1、c “ N o* *o N PC: U I e ASA CONTRA REPORT CTO REFLECTION-PLANE TESTS OF SPOILERS ON AN ADVANCED TECHNOLOGY WING WITH A LARGE FOWLER FLAP I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM 1. Repat No. 2. Govmnment Accession No. N
2、ASA CR-2696 4. Title and Subtitle Reflection-Plane Tests of Spoilers on an Advanced Technology Wing with a Large Fowler Flap 7. Author(s) W. H. Wentz, Jr, and C. G. Volk, Jr. 9. Performing Organization Name and Address Wichita State University Wichita, KS 67208 2. Sponsoring Agency Name and Address
3、National Aeronautics and Space Administration Washington, D.C. 20546 5. Supplementary Notes Langley technical monitor: Harold Crane 3. Recipients Catalog No. 5. Report Date July 1976 6. Performing Orgsnization Code 8. Performing Organization Report No. WSU AR 75-2 10. Work Unit No. 11. Contract or G
4、rant No. NSG 1118 13. Type of Report and Period Covered Contractor Report 14. Sponsoring Agency Code 6. Abstract Wind tunnel experiments have been conducted to deter- mine the effectiveness of spoilers applied to a finite-span wing which utilizes the GA(W)-1 airfoil section and a 30% chord full-span
5、 Fowler flap. A series of spoiler cross- sectioned shapes were tested utilizing a reflection-plane model. Five-component force characteristics and hinge mo- ment measurements were obtained. Results confirm earlier two-dimensional tests which had shown that spoilers could provide large lift increment
6、s at any flap setting, and that spoiler control reversal tendencies could be eliminated by providing a vent path from lower surface to upper surface. Performance penalties due to spoiler leakage airflow were measured in the present tests. 7. Key Words (Suggested by Author(s) 18. Distribution Stateme
7、nt Spoilers, with Fowler Flap, reflection plane Unclassified - unlimited Subject Category 02 - 19. Security aarrif. (of this report). 22. Rice* 21. No. of Pages 20. Security Classif. (of this pw) Unclassified $4.75 88 Unclassified * For sale by the National Technical Information Service, Springfield
8、, Virginia 22161 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-,-,-SUMMARY Wind tunnel experiments have been conducted to deter- mine the effectiveness of
9、 spoilers applied to a finite-span wing which utilizes the GA(W)-l airfoil section and a 30% chord full-span Fowler flap. A series of spoiler cross- sectioned shapes were tested utilizing a reflection-plane model. Five-component force characteristics and hinge mo- ment measurements were obtained. Re
10、sults confirm earlier two-dimensional tests which had shown that spoilers cou1.d provide large lift increments at any flap setting, and that spoiler control reversal tendencies could be eliminated by providing a vent path from lower surface to upper surface. Performance penalties due to spoiler leak
11、age airflow were measured in the present tests. Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION Earlier reports (refs. 1,2) have documented the re- sults of two-dimensional wind tunnel tests of spoilers ap- plied to the GA(W)-l airfoil s
12、ection. These tests show that for certain spoiler configurations applied to an airfoil with a large Fowler flap, a control dead-band or reversal occurs for small spoiler deflections. These characteristics had also been reported in earlier NACA spoiler research with large Fowler flaps (ref. 3). The p
13、urpose of the present wind tunnel research is to obtain experimental information as to the effectiveness of spoilers applied to a three-dimensional wing utilizing the GA(W)-l airfoil with a large Fowler flap and to obtain spoiler hinge moment measurements. For this purpose, a reflection-plane model
14、was selected for the test configura- tion. The model was designed to represent a wing panel of the Advanced Technology Light Twin (ATLIT) research vehicle currently undergoing flight evaluation at NASA Langely Re- search Center (refs. 4,s). The model was designed to permit testing of various spoiler
15、 configurations, and flap settings from Oo to 40. SYMBOLS The lift, drag, and pitching,moment data have been referred to the mean .25c location of the exposed planform. Reference area for these data is exposed planform area. Rolling moment and yawing moment measurements have been re- ferred to an eq
16、uivalent airplane centerline location (be- neath the tunnel floor) , and are non-dimensionalized with respect to total equivalent span, including the portion lage. Figure 1 illustrates and areas described above. wing area and total equivalent of the wing covered by the fuse- the reference points, le
17、ngths 2 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Dimensional quantities are given in both International (S.1.) units and in U.S. Customary units. Conversion factors between the various units may be found in reference 6. The symbols used A bt C
18、 - C C SP cD cH cL cM cN in the present report are defined as follows: 2 aspect ratio, (span) /area model reference span, including image model chord at spoiler mid-span model mean aerodynamic chord, based upon exposed area, flap nested spoiler chord model drag coefficient, drag/(dynamic pressure x
19、Se) spoiler hinge moment coefficient, hinge moment/(dynamic pressure x S x c ) (opening moment is positive) section lift coefficient model lift coefficient, lift/(dynamic pres- sure x Se) model pitching moment coefficient, pitching moment/(dynamic pressure x Se x c) section lift curve slope, per deg
20、ree wing lift curve slope, per degree model rolling moment coefficient, rolling moment/(dynamic pressure x St x bt) model yawing moment coefficient, yawing moment/(dynamic pressure x St x bt) span efficiency factor spoiler trailing edge projection height Reynolds number based on mean aerodynamic cho
21、rd SP SP 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-e t Subscripts f i SP t model exposed planform area full-span planform area, including fuselage carryover angle of attack of wing root chord, degrees increment rotation of surface from nested
22、 position, degrees flap induced spoiler total EXPERIMENTAL INVESTIGATIONS Wind Tunnel Models and Instrumentation All tests were conducted using a 1/4 scale reflection- plane model representative of the exposed right-hand wing of the ATLIT airplane, without nacelles (ref. 5) . The model (figs. 1,2) w
23、as milled from solid aluminum to provide maxi- mum flexibility in machining spoiler and flap cutouts and attaching brackets, etc. The ATLIT wing utilizes the 17% thick GA(W)-1 airfoil section (ref. 1) at root and tip, with a taper ratio of 0.53, unswept 50% chord, 3 twist (washout) and 7O dihedral.
24、All testing was conducted in the WSU 2.13m x 3.05m (7 x 10 ) low speed tunnel. An aluminum disk of 0.76m (2.5) diameter was fitted at the wing root to act as an end-plate seal. This plate provided an offset of 1.27cm (0.50 inches) above the tunnel floor to minimize tunnel wall 4 . . . . . . . . . .
25、. . . - . . . . . . ._ . . . . Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-boundary layer effects. The model spar attached directly to the tunnel main balance for direct force measurements. All data have been corrected for end plate drag, and for
26、 wall effects as outlined in reference 7. Detailed computer program listings and sample calculations are given in the, Appendices. The model was fitted with a 30% chord Fowler flap, attached at four spanwise locations. A series of brackets were fabricated to provide various flap settings. A cavity w
27、as milled in the wing to simulate the spoiler cutout and approximate rib structure of the airplane. Two strain- gaged flexures were designed to provide for spoiler attach- ment and hinge moment measurement. Each flexure utilized a full four-gage bridge. A series of wedge blocks were fabricated to pr
28、ovide for spoiler deflections from -5O to +60. Several spoiler cross-sectional shapes were fabri- cated (fig. 3). Tests were conducted at a Reynolds number of 1.0 x lo6, based upon the wing mean aerodynamic chord length of 29.21 cm (11.50 inches) . Flap-Nested Performance Predictions of finite-span
29、wing performance may be made from two-dimensional data by applying the following corrections: (a) Adjust the angle of zero lift to account for wing twist at the M.A.C. For this model the M.A.C. twist is -1.4 relative to the wing root chord reference. (b) Correct the lift-curve slope according to the
30、 following formula from reference 8: 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-, -. . . . . , , Correct the drag by adding an induced drag term given by the following formula from reference 8: 2 C, In these equations, e is the span efficiency
31、 factor, taken as 0.8. Applying the offset and slope change calculated as described in (a) and (b) , the experimental C vs. a relation- ship gives the predicted three-dimensional relationship shown in figure 4. It is seen that this prediction agrees well with the experimental three-dimensional data,
32、 even through the stall. 1 Table 1 illustrates comparisons of some predicted and experinental aerodynamic parameters. The two-dimensional values are from reference 9. Table 1 - Predicted and Experimental Three- Dimensional Aerodynamic Parameters - 2-D Value Predicted Experimental Parameter (ref. 9)
33、3-D Value 3-D Value Zero lift angle Lift curve slope Maximum lift coefficient -3.8“ -2.4“ -2.6“ 0.112/deg. .0883/deg. .087/deg. 1.57 1.35 1.36 A predicted drag polar for the flap-nested case is developed from two-dimensional data, with the added induced drag based upon an 80% span efficiency factor.
34、 As shown in 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-the figure, the experimental threeydimensional relationship agrees well with the prediction, except for lift coeffi- cients near stall. Two- and three-dimensional pitching moment data are
35、 also compared in figure 4. In this case the pitching moment values are compared directly, based upon measurements re- ferred to the 25% mean aerodynamic chord of the three-dimen- sional planform., The comparison shows good agreement. Flap-Extended Performance A series of baseline runs were made to
36、.obtain the aerodynamic characteristics of the basic reflection-plane wing for various flap settings, with spoilers closed and sealed. These data (figs. 5, 6, and 7) show the basic lift, drag and pitching moment characteristics of the re- flection-plane model. Tabulated flap gap and overlap are show
37、n for each flap deflection. These settings, with the exception of the 40 case, are the same as those developed from the twb-dimensional GA(W)-l tests reported in reference 9. During initial force tests of the wing with 40“ flap deflection, it was discovered that expected values for CLmX were not bei
38、ng attained. Tuft studies revealed that the flow over the flap was separated at all angles of attack. The flap brackets were then modified to provide for gap and overlap adjustment. Figure 8 illustrates- definitions of gap and overlap. From the tuft studies it was determined that attached flow on th
39、e flap could be achieved with modi- fied settings. Table 2 compares the best gap and overlap values for the present three-dimensional tests and the earlier two-dimensional tests. 7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Table 2 - Best Flap S
40、ettings for 40 Deflection Source Reynolds Gap_ Overlap Number 2-D (ref. 9 tests) 2.7% -0.7%c 2.2 x 10 3-D (Present tests) 2.2%c +O .8%c 1.0 x 10 6 6 The results of the revised gap and overlap settings (fig. 8) show substantial improvement in linearity of the lift curve as well as in CLmx performance
41、. All subsequent 40 flap testing was done with the revised gap and overlap setbings. The discrepancy in optimum flap settings between two- dimensional and three-dimensional tests merits discussion. Both tests utilized models of rigid construction to avoid possibie aero-elastic deflection problems. T
42、he Reynolds numbers of the tests do differ, but only by a factor qf 2. The tunnel balance load limits prohibited testing the re- flection plane model at CL with 40 flap at RN = 2 x 10 . It was possible to test this model at RN = 2 x lo6 at zero angle of attack, however. This testing showed that the
43、flap flow was not attached, indicating that the Reynolds number change is probably not responsible for the change in separa- tion observed. 6 Three-dimensional effects on the flap slot flow are difficult to assess. However, for a wing with zero sweep (at 50% chord) little spanwise flow is to be expe
44、cted. Tuft patterns tend to substantiate this, at least in the absence of separation. The reasons for the discrepancy in gap and overlap between two-dimensional and three-dimensional tests remain unexplained. 8 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from
45、 IHS-,-,-It is noted that the lift curves with flap extended have steeper slopes than the flap-nested case. This is ex- pected, since the Fowler action provides an . increase in effective chord. By accounting for the increase in wing area for each flap deflection it is possible to predict lift curve
46、 slopes for the flap extended cases. Results of calculations of this type are shown in figure 9, along with the experimental values. Agreement between experiment and prediction is. good. Flap effectiveness in producing increments in CL and in CL at a = Oo is also shown in figure 9, along with corres
47、ponding data from the two-dimensional tests. These data show that the three-dimensional flap effectiveness is 70% to 80% of the two-dimensional values. For 40 flap, the present tests yield a CL value of 3.0 compared to a section Clmax of 3.8 from two-dimensional tests. Using the method of reference
48、10, it is possible to calculate a c,orrection factor relating two-dimensional flap effectiveness to three-dimensional effectiveness. For the ATLIT reflection-plane model, this parameter is 93% which compares ” unfavorabll with the measurements as related above. The reasons for this discrepancy are not clear. I Theoretical drag polars for various flap settings have been caiculated utilizing the experimental zero lift drag and span efficiency factors of .l.O and 0.8. For this analysis no accounting has been made of section drag in- creases with lift
