NASA-TN-D-3375-1966 Aerodynamic data on a large semispan tilting wing with 0 5-diameter chord double-slotted flap and both left- and right-hand rotation of a single propeller《带有0 5.pdf

1、NASA TECHNICAL NOTE NASA TN D-3375 - c_- -.-I - -0 “ AERODYNAMIC DATA ON A LARGE SEMISPAN TILTING WING WITH 0.5-DIAMETER CHORD, DOUBLE-SLOTTED FLAP, AND BOTH LEFT- AND RIGHT-HAND ROTATION OF A SINGLE PROPELLER by Marvin P. Fink, Robert G. Mitchell, and Lucy C. Whidi. LangZey Research Center Langley

2、Station, Hampton, Vu. I JI -_ . NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. APRIL 1966 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I1lll11Hll1 Illn Ill11 I 1111 lllll11ll Ill1 0330245 AERODYNAMIC DATA ON

3、A LARGE SEMISPAN TILTING WING WITH 0.5-DIAMETER CHORD, DOUBLE-SLOTTED FLAP, AND BOTH LEFT- AND RIGHT-HAND ROTATION OF A SINGLE PROPELLER By Marvin P. Fink, Robert G. Mitchell, and Lucy C. White Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION For sa

4、le by the Clearinghouse for Federal Scientific and Technical Information Springfield, Virginia 22151 - Price $1.50 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-AERODYNAMIC DATA ON A LARGE SEMISPAN TILTING WING WITH 0.5-DIAMETER CHORD, DOUBLIGSLOTT

5、ED FLAP, AND BOJX LEFT- AND RIGHT-HAND ROTATION OF A SINGLE PR0P-R By Marvin P. Fink, Robert G. Mitchell, 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 sem

6、ispan V/STOL tilt-wing configuration having a single propeller which was tested for both right- and left-hand rotation. The wing had a chord-to-propeller-diameter ratio of 0.5, a double-slotted flap, an aspect ratio of 4.88 (2.44 for the semispan), a taper ratio of 1.0, and an NACA 4415 airfoil sect

7、ion. The data have not been analyzed in detail but have been examined to observe the predominant trends. It was found that the direction of propeller rotation had no significant effect on the lift or descent capability attainable, although different types of flow-control devices were required to ach

8、ieve the same results with different directions of rotation. The descent capability was determined from the values of attainable drag-to-lift ratios without stalling of any part of the wing within the propeller slipstream. The use of flaps was very effective in increasing the descent capability for

9、either mode of rota- tion. tested, virtually no descent capability prior to wing stalling was achieved with Oo flap deflection, whereas, with 40, 60, or TO0 flap deflection, a descent capability of about 20 was achieved. For example, with the most favorable combination of flow-control devices INTROD

10、UCTION Most of the aerodynamic research done on the tilt-wing propeller-driven V/STOL configuration has been of an exploratory character and has been obtained with small-scale models. The interest in this type of airplane has become so substantial that there is a need for large-scale systematic aero

11、dynamic design data for this type of airplane. A program has therefore been inaugurated at the Langley Research Center to provide such information by means of a large-scale semispan tilt-wing-and-propeller model in the Langley full-scale tunnel. Ref- erences 1, 2, and 3 are concerned with this inves

12、tigation, and the results of the fourth part of the investigation are reported herein. The present series of tests were made on a model having a single propeller on the semispan wing, a Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I ,111 ,111 I -.

13、-I,.-. ._.-_- II I. 111.11 1111111 I1 1111 chord-to-propeller-diameter ratio of 0.50 (compared with a ratio of 0.60 for the three previous investigations), a 35-percent-chord double-slotted flap, and a leading-edge slat which could be located in either of two positions. The investigation covered a r

14、ange of angles of attack from -20 to goo and a range of power conditions from zero thrust to that required for hovering. Both modes of propeller rotation were tested in the present investigation. The results of previous investigations (refs. 2 and 3) show that the direction of propeller rotation has

15、 no appreciable effect, but it was believed that there might be some significant effect in the present investigation because of the shorter wing chord and increased loading due to the double-slotted flap. The lift, drag, and pitching moments of the model were measured over the range of test conditio

16、ns and the flow was observed by means of tufts on the upper surface of the wing. The results of this investigation are presented herein without detailed analysis to expedite their dissemination. SYMBOLS The positive sense of forces, moments, and angles is shown in,figure 1. The pitching-moment coeff

17、icients are referred to the wing quarter-chord line. The coefficients are based on the dynamic pressure in the propeller slipstream. Conventional lift, drag, and pitching-moment coefficients based on the free- stream dynamic pressure can be obtained by dividing the slipstream coefficients by (1 - CT

18、,s); for example, CL = CL,s/(l - CT,). The thrust coefficient CG may be obtained from the equation cG = cT,($)l/( - T,S)* Measurements for this investigation were made in the U.S. Customary System of Units. Equivalent values are indicated herein in the International System (SI) in the interest of pr

19、omoting the use of this system in future NASA reports. Factors relating the two systems for units used in this paper may be found in the appendix. L lift coefficient based on free airstream, - qs CL L lift coefficient based on slipstream, - qSS CL, s D drag coefficient based on slipstream, - qSs cD,

20、 s MY pitching-moment coefficient based on slipstream, - % s 9, sc 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-m cT, s A b C Cf cv D h L MY I thrust coefficient based on slipstream, TcD;! qs T thrust coefficient based on free airstream, - qs to

21、tal area of propeller disk, ft2 (meters2) propeller-blade chord, in. (meters); or wing span, ft (meters) wing chord, ft (meters) flap chord, 11.90 in. (14.68 cm) vane chord, 5.78 in. (14.68 em) propeller diameter, ft (meters) also, total model drag, lbf (newtons) width of slat or of flap-slot gap or

22、 thickness of propeller blade, ft (meters ) total lift of model, lbf (newtons) pitching moment, lbf -ft (newton-meters ) free-stream dynamic pressure, - 2 T 7t-D slipstream dynamic pressure, q + T, lbf/sq ft (newtons/meter2) r R S T X Y a 6f 4 radius to element on propeller blade, ft (meters) radius

23、 of propeller blade, 2.83 ft (0.86 meter) area of semispan wing, 19.60 ft2 (1.82 meted) propeller thrust, lb (newtons) longitudinal distance along chord, ft (meters ) vertical height above or below chord line, ft (meters) angle of attack, deg flap deflection, deg 3 Provided by IHSNot for ResaleNo re

24、production or networking permitted without license from IHS-,-,-6, leading-edge-slat deflection, deg P v Subscript: max maxi“ mass density of air, slugs/ft3 (kilograms/meterg) free-stream velocity, ft/sec (meters/sec ) MODEL The model used in this investigation was a semispan model which would rep-

25、resent the left panel of a full-span wing. The principal dimensions of the model are given in figure 2. The wing was mounted on the scale balance system in the tunnel so that the lift and drag measurements were read directly about the wind axis. Where the wing extended through the reflection plane,

26、a circular end plate (with a diameter equal to twice the wing chord) was fitted around and attached to the wing to prevent air from leaking through the reflection plane at the wing root. The model was constructed to allow nunerous changes to be made in the test configuration, such as: leading-edge m

27、odification, and changes of airfoil, trailing-edge flap, direction of rotation of the propeller, and wing planform. The basic structure of the wing consists of a heavy steel box-beam spar to which a power train to drive the propellers through spanwise shafting is attached and around which various ai

28、rfoil contours can be fitted. The model configuration for the present tests had a 68-inch-diameter (1.73-meters) propeller having the characteristics shown in figure 3. peller location was such that the propeller tip extended to the wing tip. In the present investigation both directions of propeller

29、 rotation were tested. The propeller 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 CT,s were set by the operator by changing the speed of the drive motor.

30、The blade angle at the O.73R station of the propeller was held constant at 17 throughout the inves- tigation. wing to correspond approximately to the zero-lift line of the airfoil section. The pro- The thrust axis was inclined upward 4 from the chord line on the The airfoil used was the NACA 4415 se

31、ction with a 34-inch (0.864-meter) chord. 0.50. (1.73 meters) was 19.60 square feet (1.82 meters2) and did not include the area of the tip fairing. This chord length gave a ratio of wing chord to propeller diameter of The reference area of the wing based on a semispan of 83 inches The model had a 33

32、-percent-chord double-slotted flap which was set at Oo, The flap was deflected to TO0 for 40, and 60 for most of the present tests. one set of tests. Figure 2(b) shows the flap deflected 60 relative to the 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IH

33、S-,-,-vane. The flap bracket was constructed so that the nose of the flap positioned for each flap setting as shown in the detail in figure 2(b). This relationship was also true for the nose of the vane for flap deflections of 60 and greater. As the flap angle was decreased the nose of the vane move

34、d forward under the skirt. Because of bracket limitations, the deflection angle of zero was not obtainable so the entire flap system was replaced with a solid trailing edge for this case. The ordinates for the vane and flap are given in table 1. 2 3 4 Two positions of a leading-edge slat were invest

35、igated in combination with the flap on this model. higher than normal slat position gave better results for some test conditions. For this reason, some of the tests in the present investigation were conducted with the slat in a “high“ position. The high and low positions of the slat with the angles

36、and slot gaps used are shown in figure 2(b). Some previous unpublished data have indicated that a 4 5 6 Fences having a height of 0.20 and extending from 0.13 on the lower sur- face of the wing around the leading edge to about 0.75 on the upper surface were installed at two spanwise locations on the

37、 wing in an attempt to confine the stall inboard of the propeller slipstream. The inboard fence was placed about where the side of a fuselage might be (20% of the semispan) and at O.?r/R of the inboard propeller blade as indicated in figure 2(c). were made with fences on, both fences were installed.

38、 When tests 5 6 7 TESTS 7 8 9 The tests were made for various deflections of the double-slotted flap and for two different positions of a leading-edge slat. The specific configurations tested, together with a list of tables and figures in which the data for each may be found, are given in the follow

39、ing table: Direction of rotation Configuration Down at tip Basic leading Basic leading Basic leading Basic leading with fences Basic leading with fences Basic leading with fences edge on cage on edge on Inboard slat, Inboard slat, Inboard slat, Flap deflection, 6f, deg 0 40 60 0 40 60 0 40 60 Table

40、Figure I T-p- 10 12 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Direct ion of rot at i on Down at tip Up at tip Configuration Inboard slat, 6, = 30 Inboard slat, 6, = 30 Inboard slat, 6, = 30 with fences on with fences on with fences on Inboard

41、 slat, 6, = 30 Inboard slat, 6, = 30 Inboard slat, 6, = 30 Inboard slat, 6, = 30 Inboard slat, 6, = 30 Inboard slat, 6, = 30 with fences on with fences on with fences on Inboard slat, 6, = loo, Inboard slat, 6, = loo, Inboard slat, 6, = loo, high position high position high position Flap deflection,

42、 6f, deg 0 40 60 0 40 60 0 40 60 .40 60 70 Table 11 12 13 14 15 16 17 18 19 20 21 22 The tests were made over a range of thrust coefficients from 0 Figure 13 14 15 16 17 18 19 20 21 . 22 23 24 ;o 1.0 for the basic wing and for CT,s = 0.90, 0.80, and 0.60 tions. For any given test the thrust coeffici

43、ent was held constant over the angle-of-attack range by adjusting the propeller speed to give the required thrust at each angle of attack. The angle-of-attack range for the tests was approximately from the angle required for zero lift to that required to stall the wing or to develop a drag-lift rati

44、o of about 0.30, whichever was lower, except for CT, = 1.0 range was from 0 to go0. The test Reynolds number, based on the wing chord length and the velocity of the propeller slipstream, was about 2.32 x 106 for thrust coefficients from 1.00 to 0.30. For the CT, = 0 condition, where the thrust was h

45、eld at zero, the Reynolds number was about 1.91 x 106. for the other configura- (the static thrust case) where the angle-of-attack No tunnel-wall corrections have been applied to the data since surveys and analysis indicate that there would be no significant correction, as explained in reference 1.

46、6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-DISCUSSION The data presented have not been analyzed in detail but have been examined to observe general trends. A few such trends predominate. One very general observation was that the force-test dat

47、a could not be used as an indication of the occurrence or extent of wing stalling. The results of the tuft tests show that the onset of stalling over significant areas of the part of the wing within the propeller slipstream frequently occurs at 20 to 30 angle of attack below or above the angle of at

48、tack for maximum lift coefficient. Effect of Variables Effect of direction of propeller rotation.- The results of the force tests show no consistent or very significant effects of the direction of propeller rotation on lift or drag. The tuft tests, however, show major effects of the direction of pro

49、peller rotation. Rotation of the propellers in the down-at-the- tip direction consistently causes stalling (of the part of the wing in the slip- stream) to start inboard of the nacelle, that is, behind the upward-going blades. The up-at-the-tip rotation, on the other hand, may result in the onset of stalling occurri