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NASA NTRSDate not supplied

Developments on Reusable TPS Materials Based Upon Shuttle Tile

The insulating tiles used on the Space Shuttler Orbiter, principally AETB (Alumina Enhanced Thermal Barrier), have been seen by new vehicle developers in the aerospace industry as a fast-track approach for fielding a reusable thermal protection system (TPS) given the extensive use by both the Orbiter program and the current Orion vehicle. NASA has, and is, supporting the industry through technology transfer and commercial crew and payloads programs. However, the reusable TPS community is rapidly encountering challenges rooted in Orbiter’s and Orion’s dependence on legacy material and processes. With the aim of alleviating these issues, improving availability of reusable TPS, and strengthening the nation’s space economy, ongoing efforts at NASA Ames have focused on understanding process-property relationships in AETB as well as novel materials and technologies for reusable TPS. The first portion of this talk will address the effects of raw materials and processing conditions on the properties of AETB-8, including the raw materials used, mixing and casting parameters, and thermal history. Understanding these relationships, and how they extend to aerothermal performance, is instrumental in alleviating supply chain constraints and improving manufacturability. The implications of these finding on development of novel TPS materials and systems will be discussed. AETB will likely continue to fill a niche of relatively high performance with associated costs. There exists, therefore, a gap in developed reusable TPS for a lower cost, moderate performing material system. The later portion of this talk will touch on current development efforts at Ames focusing on reduced cost and ease of integration, which are being pursued in addition to AETB and higher performing systems. In addition to summarizing these efforts, an outlook for these projects and collaborations will be given.

Peter Edward Marshall↗

NASA NTRSDate not supplied

Developments on Reusable TPS Materials Based Upon Shuttle Tile

Thermal Protection Systems↗

NASA NTRS2016

Orion EFT-1 Cavity Heating Tile Experiments and Environment Reconstruction

Developing aerothermodynamic environments for deep cavities, such as those produced by micrometeoroids and orbital debris impacts, poses a great challenge for engineers. In order to assess existing cavity heating models, two one-inch diameter cavities were flown on the Orion Multi-Purpose Crew Vehicle during Exploration Flight Test 1 (EFT1). These cavities were manufactured with depths of 1.0 in and 1.4 in, and they were both instrumented. Instrumentation included surface thermocouples upstream, downstream and within the cavities, and additional thermocouples at the TPS-structure interface. This paper will present the data obtained, and comparisons with computational predictions will be shown. Additionally, the development of a 3D material thermal model will be described, which will be used to account for the three-dimensionality of the problem when interpreting the data. Furthermore, using a multi-dimensional inverse heat conduction approach, a reconstruction of a time- and space-dependent flight heating distribution during EFT1 will be presented. Additional discussions will focus on instrumentation challenges and calibration techniques specific to these experiments. The analysis shown will highlight the accuracies and/or deficiencies of current computational techniques to model cavity flows during hypersonic re-entry.

Salazar, Giovanni↗

NASA NTRSDate not supplied

Arc-jet Testing of Continuously Woven Aeroshells–Spiderweave–for Adaptable Deployable Entry Placement Technology

The Adaptable, Deployable Entry Placement Technology (ADEPT) is an entry system that, when deployed, creates a large, umbrella-like drag surface that can be used to decelerate spacecraft during atmospheric entry, resulting in lower entry loads and heating than a rigid aeroshell of equivalent mass. Until needed for entry, the system can remain stowed, enabling missions to overcome launch vehicle fairing diameter constraints. 12-inch diameter umbrella-like articles, similar to a small reentry probe using ADEPT, were tested in the Ames 20-MWAerodynamic Heating Facility. Performance of the test articles built with a novel, continuously woven aeroshell construction (called Spiderweave) is compared against those built with legacy construction styles. In addition, candidate thermal protective layer treatments applied to structurally sensitive regions of the aeroshell are assessed using data from on-board instrumentation as well as facility optical instruments.

ADEPT↗

NASA NTRS2020

Thermal Protection System Materials for Sample Return Missions

This white paper from the Thermal Protection System (TPS) community to the NRC Decadal Survey Sub-Panels provides an overview of TPS materials needed for future Sample Return (SR) missions. We consider the capability of heritage TPS material used by recent SR missions and identify appropriate materials for future SR missions. A prime conclusion is that the current TPS materials, if properly maintained, offer good low-density solutions for lower velocity (<13.5 km/s) sample return missions without planetary protection back-protection concerns. Furthermore, missions that have a larger capsule, a higher entry speed (>13.5 km/s), or back-protection concerns will leverage recently developed mid-density TPS materials. To maintain NASA’s Sample Return capabilities in the coming decade, we recommend that NASA continue to invest in sustainment of relevant TPS materials, as well as ground-test facilities, predictive entry modeling, and flight instrumentation.

mission design↗

NASA NTRS1995

Development of processing techniques for advanced thermal protection materials

The main purpose of this work has been in the development and characterization of materials for high temperature applications. Thermal Protection Systems (TPS) are constantly being tested, and evaluated for increased thermal shock resistance, high temperature dimensional stability, and tolerance to environmental effects. Materials development was carried out through the use of many different instruments and methods, ranging from extensive elemental analysis to physical attributes testing. The six main focus areas include: (1) protective coatings for carbon/carbon composites; (2) TPS material characterization; (3) improved waterproofing for TPS; (4) modified ceramic insulation for bone implants; (5) improved durability ceramic insulation blankets; and (6) ultra-high temperature ceramics. This report describes the progress made in these research areas during this contract period.

Selvaduray, Guna S.↗

NASA NTRS2014

Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions: An Update of the Technology Maturation Effort

This presentation will update the community on the development of conformal ablative TPS. As described at IPPW-10, in FY12, the CA-TPS element focused on establishing materials requirements based on MSL-type and COTS Low Earth orbit (LEO) conditions (q 250 Wcm2) to develop and deliver a conformal ablative TPS. This involved downselecting, manufacturing and testing two of the best candidate materials, demonstrating uniform infiltration of resins into baseline 2-cm thick carbon felt, selecting a primary conformal material formulation based on novel arc jet and basic material properties testing, developing and demonstrating instrumentation for felt-based materials and, based on the data, developing a low fidelity material response model so that the conformal ablator TPS thickness for missions could be established. In addition, the project began to develop Industry Partnerships. Since the nominal thickness of baseline carbon felts was only 2-cm, a partnership with a rayon felt developer was made in order to upgrade equipment, establish the processes required and attempt to manufacture 10-cm thick white goods. A partnership with a processing house was made to develop the methodology to carbonize large pieces of the white goods into 7.5-cm thick carbon felt.In FY13, more advanced testing and modeling of the downselected conformal material was performed. Material thermal properties tests and structural properties tests were performed. The first 3 and 4-point bend tests were performed on the conformal ablator as well as PICA for comparison and the conformal ablator had outstanding behavior compared to PICA. Arc jet testing was performed with instrumented samples of both the conformal ablator and standard PICA at heating rates ranging from 40 to 400 Wcm2 and shear as high as 600 Pa. The results from these tests showed a remarkable improvement in the thermal penetration through the conformal ablator when compared to PICAs response. The data from these tests were used to develop a mid-fidelity thermal response model. Additional arc jet testing in the same conditions on various seam designs were very successful in showing that the material could be joined with a minimum of adhesive and required no complicated gap and gap filler design for installation. In addition, the partnership with industry to manufacture thicker rayon felt was very successful. The vendor made a 2-m wide by 30-m long sample of 10-cm thick rayon felt. When carbonized, the resulting thickness was over 7.5-cm thick, nearly 4 times the thickest off-the-shelf carbon felt. In FY14, the project has initiated a partnership with another vendor to begin the scale-up manufacturing effort. This year, the vendor will duplicate the process and manufacture at the current scale for comparison with NASA-processed materials. Properties testing and arc jet testing will be performed on the vendor-processed materials. Planning for manufacturing large, 1-m x 1-m, panels will begin as well. In FY15, the vendor will then manufacture large panels and the project will build a 2-m x 2-m Manufacturing Demonstration Unit (MDU).

thermal protection systems↗

NASA NTRS2014

An Update to a Conformal Ablative Thermal Protection System for Planetary and Human Exploration Missions

As described at IPPW-10, in FY12, the CA-TPS element focused on establishing materials requirements based on MSL-type and COTS Low Earth orbit (LEO) conditions (q 250 Wcm2) to develop and deliver a conformal ablative TPS. This involved down selecting, manufacturing and testing two of the best candidate materials, demonstrating uniform infiltration of resins into baseline 2-cm thick carbon felt, selecting a primary conformal material formulation based on novel arc jet and basic material properties testing, developing and demonstrating instrumentation for felt-based materials and, based on the data, developing a low fidelity material response model so that the conformal ablator TPS thickness for missions could be established. In addition, the project began to develop Industry Partnerships. Since the nominal thickness of baseline carbon felts was only 2-cm, a partnership with a rayon felt developer was made in order to upgrade equipment, establish the processes required and attempt to manufacture 10-cm thick white goods. A partnership with a processing house was made to develop the methodology to carbonize large pieces of the white goods into 7.5-cm thick carbon felt. In FY13, more advanced testing and modeling of the down selected conformal material was performed. Material thermal properties tests and structural properties tests were performed. The first 3 and 4-point bend tests were performed on the conformal ablator as well as PICA for comparison and the conformal ablator had outstanding behavior compared to PICA. Arc jet testing was performed with instrumented samples of both the conformal ablator and standard PICA at heating rates ranging from 40 to 400 Wcm2 and shear as high as 600 Pa. The results from these tests showed a remarkable improvement in the thermal penetration through the conformal ablator when compared to PICAs response. The data from these tests were used to develop a mid-fidelity thermal response model. Additional arc jet testing in the same conditions on various seam designs were very successful in showing that the material could be joined with a minimum of adhesive and required no complicated gap and gap filler design for installation. In addition, the partnership with industry to manufacture thicker rayon felt was very successful. The vendor made a 2-m wide by 30-m long sample of 10-cm thick rayon felt. When carbonized, the resulting thickness was over 7.5-cm thick, nearly 4 times the thickest off-the-shelf carbon felt. In FY14, the project has initiated a partnership with another vendor to begin the scale-up manufacturing effort. This year, the vendor will duplicate the process and manufacture at the current scale for comparison with NASA-processed materials. Properties testing and arc jet testing will be performed on the vendor-processed materials. Planning for manufacturing large, 1-m x 1-m, panels will begin as well. In FY15, the vendor will then manufacture large panels and the project will build a 2-m x 2-m Manufacturing Demonstration Unit (MDU).

Ablative materials↗

NASA NTRS1984

Thermal mathematical modeling and system simulation of Space Shuttle less subsystem

Applications, validation tests, and upgrades of the two- and three-dimensional system level thermal mathematical system simulation models (TMSSM) used for thermal protection system (TPS) analyses are described. The TMSSM were developed as an aid to predicting the performance requirements and configurations of the Shuttle wing leading edge (WLE) and nose cone (NC) TPS tiles. The WLE and its structure were subjected to acoustic, thermal/vacuum, and air loads tests to simulate launch, on-orbit, and re-entry behavior. STS-1, -2 and -5 flight data led to recalibration of on-board instruments and raised estimates of the thermal shock at the NC and WLE. Baseline heating data are now available for the design of future TPS.

Chao, D. C.↗

NASA NTRSDate not supplied

Variance Decomposition of MEDLI2 Reconstructed Heating Using Neural Networks

The Mars Entry, Descent, and Landing Instrumentation (MEDLI2) sensor suite collected data during entry of the Mars 2020 Perseverance rover into Mars’ atmosphere. This suite included a network of MEDLI2 Instrumented Sensor Plugs (MISPs). Each MISP was comprised of a cylinder made of Thermal Protection System (TPS) material with 1-3 embedded thermocouples (TCs), and it was flush mounted into the heatshield or backshell. Data from these in-depth TCs were used to reconstruct the aeroheating environment of the vehicle throughout entry. Surface heating was posed as an inverse problem, with the goal of estimating the surface heating by minimizing an objective function of the difference between MISP temperature measurements during flight and the temperature predictions derived from the Fully Implicit Ablation and Thermal response (FIAT) program. Given an aerothermal environment, FIAT calculates the material response and provides in-depth temperatures throughout the TPS material. To achieve the reverse, an internal tool called FIAT_Opt runs through multiple different environments until the output temperature at the TC depth closely matches the flight data. 95% confidence intervals on the reconstructed surface heating were obtained using Monte Carlo analysis, in which uncertainties in the thermocouple depth and the TPS material properties (e.g., density, thermal conductivity, heat capacity, emissivity) based on flight-lot material testing were included. A variance decomposition method using Sobol indices was employed to assess the sensitivity of the reconstructed peak heating to the TC placement and material property uncertainties. Variance decomposition was found to require tens of thousands of FIAT_Opt runs in order for the Sobol indices to converge. With a single FIAT_Opt run taking on the order of 40 minutes, the required number of computations would take months to complete, even if using multiple CPUs. To mitigate this problem, three machine learning models (ridge regression with cross-validation, random forest regression, and a deep neural network) were trained and tested using the 2000 Monte Carlo runs that were already completed. A subset of 1600 runs were used to train the model (i.e., training set), while the remaining 400 runs were used as the test set. The predictions from the deep neural network (DNN) on the test set showed nearly perfect agreement to the actual values computed with FIAT_Opt (R2 > 0.99). Using the DNN as a surrogate model, the variance decomposition using 50,000 runs was completed within minutes. The resulting Sobol indices showed that the reconstructed peak surface heating was most sensitive to the uncertainties in the thermal conductivity (ST = 0.37) and heat capacity (ST = 0.26). This method can be leveraged to provide requirements for material property measurements needed to improve the accuracy of surface heating prediction and ultimately lead to the reduction of design margins in the future. This presentation will include background on the MEDLI2 suite; the method used for inverse heating estimation; the way that material property uncertainties were accounted for using Monte Carlo analysis; a brief background on variance decomposition; the motivation for using machine learning in this context; how a neural network was trained on the data to enable variance decomposition in a fraction of the time; and the variance decomposition results for one of the MISPs.

Hannah Alpert↗

NASA NTRS2016

Post-Flight Assessment of Avcoat Thermal Protection System for the Exploration Flight Test-1

On December 5, 2014 NASA conducted the first flight test of its next generation human-class Orion spacecraft. The flight was called the Exploration Flight Test -1 (EFT-1) which lasted for 4 hours and culminated into a re-entry trajectory at 9 km/s. This flight test of the 5-meter Orion Crew Module demonstrated various sub-systems including the Avcoat ablative thermal protection system (TPS) on the heat shield. The Avcoat TPS had been developed from the Apollo-era recipe with a few key modifications. The engineering for thermal sizing was supported by modeling, analysis, and ground tests in arc jet facilities. This paper will describe a postlfight analysis plan and present results from post-recovery inspections, data analysis from embedded sensors, TPS sample extraction and characterization in the laboratory. After the recovery of the vehicle, a full photographic survey and surface scans of the TPS were performed. The recovered vehicle showed physical evidence of flow disturbances, varying degrees of surface roughness, and excessive recession downstream of compression pads. The TPS recession was measured at more than 200 locations of interest on the Avcoat surface. The heat shield was then processed for sample extraction prior to TPS removal using the 7-Axis Milling machine at Marshall Space Flight Center. Around 182 rectangular TPS samples were extracted for subsequent analysis and investigation. The final paper will also present results of sample analysis. The planned investigation includes sidewall imaging, followed by image analysis to characterize TPS response by quantifying different layers in the char and pyrolysis zones. A full postmortem of the instrumentation and sensor ports will also be performed to confirm no adverse effects due to the sensors themselves. A subset of the samples will undergo structural testing and perform detailed characterization of any cracks and integrity of gore seams. Finally, the material will be characterized with layer-by-layer density measurements and SEM investigations to evaluate material morphology at microstructural level including identification of elements and compounds.

Aerothermodynamics↗

NASA NTRS2019

Loftid Aeroshell Engineering Development Unit Structural Testing

NASA’s Hypersonic Inflatable Aerodynamic Decelerator (HIAD) technology was selected for a Technology Demonstration Mission under the Space Technology Mission Directorate in 2017. HIAD is an enabling technology that can facilitate atmospheric entry of heavy payloads to planets such as Earth and Mars using a deployable aeroshell. The deployable nature of the HIAD technology allows it to avoid the size constraints imposed on current rigid aeroshell entry systems. This enables use of larger aeroshells resulting in increased entry system performance (e.g. higher pay-load mass and/or volume, higher landing altitude at Mars). The Low Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) is currently scheduled for late-2021. LOFTID will be launched out of Vandenberg Air Force Base as a secondary payload on an Atlas V rocket. The flight test features a 6m diameter, 70-deg sphere-cone aeroshell and will provide invaluable high-energy orbital re-entry flight data. This data will be essential in supporting the HIAD team to mature the technology to diameters of 10m and greater. Aeroshells of this scale are applicable to potential near-term commercial applications and future NASA missions. Currently the LOFTID project has completed fabrication of the engineering design unit (EDU) inflatable structure (IS) and the flexible thermal protection system (F-TPS). These two components along with the rigid nose and center body comprise the HIAD aeroshell system. This EDU aeroshell is the precursor to the LOFTID aeroshell that will be used for flight. The EDU was built to verify the design given the subtle differences between the LOFTID aeroshell and past aeroshell designs that have been fabricated under the NASA HIAD project. To characterize the structural performance of the LOFTID aeroshell design, three structural tests will be performed. The first test to be conducted is static load testing, which will induce a uniform load across the forward surface of the aeroshell to simulate the expected pressure forces during atmospheric entry. The IS integrated with the rigid center body will first be tested alone to provide data for analytical model correlation, and then the F-TPS will be integrated for a second series of static load testing of the full aeroshell system. Instrumentation will be employed during the test series to measure component loads during testing, and a laser scanner will be used to generate a 3D map of the aeroshell surface to verify that the shape of the structure is acceptable at the simulated flight loads. After static load testing, pack and deployment testing will be conducted multiple times on the integrated system to demonstrate the aeroshell’s ability to fit within the required packed volume for the LOFTID mission without experiencing significant damage. Finally, the aeroshell will undergo modal testing to characterize its structural response. This presentation will discuss the setup and execution of each of the three tests that the EDU aeroshell will undergo. In addition, initial results of the testing will be presented outlining key findings as LOFTID moves for-ward with fabrication of the flight aeroshell.

Swanson, G. T.↗

NASA NTRSDate not supplied

Thermal Protection Materials and Systems at NASA Ames Research Center

Thermal Protection Systems (TPS) are critical for enabling NASA missions involving high-speed atmospheric flight where the entries usually include descending into the atmosphere followed by a trajectory that aims to burn off energy and result in a controlled landing. NASA Ames focuses on qualifying and certifying TPS for current missions, sustaining TPS for future missions, and developing new TPS for upcoming missions where a heritage solution is not viable. More recently there is also a focus on advancing and transferring technologies that can benefit both commercial and government space needs.Developing mature thermal protection systems is a lengthy process involving advanced tools, extensive research, and testing. Design and analysis tools are used to predict aerothermal environments, aid the design of test and flight hardware, and support the testing for the thermal/mechanical response of thermal protection systems. More recently, advances in computational methods help reduce the time and cost of technological advances, aid in optimized material architecture design, and improve material properties and performance. While high-enthalpy testing that simulates the conditions of space flight remains essential for the evaluation and development of TPS materials, computational tools are already showing promise in reducing the need for widespread testing and can help fast-trackthe design cycle.With the exploration of new destinations EDL instrumentation remains an important element of the heatshield and NASA Ames and partners have developed and delivered instrumentation flight hardware in support of recent Mars missions,including Mars Science Lab (MSL) and Mars 2020 (M2020),as well as Artemis Orion. Sensors installed on the heatshield and backshell of spacecraft provide coveted information about the aerodynamic and aerothermal environment during entry.Overthe years NASA Ames has brought several reusable and ablative TPS materials to a level of readiness to hand off to missions and the Thermal Protection Materials branch continues to serve as a TPS steward for the agency.

Thermal protection materials↗

NASA NTRSDate not supplied

Dragonfly Entry Aerosciences Measurements (DrEAM) Suite Science Objectives

NASA Ames and Langley are partnering with DLR to propose a comprehensive instrumentation suite known as the Dragonfly Entry Aerosciences Measurements (DrEAM). DrEAM will provide key aerothermodynamic data and performance analysis for Dragonfly’s forebody and backshell Thermal Protection System (TPS), and also includes a DLR-provided Data Acquisition System (DAS). Titan's atmosphere predominantly consists of nitrogen (~ 98% by mole) with small amounts of methane (~ 2% by mole). CN is a strong radiator and is found in nonequilibrium concentrations for Titan entry, the modeling of which has proven to be a difficult task. The DrEAM instrumentation suite will significantly advance the state-of-the-art not only by documenting the environment and performance of Dragonfly’s entry system but also by making key measurements in Titan’s atmosphere for the first time, thus providing new benchmark data applicable to entry science more generally.

Aaron Brandis↗

NASA NTRSDate not supplied

Thermal Protection Materials and Systems at NASA Ames Research Center

Thermal Protection Systems (TPS) are critical for enabling NASA missions involving high-speed atmospheric flight where the entries usually include descending into the atmosphere followed by a trajectory that aims to burn off energy and result in a controlled landing. NASA Ames focuses on qualifying and certifying TPS for current missions, sustaining TPS for future missions, and developing new TPS for upcoming missions where a heritage solution is not viable. More recently there is also a focus on advancing and transferring technologies that can benefit both commercial and government space needs. Developing mature thermal protection systems is a lengthy process involving advanced tools, extensive research, and testing. Design and analysis tools are used to predict aerothermal environments, aid the design of test and flight hardware, and support the testing for the thermal/mechanical response of thermal protection systems. More recently, advances in computational methods help reduce the time and cost of technological advances, aid in optimized material architecture design, and improve material properties and performance. While high-enthalpy testing that simulates the conditions of space flight remains essential for the evaluation and development of TPS materials, computational tools are already showing promise in reducing the need for widespread testing and can help fast-track the design cycle. With the exploration of new destinations EDL instrumentation remains an important element of the heatshield and NASA Ames and partners have developed and delivered instrumentation flight hardware in support of recent Mars missions (MSL and M2020) as well as Artemis Orion. Sensors installed on the heatshield and backshell of spacecraft provide coveted information about the aerodynamic and aerothermal environment during entry. Over the years NASA Ames has brought several reusable and ablative TPS materials to a level of readiness to hand off to missions and the branch continues to serve as a TPS steward for the agency. This presentation will cover current developments in the above areas that have enabled recent missions, and look to future TPS needs for missions such as Ice Giants.

thermal protection materials↗

NASA NTRSDate not supplied

Overview of the Dragonfly Entry Aerosciences Measurements (DrEAM) Suite

NASA Ames and Langley arepartnering with DLR to propose acomprehensive instrumentation suite known as the Dragonfly Entry Aerosciences Meas-urements (DrEAM). DrEAM being the first competed mission to fly EDL instrumentation as part of NASA’s Engineering Science Investigation (ESI).DrEAM will provide key aerothermodynamicdata and performance analysisfor Dragonfly’s forebody and backshell therma lprotection system (TPS),and also includes a DLR-provided Data Acquisition System(DAS).Titan’s atmosphere predominantly consists of nitro-gen (~98% by mole) with small amounts of methane(~2% by mole)and other trace gases. CN is a strong ra-diator and is found in nonequilibriumconcentrationsfor Titan entry. The accurate modeling of nonequilibrium CN radiation has proven to be a difficult task. Prompted by the Huygensmission, many experimental campaigns and analyses were performed to better understand the aerothermal environments experience by the probe dur-ing Titan entry[1].However, the Huygens probe carried no heatshieldinstrumentation. Therefore,the DrEAM instrumentation suite will significantly advance the state-of-the-art not only by documenting theenviron-ment and performance of Dragonfly’sentry system but also by making keymeasurements in Titan’s atmos-phere for thefirst time, thus providing new benchmark dataapplicable to entry science more generally. Current Measurement Goals: Aerothermal envi-ronments and TPS responsewill be measured using sen-sors similarto the Mars Entry, Descent, and Landing In-strumentation2 (MEDLI2) Integrated Sensor Plug(MISP) and the COMbined Aerothermaland Radiome-ter Sensor (COMARS)suite[2], with the latter supplied by DLR.For MEDLI2, MISP usedembedded thermo-couples (TCs) todirectly measure in-depth temperature of theTPS at several locations,which can also be used to infer surface environmentsvia inverse analysis. For DrEAM, the MISPstyle plugswill be known as Drag-onfly Sensors for Aero-Thermal Reconstruction (Drag-STR)plugs. On Schiaparelli, the COMARSsuite in-cludedthree total surface-mounted heatflux sensors, three pressure sensors, and one radiometer. For DrEAM, the COMARS package will be known as COmbined Sensor System for Titan Atmosphere (COSSTA). Atmospheric density measurements and capsuleaerodynamic data will be obtained throughthe onboard Inertial Measurement Unit (IMU),supple-mented by pressure transducers similar tothose used by the MEDLI Mars Entry AtmosphericData System (MEADS).The DrEAM pressure sensors will be known as Dragonfly Atmospheric Flight Transducers(DrAFT) Both DragSTR and DrAFThave flight heritagefrom MISP and MEADS on the MSLand Mars 2020mis-sions, and the COMARS suite successfullyflew on the ESA Schiaparelli EDMlander. A preliminary layout of the sensors is shown in Fig. 1. Because Dragonflyuses the same aeroshell provider (i.e., LockheedMartin) and materials for the TPS,with what are expected to be sim-ilar thicknesses as MSL and Mars 2020on both the heat shield and backshell, theDrEAMinstrumentation will look to utilize the same techniques and processes as de-veloped byMEDLI and MEDLI2for vehicle integra-tion. This commonality alsoenables DrEAMto lever-age the extensiveground test qualifications performed forMEDLI and MEDLI2 and claim substantial heritagefor this system.

A Brandis↗

NASA NTRSDate not supplied

TPS and Entry Systems Technologies for Future Mars and Titan Exploration

During the past twenty years, NASA effectively erased earlier Mars mishaps with six successful missions to the Red Planet. These missions delivered one orbiter and five payloads to the surface. Those payloads included three rovers, Spirit (2004) which roamed 11 years, Opportunity (2004) which roamed nearly 15 years, and Curiosity (2012) which is in its ninth year, along with two landers, Phoenix (2008) and InSIGHT (2018). In July, the Mars 2020 mission will send another large rover, Perseverance, which will land in 2021. The InSIGHT mission even demonstrated the capability to send CubeSats along to help with communication back to the surface. NASA has demonstrated the capability to land a metric ton of vehicles plus science instruments on Mars and expects that the same technologies will be equally successful landing Dragonfly on Titan in the 2030’s. The thermal protection systems (TPS) used on the Mars missions are sufficiently developed and matured to continue furthering science on both Mars and Titan, assuming that the TPS materials are sustained by industry. The purpose of this white paper is to encourage further exploration and science on both Mars and Titan because we have the technologies to support them. In addition, we will look forward to human exploration of Mars and identify the improvements in TPS materials required to facilitate landing the larger payloads.

Thermal Protection Systems↗

NASA NTRS2012

STS-133/ET-137 Tanking Test Photogrammetry Assessment

Following the launch scrub of Space Shuttle mission STS-133 on November 5, 2010, an anomalous condition of cracked and raised thermal protection system (TPS) foam was observed on the External Tank (ET). Subsequent dissection of the affected TPS region revealed cracks in the feet of two Intertank (IT) metallic stringers. An extensive investigation into the cause(s) and corrective action(s) for the cracked stringers was initiated, involving a wide array of material and structural tests and nondestructive evaluations, with the intent to culminate into the development of flight rational. One such structural test was the instrumented tanking test performed on December 17, 2010. The tanking test incorporated two three-dimensional optical displacement measurement systems to measure full-field outer surface displacements of the TPS surrounding the affected region that contained the stringer cracks. The results showed that the radial displacement and rotation of the liquid oxygen (LO2) tank flange changed significantly as the fluid level of the LO2 approached and passed the LO2 tank flange.

Oliver, Stanley T.↗