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At least 145 records · Page 8

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NASA NTRS1996

User's Manual for the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

This user's manual provides detailed instructions for the installation and the application of version 4.1 of the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA). Also provides simulation of flow field in thermochemical nonequilibrium around vehicles traveling at hypersonic velocities through the atmosphere. Earlier versions of LAURA were predominantly research codes, and they had minimal (or no) documentation. This manual describes UNIX-based utilities for customizing the code for special applications that also minimize system resource requirements. The algorithm is reviewed, and the various program options are related to specific equations and variables in the theoretical development.

Gnoffo, Peter A.↗

NASA NTRS1995

Aerothermodynamic Measurement and Prediction for Modified Orbiter at Mach 6 and 10

Detailed heat-transfer rate distributions measured laterally over the windward surface of an orbiter-like configuration using thin-film resistance heat-transfer gauges and globally using the newly developed relative intensity, two-color thermographic phosphor technique are presented for Mach 6 and 10 in air. The angle of attack was varied from 0 to 40 deg, and the freestream Reynolds number based on the model length was varied from 4 x 10(exp 5) to 6 x 10(exp 6) at Mach 6, corresponding to laminar, transitional, and turbulent boundary layers; the Reynolds number at Mach 10 was 4 x 10(exp 5), corresponding to laminar flow. The primary objective of the present study was to provide detailed benchmark heat-transfer data for the calibration of computational fluid-dynamics codes. Predictions from a Navier-Stokes solver referred to as the Langley aerothermodynamic upwind relaxation algorithm and an approximate boundary-layer solving method known as the axisymmetric analog three-dimensional boundary layer code are compared with measurement. In general, predicted laminar heat-transfer rates are in good agreement with measurements.

Micol, John R.↗

NASA NTRS1998

Aerothermodynamics of the Mars Global Surveyor Spacecraft

The aerothermodynamics characteristics of the Mars Global Surveyor spacecraft are investigated and reported. These results have been used by the Mars Global Surveyor mission planners to design the aerobraking phase of the mission. Analytical and Direct Simulation Monte Carlo computer codes were used with a detailed, three dimensional model of the spacecraft to evaluate spacecraft aerobraking characteristics for flight in free molecular and transitional flow regimes. The spacecraft is found to be aerodynamically stable in aerobraking and planned contingency configurations. Aerodynamic forces, moments, and heating are found to be highly dependent on atmospheric density. Accommodation coefficient. is seen to strongly influence drag coefficient. Transitional flow effects are found to reduce overall solar panel heating. Attitude control thruster plumes are shown to interact with the freestream, diminishing the effectiveness of the attitude control system and even leading to thrust reversal. These plume-freestream interaction effects are found to be highly dependent on freestream density.

Shane, Russell W.↗

NASA NTRS1985

Aerothermodynamic Heating Analysis of Aerobraking and Aeromaneuvering Orbital Transfer Vehicles

The thermal-protection requirements of two aeroassisted orbital transfer vehicles (AOTVS) are analyzed for return missions between the geosynchronous and Shuttle orbits. One of the designs is a specialized version of a previously proposed generic aerobraking vehicle that Is capable of only delivery-type operations. The other Is a high-lift aeromaneuvering vehicle that is optimized for low Earth orbit sortie missions involving large, multiple plane-inclination changes. The aerothermal environment of the aerobraking vehicle is analyzed using state-of-the-art methods for nonequilibrium-radiative and convective heating that incorporate refinements unique to the configuration. The heating analysis of the aeromaneuvering vehicle required the development of a flowfield model for rarefied-hypersonic flow over a lifting surface at incidence. The predicted aerothermodynamic heating characteristics for both vehicles are correlated with thermal-control

Menees, Gene P.↗

NASA NTRS2000

Aerothermodynamic Analysis of Stardust Sample Return Capsule with Coupled Radiation and Ablation

An aerothermodynamic analysis of the forebody aeroshell of the Stardust Sample Return Capsule is carried out by using the axisymmetric viscous shock-layer equations with and without fully coupled radiation and ablation. Formulation of the viscous shock-layer equations with shoulder radius as the length scale and implementation of the Vigneron pressure condition allow resolution of the flowfield over the shoulder. With a predominantly supersonic outflow over the shoulder, a globally iterated solution or viscous shock-layer equations can be obtained. The stagnation-point results are obtained along a specified trajectory, whereas detailed calculations along the body are provided at the peak-heating point. The equilibrium calculations with ablation injection are the focus of the present study because of the lack of a general chemical nonequilibrium analysis that accounts for both surface and flowfield effect. The equilibrium calculations also provide a simple way to conserve surface (and flowfield) elemental composition for the current small ablation injection rates, where the surface elemental composition is a mixture of freestream and ablator elements. Therefore, the coupled laminar and turbulent flow solutions with radiation and ablation are obtained by using the equilibrium flow chemistry, whereas a nonequilibrium chemistry model is used for solutions without ablation and turbulence. Various computed results are compared with those obtained by the other researchers.

Gupta, Roop N.↗

NASA NTRS2001

Aerothermodynamic Environment Definition for the Genesis Sample Return Capsule

NASA's Genesis sample return mission will be the first to return material from beyond the Earth-Moon system. NASA Langley Research Center supported this mission with aerothermodynamic analyses of the sample return capsule. This paper provides an overview of that effort. The capsule is attached through its forebody to the spacecraft bus. When the attachment is severed prior to Earth entry, forebody cavities remain. The presence of these cavities could dramatically increase the heating environment in their vicinity and downstream. A combination of computational fluid dynamics calculations and wind tunnel phosphor thermography tests were employed to address this issue. These results quantify the heating environment in and around the cavities, and were a factor in the decision to switch forebody heat shield materials. A transition map is developed which predicts that the flow aft of the penetrations will still be laminar at the peak heating point of the trajectory. As the vehicle continues along the trajectory to the peak dynamic pressure point, fully turbulent flow aft of the penetrations could occur. The integrated heat load calculations show that a heat shield sized to the stagnation point levels will be adequate for the predicted environment aft of the penetrations.

Cheatwood, F. McNeil↗

NASA NTRS2000

Aerothermodynamic Analysis of the Project FIRE II Afterbody Flow

35 years later, the Project FIRE II ballistic reentry to Earth at a nominal velocity of 11.4 km/s remains one of the best sources of heating data for the design of sample return capsules. The data from this flight experiment encompass both the thermochemical non-equilibrium and equilibrium flow regimes and include measurements of both radiative and total heating on the forebody and afterbody. Because of this, a number of researchers have performed computational fluid dynamics (CFD) simulations of the forebody of the FIRE II entry vehicle, with generally good results. In particular, Olynick et. al. coupled a Navier-Stokes solver (GIANTS) with a radiation code (NOVAR) and showed excellent agreement in surface heat transfer over the FIRE II trajectory between 1634 and 1651 seconds (77 km to 37 km). However, in most cases the primary motivation of the previous work was to understand and model the coupling between shock layer radiation and aerothermodynamics, and thus the simulations concentrated on the forebody flow only. To our knowledge there have been no prior published attempts to reproduce the afterbody heating data. However, an understanding of this data is critical to our efforts to design the next generation of Earth and planetary entry vehicles and to assess our need for additional flight data.

Wright, Micheal J.↗

NASA NTRS1995

Aerothermodynamic Facilities And Measurement: Flow Characterization in Shock Tunnels

This presentation will examine the key performance aspects of shock tunnels as they relate to their use as aerothermodynamic flow simulation facilities. Assessment of shock tube reservoir conditions and flow contaminants generated in the shock tube will be presented along with their limiting impact on viable test envelopes, Facility nozzle performance as it pertains to test time assessment and nozzle exit flow quality (survey of pressure, temperature, and species) will be addressed. Also included will be a discussion of free stream flow diagnostics, both intrusive and nonintrusive, for measurement of critical flow properties not directly inferred from surface mounted transducers. The use of computational fluid dynamics for purposes of validating experimental measurements as well as predicting performance in regimes where measurements are not feasible or possible will be discussed. The use of CFD for facility research and design will also be presented.

Cavolowsky, John A.↗

NASA NTRS2001

Generic Aerothermodynamic Database

The aerothermodynamic database consists of Navier-Stokes solutions in PLOT3D format for the spherical-like nose cap of the X-38. These solutions are chosen to follow a reentry trajectory. The Mach number varies from 7.89 to 1.72 while the angle of attack varies from 38.73 degrees to 17.08 degrees. The Reynolds number varies from 4.139E6 to 2.766E7. Coarse, Medium and Fine grid solutions are supplied for each point in the data base. In addition to the flow solutions, a table of force and moment coefficients as well as the maximum heat transfer coefficient is also stored in the database.

Olsen, Thomas H.↗

NASA NTRS2002

Constrained Aerothermodynamic Design of Hypersonic Vehicles

An investigation was conducted into possible methods of incorporating a hypersonic design capability with aerothermodynamic constraints into the CDISC aerodynamic design tool. The work was divided into two distinct phases: develop relations between surface curvature and hypersonic pressure coefficient which are compatible with CDISC's direct-iterative design method; and explore and implement possible methods of constraining the heat transfer rate over all or portions of the design surface. The main problem in implementing this method has been the weak relationship between surface shape and pressure coefficient at the stagnation point and the need to design around the surface blunt leading edge where there is a slope singularity. The final results show that some success has been achieved, but further improvements are needed.

Gally, Tom↗

NASA NTRS2002

Hypersonic Airbreathing Propulsion: An Aerodynamics, Aerothermodynamics, and Acoustics Competency White Paper

This White Paper examines the current state of Hypersonic Airbreathing Propulsion at the NASA Langley Research Center and the factors influencing this area of work and its personnel. Using this knowledge, the paper explores beyond the present day and suggests future directions and strategies for the field. Broad views are first taken regarding potential missions and applications of hypersonic propulsion. Then, candidate propulsion systems that may be applicable to these missions are suggested and discussed. Design tools and experimental techniques for developing these propulsion systems are then described, and approaches for applying them in the design process are considered. In each case, current strategies are reviewed and future approaches that may improve the techniques are considered. Finally, the paper concentrates on the needs to be addressed in each of these areas to take advantage of the opportunities that lay ahead for both the NASA Langley Research Center and the Aerodynamic Aerothermodynamic, and Aeroacoustics Competency. Recommendations are then provided so that the goals set forth in the paper may be achieved.

Drummond, J. Philip↗

NASA NTRS2003

Team Software Development for Aerothermodynamic and Aerodynamic Analysis and Design

A collaborative approach to software development is described. The approach employs the agile development techniques: project retrospectives, Scrum status meetings, and elements of Extreme Programming to efficiently develop a cohesive and extensible software suite. The software product under development is a fluid dynamics simulator for performing aerodynamic and aerothermodynamic analysis and design. The functionality of the software product is achieved both through the merging, with substantial rewrite, of separate legacy codes and the authorship of new routines. Examples of rapid implementation of new functionality demonstrate the benefits obtained with this agile software development process. The appendix contains a discussion of coding issues encountered while porting legacy Fortran 77 code to Fortran 95, software design principles, and a Fortran 95 coding standard.

Alexandrov, N.↗

NASA NTRS2004

Experimental Aerothermodynamics In Support Of The Columbia Accident Investigation

The technical foundation for the most probable damage scenario reported in the Columbia Accident Investigation Board's final report was largely derived from synergistic aerodynamic/aerothermodynamic wind tunnel measurements and inviscid predictions made at NASA Langley Research Center and later corroborated with engineering analysis, high fidelity numerical viscous simulations, and foam impact testing near the close of the investigation. This report provides an overview of the hypersonic aerothermodynamic wind tunnel program conducted at NASA Langley and illustrates how the ground-based heating measurements provided early insight that guided the direction and utilization of agency resources in support of the investigation. Global surface heat transfer mappings, surface streamline patterns, and shock shapes were measured on 0.0075 scale models of the Orbiter configuration with and without postulated damage to the thermal protection system. Test parametrics include angle of attack from 38 to 42 degs, sideslip angles of 38 to 42 degs, sideslip angles of plus or minus 1 deg, Reynolds numbers based upon model length from 0.05 x 10(exp 6) to 6.5 x 10(exp 6), and normal shock density ratios of 5 (Mach 6 Air) and 12 (Mach 6 CF4). The primary objective of the testing was to provide surface heating characteristics on scaled Orbiter models with outer mold line perturbations to simulate various forms of localized surface damage to the thermal protection system. Initial experimental testing conducted within two weeks of the accident simulated a broad spectrum of thermal protection system damage to the Orbiter windward surface and was used to refute several hypothesized forms of thermal protection system damage, which included gouges in the windward thermal protection system tiles, breaches through the wing new the main landing gear door, and protuberances along the wing leading edge that produced asymmetric boundary layer transition. As the forensic phase of the investigation developed and the condition of recovered debris was examined, increasing emphasis was placed on identifying wing leading edge damage (partially and fully missing reinforced carbon-carbon panels, and eventually holes in the wing leading edge with venting to the wing upper surface) that produced off-nominal heating trends consistent with extracted Orbiter flight recorder temperature data.

Horvath, Thomas J.↗

NASA NTRS1998

Langley Aerothermodynamic Facilities Complex: Enhancements and Testing Capabilities

Description, capabilities, recent upgrades, and utilization of the NASA Langley Research Center (LaRC) Aerothermodynamic Facilities Complex (AFC) are presented. The AFC consists of five hypersonic, blow-down-to-vacuum wind tunnels that collectively provide a range of Mach number from 6 to 20, unit Reynolds number from 0.04 to 22 million per foot and, most importantly for blunt configurations, normal shock density ratio from 4 to 12. These wide ranges of hypersonic simulation parameters are due, in part, to the use of three different test gases (air, helium, and tetrafluoromethane), thereby making several of the facilities unique. The Complex represents nearly three-fourths of the conventional (as opposed to impulse)-type hypersonic wind tunnels operational in this country. AFC facilities are used to assess and optimize the hypersonic aerodynamic performance and aeroheating characteristics of aerospace vehicle concepts and to provide benchmark aerodynamic/aeroheating data fr generating the flight aerodynamic databook and final design of the thermal protection system (TPS) (e.g., establishment of flight limitations not to exceed TPS design limits). Modifications and enhancements of AFC hardware components and instrumentation have been pursued to increase capability, reliability, and productivity in support of programmatic goals. Examples illustrating facility utilization in recent years to generate essentially all of the experimental hypersonic aerodynamic and aeroheating information for high-priority, fast-paced Agency programs are presented. These programs include Phase I of the Reusable Launch Vehicle (RLV) Advanced Technology Demonstrator, X-33 program, PHase II of the X-33 program, X-34 program, the Hyper-X program ( a Mach 5,7, and 10 airbreathing propulsion flight experiment), and the X-38 program (Experimental Crew Return Vehicle, X-CRV). Current upgrades/enchancements and future plans for the AFC are discussed.

Micol, J. R.↗

NASA NTRS1998

Multi-Component Diffusion with Application To Computational Aerothermodynamics

The accuracy and complexity of solving multicomponent gaseous diffusion using the detailed multicomponent equations, the Stefan-Maxwell equations, and two commonly used approximate equations have been examined in a two part study. Part I examined the equations in a basic study with specified inputs in which the results are applicable for many applications. Part II addressed the application of the equations in the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) computational code for high-speed entries in Earth's atmosphere. The results showed that the presented iterative scheme for solving the Stefan-Maxwell equations is an accurate and effective method as compared with solutions of the detailed equations. In general, good accuracy with the approximate equations cannot be guaranteed for a species or all species in a multi-component mixture. 'Corrected' forms of the approximate equations that ensured the diffusion mass fluxes sum to zero, as required, were more accurate than the uncorrected forms. Good accuracy, as compared with the Stefan- Maxwell results, were obtained with the 'corrected' approximate equations in defining the heating rates for the three Earth entries considered in Part II.

Sutton, Kenneth↗

NASA NTRS1998

Rarefied Aerothermodynamic Predictions for Mars Global Surveyor

Mars Global Surveyor (MGS) successfully completed its first phase of aerobraking in early 1998 and is the rst planetary mission to use aerobraking as a primary means of customizing its orbit to achieve its mission objectives. The aerobraking requirements together with post-launch anomalies presented a unique challenge to provide accurate predictions of the aerothermodynamic environment of the spacecraft in the rare transitional flow regime. Direct Simulation Monte Carlo (DSMC) and free molecular techniques were used to provide heating and aerodynamic predictions and to investigate a variety of rarefied flow phenomena across the regime; MGS is the first major planetary mission in which rare flow predictions have played such a critical role all the way through design, mission planning, and operational phases. This paper summarizes these studies with emphasis on transitional-flow and gas-surface interaction phenomena.

Wilmoth, R. G.↗

NASA NTRS1997

Distributed-Memory Computing With the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA)

The Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA), a Navier-Stokes solver, has been modified for use in a parallel, distributed-memory environment using the Message-Passing Interface (MPI) standard. A standard domain decomposition strategy is used in which the computational domain is divided into subdomains with each subdomain assigned to a processor. Performance is examined on dedicated parallel machines and a network of desktop workstations. The effect of domain decomposition and frequency of boundary updates on performance and convergence is also examined for several realistic configurations and conditions typical of large-scale computational fluid dynamic analysis.

Riley, Christopher J.↗

NASA NTRS2005

Computational Aerothermodynamic Assessment of Space Shuttle Orbiter Tile Damage: Open Cavities

Computational aerothermodynamic simulations of Orbiter windside tile damage in flight were performed in support of the Space Shuttle Return-to-Flight effort. The simulations were performed for both hypervelocity flight and low-enthalpy wind tunnel conditions and contributed to the Return-to-Flight program by providing information to support a variety of damage scenario analyses. Computations at flight conditions were performed at or very near the peak heating trajectory point for multiple damage scenarios involving damage windside acreage reaction cured glass (RCG) coated silica tile(s). The cavities formed by the missing tile examined in this study were relatively short leading to flow features which indicated open cavity behavior. Results of the computations indicated elevated heating bump factor levels predicted for flight over the predictions for wind tunnel conditions. The peak heating bump factors, defined as the local heating to a reference value upstream of the cavity, on the cavity floor for flight simulation were 67% larger than the peak wind tunnel simulation value. On the downstream face of the cavity the flight simulation values were 60% larger than the wind tunnel simulation values. On the outer mold line (OML) downstream of the cavity, the flight values are about 20% larger than the wind tunnel simulation values. The higher heating bump factors observed in the flight simulations were due to the larger driving potential in terms of energy entering the cavity for the flight simulations. This is evidenced by the larger rate of increase in the total enthalpy through the boundary layer prior to the cavity for the flight simulation.

Pulsonetti, Maria↗