1、NASA Technical Paper 1878 Experimental and Theoretical Supersonic Lateral-Directional Stability Characteristics of a Simplified Wing-Body Configuration With a Series x of Vertical-Tail Arrangements Milton Lamb, Wallace C. Sawyer, and James L. Thomas AUGUST 1981 Provided by IHSNot for ResaleNo reprod
2、uction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM NASA Technical Paper 1878 Experimental and Theoretical Supersonic Lateral-Directional Stability Characteristics of a Simplified Wing-Body Configuration With a Series of Vertical-Tail Arrangements Milton Lamb, Wallace C
3、. Sawyer, and James L. Thomas Latzgley Research Cerrter Hamptotz, Virgirtia National Aeronautics and Space Administration Scientific and Technical Information Branch 1981 L Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-SUMMARY An experimental inves
4、tigation has been conducted to provide a systematic set of lateral-directional stability data for a simplified wing-body model with a series of vertical-tail arrangements. The study was made at Mach numbers from 1.60 to 2.86 at nominal angles of attack from -8O to 12O and Reynolds number of 8.2 X 1
5、O6 per meter. Comparisons at zero angle of attack have been made with three existing theoretical methods (MISLIFT - a second-order shock expansion and panel method; APAS - a slender body and “first order“ panel method; and PAN AIR - a “higher order“ panel method) and comparisons at angle of attack h
6、ave been made with PAN AIR. The results show that PAN AIR generally provides accurate estimates of these characteristics at moderate angles of attack for complete configurations with either single or twin vertical tails. APAS provides estimates for complete configurations at zero angle of attack. Ho
7、wever, MISLIFT only provides esti- mates for the simplest body-vertical-tail configurations at zero angle of attack. INTRODUCTION Computer codes have been developed for rapid accurate estimates of the aerodynamic characteristics of aircraft and missile configurations at supersonic speeds. Much atten
8、tion has been given to the development and assessment of these methods for predicting the lift, drag, and pitching-moment characteristics of complex configurations. Many of these methods have the capability of predicting the lateral-directional characteristics of aircraft and missiles, but their uti
9、l- ity has not been evaluated by comparison with experiment. An effort has been initiated at NASA Langley Research Center with the pri- mary purpose of providing experimental data on simple wing-body-vertical-tail configurations with which to assess lateral-directional stability estimating technique
10、s at supersonic speeds. The longitudinal aerodynamic characteristics of the configurations were also recorded during the test program and are included in tabular form. These experimental data are presented along with an assessment of three of the existing methods capable of estimating lateral-direct
11、ional param- eters. The methods include a second-order shock expansion and panel method (MISLIF“, ref. 1) , a slender body and “first order“ panel method (WAS, ref. 2) , and a “higher order“ panel method for linearized supersonic flow (PAN AIR, ref. 3). “ SYMBOLS Force and mment data are referred to
12、 the body axis system except for lift and drag data which are referenced to the stability axis system. The moment ref- erence center was located at 75.6 percent of the body length. The model was Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-designe
13、d, built, and the data were reduced using the U.S. Customary Units; however, all data are presented in the SI Units. reference area, maximum cross-sectional area of body, 0.00456 m2 aspect ratio wing or tail span (exposed) , cm Drag drag coefficient, - SA Lift lift coefficient, - Rolling moment roll
14、ing-moment coefficient, SAd effective dihedral parameter (roll stabilityparameter) , Pitching moment SA1 pitching-moment coefficient, Yawing moment SAd yawing-moment coefficient, directional stability parameter , Side force side-force coefficient, SA side-force parameter , root chord tip chord body
15、diameter, 7.62 cm Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-z body length, 88.90 cm M free-stream Mach number M.S. model station (measured from nose), cm 9 free-stream dynamic pressure, Pa t panel maximum thickness, cm 01 angle of attack, deg B
16、 angle of sideslip, deg A leading-edge sweep angle, deg x Ct taper ratio, - Model B Vl v2 v3 v4 V5-2d 5-4d V6-2d V6-4d W J components: body vertical tail 1 , vertical tail 2, vertical tail 3, vertical tail 4, vertical tail 5, vertical tail 5, vertical tail 6, vertical tail 6, wing body mounted body
17、mounted body mounted body mom t ed wing mounted with 2-body-diameter spacing wing mounted with 4-body-diameter spacing wing mounted with 2-body-diameter spacing wing mounted with 4-body-diameter spacing TEST PROCEDURE, MODEL, AND DATA Test Procedure The investigation was conducted in the low Mach nu
18、mber test section of the Langley Unitary Plan Wind Tunnel, which is a variable-pressure continuous-low facility. The test section is approximately 2.1 3 meters long and 1.22 meters 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-square. For the pre
19、sent tests, the Mach number, stagnation pressure, and stag- nation temperature were as follows: I I Stagnation temperature, K pressure, kPa Stagnation - 1 .60 2.36 339 68.28 2.86 339 79.43 123.05 1 339 c - . . .“ The nominal test Reynolds number was 8.2 X 1 O6 per meter. The dew-point temperature me
20、asured at stagnation pressure was maintained below 239 K to assure negligible condensation effects. All tests were con- ducted with boundary-layer transition strips on the body 2.05 cm aft of the nose and 1.02 an aft (measured streamwise) of the leading edges of the wing and tail surfaces. The trans
21、ition strips consisted of No. 60 sand sprinkled in approxi- mately 0.1 57-cm-wide strips. Mode 1 A two-view sketch of a typical model configuration is shown in figure l(a). The body consisted of a 3.5 caliber tangent ogive nose followed by a cylindrical section with an overall fineness ratio of 11.6
22、7. The wings and vertical tails were removable to permit a wide variation of model configurations to be tested. Four of the configurations had a vertical tail mounted on the body, and the two other configurations had vertical tails mounted at two spanwise locations on the wings. Details of the verti
23、cal tails are given in table I and figure 1 (b) and those of the wing in table I and figure 1(c). The leading and trailing edges of the tails and wings were sharp wedges having total angles measured in a plane perpendicular to the edges of 24O and 20, respectively. The two spanwise loca- tions for t
24、he wing mounted tails are also shown in figure l (c) . Data Measurements and Corrections Aerodynamic forces and moments on the model were measured by means of a six-component electrical strain-gage balance which was housed within the model. The balance was attached to a sting which was rigidly faste
25、ned to the tunnel support system. Balance chamber pressure was measured by means of a static- pressure or ifice located in the vicinity of the balance. 6 “ The angles of attack were corrected for deflection of the balance and sting due to aerodynamic loads and tunnel-flow misalignment. The drag-coef
26、ficient data were adjusted to free-stream conditions acting over the model base. 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-THEOlU3TICAL METHODS MISLIFT: The MISLIFT method (ref. I), which estimates lift-induced pres- sures on a wing-body comb
27、ination, incorporates two separate and distinct theo- ries. Specifically, the second-order shock-expansion theory of reference 4 is used to obtain the lifting pressures on the body alone at small angles of attack and the lifting pressures induced by a lifting surface are evaluated by the numerical s
28、olution to the linear-theory integral equations of reference 5. The mumerical solution of these equations is effected by treating the planform as a planar composite of elemental rectangles and applying summation techniques to satisfy the necessary integral relations. A further description of the the
29、oret- ical method and comparison with experimental results can be obtained from ref- erence 1. This method predicts the aerodynamic characteristics for a planar- type configuration. And, for the present study, the model was taken to be a planar configuration in the side view in order to estimate the
30、 lateral- directional characteristics. APE: The Aerodynamic Preliminary Analysis System (ref. 2) is a “first order“ panel method solution for linearized subsonic and supersonic flight. Angle of attack and sideslip solutions are handled independently. A general slender-body theory is used to represen
31、t body-alone effects. The perturbation velocities from the body are included in a Woodward “constant pressure analysis method for thin lifting surfaces. Interference shells are used in the lifting- surface analysis to carry over the load from the wings onto the body. The lifting-surface method uses
32、a vortex singularity of constant strength in the chordwise direction on each of the panels into which the lifting surfaces are divided. The loads on the lifting surfaces are calculated directly from the solution using the first-order pressure-coefficient relation, accounting for only the perturbatio
33、n velocities in the free-stream direction. The total forces and moments are a sum of the slender-body and lifting-surface solutions. PAN AIR: The PAN AIR pilot code (ref. 3) is a “higher order“ panel method solution for linearized subsonic and supersonic flow. Combined source and doublet panels with
34、 linearly varying source and quadratically varying doublet distributions can be used. The quadrilateral panels formed from a rectangular array of input points are each divided into eight triangular flat subpanels in such a way that all panel edges are contiguous with adjacent panels. Quadratic doubl
35、et distributions over each triangular subpanel are prescribed, leading to a continuous piecewise quadratic doublet strength over the entire configuration. The source strength is not required to be continuous and a linear least-square type of distribution is used. This method allows a wide variety of
36、 singularity and boundary-condition formulations to be specified including both chord plane and surface panel model- ing options. For “thick“ body type of solutions a common approach is to use combined source and doublet panels with a perturbation potential boundary condi- tion corresponding to zero
37、 normal mass flux through the surface. For “thin“ wings, thickness effects can be considered to be secondary and doublets alone can be used on the lifting surfaces to satisfy zero normal mass flux through the surface. A combination of these two approaches was used in the present applica- tion: combi
38、ned source and doublet panels were used on the body and doublet panels were used on the wings and vertical tails. In the context of small- 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-perturbation solution, a distinction between zero normal mass
39、 flux and zero normal velocity flux can be made. Based on previous experience (ref. 6) , the formulation corresponding to zero normal velocity flux was used on the body and that corresponding to zero normal mass flux was used on the wing and tail sur- faces. Also within the context of the small-pert
40、urbation solution, compressibil- ity axes can be defined which are not aligned with free-stream direction. In all of the results presented, however, compressibility axes were aligned with the f ree-stream direction. The forces and moments are calculated by integrating the pressure coeffi- cient over
41、 the surface. The isentropic pressure relation was used to calculate the pressure coefficient from the total velocity. DISCUSSION The experimental longitudinal characteristics are presented in table 11. The experimental lateral-directional stability data are presented in figures 2 to 5. The comparis
42、ons of experimental and theoretical lateral-directional sta- bility parameters at cx = Oo are presented in figure 6 for the various configu- rations. As was previously noted, MISLIFT is a planar solution and, therefore, only configurations which have lifting surfaces in a single plane could be analy
43、zed by this method: however, both the APAS and PAN AIR codes could be used to analyze complete configurations. The B, BW, and BWVl configurations are shown in figures 6(a) to 6(c). The theoretical estimates of Cy and Cl show B B fair to good agreement with the experimental results, except for of the
44、 16 BWVl configuration (fig. 6 (c) at M = 2.86, which probably results from the scatter of experimental data as shown in figure 2(c). Fair agreement is shown for (2% for either MISLIFT or PAN AIR. Both MISLIFT and PAN AIR show a vari- ation of all parameters with Mach number, whereas APAS, which is
45、based on a slender-body theory for the body effects, does not show this variation with Mach number for the B or BW configuration but does show a variation with Mach number for the BWV1 configuration. The APAS results for the BW configuration are the same as the results for the B configuration, and i
46、ndeed the experimental data show little effect of addition of the wing at = Oo. The body-vertical- tail configurations are shown in figures 6(d) to 6 (9). Fair to good canparisons are made for and C for all configurations with only slight differences among the three theories. Fair to good agreement
47、in C is shown for MISLIFT “ 16 “B and PAN AIR, except for BV3 (fig. 6(f). This exception is probably due to the increase in manent arm which tends to magnify any error in the side force or center-of-pressure location. Canparisons for the wing-mounted vertical-tail configurations are shown in figures
48、 6(h) to 6(k). Fair to good agreement is shown for all parameters for PAN AIR. Since the results from MISLIFT and APAS are invariant with angle of attack, only the results from PAN AIR are used for comparisons of the stability param- eters at angles of attack. These comparisons are shown in figure 7
49、 for the vari- 6 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I I ous configurations. It should be noted that the theory is violated at M = 2.86 for 01 and Sawyer, Wallace C.: A Method for Calculating the Aerodynamic Loading on Wing-Body Combinations at Small Angles of Attack in Supersonic Flow. NAS
