Frank S. Milos

Ablation and Thermal Response Program for Spacecraft Heatshield Analysis

An implicit ablation and thermal response program is presented for simulation of one-dimensional transient thermal energy transport in a multilayer stack of isotropic materials and structure which can ablate from a front surface and decompose in-depth. The governing equations and numerical procedures for solution are summarized. Solutions are compared with those of an existing code, the Aerotherm Charring Material Thermal Response and Ablation Program, and also with arcjet data Numerical experiments show that the new code is numerically more stable and solves a much wider range of problems compared with the older code. To demonstrate its capability, applications for thermal analysis and sizing of aeroshell heatshields for planetary missions, such as Stardust, Mars Microprobe (Deep Space n), Saturn Entry Probe, and Mars 2001, using advanced light-weight ceramic ablators developed at NASA Ames Research Center, are presented and discussed.

Thermal structure of Jupiter's atmosphere near the edge of a 5‐μm hot spot in the north equatorial belt

Thermal structure of the atmosphere of Jupiter was measured from 1029 km above to 133 km below the 1-bar level during entry and descent of the Galileo probe. The data confirm the hot exosphere observed by Voyager (∼900 K at 1 nanobar). The deep atmosphere, which reached 429 K at 22 bars, was close to dry adiabatic from 6 to 16 bars within an uncertainty ∼0.1 K/km. The upper atmosphere was dominated by gravity waves from the tropopause to the exosphere. Shorter waves were fully absorbed below 300 km, while longer wave amplitudes first grew, then were damped at the higher altitudes. A remarkably deep isothermal layer was found in the stratosphere from 90 to 290 km with T ∼ 160 K. Just above the tropopause at 260 mbar, there was a second isothermal region ∼25 km deep with T ∼ 112 K. Between 10 and 1000 mbar, the data substantially agree with Voyager radio occultations. The Voyager 1 equatorial occultation was similar in detail to the present sounding through the tropopause region. The Voyager IRIS average thermal structure in the north equatorial belt (NEB) approximates a smoothed fit to the present data between 0.03 and 400 mbar. Differences are partly a result of large differences in vertical resolution but may also reflect differences between a hot spot and the average NEB. At 15 4 bars, probe descent velocities derived from the data are consistently unsteady, suggesting the presence of large-scale turbulence or gravity waves. However, there was no evidence of turbulent temperature fluctuations >0.12 K. A conspicuous pause in the rate of decrease of descent velocity between 1.1 and 1.35 bars, where a disturbance was also detected by the two radio Doppler experiments, implies strong vertical flow in the cloud seen by the probe nephelometer. At p < 0.6 bar, measured temperatures were ∼3 K warmer than the dry adiabat, possible evidence of radiative warming. This could be associated with a tenuous cloud detected by the probe nephelometer above the 0.51 bar level. For an ammonia cloud to form at this level, the required abundance is ∼0.20 × solar.

Structure of the Atmosphere of Jupiter: Galileo Probe Measurements

Temperatures and pressures measured by the Galileo probe during parachute descent into Jupiters atmosphere essentially followed the dry adiabat between 0.41 and 24 bars, consistent with the absence of a deep water cloud and with the low water content found by the mass spectrometer. From 5 to 15 bars, lapse rates were slightly stable relative to the adiabat calculated for the observed H2/He ratio, which suggests that upward heat transport in that range is not attributable to simple radial convection. In the upper atmosphere, temperatures of >1000 kelvin at the 0.01-microbar level confirmed the hot exosphere that had been inferred from Voyager occultations. The thermal gradient increased sharply to 5 kelvin per kilometer at a reconstructed altitude of 350 kilometers, as was recently predicted. Densities at 1000 kilometers were 100 times those in the pre-encounter engineering model.

Galileo Probe Heat Shield Ablation Experiment

The Galileo probe deceleration module contained an experiment which measured the recession of the forebody heat shield during the ablative probe entry into the Jovian atmosphere. A detailed description of the experiment, reduction of the probe data, reconstruction of the heat-shield shape, and comparisons with preflight predictions are presented. Data quality is good during the second half of recession, but excessive noise levels at the onset of ablation prevented the experiment from meeting two of three performance requirements. The recession distribution is surprisingly dissimilar from preflight predictions. Measured recession was less than predicted near the stagnation point, but exceeded predictions over most of the frustum. (Author)

Ablation and Thermal Response Property Model Validation for Phenolic Impregnated Carbon Ablator

Phenolic Impregnated Carbon Ablator was the heatshield material for the Stardust probe and is also a candidate heatshield material for the Orion Crew Module. As part of the heatshield qualification for Orion, physical and thermal properties were measured for newly manufactured material, included emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates. Based on these properties, an ablation and thermal-response model was developed for temperatures up to 3500 K and pressures up to 100 kPa. The model includes transversely isotropic and pressure-dependent thermal conductivity. In this work, model validation is accomplished by comparison of predictions with data from many arcjet tests conducted over a range of stagnation heat flux and pressure from 107 W/cm 2 at 2.3 kPa to 1100 W/cm 2 at 84 kPa. Over the entire range of test

Mars Pathfinder Entry Temperature Data, Aerothermal Heating, and Heatshield Material Response

The Mars Pathe nder probe contained instrumentation that measured heatshield temperatures during entry. A description of the experiment, the data, and an analysis of the entry environment and material response are presented. Navier ‐Stokes forebody heating calculations show a peak unblown radiative-equilibrium heat e ux of 118W/cm 2 at thestagnation point and120 W /cm 2 on theshoulderforturbulente ow. Theheatload is3.8 kJ /cm 2 on thenose,decreases along thefrustum,then increasesto 2.7 ‐3.1kJ/cm 2 on theshoulder depending on the onset time forturbulence. One-dimensional charringmaterialresponseiscalculated using threedifferentmodels.Stagnationpoint temperature data are consistent with about 85% of fully catalytic laminar heating. Shoulder temperature data are inconclusive, but are consistent with fully catalytic laminar heating or with 85% of fully catalytic heating with early onset of turbulence. Aft temperature data indicate a peak heat e ux and heat load of about 1.3 W /cm2 and 70 J/cm 2 , respectively. The aft heating proe le is about 20 s longer than the forebody heating proe le. Bondline temperaturedata, although not useful forquantitative analysis of aerothermal heating, clearly showtheheatshield had adequate thickness margins for the actual entry.

Navier-Stokes Solutions with Finite Rate Ablation for Planetary Mission Earth Reentries

A formulation of finite rate ablation surface boundary conditions, including oxidation, nitridation, and sublimation of carbonaceous material with pyrolysis gas injection, based on surface species mass conservation, has been developed. These surface boundary conditions are discretized and integrated with a Navier-Stokes solver. This numerical procedure can predict aerothermal heating, chemical species concentration, and carbonaceous material ablation rates over the heat-shield surface of reentry space vehicles. Two finite rate gas-surface interaction models, based on the work of Park and of Zhluktov and Abe, are considered. Three test cases are studied. The stream conditions of these test cases are typical for Earth reentry from a planetary mission with both oxygen and nitrogen fully or partially dissociated inside the shock layer. Predictions from both gas-surface interaction models are compared with those obtained by using chemical equilibrium ablation tables. Stagnation point convective heat fluxes predicted by using Parks finite rate model are usually below those obtained from chemical equilibrium tables and Zhluktov and Abes model. Recession predictions from Zhluktov and Abes model are usually lower than those obtained from Parks model and from chemical equilibrium tables. The effect of species mass diffusion on the predicted ablation rate is also examined.

Computational Analysis of Arc-Jet Stagnation Tests Including Ablation and Shape Change

Coupled fluid-material response analyses of arc-jet stagnation tests conducted in a NASA Ames Research Center arc-jet facility are considered. The fluid analysis includes computational Navier-Stokes simulations of the nonequilibrium flowfield in the facility nozzle and test box as well as the flowfield over the models. The material response analysis includes simulation of two-dimensional surface ablation and internal heat conduction, thermal decomposition, and pyrolysis gas flow. For ablating test articles including shape change, the material response and fluid analyses are coupled to take into account changes in surface heat flux and pressure distributions with shape. The ablating material used in these arc-jet tests was a phenolic impregnated carbon ablator. Computational predictions of surface recession, shape change, and material response are compared with the experimental measurements.

Nonequilibrium Ablation of Phenolic Impregnated Carbon Ablator

Phenolic Impregnated Carbon Ablator is the forebody heatshield material for the Mars Science Laboratory, the Stardust sample-return capsule, and the SpaceX Dragon vehicle. In previous work an equilibrium ablation and thermal response model was developed. Model predictions were compared with data from many stagnation arcjet tests conducted over a range of stagnation heat flux and pressure from 107 W/cm 2 at 2.3 kPa to 1100 W/cm 2 at 84 kPa. In general, model predictions compared well with the data for surface recession, surface temperature, in-depth temperature at multiple thermocouples, and char depth. The uncertainty of the recession predictions was greatest for test conditions with low heat flux or low pressure. Additional testing has been performed at conditions down to 40 W/cm 2 and 1.6 kPa. The new test data suggest that nonequilibrium effects become important for prediction of PICA ablation at heat flux or pressure below about 80 W/cm 2 and 10 kPa, respectively. In this work we investigate two modifications to the ablation model to account for these nonequilibrium effects. Model predictions are compared with the arcjet test data.

The calculation of anisotropic extinction coefficients for radiation diffusion in rigid fibrous ceramic insulations

Abstract The specific anisotropic extinction coefficient ρ ∗ couples the effective radiative properties of a fibrous insulation into the radiation diffusion equation. This coefficient can be calculated using electromagnetic scattering theory if fiber diameters, refractive indices and fiber orientation distributions are known. In general, fiber orientation distributions are not readily accessible and past calculations have considered fibers as either randomly distributed or normal to the heat flow direction. In certain rigid fibrous ceramic insulations neither of these cases apply well, and a simple procedure is described for approximating ρ ∗ from values calculated for the random and normal orientation cases. The intrinsic error associated with this scaling procedure is investigated. Numerical computations for several test structures and fiber materials show the average error to be less than 5% for net heat flux and radiation conductivity calculations.