Robin Beck

Sizing and Margins Assessment of Mars Science Laboratory Aeroshell Thermal Protection System

The methodology employed for the thermal design and margins assessment of the Mars Science Laboratory aeroshell thermal protection system is reviewed. A new thermal margins policy was developed in the course of this work that provides additional rigor over previous methods. Because of a late change of thermal protection materials from the heritage super lightweight ablator 561V to phenolic impregnated carbon ablator, the design of the heat shield followed a nontraditional path in which the flight thickness was selected based on a mass (rather than thermal) limit. The material switch was followed by detailed thermal analyses that demonstrated that the baselined thickness was sufficient to provide adequate thermal protection to the vehicle without violating design requirements during a 3-sigma worst-case entry condition. The backshell material thickness was also finalized before the thermal sizing was completed, and the resulting analysis showed that there was more than sufficient material on the backshell....

Development of the Mars Science Laboratory Heatshield Thermal Protection System

Early in the development of the Mars Science Laboratory thermal protection system on the heatshield, project management planned to use Lockheed Martin’s Super Light Ablator in honeycomb as the ablative material based on successful use on previous Mars entry heatshields and on stagnation arcjet tests at heating rates beyond the design levels. Because this heatshield would be the first to experience combined turbulent flow and high shear environments as it entered the Mars atmosphere, tests were performed in various arcjet facilities on flat-plate, wedge, and swept-cylinder specimen configurations in order to ascertain the effects of shear on the material. During the course of these tests, a set of conditions within the flight envelope was identified that resulted in catastrophic failure in the SLA-561V. Consequently, project management decided to replace the SLA-561V with the phenolic-impregnated carbon ablator, the material that had flown successfully on the Stardust mission and was undergoing intense tes...

Development of Thermal Protection Materials for Future Mars Entry, Descent and Landing Systems

Entry Systems will play a crucial role as NASA develops the technologies required for Human Mars Exploration. The Exploration Technology Development Program Office established the Entry, Descent and Landing (EDL) Technology Development Project to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. An assessment of current entry system technologies identified significant opportunity to improve the current state of the art in thermal protection materials in order to enable landing of heavy mass (40 mT) payloads. To accomplish this goal, the EDL Project has outlined a framework to define, develop and model the thermal protection system material concepts required to allow for the human exploration of Mars via aerocapture followed by entry. Two primary classes of ablative materials are being developed: rigid and flexible. The rigid ablatives will be applied to the acreage of a 10x30 m rigid mid L/D Aeroshell to endure the dual pulse heating (peak approx.500 W/sq cm). Likewise, flexible ablative materials are being developed for 20-30 m diameter deployable aerodynamic decelerator entry systems that could endure dual pulse heating (peak aprrox.120 W/sq cm). A technology Roadmap is presented that will be used for facilitating the maturation of both the rigid and flexible ablative materials through application of decision metrics (requirements, key performance parameters, TRL definitions, and evaluation criteria) used to assess and advance the various candidate TPS material technologies.

CFD and Material Response Framework for Wedge Testing in AEDC H2

A new analysis framework for wedge testing in the Arnold Engineering Development Center (AEDC) H2 arc-jet has been developed. The technique uses the aerothermal nonequilibrium DPLR flow solver, specialized routines, and the FIAT material response code. The framework takes advantage of null-point sweep profiles from testing, and when coupled with a new automated process using flow solver has been used to analyze the recession of ablative TPS samples. Results are shown for the recent Mars Science Laboratory (MSL) Phenolic Impregnated Carbon Ablator (PICA) shear testing at AEDC, demonstrating the good agreement between aerothermal and material response simulations and the experimental data.

Development challenges of game-changing entry system technologies from concept to mission infusion

NASAs Space Technology Mission Directorate (STMD) and the Game Changing Development Program (GCDP) were created to develop new technologies. This paper describes four entry system technologies that are funded by the GCDP and summarizes the lessons learned during the development. The investments are already beginning to show success, mission infusion pathways after five years of existence. It is hoped that our experience and observations, drawn from projects supported by the GCD program/STMD, Orion and SMD can help current and future technology development projects. Observations on fostering a culture of success and on constraints that limit greater success are also provided.

Thermal Modeling of In-Depth Thermocouple Response in Ablative Heat Shield Materials

The accurate measurement of in-depth temperatures of thermal protection system (TPS) materials, whether in ground-based tests or in flight, is imperative because it allows for validation of mathematical models that are used to predict the aerothermal environment and resulting TPS response. Existing TPS material response codes of ablative heatshield materials do not account for the presence of thermocouple wire and assume one dimensional pyrolysis gas flow. The objective of the work presented in this paper is to quantify the relative error associated with the thermal lag between the thermocouple wire and the surrounding material for the low-density ablator SLA-561V. A two-dimensional heat conduction model with time-dependent wall temperature and surface recession boundary conditions is presented. The thermocouple is modeled as a circular cross-section in perfect thermal contact with the SLA-561V. Trade studies on wire diameter and thermocouple configuration have been conducted with the two-dimensional model and results are given for two heating profiles once considered for sizing of the Mars Science Laboratory heatshield. The simulations show that the temperature measured by a near-surface thermocouple can differ from that of the thermal protection system material by up to 10% for the case of a 0.305 mm diameter bare wire and a heating profile with a peak value of 217 W/cm.

Vacuum Infusion Process Development for Conformal Ablative Thermal Protection System Materials

Conformal ablators are low density composite materials comprised of a flexible fibrous substrate and polymer matrix. These materials are fabricated to near net shape by placing the substrate in a rigid, matched mold and infusing with liquid resin in an open, vacuum– assisted immersion process. This 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 and may eliminate the need for an atmosphere-controlled oven. 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. Density, resin mass fraction and char yield are measured. Lessons learned inform subsequent plans for process scale up.

Overview of Entry Descent and landing investments in the NASA Exploration Technology Development Program

Landing humans on Mars is perhaps the largest Entry, Descent and Landing challenge for this (or any) generation of NASA engineers.12 Essentially all current Mars EDL technology is based on development efforts in support of the Viking program in the 1960s and 1970s. Since that time, these key technologies have been tweaked and some additions have been added in order to extend the landed mass and altitude capabilities, but it has become clear that this technology suite is fundamentally limited to a landed mass of ∼1.25 metric tons. In contrast, current system studies for human Mars missions indicate a 40 t required landed mass. Clearly, a revolutionary change to the current state of the art is required in order to enable such ambitious missions. The present paper briefly discusses some of the recent NASA systems analysis work that has defined candidate technologies as well as overall system requirements for human landings. Those requirements were used to define a new technology development project within the Exploration Systems Mission Directorate in 2010. This Entry Descent and Landing Technology Development Project is currently investing in three major technology areas: rigid and flexible thermal protection systems, supersonic retro-propulsion systems, and simulation models & tools. The project portfolio in each of these areas is briefly discussed and some current results obtained in the first year of execution are presented. The project will continue in 2011, and has developed technology development roadmaps that would support the maturation of all of their technology investments to a Technology Readiness Level of 5 or 6 by 2016 in order to support infusion of the technologies onto either science missions or precursor exploration missions to Mars, potentially as early as 2018 or 2020.

Thermal Protection System Aerothermal Screening Tests in the HYMETS Facility

The Entry, Descent, and Landing (EDL) Technology Development Project has been tasked to develop Thermal Protection System (TPS) materials for insertion into future Mars Entry Systems. A screening arc jet test of seven rigid ablative TPS material candidates was performed in the Hypersonic Materials Environmental Test System (HYMETS) facility at NASA Langley Research Center, in both an air and carbon dioxide test environment. Recession, mass loss, surface temperature, and backface thermal response were measured for each test specimen. All material candidates survived the Mars aerocapture relevant heating condition, and some materials showed a clear increase in recession rate in the carbon dioxide test environment. These test results supported subsequent down-selection of the most promising material candidates for further development.

Deployable Materials for Spacecraft Thermal Protection: LHMEL Tests

A set of candidate deployable TPS materials, including flexible ablator materials, were screen-tested at the Laser Hardened Materials Evaluation Laboratory. Both the 10.6 micron CO2 LHMEL I laser and the 1.07 micron Fiber laser were used, operating in the nonpulsed mode. The heat flux levels and durations represented the predicted heating for deployable vehicles for future NASA Mars missions, including dual pulse scenarios. Test measurements included surface brightness and bondline temperatures, mass loss and char depths. The test materials included a range of materials, from woven carbon cloth, as well as carbon, ceramic and organic fiber felts, incorporating phenolic and silicone as additional pyrolyzing components. Performance of these materials is compared to that of SIRCA, which was selected as a baseline standard material. As anticipated, the semi-transparent glassy or polymer materials showed a different response to the Fiber laser tests than the CO2 laser irradiation. The dual pulse behavior of charred material showed stable char and somewhat higher thermal transport, compared to uncharred or virgin material. As expected, every pre-charred material survived the second heating cycle without spallation. The thermal properties of the novel TPS materials had not been characterized at the time of testing, and the test samples included an array of different areal weights, densities and thicknesses, with different chemical compositions and pyrolyzing material loadings.