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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

Exploratory wind tunnel tests of a shock-swallowing air data sensor at a Mach number of approximately 1.83

The test probe was designed to measure free-stream Mach number and could be incorporated into a conventional airspeed nose boom installation. Tests were conducted in the Langley 4-by 4-foot supersonic pressure tunnel with an approximate angle of attack test range of -5 deg to 15 deg and an approximate angle of sideslip test range of + or - 4 deg. The probe incorporated a variable exit area which permitted internal flow. The internal flow caused the bow shock to be swallowed. Mach number was determined with a small axially movable internal total pressure tube and a series of fixed internal static pressure orifices. Mach number error was at a minimum when the total pressure tube was close to the probe tip. For four of the five tips tested, the Mach number error derived by averaging two static pressures measured at horizontally opposed positions near the probe entrance were least sensitive to angle of attack changes. The same orifices were also used to derive parameters that gave indications of flow direction.

Nugent, J.↗
NASA NTRS1975

The effect of Reynolds number on the boattail drag of two wing-body configurations

An investigation has been conducted in the Langley 1/3-meter transonic cryogenic tunnel to determine the effects of varying Reynolds number on the boattail drag of wing-body configurations at subsonic speeds. Two boattailed cone-cylinder nacelle models were tested with a 60 deg delta wing at an angle of attack of 0 deg. Reynolds number, based on model length, was varied from about 2.5 million to 67 million. Even though the presence of the wing had large effects on the boattail pressure coefficients, the results of this investigation were similar to those previously found for a series of isolated boattails. Boattail pressure coefficients in the expansion region became more negative with increasing Reynolds number, while those in the recompression region became more positive. These two effects were compensating, and as a result, there was virtually no effect of Reynolds number on boattail pressure drag.

Reubush, D. E.↗
NASA NTRS1975

Application of a transitional boundary-layer theory in the low hypersonic Mach number regime

An investigation is made to assess the capability of a finite-difference boundary-layer procedure to predict the mean profile development across a transition from laminar to turbulent flow in the low hypersonic Mach-number regime. The boundary-layer procedure uses an integral form of the turbulence kinetic-energy equation to govern the development of the Reynolds apparent shear stress. The present investigation shows the ability of this procedure to predict Stanton number, velocity profiles, and density profiles through the transition region and, in addition, to predict the effect of wall cooling and Mach number on transition Reynolds number. The contribution of the pressure-dilatation term to the energy balance is examined and it is suggested that transition can be initiated by the direct absorption of acoustic energy even if only a small amount (1 per cent) of the incident acoustic energy is absorbed.

Shamroth, S. J.↗
NASA NTRS1975

Aerodynamic characteristics of a hypersonic research airplane concept having a 70 degree swept double delta wing at Mach numbers from 1.50 to 2.86

An experimental investigation of the static longitudinal, lateral, and directional stability characteristics of a hypersonic research airplane concept having a 70 deg swept double-delta wing was conducted in the Langley unitary plan wind tunnel. The configuration variables included wing planform, tip fins, center fin, and scramjet engine modules. The investigation was conducted at Mach numbers from 1.50 to 2.86 and at a constant Reynolds number, based on fuselage length, of 3,330,000. Tests were conducted through an angle-of-attack range from about -4 deg to 24 deg with angles of sideslip of 0 deg and 3 deg and at elevon deflections of 0, -10, and -20 deg. The complete configuration was trimmable up to angles of attack of about 22 deg with the exception of regions at low angles of attack where positive elevon deflections should provide trim capability. The angle-of-attack range for which static longitudinal stability also exists was reduced at the higher Mach numbers due to the tendency of the complete configuration to pitch up at the higher angles of attack. The complete configuration was statically stable directionally up to trimmed angles of attack of at least 20 deg for all Mach numbers M with the exception of a region near 4 deg at M = 2.86 and exhibited positive effective dihedral at all positive trimmed angles of attack.

Penland, J. A.↗
NASA NTRS1975

A tabulation of pipe length to diameter ratios as a function of Mach number and pressure ratios for compressible flow

Computer programs and resulting tabulations are presented of pipeline length-to-diameter ratios as a function of Mach number and pressure ratios for compressible flow. The tabulations are applicable to air, nitrogen, oxygen, and hydrogen for compressible isothermal flow with friction and compressible adiabatic flow with friction. Also included are equations for the determination of weight flow. The tabulations presented cover a wider range of Mach numbers for choked, adiabatic flow than available from commonly used engineering literature. Additional information presented, but which is not available from this literature, is unchoked, adiabatic flow over a wide range of Mach numbers, and choked and unchoked, isothermal flow for a wide range of Mach numbers.

Dixon, G. V.↗
NASA NTRS1976

Preliminary scaling laws for plasma current, ion kinetic temperature, and plasma number density in the NASA Lewis bumpy torus plasma

Parametric variation of independent variables which may affect the characteristics of bumpy torus plasma have identified those which have a significant effect on the plasma current, ion kinetic temperature, and plasma number density, and those which do not. Empirical power law correlations of the plasma current, and the ion kinetic temperature and number density were obtained as functions of potential applied to the midplane electrode rings, the background neutral gas pressure, and the magnetic field strength. Additional parameters studied included the type of gas, the polarity of the midplane electrode rings, the mode of plasma operation, and the method of measuring the plasma number density. No significant departures from the scaling laws appear to occur at the highest ion kinetic temperatures or number densities obtained to date.

Roth, J. R.↗
NASA NTRS1976

Experimental investigation to validate use of cryogenic temperatures to achieve high Reynolds numbers in boattail pressure testing

An investigation has been conducted in the Langley 1/3 meter transonic cryogenic tunnel to validate the use of cryogenic temperatures to achieve high Reynolds numbers in nozzle boattail pressure testing. Tests were conducted at 0 degree angle of attack and at Mach numbers of 0.60, 0.85, and 0.90 on two wing-body configurations with differing boattail geometries. Test data were obtained by using two different techniques, the cryogenic method and the conventional method, to obtain the same Reynolds number. Later, the test data obtained from the two techniques on boattail pressure coefficient distributions and pressure drag coefficients were compared; results from the comparisons show excellent repeatability for all test conditions and indicate no measurable errors when using cryogenic temperatures to achieve high Reynolds numbers for nozzle boattail pressure testing.

Reubush, D. E.↗
NASA NTRS1977

Transonic high Reynolds number stability and control characteristics of a 0.015-scale remotely controlled elevon model (44-0) of the space shuttle orbiter tested in calspan 8-foot TWT (LA70)

Transonic Wind Tunnel tests were run on a .015 scale model of the space shuttle orbiter vehicle in the 8-foot transonic wind tunnel. Purpose of the test program was to obtain basic shuttle aerodynamic data through a full range of elevon and aileron deflections, verification of data obtained at other facilities, and effects of Reynolds number. Tests were performed at Mach numbers from .35 to 1.20 and Reynolds numbers from 3,500,000 to 8,200,000 per foot. The high Reynolds number conditions (nominal 8,000,000/foot) were obtained using the ejector augmentation system. Angle of attack was varied from -2 to +20 degrees at sideslip angles of -2, 0, and +2 degrees. Sideslip was varied from -6 to +8 degrees at constant angles of attack from 0 to +20 degrees. Aileron settings were varied from -5 to +10 degrees at elevon deflections of -10, 0, and +10 degrees. Fixed aileron settings of 0 and 2 degrees in combination with various fixed elevon settings between -20 and +5 degrees were also run at varying angles of attack.

Parrell, H.↗
NASA NTRS1975

Use of shock tubes in high Reynolds number transonic testing

The shock tube as a device to fulfill current needs for testing of transonic airfoils at high Reynolds numbers is considered. Performance analysis of a gas-driven shock tube shows that transonic airfoil flows with chord Reynolds numbers in the range of 100 million can be produced. A study of flow development over a simple airfoil has been carried out in a shock tube at low and intermediate Reynolds numbers to assess the testing technique. Results obtained from schlieren photography and airfoil pressure measurements show that steady transonic flows similar to those produced for the same airfoil in a wind tunnel can be generated within the available testing time in a shock tube with properly contoured test section walls. The study indicates that the shock tube is a viable alternative as a facility for studying high Reynolds number transonic airfoil flows.

Cook, W. J.↗
NASA NTRS1976

Preliminary scaling laws for plasma current, ion kinetic temperature, and plasma number density in the NASA Lewis Bumpy Torus plasma

Parametric variation of independent variables which may affect the characteristics of the NASA Lewis Bumpy Torus plasma have identified those which have a significant effect on the plasma current, ion kinetic temperature, and plasma number density, and those which do not. Empirical power-law correlations of the plasma current, and the ion kinetic temperature and number density were obtained as functions of the potential applied to the midplane electrode rings, the background neutral gas pressure, and the magnetic field strength. Additional parameters studied include the type of gas, the polarity of the midplane electrode rings (and hence the direction of the radial electric field), the mode of plasma operation, and the method of measuring the plasma number density. No significant departures from the scaling laws appear to occur at the highest ion kinetic temperatures or number densities obtained to date.

Roth, J. R.↗
NASA NTRS1976

The U.S. 2.5-Meter Cryogenic High Reynolds Number Tunnel

The U.S. 2.5-Meter Cryogenic High Reynolds Number Tunnel is a fan-driven transonic wind tunnel scheduled for operation in 1981. It will operate at Mach numbers from 0.1 to 1.2, stagnation pressures from 1 to 9 bars, and stagnation temperatures from 352 to 80 Kelvin. The maximum Reynolds number capability will be 120 million at a Mach number of 1.0 based on a reference length of 0.25 meter. This paper describes the basis for the conceptual approach, the engineering design including unique features, and the performance operating envelopes for the tunnel.

Howell, R. R.↗
NASA NTRS1976

The fast decoding of Reed-Solomon codes using number theoretic transforms

It is shown that Reed-Solomon (RS) codes can be encoded and decoded by using a fast Fourier transform (FFT) algorithm over finite fields. The arithmetic utilized to perform these transforms requires only integer additions, circular shifts and a minimum number of integer multiplications. The computing time of this transform encoder-decoder for RS codes is less than the time of the standard method for RS codes. More generally, the field GF(q) is also considered, where q is a prime of the form K x 2 to the nth power + 1 and K and n are integers. GF(q) can be used to decode very long RS codes by an efficient FFT algorithm with an improvement in the number of symbols. It is shown that a radix-8 FFT algorithm over GF(q squared) can be utilized to encode and decode very long RS codes with a large number of symbols. For eight symbols in GF(q squared), this transform over GF(q squared) can be made simpler than any other known number theoretic transform with a similar capability. Of special interest is the decoding of a 16-tuple RS code with four errors.

Reed, I. S.↗
NASA NTRS1977

Side forces on a tangent ogive forebody with a fineness ratio of 3.5 at high angles of attack and Mach numbers from 0.1 to 0.7

An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics, at high angles of attack, of a tangent ogive forebody with a fineness ratio of 3.5. The investigation included the effects of nose bluntness, nose strakes, nose booms, a simulated canopy, and boundary-layer trips. The forebody was also tested with a short afterbody attached. Static longitudinal and lateral-directional stability data were obtained at Reynolds numbers ranging from 0.3 mil. to 3.8 mil. (based on base diameter) at a Mach number of 0.25, and at a Reynolds number of 0.8 mil. at Mach numbers ranging from 0.1 to 0.7. Angle of attack was varied from 0 to 88 deg at zero sideslip, and the sideslip angle was varied from -10 to 30 deg at angles of attack of 40, 55, and 70 deg.

Keener, E. R.↗
NASA NTRS1977

Side forces on forebodies at high angles of attack and Mach numbers from 0.1 to 0.7: two tangent ogives, paraboloid and cone

An experimental investigation was conducted in the Ames 12-Foot Wind Tunnel to determine the subsonic aerodynamic characteristics of four forebodies at high angles of attack. The forebodies tested were a tangent ogive with fineness ratio of 5, a paraboloid with fineness ratio of 3.5, a 20 deg cone, and a tangent ogive with an elliptic cross section. The investigation included the effects of nose bluntness and boundary-layer trips. The tangent-ogive forebody was also tested in the presence of a short afterbody and with the afterbody attached. Static longitudinal and lateral/directional stability data were obtained. The investigation was conducted to investigate the existence of large side forces and yawing moments at high angles of attack and zero sideslip. It was found that all of the forebodies experience steady side forces that start at angles of attack of from 20 deg to 35 deg and exist to as high as 80 deg, depending on forebody shape. The side is as large as 1.6 times the normal force and is generally repeatable with increasing and decreasing angle of attack and, also, from test to test. The side force is very sensitive to the nature of the boundary layer, as indicated by large changes with boundary trips. The maximum side force caries considerably with Reynolds number and tends to decrease with increasing Mach number. The direction of the side force is sensitive to the body geometry near the nose. The angle of attack of onset of side force is not strongly influenced by Reynolds number or Mach number but varies with forebody shape. Maximum normal force often occurs at angles of attack near 60 deg. The effect of the elliptic cross section is to reduce the angle of onset by about 10 deg compared to that of an equivalent circular forebody with the same fineness ratio. The short afterbody reduces the angle of onset by about 5 deg.

Keener, E. R.↗
NASA NTRS1977

Comparison of jet Mach number decay data with a correlation and jet spreading contours for a large variety of nozzles

Small-scale circular, noncircular, single- and multi-element nozzles with flow areas as large as 122 sq cm were tested with cold airflow at exit Mach numbers from 0.28 to 1.15. The effects of multi-element nozzle shape and element spacing on jet Mach number decay were studied in an effort to reduce the noise caused by jet impingement on externally blown flap (EBF) STOL aircraft. The jet Mach number decay data are well represented by empirical relations. Jet spreading and Mach number decay contours are presented for all configurations tested.

Groesbeck, D. E.↗
NASA NTRS1976

Number of signatures necessary for accurate classification

This paper presents a procedure for determining the number of signatures to use in classifying multispectral scanner data. A large initial set of signatures is obtained by clustering the training points within each category (such as 'wheat' or 'other') to be recognized. These clusters are then combined into broader signatures by a program that considers each pair of signatures within a category, combines the best pair in the light of certain criteria, saves the combined signature and repeats the procedure until there is one signature for each category. The result is a collection of sets of signatures, one set for each number between the number of initial clusters and the number of categories. With the aid of statistics such as an estimate of the probability of misclassification between categories, the user can choose the smallest set satisfying his requirements for classification accuracy.

Richardson, W.↗
NASA NTRS1976

Modal density function and number of propagating modes in ducts

Often raised questions in duct sound propagation studies involve the total number of propagating modes, the number of propagating radial modes for a particular spinning lobe number, and the number of modes possible between two given values of cutoff ratio or eigenvalue. These questions can be answered approximately by using the modal distribution function which is the integral of the modal density function for ducts in a manner similar to that previously published for architectural acoustics. The modal density functions are derived for rectangular and circular ducts with a uniform steady flow. Results from this continuous theory are compared to the actual (discrete) modal distributions.

Rice, E. J.↗
NASA NTRS1977

Aerodynamic characteristics of an F-8 aircraft configuration with a variable camber wing at Mach numbers from 0.70 to 1.15

A 0.1-scale model of an F-8 aircraft was tested in the Ames 14-Foot Transonic Wind Tunnel at Mach numbers from 0.7 to 1.15. Angle of attack was varied from -2 deg. to 22 deg. at sideslip angles of 0 deg and -5 deg. Reynolds number, dictated by the atmospheric stagnation pressure, varied with Mach number from 3.4 to 4.0 million based on mean aerodynamic chord. The model was configured with a wing designed to simulate the downward deflection of the leading and trailing edges of an advanced-technology-conformal-variable camber wing. This wing was also equipped with conventional (simple hinge) flaps. In addition, the model was tested with the basic F-8 wing to provide a reference for extrapolating to flight data. In general, at all Mach numbers the use of conformal flap deflections at both the leading edge and trailing edge resulted in slightly higher maximum lift coefficients and lower drag coefficients than with the use of simple hinge flaps. There were also found to be small improvements in the pitching-moment characteristics with the use of conformal flaps.

Boltz, F. W.↗