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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1980

Photoelastic tests on models of thermal protection system for space shuttle orbiter

The thermal protection system (TPS) of the space shuttle orbiter vehicle, consisting of ceramic tile/adhesive/strain isolator pad/adhesive/aluminum substructure, was modeled photoelasticity. A highly sensitive photoelastic material was used in the models to show the nature of the stress-transfer between the strain isolation pad (SIP) and the ceramic tile through the RTV-adhesive layer. Isochromatic fringe patterns were obtained for models subjected to tension and combined tension and bending. Tests indicated that the load-transfer between the SIP and the photoelastic material occurred at discrete locations causing stress concentrations in the photoelastic material. Stress concentration factors of the order of 1.9 were measured, but as the observed photoelastic response was an integrated effect through the model thickness, the local stress concentration factors at the SIP/tile interface could be even higher.

Prabhakaran, R.↗
NASA NTRS1981

Evaluation of proof testing as a means of assuring mission success for the Space Shuttle thermal protection system

The reliability analysis of a thermal protection system (TPS) incorporating low density, fibrous ceramic materials must take into account the time-dependency and variability of the system's strength. Fracture mechanics concepts can be used to estimate the allowable strength and expected lifetime, as well as to define a proof test scheme for assuring mission success of TPS. The aim of this study was to evaluate the proof testing scheme used on the TPS of the Space Shuttle with these fracture mechanics concepts. The analysis confirmed that proof testing was necessary for the undensified system and that proof testing should lead to adequate reliability with respect to the design stresses. For the undensified system, fracture mechanics predictions were confirmed by measuring the strength of samples that survived proof testing. It was also found that the time-dependent nature of the system's strength was controlled mainly by the ceramic and that this time-strength dependency is similar to that observed for bulk glasses that have compositions similar to the glass fibers present in the ceramic tiles.

Green, D. J.↗
NASA NTRS1983

Shell-Tile Thermal-Protection System

Durable shell-tile thermal-protection system consists of interlocking upper and lower hard caps, incorporating appropriate stiffeners and enclosing lightweight fibrous insulation. New shell tile more durable than reusable surface insulation (RSI) currently used on Space Shuttle orbiter.

Macconochie, I. O.↗
NASA NTRS1989

Thermal Protection System of the Space Shuttle

The Thermal Protection System (TPS), introduced by NASA, continues to incorporate many of the advances in materials over the past two decades. A comprehensive, single-volume summary of the TPS, including system design rationales, key design features, and broad descriptions of the subsystems of TPS (E.g., reusable surface insulation, leading edge structural, and penetration subsystems) is provided. Details of all elements of TPS development and application are covered (materials properties, manufacturing, modeling, testing, installation, and inspection). Disclosures and inventions are listed and potential commercial application of TPS-related technology is discussed.

Cleland, John↗
NASA NTRS2024

Bringing the Heat: Thermal Protection Systems for Low Earth Orbit Transportation and Lunar Exploration

Thermal protection systems (TPS) protect spacecraft from the heating associated with hypersonic flight through a planet’s atmosphere. The materials are subjected to extreme heating and the overall system must be engineered to reliably protect the spacecraft structure during atmospheric entry. The talk will introduce TPS and describe three NASA-led efforts to develop TPS materials and manufacturing processes. The first effort studied thermally stable ceramic aerogels for use in thermal barriers and high temperature seals. A design framework for such materials is proposed, linking the thermal stability performance data with existing material property measurements. The second project, Additive Manufacturing of Thermal Protection Systems (AMTPS), developed an automated, additive approach for ablative heat shield manufacturing. Novel material systems and manufacturing methods were developed to reduce the time required to manufacture ablative TPS heat shields. The third effort tackles a critical challenge of material obsolescence for silica fiber based reusable TPS. The properties and performance of heritage silica fiber are quantified and compared to modern alternatives to support qualification for flight. The talk will also describe other work in TPS conducted by the Thermal Design Branch at NASA JSC and opportunities for students and faculty to get involved with NASA.

Nathaniel Olson↗
NASA NTRS2005

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗
NASA NTRS2006

An Approximate Ablative Thermal Protection System Sizing Tool for Entry System Design

A computer tool to perform entry vehicle ablative thermal protection systems sizing has been developed. Two options for calculating the thermal response are incorporated into the tool. One, an industry-standard, high-fidelity ablation and thermal response program was integrated into the tool, making use of simulated trajectory data to calculate its boundary conditions at the ablating surface. Second, an approximate method that uses heat of ablation data to estimate heat shield recession during entry has been coupled to a one-dimensional finite-difference calculation that calculates the in-depth thermal response. The in-depth solution accounts for material decomposition, but does not account for pyrolysis gas energy absorption through the material. Engineering correlations are used to estimate stagnation point convective and radiative heating as a function of time. The sizing tool calculates recovery enthalpy, wall enthalpy, surface pressure, and heat transfer coefficient. Verification of this tool is performed by comparison to past thermal protection system sizings for the Mars Pathfinder and Stardust entry systems and calculations are performed for an Apollo capsule entering the atmosphere at lunar and Mars return speeds.

Dec, John A.↗
NASA NTRSDate not supplied

3D Woven Mid-Density Carbon Phenolic (3MDCP) Thermal Protection System Development

3-Dimensionally Woven, Mid-Density, Carbon Phenolic (3MDCP) Thermal Protection System (TPS) material is derived from the dual layer 3D woven Heatshield for Extreme Entry Environment Technology (HEEET) material. The baseline 3MDCP design is a single piece thermal protection system that avoids the manufacturing and certification challenges associated with a tiled configuration. 3MDCP is targeted for very aggressive entry environments such as high-speed sample return missions to Earth and missions to Saturn, Venus and the ice giants. A 3MDCP heatshield begins as a flat woven preform that is formed to a given heatshield shape and then infused with phenolic resin. NASA Ames, in collaboration with TEAM Inc. (weaving) and Fiber Materials Inc. (both Spirit AeroSystems Companies) have been developing and demonstrating the manufacturing processes to fabricate a 3MDCP heatshield at a diameter of 1.25 meters. The process of forming the flat woven preform into the final heatshield shape, a sphere-cone geometry, involves local movement of the yarns in the weave. This results in a single piece heatshield with continuous fibers, albeit with property variations between different regions on the heatshield. This presentation will provide a high-level status of 3MDCP development. This will include an overview of the manufacturing processes, with an emphasis on the impact of forming on fiber orientation, material properties and performance. The presentation will layout the plan for testing to assess the impact of forming on properties and review preliminary data comparing properties of flat to formed materials.

Thermal Protection System↗
NASA NTRSDate not supplied

Numerical Implementation of Source Terms to Evaluate Active Porosity Control using a Transpiration Cooled Thermal Protection System

Transpiration cooling has renewed interest of study as a renewable system of thermal protection for atmospheric entry systems. The presence of coolant within the void space of a porous material changes its effective porosity and theorized to lend the external heatshield as a potential means of porous surface control. Effective porosity is explored in two forms: saturation of void space where exiting coolant creates ‘blowing’ over the surface, and partial evacuation of the void space where withdrawal of the coolant creates ‘suction’ on the surface. Thus this study evaluates using transpiration cooling as a as means of active aerodynamic control. The feasibility of the system is evaluated numerically using Reynolds-averaged Navier Stokes based computational fluid dynamics for the aerodynamic assessments by coding source terms of transpiration cooling products at a heat shield surface boundary at the wind side shoulder, creating an asymmetric scheme. Study results show induced moments about the pitch axis incurred following the source terms of a cooling fluid at a near-shoulder location on the heat shield. Lift and drag saw increasing modulation with increase in mass flux rate. Pitch moment was altered a total of 388 N-m between lowest and highest ‘suction’ rates and changed 194 N-m maximum between ‘blowing’ rates.

transpiration cooling↗