NASA NACA-CR-3675-1983 An aerodynamic analysis computer program and design notes for low speed wing flap systems《空气动力分析计算机程序和低速襟翼系统的设计说明》.pdf

《NASA NACA-CR-3675-1983 An aerodynamic analysis computer program and design notes for low speed wing flap systems《空气动力分析计算机程序和低速襟翼系统的设计说明》.pdf》由会员分享，可在线阅读，更多相关《NASA NACA-CR-3675-1983 An aerodynamic analysis computer program and design notes for low speed wing flap systems《空气动力分析计算机程序和低速襟翼系统的设计说明》.pdf（75页珍藏版）》请在麦多课文档分享上搜索。

1、._ _ NASA Contractor Report 367 5 An Aerodynamic Analysis Computer Program and Design Notes for Low Speed Wing Flap Systems Harry W. Carlson and CONTRACT NAS I- 16000 MARCH 1983 Kenneth .B. Walkley Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH

2、 LIBRARY KAFB, NM IlnlllIIUllillllmllllnl IlOb NASA Contractor Report 3 67 5 An Aerodynamic Analysis Computer Program and Design Notes for Low Speed Wing Flap Systems Harry W. Carlson and Kenneth B. Walkley Kentron International, Inc. Hampton, Virginia Prepared for Langley Research Center under Cont

3、ract NASl-16000 National Aeronautics and Space Administration Scientific and Technical Information Branch 1983 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- SUMMARY This report describes the expanded capabilities for analysis and design of low spe

4、ed flap systems afforded by recent modifications of an existing computer program. The program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edge flap deflection schedules. Among other new features of the program are a revised attainable thrust estimation metho

5、d to provide more accurate predictions for low Mach numbers, and a choice of three options for estimation of leading-edge separation vortex flow effects. Comparison of program results with low speed experimental data for an arrow wing supersonic cruise configuration with leading-edge and trailing-ed

6、ge flaps showed good agreement over most of the range of flap deflections. Other force data comparisons and an independent study of airfoil and wing pressure distributions indicated that wind-tunnel measurements of the aerodynamic performance of twisted and cambered wings and wings with leading-edge

7、 flaps can be very sensitive to Reynolds number effects. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION The low speed aerodynamic analysis method of reference 1 provides estimates of wing performance which include the effects of attaina

8、ble leading-edge thrust and vortex lift. The method was shown to be particularly useful in the subsonic analysis of vehicles designed for supersonic cruise. In reference 2, the computer program described in reference 1 was used as an aid in the design of low speed leading-edge flaps for a candidate

9、supersonic transport configuration. This report describes modifications and improvements to the original computer program to permit more convenient, more accurate, and more efficient treatment of simple leading and trailing-edge flap systems. In this improved program, there is provision for direct i

10、nput of flap geometry. For the study of reference 2 it was necessary to make internal program modifications. In addition, the program is arranged so that, with little additional expense, solutions may be found for various combinations of leading and trailing-edge flap deflections. Another improvemen

11、t incorporated in the present computer program is a revised attainable thrust algorithm which is more accurate at the low Mach numbers sometimes encountered in wind tunnel testing. A means of estimating the distribution of leading-edge separation vortex forces has also been provided. The applicabili

12、ty of the program results to the aerodynamic analysis of wings with flaps is demonstrated through correlations with experimental data. In addition, suggestions for use of the program in an iterative fashion for flap system design are given. 2 Provided by IHSNot for ResaleNo reproduction or networkin

13、g permitted without license from IHS-,-,-SYMBOLS AR b C z Cave CA CN ACN,v cn CR ct ct,F CA CN CD cD,O ACD CL cL,a cP Cp,lim ACP,V wing aspect ratio, b*/S wing span local wing chord mean aerodynamic chord average wing chord, S/b section axial force coefficient section normal force coefficient sectio

14、n normal force coefficient increment due to the leading edge separation vortex chord of wing section normal to local wing leading edge with maximum thickness at mid chord n = sinh 27l LL(l+n)tanAL+ntanATJ+cosAL section resultant force coefficient theoretical section leading-edge thrust coefficient t

15、heoretical section leading-edge thrust coefficient for a flat wing at lo angle of attack total axial force coefficient total normal force coefficient total drag coefficient drag coefficient at zero lift for a flat wing drag coefficient due to lift, CD-CD,0 total lift coefficient total lift curve slo

16、pe, per degree pressure coefficient limiting pressure coefficient used in definition of attainable thrust incremental pressure coefficient due to detached leading edge vortex 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-be k m M Mn R Rn r rf S s

17、S S t t max LX AU X,Y,Z X Xi 9x; X V a exponents used in Cp,lim equation arbitrary constant flap deflection multiplier free-stream Mach number Mach number normal to local wing leading-edge sweep angle, Mn = M cos AL free-stream Reynolds number based on 2 Reynolds number normal to local wing leading

18、edge, Rn =R$ cosAL wing section leading-edge radius effective leading-edge radius for attainable thrust calculation wing reference area suction parameter, CL tan (CL/CL a) - AcD CL tan (CL/CL,a) I C,*/WW distance along section camber line airfoil section local thickness airfoil section Mximum thickn

19、ess effective maximum thickness for attainable thrust calculations longitudinal perturbation velocity difference across the wing lifting surface as a fraction of the free stream velocity Cartesian coordinates distance in the x direction measured from the wing leading edge X values at leading edge an

20、d trailing edge of wing element at element semispan x value for center of detached leading-edge vortex flow angle of attack of wing, in degrees 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Aaft range of angle of attack for full theoretical thrus

21、t azt angle of attack of wing giving a theoretical leading-edge thrust of zero for a specified wing spanwise station B m Y ratio of specific heats n location of maximun wing section thickness as fraction of chord 6 flap deflection angle in degrees A sweep angle Subscripts: L leading edge n measured

22、normal to flap hinge line 0 original or first value T trailing edge 132 flap segment identifier des design condition zt zero thrust condition hl hinge line C cambered wing F flat wing 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-DEVELOPMENT OF C

23、OMPUTATIONAL SYSTEM Expanded capabilities for the analysis and design of low speed flap systems are offered by recent modifications of an existing computer program introduced in reference 1. The revised program provides for the simultaneous analysis of up to 25 pairs of leading-edge and trailing-edg

24、e flap deflection schedules. Among other new features are a revised attainable thrust estimation method to provide more accurate predictions for low Mach numbers, and a choice of three options for estimation of leading-edge separation vortex flow effects. The development of the basic computational s

25、ystem is covered in consider- able detail in reference 1. That report describes numerical methods which have been incorporated into a computer program to permit the analysis of twisted and cambered wings of arbitrary planform with attainable thrust considerations taken into account. The computationa

26、l system is based on a linearized theory lifting surface solution which provides a spanwise distribution of theoretical leading-edge thrust in addition to the surface distribution of perturbation velocities. In contrast to the commonly accepted practice of obtaining linearized theory results by simu

27、ltaneous solution of a large set of equations, a solution by iteration is employed. The method also features a superposition of independent solutions for a cambered and twisted wing and a flat wing of the same planform to provide, at little additional expense, results for a large number of angles of

28、 attack or lift coefficients. A key feature of the super- position technique is the use of leading-edge thrust singularity parameters to identify and separate singular and nonsingular velocity distributions. This separation permits more accurate determination of leading-edge thrust and more accurate

29、 integration of pressure distributions for twisted and cambered wings of arbitrary planform. The following discussions will be concerned only with significant changes and improvements to the methods described in reference 1. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted with

30、out license from IHS-,-,-Leading-Edge and Trailing-Edge Flap Surfaces The leading-edge and trailing-edge devices that may be treated by the present method are limited to “simple“ hinged flaps. There is no provision for additional slats or for any separation between the flaps and the main wing surfac

31、e. In the revised program separate solutions for the longitudinal perturbation velocity distribution are made for both a leading-edge flap surface and a trailing-edge flap surface. These two additional surfaces cover the entire wing planform as do the two original surfaces for the flat and twisted a

32、nd cambered wing. With this distinct separation of individual contributions to the overall wing loading distribution, it is a simple matter to combine loadings to cover not only a range of angles of attack but also a range of leading-edge and trailing-edge flap deflections. For input flap surfaces,

33、described as a spanwise distribution of flap chord and streamwise deflection angle, values of surface slope for each of the wing elements are determined within the program. Solutions for the longitudinal perturbation velocities corresponding to these new surfaces are performed simultaneously with th

34、e solutions for the flat and cambered wing. Because an assumption of lifting pressures proportional to the sine of the deflection angle rather than the tangent (the surface slope) is more reasonable, the lifting pressures are defined as: cP = auF sfna for the flat wing tan lo sin6 cP C = 2AUc Tandy

35、= 2Auc COS 6c for the cambered surface 7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-C sin6L P =a“LtangL= 2AUL cos so that: k = ACN, ave 2x; The k factor establishes the magnitude of the Ac, distribution acting on the wing surface. Since the surf

36、ace may be cambered and may include deflect- 14 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- ed flaps, there will be contributions to axial force as well as normal force. If the vortex center lies aft of the local chord midpoint, part of the vort

37、ex force will not affect the wing and will be lost. Only limited advice regarding the selection of the vortex options can be offered at this time. The first option, with the vortex force acting perpen- dicular to the wing reference plane at the wing leading edge, was used in the correlations with ex

38、perimental data given in reference 1. At large angles of attack, that approach seemed to overestimate the vortex effect-probably because rruch of the vortex field was actually aft of the wing surface rather than at the leading edge. The correlations with experimental data given in this report have b

39、een made using the second option. However, as discussed previously, this option is appropriate only for highly swept wings with delta or modified delta planforms. The authors have had no experience with use of the third option, the method given by ban. Angle of Attack Range for Full Thrust As will b

40、e discussed in a later section of this report, it may be desirable for design purposes to know how rmch a local leading-edge flap deflection angle may be changed from the local flow alignment condition (presumed to be defined by azt) and still retain attached flow and full theoretical thrust. This a

41、ngle of attack range may be found from the flat wing attainable leading-edge thrust calculations by setting the attainable thrust fraction Kt (defined by equation (9) in reference 3) equal to 1.0 and solving for ct,n. Since Ct,n is related to ct (equation (5) in reference 3) and ct is dependent on a

42、 (equation (15) in reference l), the flat wing angle of attack at which thrust begins to be lost at any given span station may be calculated. 15 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-PROGRAM DESCRIPTION The following description of the revi

43、sed program will cover the entire listing of input data and program results even though mrch of this information is repeated from reference 1. The computer program entitled “Aerodynamic Performance of Low Speed Wing Flap Systems“ may be obtained for a fee from: Computer Software Management and Infor

44、mation Center (COSMIC) 112 Barrow Hall University of Georgia Athens, GA 30602 (404) 542-3265 Request the program by the designation LAR 13116. This program is written in FORTRAN IV for use on the Control Data 6600 and Cyber series of computers. Data are input in namelist form under the code INPTl. T

45、he wing planform information is specified by a series of leading-edge and trailing-edge breakpoints. Up to 21 pairs of coordinates may be used to describe the leading edge and up to 21 pairs to describe the trailing edge. The planform input data in program terminology are: NLEY TBLEY TBLEX NTEY TBTE

46、Y TBTEX XMAX SREF CBAR 16 number of leading-edge breakpoints ,(limit of 21) table of leading-edge y-values in increasing order of y from wing root to wing tip table of leading-edge x-values corresponding to the TBLEY table number of trailing-edge breakpoints (limit of 21) table of trailing-edge y-va

47、lues in increasing order of y from wing root to wing tip table of trailing-edge x-values corresponding to the TBTEY table largest x-ordinate occurring anywhere on the planform wing reference.area for use in aerodynamic force and moment coefficients wing reference chord for use in aerodynamic moment

48、coefficients ._. . a. . . .,-1 . . . . , m.-I. ., ., . , .* _ _ -., ._ . . _ ,_,. -. . . . . . . ._ . . . . -.- -. -.-.-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- XMC x-location of moment reference center ELAR desired element aspect ratio (for flat and mildly cambered wings without flaps an element aspect ratio approximately one-half the full wing aspect ratio is recommended, for small chord flaps it may be necessary to use a larger element aspect rati