1、CC3NFf:D.ENPtAL RESEARCH MEMORANDUti WIND-TUNNEL INVESTIGATION AT LOW SPEED OF A WIIXG SWEPT RACK 63 AND TWISTED AND CAMBERED FOR UNIFORM LOAD AT A LIFT COEFFICIENT OF 0.5 AND WITH A THICKENED TIP SECTION- By James A. Weiber g and Hubert C . Car el Ames Aeronautical Laboratory Moffett Field, Calif.
2、ClASSIF#CATION CANCELLE A,i!y-,_,-Date*S*J dfw-rLc-;r- - - - l Byy_-n to the tip of the ting resulted in reduced tip-leaUng4ge pressure peaks with no improvement of tip lift characteristics. Thus, the early loss of lift of wing sections near the tip which resulted in the Urge variations in longitudi
3、nal stabil- ity was attributable igh angles of attack. Subsequent to the tests of reference 7 it was reasoned that the lift range for satisfactory stability characteristics might be extended to higher lift coefficients by increasing the range of usable lift of the sectians near the tip through an in
4、crease of the thickness of these seo- tions . Computations showed that the increase of drag at supersdc speeds due to the increased thickness of sections near the tip would be relatively Small. Canssquently, the mcdel used for the research reprted in reference 7 was alteredtoincmczatethicker section
5、s fromthe midsemispa;nto the tip. For expediency in model constructian, the twist of the revised por- tion of the wing was also modified from that :of the origins1 wing. Tests of.the wing were made in one of the Ames 7-by lce fences, of spoilers, of elevens, and of a leading-edge flap nn the low-spe
6、ed characteristics of the Wing. NOTATION All data are -presented as B f flap No corrections were applied to the data for the effects of model distor- tion or for possible effects of interference between the model and the tunnel floor or of-leakage through the gap between the tunnel floor and the ext
7、ension of the base of the model where it passed through the floor. These effects were discussed in reference 7 and were believed to have been small. An investigation was made to determine the effect on the wing pressure distribution of the leakage through the tunnel-floor gap. The results are discus
8、sed tithe section entitled “Pressure-Distribution Measuremnts.! MODELDESCRIETION The model used in these tests, hereinafter referred to as the revised wing (figs. 1 and 3), was the model descrtbed in reference 7, hereinafter referred to as the original wing, with the twist and thick- r tested was a
9、semispan model with 630 sweepback of the leading edge, an aspect ratio of 3.5, and a taper ratio of 0.25 (ratio of tfp chord to root chord). The thickness distributfon of the tip section of the revised wing par- allel to the plane of symmetry was that of the NACA 0012 section. !Rm camber line of the
10、 tip section a3 the ci-L passed through the floor to support the model. The wing was equipped;with pressure orifices on , sections parallel to the plane of symmetry a$ 0.200, 0.383, 0.707, sad 0.924 semispan. shown in table I. The chordwise locations of the preaw orifices am. The fuselage described
11、in reference 7 w4 for producing pitching moments and. for reduc$ng the variation of aerody- namic cent8rwithliftwas greaterwiththe eleven hinge line cathewing trailing edge thas on the ?O+ercenomputed by the methods of Weissimger as outlined in references 13 Reasonably good agreement was obta+ned be
12、tween the computed and masurid span load distributions. The results of tests of a semispan model of &twisted and cambered wing with the leading edge swept back 630 -showed that increasing the thiclmess with a small modification to the twist from midsemispan to the tip resulted in no improvement of t
13、he longitudinal characteristics of the wing at low speeds A reduction in the tip-leadiwdge pressure peaks was obtained with no improvement of tip lift chsxacteristics indi- cating that the ewly loss of lift of the tip, which rCesulted in the large variations in longitudinal stabllity,$as due more to
14、 spanwise - flow of the boundary layer than to local stall of the tip sections. !J%e change of thickness and twist had a negligible effect on the lcm+speeddrag ofthewing. The addition of stall and bouudary-lay&+ontrol devices had a COP siderable effect on the stability of the wipg. Upper-surface fen
15、ces on the inner portion ofthe yiq were nearly a$ effective as those near the , tip for controlling spamise boundary-lapr flow. Fences extending over the after 50 percent of the chord of the w$gprovided about the same improvement of wing stability as full-chord: fences. Addition of a . lead-age flap
16、 over the outer 0.22 se&span of the w+g with fences at 0.6 and 0.8 semispan resulted in a nearly limar variation of wing pitcwment coefficient with Uft coefficientup to a lift coeffi- cient of 0.7. Uppex-surface split flaps on the outer 0.37 semispan were ineffec- tive for longitudinal control at hi
17、gh lift coefficients but resulted in anapproximatelylinear pitchqbentcuqe for the wing as a result of the large variation of effectiveness of the spfit fla$ with lift _ coefficient. *The basic plus additional load is presented for a lift coefficient (a = 0.4) at which the local lift and the span 1oa
18、dFng are not aplsye- ciably affected by separation. This lift coefficient corresponds apoximate3.y to the low-speed design lift, coefficient (CL = 0.38). . . - Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA RM A50114 .- 11 The modification of t
19、he outer half of the semispan wing resulted in only small changes of the chordwise pressure distributions and lift of the wing sections. Ames Aeronautical Laboratory, National Advisory Committee for Aeronautics, Moffett Field, Calif. REFEEENCES 1. Madden, Robert T.: Aerodynamic Study of a Winuselage
20、 Conibination Employing a Wing Swept Back 63o.- Characteristics at a Mach N&er of 1.53 Ihcluding Effect of Small Variations of Sweep. NACA RM A8JO4, 1949. 2. Mas, Newton A.: Aerodynamic Study of a Wing-Fuselage Combination Employing a Wing Swept Back 63o.- Characteristics for Sgmmstrical Wing Sectio
21、ns at High Subsonic and Moderate Supersonic Mach Nunibers. NACA RMAW, 1949. 3. McCormack, Gerald M., and Walling, Walter C.: Aerodynamic Study of aWing+uselage Co&lnationEmploying awing Swept Back 63o.- Investigation of a Large-Scale Model at Low Speed. NACA RM A8BO2, 1949. 4. Hopkins, Edward J.: Ae
22、rodynamics Study of a Wing+uselage Combina- tion Fimploying a Wing Swept Back 63o.- Effects of Split Flaps, Elevens, and IeadinU.ge Devices at Low Speed. NACA RM A9C21, 1949 l 5. Madden, Robert !I?.: Aerodyna&c Study of aWing+uselage CombIna- tion Employing a Wing Swept Back 63o.- Investigation at a
23、 Mach Nmiber of 1.53 to Determine the Effects of C&ring and Twisting the Wing for UnIformLoad at a Lift Coefficient of 0.25. NACA =fAgCO7, 1949. 6. Jones, J. Lloyd, and Bemele, Fred A.: Aerodynamic Study of a Wing- Fuselage Cox6UnationEarploying a Wing Swept Back 63o.- Character- istics Throughout t
24、he Subsonic Spead Range with the Wing Cambered and Twisted for a Uniform Load at a Lift Coefficient of 0.25. NACA RM AgD25, 1949. 7. Weiberg, James A., and Carel, Hubert C.: Wind4unn81 lbvestigation at Low Speed of a Wing Swept Back 630 and Twisted and Ca&ered for a Uniform Load at a Lift Coefficien
25、t of 0.5. NACA RM A5OA23, 1950. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 -: IWARMA5OIl4 8. Swanson, Robert S., and Toll, Thomas A.: Jet+BozIndary Corrections for ReflectiopPlane Models in Rectangular Wind Tunnels. NACA Rep. 770, 1943. 9. Po
26、Umms, Edward C.: Jet43oIndced-Qwash Velocities for Swept Reflectian-Plane Models Mounted Vertically in 7- by lO+!?oot Closed, Rectangular Wind Tumels. Z&CA !ZH 1752, 1948. 10. Graham, Robert R., and Conner, D. William: Investigation of High- Lift and Stall-Control Devices on an *CA 6&Series 420 Swep
27、t- back Wing With and Without Fuselage. IiACA RM L7GO9, 1947. 11. Conner, D. Willlam, and l&8-, Robert If.: Effects of a Fuselage and Various High-Lift and Stall-Con-t&l Flaps on Aerodyne&c Characteristics in Pitch of an NACA -eriea 400 Swept+Back Wing. NACA RM L&27, 1947. 12. Jones, Robert T.: Wing
28、 PlanForms for &h-Speed Flig.&. NACA TN 1033, 1946. 13. DeYomg, John: Theoretical Additional span Loading Chszacteris- tics of Wings with Arbitrszy Sweep, Aspect Ratio, and Taper Ratio. NACA Rep. 92l, 1948. 14. Stevens, victor I.: Theoretical Basic Span Loading Chsracteris- tics of Wings with Arbitr
29、ary Sweep, Aspect Ratio, and Taper Ratio. XACA TN 1772, 1948. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NACA myx14 A . TABLE I.- CECEUWISELOCATIOmS OFTHE: EESSURE ORIFICES Orifices located m both upper a&. lower wing surfaces 1 Percent wing aho
30、rd 0 30.00 1.25 40.00 2.50 50.00 5.00 60.00 7.50 70.00 10.00 80.00 15.00 go- 20.00 95-w 25.00 13 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-14 NACA RM A50114 TABIS II.- DMNSIONSOFTHE&SPANMODEZ rea of semispm modei, F, square feet . . . . . . . .
31、 . 14.262& Semispan, feet ., . . . . . . . . .I. . . . . . . . . 5.0 Mean aerodynamic chord, feet . . . , .:. . . . . . . . . 3.20 . Aspectratio. . . . . . . . . . . . .:. . . . . . . . . 3.5b Taper ratio (ratio of tip chard to root chord) . . . . . 0.25 Sweepback of leading edge, degrees . .I. . ,
32、, . . . . . 63 SW8epbaCkof quarter-chord line,degr&s . . . . . . . . 60.8 Geometric twist, degrees . , . . . . .I. , . . . . . . . 20.5 Dihedral,degrees . . . . . . . . . . . . . . . . . . . . . 0 Fuselage Length, feet , . . . . . . . . . .I . . . . . . . . l-4.2 l+faximumdie,meter, feet. . . . . .
33、. . . . . . . . . . . . 1.36 Fineness ratio (ratio of length to maximum diameter) . . . 10.4 tiea to projected tip was 14.286 square feet. %a sed on span of 10 feet and area (to projected tip) of 28.5 sqwe feet. Provided by IHSNot for ResaleNo reproduction or networking permitted without license fro
34、m IHS-,-,-NACA RMA5OI14 15 !llABm III.- cooRDINAms OF TBE Fo-sEIAm All dimensirms in inches . station Diameter 0 0 4 2.84 8 5.34 Il.2 7.50 16 9.30 20 1080 24 u.98 28 12.88 30.6 13.26 40.8 14.28 51.0 15.20 61.2 15.82 71.4 16.20 station Mameter 81.6 16.32 91.8 16.20 102.0 15.82 112.2 15.20 x&.4 14.28
35、132.6 13.26 142.8 11.68 153.0 9.86 163.2 7.58 164.4 7.16 166.4 5.82 168.4 3.58 170.4 0 length I Fineness ratio, = 10.4 n-aximumdiameter Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Provided by IHSNot for ResaleNo reproduction or networking permitt
36、ed without license from IHS-,-,-Root chord, , Ttvihg edge r . I .- ei, I -_ boment center s Wing reference plane Dimensions in inches Figure I .- Diagram of the model. 70 1 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-f I Y Mm7n line I / v Trailin
37、g edge ROOt SSCiiOn Figure 2.- Relationship befween fhe mhg geomefry and fhe coordlnofe axes. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,- - .-,_. .w- Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,
38、-,-Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, f2 I Maximum fuselage radius I i 5 P f! $4 . t: - ckiginar $ - Revised P $2 0 Ia) Twisi. (iY Maximum camber and maxlmum thickness. Figure 4.- Comparison of original and revised twist, maximum cambe
39、r, and maximum ihickness. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Fence Fw Spoiler (a) Position. Figures Geometry of the stall-control devices. 1 , Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-
40、,-,-Tjkal section A -A , fences A, 6, D,and F I . I Typlcat section C - C. etevon 2 G i3 P c Jjykal sect/on A - A , fences C and E Typical section B- B , e&on / Typlcat section D - D , leading - edge flap hadghl of spoiler Section E - E, smiler figure 5. - Concluded. (bJ Dimensions . Provided by IHS
41、Not for ResaleNo reproduction or networking permitted without license from IHS-,-,-N .F -I6 -8 0 8 16 24 .I6 ,.08 0 48 716 724 Angle of aftack, a, deg fifcbing-moment coefficienf , Cm figure h-Effect of wing revision on the lift and pitching-moment characteristics. . . Provided by IHSNot for ResaleN
42、o reproduction or networking permitted without license from IHS-,-,-I . I !A 62 o .8 78 I I I I I I I I I I I I I I I I I I -16 -8 0 8 16 24 .I6 .08 0 378 :I6 724 Angle of attack, UY , deg Pitching-moment. ceefficient , Cm W+W with fuselage. r 8 Figure 64oncluded. Provided by IHSNot for ResaleNo rep
43、roduction or networking permitted without license from IHS-,-,-f2 -.8 .“o- ._ 4 8 I2 Lift-drag r&o, L& 16 n, figure &Effect of wing revishm on the drag and Off-drag ratio. . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I6 -8 0 8 I6 Angle of attack
44、, a: , deg 24 .I6 -08 0 -.08 ,I6 24 Pitching-moment coefficient, Cm (b)Wlirg with fuselage. Figure 6.-Concluded. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-a, 12 A Original Q Revised I I .8 LY I I I I I I I I I 8 0 D8 .I6 24 .32 .& _-_ -. -&ag.#
45、&nt r-c-, - krlwi* figure &Effect of .wing revision on the drug and fift-drag ratio. . 11 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-12 A Origihaf .V 0 .08 I6 .24 32 40 Drag coefficient, CD . I - - Origina / - Rnvicaf (b) Wing with fuselaga figure Z-Concluded -0 4 8 12 I6 20 Lift-drug ratio, Q0 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-
