NASA NACA-TR-1096-1952 Experimental determination of the effect of horizontal-tail size tail length and vertical location on low-speed static longitudinal stability and damping in ing .pdf

1、: . . a r , :- I . . _ *. , ., .j NAII-ON :. . . smupicY Atid XMPING ,IN .PITCH:OF A. j -1 . I MbDEir HAVING 454 j3WEi?i!BACk : . . , , , . WA ._ . _ ,- -. , ; ” ,. .- _:- . a I. , , . : . ; -, ., .“. / . . .I:- . - _, I :, , .*- . Provided by IHSNot for ResaleNo reproduction or networking permitted

2、 without license from IHS-,-,-TECH LIBRARY KAFB, NM REPORT 1096 EXPERIMENTAL DETERMINATION OF THE EFFECT OF HORIZONTAL-TAIL SIZE, TAIL LENGTH, AND VERTICAL LOCATION ON LOW- SPEED STATJC J,ONGITUDINAL STABILITY AND DAMPING IN PITCH OF A MODEL HAVING 45” SWEPTBACK WING AND TAIL SURFACES By JACOB H. LI

3、CHTENSTEIN Langley Aeronautical Laboratory Langley Field, Va. I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-National Advisory Committee for Aeronautics Headquarters, 1724 F Street $VV., Washington 25, D. C. Created by act of Congress approved Mar

4、ch 3, 1915, for the supervision and direction of the scientific study of the problems of flight (JJ. S. Code, title 50, sec. 151). Its membership was increased from 12 to 15 by act approved March 2,1929, and to 17 by act approved May 25,194s. The members are appointed by the President, and serve as

5、such without compensation. JEROME C. HUITSAKER, SC. D., Massachusetts Institute of Technology, Chairman ALEXAKDER WETMORE, SC. D., Secretary, Smit,hsonian Institution, Vice Chairman ALLEN V. ASTIN, PH. D., Director, National Bureau of Standards. DETLEV W. BRONK, PR. D., President, Johns Hopkins Univ

6、er- sity. THOMAS S. COMBS, Rear Admiral, United States Navy, Chief of Bureau of Aeronautics. LAURENCE C. CRAIGIE, Lieutenant General, United States Air Force, Deputy Chief of Staff (Development). HON. THOMAS W. S. DAVIS, Assistant Secretary of Commerce. JAMES H. DOOLITTLE, SC. D., Vice President, Sh

7、ell Oil Co. MATTHAS B. GARDNER, Vice Admiral, United States Navy, Deputy Chief of Naval Operations (Air). R. M. HAZEN, B. S., Director of Engineering, Allison Division, General Motors Corp. WILLIAM LITTLEWOOD, M. E., Vice President, Engineering, American Airlines, Inc. HON. DONALD W. NYROP, Chairman

8、, Civil Aeronautics Board. DONALD L. PUTT, Major General, United States Air Force, Vice Commander, Air Research and Development Command. ARTHUR E. RAYMOND, SC. D., Vice President, Engineering, Douglas Aircraft Co., Inc. FRANCIS W. REICHELDERFER, SC. D., Chief, United States Weather Bureau. HON. WALT

9、ER G. WHITMAN, Chairman, Research and Develop- ment, Board, Department of Defense. THEODORE P. WRIGHT, SC. D., Vice President for Research, Cornell University. HUGH L. DRYDEN, PH. D.: Director JOHN F. VICTORY, LL. D., Executive Secretary JOHN W. CROWLEY, JR., B. S., Associate Director for Research E

10、. H. CHAMBERLIN, Executive Ogicer HENRY J. E. REID, D. Eng., Director, Langley Aeronautical Laboratory, Langley Field, Va. SMITH J. DEFRANCE, LL. D., Director, Ames Aeronautical Laboratory, Moffett Field, Calif. EDWARD R. SHAPP, SC. D., Director, Lewis Flight, Propulsion Laboratory, Cleveland Airpor

11、t, Cleveland, Ohio LANGLEY AERONAUTICAL LABORATORY, AMES AERONAUTICAL LABORATORY, LEWIS FLIGHT PROPULSION LABORATORY, Ilangley Field, Va. Moffett Field, Calif. Cleveland Airport, Cleveland, Ohio Conduct, under uni$ed control, for all agencies, of scientific research on the fundamental problems of fl

12、ight II Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-F . (_ : . . . : i,. / . .I . . B !I I ,BR: ; . - . v NACA RFXkT 1096 ” EXPERIMENTAL DETERMINATION OF THEEFFlkT OF HtiRIZONkAkAIL TAIL IENGTH, AND VERTICAL LOCATION ONU)W-SPEED STATIC LONGITUDIN

13、AL STABILITY AND DAMPING IN PITCH OF A MODEL HAVING 45 SWEPTBACK WING AND TAIL k?FACES By Jacob H. Lichtenstein _L 1952 Page 2, first column, line 19: dynamic chord. Insert bar over symbol c for SIZE, mean aero- Page 2, first column, line 20: Insert symbol 2 for tail length. . . _ - - -. ,_. : - _-

14、j. - . .-_. -. . - .- .- . . . . - . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-REPORT 1096 EXPERIMENTAL DETERMINATION OF THE EFFECT OF HORIZONTAL-TAIL SIZE, TAIL LENGTH, AND VERTICAL LOCATION ON LOW-SPEED STATIC LONGITUDINAL STABILITY AND , .“.

15、 .- DAMPING-IN: PITCH OF A MODEL HAVING- 45O SWEPTBACK WING AND TAIL SURFACES l By JACOB H. LICHTENSTEIN SUMMARY An investigation has been conducted in the Langley stability tunnel to determine the effects of horizontal tails of various sizes and at various tail lengths (when located on the fuselage

16、 center line) and also the effects of vertica.1 location of the hori- zontal tail relative to the wing on the low-speed static longitudinal stability and on the steady-state rotary damping in pitch-for a complete-model configuration. The wing and tail surfaces had the quarter-chord lines swept back

17、45” and had aspect ratios of 4. The results of the investigation showed that, in agreement with analytical considerations, the contribution of the horizontal tail to static longitudinal stability was related directly to the tail size and length; whereas, its contribution to damping in pitch was rela

18、ted directly to tail size and the square of tail length. the designer has little assurance that the low-speed charac- teristics will be satisfactory for any specific configuration. The low-speed characteristics of wings suitable for high- speed flight have already been investigated quite extensively

19、. The contributions of other component parts of the aircraft, or of the various combinations of component parts for high-speed airplane configurations, however, are not well understood. In order to provide such information, a series of investigations of models having various interchangeable componen

20、t parts is being conducted in the Langley stability tunnel. In these investigations, the rotary derivatives are being determined by the rolling- and curved-flow techniques (see references 1 and 2) and the static stability characteristics are being determined by conventional wind-tunnel procedure. At

21、 low angles of attack, addition of the wing decreased the contribution of the horizontal tail to static longitudinal stability by about one-half to one-third depending upon the vertical posi- tion of the tail relative to the wing; the contribution of the hori- zontal tail to the rotary damping in pi

22、tch on the other hand was almost una$ected by addition of the wing, regardless of tail area or location. For con$gurations with the horizontal tail mounted along the fuselage center line, the static longitudinal stability was greater at angles of attack near the stall than at OD; the static longitud

23、inal characteristics were impaired, however, by moving the horizontal tail upward. On the other hand, for conJigurations with the horizontal tail mounted along the fuselage center line, the rotary damping in pitch was less at angles of attack near the stall than at O”, but the damping in pitch was g

24、enerally increased by moving the tail upward. The present investigation is concerned with the effects of horizontal tails of various sizes and at various tail lengths (when located on the fuselage center line) and also the effects of vertical location of the horizontal tail with respect to the wing

25、on the low-speed static longitudinal stability and the steady-state rotary damping in pitch for a swept-wing configuration. Some effects of fuselage fineness ratio and of wing-fuselage interference are also considered. The rotary damping in pitch specifies the damping resulting only from curvature o

26、f the flight path, such as that obtained during a steady pitching maneuver in which the radius of flight-path curvature is constant. For a pitching oscillation, the rotary damping derivative represents only a part of the total damping since additional contributions may result from unsteady aerodynam

27、ic phenomena such as the lag of downwash between the wing and horizontal tail (refer- ences 3 and 4). It was further indicated that, at an angle of attack of about IO”, the static longitudinal stability of the wing-juselage com- bination changed adversely and that the magnitude of this change was sl

28、ightly increased by the addition of tail area along the fuselage center line at the shortest tail length but was decreased by addition of area along the fuselage center line at the longest tail length. The model used in the present investigation had 45 sweptback wing and horizontal-tail surfaces wit

29、h aspect ratios of 4. The model configurations tested for the present investigation are generally the same as those configurations used in the investigations of static lateral stability deriva- tives reported in references 5 and 6. SYMBOLS INTRODUCTION Requirements for satisfactory high-speed perfor

30、mance of aircraft have resulted in configurations that differ in many respects from previous designs. As a result of these changes, 1 Supersedes NACA TN 2381, “Effect of Horizontal-Tail Location on Low-Speed Static Longitudinal Stability and Damping in Pitch of a Model Having 45 Sweptback Wing and T

31、ail Surfaces” by Jacob H. Lichtenstein, 1951, and NACA TN 2382, “Effect of Horizontal-Tail Size and Tail Length on Low-Speed Static Longitudinal Stability and Damping in Pitch of a Model Having 45O Sweptback Wing and Tail Surfaces” by Jacob H. Lichtenstein. 1951. The data presented herein are in the

32、 form of standard NACA coefficients of forces and moments which are referred to the stability system of axes, with the origin at the pro- jection on the plane of symmetry of the quarter-chord 232358-53-l 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-

33、,-,-2 REPORT 1096-NATIONAL ADVISORY COMMITTEE FOR AEIRONAUTICS point of the mean aerodynamic chord of the wing. The positive directions of the forces, moments, angles, and angular velocities are shown in figure 1. The coefficients and symbols are defined as follows: CD cn cn L D M N e s b c A Y x a

34、VF LF P .cv- 2v lift coefficient (L/k p V2Sw) drag coefficient (D/i p V2Sw) pitching-moment coefficient (1 M i p V2SwCw yawing-moment coefficient N U i p V2Swbw lift, pounds drag, pounds pitching moment about for example, increment resulting from interference ef- fect of wing and fuselage; for examp

35、le, A1cm,=c,+F-(c,-(c) Subscripts : W wing F fuselage V vertical tail H horizont,al tail r radian measure APPARATUS, MODELS, AND TESTS The general research model used for the present investi- gation was designed to permit tests of the wing alone, fuselage alone, or the fuselage in combination with a

36、ny of several tail configurations-with or without the wing. A sketch with some dimensions of the complete model with one particular tail configuration is shown in figure 2. A list of the pertinent geometric characteristics of various component parts is given in table I. All of the parts were constru

37、cted of mahogany. Three fuselages and three horizontal tails were used for the tests in various combinations with and without the wing. For convenience, each component is designated as follows: W_- Wing F, Fz, F3- _ -_-_-_- Fuselages V _ - _ -_-_-_- _ -_ Verticaltail H, Hz, Ha- Horizontal tails A co

38、mplete list of all t,he configurations investigated is presented in table II. The three fuselages (fig. 3) were bodies of revolution having circular-arc profiles and fineness ratios of 5 for fuselage 1, 6.67 for fuselage 2, and 10 for fuselage 3. The wing and the three horizontal-tail surfaces all h

39、ad aspect ratios of 4.0, taper ratios of 0.6, and NACA 658008 airfoil sections parallel to the plane of symmetry; the quarter-chord lines were swept back 45. Ordinates for the NACA 65A008 airfoil section are given in table III. The horizontal tails, the incidence of which was kept at O” for all test

40、s, differed from each other only in area and are designated as H, H2, and H3 (in order of increasing size) in figure 4 and table I. On each Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EFFECT OF HORIZONTAL TAILS ON LOW-SPEED STATIC LONGITUDINAL ST

41、ABILITY AND DAMPING IN PITCH 3 of the fuselages, each of the three horizontal-tail surfaces was attached along the fuselage center line and at the same longitudinal location. On fuselage 2, however, horizontal tail 2 was tested at three horizontal locations for each of three vertical locations, as i

42、llustrated in figure 5. In refer- ence to the horizontal-tail locations, the letters L, C, and U indicate the vertical position as being lower, center, or upper, respectively;-tbeletters B, M, and R indicate the horizontal location as being forward, middle, or rearward, respectively. The lower middl

43、e position is the same as that at. which the other two horizontal tails were tested. A drawing of a complete-model configuration with tbe horizontal tail in the lower position and a photograph of the model with the horizontal tail in the upper position without a wing are presented in figures 6 (a) a

44、nd 6 (b), respectively, to illustrate the test setup in the tunnel. The model was rigidly mounted on a three-support-strut system with the pivot point 4 inches rearward of the quarter-chord point of the mean aerodynamic chord. Forces and moments were measured by means of a conventional six-component

45、 balance system. The tests were made in the 6- by 6-foot test section of the Langley stability tunnel. The dynamic pressure for the tests was 24.9 pounds per square foot, which corresponds to a Mach number of 0.13 and to a Reynolds number, based upon the wing mean aerodynamic chord, of 0.71X10e. The

46、 angle of attack was varied from about -6O to about 32O for the tests. In addition to the straight-flow tests, the tunnel . -11.25 - -11.06- -Horizontal tail (see fig. 4) F6.754 FIGURE 2.-Dimensions of the complete model. AU dimensions are in inches. flow was curved to obtain values of qc/2Vof 0.008

47、,0.017, and 0.022. The method of curving the flow consists in curving the tunnel walls to obtain the proper air-stream curvature and inserting upstream of the test section screens which give the proper velocity gradierit across the test section. CORRECTIONS The angle of attack at ti 1.2.5 .96 2.50 5

48、.00 :“f: 19 7.60 2.12 10 2.43 E a:, 2.93 3.30 80 3”: 3.59 3.79 B”o 35 3.93 1: L. E. radius: 0.408 4.00 3.99 3.90 3.71 3.46 3.14 % 1.90 :Z .49 .02 cH/4 line-. +- 19.72 .-I FIGURE 4.-Dimensions of the horizontal tails tested. All dimensions are in inches. The effect of wing-fuselage interference on both the static longibudinal stability and damping. in pitch is shown in figure 14. An index to the data for the configurations investigated is given in table II. .-,- ., -