Spacecraft Thermal Protection Systems Research Articles

Laser cleaning for inorganic thermal control coatings on aluminum alloys: Microstructure evolution and mechanical properties of substrate

Thermal control coatings are crucial materials in spacecraft thermal protection systems. Investigating efficient, convenient, and environmentally friendly removal techniques for thermal control coatings is essential for their application and development. This study developed a new removal technology for thermal control coatings using nanosecond pulse laser cleaning technology. The removal effect of the thermal control coating was studied, and the microstructure evolution and mechanical properties of the substrate were analyzed. The results showed that the coating was completely removed at 5.73 and 7.00 J/cm2, while at 8.28 J/cm2, in addition to the coating, approximately 5±1.1 μm of the substrate was also removed. In the completely cleaned state, micro-melting occurred on the substrate surface, and nanostructures appeared. At 5.73 and 7.00 J/cm2, the thickness of the remelted layer is 4±0.9 μm and 5.5±0.6 μm, respectively. The grains in the remelted layer exhibit significant refinement, accompanied by the observation of numerous subgrain boundaries. Additionally, the substrate surface exhibits tensile residual stress after laser cleaning. Attributed to the influence of the remelted layer, the surface hardness and tensile properties of the substrate are slightly improved after laser cleaning compared to manual grinding.

Anisotropic and hierarchical biphasic nanorod ceramic aerogels with thermal radiation shielded, high strength and exceptional stability for thermal superinsulation

The utilization of ceramic aerogels is crucial in spacecraft thermal protection systems that necessitate exceptional high-temperature stability, mechanical robustness, and low thermal conductivity. The primary challenges hindering their implementation encompass inadequate mechanical robustness and high infrared radiation transparency. Herein, a biphasic nanorod encapsulation strategy is proposed to fabricate an Al2O3-TiO2 nanorod aerogel (ATNA) with a hierarchical and anisotropic microstructure. The biphasic nanorod structure not only reduces the radiation heat transfer but also enhances the mechanical properties simultaneously. The primary compressive strain was borne by the intersecting Al2O3 nanorods, while the interpenetrated TiO2 nanorods served to reinforce and withstand partial stress at the joints. The ATNAs obtained exhibit an ultralow thermal conductivity (0.069 W·m−1·K−1 at 1100 °C), exceptional high stiffness (a specific compressive strength of 13.4 kN·m·kg−1), and remarkable thermal stability (1400 °C in air). The robust ATNAs can effectively meet the urgent demand for thermal insulation in high-temperature environments.

Predicting high‐speed particle impact damage in spacecraft thermal protection systems

All spacecraft use some sort of thermal insulation, or thermal protection system (TPS), in their design. TPS materials vary, ranging from ceramic tiles or phenolic ablators for heatshields to lightweight multi-layer insulation (MLI) blankets. Since TPS is usually placed on the spacecraft's exterior, it is susceptible to impacts by meteoroids and orbital debris. These high-speed impacts can damage the TPS to a point where the protection it offers is below acceptable limits. As such, it is important to be able to characterize expected TPS damage levels stemming from such high-speed impacts. In this paper, we present the results of a study that sought to characterize the high-speed impact damage that would be sustained by two TPS materials that have recently gained attention for possible use in future interplanetary missions. Empirical equations were developed for TPS crater depths, as well as maximum and minimum crater mouth dimensions. In the event of TPS perforations, empirical equations were developed for the maximum and minimum through-hole dimensions. As part of the analyses performed, ballistic limit equations (BLEs) for these TPS configurations were also developed where possible. The validity of the equations developed was assessed by comparing their predictions against test data. In nearly all cases, the empirical equations developed herein were seen to adequately capture the magnitudes of the measured damage parameters.

Estimation of the Thermophysical Properties of C/C Plates with Ceramic Nanocoating at Different SiO2 Filling for Aerospace Applications

In aerospace industrial and commercial scenario, the reusable launch vehicles (RLV) evolution works constantly toward the lowering of payload conveyance expenses. The thermal protection system (TPS) preserves the integrity of the space vehicle surfaces exposed to huge thermal shock during the re-entry phase: its advanced design and manufacturing, aimed at both reusing and withstanding harsh space environment, result in increasing the production and maintenance charges. The present study introduces a cost-saving concept of TPS component made of carbon/carbon (C/C) tiles coated by a commercial refractory varnish reinforced with ceramic nanoparticles. Using a reliable computing method, known as inverse method, the thermophysical properties such as heat capacity and thermal conductivity of the manufactured materials are assessed in a broad range of temperatures, with the input aid of an in-house developed experimental setup. The described technique is especially suited for approaching such kind of issues, thanks to the capability of taking into account several physical variables simultaneously, with the aim of gaining a robust knowledge of materials’ thermal behavior for potential use in spacecraft TPS.

Carbon layer encapsulation strategy for designing multifunctional core-shell nanorod aerogels as high-temperature thermal superinsulators

Aerogels have been considered as attractive candidates for spacecraft thermal protection systems. However, constructing lightweight aerogels with better mechanical strength, higher temperature resistance and lower high-temperature thermal conductivity, whether based on nanoparticles or nanofibers, is still a great challenge. Moreover, to avoid performance degradation caused by moisture absorption, insulating aerogels usually suffer from complex post-processing to obtain superhydrophobicity, which also cannot be guaranteed once the surface breaks down. Herein, a carbon layer encapsulation (CLE) strategy is proposed to resolve the above-mentioned conundrums in a simple way. Thanks to the collaboration of structural design and theoretical simulations, the tailored Al2O3-carbon core–shell nanorod aerogels demonstrate excellent comprehensive properties of low density (as low as 0.086 g·cm−3), outstanding stiffness (a specific compressive strength of 69.83 kN·m·kg−1), bionic abrasion-durable superhydrophobicity (WCA 156° after 1000 abrasion cycles), ultra-high thermal stability (over 1500 °C in argon and over 1400 °C in air) and high-temperature thermal superinsulating performance (0.065 W·m−1·K−1 at 1200 ℃). The synergy of ultrafine Al2O3 nanorods and carbon layers with suitable thickness not only forms a robust lotus leaf-like structure, but also enables the obtained aerogels to exhibit much superior thermal insulation properties than reported Al2O3-based aerogels. The significant increase in temperature resistance induced by lattice distortion is also an interesting phenomenon that has been investigated in detail. This novel strategy provides a fresh perspective for the preparation of multifunctional thermal high-temperature superinsulators applicable to spacecraft thermal protection systems.

Thermo-mechanical characterization of carbon-based ceramic foams for high temperature space application

The successful development of aerospace reusable launch vehicles (RLV) require to realize effective thermal protection systems (TPS) for preserving spacecraft integrity from the severe thermal loads during re-entry phase. To such an aim, due to the need of reducing payload transportation costs, applied research is driven towards lightweight materials with advanced thermo-mechanical properties. Space TPS are often based on sandwich structures, where the core material has the main function of thermal insulation. Ceramic porous materials, as carbon (C) and silicon carbide (SiC) foams, represent ideal candidates for application as structural TPS component, thanks to both low density and significant thermal stability at very high temperatures. The paper presents a joint experimental study of carbon-based ceramic foams proposed as sandwich’s core for TPS design. A full thermal characterization of commercial C- and SiC-foam materials is reported, including measurements of thermo-mechanical combined stress, temperature-induced outgassing behavior and heat transfer properties. These latter, in particular, are studied by means of a robust numerical technique, known as the inverse method, which allows to evaluate materials thermal conductivity and heat capacity over a wide range of temperatures, thus establishing the required material behavior for potential use in spacecraft TPS.

Plasma simulations reveal important parameters affecting inductive discharges

Spacecraft thermal protection systems can be studied by inductively coupled plasma wind tunnels, allowing researchers to understand how plasma parameters affect heat resistance.

Double-layer gradient MoSi2-borosilicate glass coating prepared by in-situ reaction method with high contact damage resistance and interface bonding strength

High emissivity coating is an important part of spacecraft thermal protection system. MoSi2-borosilicate glass coatings prepared by the in-situ reaction method showed excellent thermal endurance and oxidation resistance. However, radial cracks were prone to appear under external forces due to high cohesive of in-situ reaction dense coatings, and few studies focused on the mechanical properties of gradient coatings. In this study, a double-layer gradient MoSi2-borosilicate glass coating with a porous interface layer and a dense surface layer was prepared on mullite fibrous insulation tiles by in-situ reaction method. The double-layer gradient coating was compared with a mono-layer dense coating and a high-density gradient coating (another gradient coating with higher interface layer density) with reference to microstructure, contact damage resistance, interface bonding strength and thermal shock resistance. Compared with mono-layer dense coating and high-density gradient coating, double-layer gradient coating showed higher bearing capacity and avoided radial crack initiation and propagation in Hertzian indentation test. Besides, the double-layer gradient coating endured the most thermal shock cycles between 1500 °C and 20 °C and presented substrate fracture instead of interface fracture. In addition, double-layer gradient coating and high-density gradient coating possessed much higher interface bonding strengths than mono-layer dense coating. The total emissivities of three coatings were all greater than 0.87. The porous interface in double-layer gradient coating presented the advantages in enhanced mechanical and thermal compatibility between the in-situ reaction dense coating and substrate through reducing elastic modulus and thermal expansion coefficient differences.

NO production on the reentry spacecraft thermal protection system surface in the direct simulation Monte Carlo method

This work is aimed at development of an approach for modeling surface chemistry with multiple reaction channels in the direct simulation Monte Carlo method. In particular, production of NO on spacecraft thermal protections system surface is considered in detail. The approach is based on the previous works of the authors where a surface chemistry molecular model based on macroscopic detailed reaction sets was developed. The new approach has been implemented in SMILE++ code and verified by comparison with theoretical values of the surface population of atomic species in thermal equilibrium. Excellent agreement between computational results and theoretical prediction has been demonstrated.

One-Dimensional Thermal Analysis Model for Charring Ablative Materials

This paper presents a one-dimensional model for the analysis of the charring ablative materials used in spacecraft thermal protection systems. The numerical method is based on an implicit finite difference formulation of the governing equations written for a system of mobile coordinates that accounts for the possible presence of surface recession. The maximum allowable operating temperature for the adhesive layer of the junction between the heat shield and the substructure is used as a design parameter for determining the minimum heat shield thickness. A case study on the re-entry of the Stardust capsule is presented. The model proposed as a useful dimensioning tool for the preliminary design phase of the heat shields of spacecraft entering the atmosphere. The model was validated through a survey of the literature related to the dimensioning of thermal shields but based on numeric programs of highly representative industrial standards.

Hypersonic Aerothermochemistry Duplication in Ground Plasma Facilities: A Flight-to-Ground Approach

High conservative safety margins, applied to the design of spacecraft thermal protection systems for planetary entry, need to be reduced for higher efficiency of future space missions. Ground testing of such protection systems is of great importance during the design phase. This study covers a methodology for simulating the complex hypersonic entry aerothermochemistry in a plasma wind tunnel for a given spacecraft geometry without any assumption on axisymmetry or bluntness. A demonstration of this proposed methodology is made on the Qubesat for Aerothermodynamic Research and Measurements on AblatioN, QARMAN mission, which is a rectangular reentry CubeSat with a cork-based ablative thermal protection system in the front unit. The reacting boundary-layer profiles of the hypersonic entry probe compare well with the ones developing at the stagnation region of the plasma test model, defined with the proposed flight-to-ground duplication method.

Simulation of Orbital Debris Impact on Porous Ceramic Tiles

Ceramic tiles provide lightweight, reusable insulation for spacecraft thermal protection systems but are vulnerable to orbital debris impact damage. Because a wide range of impact conditions that may be experienced in orbit cannot be duplicated in the laboratory, design for orbital debris impact effects must include analytical or numerical modeling. Recent research has developed and validated a new computational approach to the simulation of orbital debris impact on porous ceramic thermal protection systems. Simulations using the validated formulation suggest scaling rules for application in spacecraft design.

Spectrally and Spatially Resolved Radiance Measurement in High-Speed Shock Waves for Planetary Entry

Three-dimensional images of absolute radiance as a function of position and wavelength are obtained for shock-heated plasmas traveling in a shock tube at velocities near Mach 30. Vacuum optics coupled to a shock tube are used to image the shock onto four separate spectrometers, which, in turn, disperse the radiation in wavelength onto intensified CCD arrays covering selected spectral ranges from the vacuum ultraviolet (>;120 nm) through midwave infrared (<;5000 nm). Quantitative radiometry performed in this fashion is used to benchmark radiative heating codes used in sizing spacecraft thermal protection systems.

A probabilistic sizing tool and Monte Carlo analysis for entry vehicle ablative thermal protection systems

Implicit ablation and thermal response software was developed to analyse and size charring ablative thermal protection systems for entry vehicles. A statistical monitor integrated into the tool, which uses the Monte Carlo technique, allows a simulation to run over stochastic series. This performs an uncertainty and sensitivity analysis, which estimates the probability of maintaining the temperature of the underlying material within specified requirements. This approach and the associated software are primarily helpful during the preliminary design phases of spacecraft thermal protection systems. They are proposed as an alternative to traditional approaches, such as the Root-Sum-Square method. The developed tool was verified by comparing the results with those from previous work on thermal protection system probabilistic sizing methodologies, which are based on an industry standard high-fidelity ablation and thermal response program. New case studies were analysed to establish thickness margins on sizing heat shields that are currently proposed for vehicles using rigid aeroshells for future aerocapture missions at Neptune, and identifying the major sources of uncertainty in the material response.

Calculation of High-Enthalpy Aerothermal Environment in an Arcjet Facility

S EGMENTED-CONSTRICTOR type arc-heated wind tunnels are used to test the heat shield materials for spacecraft thermal protection systems. The arc-heated wind tunnel consists of an upstream electrode (anode) chamber, constrictor section, downstream electrode (cathode) chamber, and a diverging–converging nozzle connecting to a test chamber. In the test section, the heat shield materials are exposed to a high-enthalpy flow environment produced by the facility. The high-enthalpy environment is often such that the flowdoes not reach equilibrium condition at the edge of the boundary layer over the tested material. In such a case, we need to calculate the flow properties at the surface of the material using a computational fluid dynamics approach to understand the thermal response of the material [1]. For this purpose, the arcjet freestream conditions must be known accurately. To calculate an arcjet freestream condition, two important physical processes occurring in the arcjet wind tunnel should be accounted for: the heating process in the arc heater region upstream of the nozzle throat and the relaxing process in the expanding nozzle region downstream of the nozzle throat. The ARCFLO3 code has been developed recently to calculate the flowfield in the segmentedconstrictor type of arc heaters [2]. Unlike the arc heater flowfield code named ARCFLO developed in the 1970s [3], which is able to calculate the flow in the constrictor section, this new code calculates the flow from the anode chamber to the nozzle throat [2]. Arcjet freestream conditions can be calculated fully theoretically if a nonequilibrium expanding nozzle calculation is made with the calculated flow properties at the nozzle throat obtained by using the ARCFLO3 code. In addition, because the radial distribution of the flow properties at the nozzle throat is calculated with the ARCFLO3 code, the unified computational method can give the radial flow properties in the arcjet freestream at the test section. We tried to make such a unified calculation very recently for one operating condition in an arcjet facility [1]. However, the question remains as to how well such a computational approach predicts the flow properties in an arcjet freestream. It is the purpose of the present work to test the validity of the unified method. The method is applied to calculate the flowfield in a 0.75-MW arcjet wind-tunnel facility at the Institute of Aerospace Technology of the Japan Aerospace Exploration Agency (IAT/JAXA) in Japan. This facility was chosen for the following reasons: 1) In the recent measurement [4] in the IAT/JAXA arcjet facility, the operational characteristic parameters for a wide range of conditions were obtained. The experimental data offer an opportunity to test the current state of the computational modeling in the proposed method. 2) In our previous work, the ARCFLO3 code was applied for the arc heater flowfield calculation only in a high-power-level arc heater, such as the 20or 60-MW arcjet facility at NASA Ames Research Center [2]. The applicability of the ARCFLO3 code to submegawatt class facilities is unknown.

Gas Permeability of Rigid Fibrous Refractory Insulations

Rigid fibrous refractory insulations (TPS tiles) are integral components of many spacecraft thermal protection systems. These materials are composed of refractory fibers With diameters on the order of 1 to 15 micrometers. They are lightweight and have an open, highly porous microstructure. Typical densities are less than 500 kilograms per cubic meters, and porosities generally exceed 0.8. Because of their open porosity, these materials are permeable to gas glow. There are numerous instances in which internal gas transport in a thermal protection system could be important; examples include the penetration of hot boundary-layer gases into the insulation, the flow of decomposition (pyrolysis) products from the interior, the use of convective flows to mitigate ice formation caused by cryopumping, and the design of refractory vents for pressure equilibration during atmospheric entry. Computational analysis of gas flow through porous media requires values of permeability which have not previously been available for the rigid fibrous insulations used in thermal protection systems. This paper will document measurements of permeability for a variety of insulations from NASA's LI, FRCI, and AETB families of lightweight ceramic ablators. The directional anisotropy of permeability and its dependence on gas pressure and material density will be presented. It will be shown that rarified-flow effects are significant in the flow through such materials. Connections will be drawn between the insulation microstructure and permeability. The paper will also include representative computations of flow through rigid fibrous insulations.