1、NASA TECHNICAL NOTE NASA # I I- TN - kOAN COPY: RETURN TO KIRTLANO AFB, N MEX AFWL (WLILQ) EXPERIMENTAL PRESSURE DISTRIBUTIONS ON A 120 CONE AT MACH NUMBERS FROM 2.96 TO 4.63 AND ANGLES OF ATTACK FROM 0“ TO 20 by Robert L. StulZings, Jr., und Dorothy H. Tador Langley Research Center LangZey Stution,
2、 Humpton, Vd. NATIONAL AERONAUTICS AND SPACE ADMINISTRATION WASHINGTON, D. C. MARCH 1969 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-TECH LIBRARY KAFB, NM I Illill Ill11 11111 lllll lllll lllll11111 Ill1 Ill1 0131821 NASA TN D-5054 EXPERIMENTAL P
3、RESSURE DISTRIBUTIONS ON A 120 CONE AT MACH NUMBERS FROM 2.96 TO 4.63 AND ANGLES OF ATTACK FROM 0 TO 20 By Robert L. Stallings, Jr., and Dorothy H. Tudor Langley Research Center Langley Station, Hampton, Va. NATIONAL AERONAUT ICs AND SPACE ADMINISTRATION For sale by the Clearinghouse for Federal Sci
4、entific and Technical Information Springfield, Virginia 22151 - CFSTl price $3.00 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-EXPERIMENTAL PRESSURE DISTRIBUTIONS ON A 120 CONE AT MACH NUMBERS FROM 2.96 TO 4.63 AND ANGLES OF ATTACK FROM Oo TO 20 B
5、y Robert L. Stallings, Jr., and Dorothy H. Tudor Langley Research Center SUMMARY Pressure distributions have been experimentally determined on both a sharp and a blunt 1200 cone configuration. The blunt-cone configuration consisted of a spherical seg- ment which had a radius of 1/8 the base diameter
6、 and which faired into a 120 cone frus- tum. angles of attack from Oo to 20. The tests were conducted at Mach numbers of 2.96, 3.95, and 4.63 over a range of Pressure distributions and shock shapes obtained on both the sharp and blunt cones Mach number had no effect at an angle of attack (a) of Oo w
7、ere in good agreement with an approximate theoretical solution based on the one-strip method of integral relations. on pressure distributions expressed in the form of local measured pressures divided by the free-stream pitot pressure (pz/pt,2) for either the sharp or blunt configurations at a!= 00.
8、At a! Oo, the Mach number effect was confined to the leeward side of both con- figurations and consisted of a decrease in p p with increasing Mach number. The measured pressure distributions on the cone frustum of the blunt configuration were essentially the same as those obtained within the same re
9、gion on the sharp-cone configu- ration throughout the range of test variables of this investigation. z/ t,2 An approximate technique involving a tangent-cone concept is presented for pre- dicting the windward and leeward pressures for a! Oo. in fair agreement with experimental results both in the fo
10、rm of pressure distributions and force coefficients. For a! Oo, the local pressures around the circumference of both the sharp- and blunt-cone configurations, nondimensionalized by the pressure mea- sured along the windward ray, were in good agreement with an empirically derived second-order polynom
11、ial. Force and moment coefficients obtained from integrated pressure measurements along the windward and leeward meridians together with this empirical equation for the circumferential distributions were in excellent agreement with balance measurements. Results from this method are Provided by IHSNo
12、t for ResaleNo reproduction or networking permitted without license from IHS-,-,-INTRODUCTION The use of unmanned probes for exploring low-density planetary atmospheres, such as that of Mars, has recently stimulated interest in the aerodynamics of vehicles with a low ballistic coefficient. A configu
13、ration satisfying this requirement and one that is being considered for such missions is a 120 cone. Since the bow shock for a 120 cone is detached for all Mach numbers, the governing partial differential equations for the fIow field are of the elliptic type, and no exact analytical solutions are ye
14、t available. Several experimental investigations have been conducted to determine the aerodynamic charac- teristics of such a cone (see refs. 1 to 4); however, very little experimental pressure data exist that enable determination in detail of the local flow properties. Such flow properties are requ
15、ired in order that the designer might determine local aerodynamic heating rates and local structural loading. The purpose of this investigation was to experimentally determine detailed pres- sure distributions over a 120 cone through a range of angles of attack from Oo to 20 and Mach numbers from 2.
16、96 to 4.63. The model had interchangeable nose tips and the ratios of nose radius to afterbody radius were 0 and 0.25. SYMBOLS CA Cm CN CP 2 Forebody axial force f orebody axial-force coefficient, pitching- moment coefficient, Pitching moment 2 a%orb 2rqcorb 3 Normal force 2 normal-f orce coefficien
17、t, %orb Pz - pco q, pressure coefficient, g, - p, q, base pressure coefficient, base diameter axial length of cone local Mach number bat outer edge of boundary layer free-stream Mach number Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-I “1 1111 r
18、b rn S S (% .p. X (Y rl e base pressure local static pressure along leeward meridian (6 = OO; local static pressure free - stream stagnation pressure stagnation pressure behind normal shock local static pressure along windward meridian (e = 180; = Oo) free-stream static pressure = Oo) free-stream dy
19、namic pressure radial distance from axis of symmetry base radius nose radius surface length (see fig. 1) surface length from most forward station on model to shoulder corner nondimensionalized stagnation-point location axial distance from most forward station on model (see fig. 1) angle of attack no
20、ndimensionalized angle of attack, a goo - cTc meridian angle (see fig. 1) final circumferential integration limit initial circumferential integration limit 3 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-cone semiapex angle equivalent cone semiapex
21、 angle roll angle (see fig. 1) APPARATUS AND TEST CONDITIONS Wind Tunnel This investigation was conducted in the high Mach number test section of the Langley Unitary Plan wind tunnel described in reference 5. continuous-flow tunnel has an asymmetric sliding-block nozzle that permits a continuous var
22、iation in the test-section Mach number from 2.30 to 4.63. This variable-pressure, Models and Instrumentation The 120 cone model was constructed of aluminum and had interchangeable nose tips, as illustrated in figure 1, such that pressure distributions could be obtained for both a blunt cone and a sh
23、arp cone with minimum time required for model change. The base diameter of the model was 8.0 inches (203.2 mm) and the nose radius of the blunt cone was 1.0 inch (25.4 mm). The sting used had a diameter of 1.50 inches (38.1 mm) and was 31.0 inches (787.4 mm) in length. The cone frustum was instrumen
24、ted with 58 pressure orifices of 0.050-inch (1.27-mm) internal diameter, located as shown in figure 1. was a spherical segment, was instrumented with five orifices, orifice 1 being located at the axis of symmetry. The sharp nose tip was instrumented with only four orifices since one was not located
25、at the stagnation point in order to retain a sharp apex. The pressure tubing from all orifices was routed through a slot in the sting assembly to minimize sting effects on the base-pressure measurements. A typical model installation in the test sec- tion is shown in figure 2. The blunt nose tip, whi
26、ch Pressures were recorded by using three 48-channel pressure-sampling valves. Each valve sequentially transmits each channel of pressure information to a single elec- trical pressure transducer. This electrical information is fed to a strip-chart recorder and an analog-to-digital converter. The out
27、put in digital form is then recorded on punch- cards suitable for machine computation of final data. was measured with a precision mercury manometer. The tunnel stagnation pressure 4 Provided by IHSNot for ResaleNo reproduction or networking permitted without license from IHS-,-,-Accuracy Accuracy o
28、f the pressure-sampling valves is better than 1 percent of the full-scale range of the gage; this includes all errors of linearity, hysteresis, and repeatability. Gages with a maximum range of 7.5 lb/in2 (5.17 N/cm2) and 5.0 lb/in2 (3.45 N/cm2) were used for orifices on the model nose and base, resp
29、ectively. precision mercury manometer with which the stagnation pressure was measured is 0.0035 lb/in2 (0.0024 N/cm2). The accuracy of the The results of a test-section calibration indicate that the variation in a free-stream Mach number of 2.96 was rt0.02, of 3.95 was ,t0.06, and of 4.63 was ,t0.05
30、. I Test Conditions This investigation was conducted at Mach numbers of 2.96, 3.95, and 4.63 for a nominal Reynolds number of 2 x 106 based on model base diameter. Angle of attack was varied from Oo to 20 with an accuracy of *0.lo relative to the tunnel center line. tunnel stagnation temperature was
31、 held constant at 610 R (339O K) for M, E 2.96 and at 635O R (353O K) for M, = 3.95 and 4.63. The RESULTS AND DISCUSSION A complete tabulation of the experimental data is presented in tables I to VI. Local flow properties included in this tabulation are pressures and Mach numbers. Pressures Experime
32、ntal forebody pressures.- Pressure distributions obtained through the ranges of angles of attack and Mach numbers are presented in figure 3 for the blunt cone at = Oo. The local measured pressures have been nondimensionalized by the free- stream pitot pressure and are plotted as a function of the ra
33、tio of local surface length to base diameter. For a! = Oo, the experimental data are compared with an approximate theoretical solution obtained by using the one-strip method of integral relations as described in reference 6. For all Mach numbers tested the experimental data as shown in figure 3 at a
34、! = Oo are in good agreement with the theoretical values. in the vicinity of than the experimental value. pressures on the windward side of the model and a decrease on the leeward side, as would be expected. The stagnation point, as indicated by maximum pressure measure- The maximum disagreement occ
35、urs = 0.45 where the theoretical value is approximately 3 percent less d Increasing the angle of attack results in an increase in the ments, is located on the cone frustum for a! P 10 throughout the test range 5 Provided by IHSNot for ResaleNo reproduction or networking permitted without license fro
36、m IHS-,-,-of Mach numbers. For all angles of attack greater than Oo, the overexpanding flow on the leeward side of the spherical segment part of the nose results in pressures below those obtained on the forward part of the cone frustum. In this region an adverse pres- sure gradient, the magnitude of
37、 which increases with increasing angle of attack through- out the range of this investigation, occurs. The results shown in figure 3 are replotted in figure 4 to more clearly illustrate the effect of Mach number. For a! = Oo (fig. 4(a) the variation in pressure with Mach number is negligible. This r
38、esult should be expected since the sonic point is fixed at the sharp corner of the cone frustum for all test Mach numbers. For a!= loo (fig. 4(b) and a! = 20 (fig. 4(c), the Mach number effect on the pressures is also negligible wind- ward of the measured stagnation point. Leeward of the stagnation
39、point a Mach number effect does occur which generally consists of a decrease in pressure with increasing Mach number. The magnitude of this effect increases with increasing a. Pressure distributions obtained for the blunt cone are compared with those obtained for the sharp cone in figure 5 in order
40、to assess the extent of bluntness effects, the blunt- cone data being the same data shown previously in figure 4. For a! = Oo, the pressure distributions on the two configurations are essentially the same although the pressures obtained for the sharp cone appear to be slightly greater than those for
41、 the blunt cone at the larger values of s/d. It should be noted that the values of s for orifices located on the cone frustum of the sharp cone are slightly greater than those for the same orifices on the blunt cone inasmuch as the surface length of the sharp cone tip is slightly greater (0.0053 inc
42、h (1.35 mm) than that of the blunt cone tip. If a common coordinate system for the orifice locations on the cone frustum of the two configurations had been used, the measured pressures would have been even closer than indicated in figure 5. Pressure distributions obtained on the windward side of the
43、 two configurations for a! 00 indicate the same trends as shown for a! = 0. On the leeward side of the model, pressures associated with the flow expanding around the spherical nose segment of the blunt cone are somewhat lower than those obtained for the sharp cone. The pressures obtained downstream
44、of the adverse pressure gradient occurring on the leeward side of the blunt cone are of approximately the same magnitude as those obtained for the sharp cone. Analytical forebody pressures. - The pressure distributions for the sharp cone at angles of attack were approximated by a method similar to t
45、hat suggested in reference 6 for blunt cones. The sharp cone was selected for this comparison since measured and empirical pressures will later be used for computing force coefficients and most of the large-angle-cone force measurements are for sharp cones. The method of reference 6 simply assumes t
46、hat the windward pressures correspond to those for a tangent cone of angle gc,e = cc + a! angle cc,e = uc - a. The one-strip method of integral relations is used to determine and that the leeward pressures correspond to a tangent cone of 6 Provided by IHSNot for ResaleNo reproduction or networking p
47、ermitted without license from IHS-,-,-the pressure distributions for the various cone angles (T,. It was shown in refer- ence 6 that better agreement with experimental data was obtained by shifting the stag- nation point to the most forward point of the nose and forcing the sonic points to occur at
48、the cone shoulder by a linear transformation. For a sharp cone this method indi- cates that the stagnation point would remain at the apex for angles of attack less than 90 - (T but, as shown in figure 5, this does not occur. To more accurately define the stagnation-point locations, an attempt was ma
49、de to correlate these locations for large-angle cones with other published data (ref. 4). One might expect the stagnation-point location for a cone at such an angle of attack that the windward surface is normal to the flow to be relatively insensitive to cone angle. Also, since at an angle of attack of 0 the stagnation point is
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