SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
Data tables for optimal thermal protection of liquid hydrogen storage in nuclear interplanetary spacecraft
Many of the major issues have been identified for a futuristic capability to protect against impacting comets and asteroids, and a preliminary space-based concept has been envisioned. Some of the basic concept elements, approaches, methodologies, and features have been identified. When contemplating the ability to monitor comets and asteroids continuously, there are many trade-offs between orbiting observatories and detection systems on planetary bodies without an atmosphere. Future orbit modification techniques have the potential for rapid and controlled alteration of NEO orbits, provided that high-power, compatible thermal management systems are developed. Much additional work and analysis are required to identify a final system concept, and many trade studies need to be performed to select the best mix of system capability, reliability, maintainability, and cost. Finally, it is fully appreciated that at the present time space systems are much more costly than terrestrial-based systems. Hopefully, this will change in the future. Regardless, understanding what it would take to defend against a much wider range of the impact threat will foster ideas, innovations, and technologies that could one day enable the development of such a system. This understanding is vital to provide ways of reducing the costs and quantifying the benefits that are achievable with a system like CAPS.
Metalliding parameters developed for sequential deposition of manganese, aluminum, and tantalum alloys as protective coatings for superalloys
Study of protection requirements for resistance of meteorite penetration damage for various interplanetary spacecraft systems
Production engineering and chemical, mechanical, and thermodynamic properties of thoria dispersed nickel chrome alloys for space shuttle thermal protection applications
The Mars Science Laboratory successfully landed on the Martian surface on August 5th, 2012. The rover was protected from the extreme heating environments of atmospheric entry by an ablative heatshield. This Phenolic Impregnated Carbon Ablator heatshield was instrumented with a suite of embedded thermocouples, isotherm sensors, and pressure transducers. The sensors monitored the in-depth ablator response, as well as the surface pressure at discrete locations throughout the hypersonic deceleration. This paper presents a comparison of the flight data with post-entry estimates. An assessment of the aerothermal environments, as well as the in-depth response of the heatshield material is made, and conclusions regarding the overall performance of the ablator at the suite locations are presented.
In 2012, NASA's mission of landing the Mars Science Laboratory (MSL) on Mars was successful. MSL was protected by an ablative heatshield of tiled low density material known as PICA (phenolic impregnated carbon ablator). The heatshield was instrumented with MEDLI (MSL Entry Descent Landing Instrument) a suite of sensors & thermocouples at discrete locations in order to monitor the in-depth ablator temperature response and surface pressure. MEDLI was designed and developed by NASA Langley, in partnership with NASA Ames Research Center for the purposes of probing Mars and evaluating the performance of the spacecraft upon entry into the Martian atmosphere. The flight data reduces the uncertainty in engineer models for predicting the response of a spacecraft towards the extreme heating environment of an entry into the Martian atmosphere. MEDLI2 is a part of the Mars 2020 mission and is the next-generation sensor suite for entry, descent, and landing. This data will again help engineers validate their flight models. Additionally, the atmospheric data, can help us understand atmospheric density and winds. This is a critical study for reducing risks to both robotic and future human missions to Mars. Engineered models for Mars 2020 are dependent upon parameters related to the materials response to heating and radiation. In this work, the thermal properties and other measurements of various ablative materials are analyzed to achieve greater utility of the 2012 MEDLI flight data and more accurately determine the parameters being used in Mars 2020. The purpose of this study was to measure specific heat, thermal conductivity, char yield, reflectance as a function of wavelength, and other thermal parameters and optical properties of various ablative materials. CO2(g) at extreme temperatures emits radiation impacting MEDLI2 flight predictions. Emissivity & absorptivity as a function of temperature were calculated from FTIR and UV-Vis data as a means of investigating whether these materials might absorb CO2 radiation upon entry into the Mars atmosphere.
Brief Presenter Biography:Ruth Miller is an aer-ospace systems engineer in the Entry Systems and Ve-hicle Development Branch at NASA Ames Research Center.Introduction:The flexible thermal protection sys-tem (FTPS) on NASA’s Low-Earth Orbit Flight Test of an Inflatable Decelerator (LOFTID) vehicle will be in-strumented with thermocouples (TCs) to measure the in-depth thermal response during entry into Earth’s atmos-phere[1, 2]. Accurate flight temperature measurements are critical for verifying vehicle performance during the flight test and reducing uncertainties in the thermal models.However,the deployable nature of inflatable decelerator technology presents challengesfor integrat-ing TCs, specifically the TCsneed to be compactableand cannot damage the FTPSnor the inflatable structure(IS).Unlike traditional rigid aeroshells, routing TCsthrough the thickness of the FTPS could cause signifi-cant damage during packing of the deployable aeroshellbecause the different layers may shift small amounts in relation to each other imparting strain on the TCsand FTPS materials. The LOFTID TCleads are routed from the measurement location back to the data acquisition system in the vehicle centerbody within the same FTPS layer that they are monitoring the temperature. This ap-proach eliminates the need to put holes in the FTPS lay-ers,butas a consequence,the insulated TCleads travel for an appreciable distance through a region that will expose them to high temperaturesand large thermal gra-dients.LOFTID’s TCswere baselined to be commercially available Type K TCswith a binder impregnated glass braid insulation. These TCswere chosen because they had been used successfully on IRVE-3 and in ground-based arc jet testing. Additionally, these TCsdid notdamage the FTPS nor IS during packing and deploy-ment testing. However, the glass braid insulation is only rated to a maximum continuous use temperature of 482ºC. For reference, the duration of the heat pulse on the LOFTID vehicle is on the order of minutes and themaximum predicted temperaturebeneaththe outermost FTPS layersis 1350ºC.Ground-based testing in a tube furnace at NASA Ames Research Centerwas conducted to determine if the baseline TCsrouted through FTPS samples would survive and provide accurate temperature measure-mentsat LOFTID flight-relevant temperatures[3].The test results showed large measurement errors occurred beginning at approximately 400°C due to conductive deposits on the TCinsulation electrically shorting the TCleads.The conductive deposits and thus electrical shorting weredetermined to be caused by twoerror sources:1.The organic binder on the TCinsulation carbon-izingin a high temperature, low oxygen envi-ronment2.Decomposition products from the FTPSperme-atingthe braided TC insulationFurther testing in the tube furnace demonstrated that heat cleaning the TCinsulation effectivelyremovedthe organic binder andeliminatedthe first error source.The second error sourcewas shown to be mitigated by the addition of amica wraparound each individual TCleadtoact as an impermeable barrier.To understand the applicability of theground-based tube furnace test results to flight,anarc jet testseriesat Boeing’s Large Core Arc Tunnel (LCAT) facilitywas conducted[4].The error sourcesand mitigation strate-giesidentified in the tube furnace testing were substan-tiatedin the arc jet testing.However, the arc jet testing also revealed three new error sources:1.Glass TCinsulation meltingwhich resultsin electrical shorting of the TCeither through di-rect contact between the two leads or through the electrically conductive FTPS materials2.TCwire meltingwhich results in a noisy and/or open-loop TCresponse3.Type K TCwire green-rotwhich results in large calibration errors Scope of the Presentation:This presentation will include a brief discussion onthe effect of electrical shorting on the output of a TC(i.e. how to identify elec-trical shorting in TCdataand what the associated erroris).The tube furnace and arc jet test resultswill be dis-cussedand the solutions LOFTID is implementing to mitigate the error sourcesidentified in the tube furnace and arc jet testingwill be presented.Additionally, futureresearch and development work to eliminateTCerror sourcesfor future missionswill be recommended.
User manual for test equipment used in simulating aerodynamic heating and acoustical environments on space shuttle thermal protection specimens
This viewgraph presentation provides information on the testing of a new foam for thermal protection of the Space Shuttle External Tank. The testing was conducted on an F-15 fighter aircraft.
Explore the source record for details and available documents.
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.
Explore the source record for details and available documents.
Explore the source record for details and available documents.
During lift-off, the shuttle launch vehicle (external tank, solid rocket booster and orbiter) may be subjected to a lightning strike. Tests of a proposed lightning protection method for the external tank and development materials which were subjected to simulated lightning strikes are described. Results show that certain of the high resistant paint strips performed remarkably well in diverting the 50 kA lightning strikes.
Following the trend of creating flight code with state chart modeling and auto-code generation, NASA's Deep Impact (DI) project uses Stateflow for its Fault Protection Flight Software development.
After every Shuttle mission, the Solid Rocket Boosters (SRBs) are recovered and observed for missing material. Most of the SRB is covered with a cork-based thermal protection material (MCC-l). After the most recent shuttle mission, STS-114, the forward section of the booster appeared to have been impacted during flight. The darkened fracture surfaces indicated that this might have occurred early in flight. The scope of the analysis included microscopic observations to assess the degree of heat effects and locate evidence of the impact source as well as chemical analysis of the fracture surfaces and recovered foreign material using Fourier Transform Infrared Spectroscopy and Scanning Electron Microscopy/Energy Dispersive Spectroscopy. The amount of heat effects and presence of soot products on the fracture surface indicated that the material was impacted prior to SRB re-entry into the atmosphere. Fragments of graphite fibers found on these fracture surfaces were traced to slag inside the Solid Rocket Motor (SRM) that forms during flight as the propellant is spent and is ejected throughout the descent of the SRB after separation. The direction of the impact mark matches with the likely trajectory of SRBs tumbling prior to re-entry.
The development of a unique closed-pore ceramic foam insulation (CPI) produced from low cost fly ash cenospheres is reported for space shuttle external thermal protection. Two basic design approaches were developed: bonded and mechanically fastened. A description of the concepts is presented in addition to fabrication and test results.