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At least 235 records · Page 13
All sourcesNASA NTRSDOE OSTI
NASA NTRS2018

Lightweight, Durable, and Multifunctional Electrical Insulation Material Systems for High Voltage Applications

Newly developed multilayer structures of well-known polymer insulation materials significantly improved dielectric breakdown voltage, VB, or dielectric strength, K, if well-bonded, when compared to those of single material insulations or the commercial SOA systems, such as Teflon-Kapton-Teflon (TKT), at the same overall thickness. To date, the greatest improvement of the new structures from a few candidate materials, including various types of Kapton PIs and PFA or PET as bond layer (BL), was about 61% higher than that of the Kapton PI alone films, 40.1 vs. 24.9 kV, which was translated to 86.3% decrease in insulation thickness, thus significant volume and weight reduction of the final system. However, it was of interest to note that most improvements of the multilayer structures occurred at thicker overall thicknesses, above ~ 0.15 mm. Extensive analyses also showed that K of the multilayer structures increased with (i) decreasing individual layer thickness regardless of material type, (ii) increasing total accumulated thickness of PI or overall PI/BL ratio, and (iii) increasing number of interface or total number of layers, but only above the aforementioned overall thickness limit. Increases in VB of the multilayer structures were directly correlated with damage evolution and failure mode. With further material-design-process optimizations of the multilayer structures, it was expected to achieve other multifunctionalities, such as high partial discharge (PD) resistance, improved durability, EMI shielding, and high thermal dissipation in addition to high dielectric strength. These new structures can be used in various high voltage and high temperature applications, such as future hybrid or all electric aircraft wiring and power transmission as well as many other non-aerospace high power cables, electronic parts and components, printed circuit board, and so forth. The multilayer insulation system can be easily processed and manufactured with various conductor types via calendaring, compression-molding, stamping, laminating, vacuum-bagging and autoclaving, or 3D printing, even for complex 3-D components. Based on their unique structural configurations and potential capabilities, the new insulation system was identified as micro-multilayer multifunctional electrical insulation (MMEI). Patent application of the MMEI concept and current design configurations was filed for a 1-year provisional application (OAI-58834, Serial No.: 62/659,234), pending conversion to a U.S. utility application. This paper presents details of the MMEI structures, their dielectric performance analyses, potential mechanisms, and commercial scaleup feasibility assessment.

Multifunctional↗
NASA NTRS1975

Development program to produce mullite fiber insulation

Processing methods were utilized to form a mullite fiber-Kaowool felt. The formation of a blended felt using the Rotoformer wet-laying method was successful. Felt products were evaluated for tensile strength, thermal stability, thermal conductivity and structural integrity at 1259 C and 1371 C. Textile processing methods failed in an attempt to form a yarn from staple and multifilament mullite fiber due to fiber damage through mechanical handling. The refractoriness of pure Kaowool ceramic fiber is improved with additions of 30% or greater mullite fiber.

Long, W. G.↗
NASA NTRS1985

An assessment of wind tunnel test data on flexible thermal protection materials and results of new fatigue tests of threads

Advanced Flexible Reusable Surface Insulation (AFRSI) was developed as a replacement for the low-temperature (white) tiles on the Space Shuttle. The first use of the AFRSI for an Orbiter flight was on the OMS POD of Orbiter (OV-099) for STS-6. Post flight examination after STS-6 showed that damage had occurred to the AFRSI during flight. The failure anomaly between previous wind-tunnel tests and STS-6 prompted a series of additional wind tunnel tests to gain an insight as to the cause of the failure. An assessment of all the past STS-6 wind tunnel tests pointed out the sensitivity of the test results to scaling of dynamic loads due to the difference of boundary layer thickness, and the material properties as a result of exposure to heating. The thread component of the AFRSI was exposed to fatigue testing using an apparatus that applied pulsating aerodynamic loads on the threads similar to the loads caused by an oscillating shock. Comparison of the mean values of the number-of-cycles to failure showed that the history of the thread was the major factor in its performance. The thread and the wind tunnel data suggests a mechanism of failure for the AFRSI.

Coe, Charles F.↗
NASA NTRSDate not supplied

Summary of Testing Results for the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER)

Testing was completed on the Structural Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) between August 2019 and January 2020. SHIIVER was designed to be a test bed for the scaling of cryogenic fluid management technologies as applied to large upper stages and long duration in-space stages. The baseline SHIIVER design consists of uninsulated structural skirts attached to a propellant tank insulated with polyurethane Spray-On Foam Insulation (SOFI). The initial testing of the SHIIVER hardware was with liquid hydrogen and sought to demonstrate the use of boil-off vapor to intercept heat on a structural skirt, multilayer insulation (MLI) on the tank domes, and the radio frequency mass gauge (RFMG). Testing was completed in four stages: a baseline thermal vacuum test prior to installation of the MLI on the tank domes, a thermal vacuum test after the MLI installation, a reverberant acoustic test, and a subsequent thermal vacuum test to verify that no damage occurred during the reverberant acoustic testing. Each thermal vacuum test with chamber wall at ambient temperature and vacuum level in the 10-6 torr range was conducted continuously between approximately 90% full and 25% full. Test results showed that the vapor cooling reduced the heat load to the tank by approximately 10%, but the boil-off by less than 3% at 50% full, with and without MLI installed on the domes. The MLI installed on the domes reduced the heat load to the tank by approximately 40% at all fill levels, but the boil-off by approximately 25% at 90% full and 45% below 65% full. The RFMG performed well over all fill ranges, and several RF tank modes were used to gauge the mass of fluid in the tank. SHIIVER was then exposed to an acoustic environment of 147 dB OASPL (overall sound pressure level) in a reverberant chamber. The acoustic environment envelopes the upper stage internal acoustic level of several different modern launch vehicles. No structural or thermal performance changes were observed after exposure to the acoustic environment. Final thermal vacuum testing after the acoustic testing showed no degradation to the MLI due to the acoustic environment as measured via system heat loads

SHIIVER↗
NASA NTRS2007

NDE Methods for the External Tank Spray on Foam Insulation

The foam thermal protection system (TPS) of the space shuttle External Tank (ET) has provided some unique challenges to the nondestructive testing community. Three nondestructive evaluation methods have been developed to identify defects in the foam TPS of the ET. Terahertz imaging and backscatter radiography have been developed to identify voids in thick foam regions while shearography has been developed to identify shallow delaminations, shallow voids and crush damage in the foam. The basic theory of operation along with factors, determined from a series of designed experiments, affecting the results of these methods will be described. Results from both test panels and flight tank inspections will be provided, along with predicted probability of detection estimates to show the range in defect sizes and types that can be readily detected.

Walker, James L.↗
NASA NTRS1985

Effect of processing treatments on strength of silica thread for quilted ceramic insulation on Space Shuttle

The effects of chemical, mechanical, and thermal processing treatments on the preflight strength properties of a silica sewing thread used to fabricate an external flexible Thermal Protection System (TPS) for the Space Shuttle vehicle are discussed. Below 540 C, loss of strength was observed for all treatments studied, but above conditioning temperatures of 540 C little change in strength properties was measured, indicating no appreciable mechanical damage to the yarn as a consequence of these processing treatments.

Sawko, P. M.↗
NASA NTRS1986

Fast Glazing of Alumina/Silica Tiles

Technique for applying ceramic coating to fibrous silica/alumina insulation tiles prevents cracks and substantially reduces firing time. To reduce thermal stresses in tile being coated, high-temperature, shorttime firing schedule implemented. Such schedule allows coating to mature while substrate remains at relatively low temperature, reducing stress differential between coating and substrate. Technique used to repair tiles with damaged coatings and possibly used in heat-treating objects made of materials having different thermal-expansion coefficients.

Creedon, J. F.↗
NASA NTRS2006

Characterization of Space Shuttle Thermal Protection System (TPS) Materials for Return-to-Flight following the Shuttle Columbia Accident Investigation

During the Space Shuttle Columbia Accident Investigation, it was determined that a large chunk of polyurethane insulating foam (= 1.67 lbs) on the External Tank (ET) came loose during Columbia's ascent on 2-1-03. The foam piece struck some of the protective Reinforced Carbon-Carbon (RCC) panels on the leading edge of Columbia's left wing in the mid-wing area. This impact damaged Columbia to the extent that upon re-entry to Earth, superheGed air approaching 3,000 F caused the vehicle to break up, killing all seven astronauts on board. A paper after the Columbia Accident Investigation highlighted thermal analysis testing performed on External Tank TPS materials (1). These materials included BX-250 (now BX-265) rigid polyurethane foam and SLA-561 Super Lightweight Ablator (highly-filled silicone rubber). The large chunk of foam from Columbia originated fiom the left bipod ramp of the ET. The foam in this ramp area was hand-sprayed over the SLA material and various fittings, allowed to dry, and manually shaved into a ramp shape. In Return-to-Flight (RTF) efforts following Columbia, the decision was made to remove the foam in the bipod ramp areas. During RTF efforts, further thermal analysis testing was performed on BX-265 foam by DSC and DMA. Flat panels of foam about 2-in. thick were sprayed on ET tank material (aluminum alloys). The DSC testing showed that foam material very close to the metal substrate cured more slowly than bulk foam material. All of the foam used on the ET is considered fully cured about 21 days after it is sprayed. The RTF culminated in the successful launch of Space Shuttle Discovery on 7-26-05. Although the flight was a success, there was another serious incident of foam loss fiom the ET during Shuttle ascent. This time, a rather large chunk of BX-265 foam (= 0.9 lbs) came loose from the liquid hydrogen (LH2) PAL ramp, although the foam did not strike the Shuttle Orbiter containing the crew. DMA testing was performed on foam samples taken fiom a simulated PAL ramp panel. It was found that the smooth rind on the foam facing the cable tray did significantly affect the properties of foam at the PAL ramp surface. The smooth rind increased the storage modulus E' of the foam as much as 20- 40% over a temperature range of -145 to 95 C. Because of foam loss fiom the PAL ramp, future Shuttle flights were grounded indefinitely to allow further testing to better understand foam properties. The decision was also made to remove foam from the LH2 PAL, ramp. Other RTF efforts prior to the launch of Discovery included

Wingard, Doug↗
NASA NTRS1987

Space station integrated wall design and penetration damage control. Task 4: Impact detection/location system

A program to develop a methodology is documented for detecting and locating meteoroid and debris impacts and penetrations of a wall configuration currently specified for use on space station. Testing consisted of penetrating and non-penetrating hypervelocity impacts on single and dual plate test configurations, including a prototype 1.22 m x 2.44 m x 3.56 mm (4 ft x 8 ft x 0.140 in) aluminum waffle grid backwall with multilayer insulation and a 0.063-in shield. Acoustic data were gathered with transducers and associated data acquisition systems and stored for later analysis with a multichannel digitizer. Preliminary analysis of test data included sensor evaluation, impact repeatability, first waveform arrival, and Fourier spectral analysis.

Nelson, J. M.↗
NASA NTRS2019

Analyses and Methods of Solid Rocket Motor Material Irradiation at Marshall Space Flight Center

The search for life on other worlds is among humanity’s greatest endeavors. Europa represents the most probable location to discover extraterrestrial life in our solar system, owing to its surface composition of ice covering a liquid water ocean, warmed by the tidal forces of its orbit around Jupiter. Unfortunately, the Jovian system hosts the most intense planetary radiation environment in the solar system due to the charged particles, namely electrons and protons, trapped by Jupiter’s immense magnetic field. Any mission that attempts to approach or land on Europa must survive this radiation environment [Hand et al, 2017]. Radiation effects were identified as a priority risk to the successful development of a de-orbit stage and solid rocket motor (SRM) early in the Europa Lander De-orbit Stage project concept. The effects of primary concern tend to occur very near the outer surface of the SRM. The charged particles deposit their energy quickly and are mostly stopped in the outer metallic case, but a significant portion of radiation penetrates through the bondline and outer propellant regions. High doses of ionizing radiation are known to cause significant changes to mechanical properties of many materials, especially polymers. For polymers such as the rubber-like materials (elastomers) in a solid rocket motor, the primary damage mechanism is known as cross-linking, in which ionization causes the restructuring of the matrix of long polymer chains. Ionization energy breaks the long polymer chains and allows formation of new cross-linked bonding sites. This hardens and often strengthens the polymer, but at the cost of decreased flexibility (or modulus). Propellant, insulation, liner, and pyrotechnic materials were identified as higher risk items, and so were irradiated at Marshall Space Flight Center (MSFC) for investigation of changes in mechanical and ballistic properties. This process required significant levels of analysis to evaluate how the radiation environment evolves within the spacecraft during the mission, and also to evaluate how dose is delivered into test articles within the irradiating facilities.

Caffrey, Jarvis↗
NASA NTRS2019

Analyses and Methods of Solid Rocket Motor Material Irradiation at Marshall Space Flight Center

The search for life on other worlds is among humanity’s greatest endeavors. Europa represents the most probable location to discover extraterrestrial life in our solar system, owing to its surface composition of ice covering a liquid water ocean, warmed by the tidal forces of its orbit around Jupiter. Unfortunately, the Jovian system hosts the most intense planetary radiation environment in the solar system due to the charged particles, namely electrons and protons, trapped by Jupiter’s immense magnetic field. Any mission that attempts to approach or land on Europa must survive this radiation environment [Hand et al, 2017]. Radiation effects were identified as a priority risk to the successful development of a de-orbit stage and solid rocket motor (SRM) early in the Europa Lander De-orbit Stage project concept. The effects of primary concern tend to occur very near the outer surface of the SRM. The charged particles deposit their energy quickly and are mostly stopped in the outer metallic case, but a significant portion of radiation penetrates through the bondline and outer propellant regions. High doses of ionizing radiation are known to cause significant changes to mechanical properties of many materials, especially polymers. For polymers such as the rubber-like materials (elastomers) in a solid rocket motor, the primary damage mechanism is known as cross-linking, in which ionization causes the restructuring of the matrix of long polymer chains. Ionization energy breaks the long polymer chains and allows formation of new cross-linked bonding sites. This hardens and often strengthens the polymer, but at the cost of decreased flexibility (or modulus). Propellant, insulation, liner, and pyrotechnic materials were identified as higher risk items, and so were irradiated at Marshall Space Flight Center (MSFC) for investigation of changes in mechanical and ballistic properties. This process required significant levels of analysis to evaluate how the radiation environment evolves within the spacecraft during the mission, and also to evaluate how dose is delivered into test articles within the irradiating facilities.

Caffrey, Jarvis↗
NASA NTRS2011

Demonstration of a Particle Impact Monitoring System for Crewed Space Exploration Modules

When micrometeorite or debris impacts occur on a space habitat, crew members need to be quickly informed of the likely extent of damage, and be directed to the impact location for possible repairs. The goal of the Habitat Particle Impact Monitoring System (HIMS) is to develop a fully automated, end-to-end particle impact detection system for crewed space exploration modules, both in space and on the surfaces of Solar System bodies. The HIMS uses multiple thin film piezo-polymer vibration sensors to detect impacts on a surface, and computer processing of the acoustical signals to characterize the impacts. Development and demonstration of the HIMS is proceeding in concert with NASA's Habitat Demonstration Unit (HDU) Project. The HDU Project is designed to develop and test various technologies, configurations, and operational concepts for exploration habitats. This paper describes the HIMS development, initial testing, and HDU integration efforts. Initial tests of the system on the HDU were conducted at NASA?s 2010 Desert Research and Technologies Studies (Desert-RATS). Four sensor locations were assigned near the corners of a rectangular pattern. To study the influence of wall thickness, three sets of four sensors were installed at different layer depths: on the interior of the PEM wall, on the exterior of the same wall, and on the exterior of a layer of foam insulation applied to the exterior wall. Once the system was activated, particle impacts were periodically applied by firing a pneumatic pellet gun at the exterior wall section. Impact signals from the sensors were recognized by a data acquisition system when they occurred, and recorded on a computer for later analysis. Preliminary analysis of the results found that the HIMS system located the point of impact to within 8 cm, provided a measure of the impact energy / damage produced, and was insensitive to other acoustic events. Based on this success, a fully automated version of this system will be completed and demonstrated as part of a crew "Caution/Warning" system at the 2011 Desert-RATS, along with a crew response procedure.

Opiela, J. N.↗
NASA NTRS1998

On-Orbit Teflon FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon' FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon FEP sample evaluation and additional testing of pristine Teflon FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, LTV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon FEP.

Townsend, Jacqueline A.↗
NASA NTRS1999

On-Orbit Teflon(trademark) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(trademark) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(trademark) FEP sample evaluation and additional testing of pristine Teflon FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(trademark) FEP.

Townsend, Jacqueline↗
NASA NTRS1998

On-Orbit Teflon(trademark) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(trademark) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(trademark) FEP sample evaluation and additional testing of pristine Teflon(trademark) FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(trademark) FEP.

Townsend, Jacqueline A.↗
NASA NTRS1998

Analysis of Retrieved Hubble Space Telescope Thermal Control Materials

The mechanical and optical properties of the thermal control materials on the Hubble Space Telescope (HST) have degraded over the nearly seven years the telescope has been in orbit. Astronaut observations and photographs from the Second Servicing Mission (SM2) revealed large cracks in the metallized Teflon FEP, the outer-layer of the multi-layer insulation (MLI), in many locations around the telescope. Also, the emissivity of the bonded metallized Teflon FEP radiator surfaces of the telescope has increased over time. Samples of the top layer of the MLI and radiator material were retrieved during SM2, and a thorough investigation into the de-radiation followed in order to determine the primary cause of the damage. Mapping of the cracks on HST and the ground testing showed that thermal cycling with deep-layer damage from electron and proton radiation are necessary to cause the observed embrittlement. Further, strong, evidence was found indicating that chain scission (reduced molecular weight) is the dominant form of damage to the metallized Teflon FEP.

Townsend, Jacqueline A.↗
NASA NTRS1998

On-Orbit Teflon(R) FEP Degradation

During the Hubble Space Telescope (HST) Second Servicing Mission (SM2), degradation of unsupported Teflon(R) FEP (fluorinated ethylene propylene), used as the outer layer of the multi-layer insulation (MLI) blankets, was evident as large cracks on the telescope light shield. A sample of the degraded outer layer was retrieved during the mission and returned to Earth for ground testing and evaluation. The results of the Teflon(R) FEP sample evaluation and additional testing of pristine Teflon(R) FEP led the investigative team to theorize that the HST damage was caused by thermal cycling with deep-layer damage from electron and proton radiation which allowed the propagation of cracks along stress concentrations, and that the damage increased with the combined total dose of electrons, protons, UV and x-rays along with thermal cycling. This paper discusses the testing and evaluation of the Teflon(R) FEP.

Townsend, Jacqueline A.↗
NASA NTRS2023

Space Weathering of Genesis Mission Solar-wind Collectors with Inferences for Weathering on Airless Bodies

Samples from the Genesis Solar Wind Sample Return (NASA Discovery 5) are a unique opportunity to study the initial stages of space weathering, i.e., the physical and chemical effects of solar-wind irradiation. Arrays of collectors containing multiple materials were each exposed to a different solar-wind regime (fast, slow, bulk, or coronal mass ejection) at the L1 point for long durations (years). Materials exposed to the solar wind included metals, semiconductors, and insulators. Although the time of exposure was obviously short relative to samples having extraterrestrial origins, optical properties, surface chemical properties, and matrix structure have changed in many collectors due to exposure to solar wind. The thickness of amorphous zones, where present, appears to correspond with the depth of the peak of the solar-wind H distribution in each regime. Damage from high-energy particles was negligible because the collectors were 700 μm or less in thickness and shielded from the back by the spacecraft. Micrometeorite impacts and sputtering were also negligible because of the short exposure times. Our current results are preliminary: We hope future workers will extend this study both to support Genesis characterization efforts and to further understand space-weathering processes on a geologic timescale.

space weathering↗