size and type located along the 30-percent wing chord line had the following effects on the handling and buffeting characteristics of the D-59-1 airplane: 1. There was a reduction of separated regions over the w

38、ing at Mach numbers greater than 0.85 for level-flight normal-force coefficients. At higher normal-force coefficients andMach numbers greater than 0.85, regions of separation and forward movement of the shock were reduced. 2. The buffet bocndary and wing-dropping tendency were delayed by approximate

39、ly 0.05 in Mach number at level-flight normal-force coeffi- cients = 0.25); however, no changes in the mall-amplitude lateral oscillations could be detected. 3. Buffeting intensities, as reported by the pilot, were appreciably reduced during flight penetrating the buffet region for lift coefficients

40、 below the stall. 4. No detrimental effects on the longitudinal or lateral control were encountered for the conditions investigated. 5. The drag of the airplane was increased. Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. 1. Donaldson, Coleman duP.: I

41、nvestigation of a Simple Device for Prevent- Separation Due to Shock and Boundary-Layer Interaction. NACA RM 5002a, 1950. 2. Bazlow, William H., and Lilly, Howard C.: Stability Results Obtained with Douglas D-558-1 Airplane (BuAero No. 37971) in Flight up to a Mach Number of 0.89. NACA RM La03 , 194

42、8. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-9 TABLEI PHYSICAL CHARACTERISTICS OF DOUGLAS D-558-1 AJRPUNE Wing: Area. sq ft . Spaqft . Taper ratio . Aspect ratio . Root section . NACA Tip section WA Sweepback of 50-percent-chord line Geometri

43、c dihedral, deg Incidence at root chord, deg . Geometric twist Mean aerodynamic chord, ft . 150-7 25 65-n0 65-11-0 0 .4.0 .2.0 0 .6.u 0.54 4.17 Ailerons : Area aft of hinge line (both ailerons), sq ft . 7.94 Mean aerodynamic chord. ft . 0.772 span (one.side), ft . - . 5.19 Hinge-line location (perce

44、nt cx) . 85 Horizontal tail: Airfoil section NACA 65-008 Area. sq ft . 35.98 span. ft 12.25 Aspect ratio . 4.17 Taper ratio . 0.55 Tail length. fran 0.25 M.A.C. to elevator hinge line. ft . 16.34 Elevators : Area aft of hinge line (both sides). eq ft . 8.6 Span (one side). ft 5.91 Hinge location. pe

45、rcent horizontal-tail chord . 75 Mean aerodynamic chord. ft . 0.73 Vertical tail surface: Area. sq ft . 25.68 span. ft 5.55 Aspectratio 1.20 Taper ratio 0.56 Fin offset . 0 Tail length. from 0.25 M.A.C. to rudder hinge line. ft 17.38 Dorsd-fFn -ea. sq ft . , 9.08 I ! Provided by IHSNot for ResaleNo

46、reproduction or networking permitted without license from IHS-,-,-10 NACA RM L5m3 PHYSICAL CEARACTERISTICS OF DOUGLAS D-558-1 AIRPLANE - Concluded Rudder : Area aft of hinge line, sq ft . . . . . . . . . . . . . . . . . 7.92 Mean aerodynamic chord, ft . . . . . . . . . . . . . . . . . . . 1.44 spm,f

47、t 5.6 7 Fuselage : Fuselage length, ft . . . . . . . . . . . . . . . . . -. . . , 35 .Ob Fuselage depth (maxFzrmm), ft . . . . . . . . . . . . . . . . . . 4.0 Fuselage width (maxFmum), ft . . . . . . . . . . . . . . . . . . 4.0 Load condition: Airplane weight (full fuel without tip tanks) lb . . . .

48、 . . . 10,610 Center of gravity, percent mean aerodynamlc chord . . . . . . 23.34 Negligible movement of the c-.g.-with fuel consumption. -E57 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L5lA23 Figure 1. - Three-view drawing of the Douglas D-558-1 airplane. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-f . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . I W r Figure 3. - Vortex-generator location on the D-59-1 air