1、. RESEARCH MEMORAN.D-UM PRELIMINARY AERODYNAMIC INVESTIGATION OF THE EFFECT OF CAMBER ON A 60 DELTA WING KITE ROUND AND BEVELED LEADING EDGES By John M. Riebe and Joseph E. Fikes Langley Aeronautical Laborator FOR REFGIGENC Ll because of its Ugh taper and low aspect ratio, the triangular wing also h
2、as definite structural advantages Over the conventional swept wings. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM LgnO Theoretical calculations and experfmental studies have shown that such wing plan form will develop lift- ratioe at sup
3、ersonic Mach numbers which are sufficiently high for flight, being, howwer, generally lower than those of other plan forms, Low-epeed research has also indicated poor landing characteristics of delta plan forma because of the relatively low lift-drag ratios, -particularly in the high-lift condition.
4、 The results of a theoreticd study of triangular wings (refe- ence 1) and a pressure-distribution investi;ation (reference 2) show that hi no tmes for the effects of pw have been applied to the yaw data. The aerodynamic characteristics in pitch of the 1fft-g ratios are presented in figures 6 and 7.
5、The original data for S, = Oo (fig. 4( a) ) were rejected because of extremely large scatter of the test data resulting W L/D was about 8.2 for the wing wlth both round and beveled leading-edge flaps at Oo and about LO .5 for the optirmnn flap deflection, 20 for both leading edges at a Reynolds numb
6、er of 3 x 106. he vdws of L/D were not criticdly dependent upon flap deflection in the loo to 300 flapdeflection range and were generally about the SRB for both leading-edge shapes. However, flap deflections of 40 and greater had 1i-q ratios which were generally lower than those of the plain wing, T
7、he values of L/D for the present investigation were lcrner than, those obtained in other investigations at high Reynolde numbers. (See references 4 and 7.) However, according to reference 4, increasing Reynolds number resulted. in an increase in lift-drag ratio became of skb+Yiction drag- coefficien
8、t decrease. The lift-drag ratios of reference 4 are aclmowledged to be higher than those of reference 7, probably because of the method of testing. IT the trend of the cmes uith Reynolds nuniber is considered (decrease in L/D ratio with decrease in Reynolds number) with the lmr lift4rag mtios of oth
9、er delta4ng data (reference.7), the li-g ratios for the wing of the preeent investigation are of the right order of -tude. For several def lec tions of the beveled leadivdge flap,. higher maxinmn lift-drag ratfos were obtained at lmr Reynolds nuniber. (See fig. 7.) It is not known why the effect of
10、Reynolds nmiber on lif%-drag ratio is different from that shown in reference 4 and fram that of tb round leading-edge delta wing. (See fig. 6. ) Theoretical considerations (reference 1) and pressure distributions (reference 2) Imlicate that delta-uing plan forms have high negative peak pressures alo
11、ng the leading edge and spas“la3d distributions of elliptical shape, rasulting in mlnimuminduced drag. As msntfoned previously, separation and vortex flow originating at the apex QCCUT over the upper surface of the delta wing through a large part of the lift-coefficient range. This results in insrsa
12、sed turbulence and profile drag. =thou the characteristics of camber in increasing L/D may therefore be different on full.-rscale mobls ae . was shown in reference 8. Ai-flow stu4ies.- In order, to evEtluate the reduction and delay of vortex separation mer the delta King with flap deflection, additi
13、onal tests were made in the form of boundary-hyer flow studies, using lampblack and benzine on the wing with beveled leadin-dge flap undeflected and deflected 20. (See fig. 12.) The patterns shown are believed to be thoee produced by the bottom of the vortex flow (described in references 3 and 5) in
14、 contact with the surface of the model. The naajority of the flow photographs were taken after the benzine had evaporated or had been blawn from the model leaving a trace of the boundary air flow in the lampblack. However, in E- case8 where the vortex pattern was more clearly Indicated during evapor
15、ation there are two photpgraphs at a given angle of attack; the pwkTaJ- evaporation photograph always precedes the cmplet-vaporation phot- graph. The portion of the lampblack-benzine coating on the wiq affected by the vortex flow was alw8ys the firet to evaporate, as shown in the first photograph of
16、 some of the sets, (for example a = 80, fig. 12); this early evaporation would be veq udLikely if the pattern indicated on the photographs was merely cross flow. Vortices were also indicated by probing the model with a wool streamr. The streamer would rotate in and near the region of vortices indfca
17、ted by the lanpblack teste and in a direction tarard the wing tip near the surface of the model. The height of the ortlcee was indicated to be euch that the core would be on a line above the wing surface, when viewed frcm the side, at an angle approximately equal to one-third the angle of attack, ei
18、mllar to that sham in reference 5. In figure 12, wlth flaps undeflected, the vortex pattern is very slight at a = 3O; and at a = 5O (the angle of attack of mimum L/D, flaps 0) the vortex becomes quite evldent. As the value of a is increased, the outline of the vortices becomes wider and the inner bo
19、un-y mmes inward to the center of the model. With flaps Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LgFlO 9 deflected 20, the start of the vortex was delapd to 871 angle of attack of about 80. After the vortex began, it6 progression w the
20、 separation on the subject wing started more gradually and was not w as abrupt as that indicated for the reference -6 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 NACA RM LgFlO The mdel ma longitudinally stable up through the stall angle for
21、both nose shapes and all flap deflections, s characteristic generally founcl on deltadng models. Deflecting the flaps with either nose shape, varied the elope of the pitchi-nt-coefficient curve only slightly and resulted in positive incremnts iL pitching momnt. How- ever, the flaps would be matisfac
22、tory as a trim device because of the small increlnents and because of irregular and reversed effectiveness for large deflectiona. The Bmall positive increments in pitching moment with flap deflection probably reaulted from an unloading of the leadwage flaps; this unloading was more effective at the
23、rear part of the wing where the flap area uas larger. With som alterations, such a8 increased flap area near the apex, it may be posaible to desigp a flap givlng no trim change. Lateral Stability The aerodynamic characteristics of the delta wing in yaw showed only mall variation with flap leadingedg
24、e shape but generally large effects of flap deflection. (%e figs. 8 to U.) The lateral- stability coefficients varied fairly linearly with angle of yaw at loo angle of attack (figs. 8 end 91 and did not drop off in the yaw range tested. As is the usual case for highly swept wings of high taper, larg
25、e changes in the effective dihedral C occurred throughout the lift range (figs. 10 and 11) ; the parameters were dete-ed in the 3“ to 5 yaw range. Negative effective dihedral me present at lfft coef- ficients above 1.04 for the wing with round leadingedge flap at Oo. The values of C for the delta wi
26、ng with beveled leadingedge flap flap at Oo varied in a manner similar to those for the round leadiqp edge w5ng except that the maximum value of C ma less and negative effective dihedral occurred at a lower lift coefficient (0.92). Deflection of the nose flaps to 30 reduced the effective dihedral th
27、roughout the lift range for both leading-edge shapes. The maximum positive effective dihedral for th3 wing with 30 leadimdge flap deflection was about half that of the Oo flapdeflection condition and occurred at higher angles of attack. 2f 2* 23r The directional stability of the delta wing, as deter
28、mined by Cnq, was about the 8- with or without camker throughout the lift, range and increased slightly with lift coefficient up to about CL = 0.8. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LgFlO 11 - The paramter C increased uniformly
29、with Uft- coefficient for y* I the delta wing dth no c-r. Ths variation of C with lift coer- % ficient for both leadinwdge shapes deflected 30 showed a decrease up to about the middle of the lift range and then shared a largs increase with lift coefficient. Ths results of tests in the LangLey 300 Pl
30、IPH 7- by IO-fmt tunnel to determin8 the effect of cadmr on a flat-plate 60* apex-angle delta wing accomplished by deflection of full+- round and beveled leading- edge flaps fndicated the following conclusions: 1. In the 0.2 to 0.3 lift-coefficient range, up to 28 percent increase in lif-ag ratio wa
31、.8 o3tained for a 20 deflection of eithr the beveled or round leadiwdge flap. The values of lift-drag ratio were not critically deenbnt upon flap deflections in ths range from loo to 30. 2. The values of C vere ahost indepenbnt of nose-flap , deflection. The angle of attack for maxhm lift increased
32、with flap deflection. 3. Flap deflections resulted in small trim changes but did not greatly affect the longitudinal stability, Laagley Aeronautical Laboratory National Advisow Canrmittee for Aeronautics Langley Air Force Base, Va. b Provided by IHSNot for ResaleNo reproduction or networking permitt
33、ed without license from IHS-,-,-12 UCA RM LgnO 1. Jones, Robert T. : Properties of Lar-AspeclAhtio Pointed Wings at Speeds below and above the Speed of Sound. aACA Rep. 835, 1946. 2. Wick, Bradford H. : Chordwise and Spanwise Loam8 hbaeured at Low Speed on a Triangular Wing Having an Aspect Ratio of
34、 Two and an - “7. NACA 0012 Airfoil Section. mACA TN 1650, 1948. -.3. Wilson, Hsrbert A., JP., and Lovell, J. Calvin: FullScale Investi- gation of the plaxfmum Lift and Flow Characteristic8 of an Airplane Having Approximately Triangular Plan Form. NACA RM L6P;20, 1947. 4. Edxarde, George G., and Ste
35、pheneon, Jack D.: Tests of a Triangular Wing of Aspect Ratio 2 in the Ams 12-Foot Pressure Wind Tunnel. I - The Effect of Reynolds Number and Mach Number on the Aer- dynamic Characteristics o? the Wing with Rlap Undeflected. NACA RM AW5, 1947. 5. An?Lerson, Adrien E.: An Investigation at Law Speed o
36、f a Large-Scale Triangular Wing of Aspect Ratio Two. - 11. The Effect of Airfoil Section Modification8 and the Determination of the Wake Downwaeh. mACA RM AW8, 1947. 6. Gillis, Clarence L., Polhamus, Edward C., and Gray, Joseph L., Jr.: c-8 for Determining Jetdoundary Corractiom for Complete Models
37、in 7- by 10-Foot Cloaed Rectangular Wind Tunnels. MCA I. , - rn L5G31, 1945. 7. Anderson, Adrien E.: An Investigation at Low Speed of a Large-Scale Wing Eaving a DoublNedge Airfoil Section vlth Maximum Thickness at 20-Percent Chord. NACA RM Am6, 1947. _. Triangular Wing of Aapect Ratio Two. - I. Cha
38、racteristics of a 8. Whittle, Edward F., Jr., asd Lovell, J. Calvin: NlScale Investi- gation of an Equilateral Triangular Wing Eaving lO-Percent4hick Biconvex Airfoil Sectfona. NACA RM -5, 1948. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-MACA RM
39、 LgFlO c iQu N c Leading edge Round Round a a Figure 4.- Continued. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 HACA RM 910 c 72 0 .2 .4 .6 -8 10 12 14 LrfS coeff/cren7L, CL (b) Concluded. : Figure 4.- Concluded. Provided by IHSNot for ResaleN
40、o reproduction or networking permitted without license from IHS-,-,-XACA RM LgFlO 0 .z .4 .6 .8 IU 12 1.4 LIff coe fflcim f, c- (a) R X 1.5 x 10 . 6 Figure 5.- Aerodynamic characteristics in pitch of 60 delta wing with full”span beveled .leadingedge flaps. Provided by IHSNot for ResaleNo reproductio
41、n or networking permitted without license from IHS-,-,-26 NACA RM LgFlO (a) Concluded. Figure 5.- Continued Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LgFlO 27 (b) R z 3.0 x 10 6 . Figure 5.- Continued. Provided by IHSNot for ResaleNo re
42、production or networking permitted without license from IHS-,-,-28 mACA RM LgFlO (b) Concluded. Figure 5.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-P NACA RM LgFlO Lift coefficierd, CL (a) R Z 1.5 x 10 6 . Figure 6.- Lift-itrag ratios of 60 delta wing with full-epan round leadingedge flaps. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
