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At least 577 records · Page 32
All sourcesNASA NTRSDOE OSTI
NASA NTRS1969

Thermal convection at infinite Prandtl number

Numerical solutions are developed for the full scaled three dimensional thermal convection problem, at infinite Prandtl number, and for rigid boundaries. The procedure is designed to give an approximate account of horizontally homogeneous and isotropic flow situations. Results, which included a maximum of 36 horizontal wave number vectors and 20 vertical wave numbers, appear to adequately describe the flow fill up to R = 100,000; beyond this R the results appear to show horizontal wave number truncation error. This error seems to affect the boundary slope of the mean temperature field more than other mean quantities. Despite some numerical uncertainities, certain of the qualitative features of the flow fill are predicted with reasonable confidence.

Herring, J. R.↗
NASA NTRS1972

Effect of Reynolds number on overall performance of a 3.7-inch-diameter six-stage axial-flow compressor

A 9.4-centimeter (3.7-in.) diameter six-stage axial-flow compressor was tested in argon over a range of inlet pressures corresponding to a Reynolds number range of 30,600 to 160,000. The effect of Reynolds number on efficiency, pressure ratio, work input, maximum flow, and surge is shown. The Reynolds number effects are discussed in terms of changes in boundary-layer thickness, losses, and the resulting changes in throughflow velocity. Significant deviation was noted from the 0.2 power relation often used to express the variation of loss with Reynolds number.

Heidelberg, L. J.↗
NASA NTRS1972

Normal impingement loads due to small air jets issuing from a base plate and reflecting off a platform for various jet Mach numbers, separation distances, and ambient pressures

An investigation was conducted in a 12.5-meter-diameter vacuum sphere to determine the impingement loads due to air jets issuing from and perpendicular to a circular base and reflecting off a square platform, that is, a simulation of rendezvous maneuvering, docking, launch, impact dampers etc. The nozzles had exit Mach numbers of 1, 3, 5, and 7. The ambient pressures were 0.0006, 5, 225, and 760 torr. Under near-field separation distances and at 0.0006 torr, reflections were significant; and ratios of the impingement force to thrust on both plates in the biplane arrangement varied from about 750 for exit Mach number 1 to 120 for exit Mach number 7. The far-field force ratios were near unity for the platform and zero for the base and indicated few, if any, reflections. Some reversals and rapid changes in loads were obtained at transition distances between the near and far fields. In general, increasing the exit Mach number or ambient pressure reduced the impingement loads.

Hoffman, S.↗
NASA NTRS1972

Discrete sonic jets used as boundary-layer trips at Mach numbers of 6 and 8.5

The effect of discrete three-dimensional sonic jets used to promote transition on a sharp-leading-edge flat plate at Mach numbers of 6 and 8.5 and unit Reynolds numbers as high as 2.5 x 100,000 per cm in the Langley 20-inch hypersonic tunnels is discussed. An examination of the downstream flow-field distortions associated with the discrete jets for the Mach 8.5 flow was also conducted. Jet trips are found to produce lengths of turbulent flow comparable to those obtained for spherical-roughness-element trips while significantly reducing the downstream flow distortions. A Reynolds number based upon secondary jet penetration into a supersonic main flow is used to correlate jet-trip effectiveness just as a Reynolds number based upon roughness height is used to correlate spherical-trip effectiveness. Measured heat-transfer data are in agreement with the predictions.

Stone, D. R.↗
NASA NTRS1972

The high Reynolds number cryogenic wind tunnel.

Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in test Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies have been made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature with due regard to avoiding liquefaction and adverse real-gas effects. Practical operational procedures have been developed in a low-speed prototype cryogenic wind tunnel. Aerodynamic experiments in the facility have demonstrated the theoretically predicted variations in Reynolds number and drive power. Force and moment measurements on a wing model mounted on a water-jacketed strain-gage sting balance have demonstrated the feasibility of operation of such balances in a cryogenic environment.

Goodyer, M. J.↗
NASA NTRS1972

Space shuttle: Aerodynamic stability, control effectiveness and drag characteristics of a shuttle orbiter configuration at Mach numbers from 0.6 to 4.96

Experimental aerodynamic investigations were conducted in the NASA/MSFC 14-inch Trisonic Wind Tunnel from Sept. 27 to Oct. 7, 1972 on a 0.004 scale model of the NR ATP baseline shuttle orbiter configuration. Six component aerodynamic force and moment data were recorded at 0 deg sideslip angle over an angle of attack range from 0 to 20 deg for Mach numbers of 0.6 to 4.96, 20 to 40 deg for Mach numbers of 0.6, 0.9, 2.99, and 4.96, and 40 to 60 deg for Mach numbers of 2.99 and 4.96. Data were obtained over a sideslip range of -10 to 10 deg at 0, 10, and 20 deg angles of attack over the Mach range and 30 and 50 deg at Mach numbers of 2.99 and 4.96. The purpose of the test was to define the buildup, performance, stability, and control characteristics of the orbiter configuration. The model parameters, were: body alone; body-wing; body-wing-tail; elevon deflections of 0, 10, -20, and -40 deg both full and split); aileron deflections of plus or minus 10 deg (full and split); rudder flares of 10 and 40 deg, and a rudder deflection of 15 deg about the 10 and 40 deg flare positions.

Ramsey, P. E.↗
NASA NTRS1973

Effect of fineness ratio on boattail drag of circular-arc afterbodies having closure ratios of 0.50 with jet exhaust at Mach numbers up to 1.30

An investigation was conducted to determine the effect of fineness ratio on the drag of circular-arc boattails at subsonic and low supersonic speeds. The boattails had closure ratios of 0.50 and incorporated convergent nozzles. The investigation was conducted statically and at Mach numbers from 0.40 to 1.30 at 0 deg angle of attack with jet total-pressure ratios varying from jet off to about 6, depending on Mach number. Low-fineness-ratio boattails had large separated-flow regions and the highest drag at all Mach numbers. Subsonic pressure-plus-friction drag levels were generally similar for boattails which did not have large separated regions. Drag-rise Mach number increased as boattail fineness ratio increased.

Reubush, D. E.↗
NASA NTRS1973

The correlation structure of several popular pseudorandom number generators

One of the desirable properties of a pseudorandom number generator is that the sequence of numbers it generates should have very low autocorrelation for all shifts except for zero shift and those that are multiples of its cycle length. Due to the simple methods of constructing random numbers, the ideal is often not quite fulfilled. A simple method of examining any random generator for previously unsuspected regularities is discussed. Once they are discovered it is often easy to derive the mathematical relationships, which describe the mathematical relationships, which describe the regular behavior. As examples, it is shown that high correlation exists in mixed and multiplicative congruential random number generators and prime moduli Lehmer generators for shifts a fraction of their cycle lengths.

Neuman, F.↗
NASA NTRS1973

An investigation of a free jet at low Reynolds numbers

A comparison of the measured and calculated flow field properties of a nozzle of unusual design which was used to produce an incompressible, low Reynolds number jet is presented. The nozzle is essentially a porous metal plate which covers the end of a pipe. Results are presented for nozzle Reynolds numbers from 50 to 1000 with velocities of 100 or 200 ft/sec. The nozzle produces a uniform velocity profile at nozzle Reynolds numbers well below those at which conventional contoured nozzles are completely filled with the boundary layer. A jet mixing analysis based on the boundary layer equations accurately predicted the flow field over the entire range of Reynolds numbers tested.

Greene, G. C.↗
NASA NTRS1973

An experimental investigation of three oblique-wing and body combinations at Mach numbers between 0.60 and 1.40

An experimental investigation was conducted in an 11- by 11-foot transonic wind tunnel to determine the aerodynamic characteristics of three oblique high aspect ratio wings in combination with a high fineness-ratio Sears-Haack body. The three wings had the same elliptical planform and base line curvature but had different airfoil sections. One wing had an airfoil section designed to have a lift coefficient of 1.0 at a Mach number of 0.7, another to have shock-free supersonic flow over the upper surface, and the other to have a lift coefficient of 1.3 at a Mach number of 0.6. Longitudinal and lateral-directional stability data were obtained at wing yaw angles of 0 deg, 45 deg, 50 deg, and 60 deg over a test Mach number range from 0.6 to 1.4 for angles of attack between minus 7 deg and 9 deg. Reynolds numbers for the study were 4 and 6 million per foot. Flow-visualization studies were made to examine the nature of the flow on the wing surfaces. Notable differences were found in the aerodynamic characteristics of the three wing-body combinations, particularly in the lateral-directional characteristics.

Graham, L. A.↗
NASA NTRS1973

A study of methods which predict supersonic flow fields from body geometry, distance, and Mach number

A study of seven methods for predicting flow-field pressure signatures from the parameters Mach number, body geometry, and field-path distance has been made. The methods included the method of characteristics, which served as a standard of comparison; a shock-capturing method; three Whitham theory methods; a modified characteristics method; and a bicharacteristics method. Results from each method were also compared with recently obtained wind-tunnel data for a cone-cylinder model at Mach numbers of 2.96 and 4.63 with ratios of radial distance to cone length of 2 and 5. The comparisons at a Mach number of 2.96 showed that signatures from all the methods correlated well with wind-tunnel data and with the signatures predicted by the method of characteristics. At a Mach number of 4.63, however, the agreement between the signatures obtained in the wind tunnel and those predicted by theory varied from good to poor, as did the agreement between the signatures obtained by the method of characteristics and the other six methods. It should be noted that these results and comparisons indicate pressure prediction capabilities only for the near-field flow about bodies of revolution.

Mack, R. J.↗
NASA NTRS1973

Effect of Subsonic Inlet Lip Geometry on Predicted Surface and Flow Mach Number Distributions

The effect of subsonic inlet lip geometry on predicted surface and flow Mach number distributions is illustrated. The theoretical results were obtained from incompressible potential flow calculations corrected for compressibility. The major emphasis of this investigation is on the low-speed (takeoff and landing) operating conditions. The low-speed results were obtained for a range of three geometric variables of interest: contraction ratio, defined as the ratio of highlight area to throat area; internal lip major - to minor-axis ratio; and internal lip shape. The low-speed results were obtained at both static conditions and a free-stream velocity of 42.6m/sec, with incidence angles ranging from 0 deg to 50 deg. The results indicate that of the three geometric variables considered, contraction ratio had the largest effect on the surface Mach number distributions. The effects of inlet diameter ratio and blunting of the external forebody on maximum external surface Mach numbers are illustrated at a cruise Mach number of 0.8.

Albers, J. A.↗
NASA NTRS1974

Aerodynamic characteristics of three slender sharp-edge 74 degrees swept wings at subsonic, transonic, and supersonic Mach numbers

Slender sharp-edge wings having leading-edge sweep angles of 74 deg have been studied at Mach numbers from 0.60 to 2.80, at angles of attack from about minus 4 deg to 22 deg, and at angles of sideslip from 0 deg to 5 deg. The wings had delta, arrow, and diamond planforms. The experimental tests were made in the Langley 8-foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel test section number 1. The theoretical predictions were made using the theories of NASA TN D-3767 and NASA TN D-6243. The results of the study indicated that the lift and drag characteristics as affected by planform and Mach number could be reasonably well predicted for the delta wing in the subsonic and transonic Mach number range. In the supersonic range, the delta and diamond wings were about equally good in the degree of agreement between experiment and theory. In making drag-due-to-lift predictions the vortex lift effects must be taken into account if reasonable results are to be obtained at moderate or high lift coefficients.

Davenport, E. E.↗
NASA NTRS1974

Stability and control characteristics at Mach numbers from 0.20 to 4.63 of a cruciform air-to-air missile with triangular canard controls and a trapezoidal wing

Investigations have been conducted in the Langley 8-foot transonic pressure tunnel and the Langley Unitary Plan wind tunnel at Mach numbers from 0.20 to 4.63 to determine the stability and control characteristics of a cruciform air-to-air missile with triangular canard controls and a trapezoidal wing. The results indicate that canards are effective in producing pitching moment throughout most of the test angle-of-attack and Mach number range and that the variations of pitching moment with lift for trim conditions are relatively linear. There is a decrease in canard effectiveness with an increase in angle of attack up to about Mach 2.50 as evidenced by the beginning of coalescence of the pitching-moment curves. At a Mach number above 2.50, there is an increase in effectiveness at moderate to high angles of attack. Simulated launch straps have little effect on the lift and pitch characteristics but do cause an increase in drag, and this increase in drag induces a rolling moment at a zero roll attitude where the straps cause an asymmetric geometric shape. The canards are not suitable devices for roll control and, at some Mach numbers and roll attitudes, are not effective in producing pure yawing moments.

Graves, E. B.↗
NASA NTRS1974

Fluctuating surface pressure characteristics on slender cones in subsonic, supersonic, and hypersonic Mach-number flow

A program is presented to determine the flow-induced unsteady aerodynamic loads on space-shuttle orbiter configurations in the early reentry phase. Experiments on 8 degree half-angle cone/flat-base configurations were conducted at free-stream Mach numbers of 6 and 15 and Reynolds numbers corresponding to shuttle reentry. Data are available at free-stream Mach-numbers of 0.67, 2.5, 3, 4, 6, 15, and 22. It is shown that empirical expressions could be developed for the Mach-number dependence of the overall fluctuation pressure under an attached turbulent boundary layer and under a separated base-flow region.

Heller, H. H.↗
NASA NTRS1974

Effects of leading edge sweep angle and design lift coefficient on performance of a modified arrow wing at a design Mach number of 2.6

Wing models were tested in the high-speed section of the Langley Unitary Plan wind tunnel to study the effects of the leading-edge sweep angle and the design lift coefficient on aerodynamic performance and efficiency. The models had leading-edge sweep angles of 69.44 deg, 72.65 deg, and 75.96 deg which correspond to values of the design Mach-number-sweep-angle parameter (beta cotangent A) sub DES of 0.6, 0.75, and 0.9, respectively. For each sweep angle, camber surfaces having design lift coefficients of 0,0.08, and 0.12 at a design Mach number of 2.6 were generated. The wind-tunnel tests were conducted at Mach numbers of 2.3, 2.6, and 2.96 with a stagnation temperature of 338.7 K (150 F) and a Reynolds number per meter of 9.843 times 10 to the 6th power. The results of the tests showed that only a moderate sweeping of the wing leading edge aft of the Mach line along with a small-to-moderate amount of camber and twist was needed to significantly improve the zero-lift (flat camber surface) wing performance and efficiency.

Mack, R. J.↗
NASA NTRS1974

The cryogenic wind tunnel concept for high Reynolds number testing

Theoretical considerations indicate that cooling the wind-tunnel test gas to cryogenic temperatures will provide a large increase in Reynolds number with no increase in dynamic pressure while reducing the tunnel drive-power requirements. Studies were made to determine the expected variations of Reynolds number and other parameters over wide ranges of Mach number, pressure, and temperature, with due regard to avoiding liquefaction. Practical operational procedures were developed in a low-speed cryogenic tunnel. Aerodynamic experiments in the facility demonstrated the theoretically predicted variations in Reynolds number and drive power. The continuous-flow-fan-driven tunnel is shown to be particularly well suited to take full advantage of operating at cryogenic temperatures.

Kilgore, R. A.↗
NASA NTRS1975

Experimental aerodynamics characteristics for bodies of elliptic cross section at angles of attack from 0 deg to 58 deg and Mach numbers from 0.6 to 2.0

An experimental investigation was conducted to measure the static aerodynamic characteristics for two bodies of elliptic cross section and for their equivalent body of revolution. The equivalent body of revolution had the same length and axial distribution of cross-sectional area as the elliptic bodies. It consisted of a tangent ogive nose of fineness ratio 3 followed by a cylinder with a fineness ratio of 7. All bodies were tested at Mach numbers of 0.6, 0.9, 1.2, 1.5, and 2.0 at angles of attack from 0 deg to 58 deg. The data demonstrate that the aerodynamic characteristics can be significantly altered by changing the body cross section from circular to elliptic and by rolling the body from 0 deg to 90 deg. For example, the first elliptic body (with a constant cross-sectional axis ratio of 2) developed at zero roll about twice the normal force developed by the equivalent body of revolution. At some angles of attack greater than about 25 deg, side forces and yawing moments were measured in spite of the fact that the bodies were tested at zero angle of sideslip. The side-force and yawing-moment coefficients decreased with an increase in Mach number and essentially disappeared for all the bodies at Mach numbers greater than 1.2. From the standpoint of reducing undesirable side forces at high angles of attack, it is best to have the flattest side of the nose of the elliptic bodies pitching against the stream crossflow. The effect of Reynolds number was also the least significant for both elliptic bodies when the flattest side of the nose was pitched against the stream crossflow.

Jorgensen, L. H.↗