NASA NACA-RM-A51B21-1951 Tests in the Ames 40- by 80-foot wind tunnel of an airplane configuration with an aspect ratio 2 triangular wing and an all-movable horizontal tail longit.pdf

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1、RESEARCH MEMORANDUM TESTS IN THE AMES 40- BY 80-FOOT WIND TION WITH AN ASPECT RATIO 2 AN ALL-MOVABLE HORIZONTAL TAIL - LONGITUDINAL CHARACTERISTICS David Graham and David G. Koenig Ames Aeronautical Labor ator y Moffett Field, Calif. g NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS WASHINGTON -;. Ne_rA

2、LRAy April 23, 1951 - -fLlk*tL LaeLzBAToRy rh!mky Ftt v* Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. c NACA RM A5 TESTsINTHES4BPOTLOFANALANE CONFIGORATION KITH AN ASPECT RATIO 2 TRIANGUUR WING AND AN AOVABLE EORIZON!KL TAIL - LONGITODINAL CEUUK

3、!TE%STICS ByDavid Gratiandavid G. Koenig . An investigation has been lnade to determine the effect of I a. thin, triangular, vertical tail; and a thin, unswept, all-movable hor- L izontal tail. The wing had an NACA 0005 Modified section and was equipped with partiaLspan, slotted, trailing-e flaps, T

4、ests were made with the horizontal tail at each of thee-vertical distances above the wing chordI ptie (0, 0.25, and 0.50 wing semispan) at me lmgitudinal distance behind the wing. TheReynolds nr,basedornthewingaerodgnami chord, was approximtely 14.6 X lo6 and the Mach number was 0.13. The results of

5、 the tests of the mcdel with the horizcmtal tail at each of the three vertical positions indicated that from a standpoint of longitudinal stability the most desirable position of those tested would bethatinthe extendedwingdhordplane. Downwash studies show that destabilizing aerodynamic-center varfat

6、ims, obtained with the tail in either of the other two positions, are the result of the downwash varia- ticms with angle of attack. Further tests to investigate the trim characteristics of the model with the horizontal tail fs the extended wing-chordplane imdicated that gliding speeds at a givenwing

7、lcading calculated for air-planes with and without a horizontal tail were/ for corawable attitude and static margtis, Iowerforthe a-lane yitha horizontal tail. I INTRODUCI?ICBl f Pheoreticalandexperimental studies showthatanairplanewitha lo+aspeotiatio triangQr wing would have desirable characterist

8、ics Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-2 NACARMA5lB21 . for flight at moderate supersonic speeds. At low speeds, however, the triangular wing has severalundesfrable characteristics which if not over- come will limit its use. These undesi

9、rable characteristics include low lift-drag ratios and high angles of attack at lllaximum lift coefficients. Thus anairplane utilizinga triangularwfngwould have highsinkingsnd landing speeds and abnormally high landing attitudes. . The foregoing consideraticms have neglected the problem of trim. Bec

10、ause of the characteristics of a triangular +.g, trim can be obtgLined by the use of trailing-edge flaps (tailless airplane) as well as by other Means such a6 8 conventional horizontal tail. The tailless design, how- ever, further aggravates the low-peed problems associated with the use of 8. triang

11、ular wing. This is indicated by the data of reference 1 which show that the negative flap deflections required to trim the airplane j Increase both the drag and angle of attack 8 high-fineness-ratio fuselage; a thin vertical tin of triangular plan form; and 8 thin, unswept horizon- tal tail, The pla

12、n form of the horizontal tail was made identical, and the airfofl section similar, to that of the wing the characteristics of which, throughout the subsonic Mach nuniber range, are reported in refer- ence 2. The choice of plan form was dictated by the following considera- tions: An all-movable horiz

13、ontal tail was chosen from a consideration of stability and control. The use of a triangular, al-chord plane at a fixed langi- tudinal position of the tail. Reported herein are the longitudi=l sta- bility and control characteristics of the various model cdigurations. NOTATION a free-stream angle of

14、attack with reference to the wing-chord plane, degrees . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA . b * bt horizontal-tail span, feet C c %l L D w Ef E E av it % L L/D M 9 qz S m 5121 wing spa, feet wing chord, measured mrallel to win

15、g center line, feet mean aerodynamic chord of wing, measured prallel to wing center line feet drag coefficient lift coefficient L 0 ss pitching-moment coefficient -5 ( qsc 3 total drag, pounds flap deectian,uredperpendiculartohinge line,degrees 1oc81d ownwash angle, degrees av-erage effectfve dovnw8

16、sh angle, degrees horizontal-tail incidence relatfve to the wing-chord plane, degrees distance from center of gravity to pivot line of horizontal tail, feet total lift, pounds lift-drag ratio totalpitchingmomentaboutthe center of gravitg, foot-pounds free-stream dynamk pressure, pounds per square fo

17、ot local dy33amic pressure, pounds per sq.uare foot Ha3 a-Y square feet Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-4 St % VS X Y Z . 4 mcx m 5121 horizontal-11 area, square feet gliding speed, miles perhour sinking speed, feet per seoriztal tail

18、 8s of unswept plan form and had a modified d1acm.d sectian. The original diamond section of Y. the resulting section had a =imum thiclmess of 4.2-percent chord. Ths three positions of the horizontal tail used were, namely,.a low, middle, asd high position as sham in figure 1. The tail was pivoted a

19、bout a line connecting the leading edges of the tip secticms. In the low positia, the horizontal tail was mounted on the fuselage with its pivot line in the extended chord plane 02 the wing. In the middle and high positians, the horizontal tail was mounted on the vertica-1 tail with the pivot line l

20、ocated vertically at Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A51B21 5 . approximtely2+and 50-percentwingsemispanabove the wing-chordplam, respectively. The longitudinal location was the same for all three tail positions. (See fig. 1.)

21、 W same tail-surface panels were used. for the three positions. Consequently, the tail aspect ratio was larger with the tail at the low positian than at the other two positions (4.4 and 4.0, respectively). Force and mament data were obtained for the model with t.be horizan- tal tafl at each of the t

22、hree positians and with the horizontal tail off. The tail was set at O“ and at -6O incidence at each of the three tail positions. All tests were made at zero sideslip. Flap deflections of O“ and 40 were used. The tests were made through an angl-fdttack range of -lo to 26O, except for tests with the

23、horizontal tail in the high position where the angle of attack was limited to a maximum of 18O due to structural limita- tions of the model. Additional tests were lnade with the tail at the low position in order to determfne the longitudinal trim characteristics of the model. Data were obtained for

24、a range of horizontal-tail incidences from +l“ through -100 for flap deflecticxns of O“ and 400 , and for tail incidences of O“ and -loo for flap deflectiane of loo, 20, and 30. The Reynolds number of the tests uas 14.6 milli= based on the mean aerodynamic chord of the wing. The dynamic pressure was

25、 approxinrately 25 pounds per square foot and the Bcoefficient range, then became and remained unstable (0.4p forward shift of the aerodynamic center) through the remainder of the lift range of these tests. These changes in stability were the result of the manner in which . the downwash behind the w

26、ing varied with angle of attack, as can be seen by a ccanrparison of the pitchingaoment curves of figure 5 with the corre- ponding downwash curves of figure 6(a). These downwash variations are w substantiated by the survey downwash curves, shown in figure 6(b), since Provided by IHSNot for ResaleNo

27、reproduction or networking permitted without license from IHS-,-,-8 6 d XACAiMA5til they have nearly the same shape as the corresponding curves determined from the force-test data. It can be seen that the changes in stability are related to changes in deadda. For instance, with the tail in the low p

28、ositian, the stability of the model began to increase in the angle- of-attack range above 50 when deav/du began to decrease. With the tail in either the middle or the high positions, the instability of the model occurredwhen de v/ ne 0.59 140 45 Airplane with tail, 8f = 4o“ 1.04 104 50 %/S = 30 poun

29、ds per 6-e foot. It can be seen that, by use of trailing-edge flaps as lift-producing devices and an all-movable horizontal tail as a lcngitudinal stabilizer and control, a 2 a = 24 -96 107 56 Airplane with tail, Ef = MO, a = 16 1.06 102 49 Airplane with tail, Ef = ho, a = 240 1.44 86 54 q/s = 30 p-

30、6 per square foot. These results show that if low static margins are acceptable and if no ground-g10 limitation is placed on the tailless model, then there is little to chose between the two from low-speed considerations. Possible weight and drag penalties incurred by use of a horizontal tail may th

31、erefore dictate against its use. Finally, however, it should be noted that there is no certainty that optimum flaps were tested on the model .- with a tail; hence the possibility remains, for this model, of making , substantial improvements through refinements in flap desia. The values quoted in the

32、 foregoing tables are pletely indica- : tive of the landing characteristics, since ground effects were not taken into account. As shown in reference 5, v-erg-large ground effects can be expected for a lowdspect-ratio triangular wing. Sizable increases in both CL and L/D at a given angle of attack we

33、re obtained when the wing was within a semispanof the ground. These increases in CL and L/D would result in decreases in both the gliding and sinking speeds of both configurations. No estimation of the ground effect on the effective- ness of the slotted-trailing-edge flap can be made at this time bu

34、t, as shown in reference 5, there is a slight increase in flap effectiveness of split flaps on a triangular wing in the presence of the ground. . An estimte of the tail incidence necessary to trim the airplane, when near the ground, was made in order to determine whether the required incidence6 woul

35、d be excessive. The change in pitching moment obtained for the triangular wing with split flaps in the presence of the ground (reference 5) was assumed to apply to the wing with the slotted flaps. In addition, in order to be conservative, it was assumed that the down- wash at the tail is entirely el

36、iminated by the presence of the ground, Based on these assuurpticns and a 6-percen.t static.gin, the required incidence of the tail would be 422O when the wing is at an angle of 16O with flaps.deflacted 40 and at a distance of one semispan above the gr(Ju. This incidence.is not considered to be exce

37、ssive. The angle of attack of the tail relative to the local stream would be -6O. - 3 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-mm 533321 - - 11 . Another point worthy of note is that, although the theoretical induced drag of a wing with aspe

38、ct ratio 2 is over 15 percent greater than the induced drag of 6,n aspect ratio 2.3 wing, the total drag at trim lift coefficients above 0.4 of the configuration with the all-movable horizontal tail and aspect ratio 2 wing is less than the total drag of the tailless airplane with the aspect ratio 2.

39、3 wing, awin the cases where a 6-percent statfc margin was chosen. tail The results of the investigation of the model with the at each of the three vertical positions indicated that CONCLUDING RIENRKS point of longitudinal stability the most desirable position referring to horizontal from a s-tad- o

40、f those tested is that in the extended wing-chord plane. With the tail in this position, the model had a stable aerodgnamic-center variation throughout the lift range; whereas the model with the tail in either of the two pos- itimLs above the chord plane had large destabiatim.at Low Speed of a Iarge

41、Scale Triangular Wing of Aspect Ratio Two. - II. The Effect of Airfoil Section Modifications and the Determination of the Wake Downwash. NACA RM A7H28, 1947. 4. Anderson, Adrien E.: Chordwise and Spanwise Loadings Measured at Low Speed on Large Triangular Wings. NACA RM AgB17, 1949. 5. Rose, Lemati

42、M.: Loation of a Small Triangular Wing of Aspect Ratio 2.0. I -The Effect-of_C.qribition with a _. .- _ Body of Revolution and Height Above a Ground Plane. NACA RM A7KO3, 1948. CONFIDENTLAL Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. NACARMA5lB

43、21 13 TABLE I. GEOMIPISIC DATA OF MODEL Area.spuarefeet 3X.5 Span,fsd . 25.00 Mean aerodynamic chord, f5 . g:Ci 9.50 90.00 85.00 80.00 75.00 70.00 65.00 60.00 55.00 - T Radllls 0 .26 .42 -70 1.15 1.54 1.86 2.41 2.86 3.22 ;:72 . Z:,“ NACA RM A5lB21 . Provided by IHSNot for ResaleNo reproduction or ne

44、tworking permitted without license from IHS-,-,-Dimensions shown in feef UnhSS Otherwise noted c.g. Loeotion . Tail poaitlon x Low P.9.3 Middle 3.57 Mgh 4.43 Fure /. - Geometric details of the model. I I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-

45、,-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I . Pigure2.-Thf3mdelasmoun St, og /;, 0”. - - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, . :4 P $0 0 4 8 I2 16 20 24 28 Angle of otkck, e, d

46、og I 1 . 32 28 24 PO 16 , I IQ 0 4 8 12 16 20 24 Qt? Anqfe of attuck, e, deg 20 . 16 8 4 0 0 4 8 I2 16 PO 24 28 A&? of attack, a, deg . F&W 6 VI-IMNI d ovumpa eff&he ohwash mgib mti mgle of attack. 4, 0. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-

47、,-2 0 -2 Httio L tiiiiitiiiiiiiii&TiiiiiiiAi/9PTk”iil I I I I I IA I “l/j I I I I I I I I VI Id I/- -4 0 4 8 12 I6 20 24 L39 A5 I2 08 .04 0 -.04 -08 -.&? -/6 -.20 -24 Angfe of ottack, a, deg Fmi7g-.mament coefftim( c, 0 .I 2 3 .4 5 .6 7 8 bag iwetxicknt, c, . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-0 -4 0 4 8 I2 16 20 24 A6 J2 A9 .04 0 -.M 98 -.&? -16 20 94 Angle of attack, a, deg p-menf ctxffisn c, 0 J 2 3 .4 .5 fi 7 8 Dmg &ficient, c, Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-