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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 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Splitting of inviscid fluxes for real gases

Flux-vector and flux-difference splittings for the inviscid terms of the compressible flow equations are derived under the assumption of a general equation of state for a real gas in equilibrium. No necessary assumptions, approximations for auxiliary quantities are introduced. The formulas derived include several particular cases known for ideal gases and readily apply to curvilinear coordinates. Applications of the formulas in a TVD algorithm to one-dimensional shock-tube and nozzle problems show their quality and robustness.

Liou, Meng-Sing↗
NASA NTRS1989

Computational aerothermodynamics

Computational aerothermodynamics concerns the coupling of real gas effects with equations of motion to include thermochemical rate processes for chemical and energy exchange phenomena. These processes concern the creation and destruction of gas species by chemical reactions and the transfer of energy between the various species and between the various energy modes (e.g., translation, rotation, vibration, ionization, dissociation/recombination, etc.) of the species. To gain some insight into when such phenomena occur for current and future aerospace flight vehicles the author shows the flight regimes of some typical vehicles (e.g., Concord, aerospace plane, Space Shuttle, associated space transfer vehicles, Apollo entry vehicle, etc.) in terms of flight altitude and flight speed. Also indicated are regimes where chemical reactions such as dissociation and ionization are important and where nonequilibrium thermochemical phenomena are important.

Deiwert, George S.↗
NASA NTRS1989

Direct simulation of three-dimensional flow about the AFE vehicle at high altitudes

Three-dimensional hypersonic rarefied flow about the Aeroassist Flight Experiment (AFE) vehicle was studied using the direct simulation Monte Carlo (DSMC) technique. Results are presented for the transitional flow regime encountered between 120 and 200 km altitudes with a reentry velocity of 9.92 km/s. In the simulations, a five-species reacting real-gas model that accounts for internal energies (rotational and vibrational) is used. The results indicate that the transitional effects are significant even at an altitude of 200 km and influence the overall vehicle aerodynamics. For the cases considered, the aerodynamic coefficients, surface pressures, convective heating, and flow field structure variations with rarefaction effects are presented.

Celenligil, M. Cevdet↗
NASA NTRS1990

Experimental and predicted aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration at Mach 6 and 10

The effects of Mach number, Reynolds number, and ratio of specific heats on the aerodynamic characteristics of a proposed Assured Crew Return Vehicle (ACRV) lifting-body configuration were examined for a range of angles of attack from -5 deg to 50 deg. Predictions made with a Langley-developed, three-dimensional Navier Stokes solver known as LAURA, which was exercised as an Euler solver for the present study, are compared with the experimental results. Unlike the Shuttle Orbiter, which experienced a significant nose-up increment in pitching moment with decreasing specific heat ratio (i.e., real gas effects), the aerodynamic characteristics of this lifting-body configuration are insensitive to changes in specific heat ratio. The maximum trimmed lift-to-drag ratio achieved was about 1.5. Predicted inviscid values of aerodynamic coefficients were generally in good agreement with measurement.

Micol, John R.↗
NASA NTRS1990

Enhancements and applications of DSMC for hypersonic rarefied flows

A review of the direct simulation Monte Carlo (DSMC) method of Bird is presented. The DSMC method provides the capability of simulating real gas flows in the rarefield flow regime. Recent developments and applications of the method for hypersonic flows are reported for both ground-based tests and during entry. The flows considered were three-dimensional, and both an existing general code and a new code under development were used.

Moss, James N.↗
NASA NTRS1990

Parametric study of an ODW scramaccelerator for hypersonic test facilities

A parametric study has been conducted for an oblique detonation-wave (ODW) 'scramaccelerator' suitable for projectile aerothermodynamics studies in real gas hypersonic test facilities. The results of the present analytical design evaluation indicate that an ODW scramaccelerator using conventional gaseous propellants can accelerate projectiles of 0.1 to 1000 kg masses to speeds in the 6-10 km/sec range. Potential applications for such an accelerator encompass a hypersonic ballistic test range, kinetic energy weapon accelerators, mass drivers to LEO, projectile terminal ballistics testing, projectile/target interaction studies, inertial welders, and shock compactors.

Humphrey, Joseph W.↗
NASA NTRS1991

Simulating The Performance Of A Scramjet

SCRAM computer program determines, according to one-dimensional mathematical model, cycle performance for airframe-integrated subsonic or supersonic combustion Ramjets. Reliable, efficient, and speedy ramjet designer's software tool usable on all standard computers down to those compatible with IBM PC-AT's. Scramjet cycle analysis code simulates, from nose to tail, hydrogen-fueled, airframe-integrated scramjet in real gas flow with equilibrium thermodynamic properties. Enables rapid scramjet performance estimates. Written in FORTRAN 77.

Walton, James T.↗
NASA NTRS1989

Computational Fluid Dynamics Program at NASA Ames Research Center

The Computational Fluid Dynamics (CFD) Program at NASA Ames Research Center is reviewed and discussed. The technical elements of the CFD Program are listed and briefly discussed. These elements include algorithm research, research and pilot code development, scientific visualization, advanced surface representation, volume grid generation, and numerical optimization. Next, the discipline of CFD is briefly discussed and related to other areas of research at NASA Ames including experimental fluid dynamics, computer science research, computational chemistry, and numerical aerodynamic simulation. These areas combine with CFD to form a larger area of research, which might collectively be called computational technology. The ultimate goal of computational technology research at NASA Ames is to increase the physical understanding of the world in which we live, solve problems of national importance, and increase the technical capabilities of the aerospace community. Next, the major programs at NASA Ames that either use CFD technology or perform research in CFD are listed and discussed. Briefly, this list includes turbulent/transition physics and modeling, high-speed real gas flows, interdisciplinary research, turbomachinery demonstration computations, complete aircraft aerodynamics, rotorcraft applications, powered lift flows, high alpha flows, multiple body aerodynamics, and incompressible flow applications. Some of the individual problems actively being worked in each of these areas is listed to help define the breadth or extent of CFD involvement in each of these major programs. State-of-the-art examples of various CFD applications are presented to highlight most of these areas. The main emphasis of this portion of the presentation is on examples which will not otherwise be treated at this conference by the individual presentations. Finally, a list of principal current limitations and expected future directions is given.

Holst, Terry L.↗
NASA NTRS1989

Johnson Space Center CFD overview

Recent applications and development of CFD technology have focused on flow problems that are critically important to the operation and design of space flight vehicles. The main effort is spent on the Space Shuttle in order to provide an understanding of the cryogenic fluid in the duct connecting the External Tank and the Main Engines, the subsonic flow surrounding the Orbiter during crew egress maneuvers, the transonic aerodynamic forces on the Orbiter fuselage and wing, the high angle-of-attack abort flight, and the aerodynamic heating during entry. To provide in-depth analyses for such diverse problems within a timely schedule, matured panel codes and a state-of-the-art incompressible turbulent flow code were adapted. Collaboration with Ames Research Center has resulted in a Shuttle ascent aerodynamic code; and a viscous chemical nonequilibrium code is being developed for predicting Orbiter real-gas aerodynamics and finite-catalytic heating. The remaining activities are devoted to the prediction of the flow environment around the Aeroassist Flight Experiment vehicle at hypersonic speeds and high altitudes. A thermochemical nonequilibrium Navier-Stokes code has been developed on the basis of two- temperature and 11-species models for solving both the shock layer and near wake. After validating the code against wind-tunnel aerodynamic, pressure and heating data, the code is being used to supplement the ground test facilities in predicting a more realistic flight environment. CFD technology is being relied upon by other programs as well in the consideration of candidate configurations.

Li, Chien P.↗
NASA NTRS1989

Three-dimensional calculation of supersonic reacting flows using an LU scheme

A new three-dimensional numerical program incorporated with comprehensive real gas property models was developed to simulate supersonic reacting flows. The code employs an implicit finite volume, Lower-Upper (LU) time-marching method to solve the complete Navier-Stokes and species equations in a fully-coupled and very efficient manner. A chemistry model with nine species and eighteen reaction steps are adopted in the program to represent the chemical reaction of H2 and air. To demonstrate the capability of the program, flow fields of underexpanded hydrogen jets transversely injected into supersonic air stream inside the combustors of scramjets are calculated. Results clearly depict the flow characteristics, including the shock structure, separated flow regions around the injector, and the distribution of the combustion products.

Yu, Sheng-Tao↗
NASA NTRS1991

Local and overall aerodynamic coefficients for bodies in hypersonic, rarefied flow

A computational method is given for the prediction of local pressure and viscous shear stress on windward surfaces of bluff, convex, axisymmetric or quasi-axisymmetric, hypersonic bodies in the transitional, rarefied flow regime. Overall aerodynamic forces and moments are computed by integration of the local quantities. The method is based upon a correlation of local pressure and shear stress computed by the direct simulation Monte Carlo (DSMC) numerical technique for cold wall, real gas conditions and some supplemental data from low-density, hypersonic wind tunnels. The relative simplicity of the method makes it feasible to do the necessary calculations with a personal computer. Two-dimensional shapes and leeward surfaces are not included in the scope of the method as it is presented here. Results are compared with DSMC computations for both local and overall coefficients. The latter includes sphere and blunt cone drag as well as lift and pitching moment coefficients for the NASA AFE vehicle at various angles of attack. Very satisfactory agreement is shown.

Potter, J. Leith↗
NASA NTRS1991

PARCEQ2D heat transfer grid sensitivity analysis

The material presented in this paper is an extension of two-dimensional Aeroassist Flight Experiment (AFE) results shown previously. This study has focused on the heating rate calculations to the AFE obtained from an equilibrium real gas code, with attention placed on the sensitivity of grid dependence and wall temperature. Heat transfer results calculated by the PARCEQ2D code compare well with those computed by other researchers. Temperature convergence in the case of kinetic transport has been accomplished by increasing the wall temperature gradually from 300 K to the wall temperature of 1700 K.

Saladino, Anthony J.↗
NASA NTRS1989

Hypersonic rarefied-flow aerodynamics inferred from Shuttle Orbiter acceleration measurements

Data obtained from multiple flights of sensitive accelerometers on the Space Shuttle Orbiter during reentry have been used to develop an improved aerodynamic model for the Orbiter normal- and axial-force coefficients in hypersonic rarefied flow. The lack of simultaneous atmospheric density measurements was overcome in part by using the ratio of normal-to-axial acceleration, in which density cancels, as a constraint. Differences between the preflight model and the flight-acceleration-derived model in the continuum regime are attributed primarily to real gas effects. New insights are gained into the variation of the force coefficients in the transition between the continuum regime and free molecule flow.

Blanchard, R. C.↗
NASA NTRS1990

Aeroassisted Flight Experiment aerodynamic characteristics at flight conditions

The success of NASA's Aeroassisted Flight Experiment project depends on the suitable placement of instrumentation on the vehicle surface and the ability of the vehicle to fly the maximum science payload. The initial aerodynamic data base was established using wind tunnel data and CFD analyses, where the influence of real-gas effects precluded the use of ground-facility data. More recently, a viscous thermochemical nonequilibrium flow analysis about the complete vehicle, including the wake, has updated the vehicle aerodynamic data base.

Weilmuenster, K. James↗
NASA NTRS1990

On the Goertler vortex instability mechanism at hypersonic speeds

The linear instability of the hypersonic boundary layer on a curved wall is considered. As a starting point real-gas effects are ignored and the viscosity of the fluid is taken to be related to the temperature either by Chapman's Law or by Sutherland's Law. It is shown that the flow is susceptible to Goertler vortices. If Chapman's Law is used the vortices are trapped in the logarithmically thin adjustment layer in which the temperature of the basic flow changes rapidly to its free stream value and the nonuniqueness of the neutral stability curve associated with incompressible Goertler vortices is shown to disappear at high Mach numbers if the appropriate 'fast' streamwise dependence of the instability is built into the disturbance flow structure. If, on the other hand, Sutherland's Law is used, the vortices are found to spread into an O(1) region and the concept of a unique neutral stability curve is not tenable because of the nonparallel effects. For both laws the leading order terms in the expansions of the Goertler number are independent of the wave number and are due to the curvature of the basic state.

Hall, P.↗
NASA NTRS1991

Computational methodology for radiation heat transfer in the flowfield of an AOTV

A computational methodology is developed for the calculation of radiation heat transfer in a non-scattering medium where nonequilibrium conditions exist and give rise to complex spectra. The specific problem of radiation in an aeroassisted orbital transfer vehicle (AOTV) flowfield is addressed. Nonequilibrium radiation from the gases around a hypersonic vehicle is evaluated using the NASA-Ames NEQAIR program together with a three-dimensional geometrical flowfield model. Nonequilibrium compositions and temperatures are taken from NASA-Langley 3-D hypersonic flowfield calculations that include real-gas effects and finite-rate chemical kinetics. It is shown that the concept of a transmission path adjustment, together with precomputations of curves of growth based upon local properties at flowfield locations, facilitates radiation calculations. Sample calculations of directional and spectral distributions of gas radiation received upon a hypersonic vehicle are presented.

Fuehrer, P. L.↗
NASA NTRS1991

The effects of self-generated and applied magnetic fields on the computation of flow over a Mars return aerobrake

A CFD technique is developed to calculate the electromagnetic phenomena simultaneously with the fluid flow in the shock layer over an axisymmetric blunt body in a thermal-equilibrium chemical-nonequilibrium environment. The flowfield is solved using an explicit time-marching, first-order spatially accurate scheme. The electromagnetic phenomena are coupled to the real-gas flow solver through an iterative procedure. The electromagnetic terms introduce a strong stiffness, which was overcome by using significantly smaller time steps for the electromagnetic conservation equation. The technique is applied in calculating the flow over a Mars return aerobrake vehicle entering the Earth's atmosphere. For the case where no external field is applied, the electromagnetic effects have little impact on the flowfield.

Palmer, Grant↗
NASA NTRS1991

Trim angle measurements in free-flight facilities

The aerodynamic cofficients and trim angle for an aerobrake at Mach 9.2 and 11.8 were found using a combination of experiment and computation. Free-flight tests were performed at NASA Ames Research Center's Hypervelocity Free-Flight Aerodynamic Facility, and the forebody pressure distribution was calculated using a three-dimensional Navier-Stokes code with an effective specific heat ratio. Using the computed drag, lift, and moments to prescribe the number of terms in the aerodynamic coefficient expansions and to specify the values of the higher order terms, the experimental aerodynamic coefficients and trim angle were found using a six-degree-of-freedom, weighted, least-squares analysis. The experimental and computed aerodynamic coefficients and trim angles are in good agreement. The trim angle obtained from the free-flight tests, 14.7 deg, differs from the design trim angle, 17 deg, and from the Langley wind tunnel results, 12 deg in air and 17 deg in CF4. These differences are attributable to real-gas effects.

Yates, Leslie A.↗