1、q 9 3009 2710 fER No. 925 NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS DETERMINATION OF THE DAMl?ING MOMEXL IN YAIIING FOR Langley Memorial Aeronautical Laboratory Ia,ngley Field, Va. WASHINGTON NACA WARTIME REPORTS are reprints of papers originally issued to provide rapid distribution of advance res
2、earch results to an authorized group requiring them for the war effort. They were pre- viously held under a security status but are now unclassified. Some of these reports were not tech- nically edited. All have been reproduced without change in order to expedite general distribution. L- 395 Provide
3、d by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-f 4 c i 1. NATIONAL ADVIS9RY C0hQ.Q TTEE FOR AEIIONAUTICS z ADVANCE -EPORT . DETERMINAYION OF TKE DAYPIlTG XOI1TE;NT IN YAWING FOR TAPERED !LINGS WITH PARTIAL-SPAIJ FLAPS By Sidney I!. Harmon SUhJIAFtY A metho
4、d for determiqing the damping moment in yawing for tapered wings with partial-span flaps is presented herein. Charts are given for untwisted wings with taper ratios of 0.25, 0.50, and 1.009 with asbect ratios from 6 to 16, and with center-span fLaps extending from 25 to 100 percent of - . the wing s
5、emispan. The results are also applicable to tip- span flaps extending from 0 to 75 percent of the wing semi- span. The calculated damping monient in yawing is compared - with experimental results for a rectangular wing with a flap having a span 60 percent of the wing span. IHTRODUGTI ON ) The calcul
6、ation by Wibselsberger of the wing damping mgment for an untwisted elliptical wing in yawing is suinma- . rized and extended in reference 1 to the case of an untwisted rectangular wing. Reference 2 presents the results of cal- culations for a wide range of taper ratio for untwisted wings and also fo
7、r the special angle-of-attack distribution that I results from the deflection of partial-span flaps of constant chord ratio when the rest of the span is at zero angle Ofattack Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 The rcsul-ts Iri referen
8、ce 2 for the yawing derivative due to induced drag, howeve$ contain6:iaccuracles because of the omission of ai-, inportant ten;i in the fomula for th6 yawing moment. Also, as noted in reference 2, the results cannot c r be applied .by simple superposition to the case in which the lift is contributed
9、 simultaneously by partial-span flaps and by the plain portions of the wing. This limitation follows from the fact that the damp1n.g movent varies as the square of the angle-of -attack distrJibution; hence, separate components ox the lift distribution have interactions that contribute to the resulta
10、.nt value of the yavuing derivative i # The presznt analirsis giaes the results of calculations e for th.e yawing derivative BCrl/b (g?) for untvfis ted tapered wings viith partial-span flaps of constant chori? ratio at various angles of attack. “he results Ere pmsented for the same range of taper r
11、atio as is considered in reference 2 and for center-spar1 flags extendin?omeiit doefficient due to spanwise indwced- drag distribution angular veloclty fn yaw, radians per second wing span wind velocity along plane of symmetry of airplane local relatLve whd velocity at eny section , coordinate rriez
12、suec! alorig lateral axis of airplane circulation aro7md any sect!-on section lift coefficient wing chord at ariy sectlcn incuczd angle of attack at any section, radians coordinate indicating fl.xed spanwise position noma1 component of velocity ya7f.ring moment due to spanwise induced-drag distri- b
13、ution (JnqsSb) dynamic pressure at plane of symmetry (ipvs2) wing area density parametera defining spanwise position y = 7 b cos 8; - b vvlilen 8 = n, y = -2 - 2; when 8 = 0, parameter defining f fxud spanwise position wing chord at plane of symmetry o? airplanel slope of section lift curve at plane
14、 of symmetry of airplane, per radian Al, . An ccefficients of Fourier series (see reference 3) Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 k aspect ratio cnr yawing derivative C, span section profile-drag coefficient d.0 Acdof increment of sect
15、ion profile-drag coefficient due to dcflection of flaps AC Fncrement of yawing derivstive due to spanwise “r changes in profile drag Subscripts ar,d superscripts: W refers to cz- or -distribution resulting from CdO wing angle of attack -distribution resulting from CdO f ,refers to cb- or deflection
16、of flaps a J 1 f C center-span flaps ft tip-span flaps Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I 1 I 5 I MdTHOD AND AYALYSTS .1 I The method used in the present analysis for calculating the stability derivatives is based on the assumptions ou
17、t- lined in reference 2. Because the angular motions con- sidered in the present analysis are sinal1 (rb/ZV, Fig. 4 , Figure 4.- Coqarsion of theoretical and wind-tunnel results A = 6; A = 1.00: b.p/b = 0.60; ACLZ = 0.56: and C for yawing derivate for partial-span flapped wing. = CL - 0.56, L, Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
