Engineering topics

Gnoffo, Peter A.

Publications and source records attributed to Gnoffo, Peter A..

At least 73 records · Page 4

NASA NTRS2001

CFD Validation Studies for Hypersonic Flow Prediction

A series of experiments to measure pressure and heating for code validation involving hypersonic, laminar, separated flows was conducted at the Calspan-University at Buffalo Research Center (CUBRC) in the Large Energy National Shock (LENS) tunnel. The experimental data serves as a focus for a code validation session but are not available to the authors until the conclusion of this session. The first set of experiments considered here involve Mach 9.5 and Mach 11.3 N2 flow over a hollow cylinder-flare with 30 degree flare angle at several Reynolds numbers sustaining laminar, separated flow. Truncated and extended flare configurations are considered. The second set of experiments, at similar conditions, involves flow over a sharp, double cone with fore-cone angle of 25 degrees and aft-cone angle of 55 degrees. Both sets of experiments involve 30 degree compressions. Location of the separation point in the numerical simulation is extremely sensitive to the level of grid refinement in the numerical predictions. The numerical simulations also show a significant influence of Reynolds number on extent of separation. Flow unsteadiness was easily introduced into the double cone simulations using aggressive relaxation parameters that normally promote convergence.

Gnoffo, Peter A.↗

NASA NTRS2001

Afterbody Heating Characteristics of a Proposed Mars Sample Return Orbiter

Aeroheating wind-tunnel tests were conducted on a 0.028 scale model of an orbiter concept considered for a possible Mars sample return mission. The primary experimental objectives were to characterize hypersonic near wake closure and determine if shear layer impingement would occur on the proposed orbiter afterbody at incidence angles necessary for a Martian aerocapture maneuver. Global heat transfer mappings, surface streamline patterns, and shock shapes were obtained in the NASA Langley 20-inch Mach 6 Air and CF4 Tunnels for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 1,400 to 415,000, angles of attack ranging from -5 to 10 degrees at 0, 3, and 6 deg sideslip, and normal-shock density, ratios of 5 and 12. Laminar, transitional, and turbulent shear layer impingement on the cylindrical afterbody was inferred from the measurements and resulted in a localized heating maximum that ranged from 40 to 75% of the reference forebody stagnation point heating. Comparison of laminar heating prediction to experimental measurement along the orbiter afterbody highlight grid alignment challenges associated with numerical simulation of three-dimensional separated wake flows.

Horvath, Thomas J.↗

NASA NTRS2000

A 3-D Coupled CFD-DSMC Solution Method With Application to the Mars Sample Return Orbiter

A method to obtain coupled Computational Fluid Dynamics-Direct Simulation Monte Carlo (CFD-DSMC), 3-D flow field solutions for highly blunt bodies at low incidence is presented and applied to one concept of the Mars Sample Return Orbiter vehicle as a demonstration of the technique. CFD is used to solve the high-density blunt forebody flow defining an inflow boundary condition for a DSMC solution of the afterbody wake flow. By combining the two techniques in flow regions where most applicable, the entire mixed flow field is modeled in an appropriate manner.

Glass, Christopher E.↗

NASA NTRS1999

Numerical and Analytical Solutions of Hypersonic Interactions Involving Surface Property Discontinuities

The local viscous-inviscid interaction field generated by a wall temperature jump on a flat plate in supersonic flow and on the windside of a Reusable Launch Vehicle in hypersonic flow is studied in detail by both a Navier-Stokes numerical code and an analytical triple-deck model. Treatment of the rapid heat transfer changes both upstream and downstream of the jump is included. Closed form relationships derived from the triple-deck theory are presented. The analytically predicted pressure and heating variations including upstream influence are found to be in generally good agreement with the Computational Fluid Dynamic (CFD) predictions. These analyses not only clarify the interactive physics involved but also are useful in preliminary design of thermal protection systems and as an insertable module to improve CFD code efficiency when applied to such small-scale interaction problems. The analyses only require conditions at the wall and boundary-layer edge which are easily extracted from a baseline, constant wall temperature, CFD solution.

Gnoffo, Peter A.↗

NASA NTRS1998

Computational Aeroheating Predictions for X-34

Radiative equilibrium surface temperatures, heating rates, streamlines, surface pressures, and flow-field features as predicted by the Langley Aerothermodynamic Upwind Relaxation Algorithm (Laura) are presented for the X-34 Technology Demonstrator. Results for two trajectory points corresponding to entry peak heating and two control surface deflections are discussed. This data is also discussed in context of Thermal Protection System (TPS) design issues. The work presented in this report is part of a larger effort to define the X-34 aerothermal environment, including the application of engineering codes and wind-tunnel studies.

Kelb, William L.↗

NASA NTRS1998

Computational Aeroheating Predictions for X-34

Radiative equilibrium surface temperatures, heating rates, streamlines, surface pressures, and flow-field features as predicted by the Langley Aerothermodynamic Upwind Relaxation Algorithm (LAURA) are presented for the X-34 Technology Demonstrator. Results for two trajectory points corresponding to entry peak heating and two control surface deflections are discussed. This data is also discussed in the context of Thermal Protection System (TPS) design issues. The work presented in this report is part of a larger effort to define the X-34 aerothermal environment, including the application of engineering codes and wind-tunnel studies.

Kleb,William H.↗

NASA NTRS1998

Prediction and Validation of Mars Pathfinder Hypersonic Aerodynamic Data Base

Postflight analysis of the Mars Pathfinder hypersonic, continuum aerodynamic data base is presented. Measured data include accelerations along the body axis and axis normal directions. Comparisons of preflight simulation and measurements show good agreement. The prediction of two static instabilities associated with movement of the sonic line from the shoulder to the nose and back was confirmed by measured normal accelerations. Reconstruction of atmospheric density during entry has an uncertainty directly proportional to the uncertainty in the predicted axial coefficient. The sensitivity of the moment coefficient to freestream density, kinetic models and center-of-gravity location are examined to provide additional consistency checks of the simulation with flight data. The atmospheric density as derived from axial coefficient and measured axial accelerations falls within the range required for sonic line shift and static stability transition as independently determined from normal accelerations.

Gnoffo, Peter A.↗

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 NTRS1997

Advances in Computational Capabilities for Hypersonic Flows

The paper reviews the growth and advances in computational capabilities for hypersonic applications over the period from the mid-1980's to the present day. The current status of the code development issues such as surface and field grid generation, algorithms, physical and chemical modeling, and validation is provided. A brief description of some of the major codes being used at NASA Langley Research Center for hypersonic continuum and rarefied flows is provided, along with their capabilities and deficiencies. A number of application examples are presented, and future areas of research to enhance accuracy, reliability, efficiency, and robustness of computational codes are discussed.

Kumar, Ajay↗

NASA NTRS1997

Computational Aerothermodynamic Design Issues for Hypersonic Vehicles

A brief review of the evolutionary progress in computational aerothermodynamics is presented. The current status of computational aerothermodynamics is then discussed, with emphasis on its capabilities and limitations for contributions to the design process of hypersonic vehicles. Some topics to be highlighted include: (1) aerodynamic coefficient predictions with emphasis on high temperature gas effects; (2) surface heating and temperature predictions for thermal protection system (TPS) design in a high temperature, thermochemical nonequilibrium environment; (3) methods for extracting and extending computational fluid dynamic (CFD) solutions for efficient utilization by all members of a multidisciplinary design team; (4) physical models; (5) validation process and error estimation; and (6) gridding and solution generation strategies. Recent experiences in the design of X-33 will be featured. Computational aerothermodynamic contributions to Mars Pathfinder, METEOR, and Stardust (Comet Sample return) will also provide context for this discussion. Some of the barriers that currently limit computational aerothermodynamics to a predominantly reactive mode in the design process will also be discussed, with the goal of providing focus for future research.

Gnoffo, Peter A.↗

NASA NTRS1996

Analytic corrections to CFD heating predictions accounting for changes in surface catalysis

Integral boundary-layer solution techniques applicable to the problem of determining aerodynamic heating rates of hypersonic vehicles in the vicinity of stagnation points and windward centerlines are briefly summarized. A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (but time intensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations is described. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. It allows a designer at a workstation to address two questions, given a single CFD solution. (1) How much does the heating change for a thermal protection system with different catalytic properties than was used in the original CFD solution? (2) How does the heating change at the interface of two different TPS materials with an abrupt change in catalytic efficiency? The answer to the second question is particularly important, because abrupt changes from low to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption.

Gnoffo, Peter A.↗

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 NTRS1996

Aerothermodynamic Analysis of Commercial Experiment Transporter (COMET) Reentry Capsule

An aerothermodynamic analysis of the Commercial Experiment Transporter (COMET) reentry capsule has been performed using the laminar thin-layer Navier-Stokes solver Langley Aerothermodynamic Upwind Relaxation Algorithm. Flowfield solutions were obtained at Mach numbers 1.5, 2, 5, 10, 15, 20, 25, and 27.5. Axisymmetric and 5, 10, and 20 degree angles of attack were considered across the Mach-number range, with the Mach 25 conditions taken to 90 degrees angle of attack and the Mach 27.5 cases taken to 60 degrees angle of attack. Detailed surface heat-transfer rates were computed at Mach 20 and 25, revealing that heating rates on the heat-shield shoulder ,can exceed the stagnation-point heating by 230 percent. Finite-rate chemistry solutions were performed above Mach 10, otherwise perfect gas computations were made. Drag, lift, and pitching moment coefficients are computed and details of a wake flow are presented. The effect of including the wake in the solution domain was investigated and base pressure corrections to forebody drag coefficients were numerically determined for the lower Mach numbers. Pitching moment comparisons are made with direct simulation Monte Carlo results in the more rarefied flow at the highest Mach numbers, showing agreement within two-percent. Thin-layer Navier-Stokes computations of the axial force are found to be 15 percent higher across the speed range than the empirical/Newtonian based results used during the initial trajectory analyses.

Wood, William A.↗

NASA NTRS1996

Analytic Corrections to CFD Heating Predictions Accounting for Changes in Surface Catalysis

A new approach for combining the insight afforded by integral boundary-layer analysis with comprehensive (but time intensive) computational fluid dynamic (CFD) flowfield solutions of the thin-layer Navier-Stokes equations is described. The approach extracts CFD derived quantities at the wall and at the boundary layer edge for inclusion in a post-processing boundary-layer analysis. It allows a designer at a work-station to address two questions, given a single CFD solution. (1) How much does the heating change for a thermal protection system (TPS) with different catalytic properties than was used in the original CFD solution? (2) How does the heating change at the interface of two different TPS materials with an abrupt change in catalytic efficiency? The answer to the second question is particularly important, because abrupt changes from low to high catalytic efficiency can lead to localized increase in heating which exceeds the usually conservative estimate provided by a fully catalytic wall assumption. Capabilities of this approach for application to Reusable Launch Vehicle (RLV) design are demonstrated. If the definition of surface catalysis is uncertain early in the design process, results show that fully catalytic wall boundary conditions provide the best baseline for CFD design points.

Gnoffo, Peter A.↗

NASA NTRS1995

Convective and radiative heating for vehicle return from the Moon and Mars

The aerothermal environment is examined for two vehicle forebodies near the peak heating points of lunar and martian return-to-earth trajectories at several nominal entry velocities. The first vehicle forebody is that of a 70 deg aerobrake for entry into earth orbit; the second, a capsule of Apollo configuration for direct entry into the earth's atmosphere. The configurations and trajectories are considered likely candidates for such missions. Two-temperature, thermochemical nonequilibrium models are used in the flow field analyses. In addition to Park's empirical model for dissociation under conditions of thermal nonequilibrium, the Gordiets kinetic model for the homonuclear dissociation of N2 and O2 is also considered. Temperature and emission profiles indicate nonequilibrium effects in a 2 to 5 cm post shock region. Substantial portions of the shock layer flow appear to be in equilibrium. The shock layer over an aerobrake for return from the moon exhibits the largest extent of nonequilibrium effects of all considered missions. Differences between the Gordiets and Parks kinetic model were generally very small for the lunar return aerobrake case, the greatest difference of 6.1 percent occurring in the radiative heating levels.

Greendyke, Robert B.↗

NASA NTRS1995

Effects of Sonic Line Transition on Aerothermodynamics of the Mars Pathfinder Probe

Flow field solutions over the Mars Pathfinder Probe spanning the trajectory through the Martian atmosphere at angles of attack from 0 to 11 degrees are obtained. Aerodynamic coefficients derived from these solutions reveal two regions where the derivative of pitching moment with respect to angle of attack is positive at small angles of attack. The behavior is associated with the transition of the sonic line location between the blunted nose and the windside shoulder of the 70 degree half-angle cone in a gas with a low effective ratio of specific heats. The transition first occurs as the shock layer gas chemistry evolves from highly nonequilibrium to near equilibrium, above approximately 6.5 km/s and 40 km altitude, causing the effective specific heat ratio to decrease. The transition next occurs in an equilibrium flow regime as velocities decrease through 3.5 km/s and the specific heat ratio increases again with decreasing enthalpy. The effects of the expansion over the shoulder into the wake are more strongly felt on the fustrum when the sonic line sits on the shoulder. The transition also produces a counter-intuitive trend in which windside heating levels decrease with increasing angle of attack resulting from an increase in the effective radius of curvature. Six-degree-of-freedom trajectory analyses utilizing the computed aerodynamic coefficients predict a moderate, 3 to 4 degree increase in total angle of attack as the probe, spinning at approximately 2 revolutions per minute, passes through these regions.

Gnoffo, Peter A.↗

NASA NTRS1994

Navier-Stokes simulations of Orbiter aerodynamic characteristics including pitch trim and bodyflap

An analysis of the longitudinal aerodynamics of the shuttle orbiter in the hypersonic flight regime is made through the use of computational fluid dynamics. Particular attention is given to establishing the cause of the 'pitching moment anomaly,' which occurred on the orbiter's first flight, and to computing the aerodynamics of a complete orbiter configuration at flight conditions. Data from ground-based facilities as well as orbiter flight data are used to validate the computed results. Analysis shows that the pitching moment anomaly is a real-gas chemistry effect that was not simulated in ground-based facilities, which used air as a test gas. Computed flight aerodynamics for the orbiter are within 5% of the measured flight values and trim bodyflap deflections are predicted to within 10%.

Weilmuenster, K. James↗

NASA NTRS1994

Multiblock analysis for Shuttle Orbiter reentry heating from Mach 24 to Mach 12

A multiblock, laminar heating analysis for the shuttle orbiter at three trajectory points ranging from Mach 24.3 to Mach 12.86 on reentry is described. The analysis is performed using the Langley Aerothermodynamic Upwind Relaxation Algorithm with a seven species chemical nonequilibrium model. A finite-catalytic-wall model appropriate for shuttle tiles at a radiative equilibrium wall temperature is applied. Computed heating levels are generally in good agreement with the flight data, although a few rather large discrepancies remain unexplained. The multiblock relaxation strategy partitions the flowfield into manageable blocks requiring a fraction of the computational resources (time and memory) required by a full domain approach. In fact, the computational cost for a solution at even a single trajectory point would be prohibitively expensive at the given resolution without the multiblock approach. Converged blocks are reassembled to enable a fully coupled converged solution over the entire vehicle, starting from a nearly converged initial condition.

LANGLEY AEROTHERMODYNAMIC UPWI↗