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This poster highlights the Game Changing Development Program task, Space Act Agreement between Maxar Technologies and NASA to test 5-junction solar cells on roll out compatible blankets through qualification protocols similar to AIAA Standards S-111/112. A description of the planned work and current status is included.
Two special instruments were custom built for our critical study and manufactured and employed to test the SA materials of interest to cryogenic temperature conditions. We have procured custom manufactured two independent systems such as Thermo -Mechanical Analyzer (TMA) with Three Point Bend Probe (TPBP) module and Dilatometer. TMA can be used to measure Young’ s modulus or elastic modulus (e -modulus). Dilatometer can be used to measure Co -efficient of Thermal Expansion (CTE). We have optimized the materials configuration to measure physical properties of radiated and non -radiated materials for CTE and e -modulus, which will be useful to assess materials and processes reliability for cryogenic temperature applications. We have successfully implemented for some materials that are of interest to solar panels for the planned ECM project to measure CTE and Young's modulus. Also a separate study was undertaken to thermal cycle qualify SAs for ECM project. One can do thermal cycling to temperatures of 50 K and 133 K using liquid Helium (Lq He). This is in reality exorbitantly too expensive and is also inefficient. Furthermore, this is not technically controllable test to 133 K and 50 K temperatures. This is also too expensive to accomplish. Therefore, a vacuum set -up was made along with a cryostat successfully. This is a closed loop system where we do not lose the Helium gas. We were able to get to 35 K for a given test coupon size. These temperatures are lower than what we would intend to qualify the solar array technologies. This will allow to have some margin at cold temperatures. We successfully completed the qualification of solar array technologies down to 50 K temperatures. This is the first time we are reporting this accomplishment. In this paper, we will present the solar array materials properties for cryogenic temperatures down to 20 K temperatures and the thermal cycling qualification test results from 133 K to 50 K for three times the mission life of 120 thermal cycles to meet JPL Design.
The design of the facility is described in terms of its use as an investigation tool for evaluating crop growth in space with reference to required emerging technologies. NASA's CELSS Test Facility (CTF) is designed to permit the measurement of crop-plant productivity under microgravity conditions including biomass production, food production, water transpiration, and O2/CO2 exchanges. Crucial hardware tests and qualifications are identified to assure the operation of CTF technologies in space including the nutrient-delivery, water-condensation, and gas-liquid-mixing subsystems. The design concept and related scientific requirements are described and shown to provide microgravity crop research. The CTF is expected to provide data for plant research and for concepts for bioregenerative life-support systems for applications to Martian, lunar, and space-station missions.
NASA s Evolutionary Xenon Thruster (NEXT) project is developing next generation ion propulsion technologies to provide future NASA science missions with enhanced mission performance benefit at a low total development cost. The objective of the NEXT project is to advance next generation ion propulsion technology by producing engineering model system components, validating these through qualification-level and integrated system testing, and ensuring preparedness for transitioning to flight system development. As NASA s Evolutionary Xenon Thruster technology program completes advanced development activities, it is advantageous to review the existing technology capabilities of the system under development. This paper describes the NEXT ion propulsion system development status, characteristics and performance. A review of mission analyses results conducted to date using the NEXT system is also provided.
Surface mount technology feedthrough ceramic capacitors are gaining popularity in Hi-Rel systems due to small size and great filtering characteristics. In this work, BME feedthrough capacitors compliant with AEC-Q200 requirements have been evaluated, screened and qualified for a space project. Evaluation included analysis of the effect of post-soldering thermal shocks, temperature dependencies of leakage currents, distribution of breakdown voltages, assessments of high-current capability, and thermal resistance of the parts. The parts have been screened and qualified at conditions close to the military requirements for ceramic capacitors. No failures were detected during screening including burning-in, but two samples failed during monitored 2000 hour life testing at 125C and two times rated voltage. Failure analysis revealed manufacturing defects that required additional analysis of screening and qualification test conditions. Highly accelerated life testing (HALT) has been carried out to assess reliability acceleration factors and the probability of failure at the use conditions. An approach to selection of adequate burning-in and life test conditions is discussed.
Surface mount technology feedthrough ceramic capacitors are gaining popularity in Hi-Rel systems due to small size and great filtering characteristics. In this work, BME feedthrough capacitors compliant with AEC-Q200 requirements have been evaluated, screened and qualified for a space project. Evaluation included analysis of the effect of post-soldering thermal shocks, temperature dependencies of leakage currents, distribution of breakdown voltages, assessments of high-current capability, and thermal resistance of the parts. The parts have been screened and qualified at conditions close to the military requirements for ceramic capacitors. No failures were detected during screening including burning-in, but two samples failed during monitored 2000 hour life testing at 125C and two times rated voltage. Failure analysis revealed manufacturing defects that required additional analysis of screening and qualification test conditions. Highly accelerated life testing (HALT) has been carried out to assess reliability acceleration factors and the probability of failure at the use conditions. An approach to selection of adequate burning-in and life test conditions is discussed.
The Department of Energy, Lockheed Martin (LM), Stirling Technology Company, and NASA Glenn Research Center (GRC) are developing a high-efficiency Stirling Radioisotope Generator (SRG) for potential NASA Space Science missions. The SRG is being developed for multimission use, including providing spacecraft onboard electric power for NASA deep space missions and power for unmanned Mars rovers. NASA GRC is conducting an in- house supporting technology project to assist in developing the Stirling convertor for space qualification and mission implementation. Preparations are underway for a thermalhacuum system demonstration and unattended operation during endurance testing of the 55-We Technology Demonstration Convertors. Heater head life assessment efforts continue, including verification of the heater head brazing and heat treatment schedules and evaluation of any potential regenerator oxidation. Long-term magnet aging tests are continuing to characterize any possible aging in the strength or demagnetization resistance of the permanent magnets used in the linear alternator. Testing of the magnet/lamination epoxy bond for performance and lifetime characteristics is now underway. These efforts are expected to provide key inputs as the system integrator, LM, begins system development of the SRG. GRC is also developing advanced technology for Stirling convertors. Cleveland State University (CSU) is progressing toward a multi-dimensional Stirling computational fluid dynamics code, capable of modeling complete convertors. Validation efforts at both CSU and the University of Minnesota are complementing the code development. New efforts have been started this year on a lightweight convertor, advanced controllers, high-temperature materials, and an end-to-end system dynamics model. Performance and mass improvement goals have been established for second- and third-generation Stirling radioisotope power systems.
The Astro Edge solar array is a new and innovative reflective low concentrator power generating system which has been selected for the CTA Incorporate/Lockheed Martin Clark spacecraft under the NASA Small Spacecraft Technology Initiative (SSTI) program. In support of this program, Astro Aerospace Corporation has produced one qualification and two flight solar array wings to support a July 1996 launch. The Astro Edge solar array was selected as a new technology to benefit future NASA, military and commercial missions by providing high specific power, high deployed stiffness, low stowed volume, low risk, and cost reduction features which meet the agency's 'better, faster, cheaper' goals. This novel array accounts for five of the thirty-six advanced technologies which the Clark spacecraft will demonstrate. A brief SSTI Astro Edge solar array program overview is presented. Completed qualification and acceptance testing is discussed. Finally, the major discriminators which make the Astro Edge solar array 'better, faster, cheaper' technology are provided.
Six liquid hydrogen feedline design concepts were developed for the cryogenic space tug. The feedlines include composite and all-metal vacuum jacketed and nonvacuum jacketed concepts, and incorporate the latest technological developments in the areas of thermally efficient vacuum jacket end closures and standoffs, radiation shields in the vacuum annulus, thermal coatings, and lightweight dissimilar metal flanged joints. The feedline design concepts are evaluated on the basis of thermal performance, weight, cost, reliability, and reusability. Design concepts were proved in a subscale test program. Detail design was completed on the most promising composite feedline concept and an all-metal feedline. Three full scale curved composite feedlines and one all-metal feedline assembly were fabricated and subjected to a test program representative of flight hardware qualification. The test results show that composite feedline technology is fully developed. Composite feedlines are ready for space vehicle application and offer significant reduction in weights over the conventional all-metal feedlines presently used.
Beginning in 2014, the National Aeronautics and Space Administration (NASA) Space Technology Mission Directorate (STMD) began a project to increase the state of the art for the Hall-Effect Solar Electric Propulsion (SEP) technology. The Technology Demonstration Missions (TDM) Program Office provides programmatic oversight of SEP with the project being led by Glenn Research Center, supported by the Jet Propulsion Laboratory and development, qualification & flight hardware all provided by Aerojet Rocketdyne (AR). This technology was pursued as the propulsion system for the Asteroid Redirect and Robotic Mission (ARRM). While the concept was originally slated for ARRM, it was realigned to Gateway to support future Moon to Mars objectives. The Gateway lunar station was established and will play a key role in NASA’s Artemis Program which utilizes collaboration with the Canadian Space Agency (CSA), the European Space Agency (ESA) and the Japanese Space Agency (JAXA). The 12-kW hardware will be the primary propulsion for the Gateway element, Power & Propulsion Element (PPE), from Low Earth Orbit to a Near Rectilinear Halo Orbit around the Moon. Initial efforts began with utilizing Technology Development Units (TDU) built by NASA as the predecessor to the contract activity with AR. Over the past few years, AR has built Engineering Test Units (ETU), Engineering Development Units (EDU) and qualification & flight models (QM and FM, respectively). With requirement changes from the AARM mission to PPE, the joint team has modified the design and has started qualification activities for environment testing (shock, vibe) at multiple facilities in the US leading to eventual life testing. This paper will explore the various design changes, system modeling and the latest status for qualification testing.
An integrated, generic unibody composite pressurized structure (UCPS) combined with a positive expulsion device (PED), consisting of an elastomeric bladder for monopropellant hydrazine, has been quasi-standardized for spacecraft use. The combination functions as an all-composite, non-metallic, propellant tank with bladder. The integrated UCPS combines several previous innovations - specifically, the linerless, all-composite cryogenic tank technology; all-composite boss; resin formulation; and integrated stringer system. The innovation combines the UCPS with an integrated propellant management device (PMD), the PED or bladder, to create an entirely unique system for in-space use. The UCPS is a pressure vessel that incorporates skirts, stringers, and other structures so that it is both an in-space hydrazine tank, and also a structural support system for a spacecraft in a single, all-composite unit. This innovation builds on the progress in the development of a previous SBIR (Small Business Innovation Research) Phase I with Glenn Research Center and an SBIR III with Johnson Space Center that included the fabrication of two 42-in. (≈107-cm) diameter all-composite cryogenic (LOX and liquid methane) UCPS test tanks for a lunar lander. This Phase II provides hydra zine compatibility testing of the elastomeric bladder, a see-through PED to validate the expulsion process and model, and a complete UCPS-based PED with stringers and skirts that will be used to conduct initial qualification and expulsion tests. This extends the UCPS technology to include hydrazine-based, in-space pro - pulsion applications and can also be used for electric propulsion. This innovation creates a system that, in comparison to the traditional approach, is lower in weight, cost, volume, and production time; is stronger; and is capable of much higher pressures. It also has fewer failure modes, and is applicable to both chemical and electric propulsion systems.
Beginning in 2014, the National Aeronautics and Space Administration (NASA) Space Technology Mission Directorate (STMD) began a project to increase the state of the art for the Hall-Effect Solar Electric Propulsion (SEP) technology. The Technology Demonstration Missions (TDM) Program Office provides programmatic oversight of SEP with the project being led by Glenn Research Center, supported by the Jet Propulsion Laboratory and development, qualification & flight hardware all provided by Aerojet Rocketdyne (AR). This technology was pursued as the propulsion system for the Asteroid Redirect and Robotic Mission (ARRM). While the concept was originally slated for ARRM, it was realigned to Gateway to support future Moon to Mars objectives. The Gateway lunar station was established and will play a key role in NASA’s Artemis Program which utilizes collaboration with the Canadian Space Agency (CSA), the European Space Agency (ESA) and the Japanese Space Agency (JAXA). The 12-kW hardware will be the primary propulsion for the Gateway element, Power & Propulsion Element (PPE), from Low Earth Orbit to a Near Rectilinear Halo Orbit around the Moon. Initial efforts began with utilizing Technology Development Units (TDU) built by NASA as the predecessor to the contract activity with AR. Over the past few years, AR has built Engineering Test Units (ETU), Engineering Development Units (EDU) and qualification & flight models (QM and FM, respectively). With requirement changes from the AARM mission to PPE, the joint team has modified the design and has started qualification activities for environment testing (shock, vibe) at multiple facilities in the US leading to eventual life testing. This paper will explore the various design changes, system modeling and the latest status for qualification testing.
The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system has been in advanced technology development under the NASA In-Space Propulsion Technology project. The highest fidelity hardware planned has now been completed by the government/industry team, including a flight prototype model (PM) thruster, an engineering model (EM) power processing unit, EM propellant management assemblies, a breadboard gimbal, and control unit simulators. Subsystem and system level technology validation testing is in progress. To achieve the objective Technology Readiness Level 6, environmental testing is being conducted to qualification levels in ground facilities simulating the space environment. Additional tests have been conducted to characterize the performance range and life capability of the NEXT thruster. This paper presents the status and results of technology validation testing accomplished to date, the validated subsystem and system capabilities, and the plans for completion of this phase of NEXT development.
The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system has been in advanced technology development under the NASA In-Space Propulsion Technology project. The highest fidelity hardware planned has now been completed by the government/industry team, including: a flight prototype model (PM) thruster, an engineering model (EM) power processing unit, EM propellant management assemblies, a breadboard gimbal, and control unit simulators. Subsystem and system level technology validation testing is in progress. To achieve the objective Technology Readiness Level 6, environmental testing is being conducted to qualification levels in ground facilities simulating the space environment. Additional tests have been conducted to characterize the performance range and life capability of the NEXT thruster. This paper presents the status and results of technology validation testing accomplished to date, the validated subsystem and system capabilities, and the plans for completion of this phase of NEXT development. The next round of competed planetary science mission announcements of opportunity, and directed mission decisions, are anticipated to occur in 2008 and 2009. Progress to date, and the success of on-going technology validation, indicate that the NEXT ion propulsion system will be a primary candidate for mission consideration in these upcoming opportunities.
The Department of Energy (DOE), Germantown, Maryland, Stirling Technology Company (STC), Kennewick, Washington, and NASA Glenn Research Center are developing a free-piston Stirling convertor for a high-efficiency Stirling Radioisotope Generator for NASA Space Science missions. This generator is being developed for multimission use, including providing electric power for unmanned Mars rovers and for deep space missions. STC is developing the 55-W Technology Demonstration Convertor (TDC) under contract to DOE. Glenn is conducting an in-house technology project to assist in developing the convertor for readiness for space qualification and mission implementation. As part of this effort, a Stirling Research Laboratory was established to test the TDC's and related technologies. A key task is providing an independent verification and validation of the TDC performance. Four TDC's are now being tested at Glenn. Acceptance testing has been completed for all convertors, and in general, performance agreed well with that achieved by STC prior to the delivery of the convertors. Performance mapping has also been completed on two of the convertors over a range of hot-end temperatures (450 to 650 C), cold-end temperatures (80 to 120 C), and piston amplitudes (5.2 to 6.2 mm). These test data are available online at http://www.grc.nasa.gov/WWW/tmsb/. The TDC's can be tested in either a horizontal orientation with dual-opposed convertors or in a vertical orientation with a single convertor. Synchronized dual-opposed pairs are used for dynamically balanced operation that results in very low levels of vibration. The Stirling Research Laboratory also supports launch environment testing of the TDC's in Glenn's Structural Dynamics Laboratory and electromagnetic interference and electromagnetic compatibility characterization and reduction efforts. In addition, the TDC's will be used for long-term endurance testing, and preparations are underway for unattended operation.
In this presentation, the premise that Space Station Freedom has great utility as an engineering experiment station will be explored. There are several modes in which it can be used for this purpose. The most obvious are space qualification, process development, in space satellite repair, and materials engineering. The range of engineering experiments which can be done at Space Station Freedom run the gamut from small process oriented experiments to full exploratory development models. A sampling of typical engineering experiments are discussed in this session. First and foremost, Space Station Freedom is an elaborate experiment itself, which, if properly instrumented, will provide engineering guidelines for even larger structures which must surely be built if humankind is truly 'outward bound.' Secondly, there is the test, evaluation and space qualification of advanced electric thruster concepts, advanced power technology and protective coatings which must of necessity be tested in the vacuum of space. The current approach to testing these technologies is to do exhaustive laboratory simulation followed by shuttle or unmanned flights. Third, the advanced development models of life support systems intended for future space stations, manned mars missions, and lunar colonies can be tested for operation in a low gravity environment. Fourth, it will be necessary to develop new protective coatings, establish construction techniques, evaluate new materials to be used in the upgrading and repair of Space Station Freedom. Finally, the industrial sector, if it is ever to build facilities for the production of commercial products, must have all the engineering aspects of the process evaluated in space prior to a commitment to such a facility.
A primary focus of the Entry Systems and Technology Division at NASA Ames is design, development, qualification and certification of Thermal Protection Systems for current NASA missions. Another primary focus is the development of new thermal protection systems for upcoming missions that address shortfalls in the existing suite of TPS. Examples of such shortfalls include performance at higher capability and reduction in mass. NASA is also investing in TPS sustainability ensuring the long term availability of TPS solutions for future missions. The specific TPS selection, for a given mission , depends on a number of parameters including the missions risk posture. For all missions the goal for TPS is efficient and reliable performance and to achieve these goals an understanding of the materials (composition and architecture) is required for proper design and use of the chosen TPS. Analytic tools are used to inform on a material (systems) response to a given environment and the response itself depends on the materials properties which are driven by its composition and architecture. This presentation will review the different generic ablative TPS architectures and anticipated corresponding failure modes.
The modal response of a liquid-filled tank to external acoustic excitation can be used to infer with high resolutionthe mass of contained liquid, the mass flow rate of liquids into and out of the tank, and changes in tank pressure.Both contained liquid mass and internal ullage pressure affect the modal response of the tank walls through fluidmass-loading of the tank walls and pressure-induced wall stiffening, respectively. Modal Propellant Gauging refersto the technology that exploits these shifts in modal frequencies to infer the mass of propellant in a tank. MPG is anon-invasive gauging technology that has demonstrated gauging resolutions of 1% for settled propellants and 2-3%for unsettled, sloshing propellants. Extensive parabolic flight testing of the MPG system on model tanks has beenconducted to validate the technology in reduced gravity. MPG testing on a qualification tank for the Orion Program'sEuropean Service Module has also been conducted and is reported here. Finite element modeling of the OrionESM "upper" tank is discussed and compared with measurement data. Three computational approaches to massdetermination, Peak Tracking, Point Sensor, and Spectral Density methods, are described here. Use cases are definedand analyzed in the context of the Orion ESM Qualification tank data, and an implementation scheme for continuousmass gauging on the Orion ESM is discussed.Keywords: Propellant Mass Gauging, Modal Analysis, Orion ESM