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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 55 records · Page 3
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

High purity silica reflecting heat shield development

A reflecting heat shield composed of fused silica in which the scattering results from the refractive index mismatch between silica particles and the voids introduced during the fabrication process is developed. Major considerations and conclusions of the development are: the best material to use is Type A, which is capable of ultra-high-purity and which does not show the 0.243 micrometer absorption band; the reflection efficiency of fused silica is decreased at higher temperatures due to the bathochromic shift of the ultraviolet cut-off; for a given silica material, over the wavelength region and particle sizes tested, the monodisperse particle size configurations produce higher reflectances than continuous particle size configurations; and the smaller monodisperse particle size configurations give higher reflectance than the larger ones. A reflecting silica configuration that is an efficient reflector of shock layer radiation at high ablation temperatures is achieved by tailoring the matrix for optimum scattering and using an ultra-high-purity material.

Congdon, W.↗
NASA NTRS2009

Retro Rocket Motor Self-Penetrating Scheme for Heat Shield Exhaust Ports

A preliminary scheme was developed for base-mounted solid-propellant retro rocket motors to self-penetrate the Orion Crew Module heat shield for configurations with the heat shield retained during landings on Earth. In this system the motors propel impactors into structural push plates, which in turn push through the heat shield ablator material. The push plates are sized such that the remaining port in the ablator material is large enough to provide adequate flow area for the motor exhaust plume. The push plate thickness is sized to assure structural integrity behind the ablative thermal protection material. The concept feasibility was demonstrated and the performance was characterized using a gas gun to launch representative impactors into heat shield targets with push plates. The tests were conducted using targets equipped with Fiberform(R) and PICA as the heat shield ablator material layer. The PICA penetration event times were estimated to be under 30 ms from the start of motor ignition. The mass of the system (not including motors) was estimated to be less than 2.3 kg (5 lbs) per motor. The configuration and demonstrations are discussed.

landing system↗
NASA NTRS2016

Heat-shield for Extreme Entry Environment Technology (HEEET) Development Status

The Heat shield for Extreme Entry Environment Technology (HEEET) Project is a NASA STMD and SMD co-funded effort. The goal is to develop and mission infuse a new ablative Thermal Protection System that can withstand extreme entry. It is targeted to support NASA's high priority missions, as defined in the latest decadal survey, to destinations such as Venus and Saturn in-situ robotic science missions. Entry into these planetary atmospheres results in extreme heating. The entry peak heat-flux and associated pressure are estimated to be between one and two orders of magnitude higher than those experienced by Mars Science Laboratory or Lunar return missions. In the recent New Frontiers community announcement NASA has indicated that it is considering providing an increase to the PI managed mission cost (PIMMC) for investigations utilizing the Heat Shield for Extreme Entry Environment Technology (HEEET) and in addition, NASA is considering limiting the risk assessment to only their accommodation on the spacecraft and the mission environment. The HEEET ablative TPS utilizes 3D weaving technology to manufacture a dual layer material architecture. The 3-D weaving allows for flat panels to be woven. The dual layer consists of a top layer designed to withstand the extreme external environment while the inner or insulating layer by design, is designed to achieve low thermal conductivity, and it keeps the heat from conducting towards the structure underneath. Both arc jet testing combined with material properties have been used to develop thermal response models that allows for comparison of performance with heritage carbon phenolic. A 50% mass efficiency is achieved by the dual layer construct compared to carbon phenolic for a broad range of missions both to Saturn and Venus. The 3-D woven flat preforms are molded to achieve the shape as they are compliant and then resin infusion with curing forms a rigid panels. These panels are then bonded on to the aeroshell structure. Gaps exist between the panels and these gaps have to be filled with seams. The seam material then has to be bonded on to adjacent panels and also to the structure. The heat-shield assembly is shown in Figure 1. One of the significant challenges we have overcome recently is the design, development and testing of the seam. HEEET material development and the seam concept development have utilized some of the unique test capabilities available in the US. The various test facilities utilized in thermal testing along with the entry environment for Saturn and Venus missions are shown in Figure 2. The HEEET project is currently in it's 3rd year of a four-year development. Figure 3 illustrates the key accomplishments to-date and the challenges yet to be overcome before the technology is ready for mission infusion. This proposed presentation will cover both progress that has been made in the HEEET project and also the challenges to be overcome that is highlighted in Figure 3. Objective of the HEEET project is to mature the system in time to support the next New Frontiers opportunity and we believe we are well along the way to mission infuse HEEET.

TPS↗
NASA NTRS2014

Visualization of Capsule Reentry Vehicle Heat Shield Ablation Using Naphthalene PLIF

The Orion Multi-Purpose Crew Vehicle (MPCV) will use an ablative heat shield and improved understanding of the ablation process would be beneficial for design purposes. Given that ablation is a multi-physics process involving heat and mass transfer, codes aiming to predict heat shield ablation are in need of experimental data pertaining to the turbulent transport of ablation products for validation. At The University of Texas at Austin, a technique is being developed that uses planar laser-induced fluorescence (PLIF) of a low-temperature sublimating ablator (naphthalene) to visualize the transport of ablation products in a supersonic flow. Since ablation at reentry temperatures can be difficult to recreate in a laboratory setting it is desirable to create a limited physics problem and simulate the ablation process at relatively low temperature conditions using naphthalene. A scaled Orion MPCV model with a solid naphthalene heat shield has been tested in a Mach 5 wind tunnel at various angles of attack in the current work. PLIF imaging reveals the distribution of the ablation products as they are transported into the heat-shield boundary layer and over the capsule shoulders into the separated shear layer and backshell recirculation region. Visualizations of the capsule shear layer using both naphthalene PLIF and Schlieren imaging compared favorably. High concentrations of naphthalene in the capsule separated flow region, intermittent turbulent structures on the heat shield surface, and interesting details of the capsule shear layer structure were observed using the naphthalene PLIF technique. The capsule shear layer was also shown to generally appear to be more turbulent at lower angles of attack. Furthermore, the PLIF signal increased steadily over the course of a run indicating that during a wind tunnel run the model heated up and the rate of naphthalene ablation increased. The shear layer showed increasing signs of turbulence over the course of a wind tunnel run as well, likely because of the combination of increased surface roughness and surface blowing rate. PLIF imaging also detected regions with a relatively low concentration of naphthalene in the capsule backshell recirculation region that are most likely the result of cross-flow-induced vortices on the capsule afterbody.

Combs, Christopher S.↗
NASA NTRS1974

On thermal stress failure of the SNAP-19A RTG heat shield

Results of a study on thermal stress problems in an amorphous graphite heat shield that is part of the launch-abort protect system for the SNAP-19A radio-isotope thermoelectric generators (RTG) that will be used on the Viking Mars Lander are presended. The first result is from a thermal stress analysis of a full-scale RTG heat source that failed to survive a suborbital entry flight test, possibly due to thermal stress failure. It was calculated that the maximum stress in the heat shield was only 50 percent of the ultimate strength of the material. To provide information on the stress failure criterion used for this calculation, some heat shield specimens were fractured under abort entry conditions in a plasma arc facility. It was found that in regions free of stress concentrations the POCO graphite heat shield material did fracture when the local stress reached the ultimate uniaxial stress of the material.

Pitts, W. C.↗
NASA NTRS1974

Outer planet probe entry thermal protection. II - Heat shielding requirements

Solutions are presented for the material thermal response for heat shields of blunt probes entering Saturn and Uranus. Both monolithic and sandwich concepts were considered having exposed charring ablative or reflective materials. The solutions show that (1) Teflon heat shields are sized by mass loss considerations, other materials (unless entering the Uranus cold-dense atmosphere) by heat soak considerations, (2) steeper entries require less heat shield mass, (3) sizing requirements vary little about the bodies for heat soak dominated materials, and (4) carbon-phenolic appears superior overall, but silica is superior for the steeper entries and the high helium content atmospheres.

Nicolet, W. E.↗
NASA NTRS1974

Major uncertainties influencing entry probe heat shield design

Factors influencing the design of an outer planet probe heat shield are discussed. Major factors included are: uncertainties in composition and scale height of the planet atmospheres; the augmentation/attenuation of entry heating by ablation products requires more computer study and testing; carbon heat shields, especially carbon phenolic, possessing improved resistance to spallation need developing; and white silica reflecting heat shields with improved resistance to bulk vitrification need further developing.

Congdon, W.↗
NASA NTRS2010

Thermal Protection System (Heat Shield) Development - Advanced Development Project

The Orion Thermal Protection System (TPS) ADP was a 3 1/2 year effort to develop ablative TPS materials for the Orion crew capsule. The ADP was motivated by the lack of available ablative TPS's. The TPS ADP pursued a competitive phased development strategy with succeeding rounds of development, testing and down selections. The Project raised the technology readiness level (TRL) of 8 different TPS materials from 5 different commercial vendors, eventual down selecting to a single material system for the Orion heat shield. In addition to providing a heat shield material and design for Orion on time and on budget, the Project accomplished the following: 1) Re-invigorated TPS industry & re-established a NASA competency to respond to future TPS needs; 2) Identified a potentially catastrophic problem with the planned MSL heat shield, and provided a viable, high TRL alternate heat shield design option; and 3) Transferred mature heat shield material and design options to the commercial space industry, including TPS technology information for the SpaceX Dragon capsule.

Kowal, T. John↗
NASA NTRS1992

Heat shield erosion in a dusty Martian atmosphere

The study calculates the effects of dust particle impacts on the erosion of the forebody heat shield for a 26-m diameter aerobraking vehicle entering the Martian atmosphere at 8.6 km/s. An explicit thin-layer Navier-Stokes code is used to compute the dustless flowfield about the vehicle for the actual Martian atmospheric composition. The deceleration and melting of the dust particles within the forebody shock layer was computed for dust spherules having initial diameters from 1 to 19 microns. All the particles began vaporizing shortly after entering the shock layer, but most survived to hit the heat shield surface. The two different heat shield materials considered were Shuttle ceramic tiles with glassy surfaces and a moderate-density ablator. For a vehicle with a ballistic coefficient of 200 kg/sq m, a glassy heat shield surface experienced an average of 7 mm of surface erosion, or 20 times the thickness of the glassy layer on ceramic tiles.

Papadopoulos, Periklis↗
NASA NTRS2018

Orion Heat Shield Manufacturing Producibility Improvements for the EM-1 Flight Test Program

This paper describes how the ORION program is incorporating improvements in the heat shield design and manufacturing processes reducing programmatic risk and ensuring crew safety in support of NASA's Exploration missions. The approach for the EFT-1 heat shield utilized a low risk Apollo heritage design and manufacturing process using an Avcoat TPS ablator with a honeycomb substrate to provide a one piece heat shield to meet the mission re-entry heating environments. The EM-1 mission will have additional flight systems installed to fly to the moon and return to Earth. Heat shield design and producibility improvements have been incorporated in the EM-1 vehicle to meet deep space mission requirements. The design continues to use the Avcoat material, but in a block configuration to enable improvements in consistant and repeatable application processes using tile bonding experience developed on the Space Shuttle Transportation System Program.

Producibility↗
NASA NTRS2015

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.

Blosser, Max L.↗
NASA NTRS2011

Mars Exploration Rover Heat Shield Recontact Analysis

The twin Mars Exploration Rover missions landed successfully on Mars surface in January of 2004. Both missions used a parachute system to slow the rover s descent rate from supersonic to subsonic speeds. Shortly after parachute deployment, the heat shield, which protected the rover during the hypersonic entry phase of the mission, was jettisoned using push-off springs. Mission designers were concerned about the heat shield recontacting the lander after separation, so a separation analysis was conducted to quantify risks. This analysis was used to choose a proper heat shield ballast mass to ensure successful separation with low probability of recontact. This paper presents the details of such an analysis, its assumptions, and the results. During both landings, the radar was able to lock on to the heat shield, measuring its distance, as it descended away from the lander. This data is presented and is used to validate the heat shield separation/recontact analysis.

Raiszadeh, Behzad↗
NASA NTRS2012

Regolith-Derived Heat Shield for Planetary Body Entry and Descent System with In-Situ Fabrication

High-mass planetary surface access is one of NASA's Grand Challenges involving entry, descent, and landing (EDL). Heat shields fabricated in-situ can provide a thermal protection system for spacecraft that routinely enter a planetary atmosphere. Fabricating the heat shield from extraterrestrial regolith will avoid the costs of launching the heat shield mass from Earth. This project investigated three methods to fabricate heat shield using extraterrestrial regolith and performed preliminary work on mission architectures.

Hogue, Michael D.↗
NASA NTRS1972

Gamma heating in reflector heat shield of gas core reactor

Heating rate measurements made in a mock-up of a BeO heat shield for a gas core nuclear rocket engine yields results nominally a factor of two greater than calculated by two different methods. The disparity is thought to be caused by errors in neutron capture cross sections and gamma spectra from the low cross-section elements, D, O, and Be.

Lofthouse, J. H.↗
NASA NTRS2012

Regolith-Derived Heat Shield for Planetary Body Entry and Descent System with In-Situ Fabrication

High-mass planetary surface access is one of NASA's Grand Challenges involving entry, descent, and landing (EDL). Heat shields fabricated in-situ can provide a thermal protection system for spacecraft that routinely enter a planetary atmosphere. Fabricating the heat shield from extraterrestrial regolith will avoid the costs of launching the heat shield mass from Earth. This project will investigate three methods to fabricate heat shield using extraterrestrial regolith.

Hogue, Michael D.↗