1、RESEARCH MEMORANDUM THE EFFECTS OF HORIZONTAL-TAIL HEIGHT AND A PARTlAL-SPAN LEADING-EDGE EXTENSION ON THE STATIC LONGlTUDINAL STABILITY OF A WING-FUSELAGE-TAIL COMBINATION HAVING A SWEPTBACK WING By Angelo Bandettini and F.-L 31 -_- cala.lqi$pJ- # - - .-s-m*- -_- -_- 533*-_ fL-JI-,-ses -_ 2 . 4” -I
2、c cU”WFBD- TMslmteimloonmm Morumuoothooftlmuoltsb Of tb esplOlu has, Two lfl, U.E.C., Boa. 7m ard TM, th ba!nmiunlon or y fyt . . mPnrvrbomutk.rlsadIspromblW. J NATIONAL ADVISORY COMM-ITTEE - FOR AERONAUTICii !r.r_%Y rERO;uAfflC4L tmMAtoRY 1 ?wfiY. WLC (2) the wing, body, and low tail fairing; (3) t
3、he wing, body, and high tail; (4) the wing, body, and low tail; (5) the wing with a 42-percent-span , 15-percent-chord leading-edge extension, body, and low tail. The majority of the da;a were obtained at a Reynolds number of 2,000,OOO at Mach numbers from 0.20 to 0.92. At a Mach number of 0.20, dat
4、a were also obtained at a Reynolds number of ll,OOC,OOO. Force measurements were made through a range of angles of attack of -loo to 25 for the model with the hfgh tail and -3O to 25O for the model with the low tail, except at the higher Mach numbers where the range was reduced by the limitations of
5、 WFnd-tunnel power and by choking conditions. The model with the unmodified wing was tested with stabFlizer incidences of O“, -2-1/2O, and -5O for both the low and high t the degree to which the sheet was rolled up at the location of the tail is not known. A further source of error probably lies in
6、the failure to take into account the effect of the fuselage, except insofar as it influences the wing lift distribution. Downwash angles were not accurately predicted at the higher angles of attack. It is known from previous studies that flow separation had Provided by IHSNot for ResaleNo reproducti
7、on or networking permitted without license from IHS-,-,-10 NACA RM A53JO7 occurred at the ting tip at the higher angles of attack with an accom- panying distortion of the span loading that could not be predicted by the theory. Wing-Wake andLocal DownwashMeasurements The dynamic-pressure loss in the
8、wake of the wing-fuselage combina- tion and the angle of local do wnwash near the horizontal tail are pre- sented in figures 11 through 14. Location of the wing wake.- The location of the wake has been determined from measurements of the total pressures behind the ting- fuselage combination at a pos
9、ition corresponding to 14 percent of the chord of the tail behind the tail leading edge and at three spsnwise stations. The results of these wake measurements are presented as the ratio of the decrement in dynsmic pressure at the tail to the free- stream dynamic pressure Aq/q as a function of vertic
10、al distance from the body center line. Data are presented in figure ll for angles of attack of O“, ho, 8O, 12O, and 160 at a Reynolds number of ll,OOO,OOO and a Mach number of 0.20, and in figure I2 for a Reynolds nuziber of 2,000,OOO and at Mach numbers from 0.20 to 0.92. Thetwovertical locations o
11、f the horizontal tail are identified as well as the wing chord plane extended. Of the three survey rakes used, two were located . within the tail semispan at positions 0.18 b/2 and 0.33 b/2 from the plane of symmetry, whereas the third was at 0.47 b/2 which was beyond the tip of the tail semispan. A
12、ccordingly, the vertical locations of the high and low tail have not been indicated in the figures pertaining to the rake at 0.47 b/2. At a Reynolds number of 11,000,000 and a Mach number of 0.20, the high tail was completely above the wake at angles of attack up through 160, whereas the low tail mo
13、ved into the center of the wske at an angle of attack of l2O. Throughout the Mach number range and at a Reynolds number of 2,000,OOO (fig. l2), both the high and low tail are seen to be outside the region of large wake losses at angles of attack of 8O and below. At an angle of attack of l2O (fig. 12
14、(d), the low tail had moved into the center of the wake, whereas the high tail was stillabove the wake except at the extreme tip. At 1.6 angle of attack, the low tail had moved below the center of the wake and the tip of the high tail had moved into the wake at the higher Mach numbers. At the higher
15、 angles of attack, the thickness of the wake and its displacement above the chord plane of the wing increased markedly with lateral distsnce from the plane of symmetry, especially at the htgher Mach numbers. This was a direct result of the separation on the outer portion of the wing semispan. Provid
16、ed by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACARM A53507 11 It is obvious from the foregoing that the lower tail passed through the wake in the angle-of-attack range up to 16O whereas the high tail was effectively above the wake except at the extreme
17、tip. The improvements in longitudinal stability at the higher angles of attack as a result of lowering the horizontal tail thus came about in spite of an unfavorable wake effect rather than because of any favor- able effect. A still lower tail position would probably benefit from a favorable downwas
18、h variation with angle of attack without being penalized by moving into a region of reduced dynamic pressure at the precise angle of attack where the maximum tail contribution to stabil- ity is desired. Local downwash measurements.- Figures 13 and 14 present the varia- tion of local downwash angle w
19、ith angle of attack for three spanwise stations at a Reynolds number of ll,OOO,OOO and a Mach number of 0.20, and at a Reynolds number of 2,000,OOO and Mach numbers from 0.20 to 0.9. It may be noted that only the innermost survey station (0.25 b/2) was within the extremity of the tail span, although
20、 the middle downwash station at 0.40 b/2 was just beyond the tip of the tail. The downwash survey was slightly above the low tail postion. A detailed study of local downwash in the region of the tail was not attempted during tests of the model with the survey rake. The downwash data obtained at the
21、survey-tube locations provided some information in regard to the spanwise distribution of downwash, partic- ulsrly as this distribution of downwash varied at high angles of attack of the model. At a Reynolds number of ll,ooO,OOO and a Mach number of 0.20 (fig. 13), ds.l/da was nearly constant up to
22、angle of attack of l2O and showed , little variation with spanwise location, except at the outermost survey station at the highest angles of attack. For low angles of attack, the data presented in figure 14 show little variation in the values of ds/da with spanwise location throughout the range of t
23、est Mach numbers. At the higher angles of attack, the values of ds/da increased with increasing spanwise distance at all the Mach numbers. The angles of attack at which the increases in as/da took place corresponded fairly closely with those shown in figure 8 for the effective downwash. Effect of a
24、Leading-Edge Extension A previous wind-tunnel investigation of this same model (ref. 4) has shown the effects of vsrious leading-edge chord extensions on the static stability of the model with the high tail. The most effective of these leading-edge chord extensions (0.15 c extension from 0.58 b/2 to
25、 tip) has been tested with the tail in the low position, and its effects Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 -1 NACA RM A53J07 on the longitudinal characteristics are presented in figures 15 and 16. At a Reynolds number of ll,oOO,COO a
26、nd a Mach number of 0.20 (fig. 151, the lift curve for the model with the leading-edge extension remained linear to a higher angle of attack than for the model with the unmodi- fied wing, resulting in an increase in maximum lift coefficient. At maximum lift the stall was mild uithout a large loss of
27、 lift. The effect of the leading-edge extension on the lift, drag, and pitching-moment characteristics of the model at various Mach numbers is shown in figure 16. The pitching-moment data (fig. 16(b) indicate that addition of the leading-edge chord extension eliminated or delayed to higher lift coef
28、ficients the forward shift of the center of pressure at Mach numbers below O.gO. The lift and drag data in figure 16 indicate an increase of lift-curve slope and a decrease of drag coefficient at the higher lift coefficients for the same range of Mach numbers. A com- parison of the pitching-moment d
29、ata for this model at these Mach numbers with data presented in reference 4 for the model with the high tail indicates a greater static margin at all the higher lift coefficients when the tail was in the low position. However, the effect of the leading-edge extension on the stability in the high-lif
30、t-coefficient range was slightly smaller with the tail in the low position (fig. 16(b) than in the high position (ref. 4). Similar observations were made from another investigation of leading-edge extensions with variable tail height (ref. 5). At a Mach number of 0.90, the lift coefficient at which
31、a forward shift of the aerodynamic center occurred was decreased slightly by lowering the tail although the total center-of-pressure movement was not as large as with the high tail. At a Mach number of 0.92, the pitching-moment characteristics remained essentially unaltered with addition of the lead
32、ing-edge extension. CONCLUSIONS An investigation has been made of the effects of horizontal-tail height and of a 42-percent-semispan, leading-edge chord extension on the longitudinal characteristics of a model with a 35O sweptback wing. The results of these tests and of air-stream surveys in the reg
33、ion of the horizontal tail indicate the following: 1. Lowering the tail from 22 percent to 8 percent of the wing semispan above the wing chord plane extended reduced the forward move- ment of the center of pressure of the model with the unmodified wing at moderate to high lift coefficients at all Ma
34、ch numbers up to 0.90 and at a Reynolds number of 2,000,COO. At a Reynoids number of 11,000,000 and a Mach number of 0.20, lowering the tail practically eliminated the for- ward center-of-pressure movement. The variation of calculated downwash angles with angle of attack indicated that there were ad
35、verse stability Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A53J07 13 effects due to a region of Large downwash at the position of the high tail and that they could be partially avoided by locating the tail in the low position. However, a
36、t Mach numbers of 0.90 and above, no large adverse effects of downwash were observed and the mriations of sta- bility with lift coefficient could be attributed largely to the longi- tudinal characteristics of the wing 8nd fuselage. Lowering the tail at these Mach numbers had little effect on the sta
37、bility. 2. Wake surveys in the region of the tail indicated that the efficiency of the low tail was reduced somewhat by the fact that it moved into the center of the wing wake at moderate angles of attack. 3. Addition of the wing leading-edge extension to the model with the low tall eliminated the f
38、orward movement of the aerodynamic center at moderate lift coefficients for Mach numbers up to 0.90, but prodded little change at Mach numbers of 0.90 8nd 0.92. Ames Aeronautic8l L8boratory National AdvAsory Comm%ttee for Aeronautics Moffett Field, Calif., Oct. 7, 1953 REFERENCES 1. Anderson, Seth B
39、., and Bray, Richard S.: A Flight Evaluation of the Longitudinal Stsbility Characteristics Associated With the Pftch-Up of a Swept-Wing Airplane in Maneuvering Flight 8-h Transonfc Speeds. NACA RM A5lEt2, 1951. 2. McFadden, Norman M., Rather-t, George A., Jr., 8nd Br8y, Richard S.: The Effectiveness
40、 of Wing Vortex Generators in Improving the Maneuvering Characteristics of a Swept-Wing Airplsne at Transonfc Speeds. NACARM A51Jl8,1952. 3. Bray, Richard S.: The Effects of Fences on the High-Speed Longitu- dinal Stability of a Swept-Wing Airplsne. NACA RM A5jF23, 1953. 4. Selan, Ralph, and %ndetti
41、ni, Angelo: The Effects of Lesding-Edge Extensions, a Trailing-Edge Extensfon, and a Fence on the Static Longitudinal Stability of 8 Wing-Fuselage-Tail CoIlibimtiOn mting a Wing With 35O of Sweepback asd 8n Aspect Ratio of 4.5. NACA RM A53E12, 1953. Provided by IHSNot for ResaleNo reproduction or ne
42、tworking permitted without license from IHS-,-,-NACA RM A53JO7 5. Morrison, William D., Jr., and Alford, William J., Jr.: Effects of Horizontal-Tail Height and a Wing Leading-Edge Modification Consisting of a Full-Span Fl8p and a Partial-Span Chord-Extension on the Aerodynamic Characteristics in Pit
43、ch at High Subsonic Speeds of a Model with 8 45O Sweptback Wing. NACA RM L53EO6, 1953. 6. Spooner, Stanley H., and Martina, Albert P.: Longitudinal Stability Characteristics of a 42O Sweptback Wing and Tail Combination at a Reynolds Number of 6.8 x 10s. NACA RM L8El2, 1948. 7. QueiJo, M. J., and Wol
44、hart, Walter D.: Wind-Tunnel Investigation of the Effects of Horizontal-Tail Position on the Low-Speed Longitu- dinal Stability Characteristics of an Airplane Model with a 35O Sweptback Wing Equipped with Chordwise Fences. NACA RM L5lEl7. 8. Sivells, James C., 8nd Salmi, Rachel M.: Jet-Boundary Corr
45、ections for Complete and Semispan Swept Wings in Closed Circular Wind Tunnels. NACA TN 2454, 1951. 9. Eerriot, John G.: Blockage Corrections for Three-Dimensional-Flow Closed-Throat Wind Tunnel 0.W . 2.24 E elttended 0.22 b/2 0.08 b/2 . . O0 9 -2-1/2O, and -50 0.273 Survey rake Total-pressure-tube l
46、ocations Longitudinal distance from quarter-chord point of wing to total-pressure tubes . . . . . . . . . . . . . . . . . 2.x) E Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-16 NACA RM A53507 TABLE I.- GEOMETRY OF TEE MODEL - Concluded Extent of v
47、ertical distance covered by total-pressure tubes in reference to wing chord plane extended Below wing chord plane extended . . , . . . . . . . 0.04 b/2 Above wing chord plane extended . . . :. . . . . 0.28 b/2 Spanwise positions of total-pressure tubes . 0.18 b/2(0.53 bt/2) 0.33 b/2 (0.95 bt/2) 0.47
48、 b/2(1.38 bt/2) Downwash-tube locations Longitudinal dfstance from quarter-chord point of wing to suryeytube 1.92 E Vertical distance of survey tubes above extended chordplane. 0.12 b/2 Spanwise stations of survey t M = 0.20. E Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure 4.- The lift, drag, and pitching-moment characteristics of the model with the tail in the low position at a Reynolds number of 11,000,000; M = 0.20. jll , ?8 1 . L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,
