NASA NACA-RM-A9C21-1949 Aerodynamic study of a wing-fuselage combination employing a wing swept back 63 degrees effects of split flaps elevons and leading-edge devices at low spee.pdf

1、L RM No. A9CZi RESEARCH MEMORANDUM AERODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING SWEPT BACK 63O. - EFFECTS I I OF SPLIT FLAPS, ELEVONS, AND LEA EDGE DEVICES AT LOW SPEED By Edward J. Hopkins Arne s Aeronautical Laboratory Moffett Field, Calif. c I I .DING- ! NATIONAL ADVISORY COM

2、MITTEE FOR AERONAUTICS WASHINGTON my 19, 1949 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NATIONAL ADVISORY COMMITTEE FOR AERONAUTICS By Edward J. Hopkins An investigation was conducted to evaluate the effects of split flaps, elevons, sharp leadi

3、ng edges, drooped-aose flaps, and extended- nose flaps on the lift, drag, and pitching-moment characteristics at low speed of a wing-fuselage codination having a wing with the lead- ing edge swept back 63O and having an aspect ratio of 3.5. Measure- ment s were also made of the rolling moments produ

4、ced by the-elevons . In addition, 8 study wa6 made to evaluate the effects of the fuselage and possible Reynolds nuniber effects on the characteristics of tkre wing. The optimum chordwise posttion of the split flap for increasing the lift coefficient attained before the occurrence of longitudinal Fn

5、stability and for reducing the drag at high lift coefficients was the position with the split flap hinge line coincident with the trailing edge of the wing. The effectiveness of the elevons for producing rolling moments was nearly constant up .to an angle of - attack of go, but decreased at greater

6、asgles of attack. The full- span leading-dge flaps Increased the lift coefficient attabed before the occurrence of longitudinal instability considerably more than did the 50”percent span leading-edge flaps The extended-nose flap was about twice as effective as the drooped-nose flap in reducing the d

7、rag of the model at the higher lift coefficients. INTRODUCTION A coordimted program is being conducted st -8 Aeronautical Laboratory to provide information throughout an extensive range of Mach and Repolas nmibers on a wing-fuselage conkination employing Provided by IHSNot for ResaleNo reproduction

8、or networking permitted without license from IHS-,-,-2 - HACA RM NO. 921 a wing with the leading edge swept back 63 asd having an aspect ratio of 3.5. According to the tboretical considerations of refer- ence 1, a wing of this plan form should be ca3able of .efficient flight at supersonic Mach numbe

9、rs up to 1.5. Experimental results from tests of wings of this pk form at high Mach or Remolds nders Elre presented in references 2, 3, and 4. A wing-fuselage conibination having a wing of the plan form just described was investigated in one of the Am36 7- by 10-foot wind tunnels to evaluate the eff

10、ectiveness of varfous flaps and particu- larly their capacity for e1“ting the Imge changes in the longi- tudinal stability which have been found to OCCUT above a lift coef- ficient of 0.4 (reference 4). In this connection, a drooped-nose flap and 811 extended-ose flap were tested in conjunction with

11、 trailing-dge flaps. Furthermore, an investigation wa6 made to determine the optimum chordwise position of split flaps and the effectiveness of elevons of two different plan forms. NOTATION All forces and moments are referred to the wind axes with the origin on an extension of ,ths wing root chord s

12、t the same longi- tudinal position 8s a point at 25 percent of the wing mean aero- dynamic chord. coefficient lift coefficient (p) rollingament coefficient pitching-moment . coefficient “I . . (pit chiy=-nt) , . aspect ratio (% span of semispan wing perpendicular to the pke of sptnetry, feet wing ch

13、ord paraillel to plane os symmetry, feet X Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA m NO. 921 3 L . ZW wing loading, pounds per square foot g free-stream -ic pressure (sV2) , pounds per square foot R r s V VS X Y a 6 v P Reynolds nuniber

14、fuselage radius, feet axe8 of semispan win;, squaxe feet free-stream velocity, feet per second sinking speed, feet per second longitudinal distance, feet lateral distance, feet angle of attack of the wFng chord plsne, demees control-surface deflection lneasured in a plane norms1 to the hinge line (F

15、or positive deflectiona, the flap is below the wing-chord plane. ) degrees kinematic viscosity of sir, feet aquared per second mass density of air, slugs per cubic foot Subscripts d drooped-ose flap e elevon Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IH

16、S-,-,-4 f split flap i induced NACA RM No. A9623 n extended-nose flap u uncorrected CORRECTIONS An expkumtion of the mthod uS8d in Calculating the Wind- tunnel- corrections which were applied to the data ie given in the appendix. The equations used in correcting the data are as follows: c, = + 0.001

17、0 C however, only a negligible change in angle of attack of the wing tip was measwed. Evidence that the effecte of model distort-lon were negligible was also obtained from tests of this model in the Anms 32“foot pressure wisd tunnel (reference 3) at dynamic pressures of 53 and 105 pounds per square

18、foot for a constant Reynolds nmiber of 9.75 x los. Only small affects on ths aerodynamic characteristics of the wing were produced by this dyaamicesme variation. Hence, no corrections have been applied to the data of the- present tests for the effects of model distortion. Provided by IHSNot for Resa

19、leNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. AgC21 5 The semispan wing used for this investigation had its leading edge swept back 630, an aspect ratio of 3.5 based on the geometry of the complete wing, a taper ratio (ratio of tip chord to root chord) of 0.25, n

20、o twist, no dihedral, and the NACA however, to investigate possible dynamic- scale effects the data preeented in figure.8 were obtained throughout a Reynolds rimiber range of 2.5 to 7.2 million. Increasing the Reynolds nurdber from 2.5 to 4.2 million increased the lift coef- ficient attained before

21、the occurrence of longitudhl instability of the wing with the long fuselage from about 0.4 to 0.5, but had a negligible effect on this lift coefficient of the plain wing. However, a further increase of Reynolds number to 7.2 million resulted in no improvement of this lift coefficient. The drag coef-

22、 ficients were reduced slightly for all lift coefficients between 0.1 and 0.8, but the lift-curve slope was not greatly affected by increasing the Reynolds number from 2.5 to 7.2 million. The effect of the 0.25-chord split flap in several chordwise positions on the characteristics of the model is sh

23、own in figure 9. The split flap with its hinge line at the trailing edge of the wing yielded the largest increment of lift coefficient for all angles of Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-8 I CA RM NO. 921 attack and flap deflections inv

24、estigated (an increment of at least 0.4 up to an angle of attack of 24O) and increaaed the lift coef- ficient attained before the occurrence of longitudinal instability from about 0.5 to 0.8. As the hinge line of the split flap was moved forward from 100 to 40 percent of the wing chord, the flap eff

25、ectiveness decreased rapidly. Although the split flap with its hinge line at the wing trailing edge produced the largest lift increaees, this flap also produced the largest changes in longi- tudinal balance. The split flaps with their binge lines normal to the air dream provided less negative pitchi

26、ng momnts and smaller lift increnrents than did the split flap with its hinge line at the trailing edge of the wing (fig. g(a) . With the split flap deflected 60 in the forward position with its hinge normal to the .air stream the lFft coefficient attained before ths occurrence of longitudinal insta

27、bility w88 increased from about 0.5 to 0.6. Surface tufts tndicated that the split flaps with their hinge lines norm1 to the air stream caused flow separation to occur initially near the midsemispan of the wing at an angle of attack of 00. At angles of attack greater than U0, these split flaps cause

28、d a larger portion of the wing to stall, which is probably responsible for the decreased lift-curve slope and the decreased maxirmun lift coefficient (fig. 9 (a). Increase of the deflection of the split flaps from 45 to 75O caused relatively small changes in the lift and pitchingdloment characterist

29、ics (fig. 9( a). Deflecting some of the split flape more than 45O, for example, the flap hinged at 100 percent of the wing chord, decreased the. maximum lift coefficient. Only the split. flap at the trailing edge of the wing greatly reduced the drag of the model at high lift coefficients (fig. 9( b)

30、 ) . The data for the del with the split flap of triangular plan form and with the split flap of constant;percent chord (both hinged dong the wing trailing edge) are presented in figure 10. At high angles of attack the split flap of triangular plan form produced slightly larger increments of lift co

31、efficient %has. the eplit flap of constant-percent chord of the ame area. Longitudinal insta- bility occurred at approximately the same lift coefficient with the same deflection of either flap, but the split flap of triangular plan form deflected 45O produced slightly less negative pitching moments

32、at smll angles of attack. With either flap at Oo angle in the extended position, the lift-curve slope was increased from 0.046 to 0.052 per degree and the aerodynamic center was shifted rearward 1.5 percent of the man aerodynamic chord at -11 lift cpefficients. Provided by IHSNot for ResaleNo reprod

33、uction or networking permitted without license from IHS-,-,-The drag characteristics of the model with tlu3 triangular flap were similar to those of the model with the 25-percenkhord split flap for the sam= flap deflections (fig. 10(b). Elevans The characteristics of the model with various deflectio

34、ns of the constant-percent-chord elevon and the constast-chord elevon are presented in figure ll. The pitching mnts with the canstant-chord elevon undeflected were slightly different from the pitching moments with the constant-percent-chord elevon undeflected. Similar discrep- ancies may be found in

35、 other figures of this report. These discrep escies are believed to have been caused by slnsll differences in the contour or in the Oo settings of the various controls. At snadl angles of attack, the rates of change of pitch“ and rollie mmnt coefficients with elevon deflection for the two elevons we

36、re approximately in proportion to their area mmsnts about the pitch or roll axes. The rate of change of lift, pitching-moment , and rolling-mnt coefficient with elevon deflection remained nearly conatant up to an angle of attack of go, decreased between angles of attack of go and 17O, but increased

37、st higher angles of attack for negative deflection of the elevons. In the low lift range, the rate of change of pitching-nent and rolling+uoxnent coefficients with elevon deflection decreased as the negative deflection of the elevon exceeded 30. The cbacteristics of the model with the constan-hord e

38、levon and the 0.25-chord split flap deflected 45O at the wing trailing edge are presented in figure 12. This was the elevon split-flap conibina- tion tested with the leading+dge flaps which will be discussed in the succeeding sections of this report. “F+e rate of change of lift and rolling-naoment c

39、oefficients with elevon deflection remained nearly constant to an angle of attack of 5O, but decreased at hfgher angles of attack. Therefore, with the split flap the effectiveness of the constant-chord elevon began to decrease at a smaller erngle of attack than without the split flap (figs. ll and 1

40、2) . As mentioned herehibefore the split flap hinged at the wing trailing edge produced large changes in belance; therefore, the SThe mmnt of the area of the consbt-chord elevon about either the pitch or the roll aria x88 approximately 1.5 thee that of the constan+percenbhord elevon. Provided by IHS

41、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-10 L HclCA RM No. AgC21 longitudinal-stability -gin should be considered in choosing the type of control surface to be used for balance. With the negative deflection of the elevon limfted to bo, the split flap deflect

42、ed 45 and the center of gravity at 0.25F, the wing-fusela conibimtion could be balanced only for lift coefficients zrp to 0.8fig. 12). However, it appears possible to we a mre resl-ard genter of gravity and still to maintain adequate longitudinal stability at the lower lift coefficients. With a more

43、 rearward center of gravity, the wing-fuselage combination could be balanced at all lift coefficients with ,smaller elevon deflections, thus allowing more elevon effec- tiveness for lateral control. Leading4dge Devices The model was tested wtth both the drooped-noee flap snd the extended-ose flap de

44、flected 30, 35O, bo, and wo. The optimum deflection for either flap was found to be aboyt 40. A6 only slight differences were noted Fn the result8 for the eeveral deflections, only the results for the bo deflection are presented. The model wa8 alao investigated with each of the leading+dge flaps in

45、arioue conibkations with the constanhhord elevon mdeflected and deflected negatively 20, and the 0.2-hord split flap undeflected and deflected 45O at the trailing edge of the wing. The characteristics of the model with the drooped-nose flap of 50-percent wing span and of full wing span are presented

46、 in figures 13 and 14, respectively. The drooped-se flap of 50-percen-t wing span decreased the lift at all angles of attack and failed to improve the pitching*oment characteristics of the del (fig. 13). However, the drooped-nose flap of full wing span gave slightly better results, increasing the li

47、ft coefficient at which long-ltudinal instability occurred about 0.15 with the split flap retracted and about 0.04 with the split flap extended (fi . 14). With the elevon deflected -40, the split flap deflected 45! and the drooped-nose flap of full wing span deflected bo, the lift coefficient attain

48、ed before the occurrence of longitudinal instability was greater than 1.0 (fig. 14). The characteristics of the model with ths extendedilose flap of 5-rcent wing span and full wing span are presented in figures 15 and 16, respectively. Th.e extended-nose flap of 50”percen-b wing span proved to be as

49、 ineffective a8 the drooped-nose flap of 50- percent whg spsn for increasing the lift coefftcient of the model before the OccuITence of longitudinal instability ( f ige. 13 and 15) . However, with the split flap deflected 45 and the elevon deflected -POo, the extended-nose flap produced a more nearly linear variation I Provided by I