Effect of Phenolic Matrix Microcracking on the Structural Response of a 3-D Woven Thermal Protection System

The effect of microcracking in the phenolic matrix of a three-dimensional woven thermal protection system (TPS) and the resulting material stiffness reduction was studied via a comparison of finite element analysis results from a linear analysis and an iterative linear analysis. A TPS is necessary to protect space vehicles from the aerodynamic heating of planetary entry. The Heatshield for Extreme Entry Environment Technology (HEEET) project has developed a TPS for use in high heat-flux and pressure missions. The material is a dual-layer continuous dry weave, which is then infiltrated with a low-density phenolic resin matrix to form a rigid ablator. The phenolic resin matrix does not have structural load transfer requirements, and testing has shown that the phenolic resin can fully satisfy thermal requirements when the matrix contains microcracks. Due to high stresses in the through-the-thickness direction of the material, phenolic microcracks may form in the matrix material, which would result in a reduction of stiffness. An exploratory study was conducted to determine if reduction in material stiffness would change the load paths and/or decrease the structural margins. A comparison was performed between a linear finite element analysis that did not take into account phenolic microcracking and an iterative linear finite element analysis that accounted for propagation of phenolic microcracking. Four subcases using varying assumptions were analyzed and the results indicate that the assumed strength at which the phenolic microcracking propagates was the critical parameter for determining the extent of microcracking in the phenolic matrix. Phenolic microcracking does not have an adverse effect on the structural response of the test article and is not a critical failure

Space Vehicle Heat Shield Having Edgewise Strips of Ablative Material

A heat shield for a space vehicle comprises a plurality of phenolic impregnated carbon ablator (PICA) blocks secured to a surface of the space vehicle and arranged in a pattern with gaps therebetween. The heat shield further comprises a plurality of PICA strips disposed in the gaps between the PICA blocks. The PICA strips are mounted edgewise, such that the structural orientation of the PICA strips is substantially perpendicular to the structural orientation of the PICA blocks

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

System Level Aerothermal Testing for the Adaptive Deployable Entry and Placement Technology (ADEPT)

The Adaptive Deployable Entry and Placement Technology (ADEPT), a mechanically deployable entry vehicle technology, has been under development at NASA since 2011. As part of the technical maturation of ADEPT, designs capable of delivering small payloads (10 kg) are being considered to rapidly mature sub 1 m deployed diameter designs. The unique capability of ADEPT for small payloads comes from its ability to stow within a slender volume and deploy to achieve a mass efficient drag surface with a high heat rate capability. The low ballistic coefficient results in entry heating and mechanical loads that can be met by a revolutionary three-dimensionally woven carbon fabric supported by a deployable skeleton structure. This carbon fabric has test proven capability as both primary structure and payload thermal protection system. In order to rapidly advance ADEPTs technical maturation, the project is developing test methods that enable thermostructural design requirement verification of ADEPT designs at the system level using ground test facilities. Results from these tests are also relevant to larger class missions and help us define areas of focused component level testing in order to mature material and thermal response design codes. The ability to ground test sub 1 m diameter ADEPT configurations at or near full-scale provides significant value to the rapid maturation of this class of deployable entry vehicles. This paper will summarize arc jet test results, highlight design challenges, provide a summary of lessons learned and discuss future test approaches based upon this methodology

Overview of Heatshield for Extreme Entry Environment Technology (HEEET) Project

The objective of the Heatshield for Extreme Entry Environment Technology (HEEET) projects is to mature a 3-D Woven Thermal Protection System (TPS) to Technical Readiness Level (TRL) 6 to support future NASA missions to destinations such as Venus and Saturn. Destinations that have extreme entry environments with heat fluxes up to 5000 watts per square centimeter and pressures up to 5 atmospheres, entry environments that NASA has not flown since Pioneer-Venus and Galileo. The scope of the project is broad and can be split into roughly four areas, Manufacturing/Integration, Structural Testing and Analysis, Thermal Testing and Analysis and Documentation. Manufactruing/Integration covers from raw materials, piece part fabrication to final integration on a 1-meter base diameter 45-degree sphere cone Engineering Test Unit (ETU). A key aspect of the project was to transfer as much of the manufacturing technology to industry in preparation to support future mission infusion. The forming, infusion and machining approaches were transferred to Fiber Materials Inc. and FMI then fabricated the piece parts from which the ETU was manufactured. The base 3D-woven material consists of a dual layer weave with a high density outer layer to manage recession in the system and a lower density, lower thermal conductivity inner layer to manage the heat load. At the start of the project it was understood that due to weaving limitations the heat shield was going to be manufactured from a series of tiles. And it was recognized that the development of a seam solution that met the structural and thermal requirements of the system was going to be the most challenging aspect of the project. It was also recognized that the seam design would drive the final integration approach and therefore the integration of the ETU was kept in-house within NASA. A final seam concept has been successfully developed and implemented on the ETU and will be discussed. The structural testing and analysis covers from characterization of the different layers of the infused material as functions of weave direction and temperature, to sub-component level testing such as 4-pt bend testing at sub-ambient and elevated temperature. ETU test results are used to validate the structural models developed using the element and sub-component level tests. Given the seam has to perform both structurally and aerothermally during entry a novel 4-pt bend test fixture was developed allowing articles to be tested while the front surface is heated with a laser. These tests are intended to establish the system's structural capability during entry. A broad range of aerothermal tests (arcjet tests) are being performed to develop material response models for predicting the required TPS thickness to meet a mission's needs and to evaluate failure modes. These tests establish the capability of the system and assure robustness of the system during entry. The final aspect of the project is to develop a comprehensive Design and Data Book such that a future mission will have the information necessary to adopt the technology. This presentation will provide an overview and status of the project and describe the status of the tehnology maturation level for the inner and outer planet as well as earth entry sample return missions

L.: Improving Metallic Thermal Protection System Hypervelocity Impact Resistance Through Design of Experiments Approach

A design of experiments approach has been implemented using computational hypervelocity impact simulations to determine the most effective place to add mass to an existing metallic Thermal Protection System (TPS) to improve hypervelocity impact protection. Simulations were performed using axisymmetric models in CTH, a shock-physics code developed by Sandia National Laboratories, and validated by comparison with existing test data. The axisymmetric models were then used in a statistical sensitivity analysis to determine the influence of five design parameters on degree of hypervelocity particle dispersion. Several damage metrics were identified and evaluated. Damage metrics related to the extent of substructure damage were seen to produce misleading results, however damage metrics related to the degree of dispersion of the hypervelocity particle produced results that corresponded to physical intuition. Based on analysis of variance results it was concluded that the most effective way to increase hypervelocity impact resistance is to increase the thickness of the outer foil layer. Increasing the spacing between the outer surface and the substructure is also very effective at increasing dispersion

Improving Metallic Thermal Protection System Hypervelocity Impact Resistance Through Design of Experiments Approach

A design of experiments approach has been implemented using computational hypervelocity impact simulations to determine the most effective place to add mass to an existing metallic Thermal Protection System (TPS) to improve hypervelocity impact protection. Simulations were performed using axisymmetric models in CTH, a shock-physics code developed by Sandia National Laboratories, and validated by comparison with existing test data. The axisymmetric models were then used in a statistical sensitivity analysis to determine the influence of five design parameters on degree of hypervelocity particle dispersion. Several damage metrics were identified and evaluated. Damage metrics related to the extent of substructure damage were seen to produce misleading results, however damage metrics related to the degree of dispersion of the hypervelocity particle produced results that corresponded to physical intuition. Based on analysis of variance results it was concluded that the most effective way to increase hypervelocity impact resistance is to increase the thickness of the outer foil layer. Increasing the spacing between the outer surface and the substructure is also very effective at increasing dispersion

Preliminary Thermal-Mechanical Sizing of Metallic TPS: Process Development and Sensitivity Studies

The purpose of this research was to perform sensitivity studies and develop a process to perform thermal and structural analysis and sizing of the latest Metallic Thermal Protection System (TPS) developed at NASA LaRC (Langley Research Center). Metallic TPS is a key technology for reducing the cost of reusable launch vehicles (RLV), offering the combination of increased durability and competitive weights when compared to other systems. Accurate sizing of metallic TPS requires combined thermal and structural analysis. Initial sensitivity studies were conducted using transient one-dimensional finite element thermal analysis to determine the influence of various TPS and analysis parameters on TPS weight. The thermal analysis model was then used in combination with static deflection and failure mode analysis of the sandwich panel outer surface of the TPS to obtain minimum weight TPS configurations at three vehicle stations on the windward centerline of a representative RLV. The coupled nature of the analysis requires an iterative analysis process, which will be described herein. Findings from the sensitivity analysis are reported, along with TPS designs at the three RLV vehicle stations considered

Hypervelocity Impact Test Results for a Metallic Thermal Protection System

Hypervelocity impact tests have been performed on specimens representing metallic thermal protection systems (TPS) developed at NASA Langley Research Center for use on next-generation reusable launch vehicles (RLV). The majority of the specimens tested consists of a foil gauge exterior honeycomb panel, composed of either Inconel 617 or Ti-6Al-4V, backed with 2.0 in. of fibrous insulation and a final Ti-6Al-4V foil layer. Other tested specimens include titanium multi-wall sandwich coupons as well as TPS using a second honeycomb sandwich in place of the foil backing. Hypervelocity impact tests were performed at the NASA Marshall Space Flight Center Orbital Debris Simulation Facility. An improved test fixture was designed and fabricated to hold specimens firmly in place during impact. Projectile diameter, honeycomb sandwich material, honeycomb sandwich facesheet thickness, and honeycomb core cell size were examined to determine the influence of TPS configuration on the level of protection provided to the substructure (crew, cabin, fuel tank, etc.) against micrometeoroid or orbit debris impacts. Pictures and descriptions of the damage to each specimen are included