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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 289 records · Page 16
All sourcesNASA NTRSDOE OSTI
NASA NTRS1975

The Shuttle Orbital Maneuvering System P-V-T Propellant Quantity Gaging Accuracy and Leak Detection Allowance for Four Instrumentation Conditions

The shuttle orbital maneuvering system (OMS) pressure-volume-temperature (P-V-T) propellant gaging module computes the quantity of usable OMS propellant remaining based on the real gas P-V-T relationship for the propellant tank pressurant, helium. The OMS P-V-T propellant quantity gaging error was determined for four sets of instrumentation configurations and accuracies with the propellant tank operating in the normal constant pressure mode and in the blowdown mode. The instrumentation inaccuracy allowance for propellant leak detection was also computed for these same four sets of instrumentation. These gaging errors and leak detection allowances are presented in tables designed to permit a direct comparison of the effectiveness of the four instrumentation sets. The results show the magnitudes of the improvements in propellant quantity gaging accuracies and propellant leak detection allowances which can be achieved by employing more accurate pressure and temperature instrumentation.

Duhon, D. D.↗
NASA NTRS1977

Effect of LN2 injection station location on the drive fan power and LN2 requirements of a cryogenic wind tunnel

A theoretical analysis comparing the fan power and coolant (LN2) flow rates resulting from injection of the LN2 either upstream or downstream of the drive fan of a closed circuit transonic cryogenic tunnel is presented. The analysis is restricted to steady state tunnel operation and to the condition that the tunnel walls are adiabatic. The stagnation pressure and temperature range of the tunnel is from 1.0 to 8.8 atm and from 300 K to liquefaction temperature, respectively. Calculations are made using real gas properties of nitrogen. Results show that the fan power and LN2 flow rates are lower if the LN2 is injected upstream of the fan. The lower fan inlet temperature resulting from injecting upstream of the fan has a greater influence on the power than does the additional mass flow going through the fan.

Adcock, J. B.↗
NASA NTRS1977

Operational experience in the Langley expansion tube with various test gases

The Langley Expansion Tube is an operational facility capable of producing good quality, highly repeatable, quasi-steady flow for test times sufficient to establish flow about blunt axisymmetric and two-dimensional models. Due to the capability of testing with arbitrary test gases, a wide range of real-gas, hypersonic-hypervelocity flow conditions may be generated. However, for a given test gas, the range of operating conditions producing useful flow is shown to be rather limited; hence, the facility yields a given flow condition for a given test gas, and variation in flow conditions comes about by using different test gases. Data are presented for operations using air and carbon dioxide as test gases. The driver gas was unheated helium at a nominal pressure of 5000 psi (34.5 MN sq m).

Miller, C. G.↗
NASA NTRS1978

Computational aerodynamics and the numerical aerodynamic simulation facility

Technical and economic reasons for accelerating the maturation of the discipline of computational aerodynamics include the cost of conducting the experiments required to provide the empirical data base for new aeronautical vehicles and the limitations in test facilities (Reynolds number, wall and support interferences, aeroelastic distortions, real-gas effects, etc.) for simulating the full-scale vehicle environment. General purpose computers do not have the necessary capability for the next stage of development. Solution of the three dimensional Reynolds averaged Naiver-Stokes equations in a short time to be practical for design purposes will require 40 times the power of current supercomputers. However, it is feasible to construct a special purpose processor that will meet these requirements to enhance the nation's aerodynamic design capability in the 1980's.

Peterson, V. L.↗
NASA NTRS1979

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

The past 6 years of operation with the NASA Langley 0.3 m transonic cryogenic tunnel (TCT) show that there are no insurmountable problems associated with cryogenic testing with gaseous nitrogen at transonic Mach numbers. The fundamentals of the concept were validated both analytically and experimentally and the 0.3 m TCT, with its unique Reynolds number capability, was used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects were developed and cryogenic tunnel conditions can be set and maintained accurately. Cryogenic cooling by injecting liquid nitrogen directly into the tunnel circuit imposes no problems with temperature distribution or dynamic response characteristics. Experience with the 0.3 m TCT, indicates that there is a significant learning process associated with cryogenic, high Reynolds number testing. Many of the questions have already been answered; however, factors such as tunnel control, run logic, economics, instrumentation, and model technology present many new and challenging problems.

Ray, E. J.↗
NASA NTRS1979

Full scale aircraft simulation with cryogenic tunnels and status of the National Transonic Facility

The paper reviews the results of some of the real-gas studies made at Langley that are directly related to establishing the range of operating conditions that can be used in a cryogenic nitrogen wind tunnel and still be assured of valid full-scale simulation. Consideration is given to the important aerodynamic features, anticipated performance capability, status of construction, and projected operating data for the National Transonic Facility.

Kilgore, R. A.↗
NASA NTRS1979

Modeling and control of a LN2-GN2 operated closed circuit cryogenic wind tunnel

An explicit but simple lumped parameter nonlinear multivariable model of a LN2-GN2-operated closed circuit cryogenic wind tunnel has been developed and its basic features have been experimentally validated. The model describes the mass-energy interaction involved in the cryogenic tunnel process and includes the real gas properties of nitrogen gas.

Balakrishna, S.↗
NASA NTRS1979

Review of design and operational characteristics of the 0.3-meter transonic cryogenic tunnel

The fundamentals of cryogenic testing are validated both analytically and experimentally employing the 0.3-m transonic cryogenic tunnel. The tunnel with its unique Reynolds number capability has been used for a wide variety of aerodynamic tests. Techniques regarding real-gas effects have been developed and cryogenic tunnel conditions are set and maintained accurately. It is shown that cryogenic cooling, by injecting nitrogen directly into the tunnel circuit, imposes no problems with temperature distribution or dynamic response characteristics.

Ray, E. J.↗
NASA NTRS1981

Holographic flow visualization at the Langley Expansion Tube

A holographic system used for flow visualization at the Langley Expansion Tube is described. A ruby laser which can be singly or doubly pulsed during the short run time of less than 300 microns is used as the light source. With holography, sensitivity adjustments can be optimized after a run instead of before a run as with conventional flow visualization techniques. This results in an increased reliability of the flow visualization available for the study of real-gas effects on flow about models. Holographic techniques such as single-plate schlieren and shadowgraph, two plate interferometry, double pulse interferometry for perfect infinite-fringe interferograms, and double-pulse interferometry used to examine changes in the flow over a short time period are described and examples presented.

Goad, W. K.↗
NASA NTRS1981

Effect of choked flow on terminal characteristics of MPD thrusters

A one-dimensional steady real gas MPD accelerator model based on data for various electrode lengths and anode orifice radii is presented. The model relates the electric field to the boundary conditions in a form that agrees with experiment over parametric variations in current and mass flow. The electric field is sensitive to the formulations and assumptions bound into the analysis, and is useful in the modeling of outflow properties. It is shown that the location of the sonic point in the arc chamber shifts upstream for high current operation, and the relation between terminal voltage and current is demonstrated. Experiments using an anode suggested by the theory yield a thrust efficiency 1.5 times that of the unmodified thruster.

King, D. Q.↗
NASA NTRS1982

Measured pressure distributions, aerodynamic coefficients and shock shapes on blunt bodies at incidence in hypersonic air and CF4

Pressure distributions, aerodynamic coefficients, and shock shapes were measured on blunt bodies of revolution in Mach 6 CF4 and in Mach 6 and Mach 10 air. The angle of attack was varied from 0 deg to 20 deg in 4 deg increments. Configurations tested were a hyperboloid with an asymptotic angle of 45 deg, a sonic-corner paraboloid, a paraboloid with an angle of 27.6 deg at the base, a Viking aeroshell generated in a generalized orthogonal coordinate system, and a family of cones having a 45 deg half-angle with spherical, flattened, concave, and cusp nose shapes. Real-gas effects were simulated for the hperboloid and paraboloid models at Mach 6 by testing at a normal-shock density ratio of 5.3 in air and 12 CF4. Predictions from simple theories and numerical flow field programs are compared with measurement. It is anticipated that the data presented in this report will be useful for verification of analytical methods for predicting hypersonic flow fields about blunt bodies at incidence.

Miller, C. G., III↗
NASA NTRS1982

Experimental investigation of gamma effects on heat transfer to a 0.006 scale Shuttle Orbiter at Mach 6

Real-gas effects on heating to the windward centerline of the Shuttle Orbiter and shock shape were simulated by testing a 0.006-scale Orbiter model in Mach 6 air and CF4 flows. The range of free-stream Reynolds number, based on model length, was 0.4 x 10 to the 6th to 5.3 x 10 to the 6th for air, resulting in laminar and turbulent heating. The Reynolds number for CF4 was 0.3 x 10 to the 6th and gamma was equal to 1.13, corresponding to a normal shock density ratio of 12 as compared to 5.25 in air. Angle of attack was varied from 15 deg to 45 deg in 5 deg increments. Centerline heating distributions were compared for the two gases, and to Mach 8 wind-tunnel data, predictions, and flight (STS-1 and STS-2) data. The present results revealed a significant increase in heating with increasing density ratio and decreasing Reynolds number at the lower angles of attack; the effects of density ratio and Reynolds number diminished with increasing angle of attack and were relatively small at the angles of attack for the hypersonic portion of Space Shuttle Orbiter flights STS-1 and STS-2. Wind-tunnel and flight heating distributions were in relatively good agreement, and the present heating rates in air were accurately predicted by a local infinite swept cylinder analysis.

Miller, C. G.↗
NASA NTRS1982

Plume/flowfield jet interaction effects on the Space Shuttle Orbiter during entry

A comparison of wind tunnel test results with actual flight test data from the performance of the Orbiter Reaction Control System during the STS-1 and -2 flights is presented. Jet interaction phenomena were modeled analytically and a decision was made to develop a jet interaction scaling parameter from empirical measurements wherever possible. A number of different scaling parameters resulted, and errors bounds were determined for tunnel-to-tunnel differences, model to tunnel differences, balance accuracy, real gas effects, varying nozzle configurations, and angle of sideslip. Flight test data has concentrated on the lateral/directional jet interaction of the aft yaw jets, which operate almost continuously from entry down to Mach 1 conditions. A discussion of rolling moments, side forces, and results of usage of the aft pitch and roll jets indicates an underprediction of the side force and yawing moment and overprediction of pitch and roll jet effects.

Kanipe, D. B.↗
NASA NTRS1983

Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows

Laminar heat transfer rates were measured on spherically blunted, 13 deg/7 deg on axis and bent biconics (fore cone bent 7 deg upward relative to aft cone) at hypersonic hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from 0 to 20 deg in 4 deg increments. The effect of nose bend, angle of attack, and real gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three dimensional parabolized Navier-Stokes equations.

Miller, C. G.↗
NASA NTRS1983

Transient flow analysis of the AEDC/HPDE MHD generator

A hybrid Lax-Wendroff/Method of Characteristics computer code has been developed for numerical simulation of flow transients associated with the operation of MHD generator facilities. The code employs the shock-fitting method, with an Eulerian formulation of the basic conservation equations and explicit tracking of shock waves. Pressure, temperature, and velocity are used as primary integration variables to simplify interfacing of the code with real-gas thermodynamic and transport property tables. Application of the code to the simulation of selected transients for the AEDC/HPDE MHD generator produced results that are in good agreement with experimental observations.

Wilson, D. R.↗
NASA NTRS1983

Heat-transfer distributions on biconics at incidence in hypersonic-hypervelocity He, N2, air, and CO2 flows

Laminar heat-transfer rates were measured on spherically blunted, 13 degrees/F degrees on-axis and bent biconics (fore cone bent 7 degrees upward relative to aft cone) at hypersonic-hypervelocity flow conditions in the Langley Expansion Tube. Freestream velocities from 4.5 to 6.9 km/sec and Mach numbers from 6 to 9 were generated using helium, nitrogen, air, and carbon dioxide test gases, resulting in normal shock density ratios from 4 to 19. Angle of attack, referenced to the axis of the aft cone, was varied from zero to 20 degrees in 4 degree increments. The effect of nose bend, angle of attack, and real-gas phenomena on heating distributions are presented along with comparisons of measurement to prediction from a code which solves the three-dimensional 'parabolized Navier-Stokes' equations.

Miller, C. G.↗
NASA NTRS1983

Stability and control over the supersonic and hypersonic speed range

The performance and aerodynamic characteristics of the Space Shuttle Columbia are analyzed over the speed range from Mach 2 to 26 using flight data taken from the first five Space Transportation System flights. These data are used to reconstruct the entry trajectory, calculate vehicle performance, and estimate lateral stability and control derivatives including those assocated with the onboard Reaction Control System. The trajectory reconstruction process is discussed in its relationship to the determination of the vehicle flight characteristics. Vehicle performance results are presented which show that lift and drag were generally overpredicted by 3 percent and that the life-to-drag ratio was underpredicted by 1 percent. Anomalies in pitching-moment trim characteristics are shown and noted to be due to real-gas and Mach-number effects. Lateral stability and control derivatives estimated using programmed test input maneuvers are corrected with preflight predictions and are usually within the uncertainties associated with the predictions. The lateral reaction-control system effectiveness is evaluated.

Compton, H. R.↗
NASA NTRS1983

Enhancement of the CAVE computer code

The computer code CAVE (Conduction Analysis via Eigenvalues) is a convenient and efficient computer code for predicting two dimensional temperature histories within thermal protection systems for hypersonic vehicles. The capabilities of CAVE were enhanced by incorporation of the following features into the code: real gas effects in the aerodynamic heating predictions, geometry and aerodynamic heating package for analyses of cone shaped bodies, input option to change from laminar to turbulent heating predictions on leading edges, modification to account for reduction in adiabatic wall temperature with increase in leading sweep, geometry package for two dimensional scramjet engine sidewall, with an option for heat transfer to external and internal surfaces, print out modification to provide tables of select temperatures for plotting and storage, and modifications to the radiation calculation procedure to eliminate temperature oscillations induced by high heating rates. These new features are described.

Rathjen, K. A.↗