1、RESEARCH MEMORANDUM EFFECTS OF WCH NUMBER AND SWEEP ON TAE DAMPING-IN-ROLL CHARACTERISTICS OF WINGS OF ASPECT RATIO 4 BY Richard E. Kuhn and Boyd C. Myers, II Langley Aeronautical Laboratory - Langley Air Force Base, Va. . - . “ . . . . . I C“ “ Provided by IHSNot for ResaleNo reproduction or networ
2、king permitted without license from IHS-,-,-NATIONAL AWLSORY COMMITTEX FOR AESONAUTICS The damping-in-roll characteristics of three wings with an.aspect ratio of 4, a taper ratio of 0.6, sweep angles of 3.60, 3.6O, and 46.7O at the quarter-chord line, and with theWACA 65006 section have been detemhe
3、d through the Mmh nhe increase in magnitude. of the damglng- in-roll coefficient C zp with Mach number and the decrease with sweep angle, at low angles of-attack, agreed well markedly with angle of attack (in the test range) particularly at the higher Mach nunibera investigated. . with the theoretic
4、al variations. The dampin; coefficient Fncreaeed . INTRODUCTION Low-speed experimental data and theory (references 1 and 2) indicate an appreciable reduction in the. damphg-in-roll properties of a wing as the sweep angle is increased. The theoretical manner jn which these effects are affected by com
5、pressibility is treated in references 2 and 3. Little experimental data, however, are available at high- subsonic Mach nmbers for cmpariscm with theory. Accordingly, an extensive investigation is being conducted in the Langley high-sped 7- by 10-foot tunnel to detersnine the effects of smog angle an
6、d Mach n-in-roll charachrietics of throe wings of aspect referred to the quarter-chord ulle. The investigation utilized the froe-roll technique doscribed Fn reference 4 and the testa were mde at angles of attack of 0 40, 3 .45O, and 6.5 o .765 ft on model) rolling-mrrment coefficient (L/qSb) I. I, r
7、ate of roll, radians per second mc preseure , porn, degrees . . - “ . “ . - . “ - . , n d Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA R? LgElO 3 E angle of attack of -tip chord relative to root chord, radians K correction factor for WFng dis
8、tortion due to bending Subscripts : 82 left aileron ar right aileron test measured alw, uncorrected for distortion due to bending The pertinent dimsneions of the three wings ueed In the present investigation are given in flgure 1. The wings were constructed of an altnninum alloy. The sweptback was w
9、ere desised by shearing the -wept wing; that is, the chordwise elements of the Mswept wing parallel to the plane of s-try were moved rearward until the desired sweep ane of the 25-percent-chord line wm obtained. Thus, all wing sections parallel to the plane of symmetry are EACA 6506 sections. The ai
10、lerons were true-contour, sealea-gap, plan flaps of 20 percent chord and 4-0 percent Span. The wings were supported by a st- exbndb43 fmard into the test section from a vertical 8k-b located behind the model. The vertical s-. A photograph Of th8 inetdhtim is Shorn In f igwe 3 . The Provided by IHSNo
11、t for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 MACA RM LgElO roll freely under the moment created by the deflected ailerona, the rate of roll was recorded electrically- -. For each wing, static rolling-mgnent data and rates of roll were obtaine; thmugh a Mach ripe
12、r range of 0.4- to 0.91 at -8 of attack of 0.30 , 3.45O, and 6.50 and for aileron deflections of o , 44O, and *8O in a plane parallel to the plane of symuetry. The ailerans were deflected oppositely so that the total dlfferantial aileron deflections uGed were Oo, 80, and 16O. The size of the model u
13、sed in the present taves.tigatian reaulted In an estimatkh chokGzMachn-in-roll coefficient equal to a value of CZ$ = -0.0d5. . -. ._I The rolling kment and-Mach ndere have “been corrected for blockhg by the model and its wake by the method of reference 5. The jet-boundarg effects were estimated and
14、found to pe negligible. “ “2 The almlnm-alloy wings. were known to bend .kder load. Accordingly, the effect of wing distign-m the test; results wae lnvesti;ated. The possible soUrces of error. considered were: (1) deflection of the ailerom under load; (2) twist ofthe wing about Its ehstic axis due t
15、o the aerodynamic forces being applied at sane distance from the elastic axis; ad, (3) the spanwise ,change in angle of- attack due to bending of the wing panel under .tihe span-load dlstri%utim. The error due to this last consideration is eseentially zero, of course, for an unswept wing but increas
16、es very marbdly ae the sweep angle -is increaeed. .- w “ . . ,- -. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NAGA RM LgElO - 5 Static loading of the ailera and calculations of the twist of the wing indicated Ctmt erroh arig- potnts (1) and (2
17、) are negligible; however, calculations of the mxITsIoNs 1. The damging-in-roll coefficient C2 increased in magnitude with Mach nuniber and decreased with sweep angle at low angles of attack (0 .30 and 3.43-O) Fn the sane manner as that predicted by theory- P - “. 2. The Mgnitde of the dabgFng-in-ro
18、u coefficient 2, iircreased Y markedly with angle of attack (in the test range from 0.30 to 6 -5) particularly at the higher Mach nders. 4 Langley Aeronautical Laboratory . #I National Advleor Camdttee for Aeronautics Langley Air- Force Base, Va. - c . . . - I. “ I I . Provided by IHSNot for ResaleN
19、o reproduction or networking permitted without license from IHS-,-,-NACA RM LgElO . 9 1. Toll, Thomas A=, and QueiJo, M. J.: Approximate Relatiom and Charts for Low-Speed Stability Derivatives of Swept Wing8 NACA m 1581, 1948. 2 Bird, John D. : Saane Theore4ical Low-Speed Span Loading Character- ist
20、ics of Swept Wings in Roll Elnd Sideslip. RACA TN 1839, 1949 ,- 3. Fisher, Lewis R : ApproxTmate Correctians for the Hfects of Compessibiliiq on the Subsonic Stability Derivatives of Swept Wings . RACA TN 1854, 1949. 4 Myers, Boyd C m, 11, Khn, Richard E : High-S sonic DsmpFng- In-Roll Characteristi
21、cs of a Whg with the Qmrbr-Chord Line Swept Back .35O and with Aepect Ratio 3 and Wper Ratio 0.6 NACA RM Lgc23, 1949. 5. Herriot, John G.: Blockage Corrections for Three-Dimsnsianal-Flow Closed-Throat Wind Tmels, with Cansideration of the Effect of - Compressibility. NACA RM A7B28, 1947. 6- Peareon,
22、 Henry A., and Jonee, Robert T.: Theoretical Stability and Control Characteristics of Wings with Various Amounts of c Taper and Twist . NACA Rep. 635, 1938. 7 MacLachlan, Robert, and Leeo, William: Correlation of Two Xxprlmental Methods of Determining the Rolling Charac,teristics of Unswept W-6. NAC
23、A TN 1309, 1947. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 “ . .“ - . , “ I Tabulated Data Wing Aileron Area 2.25 $q.ft. Type True contour, sealed gap Aspect ratio 40 Chord 20 %c Airfdl seZfion N ACA 65A 606 Span 40% b/2 . Mean aerodynamic c
24、hord 0.765 ft. Outboard station 95% b/2 Tdper rotio 0.6 0 Root chord I1;25in Tip chord 6.75in. span 3.0ft. lnboard station 55 % b/2 .“ -_ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I . I t . t- Tunnel ceiling Balance strut- Strut fairing - Sting
25、 fairing 7 Tunnel center line 1“ . Tunnel floor I. “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- .- . “ -. . . - , :. -% . . :- . .L . -. . .- .i . ,- - “ .- ._ ., 1. - “ “ . -, Provided by IHSNot for ResaleNo reproduction or networking permitte
26、d without license from IHS-,-,-. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3 , “ , . -? .- . . .: I -, . . “ - .- *. - . .- :i ,. . . -. I. I I . . . . . . .r I “ . “ . “
27、 . . . . .“ . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . 4 3 2 I 0 4 5 .6 .7 .a .9 I .o Mach number, M 11 Plgure 4.- The variation of teat Reynolds number with Mach number based on the mean aerodpmlc chcxrd p of 0.765 foot. ul . .
28、. . -. . . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . . . . . . . . . . . . . . . . . , . . . ,. . . l2 I .I I .o 3.6 .30,3.45,86.50 - - - - 326 .30,3.45, ,08 .04 . 0 I. -. -.04 -.08 .08 .04 .O - .04 “08 08 .04 . . . 0 - . -. - - . -.0
29、4 -138 3 4 .5. .6 7. .8 .9 I .o Mach number, M I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.08 04 0 I- .04 -08 .08 .04 0 -.04 -.08 .08 .04 0 - .04 - .08 .3 4 5 Figure 10.- The variation angle - for various Pb 2v attack. Ac/4 = 32-60. 21 .6 7 .8
30、 9 1.0 Mach number, M - with Mach number of the wing-tip helix aileron deflections at several angles of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-22 Mach number, M . ,. 1 :=. with Mach number of the wing-tip helix aileron defleotiom at several
31、angles of . “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WA RM LgElo 0 c -. 1 -. 2 P -.3 -.4 -5 -1302 DO I .O (&I 0.30 3.45 “ -“ 6.50, 0 Theory, ref. 2 0 Theory, ref3 “_ “ .5 .6 .7 .8 .9 I .o Mach number, M Figure 12.- The variation with Mach nu
32、mber of the parameters czp CZ6J and E) at seyeral anglee of attack for the 3.69 sweptback wing. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-24 NACA RM LgElO 0 -.I. -2 -.3 i m 0.30 3.45 6.50 “ T“ “- 0 Theory, ref. 2 “ 0 Theory, ref. 3 I cb -4 -5
33、 006 004 1302 0 002 .oo I n 4 5 .6 f .8 .9 Mach number, M I .o Figure 13.- The variation with Mach nlllTiber of the parameters .C Ip, CzBJ . . “_ “ 1- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM LQElO 0 -.I -.2 clP . -.3 -5 -002 c -001 8
34、0 .006 .004 .002 0 4 .5 .6 .7 , .8 .9 LO Mach number, M Figure 14.- The variation with Mach nmiber of the parameters C 2$ C 28 at several angles of attack for the 4-6.7O sweptback WIG. - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-26 0 -. I -. 2
35、.cb -.3 -.4 -.5 -002 I I (2 H NkA RM L9E10“ 0 0. .o to 20 1 30 40 50 Angle of sweep, A, deg 1 .*. . “ (a) M = 0.40. Figure 15.- The effect of sweep angle on the parametera Clp, el8, at eeveral wee of attaok and for various Mach ILzrmbers. Provided by IHSNot for ResaleNo reproduction or networking pe
36、rmitted without license from IHS-,-,-. -.I -.2 -.3 -4 d.5 .002 % .oo 1 0 0 IO 20 30 40 50 Angle of sweep,&., deg (de 0.30 345 6.50 0 Theory, ref. 2 .006 002 0 Figure 15.- Continued. 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-,-,-
