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At least 1,171 records · Page 65
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

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

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

Recession↗
NASA NTRSDate not supplied

Status of the 3-MDCP, the 3-D Woven TPS enabling MSR Earth Entry System Development

While the HEEET project established the dual-layer, 3-D Woven TPS at TRL 6 for extreme entry environment missions such as Saturn and Ice-Giant Probe missions, subsequent to the completion of HEEET project in 2018, further development was undertaken to assess the viability of a variant of the 3-D Woven, a single, insulating layer alone as a viable TPS for Mars Sample Return mission. MSR Earth Entry System (EES) baselined it and named the TPS as 3D, mid-density, carbon-phenolic (3MDCP). MSR program invested in building a new loom and the weaver T.E.A.M. Inc is currently weaving 80” wide 1.2” thick insulating layer. Unlike HEEET, where a set of tiles and seams are integrated to form the heat-shield, 3MDCP is a seamless heatshield by wrapping and shaping a flat panel into the heatshield shape and then rigidizing the preform through resin infusion. MSR EES is investing in technology maturation of 3MDCP for the current mission launch in 2027. This lighting talk will provide key developments and time line for 3MDCP technology maturation to show that 3MDCP will be ready to support future UOP or Saturn Probe mission. In addition, the limitations of the capability will be presented so that Saturn mission proposers as well as UOP flagship mission study can be fully informed as to the 3MDCP capability.

HEEET↗
NASA NTRSDate not supplied

A Combined CFD/Material Response Analysis of 3MDCP Arcjet Experiments

The Mars Sample Return Earth Entry System (MSR-EES) project has selected 3-D Woven Mid-Density Carbon Phenolic (3MDCP) as the baseline thermal protection system (TPS) material for the capsule that will return Martian soil samples to Earth sometime in the 2030’s. A series of experiments in the NASA Ames AHF and IHF arcjets will be used to characterize the performance of 3MDCP and to develop and refine material response models for it. The test objectives for these experiments include obtaining in-depth and surface temperature data, mass loss and surface recession measurements, and char depth measurements. To provided pre- and post-test support for the arcjet experiments, a combined CFD/material response was performed using the DPLR and Icarus code. Pre-and post-test analysis on the 3MDCP arcjet experiments was performed using the DPLR CFD and Icarus material response codes. The codes were not tightly coupled, but boundary condition data required by Icarus (e.g., heat transfer coefficients and surface pressures), were extracted from the CFD solutions. This poster shows comparisons between DPLR/Icarus and the experimental data taken during arcjet experiment AHF-348 where 4-inch diameter iso-q models made from 3MDCP were tested in the 12-inch nozzle of the AHF arcjet. This poster also includes a top-level overview of the CFD/material response arcjet simulation process. Best-practices for running CFD solutions of arcjet experiments including how to determine the inflow conditions for the CFD solution from the arc heater settings are discussed. The required inputs to the material response solver are presented, as well as a discussion of how they can be extracted from the CFD solution.

thermal protection systems↗
NASA NTRSDate not supplied

Investigation of the Effect of Etch Pits on the Material Properties of Carbon Fiber Structures

It is well known that etch pits are formed on carbon surfaces during their consumption in reactive gas environments such as during ablation of carbon-based thermal protection system (TPS) materials. These pits start out from atomic defects on the carbon surface, grow in size and then ultimately lead to the complete removal of the solid carbon material. However, the effect of such etch pits on the material properties of carbon ablators such as FiberForm are poorly understood, as the maximum size of pits are on the order of microns in size, which makes them hard to observe experimentally. Hence, we have developed a module within the Porous Microstructure Analysis code PuMA that can generate etch pits on an arbitrary material surface, with a prescribed distribution of pit size and density. We use this newly developed module to calculate material properties of FiberForm, which is the base material for one of the most commonly used TPS materials - Phenolic Impregnated Carbon Ablator (PICA). The detailed microstructure of FiberForm in PuMA will either be generated synthetically or obtained from X-ray microtomography. Our simulations show that material properties such as the thermal conductivity and tortuosity decrease with increasing degree of pitting. This will be helpful to more accurately predict the degradation of carbon-based TPS during ablation, and furthermore may allow us to identify previously unknown TPS failure modes due to pitting.

PuMA↗
NASA NTRSDate not supplied

Material Response Analysis of PICA-NuSil at the Hypersonic Materials Environmental Test System

Phenolic impregnated carbon ablator (PICA) is a thermal protection system (TPS) that gained heritage during the Stardust mission and is the baseline TPS material for expeditions to Mars (e.g., Mars Science Laboratory and Mars 2020).1 PICA is friable; therefore, a polysiloxane resin (NuSil CV-1144-0) is deposited onto the surface of flight hardware before assembly, test, and launch operations (ATLO) to mitigate particle shedding. Furthermore, the Mars Science Laboratory (MSL) and Mars 2020 heatshields were instrumented with an array of thermocouple plugs (MEDLI Integrated Sensor Plugs or MISP) designed to measure the in-depth thermal response of PICA during flight. Post-flight analysis of MISP data suggests that recession was lower than predicted, and NuSil is suspected as a reasonable cause. NuSil primarily consists of a copolymer blend of dimethyl and diphenyl-substituted polysiloxane resin. Four decades of research in the field of polymer-derived ceramics demonstrate that polysiloxane resins form oxidation-resistant silicon oxycarbide materials at pyrolysis temperatures as low as 800 °C. -2 Furthermore, silicon oxycarbide materials exhibit excellent thermal stability at relatively high-temperatures (< 1200 °C) but eventually decompose via carbothermal reduction. -3 Therefore, it is reasonable to suggest that NuSil forms an oxidation-resistant thermal barrier coating on the TPS surface and influences the material response during atmospheric entry. Evidence for this notion was observed during prior arc-jet testing at both the Aerodynamic Heating Facility (AHF) and the Panel Test Facility (PTF) at the NASA Ames Research Center. Articles of PICA were subjected to testing in airflow, and a temperature discontinuity was recorded at the surface of coupons coated with NuSil. A glassy coating was observed to form at the surface in shear flow configurations (wedge and panel) under moderate heat fluxes. Furthermore, state-of-the-art material response models have yet to account for the underlying physics and chemistry of the temperature discontinuity. To this end, a pathfinder test campaign was executed at the Hypersonic Materials Environmental Test System (HyMETS) to enhance the fundamental understanding of PICA-NuSil ablation phenomena and provide a preliminary data set to aid in the development of material response models.

PICA↗
NASA NTRSDate not supplied

Status of the Development of 3MDCP, the 3-D Woven TPS enabling the Mars Sample Return Earth Entry System

- The Mars Sample Return Earth Entry System is being designed to withstand entry environments of: - Peak Heat Flux of > 3000 W/cm2 - Peak Dynamic Pressure of > 2 atmospheres - Peak Shear of > 3000 Pa. - The TPS Material System Selected to support MSR EES is “3MDCP”, 3D Mid Density Carbon Phenolic. 3MDCP is a derivative of the HEEET material system, using the “Insulation Layer” only. - Implementation of 3MDCP on EES is monolithic instead of a tile architecture in order to meet MSR reliability demands. - The EES TPS Team has recently completed a number of major engineering development milestones in support of delivering the flight 3MDCP TPS hardware: - Weaving Scale-Up completed, > 80” width woven preform production underway - Full Scale Forming & Manufacturing Attrition completed to verify delivered thickness capabilities. - Recent Design Change: EES design has been changed from a 45 degree sphere cone to a 52.5 degree sphere cone to allow for additional system mass.

Z. Young↗
NASA NTRSDate not supplied

Subscale Motor Test Bed Development to Simulate Solid Rocket Motor Internal Environments for Material Characterization

NASA Marshall Space Flight Center’s Materials and Processes Department, with support from the Propulsion Systems Department has developed a test bed for exposing ablative thermal protection materials to an environment similar to that of solid rocket motors (SRMs) by using a hybrid-propellant combustor. The Solid Fuel Torch (SFT) is a pressurized. small-scale hybrid rocket motor which utilizes an aluminized solid fuel and a gaseous oxygen (GOx) oxidizer to vary internal environments which can be tailored to specific program needs. The baseline design for the SFT was developed to support the Space Shuttle Program and was redesigned to support Space Launch System (SLS) Booster program requirements. There are four configurations of the SFT testbed: 1.) Insulation configuration with the option of a “dome” or “converging cone” insulation test section, 2.) Super Sonic Blast Tube configuration to simulate nozzle exit cone environments, 3.) Flow Straightened configuration which includes an additional instrumented length between the fuel grain and the insulation, and 4.) Constant Diameter Throat configuration which includes a cylindrical phenolic test section to test samples with different ply angles under the same environment simultaneously. Motor pressures for each configuration can range from 150 to 1500 psi with heat fluxes up to 1200 BTU/sq.ft/s. This paper discusses the development of each configuration and the results obtained during the initial test series.

C.L. Bower↗
NASA NTRSDate not supplied

Enabling Planetary In-Situ and Sample Return Missions: NASA Developed 3-D Woven Thermal Protection / Heatshield Technology for Extreme Entry Environment

Heatshield for Extreme Entry Environment Technology (HEEET) and 3-D Mid-Density Carbon-Phenolic (3MDCP) are two variants of 3-D Woven. TPS funded by STMD and SMD. HEEET was matured to TRL 6 in 2019, and SMD offered it as incentivized technology. Mars Sample Return Earth Entry System has baselined a seamless 3MDCP heatshield, and it will reach TRL 6 by 2025. HEEET and 3MDCP have been tested to withstand entry peak heat-flux ~ 4500 W/sq.cm, pressure ~ 6.5 atm, and shear ~ 4000 Pa. Bally Ribbon Mills and T.E.A.M. Inc developed looms and demonstrated their capability to weave HEEET and 3MDCP respectively. F.M.I. has demonstrated capability to form and resin infuse the preforms, and machine the TPS for final assembly. Expertise to integrate the parts reside at NASA, and NASA can transfer the knowhow to interested parties via technology transfer.

Ethiraj Venkatapathy↗
NASA NTRSDate not supplied

Development of Advanced Conformal and Soft Pica for Future Nasa Missions and Commercial Space

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 MSL and Mars 2020. Continued development of a more mass efficient and compliant versions of PICA that use a carbon felt as the substrate will be presented. The felt based versions of PICA known as Conformal PICA (C-PICA) and Soft PICA (S-PICA), are very comparable in terms of peak heat-flux capabilities compared to standard PICA. The lower density and of C-PICA and S-PICA further show a much higher thermal protection efficiency and hence more mass savings. The compliant nature of these materials allows for use on a wide variety of substrates and therefore makes it very attractive for commercial low earth orbit missions as well as NASA’s aerocapture missions to Ice Giants. NASA is currently working with several commercial partners (Varda and Inversion Space) to Technology Transfer C/S-PICA variants for commercial LEO missions. Thermal, mechanical, and representative environment arc-jet tests have been conducted comparing the performance of Conformal and Soft-PICA materials.

Matthew J Gasch↗
NASA NTRS2019

Experimental Measurements of the High-Temperature Oxidation of Carbon Fibers

We carried out laboratory experiments to study the decomposition of carbon fibers at high temperature and studied the material used in charring carbon phenolic ablators for planetary probe heatshields. Porous plug samples were exposed to controlled rates of molecular oxygen, carbon dioxide, and carbon monoxide at furnace-heated flow-tube reactor temperatures between 700 and 1500 K and pressures between 2000 and 6200 Pa. We used calibrated mass spectroscopy to make time-resolved quantitative measurements of decomposition products of the gases downstream of the test sample. Absolute and differential pressure measurements were used to determine the high-temperature permeability of the material and to monitor changes during material decomposition, and we characterized the microstructure of the porous sample using scanning electron microscopy. We found substantial carbon fiber oxidation for O2 flows, resulting in increasing CO/CO2 production ratios with increasing temperature. The CO2 was shown to react with carbon fibers following a Boudouard reaction at temperatures above 1200 K.

Francesco Panerai↗
NASA NTRS2023

STXM-C-XANES, IR Microspectroscopy, and Raman Microspectroscopy Analysis of Organic Matter in the Chwichiya 002 C3.00-Ungrouped Chondrite

Chwichiya 002 is a C3.00-ungrouped carbonaceous chondrite discovered in Western Sahara in 2018. The presence of hydrated phases in Chwichiya 002 was not confirmed by XRD analysis, and only a faint water and Si-O band attributed to phyllosilicate were observed by infrared transmission spectroscopy, indicating a very low degree of aqueous alteration (Meteoritical Bulletin, 2020). The structural order of polycyclic aromatic hydrocarbons evaluated by Raman spectroscopy indicates that the effect of thermal metamorphism is less severe than in the Semarkona meteorite, an LL3.00 ordinary chondrite (Meteoritical Bulletin, 2020). Chiwichiya 002 is a very primitive carbonaceous chondrite and an important information source for the early stage of the evolution of organic matter in the solar system. In this study, we investigated the molecular structure and distribution of organic matter in Chiwichiya 002 by Carbon X-ray Absorption Near Edge Structure (C-XANES) analysis using a Scanning Transmission X-ray Microscope (STXM). The STXM-C-XANES analysis has a spatial resolution of 20~100 nm and can observe the spatial distribution of organic materials in microstructures. In addition, to investigate the degree of thermal metamorphism of the organic matter, micro-IR, and micro-Raman spectroscopic analyses were performed. In addition, the Murchison CM2 chondrite was also analyzed for comparison. To evaluate contamination during analytical procedures, antigorite, which was heated in air at 500°C for 5 hours to eliminate organic matter was also analyzed in the same manner. The samples were milled into 100 nm-thick, thin sections using a Focused Ion Beam (FIB), and C-XANES spectra were obtained using the STXM at the High Energy Accelerator Research Organization, Photon Factory (PF), BL-19B. The obtained C-XANES spectra were smoothed and then fitted with Gaussian functions using the method of Le Guillou et al. (2018). The C-XANES showed that the ratio of oxygenated functional group (ketone, phenolic) carbons in Chwichiya 002 was lower than in Murchison, whereas the ratio of aliphatic carbons was almost the same. This may reflect the low degree of aqueous alteration of Chwichiya 002. In addition, contamination peaks were detected at 287.3-287.7 eV and 288.3-288.7 eV in the C-XANES spectra of some of the antigorite samples. The source of these peaks is under investigation. Raman spectroscopic analysis showed that the intensity ratio of the D band to the G band of Chwichiya 002 was slightly higher than that of Murchison. IR absorption spectra showed aromatic absorption at 1630 cm -1 , but there was little absorption due to aliphatic C-H around 2900 cm -1 . These results suggest that the organic matter in Chwichiya 002 may be more heated than in Murchison.

Chwichiya 002↗
NASA NTRSDate not supplied

The Material Response of PICA-NuSil to a Hypersonic Flow Environment

Phenolic Impregnated Carbon Ablator (PICA) is NASA’s baseline thermal protection system for missions to Mars. PICA is friable; therefore, NuSil ® CV-1144-0 is applied to the surface of flight hardware to mitigate the contamination of spacecraft sensors. NuSil is composed of a polysiloxane resin system that can transform into an oxidation-resistant coating upon heating. Therefore, a pathfinder campaign was conducted at the Hypersonic Materials Environmental Test System (HyMETS) at the NASA Langley Research Center to investigate the material response of instrumented sphere-cone models made from PICA coated with NuSil ® . Analysis of the post-test data is presented and suggests that NuSil ® impacts the material response of test articles subjected to oxidizing atmospheres, as evidenced by a reduction of the measured surface temperature and in-depth temperature response. Finally, a four-step ablation mechanism is proposed based on observations from high-speed video, optical emission spectroscopy, thermal analysis, and post-test surface characterization.

PICA PICA-NuSil Siloxane Silicone↗
NASA NTRSDate not supplied

Parametric Analysis of Entry Vehicles for Giant Planet Missions

The Planetary Science Decadal Survey has identified Uranus and Saturn as high priori-ty destinations for a flagship and New Frontier missions respectively in the decade 2023-2032. The pro-posed presentation will focus on the entry and descent aspects of the entry vehicles design, considered as part of Giant Planet probe mission concepts, and associated trades for viable trajectory options. Giant Planet Entry Vehicle Parametric Study: Launch vehicle capabilities are evolving and provide an opportunity to increase instrumented probe dimensions. To assess the impact of larger aeroshell designs, a parametric study was conducted to understand the impact on aerothermal environments, TPS options, and TPS mass over a range of 1.0m to 2.0m aeroshell diameters for Uranus and Saturn probe concept missions. The 45° sphere-cone geometry is a legacy configuration that has demonstrated static stability and been used successfully in missions to Venus (Pioneer-Venus) and Jupiter (Galileo). A nose radius of 0.4 m was considered primarily to reduce the heat flux at the stagnation point compared to the smaller radii used in the Venus and Jupiter missions. Representative inertial velocities are chosen from a prior NASA Ames study. Viable entry trajectories to meet concept mission and science objectives were developed using the tool POST2. The newly developed thermal protection material called HEEET (Heatshield for Extreme Entry Environment Technology) was considered in the study. This material, which is at a technology readiness level (TRL) of 6, is highly customizable and available in two varieties: (i) a dual-layer version consisting of recession layer on top of an insulative layer, and (ii) a single-layer version consisting of the insulative layer alone, termed 3- dimensional Mid-Density Carbon Phenolic (3MDCP). Both options were considered for the forward heatshield (the sphere-cone part) in the pre-sent study.

Thermal Protection System↗
NASA NTRSDate not supplied

Impact Forces From Artificial and Real Birds on a Large Diameter Hopkinson Bar

Forces were measured from normal impacts with two different artificial bird recipes and two species of real birds on a very rigid flat surface. The tests were conducted by launching the soft body projectiles axially into a Hopkinson bar in accordance with the SAE AS6940 test standard at three different nominal impact conditions: (a) a 1 kg projectile at an impact velocity of 49 m/sec; (b) a 1.8 kg projectile at an impact velocity of 110 m/sec; and (c) a 1.8 kg projectile at an impact velocity of 310 m/sec. At each condition two artificial birds and one real bird projectile were tested with at least three test repeats. Simulations were conducted to assess the effects of projectile orientation and velocity on predicted forces. The Hopkinson bar consisted of an Aluminum 6061 solid cylindrical bar with a diameter of 305 mm and a length of 7.31 m, made up of two 3.66 m long sections that were in axial contact with each other. The bar was instrumented for strain measurement at two locations, 457.2 mm and 609.6 mm (1.5 and 2 diameters) from the impacted face. Forces were calculated from the measured strain. In addition, digital image correlation (DIC) was used to measure the velocity of the free end of the bar. The real bird projectiles were 1 kg Mallard ducks and 1.8 kg Golden Comet chickens that were prepared per ASTM F330-21. Prior to testing, the real birds were placed in muslin bags to minimize tumbling during flight. Two artificial bird formulations were tested, one produced by the University of Dayton Research Institute (UDRI) and the other by the German Aerospace Center, DLR. The UDRI projectile was a homogeneous, relatively soft solid, formulated with a mixture of water, gelatin and phenolic micro-balloons to achieve a nominal density of 0.95 g/cc. The DLR projectile consisted of a printed plastic shell with internal plastic ribs and filled with a gel material. The projectiles were designed such that the overall average density was the desired value of 0.95 g/cc. Initial testing showed that contact between the two Hopkinson bar segments was not sufficient to allow full transmission of the waves through the interface. For the 49 m/sec and 110 m/sec tests this resulted in reflections returning to the strain measurement site prior to the end of the impact. For these tests only a portion of the force pulse could be accurately measured. For the highest velocity tests at 310 m/sec the force pulse was complete by the time reflected waves interfered with the response. Despite this limitation, useful force data were obtained for all velocities. Impact forces were very sensitive to the impact orientation and velocity. The force history generally consisted of an initial region with varying amplitude followed by a relatively constant amplitude response. The steady state region was similar among all projectiles and similar to what would be predicted for an inviscid fluid with a density of 0.95 g/cc. The initial part of the force response exhibited varying degrees of test-to-test repeatability associated with each projectile type. Forces calculated from DIC velocity measurements of the free end of the bar compared favorably those from the strain measurements. In general, the artificial projectile impact forces were quite repeatable from test to test, with some exceptions which could be explained by issues associated with either the projectile itself or with how the projectile exited the gun barrel. For all impact velocities the initial portion of the impact force from the bird projectiles exhibited large test-to-test variability. In this paper, impact forces for three types of projectiles at three impact velocities will be presented with an emphasis on the test-to-test repeatability of the results. Based on test and simulation results a proposed method for normalizing the impact force to minimize effects of differences in projectile orientation, impact velocity and density will also be discussed.

Bird strike↗
NASA NTRSDate not supplied

A Multi-Probe Mission at Jupiter is Within our Reach!!

Soon after Galileo Probe mission, capabilities to manufacture and test heritage carbon-phenolic (HCP) TPS atrophied. The impact of capability loss to Outer Planet probe missions, Venus in-situ and Mars Sample Return (MSR) missions was realized after two decades. MSR (1997 – 2006) initiated a search and recover of HCP. Decadal Survey (2003) recommended JPOP mission (Jupiter Polar Orbiter and Probe), but lack of HCP TPS resulted in Juno being proposed without a probe. Decadal Survey (2013) recognized and advised NASA to close the TPS gap. With the community support, NASA initiated a technology development effort to close this gap in 2019 and today, missions to Venus, Saturn, Uranus, and MSR are enabled by new 3-D Woven TPS. Jupiter, the most demanding mission from an entry and TPS perspective ever, still lacks a viable qualified TPS. While the Science community needs to make the case for going back to Jupiter with a Probe, this work is an assessment of emerging capabilities to support such a mission when needed. Recent SOA developments are very promising to close this technology gaps. Encourage the Science Community about the feasibility of future Jupiter Multi-Probes Mission Engage the Entry System/TPS Technology Community as to the emerging capabilities and need to preserve them Educate and excite the early career and next generation students interested in future Science and technology challenge.

Ethiraj Venkatapathy↗
NASA NTRSDate not supplied

Development of Additive Manufacturing Technologies for 3D Printing of Spacecraft Heat Shields

Ablative heat shields are an enabling technology for entry into planetary atmospheres. From the PICA heatshields used for several Mars rovers to the carbon phenolic material used for Galileo’s Jupiter entry probe, the heat shield manages the heat load transferred to the payload, protecting the sensitive scientific instruments carried on entry probes. The Additive Manufacturing of Thermal Protection Systems (AMTPS) project, an Early Career Initiative (ECI) funded by NASA’s Space Technology Mission Directorate and led by NASA Johnson Space Center, seeks to develop materials and processes for 3D printing ablative heat shields for spacecraft. Current methods for producing ablative heat shields are extremely labor intensive and require extensive hands-on processes and quality control characterization. Additive manufacturing (AM) offers the possibility of reduced production times, improved reliability, and enhanced performance via graded compositions. Costs will also be reduced by reducing the time and labor required for heat shield production. Direct integration of the heat shield onto the structure during processing simplifies integration and reduces risk.

Nathaniel Olson↗
NASA NTRSDate not supplied

Thermostructural Qualification of PICA-D for NASA Planetary Science Missions

Phenolic Impregnated Carbon Ablator (PICA) has been used as the heatshield TPS material for many NASA planetary science missions. Due to a supply issue with the heritage Rayon fibers used in PICA fabrication, NASA started a project in 2017 to develop a sustainable replacement for PICA using a domestic fiber source, named PICA-D. This material has been selected as the heatshield TPS material for two upcoming NASA planetary science missions, Dragonfly mission to Titan and Mars Sample Retrieval Lander (SRL). While most PICA-D thermal and mechanical properties are very similar to PICA, early test data suggest that PICA-D is stiffer in tension in the in-plane (IP) direction, with higher strength but lower strain-to-failure. Preliminary thermostructural analysis using Finite Element Methods (FEM) tools predict IP compressive failure in the near-surface layers of PICA-D, for both Dragonfly and SRL flight environments. This talk provides an overview of the thermostructural qualification plans for PICA-D. A two-pronged approach is being pursued to demonstrate that the material and heatshield design has adequate capability for both SRL and Dragonfly design environments. The first part is to improve the accuracy of analytical predictions through a comprehensive characterization of PICA-D mechanical properties. The second, and more critical, part of the qualification approach is a subsystem-level test of PICA-D to demonstrate material capability under flight-relevant combined mechanical and thermal loads. The team is currently preparing for two test series in Sandia’s Solar Tower facility, scheduled in September 2023 and focusing on PICA-D acreage, followed by additional tests in the Solar Tower and LHMEL facilities focusing on design features and repair methods.

TPS↗
NASA NTRSDate not supplied

Simulations of Coupled Shoulder Ablation of A Conceptual Aeroshell

The current study explores coupled numerical simulations of flow dynamics and material response affect ablation of the shoulder region of an entry capsule used in space missions like the Mars Sample Return (MSR) project. A new computational tool named Ares, developed at NASA Ames Research Center, is applied to carry out the coupled simulations by interconnecting well-established US3D (flow solver) and Icarus (material response solver), and the material considered is three-dimensional mid-density carbon phenolic (3MDCP). Particularly, this work extends the uncoupled analysis of Shrestha et al. 2022 by incorporating the coupling mechanism across the ablating surface between fluid and the capsule geometry. For simplicity, an axisymmetric geometry of the entry system is considered. A few mission-relevant parametric studies such as angles of orientation of the material layers with respect to the streamwise direction, backshell boundary conditions, and alternation of convective heat-transfer coefficients are investigated to essentially explore Ares capability with the primary focus on shoulder location.

Coupled Ablation Modeling↗