1、RESEARCH MEMORANDUM INVESTIGATION AT LOW SPEED OF THE EFFEETIVENESS AND HINGE MOMENTS OF A CONSTANT-CHORD AILAVATOR ON A URGE-SCALE TR,IANGULAR WING WITK SECTION MODIFICATIGN By fohn G. Hawes and Ralph W. May, Jr. NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS uNCLAsslFiEL WASHINGTON April 24, 1951 Pro
2、vided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 6 NACA RM 5126 “ . - NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS EIHGE MOMZWTS OF A CONSTANT-CHORD AILAVATOR WITH SECTION MODIFICATION By John G. Hawcs and Ralph W. May, Jr. An investigation has been con
3、ducted in the Langley full-scale tunnel to determine the low-speed longitudinal, lateral, and hinge- moment control characteristics of a basic 60 delta wing of aspect ratio 2.31 with 10-percent-thick biconvex symmetrical airfoil sections. The wing was also tested in an altered condition with a nose
4、glove employing NACA 65-010 section ordinates. The wing was equipped on the left semispan with a constant-chord plain semispan ailavator having two sewent 8. The results indicate that the Characteristic force breaks caused by a separation vortex on the basic sharp-edged airfoil were eliminated by in
5、stalling an HACA 65-010 nose glove. The effectiveness and hinge moments for the fill semispan ailavator for both wings represent the sum of the chasacteristics of the two segments. The leading-edge separation vortex on the sharp-edged wing introduced large hinge- moment discontinuities with large ai
6、lavator deflections. INTRODUCTION Previous pressure-distribution and flow investigations of triangular wings (references 1, 2, and 3) have sham leading-edge separation with an accompanying strong vortex flow for wings with sharp-edged airfoils, but the effect of the vortex decreased for the wings ha
7、ving airfoil sections with increasing nose radtii. In fact, the large-scale triangular wing of reference 1 and the small-scale triangular wing of reference 4, both with munded leading edges, showed trailhg-edge separation of the type normally expected for conventional wings. Provided by IHSNot for R
8、esaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM 5126 In the flow inveatigatdon of a zero-taper-ratio wing, reported in reference 3, it wa8 shown that the separation vortices increased in size and intensity as they swept progressively frm the leading edge inboard t
9、oward the plane of symmetry with increased angle of attack. The progression of this tyge of flow over the tip sections and wing trailing edge Qmld be expected to influence the control character- istics of trailing-edge flaps or ailavators, in view of the varied loading of the sections. The preeent t
10、ests were conducted- in the Langley full-scale tunnel to investigate the effects of the vortex flow on the effectiveness and hinge-moment characteristics of outboard, fnboard, and full-semispan constant-chord ailavators on the large-scale triangular wlng of reference 3. In an attempt to alleviate th
11、e leading-edge separation and vortex flow, the nose section of the basfc wing ww altered by installing a glove incorporating NACA 65-010 section ordinates parallel to the free stream over the forward 10 percent of the chord and faired to the wing at approximately the 50-percent-chord line. The data
12、are presented as standard NACA coefficients of forces and moments. The data are referred to a set of axes cofnciding with the wind axes, and the origin waa located at the quarter chord of the mean aerodynamic chord. c, ch aCL (“. = as, ving lift coefficient (2) drag coefficient (s) . pitching-moment
13、 coefficient hinge-moment coefficient (ure 2. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA EM 51A26 TESTS 5 Tests were made on the sharp-leading-edge and round-nose configura- tions to determine lift, drag, rolling, y-awing, and hinge momen
14、ts at zero yaw through an angle-of-attack range from -5.2 to +33.3O for ailavator deflections in increments of 9 from -30 to +30. Hinge moments of each segment of the ailavators were measured for the sepents deflected individually and for the segments .deflected together in order to evaluate the int
15、eraction between them. No measurements were made of the hinge moments of the full semispan ailamtor as a unit, but the sum of the hinge moments of the indivfdual segments measured with the sesents deflected together should be identically equal to the hinge moments of the unit. For all tests the dyna
16、mic pressure was 7.3 pounds per square foot, resulting in 8. Reynolds number based on the mean aerodynamic chord of 6.00 x 106. The airspeed was approximately 55 miles per hour, corre- sponding to a Mach number of about 0.07. The data were corrected for effecta of jet-boundary interference, air-stre
17、am misalinement, buoyancy, and blocking. Support tare correc- tions were not investigated since they were found to be negligible in reference 5 for an identical wing and support setup. RESULTS AmD DISCUSSION Presentation of Data To facilitate discussion of the results, the presentation of data is ou
18、tlined below. The longitudinal characteristics including lift, drag, and pitching- moment coefficients- of the basic and round-nose wings as affected by angle-of-attack change and outboard, inboard, and semispan ailavator deflections are sham in figure6 3 to 10. Figures 6 and 10 are summary figures
19、of the variations of Cm and CL with fjL at a = Oo, and The lateral characteristics including rolling and yawing-moment coefficients for both wing configurations are shown in figure 11 and 12. Figure 13 is a summary figure of the variation of with angle of attack. Provided by IHSNot for ResaleNo repr
20、oduction or networking permitted without license from IHS-,-,-6 NACA RM 5126 The hinge-moment characteristics of each ailavator segment with the segnents deflected individually or together are shown in figures 14 and 15. Figure 16 presents the variation of hinge moment with ailavator deflection for
21、zem angle of attack, and figure 17 summarizes the variation of with angle of attack. p6)L In some instances the ailavators were not set precisely at the desired angle and results for constant-deflection angles of the ail.- As discussed in detail in reference 3, the existence of a separation vortex,
22、which ia characteristic of highly swept wings having small leading-edge radii, has a tendency to increase the lift on outboard portions of the wing. AB the angle of attack is increased, however, the vortex sweeps inboard towards the plane of symmetry, and, as a result, the outboard portion becomes c
23、ompletely stalled. As seen in figures 3(a) and 5(a), pOsit1x-e deflections of the outboard segment produced rather large increases in lift-curve slope and rearward shifts in center of pressure at lift coefficients from approximately 0.3 to 0.6. At lift coefficients just above 0.6, the outboard porti
24、ons become completely stalled; hence, a decrease in lift-curve slope and an abrupt unstable change in pitching moment resulted. These changes were intensified with increased outboard ailavator deflections. With an increase in angle of attack the stall spread8 farther over the inboard portions and th
25、e pitching-moment variation becomes stable for all flap deflections. As a result of this separation progression over the wing, the effectiveness then decreases as the angle of attack is increaeed (see fig. 6). (“9L of the outboard ailavator first increases and The maxirmun CL for the basic wing with
26、 ailavators undeflected was 1.08 and was reached at an angle of attack of 33.3. As the ailavators were deflected to angles over loo, the minimum drag began to increase appreciably and the variation of drag with lift became greater. Effect of adding nose glove.- As shown in reference 2, rounding the
27、sharp leading edge of a full-scale triangular wing to 0.0025 removed the force breaks with the flaps mdeflected but not with the flaps deflected. Increasing the leadingedge radius to 0.011 had no further significant effect. In the present tests, installing the NACA 65-010 nose glove with a leading-e
28、dge radfus of 0.00687 resulted in removing all irregularities in the lift-curve slopes and unstable breaks in the pitching nmoments (figs. 7 to 9). Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-7 NACA RM L51A26 As a result in tne variation were obs
29、erved in of the ab Isence of -th .e vortex flow, the irregularities of (“.Sx and (cmE)L with angle of attack, which the sharp-leadingedge configuration (fig. 61, were removed by changing to the NACA 65-010 leading edge (fig. 10). With the nose glove installed, the outboard ailavator shows a gradual
30、decrease in both lift and mment effectiveness and the inboard ailavator shows practically no change as the angle of attack was increased. It is of interest also to note that at all but the lowest angles of attack the flap effectiveness is appreciably greater for the basic wing than for ,the round-no
31、se wing; possibly positive flap deflection at a fixed angle of attack tends to increase the size of the vortex and provide, thereby, an additional lift increment. For a.given high-lift coefficient positive ailavator deflections produced smaller Q values than negative deflections for all ailavator co
32、nfigurations tested. For a positive deflection of 300, the inboard ailavator generally produced the least drag. From data obtained but not presented, it was found that sealing the ga.p at the ailavator leading edge with the ailavator at zero deflection had a negligible effect on the longitudinal cha
33、racteristics of the wing. ateral Characteristics Basic winR.- For moderate ailavator deflections, the variation of rolling moment with deflection, as shown in figure 11, was fairly linear at every angle of attack. As would be expected, the semispan ailavator produced the greater rolling momnt. The p
34、oint for which the lif% coefficient is 85 percent of chax which is considered representative of the usable CL for the landing condition, is indicated on the ailavator-effectiveness parameter (“ zgL curve in figure 13( a) . For the outboard, inboard, and semispan aihvators, the values of (cz for the
35、round-nose %lax wing the values were 0.00050, Ot OC055, and 0.00115, respectively. 3. The effectiveness and hinge moments of the full semipm ailavator for both wing8 represent the sum of the characteristics of the tFT0 sement s. 4. The leading-edge separation vortex on the sharp-edged wing introduce
36、d large hinge-moment discontinuities with large ailavator deflections . Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. e Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 NACA m 5126 1. Wilson, Herbert
37、 A., Jr., and Lovell, J. Calvin: Full-Scale Investi- gation of the Maximum Lift and Flow Characteristics of an Airplane Having Approximately Triangular Plan Form. NACA RM 620, 1946. 2. Anderson, Adrien E. : An Investigation at Low Speed of R Large-Scale Triangular Wing of Aspect Ratio TKO. 11. The E
38、Pfect of Airfoil Section Modifications and the “bermination of the Wake Dawnwash. I?AC!A RM AP8, 1947. 3. May, Ralph W. , Jr. , and Hams, John G. : Low-Speed Pressure- Distribution and Flow Investigation for a Large Pitch and Yaw Range of Three Low-Aspect-Ratio Pointed Wings Having Leading Fdge Swep
39、t Back 60 and Biconvex Sections. NACA RM LW07, 1949. 4. Orlik-ackeman, K. : Experimental Determination of Pressure Distribu- tions and Transition Lines of Plane Delta Winge at Low Speeds and Zero Yaw. KT - Aero TN 3, Roy. Inst. of Technology, Div. of Aero., Stockholm, Sweden, 1948. 5. Whittle, Edwar
40、d F., Jr., and Lovell, J. Calvln: Full-Scale Investi- gation of an Equilateral Triangular Wing HavLng 10-Percent-Thick Biconvex Airfoil Sections. NACA RM L8GO5, 1948. 6. Langley Research Staff (Compiled By tho ma A. Toll) Suuunary of Lateral-Control Research. NACA Rep. 868, 1947. Provided by IHSNot
41、for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM 5126 11 Station 0 .50 75 1.25 -2.5 5.0 7.5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 85 90 95 100 TABIX 1,- AIRFOIL ORDINATES PARALUX TO PLAME OF SYM“RY OF WING COmFIGURATIONS TESTED dimensions in percent chord
42、 1 Ordinate Basic wing (10-percent-thick biconvex) “ “ “ 0.25 . 49 96 1.81 1.40 2.56 3.21 3.75 4.21 4.55 4.80 4.95 . 5.00 4.95 4.80 4.55 4.21 3.75 3.21 2.56 1.81 96 “ Wing with NACA 65-010 glove “ 0.77 93 1.17 1.57 2.18 2.65 3.66 4.07 4.42 4.67 4.81 4.92 4.98 . 5-00 4.95 4.80 4.55 4.21 3.75 3.21 2.5
43、6 1.81 -* 3.04 “_ L. E. radius = 0.00687 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 mAcA RM L5lA26 J 7 Section A-A p-. 120 . Figure 1.- Geometric characteristics of wing tested with and without an NACA 63-010 nose glove. All. dimensions are i
44、n inches. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-t 1 F (a) Basic sharp-edged wing configuration. Figure 2.- The low-aspect-ratio triangular wing mounted in the Langley full-scale tunnel. Provided by IHSNot for ResaleNo reproduction or networ
45、king permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . . . . . . . . . . . . . . . . . . I - . . . . . . . . . . . . . . . . . . . . . I r (b) mACA 65-010 round-nose glove configuration. Figure 2
46、.- Continued. P F Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3 . 17 . (c) Inboard left ailavator deflected upward. L-&;201J 6- . (a) Combination in
47、board and outboard left semispan aihvator deflected downward. Figure 2. - Conciudea. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . .
48、 . . . . . . . - . . . . . . . . . . . . . . . . . . , , . . ., , I 1.8 . . . . - . . . . . . . . . . . . . . . I “ I (a) Variation of a and C, wtth CL. Figure 3.- Effect of left outboard ailamtor aeeflectLon on the lons;ltuainal characteristics of the basic wing configurntion. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,
