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All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

Temperature and Frequency Dependent Electrical Properties of AVCOAT and Shuttle Tile Material

This technical memorandum presents the data and results from a series of tests performed on AVCOAT, charred AVCOAT, and different Shuttle Thermal Protection System (TPS) tile coatings to determine their electrical properties as a function of temperature and frequency. Initially, the samples were assumed to be insulators and the complex permittivities were measured from data collected by sandwiching each sample between two conductors to form a sandwich. However, the AVCOAT sample data showed that the material becomes conductive as it begins to char. To get an accurate conductivity reading on these charred samples, a four-point resistance probe was used to collect data. Overall, this testing showed that the virgin AVCOAT starts out as a material that is a good insulator but transitions to one with a conductivity on the order of 1860 S/m as it chars. The deeper layers of AVCOAT (un-charred) all exhibited a temperature and frequency dependence that is possibly indicative of a dielectric relaxation in the material. As the sample temperatures were increased, a knee appears in the real permittivity data sets below 1 kHz. Both the real permittivity and resistivity drop up to several orders of magnitude yet show that the material remains an insulator. The Shuttle High Temperature Reusable Surface Insulation (HRSI) tile coatings studied included Reac- tion Cured Glass (RCG), Aluminum Enhanced Thermal Barrier (AETB), and Toughened Unipiece Fibrous Insulation (TUFI). The purely glass fiber parts of the tile were removed leaving only the outer layers con- taining just the coating. The data collected shows that the RCG, AETB, and outermost layer of the TUFI have similar electrical properties. The real permittivity plots show that there appears to be a relaxation that shows up at frequencies below 1 kHz. The relaxation frequency increases with increasing temperature. Also, the real permittivity increases up to about a factor of four between room temperature and 650K. Likewise, the resistivity (related to the imaginary permittivity) decreases up to five orders of magnitude over the same temperature interval, but still remains an insulator. This data was collected to support a NASA Engineering and Safety Center (NESC) assessment team investigating the source of thermocouple anomalies in both Shuttle TPS material and AVCOAT used for on the Orion spacecraft for Exploration Flight Test 1 (EFT-1). The details of their findings will be documented in a separate report with this document serving as a reference.

Mark A Nurge↗
NASA NTRS2014

AVCOAT Density Characterization for Orion Multi-Purpose Crew Vehicle

The Orion Multi-Purpose Crew Vehicle (MPCV) will transport four crew members to and from lunar-class orbital destinations. The first orbital Exploration Flight Test (EFT-1) is scheduled for December 2014 and will provide valuable data on several systems, including the heat shield. The heat shield material is AvcoatTM, a mid-density ablator. Specifically, the heat shield contains a fiberglass-phenolic honeycomb structure filled with an ablative epoxy novolac resin. Post-flight characterization of the EFT-1 thermal protection system will be conducted in order to study the material response and measure the char front of the material. Avcoat density profiling will be conducted to study three critical material zones: char, pyrolysis, and virgin states as a function of material depth. As part of a ground test campaign, Avcoat coupons are tested at the NASA Ames Research Center Arc Jet Complex. The test campaign is set forth to study material response based on environments, perform Avcoat material density characterization, and compare the char depths from the HEAT sensor and density profiling studies. These investigations will then be compared to flight data in order to improve upon TPS material response models.

AVCOAT↗
NASA NTRS2016

Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1

On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer-by-layer density measurements and SEM investigations to evaluate material morphology at microstructural level including identification of elements and compounds.

Aerothermodynamics↗
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 NTRS2019

Laser-Enhanced Arc-Jet Facility Wedge Tests: Avcoat Material Performance Under Convective and Radiative Heating Environments

This paper presents the first set of experimental results from Laser Enhanced Arc-Jet Facility (LEAF-Lite) tests that were conducted shortly after the radiative LEAF-Lite system was added to the 60-MW Interaction Heating Facility at NASA Ames Research Center. Results were gathered to characterize the new radiative and combined heating capabilities as well as the convective heating resulting from the new IHF nozzle that was required for combined heating operations. Tests were ultimately conducted at several combinations of radiative and convective heating prompted by the need to understand the effect of combined heating on the Orion heatshield material prior to pursuing combined heating tests of the more complex block architecture.

Laser-Enhanced Arc-Jet Facility↗
NASA NTRS2015

Thermochemical Ablation Analysis of the Orion Heatshield

The Orion Multi-Purpose Crew Vehicle will one day carry astronauts to the Moon and beyond, and Orion's heatshield is a critical component in ensuring their safe return to Earth. The Orion heatshield is the structural component responsible for absorbing the intense heating environment caused by re-entry to Earth's atmosphere. The heatshield is primarily composed of Avcoat, an ablative material that is consumed during the re-entry process. Ablation is primarily characterized by two processes: pyrolysis and recession. The decomposition of in-depth virgin material is known as pyrolysis. Recession occurs when the exposed surface of the heatshield reacts with the surrounding flow. The Orion heatshield design was changed from an individually filled Avcoat honeycomb to a molded block Avcoat design. The molded block Avcoat heatshield relies on an adhesive bond to keep it attached to the capsule. In some locations on the heatshield, the integrity of the adhesive bond cannot be verified. For these locations, a mechanical retention device was proposed. Avcoat ablation was modelled in CHAR and the in-depth virgin material temperatures were used in a Thermal Desktop model of the mechanical retention device. The retention device was analyzed and shown to cause a large increase in the maximum bondline temperature. In order to study the impact of individual ablation modelling parameters on the heatshield sizing process, a Monte Carlo simulation of the sizing process was proposed. The simulation will give the sensitivity of the ablation model to each of its input parameters. As part of the Monte Carlo simulation, statistical uncertainties on material properties were required for Avcoat. Several properties were difficult to acquire uncertainties for: the pyrolysis gas enthalpy, non-dimensional mass loss rate (B´c), and Arrhenius equation parameters. Variability in the elemental composition of Avcoat was used as the basis for determining the statistical uncertainty in pyrolysis gas enthalpy and B´c. A MATLAB program was developed to allow for faster, more accurate and automated computation of Arrhenius reaction parameters. These parameters are required for a material model to be used in the CHAR ablation analysis program. This MATLAB program, along with thermogravimetric analysis (TGA) data, was used to generate uncertainties on the Arrhenius parameters for Avcoat. In addition, the TGA fitting program was developed to provide Arrhenius parameters for the ablation model of the gap filler material, RTV silicone.

Sixel, William↗
NASA NTRS2016

Thermal Management and Thermal Protection Systems

During my internship in the Thermal Design Branch (ES3), I contributed to two main projects: i) novel passive thermal management system for future human exploration, ii) AVCOAT undercut thermal analysis. i) As NASA prepares to further expand human and robotic presence in space, it is well known that spacecraft architectures will be challenged with unprecedented thermal environments. Future exploration activities will have the need of thermal management systems that can provide higher reliability, mass and power reduction and increased performance. In an effort to start addressing the current technical gaps the NASA Johnson Space Center Passive Thermal Discipline has engaged in technology development activities. One of these activities was done through an in-house Passive Thermal Management System (PTMS) design for a lunar lander. The proposed PTMS, functional in both microgravity and gravity environments, consists of three main components: a heat spreader, a novel hybrid wick Variable Conductance Heat Pipe (VCHP), and a radiator. The aim of this PTMS is to keep electronics on a vehicle within their temperature limits (0 and 50 C for the current design) during all mission phases including multiple lunar day/night cycles. The VCHP was tested to verify its thermal performance. I created a thermal math model using Thermal Desktop (TD) and analyzed it to predict the PTMS performance. After testing, the test data provided a means to correlate the thermal math model. This correlation took into account conduction and convection heat transfer, representing the actual benchtop test. Since this PTMS is proposed for space missions, a vacuum test will be taking place to provide confidence that the system is functional in space environments. Therefore, the model was modified to include a vacuum chamber with a liquid nitrogen shroud while taking into account conduction and radiation heat transfer. Infrared Lamps were modelled and introduced into the model to simulate the sun's rays directly impinging on the system. Heating rate of the lamps were calculated by knowing fraction of emitted energy in a wavelength interval and the filament temperature. This version of the model can be used to predict performance of the system under vacuum with extreme cold or hot conditions. Initial testing of the PTMS showed promise, and the thermal math model predicts even better performance in thermal vacuum testing. ii) Thermal Protection Systems (TPS) are required for vehicles which enter earth's atmosphere to protect from aerodynamic heating caused by the friction between the vehicle and atmospheric gases. Orion's heat shield design has two aspects which needed to be analyzed thermally: i) a small excess of adhesive used to bond the outer AVCOAT layer to the inner composite structure tends to seep from under the AVCOAT and form a small bead in between two bricks of AVCOAT, ii) a silicone rubber with different thermophysical properties than AVCOAT fills the gap between two bricks of AVCOAT. I created a thermal model using TD to determine temperature differences that are caused by these two features. To prevent false results, all TD models must be verified against something known. In this case, the TD model was correlated to CHAR, an ablation modelling software used to analyze TPS. Analyzing a node far from the concerning features, we saw that the TD model data match CHAR data, verifying the TD model. Next, the temperature of the silicone rubber as well as the bead of adhesive were analyzed to determine if they exceeded allowable temperatures. It was determined that these two features do not have a significant effect on the max temperature of the heat shield. This model can be modified to check temperatures at various locations of the heat shield where the composite thickness varies.

Hasnain, Aqib↗
NASA NTRS2008

NASA Crew Exploration Vehicle, Thermal Protection System, Lessons Learned

The Orion (CEV) thermal protection system (TPS) advanced development project (ADP) was initiated in late 2006 to reduce developmental risk by significant investment in multiple heat shield architectural solutions that can meet the needs both the Low Earth orbit (LEO) and Lunar return missions. At the same time, the CEV TPS ADP was also charged with developing a preliminary design for the heat shield to meet the PDR requirement and at the time of the PDR, transfer the design to Lockheed- Martin, the prime contractor. We reported on the developmental activities of the first 18 months at the IPPW5 in Bordeaux, France, last summer. In June 08, at the time of the IPPW6, the CEV TPS ADP would have nearly completed the preparation for the Orion PDR and would be close to the original three-year mark. We plan to report on the progress at the Atlanta workshop. In the past year, Orion TPS ADP investment in TPS Technology, especially in PICA ablative Heat-shield design, development, testing and engineering (DDTE) has paid off in enabling MSL mission to switch from SLA 561 V heat shield to PICA heat shield. CEV TPS ADP considered SLA 561 V as a candidate for LEO missions and our testing identified failure modes in SLA and as a result, we dropped SLA for further evaluation. This close synergy between two projects is a highly visible example of how investment in technology areas can and does benefit multiple missions. In addition, CEV TPS ADP has been able to revive the Apollo ablative system namely AVCOAT honeycomb architecture as an alternate to the baseline PICA architecture and we plan to report the progress we have made in AVCOAT. CEV TPS ADP has invested considerable resources in developing analytical models for PICA and AVCOAT, material property measurements that is essential to the design of the heat-shield, in arcjet testing, in understanding the differences between different arc jet facilities, namely NASA Ames, NASA JSC and Air Force's AEDC, and in Non-Destructive Evaluation (NDE), and in integration of and manufacturing heat shield as a system. The capabilities of the two heat shield systems including failure modes via testing and analysis, once established, can serve the Probe Community and future mission designers to inner and outer planetary exploration very well. For example, missions to Venus, Mars and Titan can use either one of the system by selecting the mission design parameters that utilizes the full characteristics of these system to make use of system efficiency that will result in reduced heat shield mass, system robustness that will enhance mission success and cost. We plan to present significant progresses of the past three years and highlight the significant contributions CEV TPS ADP Project has made to advance the state of the art in Thermal Protection System technology that has and will continue to benefit future entry probe missions.

NASA Crew Exploration Vehicle↗
NASA NTRS2018

Thermal Protection Systems Technology Transfer from Apollo and Space Shuttle to the Orion Program

This paper describes how the Orion program is utilizing the Thermal Protection System (TPS) experience from the Apollo and Space Shuttle programs to reduce program risk and improve affordability to meet NASA's future manned exploration missions. The Orion program successfully completed the Exploration Flight Test (EFT-1) mission in 2014 and is currently assembling, integrating, and testing the next spacecraft for the Exploration Mission (EM-1) to meet the flight test objectives of an unmanned orbital mission to the moon and return to earth in 2019. The Orion spacecraft production operations are located in the Neil Armstrong Operations and Checkout (O&C) facility at the Kennedy Space Center (KSC) providing an affordable and seamless delivery approach of vehicles directly to the launch site eliminating spacecraft transportation and additional checkout testing. Innovative vehicle design, manufacturing and test operations approaches are maturing and evolving with each Orion vehicle build to support the challenging NASA exploration mission requirements beyond Low Earth Orbit (LEO) while reducing program cost and schedule impacts. An example of Orion's evolution is the incorporation of an improved heat shield design, assembly and testing approach to meet the higher re-entry velocities for a lunar return for the EM-1 mission. The EFT- 1 heat shield was based on the Apollo heat shield manufacturing processes and was assembled at a supplier location and then transported to KSC for final integration. The EM-1 heat shield is now manufactured, assembled, tested, and installed into the spacecraft at the O&C facility reducing program cost and production schedules. The transition of the Space Shuttle TPS capabilities has enabled Orion to provide a human rated capsule design using proven materials and processes established over years of orbiter re-entry missions. The Orion Crew Module (CM) TPS configuration is derived from the Apollo CM approach utilizing improved materials and processes developed from the Space Shuttle program. The Orion EFT-1 heat shield utilized the Avcoat ablative material from Apollo which was injected into a honeycomb substrate and has been updated for EM-1 incorporating a block configuration bonded to a composite shell structure. This approach utilizes the proven Avcoat material for the heat shield ablator and is utilizing derived bonding and inspection methods and techniques from the Space Shuttle tile experience. The Orion back shell TPS configuration is based on Space Shuttle tile designs using proven tile materials and coatings. The Orion forward bay cover utilizes the high temperature tiles similar to the back shell tiles and low temperature blankets derived from the Space Shuttle program reducing weight impacts. Space Shuttle Multi-Layer Insulation (MLI) is installed in the Orion capsule to control the interior temperature environment providing a light weight design. These TPS design approaches have performed successfully on the Orion first flight test on EFT-1 and are incorporated in the configuration for the next flight test for EM-1. Completion of these two flight tests will certify the TPS for the Orion program for human rated exploration missions and has reduced the development cost to the Orion program. In addition to transitioning the Space Shuttle TPS design configurations to Orion, the supporting manufacturing infrastructure, manufacturing processes, and inspection methods are also incorporated into the Orion assembly operations at KSC. This has avoided significant startup schedule and costs impacts of new capabilities and development of support operations necessary to fabricate, install, inspect, and validate the TPS installations for the Orion spacecraft. The Thermal Protection System Facility (TPSF) which supported the Space Shuttle program provides the heat shield Avcoat blocks and the back shell AETB-8 tiles for Orion program. Extending the existing Space Shuttle TPS operations at KSC provides onsite support to the Orion spacecraft assembly operations. A significant benefit to the Orion program is the transition of the Space Shuttle technician work force at KSC. This highly skilled workforce was able to transition to the Orion program and immediately support the TPS installation operations. They brought with them the human rated manufacturing and assembly operations culture that was critical to the Space Shuttle success avoiding the retraining cost and schedule impacts of a new workforce to Orion. The technician skills certification program from the Space Shuttle was found to be compatible with the Lockheed Martin technician certification programs enabling a seamless training process utilizing process similarity avoiding the cost of retraining the work force to Orion standards. The existing workforce is certificated for multiple processes enabling reassignment of technicians to other tasks in the Orion AI&P operations reducing the overall touch labor manpower requirements. Additional benefits of using the Space Shuttle workforce is the incorporation of the lessons learned from Space Shuttle processes to improve the Orion TPS processes. This has resulted in a continuous evolution of TPS processes to improve the producibility and reduce the program cost for the TPS for the Orion program. Transitioning the Apollo and Space Shuttle TPS designs, processes, and technician workforce has been instrumental in enabling Orion to successfully meet the program challenges for NASA's exploration missions of the future.

Stewart, Michael↗
NASA NTRS2018

Determination of Uncertainties for Analytically Derived Material Properties to Be Used in Monte Carlo Based Orion Heatshield Sizing

Ablative materials are often used for spacecraft heatshields to protect underlying structures from the extreme environments associated with atmospheric reentry. NASA's Orion EM-1 capsule has been designed to use a molded Avcoat material system. In order to determine the required heatshield thickness, a Monte Carlo approach to the sizing process was proposed. To perform the Monte Carlo simulation, statistical uncertainties on all material property input parameters were required. Obtaining these values for measured properties is straightforward, however input parameters that are derived analytically have historically used uncertainties based on engineering judgment. A MATLAB program was created to use laboratory generated thermogravimetric analysis (TGA) data to calculate uncertainties on the Arrhenius parameters for molded Avcoat. Uncertainties associated with the normalized ablation rate and pyrolysis gas enthalpy were also generated using a wrapper script and the ACE code. These uncertainties could then be tied directly to measured values of individual elemental constituents. The resulting uncertainty values will allow for a probabilistic sizing approach on molded Avcoat with a higher level of confidence in the input parameters.

material property uncertainties↗
NASA NTRS1972

Demonstration of the range over which the Langley Research Center digital computer charring ablation program (CHAP) can be used with confidence: Comparisons of CHAP predictions and test data for three ablation materials

Comparisons of ablation calculations with the charring ablation computer code and ablation test data are presented over a wide range of environmental conditions in air for three materials: low-density nylon phenolic, Avcoat 5026-39HC/G, and a filled silicon elastomer. Heat fluxes considered range from over 500 Btu/sq ft-sec to less than 50 Btu/sq ft-sec. Pressures range from 0.5 atm to .004 atm. Enthalpies range from about 2000 Btu/lb to 18000 Btu/lb. Predictions of recession, pyrolysis penetration, and thermocouple responses are considered. Recession predictions for nylon phenolic are good as steady state is approached, but strongly transient cases are underpredicted. Pyrolysis penetrations and thermocouple responses are very well predicted. Recession amounts for Avcoat and silicone elastomer are less well predicted, although high heat flux cases near steady state are fairly satisfactory. Pyrolysis penetrations and thermocouple responses are very well predicted.

Moyer, C. B.↗
NASA NTRS2013

Testing of Advanced Conformal Ablative TPS

In support of the CA250 project, this paper details the results of a test campaign that was conducted at the Ames Arcjet Facility, wherein several novel low density thermal protection (TPS) materials were evaluated in an entry like environment. The motivation for these tests was to investigate whether novel conformal ablative TPS materials can perform under high heat flux and shear environment as a viable alternative to rigid ablators like PICA or Avcoat for missions like MSL and beyond. A conformable TPS over a rigid aeroshell has the potential to solve a number of challenges faced by traditional rigid TPS materials (such as tiled Phenolic Impregnated Carbon Ablator (PICA) system on MSL, and honeycomb-based Avcoat on the Orion Multi Purpose Crew Vehicle (MPCV)). The compliant (high strain to failure) nature of the conformable ablative materials will allow better integration of the TPS with the underlying aeroshell structure and enable monolithic-like configuration and larger segments to be used in fabrication.A novel SPRITE1 architecture, developed by the researchers at NASA Ames was used for arcjet testing. This small probe like configuration with 450 spherecone, enabled us to test the materials in a combination of high heat flux, pressure and shear environment. The heat flux near the nose were in the range of 500-1000 W/sq cm whereas in the flank section of the test article the magnitudes were about 50 of the nose, 250-500W/sq cm range. There were two candidate conformable materials under consideration for this test series. Both test materials are low density (0.28 g/cu cm) similar to Phenolic Impregnated Carbon Ablator (PICA) or Silicone Impregnated Refractory Ceramic Ablator (SIRCA) and are comprised of: A flexible carbon substrate (Carbon felt) infiltrated with an ablative resin system: phenolic (Conformal-PICA) or silicone (Conformal-SICA). The test demonstrated a successful performance of both the conformable ablators for heat flux conditions between 50-500 W/sq cm. The recession and temperature profile for these materials were comparable to PICA proving them to be viable alternatives for TPS technology development for future missions.

Gasch, Matthew↗
NASA NTRS2018

Orion Heat Shield Manufacturing Producibility Improvements for the EM-1 Flight Test Program

This paper describes how the ORION program is incorporating improvements in the heat shield design and manufacturing processes reducing programmatic risk and ensuring crew safety in support of NASA's Exploration missions. The approach for the EFT-1 heat shield utilized a low risk Apollo heritage design and manufacturing process using an Avcoat TPS ablator with a honeycomb substrate to provide a one piece heat shield to meet the mission re-entry heating environments. The EM-1 mission will have additional flight systems installed to fly to the moon and return to Earth. Heat shield design and producibility improvements have been incorporated in the EM-1 vehicle to meet deep space mission requirements. The design continues to use the Avcoat material, but in a block configuration to enable improvements in consistant and repeatable application processes using tile bonding experience developed on the Space Shuttle Transportation System Program.

Producibility↗
NASA NTRS2019

Orion Heat Shield Manufacturing Producibility Improvements for the EM-1 Flight Test Program

This paper describes how the Orion program is incorporating improvements in the heat shield design and manufacturing processes reducing programmatic risk and ensuring crew safety in support of NASA’s Exploration Missions. The Orion program successfully completed the EFT-1 flight test in 2014 and is currently developing the EM-1 spacecraft to meet the test objectives of an orbital mission to the moon and return to earth in 2019. Lessons learned from the EFT-1 manufacturing and flight test experience are being incorporated into a wide variety of vehicle systems and manufacturing processes to reduce risk to the Orion missions and flight crew. A critical contributor to crew safety is the heat shield that protects the crew capsule during re-entry through the earth’s atmosphere for return from deep space. The first flight test vehicle, EFT-1, was manufactured and tested in the Neil Armstrong Operations and Checkout (O&C) facility at KSC to demonstrate early risk reduction including the functionality of the Thermal Protection Systems (TPS) for capsule safe return to earth. The approach for the EFT-1 heat shield utilized a low risk approach using Apollo heritage design and manufacturing processes using an Avcoat TPS ablator with a honeycomb substrate to provide a one piece heat shield to meet the mission re-entry heating environments. The manufacturing processes used honeycomb cell injection guns which were redeveloped from the Apollo Lunar Program processes to build the EFT-1 heat shield. The completed heat shield was transported across the country by aircraft to the O&C at KSC for installation onto the capsule. The EFT-1 heat shield successfully performed its mission and experienced ~80% of the re-entry velocity (50% heating rate) for a lunar return for an Exploration Mission. The second flight test vehicle is the EM-1 mission which will have additional flight systems installed to fly to the moon and return. Heat shield design and producibility improvements have been incorporated in the EM-1 vehicle to meet deep space mission and programmatic requirements. The design continues to use the Avcoat material, but in a “block” configuration to enable improvements in the application processes as well as additional improvements in the carrier structure design and manufacturing operations. Incorporating flight test results and producibility improvements from EFT-1 for the heat shield system design and processes have improved the thermal protection capability, improved the producibility, and cost for EM-1 flight test.

Producibility↗
NASA NTRSDate not supplied

Under Pressure: The Artemis I Heatshield Char Loss Investigation and Artemis II’s Successful Entry

Following the successful Artemis I skip reentry on December 11, 2022, unexpected char liberation from the Orion heat- shield prompted formation of an Anomaly Response Team to determine root cause and establish flight rationale and corrective actions for subsequent missions. Through a comprehensive investigation including detailed hardware analysis, modeling & simulation, and ground testing, the team determined that the low permeability Avcoat experienced extreme internal gas pressure buildup during the skip entry that could not adequately outgas, leading to crack formation and char liberation. Key discoveries included significant performance differences between permeable and impermeable heatshield regions and successful replication of the anomaly through ground testing. Based on extensive ground testing and analysis, Artemis II flew a modified trajectory without the skip entry to minimize crack-inducing conditions, and successfully splashed down on April 10, 2026 with significantly reduced char loss. For Artemis III and beyond, a corrective action was implemented to produce a more permeable version of Avcoat that will substantially reduce internal pressure accumulation during entry.

Artemis II↗
NASA NTRS2016

Estimating Orion Heat Shield Failure Due To Ablator Cracking During The EFT-1 Mission

The Orion EFT-1 heatshield suffered from two major certification challenges: First, the mechanical properties used in design were not evident in the flight hardware and second, the flight article itself cracked during fabrication. The combination of these events motivated the Orion Program to pursue an engineering-level Probabilistic Risk Assessment (PRA) as part of heatshield certification rationale. The PRA provided loss of Mission (LOM) likelihoods considering the probability of a crack occurring during the mission and the likelihood of subsequent structure over-temperature. The methods and input data for the PRA are presented along with a discussion of the test data used to anchor the results. The Orion program accepted an EFT-1 Loss of Vehicle (LOV) risk of 1-in-160,000 due to in-mission Avcoat cracking based on the results of this analysis. Conservatisms in the result, along with future considerations for Exploration Missions (EM) are also addressed.

Avcoat↗
NASA NTRS1990

Aerobraking in a dusty Martian atmosphere

The effects of dust particle impacts on the erosion of the heat shield were estimated for a 26 m diameter aerobraking vehicle entering the Mars atmosphere at 8.6 km/sec. An explicit Navier-Stokes code was used to compute the flow field about the vehicle for the actual Martian atmospheric composition at a speed of 7.5 km/sec. The deceleration and melting of the dust particles within the forebody shock-layer was computed for dust spherules having initial diameters from 3 to 10 microns. Two heat shield materials were considered: Shuttle tiles with glassy surfaces and a low-density ablator known as Avcoat.

Chang, I-Dee↗