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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2012

Performance of Conformable Phenolic Impregnated Carbon Ablator in Aerothermal Environments

Conformable Phenolic Impregnated Carbon Ablator, a cousin of Phenolic Impregnated Carbon Ablator (PICA), was developed at NASA Ames Research Center as a lightweight thermal protection system under the Fundamental Aeronautics Program. PICA is made using a brittle carbon substrate, which has a very low strain to failure. Conformable PICA is made using a flexible carbon substrate, a felt in this case. The flexible felt significantly increases the strain to failure of the ablator. PICA is limited by its thermal mechanical properties. Future NASA missions will require heatshields that are more fracture resistant than PICA and, as a result, NASA Ames is working to improve PICA's performance by developing conformable PICA to meet these needs. Research efforts include tailoring the chemistry of conformable PICA with varying amounts of additives to enhance mechanical properties and testing them in aerothermal environments. This poster shows the performance of conformable PICA variants in arc jets tests. Some mechanical and thermal properties will also be presented.

ablative materials↗
NASA NTRS2012

Aerothermal Testing of Woven TPS Ablative Materials

Woven Thermal Protection Systems (WTPS) is a new TPS concept that is funded by NASAs Office of the Chief Technologist (OCT) Game Changing Division. The WTPS project demonstrates the potential for manufacturing a variety of TPS materials capable of wide ranging performances demanded by a spectrum of solar system exploration missions. Currently, missions anticipated to encounter heat fluxes in the range of 1500 4000 Watts per square centimeter are limited to using one proven material fully dense Carbon Phenolic. However, fully dense carbon phenolic is only mass efficient at heat fluxes greater than 4000 Watts per square centimeter, and current mission designs suffer this mass inefficiency for lack of an alternative mid-density TPS. WTPS not only bridges this gap but also offers a replacement for carbon phenolic, which itself requires a significant and costly redevelopment effort to re-establish its capability for use in the high heat flux missions recently prioritized in the NRC Decadal survey, including probe missions to Venus, Saturn and Neptune. This poster will summarize some recent arc jet testing to evaluate the performance of WTPS. Both mid density and fully dense WTPS test results will be presented and results compared to heritage carbon phenolic where applicable.

Stackpoole, Mairead↗
NASA NTRS1993

Properties and processing characteristics of low density carbon cloth phenolic composites

Ply-lift and pocketing are two critical anomalies of carbon cloth phenolic composites (CCPC) in rocket nozzle applications. Ply lift occurs at low temperatures when the A/P and in-plane permeabilities of the composite materials are still very low and in-plane porous paths are blocked. Pocketing occurs at elevated temperatures when in-plane permeability is reduced by the A/P compressive stress. The thermostructural response of CCPC in a rapid heating environment involves simultaneous heat, mass, and momentum transfer along with the degradation of phenolic resin in a multiphase system with temperature- and time-dependent material properties as well as dynamic processing conditions. Three temperature regions represent the consequent chemical reactions, material transformations, and property transitions, and provide a quick qualitative method for characterizing the thermostructural behavior of a CCPC. In order to optimize the FM5939 LDCCP (low density carbon cloth phenolic) for the nozzle performance required in the Advanced Solid Rocket Motor (ASRM) program, a fundamental study on LDCCP materials was conducted. The cured composite has a density of 1.0 +/- 0.5 gm/cc which includes 10 to 25 percent void volume. The weight percent of carbon microballoon is low (7-15 percent). However, they account for approximately one third of the volume and historically their percentages have not been controlled very tightly. In addition, the composite properties show no correlation with microballoon weight percent or fiber properties (e.g. fiber density or fiber moisture adsorption capacity). Test results concerning the ply-lift anomaly in the MNASA motor firings were: (1) Steeper ply angle (shorter path lenght) designs minimized/eliminated by lifting, (2) material with higher void volume ply lifted less frequently, (3) materials with high (greater than 9 percent) microballoon content had a higher rate of ply lifting, and (4) LDCCP materials failed at microballoon-resin interfaces. The objectives of this project are: (1) to investigate the effects of carbon microballoon and cabosil fillers as well as fiber heat treatment on plylift-related mechanical properties, (2) to develop a science-based thermostructural process model for the carbon phenolics. The model can be used in the future for the selection of the improved ASRM materials, (3) to develop the micro-failure mechanisms for the ply-lift initiation and propagation processes during the thermoelastic region of phenolic degradation, i.e. postcuring and devolatilization.

Wang, C. Jeff↗
NASA NTRSDate not supplied

Qualification of Domestic Lyocell Based Phenolic Impregnated Carbon Ablator (PICA-D) for NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90’s. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions, from Sample Return missions such as Stardust, OSIRIS-REx to large Mars Lander missions such as Mars Science Lab (MSL) and Mars 2020. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues. The challenge involved in finding a replacement fiber source is in processing as well as in the final performance of the ablator. Each replacement necessitates the requalification of the PICA. This has happened at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Due to its flight heritage and proven capability, PICA is baselined as the Thermal Protection System (TPS) for Dragonfly and Mars Sample Return Sample Retrieval Lander mission and is being considered for the backshell of the Mars Sample Return Earth Entry System. All three missions are due to be launched between 2026 and 2028. The challenge this time is to ensure the PICA made with domestic material is a suitable replacement to the heritage PICA used in MSL and Mars 2020 so that the design of the heatshield can be matured without much risk. Results are presented from the recent efforts of 22 PICA-D billets that were Lot Acceptance Tested. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Testing of PICA-Domestic (PICA-D) indicates very comparable performance with respect to “heritage” PICA materials and thus PICA-D is expected to be a sustainable and nearly a “drop-in” replacement solution for future NASA missions.

PICA↗
NASA NTRS2015

Thermal Protection System Mass Estimating Relationships for Blunt-Body, Earth Entry Spacecraft

System analysis and design of any entry system must balance the level fidelity for each discipline against the project timeline. One way to inject high fidelity analysis earlier in the design effort is to develop surrogate models for the high-fidelity disciplines. Surrogate models for the Thermal Protection System (TPS) are formulated as Mass Estimating Relationships (MERs). The TPS MERs are presented that predict the amount of TPS necessary for safe Earth entry for blunt-body spacecraft using simple correlations that closely match estimates from NASA's high-fidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT). These MERs provide a first order estimate for rapid feasibility studies. There are 840 different trajectories considered in this study, and each TPS MER has a peak heating limit. MERs for the vehicle forebody include the ablators Phenolic Impregnated Carbon Ablator (PICA) and Carbon Phenolic atop Advanced Carbon-Carbon. For the aftbody, the materials are Silicone Impregnated Reusable Ceramic Ablator (SIRCA), Acusil II, SLA-561V, and LI-900. The MERs are accurate to within 14% (at one standard deviation) of FIAT prediction, and the most any MER under predicts FIAT TPS thickness is 18.7%. This work focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

mass estimating relationship (MER)↗
NASA NTRS2015

Thermal Protection System Mass Estimating Relationships For Blunt-Body, Earth Entry Spacecraft

Mass estimating relationships (MERs) are developed to predict the amount of thermal protection system (TPS) necessary for safe Earth entry for blunt-body spacecraft using simple correlations that are non-ITAR and closely match estimates from NASA's highfidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT). These MERs provide a first order estimate for rapid feasibility studies. There are 840 different trajectories considered in this study, and each TPS MER has a peak heating limit. MERs for the vehicle forebody include the ablators Phenolic Impregnated Carbon Ablator (PICA) and Carbon Phenolic atop Advanced Carbon-Carbon. For the aftbody, the materials are Silicone Impregnated Reusable Ceramic Ablator (SIRCA), Acusil II, SLA- 561V, and LI-900. The MERs are accurate to within 14% (at one standard deviation) of FIAT prediction, and the most any MER can under predict FIAT TPS thickness is 18.7%. This work focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

Spacecraft↗
NASA NTRS2014

Development Of FIAT-Based Parametric Thermal Protection System Mass Estimating Relationships For NASA's Multi-Mission Earth Entry Concept

An integrated tool called the Multi Mission System Analysis for Planetary Entry Descent and Landing (M-SAPE) is being developed as part of NASAs In-Space Propulsion Technology (ISPT) program. Part of M-SAPEs development requires the formulation of mass estimating relationships (MERs) to determine the vehicle's Thermal Protection System (TPS) material and required thickness for safe Earth entry. The objective of this study was to develop MERs using simple correlations that were non-ITAR and matched as accurately as possible NASAs high-fidelity ablation modeling tool, the Fully Implicit Ablation and Thermal Analysis Program (FIAT ). These MERs would be a first-estimate for feasibility studies; it is understood that higher-fidelity modeling like FIAT would be necessary once a proposed trajectory was down-selected. The trajectory space for these MERS consisted of 840 different trajectories, and a materials heating limit was the main constraint for an allowable trajectory. MERs for the vehicle fore body included the ablating materials Phenolic Impregnated Carbon Ablator (PICA ) and Carbon Phenolic atop Advanced Carbon-Carbon. For the backshell the materials were Silicone Impregnated Reusable Ceramic Ablator (SIRCA ), Acusil II, SLA-561V, and LI-900. The MERFIAT ratio indicates MERs are accurate to within 14 percent (at one standard deviation) of FIAT prediction, and the most any MER can under-predict TPS thickness is 18.7 percent of FIAT prediction. This poster focuses on the development of these MERs, the resulting equations, model limitations, and model accuracy.

thermal protection system↗
NASA NTRSDate not supplied

Robust and Mass Efficient Thermal Protection Systems for Future Venus Missions

Several international Venus missions, both orbiter, and in-situ probe are in the works after a prolonged absence. Future missions are expected to be long-duration lander missions and balloon missions to investigate the cloud layer. The entry mass for these in-situ missions will be significantly higher, and the aeroshell will be larger, than previous missions. Two cardinal requirements govern the selection and use of thermal protection systems (TPS), namely, robustness to assure mission safety during entry and mass efficiency so that the useful mass for science is maximized. One cannot trade mission safety for mass when it comes to TPS. The robustness of the TPS is a paramount requirement as it is a single point of system failure. At the same time, TPS mass is carried for the entire mission duration prior to entry, and any excess TPS mass is at the cost of science payload. Future missions will benefit enormously from TPS options with mass and performance benefits far beyond advanced carbon-carbon (ACC) that is currently proposed for the DAVINCI mission. NASA STMD, and SMD jointly invested in the development of 3-D woven thermal protection systems in the last decade, and as a result, the Heat-shield for Extreme Entry Environment Technology (HEEET) TPS has been matured to TRL 6. It is ready for future mission use including Venus. The HEEET project focused on developing a broad technology base, applicable for missions not only to Venus, but also to Saturn, the Ice Giants, and higher speed sample return missions such as Mars Sample Return requiring extreme robustness due to bio-hazards. The dual-layer HEEET (DL-HEEET) is proven to be extremely robust. DL-HEEET performed well in arc jet and laser testing conditions where heritage carbon-phenolic failed. HEEET also proved to be more mass efficient compared to Carbon-Phenolic. Recently, the principal author was invited to present his perspectives to the 2023-2032 Planetary Sciences Decadal Venus subcommittee on the current state of TPS for future Venus missions [1]. In addition to pointing out successful TPS and other developments in the last decade in support of entry systems, the principal author made two findings: 1) importance of and need to sustain TPS capabilities that took nearly a decade to develop and 2) the opportunity to further optimize TPS mass without sacrificing robustness to further enable future in-situ missions. This talk will focus on the above two recommendations and provide the rationale for them. The intent of the talk is to seek advocacy from the VEXAG community for the criticality of TPS sustainment and the benefits of 3D woven TPS optimization. TPS Capability Sustainment: HEEET development was necessitated by the atrophy of heritage carbon-phenolic. Atrophy has impacted other TPS materials as well. Avcoat, the TPS that successfully allowed astronauts to explore the Moon in the 1960s and 70s, took nearly a decade and $10’sM to recover in support of Orion/Artemis. One of the most used TPS materials, SLA 561V, had to be recovered prior to its use on Mars Pathfinder. Phenolic Impregnated Carbon Ablator (PICA), the TPS that replaced SLA 561V on MSL, has undergone multiple replacements and recovery due to constituent rayon changes. Recently, NASA invested in a domestic rayon replacement program and also invested in FMI to consolidate PICA capability for NASA missions, as a result of FMI’s decision to discontinue commercial FiberForm, which is needed for PICA. TPS used for planetary missions are unique and have no other commercial or defense use. In addition, low mission cadence is also a driver for TPS atrophy. Hence, NASA, as the steward, must take steps to ensure and sustain TPS capabilities. In this talk, we outline steps NASA can take to keep abreast of emerging risks and target risk mitigation steps to ensure TPS capability sustainment for Venus and other extreme environment missions. Next Generation of Mass efficient and Robust TPS: NASA invested in an alternate TPS to PICA based on felt-technology called Conformal-PICA which has the potential to save 30% - 50% mass over PICA. The development was discontinued at a Technology Readiness Level (TRL) of ~ 5. 3D Mid-Density Carbon Phenolic (3MDCP) is a single layer variant of HEEET, based on the insulating layer only (SL-HEEET) and currently baselined for the Mars Sample Return Mission Earth Entry System due to its mass efficiency (30% more mass efficient than DL- HEEET). Currently, SL-HEEET is limited to aeroshell diameters of < 1.3m. The SL-HEEET was compared to DL-HEEET in the recent ADVENT flag-ship class mission study in support of the Planetary Science Decadal. SL HEEET was the recommended TPS based on 30% mass savings for both the balloon and lander missions. Given C-PICA and SL-HEEET have superior reliability and mass efficiency, advocacy from VEXAG is sought for completing their further development to TRL 5/6 in this decade so as to enable TPS readiness for future missions.

Thermal Protection Systems↗
NASA NTRSDate not supplied

Qualification of Domestic Lyocell Based Phenolic Impregnated Carbon Ablator (PICA-D) for NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90’s. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions, from Sample Return missions such as Stardust, OSIRIS-REx to large Mars Lander missions such as Mars Science Lab (MSL) and Mars 2020. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues. The challenge involved in finding a replacement fiber source is in processing as well as in the final performance of the ablator. Each replacement necessitates the requalification of the PICA. This has happened at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Due to its flight heritage and proven capability, PICA is baselined as the Thermal Protection System (TPS) for Dragonfly and Mars Sample Return Sample Retrieval Lander mission and is being considered for the backshell of the Mars Sample Return Earth Entry System. All three missions are due to be launched between 2026 and 2028. The challenge this time is to ensure the PICA made with domestic material is a suitable replacement to the heritage PICA used in MSL and Mars 2020 so that the design of the heatshield can be matured without much risk. Results are presented from the recent efforts of 22 PICA-D billets that were Lot Acceptance Tested. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Testing of PICA-Domestic (PICA-D) indicates very comparable performance with respect to “heritage” PICA materials and thus PICA-D is expected to be a sustainable and nearly a “drop-in” replacement solution for future NASA missions.

Matt Gasch↗
NASA NTRS2019

Development of Lyocell Based Phenolic Impregnated Carbon Ablators (PICA-D) for Future NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since 1999. Due to the obsolescence of the input fiber source, new PICA materials were developed using Lyocell, a domestic rayon fiber source. Results are presented from this effort. Manufacturing included fiber conversion, fabrication of tile component and near net shaped heatshield preforms, and conversion to PICA materials. Thermal, mechanical, and representative environment arc-jet testing have been conducted. Initial testing indicates comparable performance with respect to heritage PICA material, and likely "drop-in" replacement for future NASA mission needs.

Thermal Protection↗
NASA NTRS2019

Development of Lyocell Based Phenolic Impregnated Carbon Ablator (PICA-D) for Future NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since 1999. Due to the obsolescence of the input fiber source, new PICA materials were developed using Lyocell, a domestic rayon fiber source. Results are presented from this effort. Manufacturing included fiber conversion, fabrication of tile component and near net shaped heatshield preforms, and conversion to PICA materials. Thermal, mechanical, and representative environment arc-jet testing have been conducted. Initial testing indicates comparable performance with respect to heritage PICA material, and likely "drop-in" replacement for future NASA mission needs.

thermal protection↗
NASA NTRS2019

Development of Domestic Lyocell Based Phenolic Impregnated Carbon Ablator (PICA-D) for Future NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90's. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues and required replacement and requalification at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Results are presented from this effort including fiber conversion, fabrication of tile component and near net shaped heatshield preforms, and conversion to PICA materials. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Initial testing of PICA-Domestic (PICA-D) indicates comparable performance with respect to "heritage" PICA materials and thus PICA-D is expected to be a sustainable solution for future NASA missions.

Drop-In Replacement↗
NASA NTRS2019

Development of Domestic Lyocell Based Phenolic Impregnated Carbon Ablator (PICA-D) for Future NASA Missions

Phenolic Impregnated Carbon Ablator (PICA) is a low-density ablator that has been used as the planetary entry heatshield for several NASA missions since the late 90's. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flag-ship class missions. The rayon-based carbon precursor raw material used in PICA preform manufacturing has experienced multiple supply chain issues and required replacement and requalification at least twice in the past 25 years, and a third substitution is now needed. Due to the obsolescence of the input foreign rayon fiber source, a new variant of PICA has been developed using a domestic rayon-like fiber source, Lyocell. Results are presented from this effort including fiber conversion, fabrication of tile component and near net shaped heatshield preforms, and conversion to PICA materials. Thermal, mechanical, and representative environment arc-jet tests have been conducted. Initial testing of PICA-Domestic (PICA-D) indicates comparable performance with respect to "heritage" PICA materials and thus PICA-D is expected to be a sustainable solution for future NASA missions.

Drop-In Replacement↗
NASA NTRS2008

Evaluation of the Effect of Silicone Contamination on Various Bond Systems and the Feasibility of Removing the Contamination

Silicone is a contaminant that can cause catastrophic failure of a bond system depending on the materials and processes used to fabricate the bond system, Unfortunately, more and more materials are fabricated using silicone. The purpose of this testing was to evaluate which bond systems are sensitive to silicone contamination and whether or not a cleaning process could be utilized to remove the silicone to bring the bond system performance back to baseline. Due to the extensive nature of the testing attempts will be made to generalize the understanding within classes of substrates, bond systems, and surface preparation and cleaning methods. This study was done by contaminating various meta! (steel, inconel, and aluminum), phenolic (carbon cloth phenolic and glass cloth phenolic), and rubber (natural rubber, asbestos-silicone dioxide filled natural butyldiene rubber, silica-filled ethylene propylenediene monomer, and carbon-filled ethylene propylenediene monomer) substrates which were then bonded using various adhesives and coatings (epoxy-based adhesives, paints, ablative compounds, and Chemlok adhesives) to determine the effect silicone contamination has on a given bond system's performance. The test configurations depended on the bond system being evaluated. The study also evaluated the feasibility of removing the silicone contamination by cleaning the contaminated substrate prior to bonding. The cleaning processes also varied depending on bond system.

Stanley, Stephanie D.↗
NASA NTRS2013

Parametric Thermal Soak Model for Earth Entry Vehicles

The analysis and design of an Earth Entry Vehicle (EEV) is multidisciplinary in nature, requiring the application many disciplines. An integrated tool called Multi Mission System Analysis for Planetary Entry Descent and Landing or M-SAPE is being developed as part of Entry Vehicle Technology project under In-Space Technology program. Integration of a multidisciplinary problem is a challenging task. Automation of the execution process and data transfer among disciplines can be accomplished to provide significant benefits. Thermal soak analysis and temperature predictions of various interior components of entry vehicle, including the impact foam and payload container are part of the solution that M-SAPE will offer to spacecraft designers. The present paper focuses on the thermal soak analysis of an entry vehicle design based on the Mars Sample Return entry vehicle geometry and discusses a technical approach to develop parametric models for thermal soak analysis that will be integrated into M-SAPE. One of the main objectives is to be able to identify the important parameters and to develop correlation coefficients so that, for a given trajectory, can estimate the peak payload temperature based on relevant trajectory parameters and vehicle geometry. The models are being developed for two primary thermal protection (TPS) materials: 1) carbon phenolic that was used for Galileo and Pioneer Venus probes and, 2) Phenolic Impregnated Carbon Ablator (PICA), TPS material for Mars Science Lab mission. Several representative trajectories were selected from a very large trade space to include in the thermal analysis in order to develop an effective parametric thermal soak model. The selected trajectories covered a wide range of heatload and heatflux combinations. Non-linear, fully transient, thermal finite element simulations were performed for the selected trajectories to generate the temperature histories at the interior of the vehicle. Figure 1 shows the finite element model that was used for the simulations. The results indicate that it takes several hours for the thermal energy to soak into the interior of the vehicle and achieve maximum payload temperatures. In addition, a strong correlation between the heatload and peak payload container temperature is observed that will help establishing the parametric thermal soak model.

Agrawal, Parul↗
NASA NTRS2018

Sustaining PICA TPS for Future NASA Robotic Science Missions

Phenolic Impregnated Carbon Ablator (PICA), invented in the mid 1990's, is a low-density ablative thermal protection material proven capable of meeting sample return mission needs from the moon, asteroids, comets and other "unrestricted class V destinations" as well as for Mars. Its low density and efficient performance characteristics have proven effective for use from Discovery to Flagship class missions. It is important that NASA maintain this TPS material capability and ensure its availability for future NASA use. The rayon based carbon precursor raw material used in PICA preform manufacturing required replacement and requalification at least twice in the past 25 years and a third substitution is now needed. The carbon precursor replacement challenge is twofold – the first involves finding a long-term replacement for the current rayon and the second is to assess its future availability periodically to ensure it is sustainable and be alerted if additional replacement efforts need to be initiated. Rayon is no longer a viable process in the US and Europe due to environmental concerns. In the early 80's rayon producers began investigating a new method of producing a cellulosic fiber through a more environmentally responsible process. This cellulosic fiber, lyocell, is a viable replacement precursor for PICA fiberform. This presentation reviews current SMD-PSD funded PICA sustainability activities in ensuring a rayon replacement for the long term is identified and in establishing that the capability of the new PICA derived from an alternative precursor is in family with previous versions of the so called "heritage" PICA.State of the Art Low Density Carbon Phenolic AblatorsStardust SRC post flight withPICA forebody heat shield(0.8m max. diameter)PICA Processing StepsRole of Rayon/Lyocellin PICA.

Stackpoole, Mairead↗
NASA NTRSDate not supplied

Thermal Protection System to Enable Ice Giant Aerocapture Mission for Delivering Both an Orbiter and an In Situ Probe

The Ice Giants have been identified as high priority science destinations in the last Decadal Survey [1] and could benefit from aerocapture as the primary method for orbit insertion [2]. A mass-efficient aerocapture system will enable the delivery of an orbiter along with an atmospheric probe (for in situ measurements to anchor global data collected by the orbiter) and possibly a lander at Triton [3]. Aerocapture could be executed either using low L/D rigid aeroshell with lift modulation (LMA) [4] or using deployable aeroshell using drag modulation (DMA) [3]. Nearly two decades ago, a NASA-funded team performed Neptune-Triton aerocapture studies with a mid-L/D lifting configuration [5] for achieving orbit using LMA. This study showed aerocapture challenges. Due to very high peak entry conditions combined with very high heat-load, a suite of TPS materials was required and this suite was deemed problematic from a qualification perspective, due test facility limitations. In the past 20 years, progress made in GN&C for lift-guided entry missions such as MSL, Orion EFT1, Mars 2020 and the upcoming Artemis missions, and the expertise in blunt body aerodynamics at large scale (~ 5m) has led the EDL community to conclude that aerocapture is a “go do” engineering activity and most technologies are in hand to propose missions with aerocapture [6] [7]. Aerocapture using DMA, currently in development, is an option for Ice Giant Missions. While DMA is simpler in some sense, due to ballistic entry and no need for lift-guided maneuvering, it has challenges and it’s maturity is lower. LMA and DMA both require one or more ablative Thermal Protection System (TPS) materials for the rigid aeroshell element. The ablative TPS needs to be robust and mass efficient due to the high heat loads and size of the rigid aeroshell. Currently, there are capable ablative thermal protection materials, e.g., Heatshield for Extreme Entry Environments Technology (HEEET), 3-D woven Mid-Density Carbon- Phenolic (3MDCP), and PICA (Phenolic-Impregnated Carbon Ablator) that are mature, i.e., at TRL 6 or higher. NASA also invested in Conformal PICA that was matured to TRL 5. Our goal is to evaluate the applicability of high TRL TPS and consider other design options. We first establish bounding aerocapture trajectories for a wide range of arrival conditions and the associated aerothermal environment. Based on the environments, we then determine the predicted TPS mass for the aeroshell [4]. In this presentation, we will outline the process by which we establish bounding aerocapture trajectories for hyperbolic excess velocities ranging from 27 km/s to 35 km/s, which are shown to be a range of velocities that can reduce the trip time from ~14 years to 8 years. The above velocity range translates to ~12 km/s to ~24 km/s arrival velocities at the planetary entry interface [2]. The velocity reduction required to achieve orbit ranges between ~2.5 km/s to 9.5 km/s for both Neptune and Triton. Propulsive insertion alone, due to the amount of fuel required to achieve the required velocity reduction, limits the science returned [2]. We establish the bounding aerocapture trajectories for a low L/D (~ 0.4) configuration for three different ballistic coefficients. The ballistic coefficient range is determined from three different aeroshell diameters of 3m, 4m and 5m and with an entry system mass of 2200 kg. With the above range of design parameters, we then determine conservative/bounding estimates of aerothermal environments by using a combination of CFD simulations and stagnation point heating estimates [7]. This engineering approach allows us to first assess the TPS need vs. TPS capability and determine the applicability of existing TPS. Once an applicable suite of TPS is determined, the TPS thickness and mass are computed. We show that the TPS mass fraction can be as low as 5% to as high as 20%, depending on total trip time reduction and other design parameters for a range of TPS. This is a large range for TPS mass fraction. We show PICA and HEEET can indeed enable aerocapture missions, but the missions incur a mass penalty. TPS mass savings, can be further reduced with the use of conformal PICA. Advancing the development of Conformal PICA to make it robust across the entire aerothermal environment (peak heat-flux, pressure and shear) range will result in TPS mass fractions of < 10% for Ice Giant aerocapture missions such as the Neptune-Triton mission. Aerocapture allows for not only shortening the trip time but enables larger mass to be placed in orbit. Furthermore, probes deployed from orbit will benefit in reduced entry environments allowing for a lower risk TPS implementation as compared to mission designs where the probe is released prior to orbit insertion. One of the challenges for the Ice Giant community is to ensure mission designs that maximize science and allow flexibility in the placement of the entry probe. The traditional approach to release the probe ahead of the orbiter may not optimize returned science. In this presentation, we will make the case for mature TPS such as HEEET and PICA. While these materials can enable aerocapture missions, completing the development of conformal PICA and extending Conformal PICA to be more robust, will have significant impact to TPS mass efficiency and significantly enhance science return for future Gas- and Ice-Giant missions.

E Venkatapathy↗
NASA NTRSDate not supplied

Advances in Design Capabilities for Planetary Missions from the NASA Entry Systems Modeling and Instrumentation Portfolio

The Entry Systems Modeling project (ESM) is supported by both the NASA Space Technology and the Science Mission Directorates and focuses on developing simulation tools and validated models for characterizing the performance of entry systems tailored to planetary destinations across the Solar System. ESM is organized into six technical capability areas that together address all relevant factors related to spacecraft entry, as well as some aspects of descent: Thermal Protection System (TPS) Materials; Aerothermodynamics; Entry & Descent Vehicle Dynamics; Guidance, Navigation, and Control; Vehicle Systems Analysis; and Advanced Tools and Numerical Methods. Development within the capability areas is undertaken explicitly with a focus on transition and infusion to science missions, human exploration missions, and commercial space activities. The present talk details developments that specifically impact science missions, including simulation tool capabilities that aid in mission design and model development to understand entry system performance at a given destination. Examples of the successful infusion and transition of such project outcomes to science missions also are provided. Several simulation tool development efforts within ESM have resulted in new design capabilities for missions. One such outcome is improved toolsets for mission trajectory and concept of operations design. Specifically, an initiative to couple a leading tool for entry, ascent/descent, and orbital trajectory optimization (Program to Optimize Simulated Trajectories II or POST2) to those used within the Agency for interplanetary trajectory optimization (Copernicus and Monte) has made substantial progress, with the outcomes to date promising to allow efficient trajectory optimization across mission phases. Additionally, toolchains for the evaluation of vehicle performance during entry and descent have been developed that allow assessment of multi-dimensional aeroheating on detailed vehicle geometries, characterization of deployment and inflation of parachutes, and assessment of vehicle dynamic stability during descent. These capabilities are achieved by coupling diverse sets of physics together – material response, computational fluid dynamics, radiation, and vehicle dynamics – to suitably describe complex entry and descent phenomena. Several model development and validation efforts for specific destinations and entry regimes also are underway within the ESM project. For instance, new experimental capabilities to validate radiation models at low densities/high altitudes recently have been established with project support, specifically the Low-Density Shock Tube (LDST) at the NASA Ames Research Center Electric Arc Shock Tube (EAST) facility. The LDST is being leveraged to develop improved models of shock layer kinetics and radiation in Titan atmospheres, while future studies will be conducted in the LDST and the existing high velocity shock tube to provide validation data for radiation models of Venus, Ice Giants, and Mars atmospheres. Models describing the aerothermal and thermo-structural performance of Thermal Protection System (TPS) materials has been another focus, with multiscale modeling activities on-going for the two leading TPS materials applicable to a range of entry conditions and science missions: the Phenolic-Impregnated Carbon Ablator (PICA) and woven materials like 3D Mid-Density Carbon Phenolic (3MDCP). A continual effort is made to infuse and transition outcomes from ESM simulation tool and model development activities into relevant science missions. Significant progress has been made on this front, with missions such as Dragonfly, DAVINCI, and Mars Missions benefitting from project outcomes. The groundwork also is being laid to provide insights into forward looking missions to Gas/Ice Giants as well as for potential sample returns.

Justin Haskins↗