Spacecraft Heat Shields Research Articles

Computation of effective thermal conductivity of carbon fiber felts through numerical simulation and development of reduced order models

One of the major drivers of heat transfer within a spacecraft heatshield is the effective thermal conductivity of the thermal protection material. For carbon-based thermal protection materials, this quantity is strongly connected to the thermal conductivity of the carbon fiber felt (preform) used to create the final composite. Use of effective thermal conductivity data from experiments in volume averaged numerical models is complex because: i) they are carried out in different media than its actual application, ii) changes on geometry or material properties require repetition of the experiments. Simulation of effective thermal conductivity using high fidelity simulations can overcome those problems. In this work, we propose a synthetic model based of the carbon fiber felt of Calcarb® using Porous Microstructure Analysis (PuMA) software. This model is first validated with available experimental data and applied to new experiments and to study the effect of pyrolysis gas mixture on the effective property. Simplified correlations are then developed, which predict with remarkable accuracy the effective thermal conductivity of the Calcarb® material. The developed model can be used to explore and predict effective conductivity at different conditions. In addition, the simplified algebraic correlations can be easily implemented in volume averaged material response codes to account in a more appropriate way for the effect of the percolating pyrolysis/atmospheric gas.

Arc-Jet Tests of Carbon-Phenolic-Based Ablative Materials for Spacecraft Heat Shield Applications.

We developed and tested two carbon-phenolic-based ablators for future Korean spacecraft heat shield applications. The ablators are developed with two layers: an outer recession layer, fabricated from carbon-phenolic material, and an inner insulating layer, fabricated either from cork or silica-phenolic material. The ablator specimens were tested in a 0.4 MW supersonic arc-jet plasma wind tunnel at heat flux conditions ranging from 6.25 MW/m2 to 9.4 MW/m2, with either specimen being stationary or transient. Stationary tests were conducted for 50 s each as a preliminary investigation, and the transient tests were conducted for ~110 s each to stimulate a spacecraft's atmospheric re-entry heat flux trajectory. During the tests, each specimen's internal temperatures were measured at three locations: 25 mm, 35 mm, and 45 mm from the specimen stagnation point. During the stationary tests, a two-color pyrometer was used to measure specimen stagnation-point temperatures. During the preliminary stationary tests, the silica-phenolic-insulated specimen's reaction was normal compared to the cork-insulated specimen; hence, only the silica-phenolic-insulated specimens were further subjected to the transient tests. During the transient tests, the silica-phenolic-insulated specimens were stable, and the internal temperatures were lower than 450 K (~180 °C), achieving the main objective of this study.

An optimal heat-flux targeting procedure for LEO re-entry of reusable vehicles

This paper addresses the feasibility analysis of a re-entry flight strategy from low Earth orbit, developed for an innovative re-usable wing-body concept, having a rather high surface-to-mass ratio. The re-entry procedure envisages a longer and gradual dissipation of vehicle's total energy, carried out at a flight path angle lower than typical Shuttle-like values. Hence, the heat flux is kept constant at a predicted threshold, thus allowing the radiative cooling of the spacecraft heatshield at a suitable temperature. The selected flight approach relies on a prescribed guidance law which modulates the glider angle of attack as function of Mach number, to assure the desired convective heat flux level for a certain time. Based on a low-flight path assumption, analytical considerations are derived to evaluate geometric, aerodynamic, and flight parameters that influence the convective heat flux. Furthermore, candidate guidance laws are determined formulating an optimization-based heat flux targeting procedure. The relevance of aerodynamic performances is also underlined, provided that, by results high lift configurations dictate lower angles of attack and vice versa. Analytical and numerical results evidence existing relation between surface-to-mass ratio, lift-to-drag ratio and flight duration time, that can be used “on design” to modulate the aerothermal loading. The current approach is suitable to explore less challenging aerothermal environments in order to improve reliability and affordability of space operations in the near Earth space.

Mixed Mode Crack Growth in Functionally Graded Material Under Three-Point Bending

Functionally graded materials, FGM, permit tailoring the material volume fractions to extract maximum benefit from their in-homogeneity. Such materials offer great potential for components, which operate under severe thermal or mechanical loadings, such as spacecraft heat shields, plasma facings for fusion reactors, crucial jet fighter structures and engine components. In the present work, polyester matrix reinforced by varying glass fiber distribution was used to obtain FGM polymeric composites. The hand lay-up technique was adopted to manufacture the FGM. The mixed mode crack path of glass fiber/polyester FGM under three point bend, 3PB, was observed. The mechanical properties, such as tensile and flexural strength were determined experimentally and numerically. Pre-cracks with different lengths and locations with seven different inclination angles were considered to study the mixed mode crack growth of FGM. For the same crack length, the present results showed that the maximum applied load increased with increasing the crack inclination angle, i.e. the mode of mixity increased to reach pure mode II. Generally, the crack path initiated from the original inclination angle and grew along a curvilinear trajectory to extend toward the upper loading point.

Fiber-optic temperature profiling for thermal protection system heat shields

To achieve better designs for spacecraft heat shields for missions requiring atmospheric aero-capture or entry/reentry, reliable thermal protection system (TPS) sensors are needed. Such sensors will provide both risk reduction and heat-shield mass minimization, which will facilitate more missions and enable increased payloads and returns. This paper discusses TPS thermal measurements provided by a temperature monitoring system involving lightweight, electromagnetic interference-immune, high-temperature resistant fiber Bragg grating (FBG) sensors with a thermal mass near that of TPS materials together with fast FBG sensor interrogation. Such fiber-optic sensing technology is highly sensitive and accurate, as well as suitable for high-volume production. Multiple sensing FBGs can be fabricated as arrays on a single fiber for simplified design and reduced cost. Experimental results are provided to demonstrate the temperature monitoring system using multisensor FBG arrays embedded in a small-size super-light ablator (SLA) coupon which was thermally loaded to temperatures in the vicinity of the SLA charring temperature. In addition, a high-temperature FBG array was fabricated and tested for 1000°C operation, and the temperature dependence considered over the full range (cryogenic to high temperature) for which silica fiber FBGs have been subjected.

L1 Regularization-Based Model Reduction of Complex Chemistry Molecular Dynamics for Statistical Learning of Kinetic Monte Carlo Models

ABSTRACTKinetic Monte Carlo (KMC) methods have been a successful technique for accelerating time scales and increasing system sizes beyond those achievable with fully atomistic simulations. However, a requirement for its success is a priori knowledge of all relevant reaction pathways and their rate coefficients. This can be difficult for systems with complex chemistry, such as shock-compressed materials at high temperatures and pressures or phenolic spacecraft heat shields undergoing pyrolysis, which can consist of hundreds of molecular species and thousands of distinct reactions. In this work, we develop a method for first estimating a KMC model composed of elementary reactions and rate coefficients by using large datasets derived from a few molecular dynamics (MD) simulations of shock compressed liquid methane, and then using L1 regularization to reduce the estimated chemical reaction network. We find that the full network of 2613 reactions can be reduced by 89% while incurring approximately 9% error in the dominant species (CH4) population. We find that the degree of sparsity achievable decreases when similar accuracy is required for additional populations of species.

Effects of thermal boundary conditions on the hypersonic shock layer flow over a blunt body

This paper presents a numerical study of the effects of thermal boundary conditions on the shock layer flow over a spacecraft entering a planetary atmosphere at hyper-velocity, and on the resulting convective heat fluxes to its heat shield surface. This study is motivated by the fact that spacecraft heat shields are covered with materials that ablate to reduce the spacecraft wall temperature. The ablation is not taken into account, but the effects of changes in the temperature of the vehicle surface are thoroughly studied. The results show that the flow field solutions within the shock layer and the convective heat fluxes are sensitive to the wall thermal boundary conditions. A computation of the flow field solution for an adiabatic wall is also performed. In this case, the gas recovery temperature along the wall increases as the gas moves away from the stagnation point. Computation convergence is also analysed.

Effect of Spacecraft Aerodynamics and Heat Shield Characteristics on Optimal Aeroassisted Transfer

A spacecraft designed to operate in a planetary atmosphere must have an adequate heat shield to withstand the high heat fluxes and heat loads that are generated by aerodynamic heating. Very often, the mass of the thermal protection system is a significant fraction of the total mass of the vehicle. In contrast, performing maneuvers in the atmosphere, that would be very costly in terms of propellant consumption if they were performed completely outside of the atmosphere in a classic way, is a very attractive prospective technique. The advantages and disadvantages in terms of total mass spared must be determined. The mission investigated involves an aeroassisted coplanar transfer from a high to a low Earth orbit. The approach uses a combination of three propulsive impulses in space together with an aerodynamic maneuver in the atmosphere. The heat shield adopted is fully ablative, given the expected high values of the entering heat flux. The convenience of the aeroassisted maneuver and the influence of the parameters involved are evaluated in comparison to a conventional Hohmann transfer. In particular, a parametric analysis is performed by varying the following characteristics of the vehicle: aerodynamic efficiency, mass-to-surface ratio, deorbit impulse, and initial altitude of the orbit. The influence of the thermal protection system is examined by assessing the impact of the type of ablative material employed, the thermal safety factor, and the allowable temperature for the adhesive layer on the substructure. The analysis is conducted with a highly representative thermal model by coupling the dynamic and thermal analyses and using a genetic optimizer. The optimization methodology and the thermal model are completely original. The results indicate the importance of choosing low-density ablative materials, of adopting a suitable thermal safety factor, and of choosing high-performance adhesives. The optimal trajectories obtained correspond to a zero second propulsive impulse.

Meteorite fusion crust variability

Abstract— Two assumptions commonly employed in meteorite interpretation are that fusion crust compositions represent the bulk‐rock chemistry of the interior meteorite and that the vesicles within the fusion crust result from the release of implanted solar wind volatiles. Electron microprobe analyses of thin sections from lunar meteorite Miller Range (MIL) 05035 and eucrite Bates Nunataks (BTN) 00300 were performed to determine if the chemical compositions of the fusion crust varied and/or represented the published bulk rock composition.It was determined that fusion crust compositions are significantly influenced by the incorporation of fragments from the substrate, and by the composition and grain size of those minerals. Because of compositional heterogeneities throughout the meteorite, one cannot assume that fusion crust composition represents the bulk rock composition. If the compositional variability within the fusion crust and mineralogical differences among thin sections goes unnoticed, then the perceived composition and petrogenetic models of formation will be incorrect.The formation of vesicles within these fusion crusts were also compared to current theories attributing vesicles to a solar wind origin. Previous work from the STONE‐5 experiment, where terrestrial rocks were exposed on the exterior of a spacecraft heatshield, produced a vesicular fusion crust without prolonged exposure to solar wind suggesting that the high temperatures experienced by a meteorite during passage through the Earth's atmosphere are sufficient to cause boiling of the melt. Therefore, the assumption that all vesicles found within a fusion crust are due to the release of implanted volatiles of solar wind may not be justified.

Simulation of heterogeneous atom recombination on spacecraft heat shield coatings using the methods of molecular dynamics

An efficient method of investigating the processes of interaction between gas mixtures and catalytic surfaces is developed within the framework of classical molecular dynamics. The recombination and chemical-energy accommodation coefficients on the catalytic surface can be determined with fewer computational resources than in the quantum-mechanical and semiclassical approaches. Oxygen atom recombination on a β-cristobalite surface of the type frequently used in spacecraft heat shield systems is investigated. The probability of atom recombination and the recombination energy accommodation coefficient obtained are in satisfactory agreement with the available experimental data and calculations made by means of the semiclassical method. The hypothesis that the probability of the Eley-Rideal reaction decreases and the probability of atom adsorption increases with increase in the atom collision energy is confirmed. It is attributable to the tendency of atoms to be trapped in the potential well and be desorbed in the atomic state when the surface collision energy is high instead of entering into a recombination reaction and then being desorbed in the molecular state.

Thermal Response and Ablation Characteristics of Carbon Fiber Reinforced Composite with Novel Silicon Containing Polymer MSP

Polymer composites are widely used for high temperature thermal protection (TPS)materials for spacecraft heat shields, and nozzle liners for solid propellant rocket motors. Newly developed silicon containing polymer, abbreviated MSP (Mitsui Silicon Containing Polymer), possesses high decomposition temperature (500 C), and high char yield (>94% at 1000 C)after pyrolysis. Therefore, the MSP polymer has potential as matrix resin of high performance ablative composites compared with conventional phenolic resins. In this study, thermal performance and ablation characteristics of a carbon composite with the MSP polymer (CF/MSP) were investigated under supersonic plasma air stream. The cold wall heat fluxes ranged from 1.2 to 11.0MW/m2. Conventional carbon fiber phenolic (CF/PR) composite was tested at the similar conditions for direct comparison. The mass loss of the CF/MSP composite was far less than that of the CF/PR composite, which is due to higher char yield and consequent high density of char layer. On the other hand, higher internal temperature was measured in the CF/MSP composite.

Reviving the Mars Polar Lander

Although the real‐life Mars Polar Lander was lost shortly after entry into the Martian atmosphere, children above age 10, or adults,can build a detailed, l/24th scale, paper model of the Polar Lander, thanks to the Jet Propulsion Laboratory.The scale model includes the spacecraft's heat shield, lander, microprobes, and other instrumentation.

Experimental Investigation of Surface Reactions in Carbon Monoxide and Oxygen Mixtures

During hypersonic entry into the CO 2 atmosphere of Mars, competing exothermic chemical reactions may oceur on a spacecraft heatshield surface. Two possible surface reactions are O+O → O 2 and CO+O → CO 2 . The relative importance of these reactions on quartz is investigated using a diffusion tube side-arm reactor together with two-photon laser-induced fluorescence for both O and CO species detection. The experiments show 1) that the presence of CO in the gas phase does not -significantly affect the oxygen recombination reaction on quartz and 2) that the gas-phase CO concentration is not significantly altered by the presence of atomic oxygen. These results indicate that for our experimental conditions the dominant surface reaction on quartz in oxygen-carbon monoxide mixtures is O+O → O 2 . Current heating computations for Martian entries assume CO oxidation to be fully catalytic. The resulting entry heating values are significantly higher than those computed using the assumption of fully catalytic oxygen recombination. The data presented here indicate that the assumption of fully catalytic CO oxidation may be overly conservative for heatshleld sizing purposes

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.

Structural analysis of axisymmetric solids.

N the aerospace industry, the stress analysis of complex axisymmetric structures of arbitrary shape subjected to thermal and mechanical loads is of considerable interest. Rocket nozzles and cases, solid-propellant grains, and spacecraft heat shields are practical examples of such structures. Although the governing differential equations for solids of revolution have been known for many years, closed form solutions have been obtained for only a limited number of structures; thus, the stress analyst must rely on experimental or numerical techniques. The finite difference method has been the most popular of the numerical techniques; however, for structures of composite materials and of arbitrary geometry, this procedure is difficult to apply. In the present investigation, the finite element idealization is used as the basic numerical procedure. This technique has been applied successfully in the stress analysis of many complex structures.13 In Ref. 4 impressive results were obtained in the analysis of axisymmetric shells approximated by a series of truncated cone elements. The approach, which is presented here, is similar in many respects to existing methods used in the analysis of two-dimensional stress problems.57 Recently, the finite element method was applied to the structural analysis of axisymmetric solids subjected to axisymmetric loads.8 In the present paper, the finite element method is used in the determination of stresses and displacements developed within elastic solids of revolution which are subjected to axisymmetric or nonaxisymmetr ic loads. Emphasis is placed on the application of the technique to complex aerospace structures.