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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 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

Structural Testing of a 6m Hypersonic Inflatable Aerodynamic Decelerator System

NASA is developing low ballistic coefficient technologies to support the Nations long-term goal of landing humans on Mars. Current entry, decent, and landing technologies are not practical for this class of payloads due to geometric constraints dictated by current and future launch vehicle fairing limitations. Hypersonic Inflatable Aerodynamic Decelerators (HIADs) are being developed to circumvent this limitation and are now considered a leading technology to enable landing of heavy payloads on Mars. At the beginning of 2014, a 6m diameter HIAD inflatable structure with an integrated flexible thermal protection system (TPS) was subjected to a static load test series to verify its structural performance under flight-relevant loads. The inflatable structure was constructed into a 60 degree sphere-cone configuration using nine inflatable torus segments composed of fiber-reinforced thin films. The inflatable tori were joined together using adhesives and high-strength textile woven structural straps. These straps help distribute the load throughout the inflatable structure. The 6m flexible TPS was constructed using multiple layers of high performance materials that are designed to protect the inflatable structure from heat loads that would be seen in flight during atmospheric entry. A custom test fixture was constructed to perform the static load test series. The fixture consisted of a round structural tub with enough height and width to allow for displacement of the HIAD test article as loads were applied. The bottom of the tub rim had an airtight seal with the floor. The rigid centerbody of the HIAD was mounted to a pedestal in the center of the structural tub. Using an impermeable membrane draped over the HIAD test article, an airtight seal was created with the top rim of the static load tub. This seal allowed partial vacuum to be pulled beneath the HIAD resulting in a uniform static pressure load applied to the outer surface. Using this technique, the test article was subjected to loads of up to 50,000lbs. During the test series an extensive amount of instrumentation was used to provide a rich data set, including deflected shape, structural strap loads, torus cord loads, inflation pressures, and applied static load. In this paper the 2014 6m HIAD static load test series will be discussed in detail, including the design of the 6m HIAD test article, the test setup, and test execution. Analysis results will be described supporting the conclusions that were drawn from the test series..

Hypersonic↗
NASA NTRS1990

Design, analysis, and testing of a reusable wingbox for ballistic evaluation of composite structures

Future fighter aircraft will continue to be subjected to severe mechanical loading and ballistic impact. In the past aircraft primary wing structure was constructed predominantly with aluminum, but with increasing aircraft performance the need for more structurally efficient materials is necessary. In an attempt to achieve structural efficiency requirements, aircraft designers have begun using an increasing amount of composite materials. Composite materials were indicated to have a low level of ballistic tolerance in some applications. Composite materials possess the characteristic of structural tailoring, which was utilized to demonstrate that survivability can be designed into a structure. The combat flight environment subjects fighter aircraft wing structures to the effects of combined loading during ballistic impact. These loads were previously demonstrated independently, i.e., the application of structural load, hydrodynamic ram damage, and the effects of airflow. The objective was to design a reusable test fixture for the ballistic evaluation of advanced composite structures under all the effects of combined loading. The fixture that was developed provides a test bed that can completely simulate the flight environment of a generic fighter aircraft wing, but can be used for an indefinite number and variety of tests. It is intended that this fixture will provide a standardized method of ballistically evaluating an advanced composite primary wing structure. This fixture will be used to accomplish the development of designs and/or advanced concepts that lead to a survivable composite wing structure.

Baron, W. G.↗
NASA NTRS1967

Portable fixture facilitates pressure testing of instrumentation fittings

Portable fixture facilitates pressure testing to detect possible leaks in instrumentation fittings mounted on tank bulkheads. It uses a vacuum cup which seals a pressure regulator adapter around one side of the fitting to be pressure tested. Leakage is detected with a gas sniffer.

Olson, G. A.↗
NASA NTRS1974

Laser geodynamic satellite/thermal/optical vibrational analyses and testing. Volume 2: Technical report, book 3

The overall plan for accomplishing the various tests required to achieve the program objectives and tasks of the LAGEOS thermal/optical/vibration analysis and test program is described. The requirements for each test are identified for the purpose of program review for the design and fabrication of the required test article and test fixtures, for the allocation of test facilities, equipment and expendables, and for the generation of the detail test procedures by which the specified data will be obtained at specified test conditions. The current internal test program schedules, by which the various individual preparatory subtasks are to be accomplished to ensure that the individual test will be conducted within the time period required to meet program objectives are included.

Source record↗
NASA NTRS2011

Using Innovative Technologies for Manufacturing Rocket Engine Hardware

Many of the manufacturing techniques that are currently used for rocket engine component production are traditional methods that have been proven through years of experience and historical precedence. As the United States enters into the next space age where new launch vehicles are being designed and propulsion systems are being improved upon, it is sometimes necessary to adopt innovative techniques for manufacturing hardware. With a heavy emphasis on cost reduction and improvements in manufacturing time, rapid manufacturing techniques such as Direct Metal Laser Sintering (DMLS) are being adopted and evaluated for their use on NASA s Space Launch System (SLS) upper stage engine, J-2X, with hopes of employing this technology on a wide variety of future projects. DMLS has the potential to significantly reduce the processing time and cost of engine hardware, while achieving desirable material properties by using a layered powder metal manufacturing process in order to produce complex part geometries. Marshall Space Flight Center (MSFC) has recently hot-fire tested a J-2X gas generator (GG) discharge duct that was manufactured using DMLS. The duct was inspected and proof tested prior to the hot-fire test. Using a workhorse gas generator (WHGG) test fixture at MSFC's East Test Area, the duct was subjected to extreme J-2X hot gas environments during 7 tests for a total of 537 seconds of hot-fire time. The duct underwent extensive post-test evaluation and showed no signs of degradation. DMLS manufacturing has proven to be a viable option for manufacturing rocket engine hardware, and further development and use of this manufacturing method is recommended.

Betts, E. M.↗
NASA NTRS2018

ORION - Crew Module Side Hatch: Proof Pressure Test Anomaly Investigation

The Orion Multi-Purpose Crew Vehicle program was performing a proof pressure test on an engineering development unit (EDU) of the Orion Crew Module Side Hatch (CMSH) assembly. The purpose of the proof test was to demonstrate structural capability, with margin, at 1.5 times the maximum design pressure, before integrating the CMSH to the Orion Crew Module structural test article for subsequent pressure testing. The pressure test was performed at lower pressures of 3 psig, 10 psig and 15.75 psig with no apparent abnormal behavior or leaking. During pressurization to proof pressure of 23.32 psig, a loud 'pop' was heard at ~21.3 psig. Upon review into the test cell, it was noted that the hatch had prematurely separated from the proof test fixture, thus immediately ending the test. The proof pressure test was expected be a simple verification but has since evolved into a significant joint failure investigation from both Lockheed Martin and NASA.

Evernden, Brent A.↗
DOE OSTI2021

Design and Analysis of an Advanced Three-Point Bend Test Approach for Miniature Irradiated Disk Specimens

Candidate tungsten armor materials in a magnetic confinement fusion device must be able to withstand thermal variation that leads to internal stresses caused by the impinging heat load. In addition, the thermomechanical properties of these materials are degraded by irradiation-induced defect accumulation. Fission reactor–based irradiation data are used to predict the fusion neutron damage and property change. This study examines the motivation and design of a custom-designed three-point bend test for neutron-irradiated disk specimens that are 3 mm in diameter to be able to define the flexural strength of advanced tungsten materials, alloys, and composites—and to the extent that embrittlement occurs after neutron irradiation. The theory provided shows a calculation for the flexural deflection and shear deflection due to the small-geometry constraints. A finite element deformation analysis is performed to evaluate the mechanical stress field of disk bend specimens. The stress values above 80% of the maximum stress are concentrated in 2.4 mm of the 3.0-mm length of the centerline across the tungsten disk diameter. A bend test fixture has been designed and fabricated to enable testing of these specimens with precisely engineered tolerance and minimal machine compliance. This fixture will be able to be placed inside a universal testing frame at elevated temperatures for the mechanical property evaluation of future neutron-irradiated disk specimens.

36 MATERIALS SCIENCE↗
NASA NTRS1964

Study Program to Determine the Acceleration Environment Capability of the GG159C Gas- Bearing Spinmotor Final Report

The goal of this contract was to determine the g environment under which the GC159C Gas-Bearing Spinmotor (GBSM) could reliably operate. This was fulfilled by building and testing of four GBSM's, a test fixture, and a "dummy" gyro. The test program was divided into two phases when a gas bearing improvement was required to withstand JPL shock requirement of 200 g. Phase I determined existing g capabilities and performance of the GC159C GBSM and gimbal-case structure. Phase II increased GBSM capability to meet required JPL g environments. Life tests were run on two GBSM's which were shocked at a high level to obtain bearing contact while rotating at their operating speed of 23,000 rpm. A third (nonoperating) GBSM was exposed to JPL maximum shock levels, and a fourth (nonoperating) GBSM was exposed to random vibration. Both nonoperating GBSM's were then subjected to life testing.

SPIN↗
NASA NTRS1976

Specification of pyro shock tests

Specifications of pyro shock tests are treated as the interconnection between raw measurements (or estimates) of the environment and the various test methods available. After a definition of the environment, various aspects of test method specification are discussed including one-sided pulses, test simulation parameters, and the effects of Q. Different test configurations are considered including rigid fixture tests, pseudo-real structure, and real structure.

Trummel, M.↗
NASA NTRS1989

Test methods and design allowables for fibrous composites. Volume 2

Topics discussed include extreme/hostile environment testing, establishing design allowables, and property/behavior specific testing. Papers are presented on environmental effects on the high strain rate properties of graphite/epoxy composite, the low-temperature performance of short-fiber reinforced thermoplastics, the abrasive wear behavior of unidirectional and woven graphite fiber/PEEK, test methods for determining design allowables for fiber reinforced composites, and statistical methods for calculating material allowables for MIL-HDBK-17. Attention is also given to a test method to measure the response of composite materials under reversed cyclic loads, a through-the-thickness strength specimen for composites, the use of torsion tubes to measure in-plane shear properties of filament-wound composites, the influlence of test fixture design on the Iosipescu shear test for fiber composite materials, and a method for monitoring in-plane shear modulus in fatigue testing of composites.

Chamis, Christos C.↗
NASA NTRS2001

Friction Factor Characterization for High-Porosity Random Fiber Regenerators

NASA Glenn Research Center, the Department of Energy (DOE), and Stirling Technology Company (STC) of Kennewick, Washington are developing a Stirling convertor for a high-efficiency Stirling Radioisotope Power System to provide electric power for NASA Space Science Missions. STC is developing the 55-We Technology Demonstration Convertor (TDC) under contract to DOE. Steady-flow tests were completed to determine the friction factor for the high-porosity regenerators that are used in the TDC. STC fabricated a flow test fixture and three random fiber regenerator test samples, one each at approximately 80, 88, and 96 percent porosities. The flow tests were then completed by the NASA Glenn Flow Calibration Laboratory, and the data reduced to Reynolds number and friction factor. The results showed that the 80 and 88 percent porosity samples had similar characteristics while the 96 percent porosity sample had significantly higher friction factors for given Reynolds numbers compared to the samples with lower porosities. Comparisons were also made between the test data and existing correlations. STC used this data to derive a modified regenerator friction factor correlation for use in the Stirling design code GLIMPS for porosities greater than 88 percent. Using this new correlation, the final optimized regenerator design porosity was reduced from 96 to 90 percent.

Thieme, Lanny G.↗
NASA NTRS2006

Optical and Mechanical Properties of Glass Blown In Vacuo

Theoretically, the strength of glass processed in vacuum should be higher due to outgassing of contaminants normally present in the glass, such as bulk water in the form of OH bonds that tends to weaken the glass structure. In this research, small discs of a few types of glass have been subjected to various temperatures for extended periods of time in vacuum. Their strength was then tested using a standard flexure technique, facilitated by a custom-designed test fixture, and the results were compared to glass tested in air using the same fixture. The purpose of the glass blowing investigation was to prove the basic feasibility of a high-level concept for in-space manufacture of optical elements. The central requirement was that the glass bubble had to be blown into a support structure such that the bubble could be handled by manipulation of the structure. The blown bubble attached itself to a mullite ring geometrically and mechanically, as a demonstration in the initial experiments described here, by expanding through and around it. The vacuum system used was custom made, as were most of the components of the system, such as the heating element, the glass and ring support structure, and the gas inlet system that provided the pressure needed to blow the glass.

Manning, andrew↗
NASA NTRS1975

Viking 1975 Orbiter Development Test Model/Lander Dynamic Test Model dynamic environmental testing - An overview

The current work presents an overview of the Viking 1975 environmental testing from an engineering standpoint. An extremely large vibration test fixture had to be designed, analyzed, and integrated into a test setup that employed hydrostatic bearings in a new fashion. A vibration control system was also required that would allow for thirty-six channels of sine-wave peak select control from acceleration, force-of-strain transducers. In addition, some 68 channels of peak limiting shutdown capability were needed for backup and monitoring of other data during the forced vibration test. Pretesting included analyses of the fixture design, overturning moment, control system capabilities, and response of the entire spacecraft/fixture/exciter system to the test environment. Closed-loop control for acoustic testing was a necessity due to the fact that the Viking spacecraft took up a major portion of the volume of the 10,000 cu ft chamber. The spacecraft emerged from testing undamaged.

Milder, G.↗
NASA NTRS1997

Fatigue Debonding Characterization in Composite Skin/Stringer Configurations

The objective of this work was to investigate the fatigue damage mechanisms and to identify the influence of skin stacking sequence in carbon epoxy composite bonded skin/stringer constructions. A simple 4-point-bending test fixture originally designed for previously performed monotonic tests was used to evaluate the fatigue debonding mechanisms between the skin and the bonded frame when the dominant loading in the skin is flexure along the edge of the frame. The specimens consisted of a tapered flange, representing the stringer, bonded onto a skin. Based on the results of previous monotonic tests two different skin lay-ups in combination with one flange lay-up were investigated. The tests were performed at load levels corresponding to 40%, 50%, 60%, 70%, and 80% of the monotonic fracture loads. Microscopic investigations of the specimen edges were used to document the onset of matrix cracking and delamination, and subsequent fatigue delamination growth. Typical damage patterns for both specimen configurations were identified. The observations showed that failure initiated near the tip of the flange in the form of matrix cracks at one of two locations, one in the skin and one in the flange. The location of the 90 deg flange and skin plies relative to the bondline was identified as the dominant lay-up feature that controlled the location and onset of matrix cracking and subsequent delamination. The fatigue delamination growth experiments yielded matrix cracking and delamination onset as a function of fatigue cycles as well as delamination length as a function of the number of cycles.

Cvitkovich, Michael K.↗
NASA NTRS2018

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.

Ellerby, Don↗
NASA NTRS2019

Overview of Heatshield for Extreme Entry Environment Technology (HEEET)

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 > 3500 W/cm2 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. Manufacturing/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. 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 4pt bend testing at sub-ambient and elevated temperature and culminates in testing of the ETU results from which will 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 4pt bend test fixture was developed allowing articles to be tested while the front surface is heated with a laser. These tests are being utilized to establish the systems structural capability during entry.A broad range of aerothermal tests (arcjet tests) were performed to develop material response models for predicting the required TPS thickness to meet a missions needs and to evaluate failure modes and establish the capability of the system. 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 for each of these areas and argue that HEEET has successfully achieved TRL 6.

HEEET↗