REG NACA-RM-L51C30-1951 Low-speed longitudinal and wake air-flow characteristics at a Reynolds number of 5 5 x 10(exp 6) of a circular-arc 52 degrees sweptback wing with a fuselagesiti.pdf

1、RESEARCH MEMORANDUM LOW-SPEED LONGITUDINAL AND WAECE AIR-FLOW CHARACTERLSTICS AT A REYNOLDS MJMBER OF 5.5 ,IO6 OF A . - “ By Gerald V. Foster and Roland F. Griner Langley Aeronautical Laboratory Hm.ED.-.“ - “. A -9 Provided by IHSNot for ResaleNo reproduction or networking permitted without license

2、from IHS-,-,-c . V NATIONAL ADVISORY COMMITTEZ FOR AERONcluTICS - RESEARCH MEMORANDLM . LOW-SPEED LONCZIUDINAL AKO WAKE AIR-FW CHARACmISTICS AT A REXNOLDS muMBER OF 5.5 x 106 OF A HORIZOIWAL TAIL AT VARIOUS VXRTICAL POSITIONS By Gervlld V. Foster and Roland F. Griner An investigation has been conduc

3、ted in the Langley 19-foot pressure tunnel to determine the effects of a fuselage and a horizontal tail located at various vertical positions on the low-speed longitudinal characteristics of 8 circulardrc 520 sweptback wing. Air-flow surveys were made in a vertical plane at a position which correspo

4、nded approxi- mately to the longitudinal location of the horizontal tail. The results were obtained at a Reynolds number of 5.5 x 106 with and without leading- edge and trailing-edge flaps. The low tail (located 0.132 semispan below the wing-chord plane) was situated below the vRke center for modera

5、te and high angles of attack and had a stabilizing influence through the angle-of-attack range because of a favorable rate of change of downwash angle with angle of attack. The intermediate and high tails (located 0.136 and 0.442 semi- span above the wing-chord plane) had a stabilizing influence at

6、low angles of attack; however, at moderate and high angles of attack large increases in the rate of change of downwash with angle of attack came 8 decrease in the stabilizing effect of these tails. The effect of the high tail actually became destabilizing at high angles of attack. The most favorable

7、 fmprovements in d however, in general, the effects of the fuselage on the stability of the wing were small. * The stabilizing contribution of the horizontal tail. can be predicted with a fair degree of accuracy from the air-flow survey data. 4 As part of a general study st the Langley 19-foot press

8、ure tunnel to determine the effect of a horizontal tail on the longitudinal sta- bili*y characteristics of swept wings, a low-speed investigation bs been made of a 520 sweptback wing in combination with a fuselage and a horizontal tail. The wing had symmetrical circular-arc sections, an aspect ratio

9、 of 2.84, and a taper ratio of 0.616. The longitudinal characteristics of the wing alone, wfth and without lesding-edge and trailing-edge flaps, are presented in reference 1. This paper presents results which show the effects of a fuselage and a horizontal tail (at various vertical positions) on the

10、 longitudinal characteristics of the wing with and without leading-edge and trailing- edge flaps. Results are also included of air-flow surveys made behind the wing at a longitudinal location which corresponded approximately to the longitudinal location of the tail. The data presented herein were ob

11、tained at a Reynolds number of 5.3 x 106 and a Mach number of 0.U. L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-3 c, - C Y S b C Y 9 P a v qt/ Q E a lift coefficient (7) drag coefficient e) pitching-moment coefficient, moment about 0.25F (z9 mea

12、n aerodynamic chord (M.A.C.) measured parallel to the plane area (wing unless otherwise noted), square feet span (wing unless otherwise noted), feet local chord (wing unless otherwise noted), feet spanwise ordinate, feet free-stream dynamic pressure, pounds per square foot mass density of air, slugs

13、 per cubic foot angle of attack (of wing chord unless otherwise noted). degrees free-stream velocity, feet per second ratio of local dynamic pressure at horizontal tail to free- stream dynamic pressure (unless otherwise noted) local downwash angle (unless otherwise noted), degrees local sidewash ang

14、le (inflow negative), degrees angle of incidence of horizontal tail measured with respect to wing-chord plane, positive when trailing edge is down, degrees Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 7 T tail stability parameter tall efficiency

15、 factor, ratio of position to q rate of change of pitching-moment coefficient with lift Coefficient - rate of change of pitching-moment coefficient due to tail da with angle of attack G)t lift-curve slope of isolated tail C qt rate of change of pitching-moment coefficient with tail incidence angle v

16、alue of c at zero lift for high tail position with mit flaps neutral 2 txtl length, distance from 0.aF to 0 .art, feet Z vertical distance, feet Subscripts : e effective t tail av average 0 value at zero lift of the wing MODEL AND APPARATUS The wFng plan form Etnd some of the pertinent dimensions of

17、 the wing are given in figure 1. The wing had an aspect ratio of 2.84, a taper ratio of 0.616, and symmetrical circular-arc airfoil sections prepen- dicular to the maximum thickness line. A straight line connecting the I Provided by IHSNot for ResaleNo reproduction or networking permitted without li

18、cense from IHS-,-,-* leading edge of the root and theoretical tip chord was swept back 52,05O. The maximum thickness of the airfoil sections parallel to the plane of symmetry was 6.5 percent chord at the root and 4.1 percent chord at the tip. The wing had neither geometric twist nor dihedral. The wi

19、ng was combined at zero Fncidence in a midposition with a fuselage of circular cross section (fig. 1). me fuselage had a fine- ness ratio of 10.2 and a llaximum diameter of 34.6 percent of the wing- root chord. The ordinates of the Fuselage are given in reference 2. Ftllets were not employed at the

20、juncture of the wing and fuselage. The model was tested with round-nose, extensible, leadingedge flaps which had a constant chord of 3.80 inches and extended inboard from O.9Bb/2 to 0.7=b/2 (fig. 2). These flaps were deflected 370 from the wing-chord plane in a plane perpendicular to a line Joining

21、the leading edges of the root and tips chords. Two typs of trailingedge flaps were used: Qle set located at the 80-percent-chord line are referred to as “split flags“ and the other set located at the 100-percent-chord line are referred to as “extended trailing-edge flaps .I Both types of trailing-ed

22、ge flaps were 20 percent of the wing chord and were deflected 600, as shown in figure 2. The split flaps and extended trailing-edge flaps extended outward approxi- mately 2j and 35 percent of the wing span, respectively, from the juncture of the wing and fuselage. The horizontal tail had 42.050 swee

23、pback at the leading edge, an aspect ratio of 4.01, a taper ratio of 0.625, and WCA 0012-64 airfoil sections -parallel to the plane of symmetry. The mounting arrangement of the tail allowed the tail to be secured at variaus vertical positions. The tail positions 0.44223/2 above, 0.136b/2 above, and

24、O.l32b/2 below the wing-chord plane (fig. 1) are referred to, respectively, as high, intermediate, and low. The vertical position of the tail ie deffned a8 the perpendicular distance from the wing-chord plane to the quarter- chord point of the mean aerodynamic chord of the tail. The incidence of the

25、 tail was measured with reference to the wing-chord plane and was changed by rotding it about the quarter-chord point of the mean aerodynamic chord of the tail. The accuracy of the measurement of the tail incidence angle is believed to be withfn i0.20. The air-stream survey rake of the Langley 19-fo

26、ot pressure tunnel was employed to obkin sidewash, dmwash, and dynamic pressure. The rake is composed of six Pitot-static tubes incorporating pitch and yaw orifices which were previously calibrated through a pitch range of f180 and a yaw range of *lP. A description of the rake is given in reference

27、3. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-6 TESTS NACA RM L51C30 Tests were made in the Langley 19-foot pressure tunnel with the madel mounted on the two-support system. All tests were de at a tunnel air pressure of approximately 33 pounds p

28、er square inch, abso- lute. The Reynolds number (based on the M.A.C. of the wing) was 5.5 x 10 6 and the Mach number was approxlmately O.U. Measurements of lift, drag, and pitching moment were made through a range of angles of attack *OM -bo to 320 and air-flow surveys were made at angles of attack

29、of approximately 3O, 8O, 13O, 16O, and lgo. The air-flow surveys were made in a vertical plane normal to the tunnel center line and were l.7lFbehind the quarter-chord point of the wing at 00 angle of attack. The plane of survey was selected as a compromise on the basis of the fore and aft variation

30、with angle of attack of the quarter-chord polat of the mean aerodynamic chord of the tail at various vertical positions. The maximum deviation of the tail quarter-chord point from the plane of survey occurred at high angles of attack. At the highest angle of attack (19O) the plane of survey correspo

31、nded to a tail length of 137F for the high tail and a tail length of 1.87E for the low tail. REDUCTION OF DATA Longitudinal characteristics.- The force and mment data prersented have been reduced to.nondimensiona1 coefficient form and have been corrected for the support tare and strut interference.

32、A correction for air-stream dsalinement has been applied to the values of angle of attack ad drag coefficient. Jet-boundary corrections to the angle of attack, drag coefficient, and pitching-saoment coefficient were deter- mined from reference 4 and are as follows: For configurations with horizontal

33、 tail off Em = 0.0024CL Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-WAC2 RM L51C30 . and for configurations with the horizontal tail on Ll hence, the method .of determining e and (qt/qIe was stmplified to where and C“it (%)o The effective values

34、of downwash and dynamic-pressure ratio for the low tail were not computed above an angle of attack of 200 because the reliability of the pitching-moment data for this particular angle- of-attack range and tail configuration are considered doubtful. Horizontal tail efficiency.- The tail efficiency fa

35、ctor is based on a value of I obtained in the region of zero lift with flaps neutral and with the tail in the high position. For this condition it %)O is on as assumed that the wing-fuselage combination has a negligible effect the flow over the tail. The tail efficiency factor q was obtained follows

36、 : Tail stability parameter.- The combined effects of downwash angle and dynamkpressure on the stabilizing contribution of the horizontal Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 NACA RM L51C30 tail is defined by the tail stability parameter

37、 7. The derivation of T is as follows: differentiating with respect to a where The values presented were obtained using the rehtlonskip . where (“Lacrf-vity location rear- ward of the quarter-chord pofnt of the mean aerodynamic chord. An analysis was. made to determine the effects of trim on the val

38、ues of T with the center of gravfty located at the quarter-chord point of the aa of the present investigation is mucimum, values of T are .- mean aerodynamic chord. It was found that, when values of . da significant, the changes in .t required to provlde trim were such that the product of these term

39、s produced only minor effects on the trends indicated by the variatione of T preeented. The values of T for the low tail are presented for nearly the entire angle-of-attack range; however, those values of T at angles of attack greater than approximately 20 need qualifying. It was not possible to det

40、ermine accurately - in the range above an angle of % aa attack of approximately 20; hence, the absolute values of T given for that range are open to question. It is believed, however, that the values presented reliably indicate the trends that exist. Ucal air-flow characteristics.= The air-stream su

41、rvey data have been corrected for jet-boundary effects by an angle change to the down- I wash and a downward displacement of the flow field relative to the wing- chord plane. c Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a Some air-flow condition

42、s were encountered which exceeded the limits of the survey-rake calibration. Data for these conditions were obtained from a linear extrapolation of the rake calibration. The inaccuracies introduced by extrapolating are believed to be relatively small for values Qf downwash angle less than 2p. The fa

43、ct-that the dynamic pressures measured outside of the wake at the highest angles of attack slightly exceeded unity may be attributed to the wake blockage in the closed tunnel. No corrections have been made for this blockage. I Average values of dynamic pressure and downwash.- For purpose of evaluati

44、ng air-flow survep at a particular vertical position, average weighted values of dynamic-pressure ratio and downwash were determbed for tail positions corresponding to those considered in the tail-on tests. The following were obtained : eqGtions-define the her in which theee values Pt/* Presentation

45、 of Results The longitudinal characteristics of the wing and fuselage are presented in figures 3 to 5. The results of tests of several wing configurations with a horizontal tail located at various vertical posi- tions .are presented in figures 6 to 8. A summary of the longitudinal stability characte

46、ristics of the wing with and without the horizontal tail is given in table I. The results concerning the characteristics of the air flow behind the wing are presented in figures 9 to 12 and table 11. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-.

47、NACA RM L51C30 11 Effect of Fuselage on Longitudinal Characteristics of the Wing me data presented Fn figure 3 show that the addition of a fuselage to the plain wing resulted in an increase in the value of mximrrm lift coefficient from 1.04 to 1.16. The fuselage caused only sml changes in the pitchi

48、ng mment through the lift range of the plain wing. With the 0.25b/2 leadfng-edge-flap configuration (fig. 4) , the nvsximum lift coefficient obtained with the fuselage on vas 1.26 as compared to 1.06 with the fuselage off. With the fuselage on, as with the fuselage off, a rearward shift of the aerod

49、ynamic center occurred at a lift coeffi- cient of approxhmtely 0.9. 5s shift of the aerodynamic center was less with the PJselage on than with the fuselage off. An increase in the maximum lift coefficient of the configuration with 0.23/2 leading- edge and 0.50b/2 extended trailing-edge flaps was realized with the additi