1、Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-NASA TM X-2284 7. Author(s) Samuel M. Dollyhigh I 4. Title and Subtitle STABILITY AND CONTROL CHARACTERISTICS AT MACH NUMBERS 1.60 TO 2.86 OF A VARIABLE-SWEEP FIGHTER CONFIGURATION WITH SUPERCRITICAL AI
2、RFOIL SECTIONS (U) 8. Performing Organization Report No. 1 L-1583 I 5. Report Date June 1971 6. Performing Organization Code 10. Work Unit No. NASA Langley Research Center National Aeronautics and Space Administration An experimental investigation has been made in the Mach number range from 1.60 to
3、2.86 to determine the longitudinal and lateral aerodynamic characteristics of a variable- relatively high levels of zero-lift pitching moment, results in a high instantaneous normal acceleration capability for the configuration at a Mach number of 1.60 and an altitude of 10 668 m (35 000 ft). Indica
4、tions were that the normal-acceleration capability could be increased somewhat by reducing the longitudinal stability of the configuration. The lateral- stability results, however, indicated rather poor directional characteristics for an angle of attack greater than 10. For the Mach number range of
5、the tests, static directional stability was maintained only to angles of attack of 10 to 12O, which were well below the angles of attack at which the configuration had longi 17. Key Words (Suggested by Author(s) Fighter configuration Supercritical airfoil 19. Security Classif. (of this report) 20. s
6、ecvL “ Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-STABILITY AND CONTROL CHARACTERISTICS AT MACH NUMBERS 1.60 TO 2.86 OF A VARIABLE-SWEEP FIGHTER CONFIGURATION WITH SUPERCRITICAL AIRFOIL SECTIONS * By Samuel M. Dollyhigh Langley Research Center S
7、UMMARY An experimental investigation has been made in the Mach number range from 1.60 to 2.86 to determine the longitudinal and lateral aerodynamic characteristics of a variable- sweep fighter configuration with supercritical airfoil sections, twin rectangular inlets, twin vertical tails, and boom-m
8、ounted aft horizontal tails. The results of the investiga- tion indicated that the configuration had good longitudinal stability characteristics and a relatively high horizontal-tail control effectiveness. This horizontal-tail effectiveness, coupled with relatively high levels of zero-lift pitching
9、moment, results in an instanta- neous normal-acceleration capability of approximately log for the configuration at a Mach number of 1.60 and an altitude of 10 668 m (35 000 ft). Indications were that the normal- acceleration capability could be increased somewhat by reducing the longitudinal stabili
10、ty of the configuration. The lateral-stability results, however, indicated rather poor direc- tional characteristics for an angle of attack greater than 10. For the Mach number range of the tests, static directional stability was maintained only to angles of attack of lo0 to 12O, which were well bel
11、ow the angles of attack at which the configuration had longitudinal trim capability. INTRODUCTION The National Aeronautics and Space Administration has recently completed a study of advanced configuration concepts applicable to fighter aircraft to assess attainable per- formance levels and to provid
12、e a focal point for current and future research programs. One of the configurations studied was a variable-sweep fighter, designated LFAX 4, which incorporated a supercritical-wing airfoil section in an effort to provide the good subsonic and supersonic aerodynamic characteristics required for multi
13、mission capability. In order to minimize the shift in aerodynamic center with wing sweep and thus avoid large static margins at supersonic speeds, an outboard wing-pivot location was selected. The * Title, Unclassified. Provided by IHSNot for ResaleNo reproduction or networking permitted without lic
14、ense from IHS-,-,-twin horizontal tails were mounted low and outboard on booms to provide a uniform varia- tion of pitching moment with lift and to provide good control effectiveness and high instan- taneous normal-acceleration capability. The LFAX 4 was designed to have minimum wave drag at a Mach
15、number of 1.60 by a computer program that evolved from the method discussed in reference 1 and good drag due to lift and pitching-moment characteristics at a Mach number of 1.60 by modifying the wing-twist distribution by the method of reference 2. Because of the importance of good subsonic performa
16、nce in a fighter with multimission capability, the wing-camber distri- bution was designed from subsonic considerations. The purpose of the present investigation was to determine the experimental longi- tudinal and lateral aerodynamic characteristics of the LFAX 4 configuration at supersonic speeds.
17、 Wind-tunnel tests of a scaled model of the LFAX 4 were conducted in the Langley Unitary Plan wind tunnel at Mach numbers from 1.60 to 2.86. The results of the wind- tunnel investigation are reported herein. SYMBOLS The longitudinal results (P = 0) are referred to the wind-axis system, and the later
18、al results are referred to the body-axis system. All coefficients are based on the geometric characteristics of the wing in the 30 leading-edge-sweep position (with leading and trailing edges projected into plane of symmetry). The moment reference point was located at fuselage station 54.935 cm (21.
19、628 in.). Measurements and calculations were made in U.S. Customary Units. They are presented herein in the International System of Units (I) with the equivalent values given parenthetically in the U.S. Customary Units. A aspect ratio b wing span c aerodynamic chord - c mean aerodynamic chord drag c
20、oefficient, Drag qs base-drag coefficient, Base drag qS Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Chamber drag chamber-drag coefficient, qs internal-drag coefficient, Internal drag CIS drag coefficient at zero lift drag-due-to-lift parameter Li
21、ft lift coefficient, - qs slope of lift curve, per deg Rolling moment rolling-moment coefficient, qsb AC effective-dihedral parameter, - , per deg (where P = O0 and 3) A6 Pitching moment pitching-moment coefficient, qSF longitudinal stability parameter pitching effectiveness of horizontal tail Yawin
22、g moment y awing-moment coefficient, qsb AC directional-stability parameter, - “, per deg (where P = 0 and 3) A6 side-force coefficient, Side force qs side-force parameter, - per deg (where 6 = 00 and 3O) A6 lift-drag ratio free-stream Mach number Provided by IHSNot for ResaleNo reproduction or netw
23、orking permitted without license from IHS-,-,-cd. free-stream dynamic pressure X reference area of wing including fuselage intercept x longitudinal direction Y lateral direction z vertical direction a angle of attack, deg P angle of sideslip, deg r dihedral angle, deg 6h horizontal-tail deflection a
24、ngle, positive when trailing edge is down, deg A leading-edge sweep angle, deg Model components: B body H horizontal tail U ventral fin V vertical tail W wing Subscript: max maximum DESCRIPTION OF MODEL A three-view drawing of the complete model configuration is shown in figure 1, and some geometric
25、 characteristics are given in table I. A photograph of the model is Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-presented in figure 2. The configuration incorporates a fuselage; a variable-sweep wing with horizontal ramp-type inlets located ahead
26、 of the wings; twin horizontal tails, twin vertical tails, and twin ventral fins mounted on booms; and lateral strakes on the forebody. The wing on this configuration utilized the variable-sweep concept with an outboard pivot and had a supercritical airfoil section. The streamwise airfoil ordinates
27、are defined with the wing in the 30 sweep position. The airfoil ordinates are presented in table I1 for three spanwise stations. The wing had a thickness ratio of 11.0 percent at the wing pivot station and 10.0 percent at the wing tip. The range of wing leading-edge sweep angle was from 23O (forward
28、 position) to TI0 (aft position). There was an intermediate sweep angle of 30 for subsonic cruise and maneuver that was used as the reference sweep angle. The model was tested with the wing fixed in the swept mode. The wing in the aft sweep position (A = ?I0) had optimum twist for a Mach number of 1
29、.6. The configuration employed low twin horizontal tails with an airfoil section that was 4 percent biconvex. The horizontal tails had a negative dihedral of 9.3O. The twin verti- cal tails were toed-in at an angle of 12O and had an airfoil section that was of a 3.5 percent half-biconvex. The ventra
30、l fins were made from a 0.318-cm-thick (0.125-in.) flat plate with a beveled-edge angle of 30. The tail surfaces and ventral fins could be removed from the model, and the horizontal tail could be deflected over a range from 5 to -20. TESTS AND CORRECTIONS The tests were conducted in the Eangley Unit
31、ary Plan wind tunnel at Mach numbers of 1.60, 2.00, 2.50, and 2.86. The Reynolds number was 9.8 X lo6 per meter (3.0 X 106 per foot) for all tests except those at the higher angles of attack at Mach numbers of 1.60, 2,00, and 2.50. For these tests, the Reynolds number was reduced to 8.2 X 106 per me
32、ter (2.5 X per foot) and 6.6 X lo6 per meter (2.0 X lo6 per foot) in order to stay within balance load limits. The dewpoint was maintained sufficiently low to prevent measurable condensation effects in the test section. The angle-of-attack range was from approxi- mately 6 to 23O, and the angle of si
33、deslip was from approximately -4 to 6. In order to assure boundary- layer transition to turbulent flow, 0.16-cm-wide (1/16 in.) transition strips of No. 50 carborundum grit were placed on the body 3.05 cm (1.2 in.) aft of the nose of the model and 1.02 cm (0.40 in.) streamwise on the wings, tails, v
34、entral fins, inlet ram,ps , and external inlet surfaces. Aerodynamic forces and moments on the model were measured by means of a six- component strain-gage balance which was housed within the model. The balance was attached to a sting which in turn was rigidly fastened to the tunnel support system.
35、Balance-chamber static pressures were measured with pressure tubes located in the vicinity of the balance. The internal-flow total and static pressures were measured in Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-each duct with a rake consisting
36、of 13 total-pressure tubes and 5 static-pressure tubes; the rakes were placed 0.25 cm (0.10 in.) inside the exit of the ducts. The base pressure was measured by two pressure tubes which were fastened to the sting and held in proximity to the base of the model. The internal-flow pressure rakes and ba
37、se-pressure tubes were removed during the force-measurement tests. Corrections to the indicated model angles of attack and sideslip have been made for both tunnel-airf low misalinement and deflection of the balance and sting under load. Figure 3 documents that the aerodynamic character- istics for t
38、he model upright and inverted are the same. The drag data presented herein have been corrected for internal drag and also corrected to the condition of free-stream static pressure at the model base and in the balance chamber. Figures 4 to 6 show the values of chamber, base, and internal drag which w
39、ere used to correct the drag data. Fluorene sublimation material was used to determine the natural-transition loca- tions on the clean model at zero lift coefficients. An increment in skin friction was ana- lytically calculated between the data for natural transition and the condition of all-turbule
40、nt flow behind the transition strips. This correction was added to the natural-transition data. The trip drag was considered to be the difference between the drag data with transition strips on the model and the corrected natural-transition data. The trip drag is presented, but the wind-tunnel data
41、are not corrected for trip drag. The geometry of the aft end of the LFAX 4 configuration was such that with internal flow there was interference between the sting and model booms. In order to estimate the effects of this interference, the aft end of the model was “extended“ to the end of the tail by
42、 using a sleeve attached to the sting. The sleeve as pictured in figure 7 was a contin- uation of the base of the model and was attached to the sting with a gap approximately 0.102 cm (0.040 in.) between the model and sleeve. The gap between model and sleeve was sufficient to insure that no forces o
43、n the sleeve were measured by the model balance. The sleeve prevented the internal flow from expanding and resulting in interference between the sting and model booms; thus, the actual flight condition was more nearly represented. The effects on longitudinal characteristics of adding the aft body ex
44、tension are presented, but no corrections due to removing the estimated sting interference are applied to the remaining data. PRESENTATION OF RESULTS The moment reference point is on the body axis 54.935 cm (21.628 in.) from the nose of the model. The results are presented in the following figures:
45、Figure Upright and inverted characteristics, sleeve off 3 Base-drag coefficient 4 Chamber-drag coefficient 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Figure Internal-drag coefficient . 6 Longitudinal control characteristics 8 Summary of longit
46、udinal control characteristics 9 Longitudinal trim characteristics 10 Effects of components on longitudinal characteristics . 11 Effects of afterbody extension on longitudinal characteristics . 12 Tripdrag . 13 Sublimation photographs . 14 Oil-flow photographs . 15 Lateral characteristics . 16 Effec
47、ts of vertical components on lateral characteristics at a! = -I0 . 17 Effects of vertical components on lateral characteristics at a! 6 . 18 Effects of vertical components on sideslip derivatives . 19 Effects of tail deflection on sideslip derivatives 20 DISCUSSION Longitudinal Characteristics The l
48、ongitudinal aerodynamic and control characteristics of the basic LFAX 4 con- figuration are shown in figure 8 for the test Mach numbers from 1.60 to 2.86. The con- figuration appears to have good stability characteristics and good horizontal-tail control effectiveness throughout the Mach number rang
49、e of the tests. The linearity of the pitching- moment variation uith lift and the control-surface effectiveness are due principally to the low, outboard locatilm of the twin horizontal tails. As presented, the configuration prob- ably has excessive stability for the supersonic conditions. The center-of-gravity location upon which the dat:. of f
