1、NASA TECHNICAL NOTE NASA TN D-1 586 AERODYNAMIC DATA ON LARGE SEMISPAN TILTING WING WITH 0.6-DIAMETER CHORD, SINGLE SLOTTED FLAP, AND SINGLE PROPELLER ROTATING UP AT TIP by Maruin P. Fink, Robert G. Mitchell, and Lwy C. White Langley Research Center Lungley Station, Humpton, Va. , NATIONAL AERONAUTI
2、CS AND SPACE ADMINISTRATION WASHINGTON, D. C. OCTOBER 1964 j4 / Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I Illlll 11111 lllll IIIH lllll INII lllll Ill1 llll 0353773 AERODYNAMIC DATA ON LARGE SEMISPAN TILTING WING WITH 0.
3、6-DIAMETER CHORD, SINGLE SLOTTED FLAP, AND SINGLE PROPELLER ROTATING UP AT TIP By Marvin P. Fink, Robert G. Mitchell, and Lucy C. White Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sale by the Office of Technical Services, Department of Comm
4、erce, Washington, D.C. 20230 - Price $2.75 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AERODYNAMIC DATA ON LARGE SEMISPAN TILTING WING WITH O.-DTER CHORD, SINGIX SLOTmD FLAP, AND SINGIZ PROPELIXR ROTATING UP AT TIP By Marvin P. Fink, Robert G. Mi
5、tchell, and Lucy C. White Langley Research Center SUMMARY An investigation has been made in the Langley full-scale tunnel to deter- mine the longitudinal aerodynamic characteristics of a large-scale semispan V/STOL tilt-wing configuration having a single propeller with propeller rotation such that t
6、he blades rotated upward at the wing tip and downward near the root. The wing had a ratio of chord to propeller diameter of 0.6, a single slotted flap, an aspect ratio of 4.05 (2.025 for the semispan), a taper ratio of 1.0, and an NACA 4415 airfoil section. The data have not been analyzed in detail
7、but have been examined to observe general trends. A few such trends predominate. The basic leading-edge configu- ration had practically no stall on that portion of the wing immersed in the propeller slipstream at angles well above those corresponding to the peak of the lift curve for the high thrust
8、 conditions corresponding to operation in the STOL range of flight; and, in general, the stall on the wing center section coincides with the angle of attack for maximum lift for the low thrust coefficients. The use of a leading-edge slat on the outboard wing section had virtually no effect on the ae
9、rodynamic characteristics of the wing since there was no stalling on the outboard section of the wing without the slat. The use of an inboard slat had no effect on the tip section. Nl-span slat reduced stall on the inboard section of the wing and increased both the angle of attack and drag at maximu
10、m lift, but did not increase the value of maximum lift. Neither the flow in the slipstream nor the force data was improved by the Krueger flap, but the Krueger flap did improve the flow on the part of the wing center section inboard of the propeller slipstream for the higher thrust coefficients. INT
11、RODUCTION Most of the aerodynamic research that has been done on the tilt-wing propeller-driven V/STOL configuration in the past has been of an exploratory character and has been obtained with small-scale models. The interest in this type of airplane has now become so substantial, however, that ther
12、e is a need for large-scale systematic aerodynamic design data for this type of airplane. A program has therefore been inaugurated at the Langley Research Center to Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-The model configuration for the prese
13、nt tests had a 68-inch-diameter pro- peller having the characteristics shown in figure 2(b). The propeller location was such that the propeller tip extended out to the wing tip. The direction of propeller rotation was up at the wing tip and down at the root. rotation is sometimes referred to as “wit
14、h the tip vortex.“ thrust was measured by a strain-gage balance which was a part of the propeller shaft. The output was fed through sliprings to an indicating instrument. The required values of thrust for each were set by the operator by changing the speed of the drive motor. peller was held constan
15、t at 170 throughout the investigation. The thrust axis was inclined upward bo from the chord line of the wing to correspond approxi- mately to the zero-lift line of the airfoil section. This mode of The propeller CT, The blade angle at the 0.75R station of the pro- The airfoil used was the NACA 4415
16、 section with a 41-inch chord. This chord length gave a ratio of wing chord to propeller diameter of 0.6. The ref- erence area of the wing based on a semispan of 83 inches was 23.62 square feet, and did not include the area of the tip fairing. The model had a 40-percent-chord single slotted flap whi
17、ch had a deflection Figure 3 shows the flap in the 50 deflected position and range from 0 to 50. also shows the slot geometry. The two leading-edge flow-control devices shown in figure 3 were investi- gated in combination with the flap on this model. flap and a leading-edge slat. The Krueger flap, w
18、hich in the retracted position in actual use would form the bottom contour of the nose section, was constructed of sheetmetal and was hinged at the 0.017 station. Its deflection could be varied from 30 to 90 in increments of loo. However, previous investigations covering a large range of deflections
19、 showed that a 50 deflection proved near optimum for this wing; therefore, for these tests, only the 50 deflection was used. In one test the Krueger flap was faired straight from the end of the flap to the leading edge of the basic airfoil nose as indicated in figure 3. For the leading-edge slat, tw
20、o deflection angles (20 to 30“) and two slot gaps (0.0244 and 0.0122) were originally provided. Test data presented in refer- ence 1 showed little change in the results with variation of slat angle and gap; consequently, the present tests were made only with a 20 deflection and an 0.244 gap. The sec
21、tion designated as the inboard section extended from the wing root to the nacelle and that section designated as the outboard section extended from the nacelle to the wing-tip fairing. These devices were a Krueger TESTS, RESULTS, AND DISCUSSION The tests were made for a range of single slotted flap
22、deflections and a combination of leading-edge flow-control devices. The specific configuration tested, together with a list of tables and figures in which data for each may be found, are given in the following table: 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without lice
23、nse from IHS-,-,-Leading-edge configuration Basic leading edge Leading-edge slat: Outboard section; 6, = 20 Outboard section; 6, = 20 Outboard section; 6, = 20 Outboard section; 6, = 20 Inboard section; 6, = 20 Inboard section; 6, = 20 ull span; 6s = 20 Full span; 6, = 20 Krueger flap: Outboard sect
24、ion; % = 50 Full span; 6 = 500 Inboard section (faired to leading edge); S, = 50 Flap deflection, deg 6f = 0 6f = 20 Sf = 40 6f = 50 6f = 0 6f = 20 6f = 50 6f = 50 6f = 50 6f = 50 6f = 50 6f = 50 6f = 40 6f = 40 6f = 40 Pable 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 pigure 8 9 10 11 12 13 14 15 16 17 18
25、The tests were made over a range of thrust coefficient from 0 to 1.0, and for any given test the thrust coefficient was held constant over the angle-of- attack range by adjusting the propeller speed to give the required thrust at each angle of attack. from the angle required for zero lift to that re
26、quired to stall the wing or develop a drag-lift ratio of about 0.3, whichever was lower. except for CT,s = 1.0 900. The test Reynolds number, based on the wing chord length and the velocity of the propeller slipstream, was about 2.8 X 106 for thrust coefficients from 1.00 to 0.30. For the condition
27、where the thrust was held at zero, the Reynolds number was about 2.3 x 10 6 . The angle-of-attack range for the tests was approximately (the static thrust case) where the angle-of-attack range was Oo to CT, = 0 No tunnel-wall corrections have been applied to the data since surveys and analysis had i
28、ndicated that there would be no significant correction as explained in reference 1. The data presented have not been analyzed in detail, but have been examined to observe general trends. A few such trends predominate. For all the various leading-edge configurations, the trailing-edge flap was For an
29、gles of attack stalled over most of its area for deflections of 40 and 50 (but not 200) at angles up to approximately that required for maximum lift. above that for maximum lift, however, the stalling on the flap disappeared. Flap deflections of 400 and 50 were found to give almost exactly the same
30、lift 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-and drag characteristics, and they both gave higher lift and drag than that for the 20 deflection. For the basic leading edge, the wing-flow photographs show that there is practically no stall on
31、 that portion of the wing immersed in the propeller slip- stream at angles of attack well above that corresponding to the ped of the lift curve for the high thrust conditions corresponding to operation in the STOL range of flight (CT, = 0.060 to 1.00) and that, in general, the stall on the wing cent
32、er section coincides with the angle of attack for maximum lift for the low thrust coefficients (CT, = 0 and 0.30). Similar stall characteristics were also noted in the results of reference 1 in which the same model was tested with an extended Fowler type of trailing-edge flap with the same mode of p
33、ropeller rotation. The result of tests reported in reference 2 where only the propeller rotation was different (down at the tip) from the present tests the tufts showed stall starting at the wing root and progressing smoothly outboard onto the pos- tion of the wing in the propeller slipstream inboar
34、d of the nacelle. Evidently, the direction of propeller rotation which has the effect of increasing the angle of attack of the portion of the wing behind “upgoing“ blade causes this change in stall characteristics. That portion of the center section which is not in the slipstream does not appear to
35、be affected by the direction of propeller rotation. The use of the leading-edge slat on the portion of the wing outboard of the nacelle had virtually no effect on either the lift and drag or wing-flow charac- teristics - evidently because there was no significant wing stalling on the out- board sect
36、ion of the wing which might be affected by the slat. The inboard slat, however, delayed the stall of the center section (inboard of the slipstream) for all conditions tested; and it caused significant increases in the angle of attack for maximum lift and the drag at maximum lift at low thrust coeffi
37、cients (CT, = 0 to 0.60) where the center-section lift was an appreciable part of the total lift. The inboard slat did not, however, significantly increase the value of the maximum lift coefficient. The sharp break in the lift curve at stall for the basic leading edge at low thrust coefficients was
38、reduced to a gradual decline by the inboard slat. The full-span slat gave almost exactly %he same results as the inboard slat alone, as would be expected, since the outboard slat was not effective. None of the Krueger flap configurations gave any appreciable improvement in the flow on that portion o
39、f the wing in the propeller slipstream or in the force data over that of the basic leading-edge configuration. The inboard portion of the Krueger flap did, however, improve the flow on the portion of the wing inboard of the slipstream for thrust coefficients from 0.95 to 0.60, but for the low thrust
40、 conditions from 0.30 to 0, the Krueger flap did not have this effect since no stall occurred in this area for angles of attack up to maximum lift. CONCLUSIONS An investigation to obtain large-scale aerodynamic data and flow studies on a semispan wing for an angle-of-attack range from -20 to 900 for
41、 thrust coeffi- cients from 0 to 1.0 indicates the following conclusions: 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-1. The basic leading-edge configuration had practically no stall on that portion of the wing immersed in the propeller slipstr
42、eam at angles well above those corresponding to the peak of the lift curve for the high thrust conditions corresponding to operation in the STOL flight range; and, in general, the stall on the wing center section coincides with the angle of attack for maximum lift for the low thrust coefficients. 2.
43、 A leading-edge slat on the outboard wing section had virtually no effect on the aerodynamic characteristics. The slat on the inboard wing section had almost the same characteristics as the full-span slat. A full-span slat reduced stall on the inboard section of the wing and increased both the angle
44、 of attack and drag at maximum lift but did not increase the value at maximum lift. 3. Neither the flow in the slipstream nor the force data was improved by the Krueger flap, but the Krueger flap did improve the flow on the part of the wing center section inboard of the propeller slipstream for, the
45、 higher thrust coefficients. Langley Research Center, National Aeronautics and Space Administration, Langley Station, Hampton, Va., July 1, 1964. REFERENCES 1. Fink, Marvin P., Mitchell, Robert G., and White, Lucy C. : Aerodynamic Data on a Large Semispan Tilting Wing With 0.6-Diameter Chord, Fowler
46、 Flap, and Single Propeller Rotating Up at Tip. NASA TN D-2180, 1964. 2. Fink, Marvin P., Mitchell, Robert G., and White, Lucy C. : Aerodynamic Data on a Large Semispan Tilting Wing With 0.6-Diameter Chord, Single-Slotted Flap, and Single Propeller Rotating Down at Tip. NASA TN D-2412, 1964. 7 Provi
47、ded by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I IIIIIIIIIIII I1 I1 I TAELE I.- TAUIAl!ED AER0IMUMJ.C MTA FOR sf = Oo 0.1 CmJ 6 -0.083 -e073 -.061 -.044 - .017 -.m2 .030 .0b1 -047 .044 .035 .033 .028 .025 .024 .022 .025 .029 -0.724 - .7a2 -.as -.a= - .82
48、1 - .793 -.748 - .692 - .613 - * 517 - .394 - * 279 -0173 - .014 . U8 .227 .333 .403 -0.482 - .353 - .190 -.Ox, .I62 * 329 .w .673 .goo 1.005 1 * 097 1.172 1.224 1.252 1.259 1.209 .a35 1.247 -20 -15 -10 -5 0 5 10 15 20 25 30 35 40 45 50 55 60 65 70 75 80 90 -20 -15 -10 -5 0 5 10 15 20 25 30 35 40 45
49、 50 55 60 65 - _ -0.583 -.512 - .269 - .061 .176 .400 .654 ,850 1.074 1.171 1.234 1.276 1.297 1.286 1.202 1 * 033 -0.456 -.550 -.585 - .595 -.97 - .561 -.p8 -.443 -0350 -.248 -.8 .OW .182 .327 .440 * 523 -0.109 -.=7 -. - 50 Trailing-edge flap Figure 3.- Sectional views of various leading-edge devices and trailing-edge flap. Provided by IHSNot fo
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