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NASA NTRS2014

Woven Thermal Protection System (WTPS) a Novel Approach to Meet Nasa's Most Demanding Reentry Missions

NASA's future robotic missions to Venus and other planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heatshield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic is a robust TPS, however, its high density and thermal conductivity constrain mission planners to steep entries, high fluxes, pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose certification challenges in existing ground based test facilities. In 2012 the Game Changing Development Program in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System to meet the needs of NASA's most challenging entry missions. This presentation will summarize the maturation of the WTPS project.

thermal protection↗

NASA NTRS2014

Woven Thermal Protection System (WTPS) a Novel Approach to Meet NASA's Most Demanding Reentry Missions

NASA's future robotic missions to Venus and outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of current mid-density ablators (PICA or Avcoat). Therefore mission planners assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic (CP) is a robust Thermal Protection System (TPS) however its high density and thermal conductivity constrain mission planners to steep entries, high heat fluxes, pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose certification challenges in existing ground based test facilities. In 2012 the Game Changing Development Program in NASA's Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASA's most challenging entry missions. This presentation will summarize maturation of the WTPS project.

carbon phenolic↗

NASA NTRS2016

Heatshield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASAs high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50 mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in its 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

Thermal Protection System↗

NASA NTRS2016

Heat-shield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASA's high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50% mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in it's 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to-date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

TPS↗

NASA NTRS2017

Ablative Thermal Protection Systems Fundamentals

This is a presentation of the fundamentals of ablative TPS materials for a short course at TFAWS 2017. It gives an overall description of what an ablator is, the equations that define it, and how to model it.

TPS↗

NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bondline temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

margins↗

NASA NTRS2018

A Reliability Comparison of Classical and Stochastic Thickness Margin Approaches to Address Material Property Uncertainties for the Orion Heat Shield

The Orion Thermal Protection System (TPS) margin process uses a root-sum-square approach with branches addressing trajectory, aerothermodynamics, and material response uncertainties in ablator thickness design. The material response branch applies a bond line temperature reduction between the Avcoat ablator and EA9394 adhesive by 60 C (108 F) from its peak allowed value of 260 C (500 F). This process is known as the Bond Line Temperature Material Margin (BTMM) and is intended to cover material property and performance uncertainties. The value of 60 C (108 F) is a constant, applied at any spacecraft body location and for any trajectory. By varying only material properties in a random (monte carlo) manner, the perl-based script mcCHAR is used to investigate the confidence interval provided by the BTMM. In particular, this study will look at various locations on the Orion heat shield forebody for a guided and an abort (ballistic) trajectory.

FIAT↗

NASA NTRS2018

Shock Layer Radiation Measurements for Planetary Probes

Shock layer radiation to a planetary probe's heat shield has been investigated since the 1960's, using ground tests, flight tests, and theoretical modelling. Radiometers and a spectrometer were embedded in the TPS of NASA's earlier Project FIRE II, Apollo 4 and 6, as well as PAET. PAET is particularly relevant to planetary probes since it successfully demonstrated detecting the atmospheric composition using shock-layer radiometry. The NASA Orion program and Mars 2020 include and propose for future use embedded small scale, low-mass radiometers and mini-spectrometers. Recent work used a specific COTS fiber-optic mini-spectrometer, selected for wide wavelength range for testing flexibility. In proof of concept tests, these mini-spectrometers detected the strongest Na and K spectral lines in both Arcjet test and Laser tests, with thermal radiation. Characterization and tests of the radiometer and spectrometer devices components is described.

radiation↗

NASA NTRS2019

TUFROC Thermal Protection System

Toughened Unipiece Fibrous Reinforced Oxidation-resistant Composite (TUFROC) is a tiled Thermal Protection System (TPS) suitable for reusable entry heating at 2900+ °F and with single use potential up to at least 3600 °F. TUFROC was initially developed for NASA's X-37 project and ultimately resulted in use on the Air Force X-37B as the wing leading edge (WLE) of the vehicle. TUFROC has similar high temperature capability compared with carbon/carbon, but is manufactured at an order of magnitude lower cost & faster schedule.

thermal protection system↗

NASA NTRSDate not supplied

Overview of Ablation Modeling at NASA

The ambitious scientific payload and crew delivery goals of imminent and future NASA missions are associated with challenging and complex vehicle entries. Advanced ablative Thermal Protection System (TPS) materials will be required for such missions, and, as such, a robust ablation modeling capability is critical to assessing performance by bridging the wide gap between ground testing and entry conditions. The traditional ablation modeling toolset - continuum thermal/materials response analysis - has grown recently to include high-fidelity, multi-scale, and multi-physics techniques that provide a more complete description of the rich physics and chemistry of ablation to better drive down risks related to extreme entries. The present talk provides a snapshot of ongoing NASA activities in ablation modeling, including a review of the current technical capabilities and tools at play within the Agency, the important role academia plays in supporting technical area advancements, and how such internal and external investments intersect with upcoming missions to drive down risks.

TPS↗

NASA NTRSDate not supplied

Comparison Between DSMC and CFD for Hypersonic Planetary Entry Simulations

Hypersonic planetary entry flows span a wide range of Knudsen numbers between rarefied and continuum flows. While computational fluid dynamics (CFD) techniques cannot provide an accurate solution for flows in the rarefied regime, the direct simulation Monte Carlo (DSMC) method is capable of providing accurate solutions for flows in both in the rarefied and continuum regimes but becomes prohibitively expensive as the Knudsen number decreases. For the purpose of thermal protection systems (TPS) design and post-flight reconstruction, various selected points along an entry trajectory are often solved using hypersonic solvers. The quantities of interest that are obtained from that exercise are generally surface quantities, such as pressure, heat flux and enthalpy. Then, material response solvers are used to either design the heat shield to an optimal thickness based on a choice of material, or to provide in-depth heating profiles through the material at various select locations, and compare with flight instrumentation such as the ones that flew on NASA’s two most recent Mars missions, MSL and Mars2020. While most of the heating is generally experienced during the continuum part of the entry, the heating within the rarefied regime is significant for some atmospheres, and hence the flow solutions need to be computed using the DSMC method. Ensuring consistency between hypersonic CFD and the DSMC is crucial so that reliable surface quantities can be passed to material response solvers. Studies were performed to compare the two methods at various select locations, for both non-reacting argon flows as well reacting CO2/N2 flows. Preliminary conclusions show that, for non-reacting flows, the agreement between the two methods for surface heating is excellent (within expected uncertainties) for a freestream Knudsen number of 0.0006, and gets progressively worse as the Knudsen number increases to 0.06. Continuum breakdown analyses were performed and showed that, in general, the Gradient Length Local Knudsen number (KnGLL) from Boyd [1] and associated criteria (KnGLL > 0.05) can be accurately used to predict zones of breakdown in the flow, and associated errors for surface quantities. Furthermore, when studying reacting flow, our studies showed that while flow quantities are highly dependent on relaxation parameters and chemistry rates, it is possible to obtain a good agreement for surface heating, as long as the continuum breakdown in minimal (see Fig. 1). Finally, we studied the impact of the grid refinement on the surface quantities in DSMC.

DSMC↗

NASA NTRSDate not supplied

Thermal Protection Design for Uranus Orbiter Flagship Mission: Probes and Aerocapture

Introduction: The Decadal Strategy for Plane-tary Science and Astrobiology prioritized a Uranus Orbiter and Probe (UOP) as the highest-priority new Flagship mission for the 2023 – 2032 decade. Missions to the outer solar system require significant interplanetary cruise durations on the order of 12 – 15 years while carrying propellant to reduce velocity by several km/s to achieve a desired orbit. Parasitic mass and travel time can be traded where faster arrival velocities require more mass expensive capture burns. Recent advancements in Thermal Protection Systems (TPS) for planetary entry enable mass-efficient designs for atmospheric probes and missions that apply aerocapture for orbit insertion.

Uranus TPS Aerocapture Probe↗

NASA NTRSDate not supplied

Arc Jet Testing of 3D Mid-Density Carbon Phenolic (3MDCP) for Mars Sample Return

When accounting for the highest-heating trajectory with margin, dispersion, and greatest system mass, MSR-EES is predicted to experience the highest heat flux and pressure (approximately 3300 W/cm2 hot-wall, 200 kPa) of any earth entry vehicle todate. To verify performance requirements are met by the forebody TPS, the material must be tested to validate model predictions and give confidence to stakeholder’s expectation of performance. 3D Mid-Density Carbon Phenolic (3MDCP) is NASA’s baseline material for the forebody heatshield of the MSR-EES. As part of the arc jet campaign to evaluate material performance, recent testing has completed in the Interaction Heating Facility (IHF) using a new facility setting to achieve the necessary environments.

arcjet↗

NASA NTRSDate not supplied

Improved Aerothermal Reliability Analysis Enabled by the Mars Sample Return Earth Entry System Aerothermal Database

The Mars Sample Return (MSR) campaign is a series of missions designed to retrieve Martian rock and soil samples for detailed study on Earth. The campaign is split into three primary phases: sample collection with the Mars2020 rover, retrieval with the Sample Return Lander (SRL) and Mars Ascent Vehicle (MAV), and then return to Earth with the Earth Return Orbiter (ERO) and Capture, Containment, and Return System (CCRS) [1]. The final sequence in the Earth return phase is the delivery and entry of the Earth Entry System (EES) sample return capsule. Due to unprecedented planetary protection concerns, the sample return capsule is subject to strict reliability requirements. To this end, the MSR-EES aerothermal team has implemented a flexible aerothermal database architecture capable of integration with state-of-the-art trajectory codes to provide a more rigorous aerothermal reliability analysis. The EES database enables the generation of environments at any location on the heatshield and can incorporate trajectory uncertainties to both statistically quantify aerothermal environments for arcjet testing and produce material response boundary conditions to rigorously select thermal protection system (TPS) sizing environments. This poster will not discuss the fundamental modeling assumptions included in the database, and will instead focus on the downstream reliability analyses that can be performed with a database of this architecture.

MSR-EES↗

NASA NTRSDate not supplied

Polymeric Materials Application and Testing for Space Missions

The EM41 Non-metallics polymers team performs polymeric materials screening, characterization, development, qualification, verification and validation testing for flight hardware and in-space applications. They perform insight, testing and application of polymeric materials for launch vehicles and space environments. Personnel have a deep understanding of the technologies associated with cryogenic insulation, high temperature thermal protection materials, paints, primers, coatings, and adhesive materials. Numerous TPS material characterization tests can also be performed by the Lab. The team has both internal and partner test capabilities to subject polymeric materials to Aerospace and Space environments Material properties, response and performance are tested within environments from the launch pad, through the launch environment and into space. A primary role of the team is to select materials and organize test regimen then analyze test results. Test regiment are established to confirm the materials meet the performance requirements levied by customers within the environments that are expected for the mission.

Thermal↗

NASA NTRSDate not supplied

Material Characterization and Modeling of Room Temperature Vulcanizing Silicone

Room Temperature Vulcanizing silicone (RTV) is a high-temperature adhesive that has successfully been used as a gap-filler between Thermal Protection System (TPS) tiles for heatshields on numerous missions. It is also used to bond instrumentation plugs such as temperature and pressure sensors into the heatshields. While RTV has been traditionally assumed to be a non-porous and non-ablating material, numerous experiments have shown that RTV pyrolyzes and becomes highly porous as it is heated. Heating RTV has also shown swelling, or intumescence, which can pose unique problems that lead to roughness induced boundary-layer transition, surface oxide formation and contamination of heat shield sensors. Therefore, it is crucial to understand and model the intumescence phenomenon of RTV. As data for RTV material properties is limited, the first step in modeling RTV is to collect material properties such as pyrolysis mass-loss, microstructure change, virgin and char porosity, etc. which was performed in our initial study. Additionally, thermomechanical properties such as Young’s modulus and Poisson ratio are required for modeling the intumescence of RTV, which were taken from literature and the coefficient of thermal expansion was collected using in-situ heating and Micro Computed Tomography (µ-CT) in previous studies. Finally, numerous other properties such as pyrolysis gas properties, virgin and char thermal conductivity and specific heat were compiled from previous experiments and literature into a material database that can be used for simulations. In Porous Material Analysis Toolbox based on OpenFOAM (PATO) [4], structural mechanics coupled with material response was used for simulating the intumescence of RTV as it is heated. However, since the permeability of the material is very low, the pyrolysis gas creates an internal pressure build-up as the material is being heated, significantly contributing to the deformation of the material. To correctly characterize this phenomenon, additional physics models were implemented into PATO's stress analysis solver, and results were compared with RTV dilatometry test data as a preliminary verification case. Future work will include experiments of RTV at the Plasmatron X facility and the in-situ heating cell with µ-CT, and improvement of simulation tools to more accurately model RTV intumescence.

TPS↗

NASA NTRSDate not supplied

Material Properties and Modeling of Room Temperature Vulcanizing Silicone

PATO↗

NASA NTRSDate not supplied

Developing and Sustaining Thermal Protection Systems for Uranus Flagship Mission

NASA-developed TPS (3D Woven and C-PICA) needed for Uranus Orbiter and Probe missions are mature and ready today. They also enable other outer planet and sample return missions. We want to raise the awareness of the Science community that given the potential for a ~ 5 years of gap between now and NF-5/NF-6 or UOP mission, the potential for capability atrophy is a high risk that needs to be mitigated.

Uranus Orbiter and Probe↗