Thermal Protection System Research Articles

Unique Carbon Fibers for High-Temperature Applications

Carbon-carbon composites have been used for many decades in applications, such as rocket nozzles, brakes, and thermal protection systems (TPS). The material consists of carbon fiber in a matrix of carbon/graphite. Carbon-carbon provides structural performance while withstanding extremely high temperatures. The material has a low coefficient of thermal expansion and performs well under thermal shock. Interest in carbon-carbon has grown recently with the focus on hypersonic weapon systems that undergo high thermal shock. The challenge with hypersonics is to develop a thermal protection systems that does not ablate or erode so the control surfaces of a hypersonic weapon remain intact to allow for manuevering of the weapon in flight. Most carbon-carbon composites produced to date do ablate at elevated temperatures and there continues to be many applications for this material. Rocket nozzles and throats are a primary example and with the increases in launches of rockets, ablative carbon-carbon will still be key component in launch systems.

Velocity measurements in particle-laden high-enthalpy flow using non-intrusive techniques

This research aims at analysing the particle-laden flow of the hypersonic high-enthalpy wind tunnel L2K, situated in Köln at the German Aerospace Center (DLR). In the L2K wind tunnel, Martian atmosphere can be created, and the facility can simulate heat load conditions encountered during atmospheric entry of Martian missions. In the tests, a simplified Martian atmosphere (97% CO2 and 3% N2) was used. The high-enthalpy flow was loaded with micrometric particles of magnesium oxide. The particles’ mean velocity was measured with a 2D–2C particle image velocimetry (PIV) system, in the region right downstream the nozzle expansion of the wind tunnel. The work proves the possibility of creating a high-enthalpy particle-laden flow for thermal protection systems (TPS) testing with simulated Martian atmosphere. Average particle velocities of around 2000 m/s are measured and compared with the numerical simulation of the wind tunnel’s particle-free flow, and with the flow velocity measured with tunable diode laser absorption spectroscopy (TDLAS). The study also highlights some unexpected results and features of the high-enthalpy particle-laden flow and proposes some theories for the causes of such effects, which include agglomeration due to melting, and gravitational effect.

Piezoelectric resonant ice protection systems - Part2/2 : benefits at aircraft level

Piezoelectric resonant de-icing systems are attracting great interest. This paper aims to assess the implementation of these systems at the aircraft level. The article begins with the model to compute the power requirement of a piezoelectric resonant de-icing system sized from the prototype detailed in Part 1/2 of this article. Then the mass, drag, and fuel consumption of this system and the subcomponents needed for its implementation are assessed. The features of a piezoelectric resonant de-icing system are finally computed for aircraft similar to Airbus A320 aircraft and aircraft of different categories (Boeing 787, ATR 72 and TBM 900) and compared with the existing thermal and mechanical ice protection systems. A sensitivity analysis of the main key sizing parameters of the piezoelectric de-icing system is also performed to identify the main axes of improvement for this technology. The study shows the potential of such ice protection systems. In particular, for the realistic input parameters chosen in this work, the electro-mechanical solution can provide a 54% reduction in terms of mass and a 92% reduction in terms of power consumption for an A320 aircraft architecture, leading to a 74% decrease in the associated fuel consumption compared to the actual air bleed system.

Preparation and characterization of novel lightweight Y2Si2O7 fibrous porous ceramics for high temperature thermal insulation

Fibrous porous ceramics play a vital role in the field of thermal protection systems. In this work, novel lightweight Y2Si2O7 fibrous porous ceramics were successfully prepared based on Y2Si2O7 fibrous membranes by molding method. The effects of mechanical dispersion speed and heat treatment temperature on the properties of Y2Si2O7 fibrous porous ceramics were characterized. The results revealed that benefiting from the appropriate fiber length, the ceramics had a better performance at a mechanical dispersion speed of 4000 revolutions per minute. The density, shrinkage, and compressive strength of the ceramics increased, but the porosity decreased slightly with the increasing heat treatment temperature. The prepared Y2Si2O7 fibrous porous ceramics exhibited favorable properties like low density (0.37 g/cm3), high porosity (90.56%), good high-temperature resistance (∼1350 °C), and low thermal conductivity (∼0.12 W m−1 K−1 at 27 °C, ∼0.566 W m−1 K−1 at 1000 °C). These properties indicate that Y2Si2O7 fibrous porous ceramics could be a promising candidate as thermal insulation materials in high-temperature environments.

Experimental investigation and tomography analysis of Darcy-Forchheimer flows in thermal protection systems

In thermal protection systems (TPS), Darcy’s law or Darcy-Forchheimer’s law is employed to model the pyrolysis gas flow within the anisotropic porous ablator depending on the flow regime considered. A key challenge with using these laws is first, the knowledge of the validity domain of each flow regime in terms of a critical Reynolds number (Rec). Secondly, the lack of data on macroscopic properties, namely, the permeability and Forchheimer tensors is particularly challenging for the relevance of the models. The objective of this work is to contribute to overcoming these challenges by performing experimental and X-ray tomographic image-based characterization of Calcarb, a commercial carbon preform used for manufacturing TPS. For this purpose, fluid flow within Calcarb was studied experimentally in the Through-Thickness (TT) and the In-Plane (IP) directions for Reynolds numbers of 0.05 to 10.46 -representative of the TPS application. Tomography image-based micro-scale simulations, involving the direct resolution of the Navier–Stokes equations under both flow regimes, were also performed. Experimental results exhibit the anisotropic nature of Calcarb, namely through Rec values, corresponding to the Darcy flow regime limit, slightly different in the two directions (Rec of 0.31 and 0.43) with measured permeability values of 1.248×10−10 m2 and 1.615×10−10 m2 for TT and IP directions respectively. In the Forchheimer regime, experimental Forchheimer coefficients β were 2.0010×105 m−1 (TT) and 1.4948×105 m−1 (IP). During the simulation process, a numerical strategy was defined to obtain the permeability tensor yielding values within 8% of the experimental ones. The comparison of experimental results vs simulation results in the Forchheimer regime was performed through the analysis of the pressure gradients as functions of Re in the x, y, and z directions. The numerical estimations were compared successfully with experimental measurements, with a discrepancy of 5.2%, for Re values up to 2.4.

Probabilistic Design Method for Aircraft Thermal Protective Layers Based on Surrogate Models

In this study, a probabilistic method was proposed for an aircraft’s thermal protective layers. The uncertainties of material properties, geometric dimensions, and incoming flow environments were considered for the design inputs. To accelerate the design efficiency, Latin hypercube sampling and surrogate models were built based on finite element method calculations to enhance the simulation efficiency. Thus, the Monte Carlo method can be implemented with such a fast simulation method and produce a massive number of samples for the uncertainty quantification and sensitivity analysis, exploring their impact on the back temperature of the thermal protection layer. Compared to the deterministic method with the extreme deviation design, the probabilistic design yields a weight reduction of 15.61%. This indicates that probabilistic design is an efficient approach to enhance the performance of aircraft and reduce the overall weight of the aircraft. The general goal of this study is to provide a new design method for the coating film of thermal protection systems by considering multiple sources of uncertainties.

Experimental study of the effects of isolated roughness elements on the stability and transition of a hypersonic boundary layer on a flat plate

Surface roughness elements on hypersonic vehicles can cause early boundary layer transition, increasing wall skin friction and heat flux and affecting aircraft range and thermal protection systems. Accurate prediction of the transition caused by these roughness elements is crucial for the design of hypersonic vehicles. In this work, wind tunnel experiments on isolated roughness-induced boundary layer transition at Ma = 6 are conducted. Infrared thermography and high-frequency pressure sensors are utilized to investigate the effects of different roughness element configurations (cylindrical, diamond, ramp) on the hypersonic boundary layer instability and transition. The experimental results show that all three roughness elements can effectively enhance the generation of second mode waves and promote boundary layer transition. Compared to smooth surfaces, they exhibit similar frequency band range, faster growth, and earlier saturation. Among them, the ramp roughness element most effectively triggers the boundary layer transition, with a relatively small heat flux increase. Furthermore, bispectral analysis illustrates that all three roughness elements undergo self-interactions that lead to spectral broadening, ultimately resulting in boundary layer transitions.

Numerical Investigation on Transpiration Cooling Performance of Oriented Porous Structure Under Local Block Conditions

With the development of an active thermal protection system (TPS) for hypersonic vehicles, reliability improvement and weight reduction of TPS are key issues to be considered. In this paper, an oriented porous structure (OPS) is proposed to solve the local overheating problem in transpiration cooling. The transpiration cooling performances are investigated in a two-dimensional computational domain using a local thermal nonequilibrium model. The cooling mechanisms of OPS and uniform porous structure (UPS) under local block conditions have been numerically analyzed and compared. The results show that OPS leads to effective cooling of the overheated region by regulating the coolant flow behavior, especially in the case of serious block conditions. The maximum temperature reduction can be up to more than 200 K, and the maximum optimization rate can reach up to 23% compared to UPS. Moreover, OPS consumes less coolant and uses less driving force to deliver coolant when reaching the same peak temperature as OPS. This study shows that OPS has a better thermal protection effect in transpiration cooling, which can improve the reliability of TPSs, and that OPS can also achieve the lightweight of the aircraft.

Flow Mechanics in Ablative Thermal Protection Systems

Ablative thermal protection systems have experienced renewed interest in the past decade owing to the retirement of NASA's Space Shuttle fleet and the US presidential mandate to develop technologies that enable humans to explore space beyond low Earth orbit. Blunt body architecture for spacecraft and the use of ablators for thermal protection systems returned as the primary choice in mission planning. This review addresses current progress in modernizing predictive tools for ablative material response. Current theory development leverages progress made in the theory of flows in porous media. This development, combined with progress in experimental techniques and high-end computing, is enabling the development of 3D macroscale models with realistic closure coefficients derived from direct numerical simulations of 3D microscale geometries of actual materials. While flight data quantifying ablative material response remain sparse, the next decade will be one of exploration in which heatshield instrumented spacecraft will provide crucial flight data for refining and validating closure models.

ZrO2-reinforced polymer-matrix composites used for thermal protection systems of ultra-high temperature aerospace propulsion

Owing to the rapid development of high-energy propellants, the temperature of the solid rocket motor (SRM) combustion chamber has increased, necessitating the development of high-performance polymer matrix composites (PMCs) that meet the requirements of thermal protection systems (TPSs) for use in the ultra-high temperature combustion chamber. ZrO2 is an excellent and potential ultra-high temperature filler, the ablation properties of ZrO2-reinforced PMCs were investigated. ZrO2 reduced the ablation rate of the PMCs by up to 30.6 %. In order to optimize the design of ZrO2 filler, the reinforcement mechanism of ZrO2 on the ablation resistance of the PMCs was explained from the perspective of the properties and chemical reaction in the char layer. The results showed that ZrO2 can increase the residue rate and promote the densification of the char layer. In addition, ZrO2 promoted the formation of ZrC in the char layer, thereby improving the ablation resistance of the ZrO2-reinforced PMCs. The ZrO2-reinforced PMCs are expected to meet the requirements of the TPS used in the ultra-high temperature combustion chamber. Furthermore, the reinforcement mechanism of ZrO2 is conducive to the design of TPSs materials with excellent ablation performance and development of predictive models for the ablation performance of composites.

In Situ Ceramization of Nanoscale Interface Enables Aerogel with Thermal Protection at 1950 °C.

Conventional carbon fiber felt-reinforced aerogel composites are often used as lightweight thermal protection systems (TPSs) for aerospace craft. However, due to their poor oxidation resistance, they have gradually failed to handle increasingly harsh thermal environments. In this work, a nanoscale composite coating interface of SiC-ZrC ceramic precursor is first constructed on the fiber surface. Subsequently, using the coated fiber felt as a three-dimensional skeleton and through polymerization-induced phase separation, an aerogel composite with excellent thermal protection in extreme thermal environments is prepared. Owing to the in situ ceramization of this nanoscale interface at ultrahigh temperatures, the back temperature of the 12 mm thick aerogel is only 147 °C after exposure to an oxyacetylene flame at 1950 °C for 70 s. Meanwhile, the central region of the aerogel recedes by only 7%. Not only does this work provide a way to enhance aerogels by constructing a self-ceramizable nanoscale interface it is also expected that the developed aerogel composite can be applied in the ultrahigh-temperature thermal protection of future aerospace craft.

Chemical and structural degradation of room temperature vulcanizing (RTV) silicone at high temperatures

Room temperature vulcanizing (RTV) silicone is an insulating adhesive used on thermal protection systems of space capsules. RTV materials experience significant heating during entry into an atmosphere. To understand the response of RTV materials in entry environments, RTV 511 and RTV 560 are heated in a mass spectrometer, a thermogravimetric analyzer, and a high-temperature flow tube furnace to understand the underlying chemical and structural degradation processes occurring during the decomposition of the two RTV materials. In addition, infrared spectroscopy is used to identify the functional groups and the polymer backbone configuration. X-ray computed tomography is used to create three-dimensional volumes of the decomposed samples to examine the structure of the degraded RTV materials. The three-dimensional volumes are also used to evaluate the permeability of degraded RTV materials. It is observed that both RTV materials decompose in three stages. Analysis of the three-dimensional volumes generated through X-ray computed tomography indicates that the decomposition begins with the formation of small cavities in both RTV variants. However, as the materials are subjected to higher temperatures, RTV 511 evolves into an outer solid layer with cavities inside, while the solid material in RTV 560 is distributed more homogeneously throughout the material, resulting in a structure with better mechanical integrity. Although most regions of both RTV variants are impermeable, certain regions of RTV 511 have a permeability that is two orders of magnitude higher than the permeability of the base heat shield material. The overall analysis indicates that the response of RTV materials to heat must be considered in the design and analysis of thermal protection systems.

Development of advanced materials for thermal protection systems in spacecraft

Development of advanced materials for thermal protection systems in spacecraft

Dynamic Thermal Response of Multiple Parallel Cracks in a Half Plane under General Transient Thermal Loading

Understanding the dynamic thermal response of materials is crucial for designing effective thermal protection systems, particularly in extreme thermal environments such as transient thermal loading, extremely low/high temperature, etc. This study investigates the dynamic thermal response of a half plane containing multiple parallel cracks under transient thermal loading using a non‐Fourier, hyperbolic heat conduction model. Our findings highlight significant deviations from traditional Fourier models, enhancing the predictive capabilities for designing thermal protection systems in extreme thermal environments. The cracks are modeled as distributions of thermal dislocations, with densities determined through Fourier and Laplace transforms. By solving the resulting singular integral equations, we calculate the temperature gradient intensity factors across multiple scenarios. Additionally, we examine the effects of thermal relaxation time, loading parameters, and crack spacing on the thermal intensity factors, both in transient and steady‐state conditions. This research confirms the robustness of the hyperbolic model and its practical implications for thermal analysis in engineering applications.

Enhancement of mechanical and thermal properties of carbon fiber phenolic resin composites using silicon carbide filler for thermal protection system applications

Thermal protection systems (TPS) are vital for re-entry vehicles for their safe passage into the atmosphere from space. Hence, researchers took a keen interest in improving the thermal and ablative properties of composites to be used in making thermal protection systems. Therefore, an attempt was made to improve the thermal and ablative properties of composites made of carbon fibers (Cf) and resorcinol formaldehyde phenolic (Ph) resin with the incorporation of silicon carbide (SiC) particles. The filler was added in various percentages (0 wt% - blank, 1 wt%, 3 wt%, and 5 wt%), and the composites were tested for ablative, thermal and mechanical properties. The results demonstrate that the SiC-modified PAN-based carbon fiber reinforced phenolic (SiC-PANCf-Ph) composite with 3 wt% SiC enhancement exhibited ideal properties. The post-ablation phase composition and microstructure were examined through X-ray diffraction (XRD), Scanning electron microscopy (SEM), and energy-dispersive X-ray spectroscopy (EDS). The surface morphology evidences the formation of a silicon dioxide (SiO2) layer on the composites. The SiC-PANCf-Ph composites demonstrated the lowest ablation rate, enhancing their potentiality for effective TPS applications.

A stability-enhanced approach for aerodynamic heating computation in high-speed flows

Near-space vehicles face severe aerothermal environments due to hypersonic speeds. Accurately predicting hypersonic aerodynamic heating is crucial for the design of spacecraft’s thermal protection systems, aerodynamic shapes, and propulsion systems. Numerical simulations have become an important tool for computing aerothermal environments under extreme conditions. However, there is a contradiction between the robustness and low dissipation characteristics of existing numerical methods for hypersonic flow fields. This paper proposes an effective dimension hybrid approach that enhances the robustness of AUSM-family schemes while maintaining their low dissipation characteristics. Numerical experiments demonstrate the proposed approach’s effectiveness and potential for practical engineering applications.

Research Progress on Carbide Ultra-high Temperature Ceramic Anti-ablation Coatings for Thermal Protection System

Research Progress on Carbide Ultra-high Temperature Ceramic Anti-ablation Coatings for Thermal Protection System

Design of Planar and Curved Spiral Antennas Covered by 2-D Conformal Thermal Protection System

Design of Planar and Curved Spiral Antennas Covered by 2-D Conformal Thermal Protection System

Preparation of ceramifiable and ablatively resistant epoxy resins by microphase structure modulation of silicones

The development of ceramic-polymer composites is a focal point in the research of thermal protection systems, and the control of microstructure is crucial for fabricating high-performance resins. In order to investigate the effect of the silicon phase structure on the properties of epoxy resins and preparate the ablation-resistant ceramifiable resins, the homogeneous and “sea-island” phases of epoxy resins with different amounts of silicones were prepared by ring-opening and grafting reactions. The reaction mechanism and the microphase structure of silicone were verified by SEM and TEM. It was found that the homogeneous silicone-epoxy resins exhibited exceptional ablation resistance, with mass and linear ablation rates of 0.0646 g/s and 0.0073 mm/s, which were 78.61 % and 98.36 % lower than those of epoxy resins, and exceeded most of the resin-based ablation-resistant materials reported in the literatures. The homogeneous silicone further slows down the pyrolysis rate of the epoxy resin and inhibits the release of pyrolysis products. During the carbonization and ceramization, the carbothermal reduction reaction between silicon and carbon occurs sufficiently in the homogeneous system, which enhanced the content of SiC ceramics in the carbon layer. Eventually, a dense char layer with low porosity and a highly graphitized structure constructed in situ on the ablative surface, which greatly improves the ablative resistance. Furthermore, its exceptional mechanical properties render it capable of substituting conventional resin-based thermal protection materials, thereby meeting the stringent demands of challenging environments.

Numerical analysis on thermal environment of reusable launch vehicle during supersonic retropropulsion

Developing reusable launch vehicles (RLV) is an effective way to significantly reduce space transportation costs. However, in the return stage of the vehicle, especially the supersonic retropropulsion (SRP), the first stage is partially submerged in the hot exhaust plume and faces severe convective heating at the base plate and sidewalls. As a rule of thumb, the layout of nozzles at the rocket base significantly affects the heat load. In order to evaluate the influence of heat load under different nozzle configurations during SRP (33 km–22 km), the thermal environment of RLV was studied through numerical investigation for three typical nozzle configurations. Using Reynolds-averaged Navier–Stokes equations (RANS)-based computational methods and the k−ω SST turbulence model, the flowfield is characterized at different trajectory points during retropropulsion. Results showed that the high heat flux is mainly concentrated in the baseplate's center area and the sidewall's near base area during SRP. In general, activating engines in the radial position is beneficial for the thermal protection system of a reusable launch vehicle while the thrust is identical. Base heating primarily results from high-temperature counterflows due to plume interactions, while sidewall heating is predominantly a result of high-temperature recirculation zones generated by plume-induced flow separation. During the supersonic retropropulsion phase, the 3-nozzle-180°, a non-closed nozzle configuration, has significantly lower heat loads on the baseplate and the sidewall than that of the 4-nozzle-90° and the 3-nozzle-120°. In addition, reducing the number of engines is also an effective solution under the same conditions of thrust. Studying the layout of the working engine without changing the original layout has important guiding significance for improving the reusability and safety factor of the RLV.