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At least 271 records · Page 15

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NASA NTRSDate not supplied

Reliability Issues with Polymer and MnO2 Tantalum Capacitors for Space Applications

This work gives a comparative analysis of degradation processes, failure modes and mechanisms in MnO2 and polymer technology capacitors. Analyzed conditions include effects of vacuum and radiation, soldering (pop-corning), long-term storage and operation at high temperatures, stability at low and high temperatures, and anomalous transients. Screening and qualification procedures to assure space-grade quality of CPTCs are suggested.

tantalum capacitors↗

NASA NTRSDate not supplied

Polymeric Materials Application and Testing for Space Missions

The EM41 Non-metallics polymers team performs polymeric materials screening, characterization, development, qualification, verification and validation testing for flight hardware and in-space applications. They perform insight, testing and application of polymeric materials for launch vehicles and space environments. Personnel have a deep understanding of the technologies associated with cryogenic insulation, high temperature thermal protection materials, paints, primers, coatings, and adhesive materials. Numerous TPS material characterization tests can also be performed by the Lab. The team has both internal and partner test capabilities to subject polymeric materials to Aerospace and Space environments Material properties, response and performance are tested within environments from the launch pad, through the launch environment and into space. A primary role of the team is to select materials and organize test regimen then analyze test results. Test regiment are established to confirm the materials meet the performance requirements levied by customers within the environments that are expected for the mission.

Thermal↗

NASA NTRS1992

Design definition of the Laser Atmospheric Wind Sounder (LAWS), phase 2. Volume 1: Executive summary

The LAWS phase 1 and phase 2 studies have been completed on schedule and have led to significant advances in CO2 laser development. The Phase 2 Design Definition Study has shown that a large scanning mirror/high pulse energy laser LAWS Instrument is feasible and within the existing technology. The capability to monitor wind velocities with backscatter ratios of 10(exp 11) m(exp -1) SR(exp -1) is feasible. The weight budget allocated for the baseline LAWS is adequate, and sufficient reserves exist with the potential downsized configuration. With the possible decrease in available power from the baseline of 2.2 kW guideline, power and shot management is critical for the baseline configuration (15 to 20 J). This is particularly true during the 100 day occultation period each year. With the downsized configurations (5 to 7 J), power management is still necessary during the occultation but is primarily limited to shot management over the polar regions. The breadboard effort has produced significant laser advances for a tight 18 month schedule and the minimum budgets available from NASA, Lockheed, and TDS. Using the NASA funds and Lockheed and TDS fixed assets budgets, the breadboard was designed, fabricated, and brought on-line with first laser light within 16 months after ATP. First laser beam was obtained on 21 April 1992 at a 5 J power level. Tests since then have been conducted at sustained, repetitive pulse levels of over 7 J and 20 Hz. This is an increase of over two or three times greater than any system previously developed from this type laser. Increased power levels and additional life tests will be accomplished in the next LAWS phase. The Lockheed LAWS design will operate in the gravity gradient mode on-orbit, and all possible instrument vibration and jitter modes have been considered. Adequate pointing stability and control is state-of-the-art technology for the critical time periods, frequency rates, and control responses required by LAWS. Lockheed recommends a 6-1/2 year phase C/D program for LAWS to provide adequate feedback from the engineering unit and the qualification unit to the final flight unit. Assuming a one year period for LAWS integration to the spacecraft, followed by a six-month period for launch vehicle integration, LAWS could be successfully developed and launched in eight years. Our baseline design or downsized design can be accommodated by either the Atlas 2AS or the Delta launch vehicles. Lockheed's recommendation is that, based on the successful phase 2 design study and breadboard program, a follow-on 18 month extended breadboard testing program and additional system engineering studies, primarily in interfacing with a to be defined platform, be initiated. This should be immediately followed by the phase C/D program, leading to a LAWS launch in late 2001 or early 2002.

Source record↗

NASA NTRS2003

Reliability and qualification of advanced microelectronics for space applications

This paper provides a discussion of the subject and an approach to establish a reliability and qualification methodology to facilitate the utilization of state-of-the-art advanced microelectronic devices and structures in high reliability applications.

advanced technologies reliability↗

NASA NTRS2006

Lightweight Ablative and Ceramic Thermal Protection System Materials for NASA Exploration Systems Vehicles

As a collaborative effort among NASA Centers, the "Lightweight Nonmetallic Thermal Protection Materials Technology" Project was set up to assist mission/vehicle design trade studies, to support risk reduction in thermal protection system (TPS) material selections, to facilitate vehicle mass optimization, and to aid development of human-rated TPS qualification and certification plans. Missions performing aerocapture, aerobraking, or direct aeroentry rely on advanced heatshields that allow reductions in spacecraft mass by minimizing propellant requirements. Information will be presented on candidate materials for such reentry approaches and on screening tests conducted (material property and space environmental effects tests) to evaluate viable candidates. Seventeen materials, in three classes (ablatives, tiles, and ceramic matrix composites), were studied. In additional to physical, mechanical, and thermal property tests, high heat flux laser tests and simulated-reentry oxidation tests were performed. Space environmental effects testing, which included exposures to electrons, atomic oxygen, and hypervelocity impacts, was also conducted.

Valentine, Peter G.↗

NASA NTRS1970

Backscatter Ultraviolet (BUV) Instrument

This final project report covers the design, development, construction, testing and qualification of the NIMBUS D BUV (Backscatter Ultraviolet) instrument. The instrument was designed and constructed in accordance with GSFC specification S-623-P-l and amendments. The design phase consisted of activity in three areas of technology: (1) Optical, (2) Mechanical, and (3) Electrical. Beckman Instruments, Inc. performed the tasks concerned with optical and mechanical design while Analog Technology Corporation was the electronics subcontractor. Initial studies resulted in the development of a breadboard model of the BUV instrument. Following extensive testing, additional mathematical analyses, and required modifications, preliminary and critical design reviews were performed. The results of these activities were the construction of an engineering model, used for extensive testing and demonstration of the final design features. Final design data resulted in the construction of two flight-rated instruments, serial numbers P103 and Fl04. On April 8, 1970, the P103 instrument was successfully launched into orbit on NIMBUS D and is presently fully operational.

Source record↗

NASA NTRSDate not supplied

Qualification of a Pulsed, Millinewton Class Metal Plasma Thruster for Broad Mission Applications

In the field of low power (<100W) electric propulsion, all thruster options demand significant trade-offs between operating parameters, reliability and scalable cost. With many systems on the market there are concerns with reliability, the need for extra considerations such as electron neutralizers for pure ion plumes, and the cost or craft compatibility of propellants. The Metal Plasma Thruster (MPT) is a new type of electric propulsion technology intended for low power applications. The system imparts momentum using inert, solid metal pucks as a propellant by using pulsed power to convert the metal into high velocity (~17km/s for Mo) jets of quasi-neutral plasma. The MPT technology does not require gas or liquid propellants, neutralizers, standby heaters, high voltage electronics, high electric or magnetic fields to operate. This comparatively simple pulsed operation is amenable to closed loop control, which provides for fine thrust control and S/C directed impulse on demand. Furthermore, the technology can use any metal as propellant, opening up unique opportunities for In-Situ Resource Utilization (ISRU) as well as customizability of performance for meeting specific mission needs. This paper describes implementation and direct measurement of this closed loop control mode as well as impulse measurement of multiple metals consistent with the aim of ISRU at NASA Glenn Research Center (GRC).

Electric Propulsion↗

NASA NTRSDate not supplied

Qualification of a Pulsed, Millinewton Class Metal Plasma Thruster for Broad Mission Applications

Electric Propulsion↗

NASA NTRS1991

Space nuclear power systems; Proceedings of the 8th Symposium, Albuquerque, NM, Jan. 6-10, 1991. Pts. 1-3

The present conference discusses NASA mission planning for space nuclear power, lunar mission design based on nuclear thermal rockets, inertial-electrostatic confinement fusion for space power, nuclear risk analysis of the Ulysses mission, the role of the interface in refractory metal alloy composites, an advanced thermionic reactor systems design code, and space high power nuclear-pumped lasers. Also discussed are exploration mission enhancements with power-beaming, power requirement estimates for a nuclear-powered manned Mars rover, SP-100 reactor design, safety, and testing, materials compatibility issues for fabric composite radiators, application of the enabler to nuclear electric propulsion, orbit-transfer with TOPAZ-type power sources, the thermoelectric properties of alloys, ruthenium silicide as a promising thermoelectric material, and innovative space-saving device for high-temperature piping systems. The second volume of this conference discusses engine concepts for nuclear electric propulsion, nuclear technologies for human exploration of the solar system, dynamic energy conversion, direct nuclear propulsion, thermionic conversion technology, reactor and power system control, thermal management, thermionic research, effects of radiation on electronics, heat-pipe technology, radioisotope power systems, and nuclear fuels for power reactors. The third volume discusses space power electronics, space nuclear fuels for propulsion reactors, power systems concepts, space power electronics systems, the use of artificial intelligence in space, flight qualifications and testing, microgravity two-phase flow, reactor manufacturing and processing, and space and environmental effects.

El-Genk, Mohamed S.↗

NASA NTRS2018

Stirling Convertor Controller Development at NASA Glenn Research Center

Over the past decade, NASA Glenn Research Center (GRC) has been supporting the development of Radioisotope Power Systems (RPS). NASA desires higher conversion efficiency RPS options that are reliable and robust with long life design. Dynamic conversion, such as Stirling and Brayton, offer the potential for higher conversion efficiencies but have yet to be demonstrated in a flight application. The RPS program sent out a solicitation to investigate options for dynamic conversion technologies. As a result of this solicitation, four dynamic power convertor (DPC) technologies were selected for design and fabrication of a prototype dynamic convertor. One lesson learned from the Advanced Stirling Radioisotope Generator (ASRG) project is that controller development should start early in the development of a dynamic convertor. As a result of this, NASA GRC has been utilizing hardware from past Stirling convertor projects including that of the ASRG to support controller development for the DPC's. NASA GRC has developed a strong knowledge base on both analog and digital Stirling dynamic power convertor controllers and will continue to expand and apply that knowledge to the four DPC's. Over the past 15 years, controllers were developed in-house at GRC, at Lockheed Martin Coherent Technologies (LMCT) and by the Johns Hopkins University/Applied Physics Laboratory (JHU/APL). Various generations of the controllers, have been developed as lessons were learned through various component and system level tests. Some of the tests performed were fault tolerance, qualification vibration level, electromagnetic interference, Radioisotope Power System Systems Integration (RSIL) tests, and extended operation. The fault tolerance test characterized the controller's ability to handle various fault conditions, including high or low bus power consumptions, total open load or short circuit, and replacing a failed controller card while the backup maintains control of the ASC. The vibration test confirms the controller's ability to control an ASC during launch. The EMI test characterized the AC and DC magnetic and electric fields emitted by the single ASC and if the controller has an impact on the radiated EMI. RSIL testing provided insight into the electrical interactions between the representative RPS, its associated control schemes, and realistic electric system loads. The extended operation test allows data to be collected over a period of thousands of hours to obtain long term performance data of the system. This paper describes the history of controller development at NASA GRC, tests performed on these controllers, and lessons learned.

Dugala, Gina M.↗

NASA NTRS1984

Ground test challenges in the development of the Space Shuttle orbiter auxiliary power unit

A conventional aircraft hydraulic system design approach was selected to provide fluid power for the Space Shuttle Orbiter. Developing the power unit, known as the Auxiliary Power Unit (APU), to drive the hydraulic pumps presented a major technological challenge. A small, high speed turbine drive unit powered by catalytically decomposed hydrazine and operating in the pulse mode was selected to meet the requirement. Because of limitations of vendor test facilities, significant portions of the development, flight qualification, and postflight anomaly testing of the Orbiter APU were accomplished at the Johnson Space Center (JSC) test facilities. This paper discusses the unique requirements of attitude, gravity forces, pressure profiles, and thermal environments which had to be satisfied by the APU, and presents the unique test facility and simulation techniques employed to meet the ground test requirements. In particular, the development of the zero-g lubrication system, the development of necessary APU thermal control techniques, the accomplishment of integrated systems tests, and the postflight investigation of the APU lube oil cooler behavior are discussed.

Chaffee, N. H.↗

NASA NTRS1984

Characterization of 8-cm engineering model thruster

Development of 8 cm ion thruster technology which was conducted in support of the Ion Auxiliary Propulsion System (IAPS) flight contract (Contract NAS3-21055) is discussed. The work included characterization of thruster performance, stability, and control; a study of the effects of cathode aging; environmental qualification testing; and cyclic lifetesting of especially critical thruster components.

Williamson, W. S.↗

NASA NTRS2002

Photovoltaic Engineering Testbed Designed for Calibrating Photovoltaic Devices in Space

Accurate prediction of the performance of solar arrays in space requires that the cells be tested in comparison with a space-flown standard. Recognizing that improvements in future solar cell technology will require an ever-increasing fidelity of standards, the Photovoltaics and Space Environment Branch at the NASA Glenn Research Center, in collaboration with the Ohio Aerospace Institute, designed a prototype facility to allow routine calibration, measurement, and qualification of solar cells on the International Space Station, and then the return of the cells to Earth for laboratory use. For solar cell testing, the Photovoltaic Engineering Testbed (PET) site provides a true air-mass-zero (AM0) solar spectrum. This allows solar cells to be accurately calibrated using the full spectrum of the Sun.

Landis, Geoffrey A.↗

NASA NTRSDate not supplied

Space-qualified Chalcogenide Materials through ISS MISSE Exposure —From Property to Application

In March 2021, twenty-four samples of various phase change materials (PCMs - Ge2Sb2Te5, Ge2Sb2Se4Te1, and Sb2Se3) along with metasurface optical components comprised of these PCMs, were delivered to the International Space Station (ISS) as part of the Materials International Space Station Experiment (MISSE-14) test campaign [1]. Although PCMs have previously been noted for their resilience to various forms of radiation [2], they had not been tested in a realistic space environment until this exposure campaign with joint NASA and MIT collaboration. During the six-month total open exposure time in low earth orbit (LEO), high-resolution cameras scanned and captured photographs of the samples to detect changes as a function of time along with on-orbit measured temperature, UV radiation, total atomic oxygen fluence, and total ionizing radiation doses. The samples were returned to NASA Langley Research Center in March 2022 for post-flight characterization. This duplicated the preflight characterization (i.e., material composition and crystallinity that limits switching speed, index contrast, loss, etc.) conducted before launch. The space sector has witnessed tremendous growth within the past decade—not only from government agencies but also entrants from the private sector. Future growth in the capabilities of Earth observation, deep space, and planetary surface missions using miniaturized spacecraft platforms can only be sustained by innovations in the design of remote sensors and other sub-systems. Active metasurface optics with enhanced tunability and reconfigurability continues to redefine the boundaries of optical science [3]. The introduction of PCM technology and associated optical devices will help to accelerate the adoption of new architectures for reduced size, weight, power, and cost (SWaP-C) platforms in space.Here we introduce results obtained from the MISSE-14 mission related to space qualification of PCM-based optic devices and constituent materials. We then discuss our recent work developing active integrated photonic devices and metasurface optics based on PCMs for space applications. This includes tunable and reconfigurable optical metasurface devices to support NASA space communication and LIDAR applications. PCMs are quickly becoming interesting photonics materials but questions related to mission suitability remain, particularly in regard to key properties like figures-of-merit (FOM, Δn/Δk), glass forming temperatures, and phase transition speeds. This talk will describe efforts to afford researchers the ability to have access to cost-effective data on exposure-induced changes to PCM fundamental physical and optical properties to assess their utility for space applications. The MISSE-14 sample exposure campaign allows a complete understanding of the limitations of the PCMs for various space-based electronic and optoelectronic applications.

Chalcogenide Phase change material↗

NASA NTRS2016

Experimental Durability Testing of 4H SiC JFET Integrated Circuit Technology at 727 Degrees Centigrade

We have reported SiC integrated circuits (ICs) with two levels of metal interconnect that have demonstrated prolonged operation for thousands of hours at their intended peak ambient operational temperature of 500 degrees Centigrade. However, it is recognized that testing of semiconductor microelectronics at temperatures above their designed operating envelope is vital to qualification. Towards this end, we previously reported operation of a 4H-SiC JFET IC ring oscillator on an initial fast thermal ramp test through 727 degrees Centigrade. However, this thermal ramp was not ended until a peak temperature of 880 degrees Centigrade (well beyond failure) was attained. Further experiments are necessary to better understand failure mechanisms and upper temperature limit of this extreme-temperature capable 4H-SiC IC technology.Here we report on additional experimental testing of custom-packaged 4H-SiC JFET IC devices at temperatures above 500 degrees Centigrade. In one test, the temperature was ramped and then held at 727 degrees Centigrade, and the devices were periodically measured until electrical failure was observed. A 4H-SiC JFET on this chip electrically functioned with little change for around 25 hours at 727 degrees Centigrade before rapid increases in device resistance caused failure. In a second test, devices from our next generation 4H-SiC JFET ICs were ramped up and then held at 700 degrees Centigrade (which is below the maximum deposition temperature of the dielectrics). Three ring oscillators functioned for 8 hours at this temperature before degradation. In a third experiment, an alternative die attach of gold paste and package lid was used, and logic circuit operation was demonstrated for 143.5 hours at 700 degrees Centigrade.

Silicon Carbide↗

NASA NTRS2011

Base Flow Investigation of the Apollo AS-202 Command Module

In recent years, both Europe and the US are developing hypersonic research and operational vehicles. These include (re)entry capsules (both ballistic and lifting) and lifting bodies such as ExoMars, EXPERT, ARV, CEV and IXV. The research programs are meant to enable technology and engineering capabilities to support during the next decade the development of affordable (possibly reusable) space transportation systems as well as hypersonic weapons systems for time critical targets. These programs have a broad range of goals, ranging from the qualification of thermal protection systems, the assessment of RCS performances, the development of GNC algorithms, to the full demonstration of the performance and operability of the integrated vehicles. Since the aerothermodynamic characteristics influence nearly all elements of the vehicle design, the accurate prediction of the aerothermal environment is a prerequisite for the design of efficient hypersonic systems. Significant uncertainties in the prediction of the hypersonic aerodynamic and the aerothermal loads can lead to conservative margins in the design of the vehicle including its Outer Mould Line (OML), thermal protection system, structure, and required control system robustness. The current level of aerothermal prediction uncertainties results therefore in reduced vehicle performances (e.g., sub-optimal payload to mass ratio, increased operational constraints). On the other hand, present computational capabilities enable the simulation of three dimensional flow fields with complex thermo-chemical models over complete trajectories and ease the validation of these tools by, e.g., reconstruction of detailed wind tunnel tests performed under identified and controlled conditions (flow properties and vehicle attitude in particular). These controlled conditions are typically difficult to achieve when performing in flight measurements which in turn results in large associated measurement uncertainties. Similar problems arise when attempting to rebuild measurements performed in "hot" ground facilities, where the difficulty level is increased by the addition of the free-flow characterization itself. The implementation of ever more sophisticated thermochemical models is no obvious cure to the aforementioned problems since their effect is often overwhelmed by the large measurement uncertainties incurred in both flight and ground high enthalpy facilities. Concurrent to the previous considerations, a major contributor to the overall vehicle mass of re-entry vehicles is the afterbody thermal protection system. This is due to the large acreage (equal or bigger than that of the forebody) to be protected. The present predictive capabilities for base flows are comparatively lower than those for windward flowfields and offer therefore a substantial potential for improving the design of future re-entry vehicles. To that end, it is essential to address the accuracy of high fidelity CFD tools exercised in the US and EU, which motivates a thorough investigation of the present status of hypersonic flight afterbody heating. This paper addresses the predictive capabilities of after body flow fields of re-entry vehicles investigated in the frame of the NATO/RTO - RTG-043 Task Group and is structured as follows: First, the verification of base flow topologies on the basis of available wind-tunnel results performed under controlled supersonic conditions (i.e., cold flows devoid of reactive effects) is performed. Such tests address the detailed characterization of the base flow with particular emphasis on separation/reattachment and their relation to Mach number effects. The tests have been performed on an Apollo-like re-entry capsule configuration. Second, the tools validated in the frame of the previous effort are exercised and appraised against flight-test data collected during the Apollo AS-202 re-entry.

Walpot, Louis M. G.↗

NASA NTRS2019

Transition of OSMA AM Responsibility from WSTF to MSFC

Additive manufacturing (AM) is a rapidly evolving, game changing technology requiring many disciplines (including NDE) to ensure mission safety and reliability. While AM is becoming increasingly democratized and pervasive, managing risks for fracture critical L-PBF spaceflight hardware must clearly come from MSFC. Therefore OSMA roles for NDE qualification space flight hardware are being transitioned to MSFC.

Waller, Jess↗

NASA NTRS2019

13kW Advanced Electric Propulsion Flight System Development and Qualification

The next phase of robotic and human deep space exploration missions requires high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdyne's Advanced Electric Propulsion System (AEPS) program is completing development and qualification of a 13kW flight EP system to support NASA exploration. The first use of the AEPS is planned for the NASA Power & Propulsion Element, which is the first element of NASA's cis-lunar Gateway. The flight AEPS system includes a magnetically shielded long-life Hall thruster, power processing unit (PPU), and xenon flow controller (XFC). The Hall thruster, originally developed and demonstrated by NASA's Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 13.3kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation beyond the orbit of Mars. The integrated system is continuously throttleable between 3 and 13.3kW. The program has completed testing of the Technology Development Units and is progressing into the Engineering Development Unit test phase and the final design phase to Critical Design Review (CDR). This paper will present the high power AEPS system capabilities, overall program and design status and the latest test results for the 13kW flight system development as well as the plans for the development and qualification effort of the EP string.

in-space propulsion↗