Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “STARDUST”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

516 records · Page 29
All sourcesNASA NTRSDOE OSTI
NASA NTRSDate not supplied

Development of Advanced Ablative Carbon Phenolic TPS for Future NASA Missions and Commercial Space

Phenolic Impregnated Carbon Ablator (PICA) is a low-density carbon phenolic 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 Flagship 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 mass efficient, compliant version 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) is comparable in terms of peak heat-flux capabilities compared to standard PICA. Additionally, a novel all felt PICA-Flex offers further reductions in density, with higher thermal protection efficiency resulting in additional mass savings beyond C-PICA. Furthermore, the compliant nature of these materials allows for application without Strain Isolation Pad to a wide variety of substrates, making 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. Preliminary thermal and mechanical tests have been conducted comparing the performance of Conformal and Soft-PICA and PICA-Flex materials.

Flexible Ablative TPS↗
NASA NTRS2024

Flight Mechanics Modeling and Simulation of the Earth Entry System

Introduction: The Mars Sample Return (MSR) Campaign being planned by NASA and ESA has the ambitious goal to return Mars samples back to Earth. This international collaboration had developed a concept of operations that included a ESA-designed Earth Return Orbiter (ERO) and NASA-designed Capture, Containment, and Return System (CCRS). The Earth Entry System (EES), consisting of a protective aeroshell that houses the samples as well as sample containment vessels, would conduct entry, descent, and landing (EDL) on a direct Earth trajectory. The EES would enter on a spin-stabilized ballistic trajectory with the goal to passively achieve aerodynamic stability throughout all regions of flight. The EDL sequence would end with the EES impacting the soft playa soil of the Utah Test and Training Range (UTTR). As of the submission of this abstract, the MSR campaign is undergoing a re-architecture leading to a pause in EES development. However, the novel approaches developed in flight mechanics modeling and simulation can significantly benefit the greater IPPW community in the development of Earth return vehicles. This paper will present the latest state of EES flight mechanics modeling and simulation. The paper will highlight the simulation architecture developed and key lessons learned from understanding of EDL trajectory sensitivities. Modeling and Simulation: Figure 1 provides a high-level concept of operations for the approach, entry, descent, and landing (AEDL) phase of the CCRS-portion of MSR. The objective of EES flight mechanics is to model and simulate the EES trajectory from ERO separation to ground impact at UTTR. A variety of flight mechanics simulation models were utilized to model both exo-atmopsheric and atmospheric portions of flight. 42, a 6-DOF simulation developed at Goddard Space Flight Center, is utilized for propagating the attitude of EES during exo-atmospheric flight. 42 allows for a variety of spin eject mechanism scenarios to be simulated for analysis. 10 minutes prior to entry, the 42 states are handed off to the EDL sims. The prime EDL sim utilized by EES is the Program to Optimize Simulated Trajectories II (POST2), a 6-DOF sim developed at Langley Research Center, and the independent verification and validation EDL sim utilized is DSENDS, a 6-DOF sim developed at Jet Propulsion Laboratory. Figure 2 provides a visualization of the flight mechanics simulation model flow through various points in the AEDL phase. Due to the existence of a variety of sim models, the EES flight mechanics team developed processes for data hand-off. These processes included the development of a centralized coordinate frame document, utilization of a single, centralized simulation input document for all sims to reference, and hand-off files containing both the technical data to be ingested by other flight mechanics sims as well as annotations of modeling assumptions utilized to generate the data. Figure~\ref{fig:post2simarchitecture} provides an overview of the POST2 sim architecture wherein POST2 ingests numerous subsystem models and input files. The dispersed state file generated by MONTE provides the position/velocity state of the trajectory while the 42 Handoff file provides the attitude. The aerodynamics database, delivered by the EES aeroscience team, is utilized to simulate the aerodynamic forces and moments experienced during EDL. A custom atmosphere model, developed by EES atmosphere team, is utilized to simulate the anticipated atmosphere environment around the region of Earth through which the EES trajectory flys. These inputs and subsystem models can be varied depending on the AEDL flight mechanics scenario being simulated. Monte Carlo simulations are utilized to generate statistical AEDL performance metrics in the form of scorecards and violin plots. Furthermore, outputs from the POST2 simulation are utilized for follow-on analyses including aerothermal and landing performance. \section{Flight Mechanics Lessons Learned} Though the EES flight mechanics team uncovered a variety of lessons learned through the analysis conducted to support CCRS through preliminary design review, this paper will highlight the most important lessons. A key AEDL performance goal is to ensure the landing footprint of EES remains on the UTTR south range. A common modeling strategy used in EDL analysis is One-Variable-At-a-Time (OVAT). OVAT analysis provides insight into the key drivers that affect AEDL performance metrics. Figure 3 shows the landing ellipses for single dispersion sources as compared to the baseline aggregate of all dispersions. The figure shows that atmosphere winds alone dominate the size of the footprint ellipse (note: EES does not use a parachute unlike previous Earth-return missions and is in wind-driven free fall for ~5min). The significance of the wind led the EES flight mechanics team to pursue the development of a Custom Atmosphere Model [4], in lieu of EarthGRAM [1], built on actual radiosonde wind measurements around the UTTR-region. This decision was driven by the realism in the generated footprint ellipses and lessons-learned from Stardust [5]. These findings will be invaluable for future Earth-return missions in providing an early understanding of the key drivers affecting footprint size and modeling considerations for which to account. Another lesson learned is tied to the AEDL performance goal of achieving passive stability throughout all regions of flight. It is well understood that blunt-body aeroshells are less stable as they transition from supersonic to subsonic. Eliminating a backshell does help improvestability; however, other phenomena such as roll-induced instability during terminal descent can still arise. The EES flight mechanics team developed stability metrics as tools to better understand the causes of and better predict the onset of dynamic instability. These tools were built upon analytical models developed by Jaffe [3] and Murphy [2]. The tools were shown to both be very accurate in correlation with actual unstable cases and useful in developing stability margin policies based on the vehicle design and simulation considerations (e.g. sphere-cone angle change, mass change, wind turbulence). These tools allowed for the current EES design to demonstrate the ability to achieve passive stability and can be an invaluable tool for consideration in the design of parachute-less Earth-return vehicles.

Rohan Deshmukh↗
NASA NTRSDate not supplied

Thermal Protection Systems (TPS) for High Velocity Earth Entry Missions

The fastest man-made object to re-enter into Earth’s atmosphere, with an entry velocity of 12.8 km/s, was the Stardust capsule, which returned samples from the comet Wild 2. Sample return missions offer high and long-lasting science yield; as such, they are discussed repeatedly in the latest planetary science decadal survey, titled Origins, Worlds, and Life. An advantage of sample return missions over in situ measurements is the ability to use state-of-the-art instruments that are not limited in power, size, or complexity; conversely, in situ measurements rely on instruments optimized to fit within the limitations of the spacecraft. Additionally, sample return provides the opportunity to revisit samples as analysis techniques improve as well as to further investigate unanticipated or ambiguous results. As launch vehicle and propulsion technologies continue to advance, it is becoming feasible to consider sample return missions from further out in the Solar System beyond Mars orbit. However, returning samples from increasingly distant destinations comes with high entry velocities, and thus requires ever more capable thermal protection systems (TPS). Evaluating the capabilities of existing mature TPS materials for a range of entry conditions and sample return aeroshell configurations is key to establishing the feasibility of future mission proposals. As such, this presentation describes the comprehensive trade study done to evaluate the feasibility of high velocity Earth entry missions with the currently available TPS materials.

Hannah S. Alpert↗
NASA NTRSDate not supplied

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

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

Chemistry↗
NASA NTRSDate not supplied

TPS Technology Status on Enabling New Frontiers Sample Return Missions from Moon, Asteroids and Comets and Probe Missions to Saturn and Uranus

STMD and SMD investment in TPS technology development such as Heatshield for Extreme Entry Environment (HEEET), 3-D woven Mid-Density Carbon Phenolic (3MDCP) and Conformal PICA over the past decade in response the the Decadal Committee Recommendation (2012). 3MDCP, a variant of HEEET, is baselined as the TPS for MSR, HEEET is enabling the Rocket Lab mission to Venus and Conformal PICA enabled VARDA as the heat shield TPS for their sample return capsule Winnebago-1. In the meantime, PICA, the TPS that enabled Stardust and OSIRIS-REx has atrophied. While these are success stories, upcoming NF-5/6 AO is anticipated to have multiple proposals that may be more demanding for sample return missions from Moon, Asteroids and Comets, and in addition probe missions to Saturn and UOP . This poster will address the state-of-the-art TPS for NF-5/6 in-situ and sample return missions, and for the UOP flagship mission. The poster will focus on steps that need to be taken to both advance and sustain NASA developed TPS and be ready to enable NF-5/6 and Flagship missions.

Todd White↗
NASA NTRS2024

Capsule_Grid User Guide

Software utility CAPSULE_GRID and several related procedures that automate grid generation for computational fluid dynamics calculations about axisymmetric atmospheric entry vehicles are described as a supplement to the documentation within the source codes. This application program and its ancillary software utilities are all in the public domain at the web site https://software.nasa.gov/software/category/all/arc/1/cfdtools with the exception of the 2D hyperbolic volume gridding step for which scripts driving either GRIDGEN1 or HYPGEN2 are available. CAPSULE_GRID was developed at NASA Ames Research Center in support of the DPLR3 real gas flow solver. Both forebodies (only) and full bodies are treated, for axisymmetric 2D flow calculations or for 3D calculations at angle of attack. The expected application is to the smooth outer mold line of a hypersonic atmospheric entry vehicle, for which a single layer of point-matched structured grid blocks normally suffices. Control files for a variety of representative cases are included with illustrations of results, and some insights into the subtleties of curvature-based grid point distribution are provided. Gridding of a capsule on a sting is supported. Additional procedures for treating arbitrarily long 2D wakes, prompted by asteroid studies and relevant to spacecraft such as Stardust, are also documented. Most recently, treating an asymmetric forebody (including an off-center nose, with or without an aft body) has been incorporated as an extra step.

spacecraft↗
NASA NTRSDate not supplied

An Overview of Experiments and Modeling of Polysiloxane-Coated Thermal Protection Systems for Missions to Mars, Titan, and Beyond.

Phenolic Impregnated Carbon Ablator (PICA) gained heritage during the historic Stardust mission, where it successfully returned samples from a comet’s tail and has since been instrumental in delivering payloads to the surface of Mars [1-3]. Most recently, PICA enabled the safe return of samples collected from asteroid Bennu as part of the OSIRIS-REx mission. This rich legacy underscores PICA’s critical role in allowing NASA’s most ambitious exploration missions. However, the friable nature of its phenolic phase presents challenges during handling and pre-launch activities. To mitigate this issue, PICA is coated with a polysiloxane resin system, which serves to suppress particulate dispersion and thereby safeguard spacecraft components. A comprehensive understanding of the polysiloxane resin’s behavior is imperative, as it profoundly shapes the material response of PICA during atmospheric entry by influencing its thermal and oxidative stability. This influence extends to thermocouple plugs embedded within thermal protection systems. These plugs have demonstrated their significance in missions such as Mars Science Laboratory (MSL) and Mars 2020, where the MEDLI and MEDLI2 instrumentation suites delivered in-valuable insights into the performance of thermal protection systems during entry into the Martian atmosphere [4]. Looking ahead, missions such as Dragonfly, set to descend into Titan’s dense atmosphere, aim to leverage advanced sensor technologies to further refine our understanding of thermal protection response [5]. Moreover, thermocouple plugs play an essential role in validating cutting-edge material response models, such as those pioneered under NASA’s Entry Systems Modeling Project (ESM), designed, in-part, to predict the operational integrity of thermal protection systems under the extreme stresses of atmospheric entry. To achieve these modeling goals, ground-based experiments are crucial to provide the foundational data necessary for developing and refining these predictive tools. To this end, an extensive test campaign was conducted at the Hypersonic Materials Environmental Test System (HyMETS) to investigate the high-temperature behavior of the polysiloxane resin in an air environment [6]. These experiments revealed critical phenomena, including the formation of a silicon oxycarbide layer that enhances oxidation resistance, moderates surface temperatures, and alters in-depth thermal response. Building on these findings, subsequent tests were designed to simulate atmospheric entry conditions in reactive gases, such as CO2 and N2, to mimic the environments of Mars and Titan, respectively, as well as non-reactive gases representing the atmospheres of the Ice Giants (Neptune and Uranus). A heating rate dependent decomposition mechanism has been identified for the polysiloxane resin under oxidizing conditions (Fig. 1). In the initial stage, the resin and the underlying thermal protection system undergo pyrolysis, rapidly generating a thin amorphous silicon oxycarbide interwoven with carbonaceous char and residual fibers from PICA. During the second stage, the nascent oxide layer establishes a robust, oxidation-resistant thermal barrier coating, which significantly impedes heat transfer to the underlying carbonaceous char, resulting in a stagnation of the surface temperature. A key factor contributing to this thermal resistance is the low recombination efficiency of atomic oxygen (γ), which further diminishes the heat load on the material’s interior layers [7]. Moreover, as the surface temperature stagnates, the silicon oxycarbide phase separates into distinct regions of silica and free graphite. Ultimately, when the heat flux reaches a critical threshold, a third stage is triggered, leading to the breakdown of the coating through carbothermal reduction, exposing the underlying char layer. This exposure leads to a dramatic surface temperature spike, driven by highly exothermic reactions between atomic oxygen and the char layer, further accelerating material degradation. A detailed mass and heat transfer model of PICA coated with polysiloxane resin was implemented in the Porous material Analysis Toolbox based on OpenFOAM, PATO [8]. The initial stage was considered negligible in this model because the resin decomposition occurs rapidly within a thin surface layer. Instead, the coating was directly considered as an oxygen-resistant thermal barrier coating. For the second stage, the thin amorphous silicon oxycarbide was treated as a pure silica surface to simplify the thermochemical behavior. The model ac-counts for surface equilibrium processes using representative elements of the coating-environment system. For the third stage, specific boundary conditions were developed to estimate the onset and progression of the coating removal. Two-dimensional material response simulations were conducted to compare uncoated and coated PICA using boundary conditions calibrated with HyMETS data. Fig. 2 illustrates that the simulations closely align with experimental data, successfully reproducing measured temperature profiles. This work will include the latest advancements in the coating model, including the calibration of recombination of atomic oxygen at the surface during the second phase. These simulated results will be further validated against additional CO2 data points from HyMETS, reinforcing the models’ predictive capabilities. These mechanisms and their effects on thermal protection systems, including thermochemical behavior and thermocouple probe performance in extreme environments, provide crucial insights for optimizing spacecraft designs that safeguard scientific payload and ensure mission success in future planetary exploration endeavors.

Active Oxidation↗
NASA NTRSDate not supplied

A Machine Learning Framework for Error Compensation in Radiative Transfer Calculations

Radiative heat transfer influences the amount of heat flux transferred to the surface of the hypersonic vehicle, which is essential to evaluate the performance of thermal protection systems. The radiative heat flux is found to be computationally prohibitive while accounting for the variation in spatial, angular, and spectral domains. A new methodology has been recently developed to alleviate the cost of computation in the spectral domain by constructing flow-agnostic reduced-order models (ROMs). The developed spectral ROM databases provide grouping strategies that account for non-equilibrium absorption and emission as well as interaction between disparate species due to spectral overlap in associated radiative processes. However, the developed ROMs need to be optimized for a specific combination of interacting gas species and would need to re-calibrated in case individual species are added/omitted. In this work, we use various machine learning (ML) techniques to approximate the radiative intensities determined by a ROM optimized for a specific gas mixture. The ML model relies on the ROM databases developed for a single species which ignores any spectral overlap. Thus, radiation evaluation starts with a simple summation of radiative intensities predicted using these non-calibrated ROMs for the contributing species. The ML framework then provides a correction to account for the interplay in the frequency, i.e., emission of photons by one species and absorption by another, and yields mixture-specific radiation fields. Once trained on the individual ROM databases, the ML framework offers instantaneous corrections that serves as a time/cost effective alternative to the optimization of ROMs for a specific gas mixture. The ML framework is trained on both the high fidelity and ROM evaluated line of sight (LOS) data from Orion, Stardust, and FIRE II cases to obtain a general purpose correction model for earth re-entry scenarios when radiation contributions from both atomic nitrogen and atomic oxygen are considered. A geometric length scale parameter is used in the training process to account for errors introduced in the ROM databases as a consequence of high optical thickness. The efficacy of the ML framework is underscored through extensive analysis of train and test errors with respect to all the re-entry scenarios. The applicability of such an ML framework was further corroborated by embedding it in a state-of-the-art US3D - NERO system for determining the radiative heat flux transferred to the hypersonic vehicle surface.

Radiation↗
NASA NTRS2026

Adaptive Curation at NASA Johnson Space Center: Preparing for Artemis Samples by Leveraging Proven Practices and Innovative Solution

The proper curation of returned astromaterial samples is essential to ensure that high-quality scientific investigations can be conducted for decades, enabling future generations to address evolving research questions. The Astromaterials Acquisition and Curation Office at NASA Johnson Space Center (hereafter JSC Curation) is responsible for curating all of NASA’s extraterrestrial samples. Under the governing document, NASA Procedural Requirement (NPR) 7100.5 “Curation of Extraterrestrial Materials”, JSC Curation is charged with “The curation of all extraterrestrial material under NASA control, including future NASA missions.” The Directive goes on to define Curation as including “...documentation, preservation, preparation, and distribution of samples for research, education, and public outreach.” JSC Curation has a long-standing legacy of curating extraterrestrial materials, including but not limited to Apollo and Luna lunar samples, Genesis solar wind samples, Stardust comet samples, asteroid samples (Hayabusa1, Hayabusa2, and OSIRIS-REx), and Antarctic Meteorite samples from a variety of parent bodies. Building on this foundation, the Artemis program introduces new challenges and opportunities for sample curation, requiring both the application of proven practices and the development of innovative solutions. The Artemis Collection will be curated using established protocols refined through decades of experience with Apollo and subsequent sample collections, including but not limited to the utilization of cleanrooms, custom nitrogen gloveboxes, specialized storage containers and tools; all of which have strict material utilization and prohibition requirements. These practices provide a robust framework for contamination control, documentation, and long-term preservation. However, Artemis samples may present unique scientific and operational requirements, including enhanced contamination control measures (relative to Apollo) driven by evolving science objectives. To meet these needs, JSC Curation is actively developing new technologies and protocols that extend beyond traditional approaches, ensuring that the integrity of samples is maintained under increasingly stringent standards. One critical area of innovation is the development of cold sample curation capabilities. Certain Artemis samples, particularly those from Permanently Shadowed Regions (PSRs) and cold environments, are expected to contain ices and volatile components that require preservation and handling at sub-zero or even cryogenic temperatures. JSC Curation is leveraging best practices from the cold and cryogenic sample industries, as well as other government agencies and academic experts, to design facilities and handling procedures that maintain sample integrity while enabling scientific access. The goal is to develop capabilities to allow researchers to investigate volatiles and other temperature-sensitive materials while minimizing chemical and physical alterations from their returned state. In summary, adaptive curation at JSC combines the reliability of proven methodologies with forward-looking innovations to meet the scientific and operational demands of Artemis. Through enhanced contamination control, advanced cold curation capabilities, and an understanding of known and evolving future science priorities, NASA is preparing to maximize the scientific return from Artemis samples and preserve their value for generations of researchers.

Andrea D Harrington↗