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NASA NTRS2015

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

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

Spacecraft↗

NASA NTRS2015

Alternative High Performance Polymers for Ablative Thermal Protection Systems

Ablative thermal protection systems are commonly used as protection from the intense heat during re-entry of a space vehicle and have been used successfully on many missions including Stardust and Mars Science Laboratory both of which used PICA - a phenolic based ablator. Historically, phenolic resin has served as the ablative polymer for many TPS systems. However, it has limitations in both processing and properties such as char yield, glass transition temperature and char stability. Therefore alternative high performance polymers are being considered including cyanate ester resin, polyimide, and polybenzoxazine. Thermal and mechanical properties of these resin systems were characterized and compared with phenolic resin.

Ablator↗

NASA NTRS2016

Conformal Ablative Thermal Protection Systems (CA-TPS) for Venus and Saturn Backshells

This poster provides an overview of the work performed to date on the Conformal Ablative TPS (CA-TPS) element of the TPSM project out of GCDP. Under this element, NASA is developing improved ablative TPS materials based on flexible felt for reinforcement rather than rigid reinforcements. By replacing the reinforcements with felt, the resulting materials have much higher strain-to-failure and are much lower in thermal conductivity than their rigid counterparts. These characteristics should allow for larger tile sizes, direct bonding to aeroshells and even lower weight TPS. The conformal phenolic impregnated carbon felt (C-PICA) is a candidate for backshell TPS for both Venus and Saturn entry vehicles.

Saturn↗

NASA NTRS2011

An Interview with Joe McMann: His Life Lessons

Pica Kahn conducted "An Interview with Joe McMann: His Life Lessons" on May 23, 2011. With over 40 years of experience in the aerospace industry, McMann has gained a wealth of knowledge. Many have been interested in his biography, progression of work at NASA, impact on the U.S. spacesuit, and career accomplishments. This interview highlighted the influences and decision-making methods that impacted his technical and management contributions to the space program. McMann shared information about the accomplishments and technical advances that committed individuals can make.

McMann, Joe↗

NASA NTRS2010

Interview with Dean Eppler

Pica Kahn interviewed Dr. Dean Eppler. The interview highlighted the personal and professional influences that have impacted Eppler's contributions to the space program.

Eppler, Dean↗

NASA NTRS2017

NASA's Advanced TPS Materials and Technology Development: Multi-Functional Materials and Systems for Space Exploration

NASA's future missions will be more demanding. They require materials to be mass efficient, robust, multi-functional, scalable and able to be integrated with other subsystems to enable innovative missions to accomplish future science missions. Thermal protection systems and materials (TPSM) are critical for the robotic and human exploration of the solar system when it involves entry. TPSM is a single string system with no back-up. Mass efficiency and robustness are required. Integration of TPSM with the aeroshell is both a challenge and an opportunity. Since 2010, NASA's Space Technology Mission Directorate has invested in innovative new materials and systems across a spectrum of game changing technologies. In this keynote address, we plan to highlight and present our successful approaches utilized in developing four different materials and system technologies that use innovative new manufacturing techniques to meet mission needs. 3-D weaving and felt manufacturing allowed us to successfully propose new ways of addressing TPSM challenges. In the 3-D MAT project, we developed and delivered a multi-functional TPS materials solution, in under three years that is an enabler for Lunar Capable Orion Spacecraft. Under the HEEET project, we are developing a robust heat-shield that can withstand extreme entry conditions, both thermally and mechanically, for entry at Venus, Saturn or higher speed sample return missions. The improved efficiency of HEEET allows science missions entry at much reduced G'loads enabling delicate science instruments to be used. The ADEPT concept is a foldable and deployable entry system and the critical component is a multi-functional fabric that is foldable and deployable and also functions as a mechanical aeroshell and a TPS. The fourth technology we will highlight involves felt to address integration challenges of rigid ablative system such as PICA that was used on MSL. The felt technology allows us to develop a compliant TPS for easy integration. The above four technology developments have focused on mission infusion as the success criteria. These technologies are in different stages of mission infusion. These innovations have led to new mission concepts to be proposed in the future. In our keynote address we will present approaches we have employed throughout the project to create the bridge to transition from low TRL to mission infusion and to overcome the traditional TRL valley of death.

TPS↗

NASA NTRS2017

Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) Sensor Suite

The Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) sensor suite seeks to address the aerodynamic, aerothermodynamic, and thermal protection system (TPS) performance issues during atmospheric entry, descent, and landing of the Mars 2020 mission. Based on the highly successful instrumentation suite that flew on Mars Science Laboratory (MEDLI), the new sensor suite expands on the types of measurements and also seeks to answer questions not fully addressed by the previous mission. Sensor Package: MEDLI2 consists of 7 pressure transducers, 17 thermal plugs, 2 heat flux sensors, and one radiometer. The sensors are distributed across both the heatshield and backshell, unlike MEDLI (the first sensor suite), which was located solely on the heat-shield. The sensors will measure supersonic pressure on the forebody, a pressure measurement on the aftbody, near-surface and in-depth temperatures in the heatshield and backshell TPS materials, direct total heat flux on the aftbody, and direct radiative heating on the aftbody. Instrument Development: The supersonic pressure transducers, the direct heat flux sensors, and the radiometer all were tested during the development phase. The status of these sensors, including the piezo-resistive pressure sensors, will be presented. The current plans for qualification and calibration for all of the sensors will also be discussed. Post-Flight Data Analysis: Similar to MEDLI, the estimated flight trajectory will be reconstructed from the data. The aerodynamic parameters that will be reconstructed will be the axial force coefficient, freestream Mach number, base pressure, atmospheric density, and winds. The aerothermal quantities that will be determined are the heatshield and backshell aero-heating, turbulence transition across the heatshield, and TPS in-depth performance of PICA. By directly measuring the radiative and total heat fluxes on the back-shell, the convective portion of the heat flux will be estimated. The status of the current tools to perform the post-flight data analysis will be presented, along with plans for model improvements.

Hwang, Helen H.↗

NASA NTRS2017

Mars 2020 Entry, Descent, and Landing Instrumentation 2 (MEDLI2) Sensor Suite

Hwang, Helen↗

NASA NTRS2017

Heatshield for Extreme Entry Environment Technology (HEEET) Enabling Missions Beyond Heritage Carbon Phenolic

Future NASA robotic missions utilizing an entry system into Venus and the outer planets, results in extremely high entry conditions that exceed the capabilities of state of the art low to mid density ablators such as PICA or AVCOAT. Previously, mission planners had to assume the use of fully dense carbon phenolic heatshields similar to what was flown on Pioneer Venus or Galileo. Carbon phenolic is a robust TPS material, however, its high density and relatively high thermal conductivity constrain mission planners to steep entries, with high heat fluxes and pressures and short entry durations. The high entry conditions pose challenges for certification in existing ground based test facilities and the longer-term sustainability of CP will continue to pose challenges. NASA has decided to invest in new technology development rather than invest in reviving carbon phenolic. The HEEET project, funded by STMD is maturing a game changing Woven Thermal Protection System technology. HEEET is a capability development project and is not tied to a single mission or destination, therefore, it is challenging to complete ground testing needed to demonstrate a capability that is much broader than any single mission or destination would require. This presentation will status HEEET progress. Near term infusion target for HEEET is the upcoming New Frontiers (NF-4) class of competitively selected Science Mission Directorate (SMD) missions for which it is incentivized.

thermal protection↗

NASA NTRS2017

Modeling the Relationship Between Porosity and Permeability During Oxidation of Ablative Materials

The ablative materials used in thermal protection systems (TPS) undergo oxidation during atmospheric entry which leads to an in-depth change in both permeability and porosity. These properties have a significant affect on heat transfer in a TPS during entry. X-ray micro-tomography has provided 3D images capturing the micro-structure of TPS materials. In this study, we use micro-tomography based simulations to create high-fidelity models relating permeability to porosity during oxidation of FiberForm, the carbon fiber preform of the Phenolic Impregnated Carbon Ablator (PICA) often used as a TPS material. The goal of this study is to inform full-scale models and reduce uncertainty in TPS modeling.

Ablativ↗

NASA NTRS2017

Vacuum Infusion Process Development for Conformal Ablative Thermal Protection System Materials

Conformal ablators are low density composite materials comprised of a flexible carbon felt based fibrous substrate and a high surface area phenolic matrix. These materials are fabricated to near net shape by molding the substrate, placing in a rigid matched mold and infusing with liquid resin through a vacuum assisted process. The open mold process, originally developed for older rigid substrate ablators, such as PICA, wastes a substantial amount of resin. In this work, a vacuum infusion process a type of liquid composite molding where resin is directly injected into a closed mold under vacuum is advanced for conformal ablators. The process reduces waste over the state-of-the-art technique. Small, flat samples of Conformal Phenolic Impregnated Carbon Ablator are infused using the new approach and subjected to a range of curing configurations and conditions. Resulting materials are inspected for quality and compared to material produced using the standard process. Lessons learned inform subsequent plans for process scale up.

TPS↗

NASA NTRS2018

Aerothermodynamics for Dragonfly's Titan Entry

Dragonfly is a proposed spacecraft and mission that would send a mobile robotic rotorcraft lander to Titan, the largest moon of Saturn, in order to study prebiotic chemistry and extraterrestrial habitability at various locations. Titan is unique in having an abundant, complex, and diverse carbon-rich chemistry on the surface of a water-ice-dominated world with an interior water ocean, making it a high priority target for astrobiology and origin of life studies. The mission was initially proposed in April 2017 to NASA's New Frontiers program by the Johns Hopkins Applied Physics Laboratory. In December 2017, it was selected as one of two finalists (out of twelve proposals) to further refine the mission's concept. NASA Ames Research Center and NASA Langley Research Center are partnering as the leads for the Dragonfly's entry system to provide the completed EDL Assembly. The aerothermal analysis for Dragonfly utilizes four simulation tools from NASA Ames Research Center. Traj for calculating the trajectory, DPLR 4.04.0 for calculating the flowfield around the vehicle and convective heating, NEQAIR V15.0 for calculating the radiative heating, and FIAT for calculating the material response and thermal protection system (TPS) sizing for the heatshield. The entry conditions are relatively benign and can readily be accommodated with a tiled PICA heatshield similar to MSL and a number of flight proven materials for the backshell. This work will demonstrate that the aerothermal entry environments can be readily solved using heritage materials and techniques.

New Frontiers↗

NASA NTRS2018

Sizing and Margins Assessment of the Mars2020 Acusil-II Thermal Protection System

The Mars2020 entry vehicle is currently being developed by NASA to safely land its next rover on the Martian surface in 2021. The vehicle will be protected from entry aeroheating using three different TPS materials: PICA tiles on the forebody, SLA-561V on the backshell and Acusil-II on the parachute close-out cone (PCC) and its backshell interface plate (BIP). Mars2020's entry vehicle and TPS design is identical to the Mars Science Laboratory, NASA's last Mars lander; therefore, the purpose of this study is to assess the adequacy of the existing TPS design and thickness for Mars2020 predicted environments. This study focuses on sizing and margin assessment of Acusil-II TPS on the PCC and BIP. The methodology and analysis techniques that were used for assessing thermal margins are reviewed. Analysis assumptions and limitations are discussed in detail. Thermal sizing is performed at different locations and results are presented.

Acusil↗

NASA NTRS2018

Full-Scale MSL Heatshield Material Response Using DSMC and CFD to Compute the Aerothermal Environments

During Mars atmospheric entry, the Mars Science Laboratory (MSL) was protected by a 4.5 meters diameter ablative heatshield assembled in 113 tiles [1]. The heatshield was made of NASA's flagship ablative material, the Phenolic Impregnated Carbon Ablator (PICA) [2]. Prior work [3] compared the traditional one-dimensional and three-dimensional material response models at different locations in the heatshield. It was observed that the flow was basically one-dimensional in the nose and flank regions, but three-dimensional flow effects were observed in the outer flank. Additionally, the effects of tiled versus monolithic heatshield models were also investigated. It was observed that the 3D tiled and 3D monolithic configurations yielded relative differences for in-depth material temperature up to 18% and 28%, respectively, when compared to the a 1D model.

Aerothermal↗

NASA NTRS2018

Significance of DSMC Computed Aerothermal Environments in the Rarefied Regime for Atmospheric Entry Material Response

During Mars atmospheric entry, the Mars Science Laboratory (MSL) was protected by a 4.5 meters diameter ablative heatshield assembled in 113 tiles. The heatshield was made of NASA's flagship ablative material, the Phenolic Impregnated Carbon Ablator (PICA). Prior work compared the traditional one-dimensional and three-dimensional material response models at different locations in the heatshield. It was observed that the flow was basically one-dimensional in the nose and flank regions, but three-dimensional flow effects were observed in the outer flank. The objective of this work is to study the effects of the aerothermal environment on the material response. We extend prior work by computing aerothermal environments using the direct simulation Monte Carlo (DSMC) code SPARTA and the CFD code Data Parallel Line Relaxation (DPLR). SPARTA is used to compute environment in the rarefied regime prior to 48.4s of entry where the Knudsen number is such that the Navier-Stokes equations can be inaccurate. Similarly to previous work, the DPLR software is used to compute the hypersonic environment for laminar then turbulent boundary layer assumptions from 48.4 s up to 100 s after Entry Interface (EI) along the MSL 08-TPS-02/01a trajectory. We observe that extending the aerothermal environments to times prior to 48.4 s modifies the thermal response of the heat shield at the surface and in-depth; however the effects on the recession are minimal. Additionally, using the assumption of a turbulent boundary layer versus a laminar one leads to higher surface and in-depth temperatures, larger recession, and a displacement of the peak heating and peak recession location.

Aerothermal↗

NASA NTRS2014

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

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

thermal protection system↗

NASA NTRS2014

Woven Thermal Protection System (WTPS) - a Novel Approach to Meet NASA's Most Demanding Missions

NASAs future robotic missions utilizing an entry system into Venus and the outer planets, namely, Saturn, Uranus, Neptune, result in extremely high entry conditions that exceed the capabilities of state of the art low to mid density ablators such as PICA or Avcoat. Therefore mission planners typically assume the use of a fully dense carbon phenolic heat shield similar to what was flown on Pioneer Venus and Galileo. Carbon phenolic is a robust TPS material however its high density and relatively high thermal conductivity constrain mission planners to steep entries, with high heat fluxes and pressures and short entry durations, in order for CP to be feasible from a mass perspective. The high entry conditions pose challenges for certification in existing ground based test facilities and the longer-term sustainability of CP will continue to pose challenges. In 2012 the Game Changing Development Program (GCDP) in NASAs Space Technology Mission Directorate funded NASA ARC to investigate the feasibility of a Woven Thermal Protection System (WTPS) to meet the needs of NASAs most challenging entry missions. This project was highly successful demonstrating that a Woven TPS solution compares favorably to CP in performance in simulated reentry environments and provides the opportunity to manufacture graded materials that should result in overall reduced mass solutions and enable a much broader set of missions than does CP. Building off the success of the WTPS project GCDP has funded a follow on project to further mature and scale up the WTPS concept for insertion into future NASA robotic missions. The matured WTPS will address the CP concerns associated with ground based test limitations and sustainability. This presentation will discuss results from the WTPS heat-shield for extreme entry environment technology (HEEET) project

woven TPS↗

NASA NTRS2019

The Dragonfly Entry and Descent System

Dragonfly is a proposed New Frontiers class mission that will send a nuclear powered octocopter to the surface of Titan for an extended science mission. This presentation will provide an overview of the Entry and Descent system that is under development to ensure the save delivery of this unique "relocatable lander" to Titan. Titan's dense atmosphere, large atmospheric scale height, and low gravity allows for a slow-paced entry and descent sequence that lasts more than 100 minutes, as opposed to the "7 minutes of terror" that is charac-teristic of landed Mars missions. This slow pace al-lows for sufficient temporal separation between critical events of the EDL sequence to minimize overall risk.The Dragonfly entry and descent system is composed of high-heritage components, minimizing overall risk. The aeroshell will be a scaled Genesis Sample Return capsule with a diameter of 3.75 meters, built by Lock-heed Martin. The thermal protection system (TPS) is made up of Phenolic Impregnated Carbon Ablator-Domestic (PICA-D) on the heatshield, SLA-561V on the backshell, and SLA-220M on the aft cover and low gain antenna. Each material has extensive heritage for the chosen application. The spacecraft will enter Titan at a velocity of 7.3 km/s, resulting in a predicted fully margined stagnation point heating environment of 254 W/cm2 heat rate and 13 kJ/cm2 heat load, well within the tested limits of the chosen materials. The aeroheat-ing environments, including the significant contribu-tion of shock layer radiation from CN on both the heatshield and backshell, are evaluated using state of the art models and codes that have been validated with appropriate ground testing.Once the deceleration pulse is complete, a disk-gap-band (DGB) drogue parachute will be deployed at ap-proximately Mach 1.5 to stabilize and further deceler-ate the spacecraft. Due to the dense atmosphere, the spacecraft will spend more than 80 minutes on this parachute, until reaching an appropriate altitude to de-ploy the subsonic main parachute. The lander is re-leased after approximately 17 minutes on the main chute before releasing and transitioning to powered flight in order to navigate to its first landing site. The release of the lander from the backshell effectively ends the entry and descent portion of the mission.The full presentation will provide additional details about the design of the EDL system hardware, engi-neering design, and overall con-ops. Preliminary aero-thermal and TPS sizing analyses will be presented, and the parachute system will be described in greater detail. In addition, the Dragonfly spacecraft will carry an En-gineering Science Investigation (ESI) package designed to obtain engineering data during EDL that will be used to validate the design methodology for future missions. An overview of the proposed ESI package will also be presented.

EDL↗