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Space exploration missions rely on ablative heat shields for the thermal protection of spacecraft during atmospheric entry flights. While dedicated research is needed for future missions, the scientific community has limited access to ablative materials typically used in aerospace. In this paper, we report the development of the HEFDiG Ablation-Research Laboratory Experiment Material (HARLEM), a carbon–phenolic ablator designed to supply the need for ablative materials in laboratory experiments. HARLEM is manufactured using polyacrylonitrile-based carbon fiber preforms and a simplified processing route for phenolic impregnation. We characterized the thermal protection performance of HARLEM in arcjet experiments conducted in the plasma wind tunnel PWK1 of the Institute of Space Systems at the University of Stuttgart. We assessed the performance of the new material by measuring surface recession rate and temperature using photogrammetry and thermography setups during the experiments, respectively. Our results show that HARLEM's thermal protection performance is comparable to legacy carbon–phenolic ablators that have been validated in different arcjet facilities or in-flight, as demonstrated by calculations of the effective heat of ablation and scanning electron microscopy of as-produced samples. In-house manufacturing of carbon–phenolic ablators enables the addition of embedded diagnostics to ablators, allowing for the acquisition of data on internal pressure and more sophisticated pyrolysis analysis techniques. [ABSTRACT FROM AUTHOR]

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. [ABSTRACT FROM AUTHOR]

Solid Rocket Motors (SRMs) are used to launch satellites and other civilian or military payloads in orbit. Due to their low density and outstanding insulation capability, EPDM/aramid fibers ablatives represent the state of the art of Elastomeric Heat Shielding Materials (EHSMs) SRMs. However, this class of polymeric ablatives tend to exhibit a limited ablation resistance induced by thermo-oxidation and mechanical erosion due to the propellant combustion byproducts. In order to increase the thermo-oxidation resistance, char retention, and insulation index, boron based fillers can be considered in the material formulation of the SRM insulation liner. In this work we systematically studied the effect of different boron containing fillers embedded in EPDM based EHSMs. The heat capacity, thermal stability, mechanical properties, ablation resistance, pre- and post-oxy-acetylene torch testing morphology of these materials were analyzed evidencing the clear benefits of these fillers. [ABSTRACT FROM AUTHOR]

Heat shields used to protect space capsules during very high-speed atmospheric entry incorporate lightweight insulating refractories based on carbon-fiber preforms. These ablators, with up to 80% porosity, present exceptional thermal and chemical properties. A joint experimental and modeling approach to study such materials is presented, which contributes to improving their design in relevant operating conditions. Samples were tested in an inductively coupled plasma at 1.5 kPa and 25 kPa pressures, with surface temperatures ranging from 1500 to 2800 K. The recession rate was measured in-situ and the local flow conditions were reconstructed numerically from experimental data. The porous medium was imaged by X-ray Computerized Micro-Tomography (CMT), and the depth affected by the gas-solid interaction phenomena due to the plasma exposure was extracted from these data. A simple analytical model has been derived to relate observable and reconstructed quantities and the fiber-scale reaction constants. Image-based numerical simulations of ablation by nitridation, with simultaneous catalytic recombination of atomic nitrogen, were compared to the analytical model and used to extract the intrinsic, fiber-scale reaction rate constants from test data. Results show, among others, that the fiber-scale nitridation rate constant is close to the rate of graphite nitridation, and that it decreases strongly with pressure. [ABSTRACT FROM AUTHOR]

Spallation is a phenomenon in which solid particles are ejected off the surface of an ablative material in a high-enthalpy, high-shear flow field. The main contributor to this phenomenon in carbon-based heat shields is the mechanical erosion of carbon fibers weakened by oxidation decomposition. The dynamics of this phenomenon, which are poorly characterized in the literature, strongly affect the ablation rate of the material. In state-of-the-art codes, ablation by spallation is modeled using a “failure” ablation rate that is empirically determined. The present study aims at understanding the rate of ablation of low-density carbon materials. Results from a test campaign at the NASA Langley Hypersonic Materials Environmental Test System (HYMETS) arc jet facility are used to examine spallation. High-speed multi-camera imagery at 44,000 fps is used to generate velocity vectors of spalled particles emitted from carbon-fiber samples exposed to an arc jet airflow. The imagery recorded approximately 4 × 10 6 unique particles, indicating that spallation is a potentially non-trivial process. The velocities of the particles ejected from the surface were found to be between 10 m/s and 20 m/s, accelerating to velocities as high as 250 m/s further away from the sample surface. Although the particle diameters were not directly observable, estimates suggest anywhere from 0.06% to 5.6% of the mass loss from the sample occurred due to spallation. [ABSTRACT FROM AUTHOR]

Ablative materials are at the base of entire aerospace industry; these sacrificial materials are used to manage the heat shielding of propulsion devices (such as liquid and Solid Rocket Motors (SRMs)) or to protect vehicles and probes during the hypersonic flight through a planetary atmosphere. Accordingly they are also known as Thermal Protection System (TPS) materials. Some non-polymeric materials have been successfully used as ablatives; however, due to their versatility, Polymeric Ablatives (PAs) represent the widest family of sacrificial TPS materials. In fact, when compared to non-polymeric ablatives such as high melting point metals, inorganic polymers (or metal oxides or carbides), PAs have some intrinsic advantages such as: tunable density, lower cost, and higher heat shock resistance. This review covers all main topics related to the science and technology of ablative materials with current and potential applications in the aerospace industry. After a short, yet comprehensive, introduction on non-ablative materials, this review paper summarizes fifty years of research efforts on polymeric ablatives, starting from the state of the art solutions currently used as TPS, up to covering the most recent efforts for nanostructuring their formulations. [ABSTRACT FROM AUTHOR]

EPDM/aramid fibers ablatives represent the state of the art Elastomeric Heat Shielding Materials (EHSMs) for Solid Rocket Motors (SRMs). Due to their mechanical properties and excellent thermal stability, aramid fibers or pulp constitute the common reinforcement of EPDM based formulations. Aramid fiber was introduced by NASA and space companies such as Orbital-ATK with the aim to replace asbestos. Space Shuttle’s reusable SRMs and many ICBMs employed Nitrile butadiene rubber (NBR) loaded with asbestos (ASNBR) as a primary internal case insulation; however, due to cancerogenicity of asbestos, NASA sought a replacement for many years, but this goal has been very challenging to achieve since the comprehensive performance of ASNBR was and is still difficult to match. Wollastonite can be considered the only non-hazardous mineral/inorganic filler able to partially replace asbestos. Wollastonite is a calcium inosilicate mineral (CaSiO 3 ) and the shape of the filler is acicular. Our study wanted to evaluate Wollastonite as a possible replacement of aramid pulp in the production of EPDM based EHSMs for solid rocket motors. The heat capacity, thermal and dimensional stability, mechanical properties, and ablation resistance of the produced materials were studied and here reported. [ABSTRACT FROM AUTHOR]

Elastomeric Heat Shielding Materials (EHSMs) are essential to manufacture Solid Rocket Motor (SRM) liners. Ethylene Propylene Diene Monomer (EPDM) represents the State Of The Art (SOTA) matrix for EHSMs and it is generally combined with aramid fibers and nanosilica to enable a sufficient ablation resistance. In addition to aramid, researches also tested other organic reinforcements such as Kynol® and Zylon®. Our team previously compared aramid (Twaron®) and phenolic filaments (Kynol®) and evidenced merits and deficiencies of both of fibers when embedded in EPDM based EHSMs. In this new step of this research we introduced a 1D model able to describe the insulation capability of the studied materials. The analytical considerations introduced in the theoretical section allowed to implement a numerical tool able to take into consideration the reversible and irreversible thermal expansion of the virgin and charred material. Furthermore, some efforts were effectively spent to introduce a model able to consider the coking process – char densification – due to the carbon deposition of the pyrolysis gases through porous carbonaceous skeleton. These considerations allowed to proper account for the in depth temperature profiles and also to produce the real char densities of the materials tested through oxy-acetylene tests. [ABSTRACT FROM AUTHOR]

We investigate here effects of adaptations like dimension reduction, omission of parameter correlation, and grid refinement on parameter influence over material temperature response and total recession. The task is accomplished by applying Morris screening and method of Sobol' sensitivity studies applied to a range of 3-D, 1-D, and 1-D TC1-driver cases for the MSL entry through the Martian atmosphere, where the material response of the MSL heat shield is simulated with NASA's PATO toolbox. Results of the investigation show that lower-dimensional approaches can, in some cases, be utilized for initial parameter screening prior to execution of a fully-detailed simulation, but in general, obtained sensitivities differ. Moreover, coarse computational grids do not impose significantly distinct input-output mechanics and can be utilized for accurate screening and quantitative sensitivity analysis exercises even with non-negligible residual values. The investigation also discusses the implications of an unknown correlation structure on parameter sensitivities while finding little evidence of significant correlation effects on parameter classification given input uncertainty bounds defined in this work. [ABSTRACT FROM AUTHOR]

The Mars Science Laboratory entry vehicle successfully landed the Curiosity rover on the Martian surface on 5 August 2012. A phenolic impregnated carbon ablator heatshield was used to protect the spacecraft against the severe aeroheating environments of atmospheric entry. This heatshield was instrumented with a comprehensive set of pressure and temperature sensors. The objective of this paper is to perform an inverse estimation of the entry vehicle's surface heating and heatshield material properties. The surface heating is estimated using the flight temperature data from the shallowest thermocouple. The sensitivity of the estimated surface heating profile to estimation tuning parameters, measurement errors, recession uncertainty, and material property uncertainty is investigated. A Monte Carlo analysis is conducted to quantify the uncertainty bounds associated with the nominal estimated surface heating. Additionally, a thermocouple driver approach is employed to estimate heatshield material properties using the flight data from the deeper thermocouples while applying the shallowest thermocouple temperature as the surface boundary condition. [ABSTRACT FROM AUTHOR]

Thermal protection system materials for the Mars Science Laboratory mission were tested in the NASA Ames 60 MW arcjet in a shear environment. Shear tests were performed on candidate ablative heat shield materials in wedge and swept cylinder test fixtures. In portions of the expected flight environment, the proposed main heat shield material, known as SLA-561V, recessed orders of magnitude faster than predicted. An alternate main heat shield material, known as phenolic impregnated carbon ablator, behaved reasonably well in all regions of the flight envelope investigated here. However, the measured recession rate of the phenolic impregnated carbon ablator was on average 50% greater than the predicted recession by the fully implicit ablation and thermal response code, and in some cases the measured recession was as much as 150% greater (uncertainties included). Phenolic impregnated carbon ablator's higher recession rate than predicted (in shear) resulted in adding thickness margin to the flight heat shield design. Other tests of the phenolic impregnated carbon ablator heat shield system, including tests of various gap fillers, damage scenarios, and repair scenarios, demonstrated the robustness of this heat shield system for use in the predicted Mars Science Laboratory flight environment. [ABSTRACT FROM AUTHOR]

The methodology employed for the thermal design and margins assessment of the Mars Science Laboratory aeroshell thermal protection system is reviewed. A new thermal margins policy was developed in the course of this work that provides additional rigor over previous methods. Because of a late change of thermal protection materials from the heritage super lightweight ablator 561V to phenolic impregnated carbon ablator, the design of the heat shield followed a nontraditional path in which the flight thickness was selected based on a mass (rather than thermal) limit. The material switch was followed by detailed thermal analyses that demonstrated that the baselined thickness was sufficient to provide adequate thermal protection to the vehicle without violating design requirements during a 3- sigma worst-case entry condition. The backshell material thickness was also finalized before the thermal sizing was completed, and the resulting analysis showed that there was more than sufficient material on the backshell. The parachute cone cover and backshell interface plate were the only major thermal protection system elements that followed a standard design process. Thermal sizing was performed for acreage and special features on the cone cover and interface plate, and the hardware was manufactured according to those analyses. [ABSTRACT FROM AUTHOR]

The aim of this study is to give information on the influence of the experimental conditions to estimate thermal diffusivity of carbon fiber reinforced β-SiAl ON ablative composite. Here, we present the inverse solution technique to estimate the parameter used to identify the effective thermal diffusivity. The transient temperature measurements, taken from the top surface and through the thickness of the samples in oxyacetylene flame test, are used in the inverse analysis. The results of this study clarify the mechanism of the ablation and thermal diffusivity of the β-SiAl ON-based composite heat shields due to the ultra high temperature ablation. [ABSTRACT FROM AUTHOR]

For atmospheric reentries at high enthalpies ablative heat shield materials are used, such as those for probes entering the atmosphere of Saturn's moon Titan, such as Cassini-Huygens in December, 2004. The characterization of such materials in a nitrogen/methane atmosphere is of interest. A European ablative material, AQ60, has been investigated in plasma wind tunnel tests at the 1RS plasma wind tunnel PWK1 using the magnetoplasma dynamic generator RD5 as plasma source in a nitrogen/methane atmosphere. The dimensions of the samples are 45 mm in length with a diameter of 39 mm. The actual ablator has a thickness of 40 mm. The ablator is mounted on an aluminium substructure. The experiments were conducted at two different heat flux regimes, 1.4 MW/m2 and 0.3MW/m2. In this paper, results of emission spectroscopy at these plasma conditions in terms of plasma species' temperatures will be presented, including the investigation of the free-stream species, N2 and N2+, and the major erosion product C2, at a wavelength range around 500 nm-600 nm. [ABSTRACT FROM AUTHOR]

Phenolic impregnated carbon ablator was the heat-shield material for the Stardust probe and was considered as a candidate heat-shield material for the Orion Crew Module. As part of the heat-shield qualification for Orion, physical and thermal properties were measured for newly manufactured material, including emissivity, heat capacity, thermal conductivity, elemental composition, and thermal decomposition rates. Based on these properties, an ablation and thermal-response model was developed for temperatures up to 3500 K and pressures up to 100 kPa. The model includes transversely isotropic and pressure-dependent thermal conductivity. In this work, model validation is accomplished by comparison of predictions with data from many arcjet tests conducted over a range of stagnation heat flux and pressure from 107 W/cm² at 2.3 kPa to 1100 W/cm² at 84 kPa. Over the entire range of test conditions, model predictions compare well with measured recession, maximum surface temperatures, in-depth temperatures, and char depth. [ABSTRACT FROM AUTHOR]

Carbon-phenolic ablative materials represent the most effective solution as thermal shield for a broad kind of space missions which require a ballistic reentry in atmosphere. These materials were successfully employed for shielding probes during the reentry from the Moon and landing on Mars. Nowadays further improvements are required and the addition of nano-fillers is a promising innovation strategy. In this work a carbon-phenolic ablative material with a density of 0.5 g/cm3 was manufactured with and without the addition of nano-Al 2 O 3. Its thermo-chemical behavior was evaluated with thermogravimetric analysis and kinetic parameters were calculated according to Ozawa, Friedman and Kissinger methods. The differences in ablation performance were enlightened by oxyacetylene flame tests. The conducted analyses show that nano-Al 2 O 3 can modify the chemical behavior of the ablative material and can improve the ablative performance. • Two different kind of carbon-phenolic ablators were manufactured with and without the addition of nano-Al 2 O 3. • Thermogravimetric analysis were carried out and kinetic parameters were calculated by isoconversion methods • Samples were tested by an oxyacetylene flame torch and, during the test, the back and frontal temperature were recorded. • Samples with nano-Al 2 O 3 addition show some difference in the kinetic parameters and better ablative performances. [ABSTRACT FROM AUTHOR]