1、RESEARCH MEMORANDUM ELXPERJMENTAL AND THEORETICAL AERODY?MMtC CHARACTERJSTICS i L,Otillllllllllllilllliililllullll OlA4B1b NACA RM L57EO2 L NATIONAL ADVISORY COMMITIEE FOR AERONAUTICS RESRARCHMEMORAND “ EXPERlIMENTALAND THEORECICALAERODYNAMIC CHARACTERISTICS OF TWO LOW-ASPFCT-RATIO DELTA WINGS ATANG
2、LESOF ATTACK To 50 AT A MACH NUBER OF 4.07 By F?cedM. Smith SIMMARY - An investigation has been conducted in the Langley 9- by g-inch Mach nmber 4 blowdown jet to determine the aerodynamic characteristics of two double-wedge-section delta wings of aspect ratio 1.3 and 2.3 to angles of attack of d“ o
3、 at a Mach number of 4.07 snd Reynolds nmbers of 6.0 and 5.3 x 10 , respectively. -The results of the investigation are ccanpared with the predictions of linear theory, two-dimensional shock-expansion theory, Newtonian-impact theory, and a method which utilizes the shock-wave and expansion-wave equa
4、tions expanded by the two-dimensional hypersonic-flow similarity parameter. Linear theory, although fortuitously, generally gives the best predictions for all components. The results of this investigation extend a trend established in lower Mach nmber tests that the max$num lift coefficient for low-
5、 aspect-ratio delta wings decreases with increasing Mach nmber. The results also extend to a Mach number of 4.07 the trend indicated by lower Mach number tests that the angle of attack for maximum lift coef- ficFent ticreases with increasing Mach number. INTRODUJTION Numerous expertients have been m
6、ade to determine the supersonic aerodynamic chsracteristics of low-aspect-ratio delta wings at low and moderate angles of attack (for example, refs. 1 to 5). Relatively few tests have been made, however, at high angles of attack (above sn angle of attack of 30) and these data sre limited to a maximu
7、m Mach number of 3.36 (refs. 6 to 8 and unpublished-data.obtained at the Ames Aeronautical Laboratory). This shortage of high-angle supersonic data leaves the designer with no empirical results upon which to base his predictions of the maxim= lift and other wing characteristics at high HADC ADJ 57-5
8、007 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACA RM L57ED2 angles. It also precludes a ccxnpsrison of theoretical methods with experimental results to determine which theoretical methods best predict the high-angle-of-attack and high Mach n
9、umber wing chsracteristics. The present investigation, instigated by the lack of experimental data, was conducted on two low-aspect-ratio (1.33 and 2.31) double-wedge- section delta wings in the Langley 9- by g-inch Mach number 4 blowdown jet. In this investigation, normal force, chord force-, pitch
10、ing mane these humidities should be low enough to eliminate water-condensation effects. The test-section static temperature-and pressure did not reach values for which liquefaction of the air would occur. .- - MODELS The models consisted of two steel semispan delta wings having, respectively, semiap
11、ex angles of 18.4 and 3o“, aspect ratios of 1.33 and 2.31, and double-wedge sections 8 percent and 5 percent thick. The wings are shown in schematic form as figure 2. PRECISION OF DATA The maximum inaccuracies of the experimental angles, forces, and moments due to balsnce and recording equipment lim
12、itations, and the t average repeatability of the system have been estimated and are presented in the following table: Value .- Accuracy FL +0.1 fl.010 CD . .- . .-. K).cqo cm ; . ; . ; .-;-i . -i-i -. i . : RM A7IlO, 1947. 3. Ellis, Macon C., Jr., and Hasel, Lowell E.: Preliminary Investiga tion at
13、Supersonic Speeds of Trianguls and Sweptback Wings. NACA 9% 1955; 1349. 4. Tniaanri;-Eaward F., and DmTning, Robert W.: Aerodynamic Characteristics of Two Delta Wings at Mach Number 4.04 and Correlations of Lift and Minimum-Drag Data for Delta Wings at Mach Numbers From 3.62 to-6.9. NACA RM L52Kl9,
14、1932. - 5. Dunning, Robert-W., and Smith, Fred M.: Aerodynamic Characteristics of Two Delta Wings and Two Trapezoidal Wings at Mach Number 4.04. NACA RM L33D3Oa, 1933. , 6. Gallagher, James J., and Mueller, James N.: An Investigation of the Maximum Lift of Wings at Supersonic Speeds. CA Rep. 1227, 1
15、935. (Supersedes NACA RM L7JlO.) 7. Raattari, George E.: Pressure Distributions on !lYiangulm and Rectangular Wings to High Angles of Attack - Mach Nmnbers 1.45 and 1.9. NACA RM A34D19, 19%. 8. Kaattszi, George E.: Pressure Distributions on!h9.angular and Rectangular Wings to High Angles of Attack -
16、 Mach Numbers 2.46 and 3.36. NACA RM A54Xl2, 1935. 9. Dorrance, William H.: Two-Dimensional Airfoils at Moderate Hypersonic Velocities. Jour. Aero. Sci., vol. 19, no. 9, Sept. 1952, PP. 393-6oo. 10. Ulmann; Edward F.,and Lord, Douglas R.: An Investigation of Flow Chsracteristics at Mach Number 4.04
17、Over 6- and g-Percent;Thick Symmetrical Circular-Arc Airfoils Having 30-Percerrt-Chord Trailing- Edge Flaps. ws F E. R.: The Turbulent Boundary Layer for Ccmpressible Fluids on a Flat Plate With Heat Transfer. Rep. No. AL-997, North American Aviation, Inc., Jan. 27, lgw. 15. Boatright, William B.: E
18、xperimental Study and Analysis of Loading and Pressure Distributions on Delta Wings Due to Thickness and to Angle of Attack at Supersonic Speeds. NACA RM 56114, 1956. . 16. Ulmann, Edward F., and B aspect-ratio-l.33 delta wing; .E = 18.4O %c %I (1) (L/D)- Experiment 0.0150 0.0093 4.78 Linear theory
19、l o177 .oll7 4.65 Shock-expansion theory .0185 .0108 4.95 Newtonian- impact theory .0050 .0052 3.35 Dorrances method (ref. 9) .o18g .0107 4.57 Wing 2; aspect-ratio-2.Xdelta wiqg; e = 200 Experiment 0.0184 0.0061 6.57 Linear theory l o177 .oo65 6.106 Shock-expansion theory 00177 .0064 6.31 Newtonian-
20、impac.t theory .002g .0040 3.81 Dorrances method (rek9) .0180 .oo60 6.25 All theoretical predictions for Cm contain a skin- friction drag coefficient computed from references 13 and 14 (O.OOb3 for wing 1 and 0.0038 for wing 2). L Provided by IHSNot for ResaleNo reproduction or networking permitted w
21、ithout license from IHS-,-,-NACA BM L57EO2 13 Figure l.- Schematic diagram of Langley g- by g-inch Mach number 4 blow- down jet showxLng:mod,el and balance arrangement. f : i Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 NACA RM L57EO2 * -b-P- m
22、 wing 1 9 l . - Leading edge wing 2 Figure 2.- Schenlatic diagrams of test wings. DFmensions are in inches. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-L-57-1568 Figure 3.- Photograph of wing 1 showing the regions of l.adnar and turbu- lent bound
23、ary hyer. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM L57EO2 Figure 4.- Normal-force coefficients of two delta wings at M = 4.07. Flagged symbols represent check points. . Provided by IHSNot for ResaleNo reproduction or networking perm
24、itted without license from IHS-,-,-.aa .72 . la 1 0 CL Expsrlment - cpJcosa f- (a) Wing 1. Figure 5.- Lift coefficients of two delta wings at M = 4.07. Fl symbols represent check points. .agged Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.S6 .a8
25、.%o .72 .a4 .56 .32 0 -.m t I c- -8 0 8 18 24 32 40 42 66 (L, ciag (b) Wing 2. Figure 5.- Concluded. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L57EO2 19 .7a. .32 II I I I (a) Wing 1. Figure 6.- Drag coefficients of two delta wFngs at
26、M = 4.07. symbols represent check points. Flagged Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-20 NACA RM L57E02 .w .86 .80 .72 .w t-t-t-m I ) , t I I I i i (b) Wing 2. Figure 6.- Concluded. Provided by IHSNot for ResaleNo reproduction or networki
27、ng permitted without license from IHS-,-,-NACA RM L57FiO2 - 21 .04 .03 / / / , / / / / / t i I I I 4 I I I L - Reference 8 -8 cl 8 16 24 32 40 48 56 =I deg Figure 7.- Pitch- mcment and chordwise center of pressure of two delta wings at M = 4.07. Flagged symbols represent check points. Provided by IH
28、SNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 NACA J3M L57F102 90 70 -8 0 B 18 24 32 40 48 58 a, deg (b) Wing 2. Figure 7.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM L57EO2 23 40 90 zl
29、 10 I I I I I I -8 0 8 16 24 32 40 48 56 a, wf (a Wing 1. Figure 8.- Wing-root bending moments and spanwise centers of pressure of two delta wings at M = 4.07. Flagged symbols indicate check points. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NAC
30、A FM 57302 -. 08 80 Eo - 40 30 20 10 -8 0 8 16 24 32 40 48 -56 a, w b) Wing 2. Figure 8.- Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-:M NACA RM L57EC2 25 , ” 2 . % 3 1 v .m .ed .66 .64 .d2 .eQ l (a) Wing 1. .m .68 .66 0, c&y 1 Center
31、I Linear I I 60 of area, / I I- _ I 24 .Z6 -28 .-SO -32 .34 -36 .38 -40 -42 Spsnwiee location, */b (b) Wing 2. Figure 9.- Center-of-pressure travel for two delta wings at M = 4.07. Flagged symbols indicate check points. Provided by IHSNot for ResaleNo reproduction or networking permitted without lic
32、ense from IHS-,-,-26 NACA IUd L57E02 1.2 A SGlrce I . h 0 1.3 Present -cl 2.3 teats mfs. 7, 8, bl. m I I AAL 0 2.0 (a) Maximum lift coefficient. . (b) Angle of attack for maximum lift. Figure lO.- Effect of Mach number on maximum lift of-several low-aspect- ratio delta wings. Provided by IHSNot for
33、ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA FM L57EO2 27 1 0 (a) Wing 1. -3 0 a Id 24 22 40 48 63 =l dt9 (b) Wing 2. Figure ll.- Lift-drag ratios of two delta wings at M = 4.07. nagged symbols indicate check points. NACA - Lnmtey Fteld, VP. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
