Flexible Thermal Protection System Research Articles

Inflatable aerodynamic decelerators for CubeSat reentry and recovery: Surface properties

In this investigation, Direct Simulation Monte Carlo Method is used to compute the surface properties of inflatable aerodynamic decelerators (IAD) applied for the reentry, recovery, and reuse of cubeSats. Surface properties calculation is of fundamental importance for flexible thermal protection system materials selection that withstand the harsh reentry environment. Furthermore, simulations are extremely important to detect possible hot spots at the IAD before design and flight. In this work, three IAD configurations are exposed to a rarefied hypersonic flow with freestream conditions that correspond to those found at reentry conditions at 105 km altitude. All geometries are assumed to be axisymmetric and have a maximum radius of 0.3 m from the center line to the shoulder edge. The main differences between the geometries are the angle between the IAD and cubeSat body and the IAD curvature for geometry 3. According to the results, all IAD geometries are effective in protecting the cubeSat during the reentry phase. It is shown that the IAD geometry has a significant influence on the heat transfer, pressure, and skin friction coefficient distribution over the flexible heat shield. The maximum heat transfer coefficient computed for geometry 1, 2, and 3 is 0.90, 0.97, and 0.84, respectively.

Arcjet Evaluation of Thermocouple Performance in Flexible Thermal Protection System Materials

The flexible thermal protection system (FTPS) on NASA’s low-Earth-orbit flight test of an inflatable decelerator vehicle was instrumented with thermocouples to measure the in-depth thermal response during entry into Earth’s atmosphere. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models, but the deployable nature of inflatable decelerator technology presents challenges. The thermocouples need to be compactable, cannot damage the FTPS or the inflatable structure, and need to be able to withstand high temperatures and large thermal gradients. The present paper details the arcjet testing that was performed to examine the potential error sources associated with instrumenting NASA’s FTPS with different thermocouple Types (Type K versus Type N), wire gauges (30 versus 28 American wire gauge), insulations (glass versus ceramic), and barriers (mica versus no mica). Several error sources were identified, including 1) conductive deposits on the thermocouple insulation electrically shorting the thermocouple leads; 2) melting of the glass thermocouple insulation, allowing for electrical shorting of the thermocouple leads; 3) thermocouple leads melting, resulting in loss of signal; and 4) oxidation of the positive thermocouple lead (known as green rot), generating a drift in the measured output. A mica tape wrap around each individual thermocouple lead was successfully used to mitigate the first error source, and switching to ceramic insulation eliminated the second error source. The third and fourth error sources were examined by exploring the effect of the thermocouple type and wire gauge. Thermocouple solutions that minimize the potential error sources at each layer in the FTPS are presented. Finally, the arcjet test results are compared to the previously reported ground-based tube furnace test results, and their applicability to the flight environment is discussed.

Gas–thermal–solid coupling analysis of a hypersonic inflatable reentry and descent system

Hypersonic reentry has a severe flight environment during spacecraft return. To analyze the influence of pneumatic loads on the flexible structure under the extreme load condition during the deceleration process of inflatable reentry and descent system, fluid–and thermal–solid couplings have been studied. In addition, the stress and structural deformation distributions and the temperature distribution of each functional layer of the flexible thermal protection system were obtained in this study. The influence of flight conditions and structural parameter changes on inflatable structure performance was studied, and the transient response mechanism of the pneumatic load to the inflatable structure was revealed. Ballistic analysis combined with engineering algorithm was used to predicted the flight envelope. The fluid-solid thermal coupling method in Workbench is utilized to realize the real-time transfer of aerodynamic and thermal loads to flexible structures. And the loose coupling method was applied to carried out the one-way transfer of aerodynamic load and thermal load to the surface of flexible structure. The results showed that the mechanical–thermal–structural coupling analysis method could predict the mechanical properties of flexible inflatable deceleration structures, which can provide references for the aerodynamic configuration and thermal protection design under extreme flight conditions.

Feasibility of a Morphing Rocket Nozzle for Thrust Vector Control based on Corrugated Composite Laminates

The paper presents technological, experimental and numerical studies aimed at the development of a morphing nozzle in composite material with a deformable divergent section, which is designed to control the vector of thrust in rockets without using mechanisms or localized flexible joints. The nozzle design is characterized by corrugated composite walls and longitudinal elements that make possible a separation of load paths involved in bending response, bearing of internal pressures and related force resultants. Mechanical requirements and conditions of pressure, temperature and gas velocity in the morphing part of the nozzle are defined by selecting and investigating two potential application scenarios. Thereafter, both technological aspects and the assessment of mechanical performances are addressed by designing, manufacturing and testing a composite demonstrator with morphing capabilities. A lightweight and flexible thermal protection system is proposed, designed and numerically analyzed. Finally, a virtual demonstrator of the composite morphing nozzle is developed. Results point out the possibility of manufacturing composite elements with axial load bearing capability that may undergo significant bending deformation without failures. The integration of a flexible thermal protection systems leads to a promising concept for the development of innovative morphing nozzle for rocket engine applications.

Electrical Shorting of Thermocouples in Flexible Thermal Protection System Materials

The flexible thermal protection system (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle will be instrumented with thermocouples to measure the in-depth thermal response during entry into Earth’s atmosphere. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models, but the deployable nature of inflatable decelerator technology presents challenges. The thermocouples need to be compactable, cannot damage the FTPS or the inflatable structure, and need to be able to withstand high temperatures and large thermal gradients. The present paper details the developmental testing that was performed using a tube furnace to examine the potential error sources associated with instrumenting NASA’s FTPS with thermocouples insulated with either glass braid or ceramic braid. Large measurement errors occurred with both thermocouple insulation materials due to conductive deposits on the insulation electrically shorting the thermocouple leads. Two mechanisms were identified that contribute to this phenomenon: 1) the organic binder on the thermocouple insulation carbonizes in a high-temperature, low-oxygen environment and 2) decomposition products from the FTPS are able to permeate the braided insulation. The effect of heat cleaning the thermocouples to eliminate the first mechanism is presented, and the possibility of using a mica wrap as an impermeable barrier to mitigate the second mechanism is explored. Finally, the applicability of the ground-based tube furnace test results to the flight environment is discussed.

Lightweight, flexible, and heat‐insulated phenolic impregnated carbon ablator (PICA) with adjustable flexibility and high compressive resilience property

AbstractPhenolic impregnated carbon ablator (PICA) as a typical lightweight ceramic ablative material is widely used in thermal protection systems (TPS). However, the rigidity and poor mechanical property of the existing PICA limit its application in flexible TPS. For the purpose of solving these problems, a novel lightweight, flexible, and heat‐insulated PICA material is first successfully fabricated using resorcinol‐formaldehyde (RF) aerogel as matrix and flexible carbon fiber felt as reinforcement through vacuum impregnating and ambient pressure drying. The prepared composites have relatively low densities (0.20 ~ 0.27 g/cm3) and low thermal conductivities (~0.11 W/[m·K]) at room temperature. The flexible PICA has adjustable flexibility through changing the RF concentration of impregnated sols and exhibits excellent folding endurance, which can be folded more than 500 times without fracture. The material can recover to its original shape without fracture when the compressive strain is up to 60%, which shows high compressive resilience properties. The result of the back‐face temperature test indicates the prepared flexible PICA composite has improved thermal insulation property at high temperature by impregnating with RF aerogels. These results demonstrate the flexible PICA prepared in this work is a promising material that can be used in the flexible thermal protection field of aerospace.

Microscale Knudsen Effect over the Transverse Thermal Conductivity of Woven Ceramic Fabrics Under Compression

Woven-fiber ceramic materials have shown remarkable results in the design of insulative lay-up structures for flexible thermal protection systems. A deeper understanding of heat transfer through the different insulation layers is key for predicting the performance of heat shields. In this article, a thermo-mechanical multiscale model is developed to predict the out-of-plane thermal conductivity at the micro- and meso-levels of transversely loaded two-dimensional woven ceramic fabrics. Knudsen effects within the multiscale structure are studied by adjusting gas pressure conditions. Alumina-based Nextel-BF20 and silicon carbide Hi-Nicalon with a 5-harness satin weave pattern are modeled by finite-element analysis. The computational results are validated experimentally by applying the anisotropic transient plane source method. We find that out-of-plane thermal conductivity decreases significantly with gas pressure due to Knudsen effects in the confined air within the fibrous structure. The dependence of thermal conductivity of fabrics on fiber volume fraction is shown to decrease markedly with pressure reduction. The proposed multiscale modeling approach yields a notable accuracy improvement, with respect to simplified series-parallel models, when compared with our experimental measurements. The FEA model is applicable to other fabric materials and loading conditions and presents an opportunity to study how changes at the fiber level affect the overall thermal behavior of woven fabrics.

Aerothermodynamics and thermal design for on-ground and in-flight testing of a deployable heat shield capsule

The article deals with the aerothermodynamic and thermal analysis of two technological demonstrators suitable for research activities on deployable heatshield entry capsules. Variable geometry bodies represent a formidable way to build low ballistic coefficient entry platforms able to limit convective heat fluxes, mechanical loads, and final descent velocity. These characteristics suggest several interesting applications for space missions such as to exploit those systems for human or large payloads landing on Mars. The main feature of the demonstrators is the umbrella-like deployable heatshield made of a flexible thermal protection system. In this framework, the work reports on the design analyses carried out for a flight and a ground demonstrator developed to prove, with suborbital flight and plasma wind tunnel test campaigns, several key technologies of the deployable heatshield. Several non-equilibrium three-dimensional flowfield simulations were performed to assess the aerothermal loading conditions expected for the demonstrators. Pressure and heat flux distributions on the deployable heatshield are provided and discussed. Finally, those surface loading conditions are set to feed a finite-element thermal analysis of the demonstrators, also detailed in the work.

Trends in the development of flexible thermal protection systems for modern aircraft

The emergence of new types of space and aviation technology necessitates the development of new types of thermal protection systems capable of operating at high temperature and long operating times. There are several types of thermal protection systems for different operating conditions: active thermal protection systems using forced supply of coolant to the protected surface, passive thermal protection systems using materials with low thermal conductivity without additional heat removal, high-temperature systems, which are simultaneously elements of the bearing structure and provide thermal protection, ablation materials. Heat protection systems in the form of rigid tiles and flexible panels, felt and mats are most common kind of heat protecting systems. This article examines the trends of development of flexible reusable heat protection systems intended for passive protection of aircraft structural structures from overheating.

Anisotropic thermal conductivity under compression in two-dimensional woven ceramic fibers for flexible thermal protection systems

Flexible thermal protection materials made from two-dimensional woven ceramics fibers are of significant interest for hypersonic inflatable aerodynamic decelerators being developed by NASA for future missions on Mars and other planets. A key component of the thermal shield is a heat-resistant outer ceramic fabric that must withstand harsh aero-thermal atmospheric entry conditions. However, a predictive understanding of heat conduction processes in complex woven-fiber ceramic materials under deformation is currently lacking. This article presents a combined experimental and computational study of thermal conductivity in 5-harness-satin woven Nextel 440 fibers, using the hot-disk transient plane source method and computational thermo-mechanical modeling by finite-element analysis. The objective is to quantify and understand the effect of compressive strain on anisotropic heat conduction in flexible two-dimensional ceramic materials. We find, both experimentally and theoretically, that thermal conductivity of woven fabrics rises in both in-plane and out-of-plane directions, as the transverse load increases. Air gap conduction and fiber-to-fiber contacts are shown to play a major role in this behavior. Our finite-element simulations suggest that the thermal conductivity anisotropy is strong because heat transfer of air confined between fibers is reduced compared to that of free air. The proposed modeling methodology accurately captures the experimental heat conduction results and should be applicable to more complex loading conditions and different woven fabric materials, relevant to extreme high temperature environments.

Surface Heating and Boundary-Layer Transition on a Hypersonic Inflatable Aerodynamic Decelerator

The aerothermodynamic environment of a hypersonic inflatable aerodynamic decelerator with a flexible thermal protection system has been investigated through wind-tunnel testing. Boundary-layer transition onset locations and surface heating distributions were measured using global phosphor thermography on deflected, solid-surface models representative of a hypersonic inflatable aerodynamic decelerator aeroshell with a flexible surface. Data were obtained at Mach 6 for a wide range of surface deflection heights, angles of attack, and freestream Reynolds numbers. The surface deflections resulted in earlier boundary-layer transition onset and higher heating levels than for a smooth-wall surface, with the greatest effects occurring on the leeward side of the aeroshell. The heating augmentation and transition onset data were correlated using flowfield properties obtained from a complementary computational study. A method was developed to estimate flexible thermal protection system deflection effects using these correlations along with smooth-wall flowfield predictions. This method was validated through comparisons to the wind-tunnel data and then used to generate postflight predictions for aeroshell heating levels of the Inflatable Reentry Vehicle Experiment flight-test program. The results of this study provide a means for estimating flexible thermal protection system deflection effects for future flight tests as well as an experimental database for use in the development and validation of computational methods for simulations of hypersonic inflatable aerodynamic decelerator aerothermodynamic environments.

Thermal Protection System Response Uncertainty of a Hypersonic Inflatable Aerodynamic Decelerator

The objective of this paper is to investigate the uncertainty in the bondline temperature response of a flexible thermal protection system subject to uncertain parameters in the hypersonic flowfield and thermal response modeling of a hypersonic inflatable aerodynamic decelerator configuration for ballistic Mars entry. An inflatable decelerator with a 10 m major diameter is selected for this study based on the forebody dimensions scaled from the 6 m test article that was tested in the NASA Ames Research Center’s National Full-Scale Aerodynamics Complex facility. A global nonlinear sensitivity analysis study for the bondline temperature uncertainty shows that the dimension of uncertain parameters can reduce from 22 to 8. An uncertainty analysis of the bondline temperature in the reduced dimensions indicates that the bondline temperature varies by as much as 125% above the nominal prediction and exceeds the temperature limit of 400°C. The largest uncertainty occurred at 70 s in the trajectory before separation of the inflatable decelerator for transition to a secondary descent technology. The main contributors to the bondline temperature uncertainty are the insulator and outer fabric conductivities, as well as the freestream density. The thickness and initial density of the insulator layer, closest to the gas barrier layer, are also shown to be significant contributors to the bondline temperature uncertainty, especially earlier in the trajectory.

Fracture Behavior of Woven Silicon Carbide Fibers Exposed to High‐Temperature Nitrogen and Oxygen Plasmas

High‐temperature aero‐thermal heating in a 30 kW inductively coupled plasma torch was used to replicate the effects of harsh oxidizing environments during hypersonic atmospheric entry on fracture behavior and microstructure of two‐dimensional woven SiC fibers. Hi‐Nicalon SiC woven cloths were exposed to surface temperatures over 1400°C with different high‐enthalpy dissociated oxygen and nitrogen plasma flows, and were subsequently deformed in pure tension at room temperature. Changes in fiber microstructure and surface chemistry after thermal exposure were examined by scanning electron microscopy. Pure nitrogen plasmas resulted in a 50% decrease of strength in woven SiC fibers with minimal effects on the fiber structure, except for highly localized surface pitting caused by partial decomposition of silicon oxycarbonitride phase at high temperature. In contrast, exposure to dissociated oxygen and air plasmas led to severe strength reduction and embrittlement over significantly short time scales, corresponding to degradation rates up to 200 times higher than those reported with static heating at equivalent temperatures. The origin of accelerated embrittlement at microscopic scale was found related to complex gas‐surface interactions and high‐temperature oxidizing processes involving the formation of SiO2 bubbles and microcracks on the surface. These findings are important for the development of outer fabric materials for new flexible thermal protection systems in space applications.

Aeroelastic Stability of Thermal Protection System for Inflatable Aerodynamic Decelerator

A theoretical aeroelastic stability analysis has been performed on the flexible thermal protection system for an inflatable aerodynamic decelerator. Structural models consist of one or more truncated conical shells of the Donnell type, which may be elastically supported along the middle surface. The aerodynamic model is first-order piston theory. The Lagrangian of the system is formulated in terms of the generalized coordinates for all shell displacements, and the Rayleigh–Ritz method is used to derive the equations of motion. The aeroelastic stability boundaries and mode shapes are found by calculating the eigenvalues and eigenvectors of a large coefficient matrix. When the thermal protection system is approximated as a single conical shell, circumferentially asymmetric coalescence flutter between the second and third axial modes is observed. When many circumferential elastic supports are included, the shell flutters symmetrically in zero circumferential waves, with the first, second, and third axial modes being the most critical. In this case, the flutter boundary, flutter mechanism, and critical modes may change significantly with the addition of structural damping. Aeroelastic models that consider the thermal protection system as multiple interacting shells tend to flutter asymmetrically at high dynamic pressures relative to the single shell models, with higher axial modes being more critical. It is also found that tension applied at the shell edges, orthotropicity, and elastic support stiffness are important parameters that can dramatically affect the shell’s flutter behavior.

Methodology for Flight-Relevant Arc-Jet Testing of Flexible Thermal Protection Systems

A methodology to correlate flight aeroheating environments to the arc-jet environment is presented. For a desired hot-wall flight heating rate, the methodology provides the arc-jet bulk enthalpy for the corresponding cold-wall heating rate. A series of analyses were conducted to examine the effects of the test sample model holder geometry to the overall performance of the test sample. The analyses were compared with arc-jet test samples, and challenges and issues are presented. The transient flight environment was calculated for the Hypersonic Inflatable Aerodynamic Decelerator Earth Atmospheric Reentry Test vehicle, which is a planned demonstration vehicle using a large inflatable, flexible thermal protection system to reenter the Earth’s atmosphere from the International Space Station. A series of correlations were developed to define the relevant arc-jet test environment to properly approximate the vehicle flight environment. The computed arc-jet environments were compared with the measured arc-jet values to define the uncertainty of the correlated environment. The results show that, for a given flight surface heat flux and a fully catalytic thermal protection system, the flight-relevant arc-jet heat flux increases with the arc-jet bulk enthalpy, while for a noncatalytic thermal protection system, the arc-jet heat flux decreases with the bulk enthalpy.

An Armadillo-Like Flexible Thermal Protection System for Inflatable Decelerators: A Novel Paradigm

Polymeric ablative materials play a strategicrole in the whole aerospace industry: they areused to produce the thermal protection system (TPS), which protects vehicles during thehypersonic flight through a planetary atmosphere. To date, TPSs are typically designed asrigid monolithic heat shields. Inflatable deployable aeroshells (IDAs) offer an alternativeto rigid TPSs: IDAs are based on the use of flexible TPSs (FTPSs). State of the art FTPSscan handle heat fluxes up to approximately 20Wcm