1、NASA Tech n ica I Paper 2772 1987 National Aeronautics and Space Administration Scientific and Technical Information Division Wind-Tunnel Investigation of a Full-Scale General Aviation Airplane Equipped With an Advanced Natural Laminar Flow Wing Daniel G. Murri and Frank L. Jordan, Jr. Langley Resea
2、rch Center Hampton, Virginia Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Summary An investigation has been conducted in the Lang- ley 30- by 60-Foot Tunnel to evaluate the perfor- mance, stability, and control characteristics of a full-scale gene
3、ral aviation airplane equipped with an advanced natural laminar flow wing. The study fo- cused on the effects of natural laminar flow and pre- mature boundary-layer transition on performance, stability, and control, and also on the effects of sev- eral wing leading-edge modifications on the stall/de
4、parture resistance of the configuration. Data were measured over an angle-of-attack range from -6 to 40 and an angle-of-sideslip range from -6 to 20. The Reynolds number was varied from 1.4 x lo6 to 2.4 x lo6 based on the mean aerody- namic chord. Additional measurements were made using hot-film and
5、 sublimating chemical techniques to determine the condition of the wing boundary layer, and wool tufts were used to study the wing stall characteristics. The investigation showed that large regions of natural laminar flow existed on the wing which would significantly enhance the cruise performance o
6、f the configuration. Also, because of the characteristics of the airfoil section, artificially tripping the wing boundary layer to a turbulent condition did not sig- nificantly affect the lift, stability, and control char- acteristics. The addition of a leading-edge droop ar- rangement was found to
7、increase the stall angle of attack at the wingtips and, therefore, was consid- ered to be effective in improving the stall/departure resistance of the configuration. Also, the addition of the droop arrangement resulted in only minor in- creases in drag. The configuration exhibited good longitudinal
8、stability and control characteristics for all test conditions and stable effective dihedral up to the angle of attack for wing stall. The directional stability characteristics were generally poor at the higher angles of attack because of the loss of vertical tail effectiveness as angle of attack inc
9、reased. The lateral-directional control characteristics were satis- factory, except near wing stall where large yawing and rolling moments were encountered as a result of asymmetric wing stall. Introduction In recent years, studies have shown that sig- nificant improvements in the performance of gen
10、- eral aviation and commuter aircraft are possible from the realization of increased amounts of nat- ural laminar flow (NLF) (refs. 1 to 5). These results have been achieved in part through ad- vanced NLF airfoil design and modern construc- tion materials and fabrication techniques such as composite
11、s and milled or bonded aluminum skins. The emphasis in airfoil design has been directed toward developing airfoils with extensive natural laminar flow in an attempt to obtain lower cruise drag coefficients while maintaining acceptable max- imum lift and stall characteristics. One airfoil designed wi
12、th these considerations is designated the NASA NLF( 1)-0414F. (See refs. 4 and 5.) The current tests were conducted in a coopera- tive program between the NASA Langley Research Center and the Cessna Aircraft Company by test- ing a full-scale modified Cessna T-210 airplane in the Langley 30- by 60-Fo
13、ot Tunnel (figs. 1 and 2). This airplane features a modified wing of increased aspect ratio and incorporates the NASA NLF( 1)-0414F air- foil. A primary objective of these tests was to doc- ument the characteristics of the airfoil in this ap plication and to determine the effects of premature bounda
14、ry-layer transition on the overall airplane per- formance, stability, and control. In addition, results are presented concerning the effects of power and flap deflections on the longitudinal characteristics and the lateral-directional stability and control, and also the effects of fairing the airfoi
15、l trailing-edge reflex. The tests with the faired trailing-edge reflex were con- ducted to evaluate the effects of changing the airfoil contour to a shape that would be much easier and less expensive to fabricate. Additional results are presented concerning the effects of several wing leading-edge m
16、odifications applied to the modified Cessna T-210. These tests were conducted to determine whether leading- edge modifications previously shown to provide ex- cellent stall/spin resistance on more conventional wing/airfoil configurations (refs. 6 to 8) could be de- veloped for application to an NLF
17、wing design of high aspect ratio. One approach recently studied in exploratory re- search (ref. 9) was to use the NASA NLF(1)-0414F airfoil for enhanced performance and to use another NLF airfoil of current interest, the NASA NLF(1)- 0215F (ref. 3), for the leading-edge droop design. A leading-edge
18、droop was developed for the current configuration in subscale tests in the Langley 12-Foot Low-Speed Tunnel using a wingtip balance to mea- sure the aerodynamics of the outer wing panel. The droop was developed from the NLF( 1)-0215F airfoil by gloving over the leading-edge outboard panel of the bas
19、ic wing. An important feature of the droop is the abrupt discontinuity of the droop inboard lead- ing edge. This discontinuity is effective in generating a vortex that acts as an aerodynamic fence to stop the spanwise flow from the inboard portion of the wing as stall progresses outward. The leading
20、-edge droop extends to near the wingtip such that the outer I Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-portion of the wing performs as a low-aspect-ratio wing with a very high stall angle of attack. This earlier research also revealed that on
21、this particular configuration the effectiveness of the outboard droop could be enhanced by the addition of a small-span droop located inboard on the wing. Results are pre- sented from the current tests which show the effects of the leading-edge modifications on the wing stall characteristics and the
22、 associated effects on stabil- ity and control, roll damping, and calculated cruise performance. Symbols All longitudinal forces and moments are referred to the wind axis system, and all lateral-directional forces and moments are referred to the body axis system. Moment data are presented with respe
23、ct to a center-of-gravity position of 25 percent of the wing mean aerodynamic chord. b wing span, ft CD drag coefficient, 9 AcD incremental drag coefficient CL lift coefficient, 3 ;i rolling-moment coef- ficient, positive with right wing down, Rolling moment Qm Sb W incremental rolling- c7n pitching
24、-moment coeffi- cient, positive with nose Pitching moment moment coefficient UP, qw SE Acm incremental pitching- cn yawing-moment coeffi- moment coefficient cient, positive with nose to right, Yawing ;merit 900 Acn incremental yawing- CT thrust coefficient, % CY side-force coefficient, pos- moment c
25、oefficient itive to right, SiTwfy coefficient AcY incremental side-force C I f h k P 900 R S V W 2 z a P P Sa 6a,eff 6Cf 6, 6, 6, local wing chord, ft mean aerodynamic chord (reference), ft frequency of oscillation, Hz altitude, ft reduced-frequency param- eter, wb/2V roll rate, rad/sec free-stream
26、dynamic pressure, lb/ft2 Reynolds number based on I wing reference area, ft2 velocity, ft/sec (or knots as indicated) weight, lb chordwise distance from wing leading edge, positive aft, ft vertical distance from wing leading edge, positive up, ft angle of attack, deg angle of sideslip, deg time rate
27、 of change of angle of sideslip, rad/sec aileron deflection, positive with trailing edge down, deg effective aileron deflec- tion, 2 , deg however, a polyester resin filler material was applied to the wing to obtain the de- sired airfoil section contours. The waviness of the wing surface when measur
28、ed in a chordwise direc- tion was maintained within f0.003 in. per 2 inches of wavelength. Drawings of the model geometry and photographs of the model mounted in the Langley 30- by 60-Foot Tunnel are shown in figures 1 and 2, respectively. Also, a summary of the model geomet- ric characteristics is
29、presented in table I. The wing incorporates a small 12.5-percent-chord trailing-edge “cruise” flap that is designed to vary the low-drag lift coefficient range with small flap deflections. This cruise flap could also be set to large trailing-edge- down deflections (to 40) to enhance the rnaximum lif
30、t characteristics. Roll control is provided by a combination of ailerons and spoilers. The NASA NLF(1)-0414F airfoil section, shown in figure 3, is designed to achieve low-profile dra ce maintaining natural laminar flow to about 70 percent chord on both upper and lower surfaces. Airfoil coor- dinate
31、s are provided in table 11, and a more-detailed description of the airfoil and its characteristics can be found in reference 5. . The control settings were tested in the follow- ing ranges: 6, = -25 to 15, however, the data of figure 13(a) indicate re- duced elevator effectiveness at angles of attac
32、k above wing stall. The directional control effectiveness of the rudder is summarized in figure 14. The data indicate that the rudder generally maintains yaw-control effective- ness up to the angle of attack for wing stall (a x 17) and exhibits reduced effectiveness at higher angles of attack. At an
33、 angle of attack of 17, however, large yawing and rolling moments are encountered that are a result of asymmetric wing stall. The data of fig- ure 14(a) indicate that these large moments at the stall are greater than the yawing moment provided by full rudder deflection. In order to determine the iso
34、lated effects of the lateral control surfaces, tests were conducted using individual aileron and spoiler deflections. Presented in figure 15 is the effect of deflecting the left aileron alone. The data show an initial reduction in aileron effectiveness at an angle of attack of about 5, prob- ably be
35、cause of wing trailing-edge separation. This reduction in aileron effectiveness may be delayed to higher angles of attack at higher Reynolds numbers because of the effect of increasing Reynolds number to delay trailing-edge separation (discussed earlier). Nonetheless, good control effectiveness is i
36、ndicated up to wing stall, and then reduced effectiveness is shown at higher angles of attack. The data of fig- ure 16 show that the spoilers were much less effective in providing roll control than the ailerons. A compar- ison of the data in figures 16(a) and 16(b) indicates that spoiler deflections
37、 of 18 or less were ineffective in providing roll control, except at negative angles of attack. At Q = O, spoiler deflections greater than Y 18 are shown to provide a fairly linear variation in rolling moment; however, all usable spoiler roll- control effectiveness is lost by an angle of attack of a
38、bout 12. Therefore, the only roll control available at the stall is provided by the ailerons. The aileron- effectiveness results of figure 15 (for the left aileron alone), however, suggest that the total roll control provided by the ailerons may become marginal at wing stall. Boundary-Layer Study Fr
39、ee-transition characteristics. The boundary- layer transition characteristics of the basic wing were measured using both the sublimating chemical tech- nique (ref. 1) and the hot-film technique (ref. 11). An example of one of the sublimating chemical tests is shown in figure 17(a) for a = 1 and R =
40、2.4 x lo6. The test results showed that laminar flow was main- tained to about 70 percent chord on both upper and lower surfaces at the cruise angle of attack. These re- sults agree well with the theoretical predictions and the two-dimensional-airfoil transition measurements from references 4 and 5.
41、 These data also agree well with the transition measurements made in flight by Cessna with the same modified Cessna T-210. (See ref. 18.) Presented in figures 17(b) and 17(c) are the upper and lower surface boundary-layer transition charac- teristics measured using the hot-film technique. The data s
42、how the movement of the transition location with changes in angle of attack and cruise flap deflec- tion. Some exceptions were noted when comparing the forward movement of transition with changes in angle of attack from present test results with those of the flight tests (ref. 18). For example, the
43、wind- tunnel data indicate earlier forward movement of transition on the upper surface as angle of attack is decreased and on the lower surface as angle of attack is increased. These differences are probably due to differences in Reynolds number and turbulence level between the conditions in the win
44、d tunnel and in cruise flight and are discussed further in reference 18. Boundary-layer transition. With such a large ex- tent of natural laminar flow occurring on the wing of this configuration, the obvious question that arises is the effect of artificially tripping the laminar bound- ary layer to
45、a turbulent condition. In an operational environment, periodic wing cleaning (possibly before each flight) would probably be required to ensure the performance benefit from natural laminar flow. Once airborne, however, premature boundary-layer transition could occur after insect contamination or dur
46、ing flight in moisture, and it could potentially result in changing the trim or stability and control 7 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-characteristics (ref. 9). Therefore, designing a wing with a natural laminar flow airfoil to minim
47、ize these characteristics is very desirable. The effects on the total airplane characteris- tics of artificially tripping the boundary layer at z/c = 0.05 are presented in figure 18. The data of figure 18(a) indicate the very desirable characteristic that boundary-layer transition results in essenti
48、ally no change in the lift and pitching-moment charac- teristics. Of particular interest is the characteristic mum lift remain unchanged for either boundary-layer condition. The total airplane drag characteristics with artificial boundary-layer transition are shown in figure 18(b). As would be expec
49、ted, the turbulent boundary-layer condition results in large increases in drag at cruise lift coefficients. The total increase in drag ACD at a cruise lift coefficient CL of 0.3 is shown to be about 0.0070. The calculated performance increases of the mod- ified Cessna T-210 due to the large extent of laminar flow are illustrated in figure 19. These performance calculations are based on trimmed values of CL and CD and are made for an altitude of 10 000 ft, a weight of 3500 lb, and at 75 perce
