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Stewart, D. A.

Publications and source records attributed to Stewart, D. A..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Graphite Ablation and Thermal Response Simulation Under Arc-Jet Flow Conditions

The Two-dimensional Implicit Thermal Response and Ablation program, TITAN, was developed and integrated with a Navier-Stokes solver, GIANTS, for multidimensional ablation and shape change simulation of thermal protection systems in hypersonic flow environments. The governing equations in both codes are demoralized using the same finite-volume approximation with a general body-fitted coordinate system. Time-dependent solutions are achieved by an implicit time marching technique using Gauess-Siedel line relaxation with alternating sweeps. As the first part of a code validation study, this paper compares TITAN-GIANTS predictions with thermal response and recession data obtained from arc-jet tests recently conducted in the Interaction Heating Facility (IHF) at NASA Ames Research Center. The test models are graphite sphere-cones. Graphite was selected as a test material to minimize the uncertainties from material properties. Recession and thermal response data were obtained from two separate arc-jet test series. The first series was at a heat flux where graphite ablation is mainly due to sublimation, and the second series was at a relatively low heat flux where recession is the result of diffusion-controlled oxidation. Ablation and thermal response solutions for both sets of conditions, as calculated by TITAN-GIANTS, are presented and discussed in detail. Predicted shape change and temperature histories generally agree well with the data obtained from the arc-jet tests.

Chen, Y.-K.↗
NASA NTRS1993

Trajectory based, 3-dimensional heating and ablation calculations for the Apollo Lunar/Earth return capsule

Three dimensional (3-D) surface heating and ablation rate distributions have been obtained based on the geometry of the project Apollo Lunar/Earth return capsule and high Reynolds number inviscid/boundary layer methods. This application is based on the principles of the pseudo-three dimensional axisymmetric streamline analogy. Inviscid surface streamlines are determined for the forebody portion of the Apollo capsule using a CFD solver. Streamlines are obtained for the 20 deg angle-of-attack flight case, and the associated streamline metrics are determined. Boundary layer heating and ablation calculations are performed along selected Apollo forebody streamlines using the non-similar boundary layer code. Surface radiative heating estimates were obtained and the validity of the streamline tracing analogy is examined. These comparisons and the Apollo trajectory heating results are discussed.

Henline, W. D.↗
NASA NTRS1992

Navier-Stokes solutions with surface catalysis for Martian atmospheric entry

In this study numerical solutions have been obtained for two-dimensional axisymmetric hypersonic nonequilibrium CO2 flow over a high angle blunt cone with appropriate surface boundary conditions to account for energy and mass conservation at the body surface. The flowfield is described by the Navier-Stokes equations and multicomponent conservation laws which account for both translational and internal vibrational nonequilibrium effects. Complete forebody solutions have been obtained for the peak heating point of the Mars entry trajectory specified in the proposed NASA MESUR (Mars Environmental Survey) project. In these solutions, radiative equilibrium wall temperature and surface heating distributions are determined over the MESUR aeroshell forebody for entry velocity equal to 7 km/sec with varying degrees of surface catalysis. The effects of gas kinetics, surface catalysis, transport properties, and vibrational relaxation times on the surface heating are examined. The results identify some important issues in the prediction of surface heating for flows in thermochemical nonequilibrium and show that the Navier-Stokes code used herein is effective for thermal protection system design and materials selection.

Chen, Y.-K.↗
NASA NTRS1986

Heating distribution comparison between asymmetric and symmetric blunt cones

An experiment was performed to compare the heating distribution between symmetric and asymmetric large-angle blunt cones, with cone angles of 100, 120, and 140 deg. These hot-wall data were obtained from models made from typical thermal protection insulation for proposed aeroassisted orbital transfer vehicles. Experimental data are compared with predictions using a boundary-layer integral matrix procedure with kinetics to determine how well the heating distribution over an asymmetric cone could be approximated using axisymmetric solutions for a cone and spherical segment. In addition, a relationship between the stagnation-point heat-transfer rate and the bow-shock standoff distance for these cones is discussed. The heat-distribution data from the symmetric and asymmetric cones were very similar. Numerical results compared well with the measured wall temperatures at the stagnation point but slightly underpredicted them over the conical portion of the models.

Stewart, D. A.↗
NASA NTRS1985

Options for improving rigidized ceramic heatshields

Studies regarding capabilities, tensile strength, and dimensional stability of fibrous ceramics, used in current thermal protection systems and those required for future NASA advanced transfer vehicles, are presented. It is shown that the use of smaller diameter (2-4 micron) aluminoborosilicate fibers (instead of conventional 11-micron size) in the FRCI composites improves the homogeneity and tensile strength of high-silica composite. Substitution of the smaller aluminoborosilicate fibers by 2-4-micron alumina fibers in the AETB composites improves the dimensional stability, at the expense of a substantial increase in the thermal expansion coefficient, with a potential decrease in tensile strength.

Leiser, D. B.↗
NASA NTRS1985

Characterization of the thermal conductivity for fibrous refractory composite insulations

Thermal conductivity for a family of fibrous, refractory composite insulations are characterized through the use of an engineering model and arc-jet data. The model defines the composite as a uniform pore material using measured physical and mechanical properties. Efficiency parameters assigned to the solid and radiation conduction terms in the thermal conductivity equation were found to be dependent only on solid volume fraction and porosity.

Stewart, D. A.↗
NASA NTRS1985

Thermal response of integral multicomponent composite thermal protection systems

Integral-multicomponent thermal-protection materials are discussed in terms of their thermal response to an arc-jet airstream. In-depth temperature measurements are compared with predictions from a one-dimensional, finite-difference code using calculated thermal conductivity values derived from an engineering model. The effect of composition, as well as the optical properties of the bonding material between components, on thermal response is discussed. The performance of these integral-multicomponent composite materials is compared with baseline Space Shuttle insulation.

Stewart, D. A.↗
NASA NTRS1985

Heat-Shield Gap Filler

Ceramic cloth strips provide flexible, easily replaceable insulating filler. Filler prevents hot gas from flowing between heat-shield tiles while allowing space for thermal expansion and contraction. Strips easily replaced when necessary.

Leiser, D. B.↗
NASA NTRS1985

Effect of variable surface catalysis on heating near the stagnation point of a blunt body

This paper describes arc-jet data obtained on the performance of glass coated thermal protection systems in a convectively heated environment. These data confirm earlier flight and arc-jet data that show an increased surface catalysis with salt contamination and a decreased surface catalysis near the softening point temperature of the glass. In addition, surface temperature distributions along sphere-cones with abruptly changing surface catalysis were measured near the stagnation point and compared well with computations using a reacting boundary layer code.

Stewart, D. A.↗
NASA NTRS1984

Two-Layer Glass Thermal-Control Coating

Optical properties endure high temperatures. Coating has outer scattering layer and inner high emissivity layer. Absorptivity/emissivity ratio less than 0.4. Coating withstands repeated exposure to temperatures in excess of 2,000 degrees F. Coating has industrial uses in solar-energy equipment, high temperature chemical processing systems, laboratory equipment and high temperature instrumentation.

Stewart, D. A.↗
NASA NTRS1983

Catalytic surface effects on space thermal protection system during Earth entry of flights STS-2 through STS-5

An on going orbiter experiment catalytic surface effects experiment being conducted on the Space Shuttle is discussed. The catalytic surface effects experiment was peformed on four of the five flights of Columbia. Temperature time histories and distributions along the midfuselage and wing of the orbiter were used to determine the surface catalytic efficiency of the baseline high temperature reusable surface insulation. Correlation parameters are shown that allow the comparison of all flight data with predictions from the design and surface emittance decreased as a result of contaminants during the five flights of the Space Shuttle.

Stewart, D. A.↗
NASA NTRS1983

Thermal response of advanced high temperature ceramic-composite insulation to a convectively heated environment

The Space Shuttle uses currently fibrous high temperature, reusable surface insulation (HRSI) as a Thermal Protection System (TPS) at 1260 C for normal-design mission operation. Higher density material of a similar type extended the maximum temperature to 1480 C. However, a surface temperature capability exceeding 1540 C is desired for future space vehicles, such as the Orbital Transfer Vehicle (OTV). A TPS providing such a capability would permit future entry vehicles to enter the atmosphere at a much higher velocity than the Space Shuttle. The present investigation is concerned with arc-jet test data obtained from a family of fibrous refractory-composite insulation (FRCI) materials developed to meet the future TPS requirements. FRCI materials containing between 20 and 80 percent aluminoborosilicate fibers with the balance composed of silica fibers were studied.

Stewart, D. A.↗
NASA NTRS1983

Thermal and mechanical properties of advanced, high temperature ceramic-composite insulation

It is found that the mechanical properties of fibrous refractory composite insulation (FRCI) can be improved if a higher calcination temperature is used before final processing. The results also reveal that a higher density FRCI containing 60 wt pct aluminoborosilicate fibers will exhibit minimal surface recession at 1480 C in a convective-heating environment. Another finding is that the material performance in the convective-heating environment is limited by the coating and the temperature capability of its emittance agent. A table is included giving the coating compositions used with the composite insulation and tested in the convectively heated environment at a surface temperature of 1480 C.

Leiser, D. B.↗
NASA NTRS1983

High temperature glass thermal control structure and coating

A high temperature stable and solar radiation stable thermal control coating is described which is useful either as such, applied directly to a member to be protected, or applied as a coating on a re-usable surface insulation (RSI). It has a base coat layer and an overlay glass layer. The base coat layer has a high emittance, and the overlay layer is formed from discrete, but sintered together glass particles to give the overlay layer a high scattering coefficient. The resulting two-layer space and thermal control coating has an absorptivity-to-emissivity ratio of less than or equal to 0.4 at room temperature, with an emittance of 0.8 at 1200 F. It is capable of exposure to either solar radiation or temperatures as high as 2000 F without significant degradation. When used as a coating on a silica substrate to give an RSI structure, the coatings of this invention show significantly less reduction in emittance after long term convective heating and less residual strain than prior art coatings for RSI structures.

Stewart, D. A.↗
NASA NTRS1983

Performance of an ablator for Space Shuttle inorbit repair in an arc-plasma airstream

An ablator patch material performed well in an arc plasma environment simulating nominal Earth entry conditions for the Space Shuttle. Ablation tests using vacuum molded cones provided data to optimize the formulation of a two part polymer system for application under space conditions. The blunt cones were made using a Teflon mold and a state of the art caulking gun. Char stability of formulations with various amounts of catalyst and diluent were investigated. The char was found to be unstable in formulations with low amounts of catalyst and high amounts of diluent. The best polymer system determined by these tests was evaluated using a half tile patch in a multiple High Temperature Reusable surface Insulation tile model. It was demonstrated that this ablator could be applied in a space environment using a state of the art caulking gun, would maintain the outer mold line of the thermal protection system during entry, and would keep the bond line temperature at the aluminum tile interface below the design limit.

Stewart, D. A.↗