1、r;.-“-:, “ ; “. . . . . . . . . . ,I RESEARCH MEMORANDUM HOMZONTAL-TAIL EFFECTIVENESS AND DOWNWASH SURVEYS FOR FUO 477O SWEPTBACK WING-FUSELAGE COMBINATIONS H ASPECT RATI3S OF 5.1 AND 6.0 A FLFXNOLDS NUMBEROF 6.0 X IO6 -P 0 By Rein0 J. Salmi Langley Aeronautical Laboratory Langley Field, Va. 14 to .
2、T NATONAL ADVISORY COMMITTEE f - FOR AERONAUTICS WASHINGTON January 12, 1951 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1 w FOR TWO 47.p SWEPTBACK WINGFUSELAGE COMBINATIONS WITH ASPECT RATIOS OF 5.1 AND 6.0 AT A REXNOLDS NUMBER OF 6.0 x lo6 By R
3、eino J. Salmi An exgerimental iwestigation of the effectiveness of a horizontal tail on two 47.7 sweptback wing-fuselage combinations of aspect ratios 5.1 and 6.0 was made at a Reynolds numbr of 6.0 x 106 and a Mach number of . . 0.14. The tests were made with various combFnatiojls of lea-edge flaps
4、, . drooped-nose, split flaps, and double Blotted flaps in addition to .the flaps-neutral configurstiona. Air-stream surveys in the region of the tail were also made. The results corroborate previous law-speed investigations in that a tail position just beloy the extended wing-chord plane maintained
5、 effec- tiveness and improved the stability at the hi thue, the computations of the effective downwash angle and dynamic-pressure ratios were simplified to and where Tail efficiency factor.- In order to compare the effective values of qt/q to the average values, the tail efficiency factor q was esti
6、mated from the pitching-moment data. The factor was based on the rate of change of pitching moment with tail incidence angle. The inter- ference effects for the hi hence the values of 7 through that angle-of- da attack range are more nearly representative of the center-of-gravfty location at which t
7、he me thereby, the stabilizing effect of the tail is delayed. Effects of Trim on T As previously indicated, the values of T were obtained with a fixed tail incidence and a large out-of-trim condition existed at the high lift coefficients. The effects on T of changes in the tail load were therefore c
8、alculated and found to be of significant magnitude. The . changes in T were greatest for the combinations with double slotted flaps because of the large tail load required for trim and the deep wake behind the flaps. The effects on T of the changes in % required for.trimming are sham in figure 19 fo
9、r the wFn; of aspect ratio 5.1 with O.b0b/2 double slotted flaps and 0.4756/2 leading-edge flaps deflected. Inasmuch as a negative value of at was required for trim, an increase in qt/q resulted in a reduction of the tail effectiveness asd a decrease in qt/q increased the tail effectiveness. The cha
10、nges in T were significant only at the high angles of attack wfiich correspond to the lift range where the lift-curve slope is small. It has been ,foyd that through the angle-of-attack range for which the values of a%/) dff of the present wing are maximum, the values of T are applicable to a trim co
11、ndition for a center-of-gravity location rearward of the 25percent E. Air-Flow Surveys - The advantages of locating the horizontal tail in the region below the wing wake for improving the low-speed stability of the wing-fuselage- tail combinations at the higher angles of attack can also be shown fro
12、m an analysis of the contour charts of the air-flow characteristics (figs. 25 to 28). In using the contour charts to obtain the average values of qt/q and E for comparison with the effective values, discrepancies in the average values may result from the follorring causes: (1) The survey plane was p
13、erpendicular to the tunnel center line (see fig. 4), and, because of the tail sweepback and the forward and rearward movement of the tail with % angle of attack, the tail sections may be located at stations in the flow Provided by IHSNot for ResaleNo reproduction or networking permitted without lice
14、nse from IHS-,-,-12 NACA RM 5006 field different from those used in the calculations so that greatly different average values result where the %/q and E gradients are large: (2) where the values of qt/q and E were very large, extrapo- lations of the survey-rake calibrations were necessary (noted by
15、shaded areas on contour charts); and (3) no survey data were obtained in the region directly below the fuselage. The coqarison of the effective and average values of E and qt/q in table I1 shows that, in general, fairly good agreement was obtainea for the high tail position and somewhat poorer agree
16、ment for the low tail. I Inasmuch as the low tail did not reduce the unstable shift of the aerodynamic center prior to the maximum lift for the flaps-neutral com- binations, a brief investigation was made by using the contour charts to determine whether a more favorable tail location than the low po
17、sition would be indicated. The results indicated that a tail position of O.O5b/2 above the extended wing-chord plane was at least as good as the low position (-0.0%/2) that was tested but that any further increases in the stability of the wing-fuselage-tail combination would be small. Fairly good ag
18、reement between the C8lCulated and experimental pitching- monent curves for the low tail position was obtained. The contour charts indicate strong vorticity in the region of the tat1 at the high angles of attack. Large negative sidewash angles which developed in the region ab0u.t 40- to 50-percent s
19、emispa above the extended wing-chord plane decreased as the extended wing-chord plane was apprmched and finally assumed large positive values. References 4 and 6 pointed out that a vortex flow developed along the leading edge of the subject wings and that, at high angles of attack, the vortex flow w
20、as concentrated I over the inboard areas of the wings. The increased vorticity near the tail may, therefore, be caused by a more rapid rolling up of the vortex sheet due to the strong vortices formed at the leading edge. The vortex flow was not evident on a wing of lower sweepback (42O, reference l)
21、, and relatively smaller sidewash angles were obtained in the region of the tail (reference 7). With the leading-edge flaps deflected, both the downwash and sidewash angles are decreased at the high angles of attack. The large depression of the wake center line due to the trailing- edge flaps for th
22、e regfon directly behind the flaps can be seen from the contour charts. As the wake center line moves abwe the wing-chord plane with increasing angle of attack, only the tip sections of the high tail may be at the wake center, but the entire tail may be adversely affected by the broad wake. The loca
23、tion of the trailing vortices from the split and double slotted flaps can also be determined from the contour charts. The vortex from the trailing-edge flaps moved from a position well below the extended wing-chord plane to a point considerably above the extended wing-chord plane when the angle of a
24、ttack was increased. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-i NACA FI 5006 13 CONCLUDING REMARKS The main results of the investigation of the effectiveness of a horizontal tail on two 47.70 sweptback wing-fuselage combinations of aspect rati
25、os 5.1 and 6.0 are summarized as follows: The low tail increased the stabflity of the flaps-neutral cambina- tions at the very hi angles of attack but it ad not reduce the large unstable change of dCddCL (rate of change of pitching-moment coefficient .with lift coefficient) .WC occurred at an angle
26、of attack of 13O. Calculations 400. semispan and were considerably reduced for trailing-edge flap spans extending to 0.500 semispan and for combinations with only the leading- edge flaps deflected. The high tailposition was destabilizing at the high angles of attack for all the combinations tested.
27、Langley Aeronautical Laboratory National Advisory Committee for Aeronautics Langley Field, Va. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 RE-CES NACA RM 5006 1. Spooner, Stanley H., and Martina, Albert P.: Longitudinal Stabflity Characteris
28、tics of a 42 Swe tback Wing and Tail Combination at a Reynolds Number of 6.8 x log NACA RM L8E12, 1948. 2. Foster, Gerald V., and Fitzpatrick, James E.: Longitudinal-Stability Investigation of High-Lift and Stall-Control Devices on a 52 Swept- back Wing with and without Fuselage and Horizontal Tail
29、at 8 Reynolds Number of 6.8 x lo6. NACA 34 L8108, 1948. 4. Sam, Reino J. : Effects of Leading-Edge Devices and Trailing-Edge Flaps on the Longitudinal Characteristics of Two 47.70 Sweptback Wings of Aspect Ratios 5.1 and 6.0 at a Reynolds Number of 6.0 X lo6. NACA RM L50F20, 1950. 5. Eisenstadt, Ber
30、tram J.: Boundary-Induced Upwash for Yawed and Swept- Back Wings in Closed Circular Wind Tunnels. NACA TN 1265, 1947. 6. Salmi, Reino J., and Carros, Robert J.: Longitudhal Characteristics of Two 47.7O Sweptback Wings with Aspect Ratios of 5.1 and 6.0 at Reynolds Numbers up to 10 x lo6. NACA RM LWAO
31、4, 19%. 7. Furlong, G. Chester, and Bollech, Thomas V.: Downwash, Sidewash, and Wake Surveys behind a 42O Sweptback Wing st a Reynolds Number of 6.8 x 10 with and without a Simulated Ground. NACA RM -22, 1948. 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license fro
32、m IHS-,-,-NACA RM 5006 15 Modal configuration and tail heleht, z. above extended ring-chord plana A = 5.1 Tall off Plain sing -1 I 0 8 16 24 32 z = -0.05bD z = 0.3 R = 6.0 X 10 6 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-28 NACA RM 5006 3 0 t
33、 (a) LOW tail. (b) High tail. Figure 7.- Variation with angle of attack of the tail effectivnees parameter, effective downwash angle, dynamic-pressure ratio and rate. of change of downwash angle with R = 6.0 x 10 6 . - Provided by IHSNot for ResaleNo reproduction or networking permitted without lice
34、nse from IHS-,-,-/ .4 /.2 /.O .2 0 -. 2 -. 2 . Figure 8.- Effects of horizontal-tail position on the variation of CL and Cm with a for a 47.70 sweptback wing-fuselage combination of aspect ratio 6.0. Flaps neutral; R = 6.0 x 10 6 . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
