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NASA NTRS1988

In-flight rain damage tests of the shuttle thermal protection system

NASA conducted in-flight rain damage tests of the Shuttle thermal protection system (TPS). Most of the tests were conducted on an F-104 aircraft at the Dryden Flight Research Facility of NASA's Ames Research Center, although some tests were conducted by NOAA on a WP-3D aircraft off the eastern coast of southern Florida. The TPS components tested included LI900 and LI2200 tiles, advanced flexible reusable surface insulation, reinforced carbon-carbon, and an advanced tufi tile. The objective of the test was to define the damage threshold of various thermal protection materials during flight through rain. The test hardware, test technique, and results from both F-104 and WP-3D aircraft are described. Results have shown that damage can occur to the Shuttle TPS during flight in rain.

Meyer, Robert R., Jr.↗

NASA NTRS2003

Thermo Physics Facilities Branch Brochure ARC Jet Complex Fact Sheets, Hypervelocity Free-Flight Aerodynamic Facility Fact Sheets, Ames Vertical Gun Range Fact Sheets

The Ames Arc Jet Complex has a rich heritage of over 40 years in Thermal Protection System (TPS) development for every NASA Space Transportation and Planetary program, including Apollo, Space Shuttle, Viking, Pioneer-Venus, Galileo, Mars Pathfinder,Stardust, NASP,X-33,X-34,SHARP-B1 and B2,X-37 and Mars Exploration Rovers. With this early TPS history came a long heritage in the development of the arc jet facilities. These are used to simulate the aerodynamic heating that occurs on the nose cap, wing leading edges and on other areas of the spacecraft requiring thermal protection. TPS samples have been run in the arc jets from a few minutes to over an hour,from one exposure to multiple exposures of the same sample, in order t o understand the TPS materials response to a hot gas flow environment (representative of real hyperthermal environments experienced in flight). The Ames Arc l e t Complex is a key enabler for customers involved in the three major areas of TPS development: selection, validation, and qualification. The arc jet data are critical for validating TPS thermal models, heat shield designs and repairs, and ultimately for flight qualification.

Fretter, E. F.↗

NASA NTRS2020

Technologies for Future Venus Exploration

The purpose of this white paper is to provide an overview to the NRC Decadal Survey Inner Planets Sub-Panel on thermal protection system (TPS) technologies required for future Venus exploration missions. It considers the capability of heritage TPS technology used by the Pioneer Venus and Galileo probe missions and identifies new technologies that could enable greater science value and more ambitious missions in the future. A prime conclusion is that there are important advances regarding the availability of forebody TPS required for Venus entry probes. Specifically, the carbon phenolic flown on the Pioneer Venus probes has been deemphasized and a new 3D woven material system, Heatshield for Extreme Entry Environment Technology (HEEET) has replaced carbon phenolic. Along with the development of HEEET, there have been test facility upgrades and design tool improvements. These new material and test methods are enabling for Venus missions. However, without a mission scheduled in the near future, even these new developments are at risk of becoming unavailable. Therefore, we recommend that NASA invest in a cross-cutting technology program that focuses on sustainment of relevant TPS materials, entry systems, test facilities, design tools, and flight instrumentation.

Thermal Protection Systems↗

NASA NTRS1972

Orbiter entry trajectory considerations

Space shuttle trajectory-shaping optimization considering the vehicle's thermal environment and the resulting requirements of the thermal protection system (TPS) are discussed. Optimization studies have been conducted yielding results which are generally applicable to shuttle orbiters and are independent of evolution and redirection of the space shuttle program. Two investigations of optimal trajectory shaping, with different methods of considering the thermal environment are presented.

Rehder, J. J.↗

NASA NTRSDate not supplied

Fracture of Charring Avcoat With Meshfree Material Response-Coupled Fracture Approach

Fracture of thermal protection systems (TPS) is typically unfavorable, but often unavoidable. TPS can fracture during entry, manufacturing or from impact. Many NASA missions use ablative materials for TPS, including the fore-body for Dragonfly (PICA) and Orion for Artemis-I (Avcoat). The material response to thermal and mechanical loads during entry is integral to TPS sizing, design and analysis. The fracture of TPS also depends, and can be caused by, the change of material properties during entry. Simulations that resolve fracture of TPS due to thermo-mechanical forces can identify TPS failure mechanisms and be applied to conditions inaccessible to ground testing. We present simulation work using a mesh-free/Lagrangian approach to solving continuum mechanics, coupled to material response. Two different materials are simulated under different model entry-like boundary conditions. The dynamic crack structure is analyzed and compared across simulations. Crack analyses inform the role of design features, such as entry trajectories and manufacturing-influenced material properties, play in TPS fracture.

Andrew P Santos↗

NASA NTRSDate not supplied

Investigating Material Behavior in Atmospheric Entry Conditions: Arc-Jet Testing Insights from Meteorite Ablation to High-Temperature Coatings.

Arc-jet testing is an indispensable tool in elucidating the complex interactions materials undergo when subjected to the extreme thermal and mechanical stresses of atmospheric entry. By replicating these harsh conditions, arc-jets enable researchers to explore critical phenomena such as ablation, melting, and spallation, generating invaluable data that informs the development of sophisticated, physics-based models. These models are pivotal not only in predicting material performance for spacecraft re-entry systems but also in assessing the threat posed by celestial bodies as they encounter Earth’s atmosphere. One of the leading facilities for such investigations is the Hypersonic Materials Environmental Test System (HyMETS) at NASA’s Langley Research Center. A pioneering test campaign conducted at HyMETS focused on unraveling the ablation mechanisms of an ordinary chondrite meteorite (Tamdakht H5) and a terrestrial analog (basalt). These studies unveiled distinct material behaviors: Tamdakht demonstrated a remarkably stable melt flow, with mass loss primarily driven by the volatilization of elements such as iron and sodium, whereas basalt exhibited more aggressive surface degradation due to the rapid decomposition of hydrated minerals, culminating in significant spallation and surface material ejection. Beyond meteorite analysis, HyMETS has also been integral in the exploration of advanced Thermal Protection Systems (TPS) such as Phenolic Impregnated Carbon Ablator (PICA), a material with extensive flight heritage in missions like Stardust, Mars Science Laboratory, Mars 2020, and Osiris-REx. PICA, when coated with NuSil CV-1144-0—a polysiloxane resin designed to prevent particle shedding—undergoes a remarkable transformation upon heating. The resin pyrolyzes to form a thin, oxidation-resistant silicon oxycarbide layer that profoundly influences the material’s thermal response through a sophisticated four-stage process. Initially, the silicon oxycarbide acts as a formidable barrier, impeding reactive interactions with the PICA char and effectively suppressing surface temperature rise and material recession. However, as the heating persists, the protective layer decomposes via carbothermal reduction, exposing the underlying char, which in turn leads to a dramatic spike in surface temperature and accelerated material erosion. In the final stage, equilibrium is reached, with recession rates aligning with those of the virgin material. This intricate understanding of PICA-NuSil behavior under extreme conditions offers valuable insights, enhancing material response models and advancing the development of next-generation TPS for space exploration.

Chemistry↗

NASA NTRS2019

Overview of Heatshield for Extreme Entry Environment Technology (HEEET)

The objective of the Heatshield for Extreme Entry Environment Technology (HEEET) projects is to mature a 3-D Woven Thermal Protection System (TPS) to Technical Readiness Level (TRL) 6 to support future NASA missions to destinations such as Venus and Saturn. Destinations that have extreme entry environments with heat fluxes > 3500 W/cm2 and pressures up to 5 atmospheres, entry environments that NASA has not flown since Pioneer-Venus and Galileo. The scope of the project is broad and can be split into roughly four areas, Manufacturing/Integration, Structural Testing and Analysis, Thermal Testing and Analysis and Documentation. Manufacturing/Integration covers from raw materials, piece part fabrication to final integration on a 1-meter base diameter 45-degree sphere cone Engineering Test Unit (ETU). A key aspect of the project was to transfer as much of the manufacturing technology to industry in preparation to support future mission infusion. The forming, infusion and machining approaches were transferred to Fiber Materials Inc. and FMI then fabricated the piece parts from which the ETU was manufactured. The base 3D-woven material consists of a dual layer weave with a high-density outer layer to manage recession in the system and a lower density, lower thermal conductivity inner layer to manage the heat load. At the start of the project it was understood that due to weaving limitations the heat shield was going to be manufactured from a series of tiles. And it was recognized that the development of a seam solution that met the structural and thermal requirements of the system was going to be the most challenging aspect of the project. It was also recognized that the seam design would drive the final integration approach and therefore the integration of the ETU was kept in-house within NASA. A final seam concept has been successfully developed and implemented on the ETU. The structural testing and analysis covers from characterization of the different layers of the infused material as functions of weave direction and temperature to sub-component level testing such as 4pt bend testing at sub-ambient and elevated temperature and culminates in testing of the ETU results from which will validate the structural models developed using the element and sub-component level tests. Given the seam has to perform both structurally and aerothermally during entry a novel 4pt bend test fixture was developed allowing articles to be tested while the front surface is heated with a laser. These tests are being utilized to establish the systems structural capability during entry.A broad range of aerothermal tests (arcjet tests) were performed to develop material response models for predicting the required TPS thickness to meet a missions needs and to evaluate failure modes and establish the capability of the system. The final aspect of the project is to develop a comprehensive Design and Data Book such that a future mission will have the information necessary to adopt the technology. This presentation will provide an overview for each of these areas and argue that HEEET has successfully achieved TRL 6.

HEEET↗

NASA NTRS1972

Thermal design of composite materials high temperature attachments

The thermal aspects of using filamentary composite materials as primary airframe structures on advanced atmospheric entry spacecraft such as the space shuttle vehicle were investigated to identify and evaluate potential design approaches for maintaining composite structures within allowable temperature limits at thermal protection system (TPS) attachments and/or penetrations. The investigation included: (1) definition of thermophysical data for composite material structures; (2) parametric characterization and identification of the influence of the aerodynamic heating and attachment design parameters on composite material temperatures; (3) conceptual design, evaluation, and detailed thermal analyses of temperature limiting design concepts; and (4) the development of experimental data for assessment of the thermal design methodologies and data used for evaluation of the temperature-limiting design concepts. Temperature suppression attachment concepts were examined for relative merit. The simple isolator was identified as the most weight-effective concept and was selected for detail design, thermal analysis, and testing. Tests were performed on TPS standoff attachments to boron/aluminum, boron/polyimide and graphite/epoxy composite structures.

Source record↗

NASA NTRS1999

Aerothermal Test of Thermal Protection Systems for X-33 Reusable Launch Vehicle

An array of metallic Thermal Protection System (TPS) panels developed for the windward surface of the X-33 vehicle was tested in the 8-Foot High Temperature Tunnel at the NASA Langley Research Center. These tests were the first aerothermal tests of an X-33 TPS array and the test results will be used to validate the TPS for the X-33 flight program. Specifically, the tests evaluated the structural and thermal performance of the TPS, the effectiveness of the high temperature seals between adjacent panels and the durability of the TPS under realistic aerothermal flight conditions. The effect of varying panel-to-panel step heights, intentional damage to the seals between adjacent panels, and the use of secondary seals were also investigated during the test program. The metallic TPS developed for the windward surface of the X-33, the blanket TPS developed to protect the leeward surfaces of the X-33, and the test program in the 8-Foot High Temperature Tunnel are presented and discussed.

Sawyer, James Wayne↗

NASA NTRS1998

Orbital Debris Impact Damage to Reusable Launch Vehicles

In an effort by the National Aeronautics and Space Administration (NASA), hypervelocity impact tests were performed on thermal protection systems (TPS) applied on the external surfaces of reusable launch vehicles (RLV) to determine the potential damage from orbital debris impacts. Three TPS types were tested, bonded to composite structures representing RLV fuel tank walls. The three heat shield materials tested were Alumina-Enhanced Thermal Barrier-12 (AETB-12), Flexible Reusable Surface Insulation (FRSI), and Advanced Flexible Reusable Surface Insulation (AFRSI). Using this test data, predictor equations were developed for the entry hole diameters in the three TPS materials, with correlation coefficients ranging from 0.69 to 0.86. Possible methods are proposed for approximating damage occurring at expected orbital impact velocities higher than tested, with references to other published work.

Robinson, Jennifer H.↗

NASA NTRS2020

Thermal Protection System Materials for Sample Return Missions

This white paper from the Thermal Protection System (TPS) community to the NRC Decadal Survey Sub-Panels provides an overview of TPS materials needed for future Sample Return (SR) missions. We consider the capability of heritage TPS material used by recent SR missions and identify appropriate materials for future SR missions. A prime conclusion is that the current TPS materials, if properly maintained, offer good low-density solutions for lower velocity (<13.5 km/s) sample return missions without planetary protection back-protection concerns. Furthermore, missions that have a larger capsule, a higher entry speed (>13.5 km/s), or back-protection concerns will leverage recently developed mid-density TPS materials. To maintain NASA’s Sample Return capabilities in the coming decade, we recommend that NASA continue to invest in sustainment of relevant TPS materials, as well as ground-test facilities, predictive entry modeling, and flight instrumentation.

mission design↗

NASA NTRS1982

Space Shuttle Orbiter - Reusable surface insulation flight performance

The first two flights of the Space Shuttle Orbiter have provided the initial data required for operational certification of the Thermal Protection System (TPS). The flight performance characteristics of the TPS reusable surface insulation (RSI) will be discussed. The discussion will be based on post-flight inspections of the RSI and post-flight interpretations of the flight instrumentation data. The flights to date indicate that the thermal and mechanical design requirements for the RSI system were met or exceeded.

Dotts, R. L.↗

NASA NTRS1983

The Shuttle Orbiter thermal protection system materials, designs, and flight performance overview

The design requirements for the Orbiter thermal protection system (TPS), the various TPS materials that are used, the different design approaches associated with each of the materials, and the performance experienced during the flight test program are described. The first five flights of the Orbiter Columbia have provided the necessary data to verify the TPS thermal performance, structural integrity, and reusability. The flight performance characteristics of each TPS material are discussed. This discussion is based on postflight inspections and postflight interpretation of the flight instrumentation data. The flights to date indicate that the thermal and structural design requirements for the Orbiter TPS have been met and that the overall performance has been outstanding.

Dotts, R. L.↗

NASA NTRSDate not supplied

Computational Investigation of Oxidative Etch Pitting in FiberForm and Its Impact on Material Properties

Oxidation-driven carbon erosion does not occur uniformly but rather through the development of localized etch pits at active surface sites. These active sites form due to atomic defects on the carbon surface, making them significantly more reactive than the surrounding, non-defective areas. As a result, these sites are the first to react during ablation, leading to their removal. This process creates new defects in neighboring atoms, increasing their reactivity and causing localized carbon removal around these active sites. In this way, the highly reactive defective areas serve as nucleation points for the formation and growth of etch pits, which can have adverse effects on structural integrity of FiberForm. To better understand how these etch pits impact the material properties of carbon fiber microstructures, we have developed a new capability within the direct simulation Monte Carlo (DSMC) framework to capture the etch pit formation process. This capability, integrated into the DSMC code SPARTA (Stochastic Parallel Rarefied-gas Time-accurate Analyzer), models material removal in the presence of active sites, leading to the formation of etch pits. The current work focuses on studying the effects of these etch pits on the material properties of FiberForm, a widely used base material in thermal protection systems (TPS). The microstructure of virgin FiberForm, obtained via X-ray microtomography, is imported into SPARTA to generate the ablated geometries with etch pits. These modified microstructures are then analyzed using the Porous Microstructure Analysis (PuMA) software to compute various material properties, including elasticity, thermal conductivity, and permeability. We investigate the variation of these properties due to the complex surface topology changes caused by etch pit formation. Additionally, we compare the effects of pitting with the conventional model of shrinking fibers, traditionally used to simulate the ablation of carbon structures. Significant differences emerge between the two approaches. Consequently, this physically realistic model of material removal through etch pit formation offers improved accuracy in predicting the degradation of carbon-based TPS during oxidation. It also provides insights into other mechanisms, such as spallation, where chunks of material are removed into the flow due to etch pit growth. Ultimately, this model enhances our understanding of failure modes in these materials during ablation.

PuMA↗

NASA NTRSDate not supplied

Minimally Intrusive Single-Chip Recession/Temperature Sensors for Spacecraft Thermal Protection Systems

We report the initial results of ongoing development of minimally intrusive, single-chip integrated capacitive recession/temperature sensors that would provide real-time monitoring of ablation recession of heatshield thermal protection systems (TPS) during spacecraft atmospheric entry. This work, which was performed using Phenolic Impregnated Carbon Ablator (PICA), is also to demonstrate the ease of sensor insertion during TPS manufacturing, thereby resulting in faster, low-cost, uniform sensor assembly, and spatial distribution, without compromising TPS structural integrity. Using oxy-acetylene flame at 500 W-cm-2 heat flux for 16 seconds, the average recession depth of 3.25 mm was obtained. Some measurement problems were encountered in this first attempt. However, the initial results obtained are in relatively good agreement with published data, and offer promise of future improvement based on lessons learned.

Recession↗

NASA NTRS1975

Phase change paint tests to investigate effects of TPS tiles on heating rates of the Rockwell space shuttle orbiter (test OH4C, model 21-0)

Information and data from wind tunnel tests conducted on 0.0175-scale models of the space shuttle orbiter are presented. The primary objective of the tests was to evaluate aerodynamic heating effects of the tiles in the thermal protection system (TPS). Tile gap depth and flow orientation effects on the TPS were investigated. Tile patterns were cut into the undersides of the orbiter models to simulate the gaps. One model was left smooth for comparison.

Quan, M.↗

NASA NTRSDate not supplied

Revisiting MEDLI with Improved Modeling and Simulation Tools

The thermal protection system (TPS) of the Mars 2020 capsule will carry a suite of sensors called MEDLI2 (Mars Entry, Descent, and Landing Instrument 2). This suite consists of pressure sensors/transducers and thermocouple plugs meant to measure atmospheric properties, vehicle attitude, TPS performance, etc. While the suite is similar to that carried on the successful MSL (Mars Science Laboratory) mission in 2012, it now includes pressure sensors, thermocouple plugs, and radiometers on the backshell as well. An analysis plan is therefore necessary to help with the interpretation of flight measurements.

Atmospheric Entry↗

NASA NTRS1977

Orbiter Thermal Protection System Development

The development of the Space Shuttle Orbiter Thermal Protection System (TPS) is traced from concept definition, through technical development, to final design and qualification for manned flight. A sufficiently detailed description of the TPS design is presented to support an indepth discussion of the key issues encountered in conceptual design, materials development, and structural integration. Emphasis is placed on the unique combination of requirements which resulted in the use not only of revolutionary design concepts and materials, but also of unique design criteria, newly developed analysis, testing and manufacturing methods, and finally of an unconventional approach to system certification for operational flight. The conclusion is drawn that a significant advance in all areas of thermal protection system development has been achieved which results in a highly efficient, flexible, and cost-effective thermal protection system for the Orbiter of the Space Shuttle System.

Greenshields, D. H.↗