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NASA NTRSDate not supplied

Atmospheric Entry Heat Flux Sensor Calibration and Flight Data Analysis

- Recent NASA missions have included total heat flux sensors (THFS) embedded in the thermal protection system (TPS) to measure the combined convective and radiative heating during atmospheric entry. These measurements are key to fundamental entry science and mission design. - The THFSs for the Mars Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite on the Mars 2020 entry vehicle and the Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) technology demonstration mission were passively cooled Schmidt–Boelter gauges. - As presented at IPPW 2021, THFS calibration is notoriously difficult and includes several sources of measurement uncertainty [1]. - Recent work has been conducted to understand and quantify the measurement uncertainty in the MEDLI2 and LOFTID Schmidt-Boelter THFSs.

R. A. Miller↗

NASA NTRSDate not supplied

Analysis of Pyrolysis Products from Ablative TPS

NASA’s state-of-the-art ablative materials are composed of a three-dimensional network of carbon fibers impregnated with polymers that dissipate thermal energy through pyrolysis. A fundamental understanding of the decomposition mechanisms and pyrolysis product distributions of various classes of polymers is instrumental in the design of new ablative materials. Furthermore, innovative experiments are essential to the continuous modernization of material response models by providing high-fidelity data. Thus, an apparatus has been designed to measure pyrolysis products from polymers and composite materials by implementing in-situ mass spectrometric techniques. Initial results from experiments performed on siloxane resins and a common phenolic resin will be discussed. Both classes of polymers exhibit heating-rate-dependent decomposition mechanisms. At the onset of heating, phenolic polymers decompose through competitive reactions to form gaseous products and a carbonaceous char. Gas phase products of phenolic resins are typically composed of molecular hydrogen, water, and aromatic hydrocarbons. Pyrolysis products from siloxane polymers include molecular hydrogen, small molecules, and cyclic oligomers from the polymer.

Pyrolysis↗

NASA NTRS1985

Definition of an entry research vehicle

Within the last year, there has been a growing interest in technology which would support the development of vehicles which can maneuver in the atmosphere while returning from orbit and vehicles which can fly in the atmosphere at hypersonic speeds for sustained periods of time. This rebirth of interest in hypersonics is partly a result of a developing awareness of the potential benefits which can be derived from vehicles with capability to operate between the limits of existing aircraft and spacecraft. Examples of the types of maneuvers which are projected for this new class of vehicles are presented. Challenges provided by the considered maneuvers are summarized, taking into account a synergetic or atmospheric plane change. The plane change maneuver will require a Thermal Protection System (TPS). Attention is given to the definition of a Shuttle launched entry research vehicle experiment. Details regarding the synergetic plane change are considered along with the maneuvering entry, vehicle heating, and experiments and instrumentation.

Freeman, D. C.↗

NASA NTRS2014

Overview of the 6 Meter HIAD Inflatable Structure and Flexible TPS Static Load Test Series

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). At the beginning of 2014, a 6m HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify the designs structural performance. The 6m HIAD structure was constructed in a stacked toroid configuration using nine inflatable torus segments composed of fiber reinforced thin films, which were joined together using adhesives and high strength textile woven structural straps to help distribute the loads throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for displacement of the inflatable structure as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The centerbody of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to provide many data sets including: deformed shape, shoulder deflection, strap loads, cord loads, inflation pressures, and applied static load.In this overview, the 6m HIAD static load test series will be discussed in detail, including the 6m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally initial results and conclusions from the test series.

inflatable aerodynamic decelerator↗

NASA NTRS2014

Overview of the 6m HIAD Inflatable Structure and Flexible TPS Static Load Test Series

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). At the beginning of 2014, a 6m HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify the designs structural performace. The 6m HIAD structure was constructed in a stacked toroid configuration using nine inflatable torus segments composed of fiber reinforced thin films, which were joined together using adhesives and high strength textile woven structural straps to help distribute the loads throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for displacement of the inflatable structure as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The centerbody of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to provide many data sets including: deformed shape, shoulder deflection, strap loads, cord loads, inflation pressures, and applied static load.In this overview, the 6m HIAD static load test series will be discussed in detail, including the 6m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally initial results and conclusions from the test series.

TPS↗

NASA NTRS2025

Expanding the Measurement Capabilities of the mARC II Arc-Jet to Map the Operating Envelope for High-Enthalpy Air Flows

The mARC II is a 30 kW arc-jet facility at NASA Ames Research Center developed to produce high-enthalpy flows for low-cost technology development purposes. In this work, we introduce new measurement capabilities following the latest facility upgrades and begin characterizing the operating envelope for the two-disk arc-heater configuration with air as the working gas. The lower bound of the envelope corresponds to the lowest set current and the farthest sensor distance from the nozzle (𝐼set = 40 A, 𝑧 = 69 mm), while the upper bound corresponds to the highest current and closest sensor distance (𝐼set = 200 A, 𝑧 = 2 mm). Stagnation point heat flux was measured using a water-cooled Gardon gauge (⌀4.76 mm, 3/16" hemispherical), with values ranging from 15 ≤ q̇₀ ≤ 900 W/cm 2 . Stagnation pressure was measured using a water-cooled Pitot probe (⌀4.76 mm, 3/16" hemispherical), with values ranging from 18 ≤ p₀ ≤ 1100 Pa. The upgraded vacuum system is demonstrated to significantly extend the lower end of the operating envelope of the mARC II. Additionally, measurements of bulk enthalpy (3 ≤ h ≤ 13 MJ/kg) and stagnation enthalpy (4 ≤ h₀ ≤ 34 MJ/kg) demonstrate the mARC II facility's high-enthalpy capabilities. Completion of the operating envelope characterization is ongoing in tandem with the development of complementary diagnostic capabilities and numerical simulations.

instrumentation↗

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

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

Hannah S. Alpert↗

NASA NTRS2012

Flowfield Analysis of a Small Entry Probe (SPRITE) Tested in an Arc Jet

A novel concept of small size (diameter less than 15 inches) entry probes named SPRITE (Small Probe Re-entry Investigation for TPS Engineering) has been developed at NASA Ames Research Center (ARC). These flight probes have on-board data acquisition systems that have also been developed in parallel at NASA ARC by Greg Swanson1. Flight probes of this size facilitate testing over a wide range of conditions in arc jets available at NASA ARC, thereby fulfilling a 'test what you fly' paradigm. As indicated by the acronym, these probes, with suitably tailored trajectories, are primarily meant to be robotic flight test beds for TPS materials, although the design is flexible enough to accommodate additional objectives of flight-testing other vehicle subsystems. A first step towards establishing the feasibility of the SPRITE concept is to arc-jet test fully instrumented models at flight scale. In a follow-on to the Large-Scale Article Tests (LSAT2) performed in the 60 MW Interaction Heating Facility (IHF) in late 2008/early 2009, a full-scale model of Deep Space-2 (DS23) made of red oak was tested in the 20 MW Aerodynamic Heating Facility (AHF). There were no issues with mass capture by the diffuser for blunt bodies of roughly 15 inches diameter tested in the 18-inch nozzle of the AHF. Building on this initial success, two identical test articles - SPRITE-T1-1 and SPRITE-T1-2 (T1 indicating the choice of back shell geometry) - were fabricated, and one of them, SPRITE-T1-1, was tested in the AHF recently. Both these test articles, 14 inches in diameter, have a 45deg sphere-cone (like DS2) made of PICA bonded on to a 1/8th inch thick aluminum shell using RTV. The aft portion of the test article is a conical frustum (15deg cone angle) with LI-2200 bonded on to the aluminum shell. Each model is fully instrumented with: (a) thermocouples imbedded in plugs in the heat shield, (b) thermocouples bonded to the aluminum substructure; the thermocouples are distributed over the entire shell, and (c) a few strain gages. Data from some of the thermocouples and gages are acquired by the on-board data acquisition system (DAS), while data from the others are routed to the facility-provided DAS, thereby enabling a cross check on the in situ measurement capability. as inputs to v2.6.1 of the in-house materials thermal response code, FIAT

atmospheric entry↗

NASA NTRS2015

Overview of the 2nd Gen 3.7m HIAD Static Load Test

To support NASAs long term goal of landing humans on Mars, technologies which enable the landing of heavy payloads are being developed. Current entry, decent, and landing technologies are not practical for human class payloads due to geometric constraints dictated by current launch vehicle fairing limitations. Therefore, past and present technologies are now being explored to provide a mass and volume efficient solution to atmospheric entry, including Hypersonic Inflatable Aerodynamic Decelerators (HIADs). In October of 2014, a 3.7m HIAD inflatable structure with an integrated flexible thermal protection sys-tem (F-TPS) was subjected to a static load test series to verify the designs structural performance. The 3.7m HIAD structure was constructed in a 70 deg sphere-cone stacked-toroid configuration using eight inflatable tori, which were joined together using adhesives and high strength textile webbing to help distribute the loads throughout the inflatable structure. The inflatable structure was fabricated using 2nd generation structural materials that permit an increase in use temperature to 400 C+ as compared to the 250 C limitation of the 1st generation materials. In addition to the temperature benefit, these materials also offer a 40 reduction in structure mass. The 3.7m F-TPS was fabricated using high performance materials to protect the inflatable structure from heat loads that would be seen during atmospheric entry. The F-TPS was constructed of 2nd generation TPS materials increasing its heating capability from 35W sq cm to over 100W sq cm. This test article is the first stacked-torus HIAD to be fabricated and tested with a 70 deg sphere-cone. All previous stacked-torus HIADs have employed a 60o sphere-cone. To perform the static load test series, a custom test fixture was constructed. The fixture consisted of a structural tub rim with enough height to allow for dis-placement of the inflatable structure as loads were applied. The tub rim was attached to the floor to provide an airtight seal. The center body of the inflatable structure was attached to a pedestal mount as seen in Figure 1. Using an impermeable membrane seal draped over the test article, partial vacuum was pulled beneath the HIAD, resulting in a uniform static pressure load applied to the outer surface. During the test series an extensive amount of instrumentation was used to characterize deformed shape, shoulder deflection, strap loads, and cord loads as a function of structural configuration and applied static load. In this overview, the 3.7m HIAD static load test series will be discussed in detail, including the 3.7m HIAD inflatable structure and flexible TPS design, test setup and execution, and finally results and conclusions from the test series.

HIAD↗

NASA NTRS2013

Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Technology Development Overview

The successful flight of the Inflatable Reentry Vehicle Experiment (IRVE)-3 has further demonstrated the potential value of Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology. This technology development effort is funded by NASA's Space Technology Mission Directorate (STMD) Game Changing Development Program (GCDP). This paper provides an overview of a multi-year HIAD technology development effort, detailing the projects completed to date and the additional testing planned for the future. The effort was divided into three areas: Flexible Systems Development (FSD), Mission Advanced Entry Concepts (AEC), and Flight Validation. FSD consists of a Flexible Thermal Protection Systems (FTPS) element, which is investigating high temperature materials, coatings, and additives for use in the bladder, insulator, and heat shield layers; and an Inflatable Structures (IS) element which includes manufacture and testing (laboratory and wind tunnel) of inflatable structures and their associated structural elements. AEC consists of the Mission Applications element developing concepts (including payload interfaces) for missions at multiple destinations for the purpose of demonstrating the benefits and need for the HIAD technology as well as the Next Generation Subsystems element. Ground test development has been pursued in parallel with the Flight Validation IRVE-3 flight test. A larger scale (6m diameter) HIAD inflatable structure was constructed and aerodynamically tested in the National Full-scale Aerodynamics Complex (NFAC) 40ft by 80ft test section along with a duplicate of the IRVE-3 3m article. Both the 6m and 3m articles were tested with instrumented aerodynamic covers which incorporated an array of pressure taps to capture surface pressure distribution to validate Computational Fluid Dynamics (CFD) model predictions of surface pressure distribution. The 3m article also had a duplicate IRVE-3 Thermal Protection System (TPS) to test in addition to testing with the Aerocover configuration. Both the Aerocovers and the TPS were populated with high contrast targets so that photogrammetric solutions of the loaded surface could be created. These solutions both refined the aerodynamic shape for CFD modeling and provided a deformed shape to validate structural Finite Element Analysis (FEA) models. Extensive aerothermal testing has been performed on the TPS candidates. This testing has been conducted in several facilities across the country. The majority of the testing has been conducted in the Boeing Large Core Arc Tunnel (LCAT). HIAD is continuing to mature testing methodology in this facility and is developing new test sample fixtures and control methodologies to improve understanding and quality of the environments to which the samples are subjected. Additional testing has been and continues to be performed in the NASA LaRC 8ft High Temperature Tunnel, where samples up to 2ft by 2ft are being tested over representative underlying structures incorporating construction features such as sewn seams and through-thickness quilting. With the successful completion to the IRVE-3 flight demonstration, mission planning efforts are ramping up on the development of the HIAD Earth Atmospheric Reenty Test (HEART) which will demonstrate a relevant scale vehicle in relevant environments via a large-scale aeroshell (approximately 8.5m) entering at orbital velocity (approximately 7km/sec) with an entry mass on the order of 4MT. Also, the Build to Print (BTP) hardware built as a risk mitigation for the IRVE-3 project to have a "spare" ready to go in the event of a launch vehicle delivery failure is now available for an additional sub-orbital flight experiment. Mission planning is underway to define a mission that can utilize this existing hardware and help the HIAD project further mature this technology.

Hughes, Stephen J.↗

NASA NTRS2019

Challenges and Opportunities for Ensuring Entry System Technology Readiness for Ice Giants Probe Missions

The Ice Giants represent a distinct class of planets within our solar system, and appear to be similar to most exoplanets that have been detected thus far. Exploring Ice Giants in our Solar System would allow us to better understand their formation and evolution processes, and thus help establish scientific links to exoplanets. In situ exploration using probes similar to Galileo, along with an orbiter or a relay spacecraft, will require entry followed by deployment of the descent probe containing science instruments into Uranus or Neptune atmosphere. The challenge is not in the deployment of the probe, but in the atmospheric entry prior to deployment. The entry system has to have a capable, robust and efficient ablative thermal protection system (TPS) designed to protect the descent probe from the thermal and mechanical entry loads. Although entries into Ice Giants may not be as demanding as the Galileo entry at Jupiter, the entry environments will be more severe than environments for Mars, Sample Return missions, and Venus, and will therefore require robust TPS. While Galileo Probe’s success, nearly 25 years ago, should give us confidence, the recession data from the Galileo entry informs us that the entry environment was under-predicted and the design thickness was barely adequate. The lesson learned from the Galileo probe for future Ice Giant missions will require us to be cautious and demand a more robust design. The TPS technology used on the Galileo entry system no longer exists due to atrophy of manufacturing processes. Instead of attempting to revive Galileo-legacy TPS technology, NASA invested in a new and innovative TPS called HEEET (Heat-shield for Extreme Entry Environment Technology). HEEET has been matured, and is now ready to support future missions not only to the Ice Giants but also for Venus, high-speed sample return, and Saturn probe missions. This lead talk, intended for the technology section of the workshop, will cover entry, descent, and deployment (EDD), with an emphasis on entry. A brief history of the TPS challenges for extreme entry missions will be given, along with a quick overview of the concept of operations for EDD. The development and maturation of HEEET system capability will be described. Data gathered in ground-test facilities in the US will be highlighted to show that the technology is mature and ready for Ice Giant missions. All thermal protection systems carry some risk as a result of ground test limitations and Ice Giant missions present some unique challenges. These challenges are not only technical, but also due to limitations in the currently established manufacturing and integration. In addition, the concerns that arise due to the potential for atrophy for future Ice Giant missions a decade or more from now will be analyzed. Plausible avenues for mitigation will be presented. There are two planned companion presentations by Dr. Prabhu and Dr. Hwang that will dive deeper into the challenges and opportunities. This intended talk will set the stage for their presentations.

Venkatapathy, Ethiraj↗

NASA NTRS2019

Challenges and Opportunities For Ensuring Entry System Technology Readiness For Ice Giants Probe Missions

The Ice Giants represent a distinct class of planets within our solar system, and appear to be similar to most exoplanets that have been detected thus far. Exploring Ice Giants in our Solar System would allow us to better understand their formation and evolution processes, and thus help establish scientific links to exoplanets. In situ exploration using probes similar to Galileo, along with an orbiter or a relay spacecraft, will require entry followed by deployment of the descent probe containing science instruments into Uranus or Neptune atmosphere. The challenge is not in the deployment of the probe, but in the atmospheric entry prior to deployment. The entry system has to have a capable, robust and efficient ablative thermal protection system (TPS) designed to protect the descent probe from the thermal and mechanical entry loads. Although entries into Ice Giants may not be as demanding as the Galileo entry at Jupiter, the entry environments will be more severe than environments for Mars, Sample Return missions, and Venus, and will therefore require robust TPS. While Galileo Probe’s success, nearly 25 years ago, should give us confidence, the recession data from the Galileo entry informs us that the entry environment was under predicted and the design thickness was barely adequate. The lesson learned from Galileo probe for future Ice Giant missions will require us to be cautious and demand a more robust design. The TPS technology used on Galileo entry system no longer exists due to atrophy of manufacturing processes. Instead of attempting to revive Galileo-legacy TPS technology, NASA invested in a new and innovative TPS called HEEET (Heat-shield for Extreme Entry Environment Technology). HEEET has been matured, and is now ready to support future missions not only to the Ice Giants but also for Venus, high-speed sample return, and Saturn probe missions. This lead talk, intended for the technology section of the workshop, will cover entry, descent, and deployment (EDD), with an emphasis on entry. A brief history of the TPS challenges for extreme entry missions will be given along with a quick overview of the concept of operations for EDD. The development and maturation of HEEET system capability will be described. Data gathered in ground-test facilities in the US will be highlighted to show that the technology is mature and ready for Ice Giant missions. All thermal protection systems carry some risk as a result of ground test limitations and Ice Giant missions present some unique challenges. These challenges are not only technical, but also due to limitations in the currently established manufacturing and integration. In addition, the concerns that arise due to potential for atrophy for future Ice Giant mission a decade or more from now will be analyzed. Plausible avenues for mitigation will be presented. There are two companion planned presentations by Dr. Prabhu and Dr. Hwang will dive deeper in the challenges and opportunities. This intended talk will set the stage for their presentations.

Venkatapathy, Ethiraj↗

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 NTRS1979

Space shuttle SRB TPS protective paint test and evaluation in NASA Hot Gas Facility and AEDC Tunnel C

The results and outcome of thermal tests conducted to evaluate the performance of the protective coat of paint on the solid rocket booster (SRB) thermal protection system are discussed. A problem was uncovered during a series of tests on the SRB instrumentation islands in AEDC Tunnel C on 13 January 1979. The white protective paint or the Turco coating on the thermal protection system panels began to flow soon after the panels were exposed to the flow. This presented a serious problem especially since the critical pressure sensing, parachute opening baroswitches located on the frustum of the SRB were most likely to be contaminated by the paint flowing down the sides of the SRB nose cone. Because the first two flight articles were already completed, it was necessary to find a solution to the existing paint problem. It was found that all the coatings tested, except the Hypalon, had similar undesirable flow characteristics. Also even the Hypalon, which did not flow, would bubble up and disintegrate when it was applied on top of the new Turco. Recently, the Turco coating was removed from an MSA-:11 panel by dissolving the paint with a certain agent. This was done in two ways, by dissolving and removing almost all of the paint on one side of the panel and dissolving and removing about 50% of the paint on the other. The panel was then coated with Hypalon and tested as before in the Hot Gas Facility. No evidence of any paint flow nor any adverse performance of MSA was observed.

Karu, Z. S.↗

NASA NTRS2011

Development of a Sheathed Miniature Aerothermal Reentry Thermocouple for Thermal Protection System Materials

The Sheathed Miniature Aerothermal Reentry Thermocouple is a micro-miniature thermocouple for high temperature measurement in extreme environments. It is available for use in Thermal Protection System materials for ground testing and flight. This paper discusses the heritage, and design of the instrument. Experimental and analytical methods used to verify its performance and limitations are described.

Reentry↗

NASA NTRSDate not supplied

Overview of NASA's Detailed Investigation into the MEDLI2 Flight Data

- MEDLI2 “Deep Dive” - Why a focused investigation? - Deeper understanding of the MEDLI2 dataset, including TPS in-depth temperatures, backshell pressure, and radiative heat flux - Validate new tools for future Mars missions including Mars Sample Return and Humans-to-Mars - MEDLI2 is considerably more complex than MEDLI - Fully leverage the significant investment in the MEDLI2 instrumentation suite - Core research areas: - PICA-NuSil (PICA-N) testing, model development, and validation - Detailed aeroheating investigations - Sensor fusion of collocated measurements - High-fidelity analyses and uncertainty quantification - Trajectory and aerodynamics investigations

Tom West↗

NASA NTRS2006

Overview of a Proposed Flight Validation of Aerocapture System Technology for Planetary Missions

Aerocapture System Technology for Planetary Missions is being proposed to NASA's New Millennium Program for flight aboard the Space Technology 9 (ST9) flight opportunity. The proposed ST9 aerocapture mission is a system-level flight validation of the aerocapture maneuver as performed by an instrumented, high-fidelity flight vehicle within a true in-space and atmospheric environment. Successful validation of the aerocapture maneuver will be enabled through the flight validation of an advanced guidance, navigation, and control system as developed by Ball Aerospace and two advanced Thermal Protection System (TPS) materials, Silicon Refined Ablative Material-20 (SRAM-20) and SRAM-14, as developed by Applied Research Associates (ARA) Ablatives Laboratory. The ST9 aerocapture flight validation will be sufficient for immediate infusion of these technologies into NASA science missions being proposed for flight to a variety of Solar System destinations possessing a significant planetary atmosphere.

Keys, Andrew S.↗