REG NACA-RM-A8D20-1948 Aerodynamic study of a wing-fuselage combination employing a wing swept back 63 degrees subsonic Mach and Reynolds number effects on the characteristics of n.pdf

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1、DUM mRODYNAMIC STUDY OF A WING-FUSELAGE COMBINATION EMPLOYING A WING SWEPT BACK 63.-SU88ONIC MACH AND REYNOLDS NUMBER EFFECTS ON THE CHAR- ACTERISTICS OF THE WING AND ON THE EFFECTIVENESS OF AN ELEVbN By Robert M. Reynolds and Donald W. Smith NATIONAL ADVISORY COMMITTEE Provided by IHSNot for Resale

2、No reproduction or networking permitted without license from IHS-,-,-NACA RM No. A8IS20 By Robert M. Reynolda and Donald W. smith A wind-tunnel Investigation ha8 been made of a semispan model . of a wing swept back 63 having an aspect ratio of 3.5 and a taper ratio (tip chordlroot chord) of 0.25. 12

3、leee tests were conducted to evaluate the effects of Reynolds mber and MEtch rnrmber on the aerodynamic characterietics of the wfng. Ihcluded In the inveeti- stion were measurements of the effectivenesa of an elevon used as a longitudinal control. The aerodynamic center of the King shifted rearward

4、near a lift coefficient of 0.2; whereas above a lift coefficient of approximately 0.4 there was an abrupt forward ehift of. the aerodpmmlc center. Increase of the Mach number frm 0.m to 0.925 at a constant Reynolds number of 3.55 million resulted in a gradual increaee of the lift-curve elope at zero

5、 lift frm 0.043 to 0.048 per degree, a rearward shift of the aerodgnamic center at zero lift frm 42.4 percent to 44.6 percent of the mean aerodynamic chord, and a decrease of the maximum lifMO-drag ratio from 18.0 to 14.7. The elevon pitch effectivenese (rate of change of pitchhg+acment coefficient

6、per degree of elevon deflection) had a value of -0.0053, and wa8 not appreciably changed by varying the Mach number fram 0.60 to 0.90 at a comtant Reynolds number of 2.26 million. An increase of Repolde number fram 4.U to 9.85 at an approxi- mately constant bkch number and at a constant dpmic pressu

7、re of 50 pounde per square foot caused a reduction of the lift-curve elope from 0 to 0.042, and an increaee of the maxirmrm lift-to-drag ratio from 17.7 to 20.6. There me little shift of the aerodynamic UNCLASSIFIED Provided by IHSNot for ResaleNo reproduction or networking permitted without license

8、 from IHS-,-,-2- - NACA RM No. Am20 center at zero lift from 42.4 percent of the mean aerodynamic chord. The elevon effectiveness was little affected by changes in the Reynoliir, nuniber. Recent developments extending the theory of supersonic flow to the analysis of swept wings of finite aspect rati

9、o (reference I) have indicated that efficient flight at kch nunibere q to 1.5 may be achieved by the use of a large sweepback angle together with the higheat posaible aspect ratio. A wing deeigned according to the indications of reference 1 ie being investigated extensively at the Ames Aercmautfcal

10、Laboratory to evaluate its behavior over a wide range of Mach and Reynolds numbers, both alone and in the presence of a slender fuselage. The series of tests performsd in the .U+foot pressure wind tunnel and reported herein is part of a coordinated program abed at the ult-te development of a configu

11、ration adaptable to efficient, eco- nomical flight at Mach nunibera ug to 1.5. This report presents the sibsonic serdpamic characteristics of a semispan model of a Xing swept back 63O as influenced by the independent variation of lkch and Reynolds nunibem. Also included are data on the effectiveneas

12、 of a constantchord elevon. SYMBOIS The following synibols are used in tus report: pitch-mnt coefficient wing mean aerodpmlc chord angle of attack of wing chord, degrees angle between wing chord and slevcrn chord, measured in a plane perpendfcular to the elevon hinge line, positive for downward defl

13、ection with respect to the wing, degrees . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EACA RM No. Am9 3 Q dynamic pressure ($1 , pounds per square foot V airspeed, feet per second P mss densitr of air, slugs per cubic foot P viscosity of air, sl

14、ugs per foot-second e. speed of sound, feet per second s wimg no provfsion W simultaneously at lift coefficients near 0.2: (1) There is a perceptible increase in the slope of the lift curve; (2) there is a sharp increase in the rate of change of drag with lift; and (3) there is a moderately abrupt r

15、earward shift of the aerodynamic center. n reference 6, this nonlinear behavior of the characteristics is attributed to separation from the airfoil leading edge, with a consequent loss of lead- suction and a rapid increase in drag. The 0.2 value of the lift coefficient at whfch the effects of separa

16、tion appear, as shown in reference 6, agrees with the indications of oblique-wln; theory wbich predicts tus behavior at a lift coefficient equiva- lent to the section nraximum lift coefficient reduced by the square of the cosine of the sweep angle. With regard to longitudinal stability, the data ind

17、icate a behavbor typical of hi hly swept wings and already reported extensively elsewhere $e.g., reference 7). This behavior, illustrated in figures 3(c) d.4(c), is ,the forward shift of the aerodynamic center which in this case occurs at lifi coefficients of the order of 0.4. As discussed in refere

18、nce 7, this longitudinal instabflity is primrily dependent upon the psrticular cordbination of sweep angle and aspect ratio. Also of interest are the drag data of figures 3(b) d 4(b) , which show the influence of the lowLdrag King section betxeen lift crxfficfents of -0.1 to 0.1. Effect of Mach numb

19、er.- DeLta for Reynolds nunibera of 2.35 and 3.55 million ad Mach numBers from 0.160 to 0.925 are presented in figures 3 and 4, and the effects of Wch nuniber are surmnecrized in figures 5 to 8. The data show na abrupt changes with increasing Mach mer. Provided by IHSNot for ResaleNo reproduction or

20、 networking permitted without license from IHS-,-,-XACA RM No. Am20 7 In figures 3(c) and 4( c), it is of interest to note that, as the Mach nmiber is increased,the chasge to a positive variation of pit-nt coefficient with lift coefficient does not occur as rapidly as at the low Mach rider, but take

21、s place ov8r an increasingly wide range of lift coefficients. In figure 5, are shown the I=, drag, and pitchinginaoment coefficients as a function of Mach rmpiber. It will be seen that there is a gradual increase of the lft coefficients at constant angles of attack with increasing Mach mer. The gene

22、ral -trend is for the drag coefficient to increase witk increasing Mch rumiber over the range of lift coefficients plotted; hrmever, there is a decrease of the drag coefficient between Mach riders of 0.7% to 0.925 for lift coefficients of 0.35 and 0.4. At lift coefficients less than 0.25, increasing

23、 Mach nuniber caused little change in the pit- moment coefficients; however, at the higher lift coefficients, the pitching-moment coefficients becams less negative as the Wch nuniber was increased. The effect of compressibility on the variation of lifMo4rag ratio with lift coeff fcient is shown in f

24、igure 6. It is evident that there is a coneiderable decrease in the value of the nmr-lmrlm lift-to- drag ratio with increasing Mach nuplber. Ih figures 7 and.8, are summarized the variations with Mach n (2) there is a progressive reduction of . drag coefficient with increasiq lift coefficient; and (

25、3) there are changes in the measured pitching-monmnt coefficients. The increase in Re-grllold.8 number -roved the lift-t-ag ratio as shown in figures 6 and 8. Lift, drag, and pitching-momsnt data for the wing at Reynolds nilmbers of 2.35, 4.10, 7.40, and 10.30 million for a conetant hkch number of 0

26、.180 are compared in figure 9. Lncrease of the Reynolds n-er nrty be seen to have the effects discussed in the previous paragraph. .More clearly shown in this figure, however, is the effect of increasing Beynolda ndber in extend“ the lfnear- variation of the pitching-moment coefficient with the lift

27、 coeffi- cient to higher lift coefficients. Whereas the tests at a Reynolds number of 2.35 mLXLion indicate separation.beginning at a lift coeffi- cient of about 0.20,for Reynolds nunibers of 4.10, 7.40, and 10.30 million, the separation is delayed progressively to llft coefficients of approximtely

28、0.25, 0.30, and 0.35, respectively. Because of its particular conformatioa of sweepback, aspect I. ratio, and thin wing section, the model was susceptible to considera- ble bending under lifting loads. Since the magnitude of the deflection is directly proportiopLl to the dynamic4 pressure, it was th

29、ought advisable to vary the Reynolds nuniber while keeping the dynamic presswe constant. Thus, for ang given lift coefficient, the lift, and hence the deflection, was the same even though the Reynolds nuniber of the test was changed. This procedure entailed a snrsll change in Mach nudber, but probab

30、ly involved no appreciable compreesi- bility effects at the low Mach numbers involved. Accordingly, tests were made at a dynamic pressure of 50 pounds per square foot in which data were obtained at Reynolds nu whereas that of the smooth wing has a eue of 0.0049. Comparisons of the wing characterbtic

31、s with and ethout roughness are made in figure 23 for tests at Reynolds nunibem of 4.10 and 7.30 million xith corresponding Mach nunibem of 0.19 and 0.109. The changes in the aerodynamic characteristics cawed by the roughness strips were mre pronounced at the low Mach nunibers (fig. 23) than at the

32、higher Madh amber8 (fig. U); the roughness strips increased the drag coefficients somewhat but caused only small chesges of the lift and pitc-nent coefficients. The ineffectivemess of the roughness strips aay have been due to their be- improperly positioned on the wing surfaces, but timrs dfd not pe

33、rmit a thorough investigation to ascertafn if other chordwise locations of the roughness strips would be =re effective. Elevon deflected - 10 *.- For a Mach number of 0.900, in f fgure 24(a) the characteristics of the wing with roughness strips at 0.03 for Remolds nuribem of 2.30 and 3.60 miUion are

34、 compared to the chmc- teristics of the smooth wing at a Reynolds e effect of doubline; the dynamic pressure), wbile the Rem5Lds nuniber wa8 held Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 - NACA RM No. -20 I constant at 9.75 million and tbe

35、lhch mer allowed to increase from 0.080 to 0.160. The data show only small effects due to distortion. CONCLUSIOHS Anoinvest fgatfon was de of a semispan model of a wing swept back 63 and having an aspect ratio of 3.5. These tests were conducted to determine the separate effects of Mach and Reynolds

36、number on the aeroQmmic characteristic8 of the wing and on the effectiveness of an elevon. 1. The aerodynamic center of the wing shifted rearward near a lift coefficient of 0.2; whereas above a lift coefficient of approxi- mately 0.4 there webs an abrupt forward shift of the aerodynamic center 2. AB

37、 the Mch number was increased From 0.160 to 0.925 at Reynolds nunibera of 2.35 and 3.55.million in the low 1if“coefficien-b range (lift coefficient less than 0.2) the following effects of compresBlbility occurred: (a) The lift-curve slope at zero l-lft increased gradually from 0.043 to 0.048 per Seg

38、rea .(R = 3.55 million). (b) There was an increase in the static longitudinal stability, the aerodynamic center at zero lift moving aft from 42.4 to 44.6 percent of the meas aerodynamic chord. (c) The nraxfmum lift-t-ag ratio decreased from 18 to 14.7 (R = 3.55 mfllim). (dl At a Mach nuniber of 0.60

39、, the pitch effectiveness (rate of change of pitching-moment coefficient per degree of elevon deflection) hEad a value of -0.0053, and the lift effectiveness. (rate of change of lift coefficient per degree of elevon deflection) had a value of 0.0045. These values were not appreciably changed by vary

40、ing the &ch n&er from 0.60 to 0.90 (R = 2.26 million). 3. The effects of increaaing Regnolds number at an approxi- mately constant Mach nuniber and at a constant dynamic pressure of 50 pounds per square foot in the law lift-coefficient mag8 (lift coefficient less than 0.2) as determined from khese t

41、ests my be swmrparized as follows : Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NIICA RM No. Am20 - 13 (a) The lift-curve slope at zero lift decreased from 0.045 to 0 .Ob2 per degree. (b) There was little chaage in the static longitudfxal stabili

42、ty, the aerodynamic center remaining at approxfmately 42.4 percent of the man serodymnh chord. (c) The mxlnnun lift-to-drag mtio increased from 17.7 to 20.6. (a) The elevon pitch effectlvenese and lift effectiveness were little changed by the variation of Reynolds nmiber. Ames Aeronautical Iaborator

43、g National Advisory Codttee for Aeronautice, Moffett Field, Calif. 2. Loftin, Lsurence K., Jr. : . Theoretical and Experimental Ikta for a number of Ni4CA &-Series Afrfoil Sectims. NclCA m no. 138, 1947. 3. Sivells, James C. , and Deters, Gwen J.: Je-KBoundary ead T- Form Corrections for Partid4pan

44、Models with Reflection Plane, En3 Plate, or no End Plate, in a Closed Circular Winit punnel. mcA TN No. 1077, 1946. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 NACA RF4 No. A&D20 6. McCormack, Gerald M., and Wall-, Walter C. : Aerodynamic Stud

45、 of a Wing-Fueelage Canbination Bnploying a Wing Swept Back 63 g .- Invefltigation of a Large-Scale Model at Low Bpeed. mACA BM No. ASDo2, 1948. 7. Shortal, Joseph A.,. and Maggin, Bernard: Effect of Sweegback and Aspect Ratio on Ipngitudlnal Stability Characterietice of Wings st Low Speeds. . WA pm

46、 No. 1093, 1946- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM No. A8D20 - Condition Elevon deflected, smooth surface Do. Do. Do. Do. Elevon deflected, smooth surface Do. Do. Elevon mdeflected, roughness strips at 0.03 Do. Elevon deflected,

47、 roughness strfps at 0.03 Do. Mach rimer 0.180 to 0.925 0.160 to 0.925 0.180 0.182 to 0.080 0.080 to 0.160 0.- to 0.m 0.190 . 0.190 to 0.080 0.800 to 0.925 o.lgo and 0.109 0.m 0.lgO 2.35 3.55 2.35 to 10.30 4.11 to 9.85 9.75 2.26 4.00 4.20 to 9.80 4.10 and 7.30 2.30 and 3.60 4.00 Provided by IHSNot f

48、or ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. . Arm of the Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-. Provided by