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All sourcesNASA NTRSDOE OSTI
NASA NTRS2021

Long-Duration Wear Testing of the ASTRAEUS Hall Thruster, Phases II & III: 72 kg Xe Throughput & 1.2 MN-s Total Impulse

The thruster component of the Ascendant Sub-kW Transcelestial Electric Propulsion System successfully completed a propellant throughput demonstration of 72 kg Xe in a long-duration wear test planned to eventually exceed 100 kg. The thruster has been operated for a total of 5450 h over three operating conditions: 3287 h at 300 V – 1000 W, 1745 h at 500 V – 1000 W, and 417 h at 400 V – 600 W. A total of 1.2 MN-s of impulse has been demonstrated to date. Thrust, telemetry, and temperature measurements periodically recorded at four operating conditions spanning 200 – 500 V and 200 – 1000 W revealed constant performance across the full test duration. The erosion of the downstream faces of the thruster was observed to be higher than anticipated after the first 5032 h of thruster operation. This issue was traced to the application of sub-nominal magnetic field strengths throughout the LDWT, caused by an electrical-related ground-support equipment issue, which resulted in a downstream shift of the discharge plasma and enhanced erosion of the thruster’s pole covers. Nevertheless, the thruster performance and operational behaviour remained unaffected. High-fidelity plasma simulations of the erosion of the thruster’s downstream faces, supported by prior short-duration wear testing, suggest a theoretical lifetime capability of >30 kh, corresponding to propellant throughput of >450 kg Xe and a total impulse of >7 MNs, at the 300 V – 1000 W operating condition. The test was voluntarily paused after 68 kg Xe and 72 kg Xe throughput, respectively, to take these erosion measurements at the 500 V and 400 V operating conditions before the test was resumed at a different operating condition.

Goebel, Dan↗
NASA NTRS1997

2 kWe Solar Dynamic Ground Test Demonstration Project: Design Report - Volume 2

Critical Design Reviews (CDR's) were held on the Solar Dynamic Ground Test Demonstrator (SDGTD). This CDR summary report will provide the following information for each of the system components and the system integration: (1) A bibliography of design/design review documentation; (2) A summary of the major discussion issues from issues from each design review; (3) A definition of the component and system detail designs along with the bottom line from the supporting analysis; (4) Status and key results from pertinent development activities on-going in the CDR time period; (5) A brief description of planned testing; and (6) A discussion of issues stiff open at the completion of CDR. Appendix 1 to this report contains a listing and status (as of 28 June 1993) of all the action items generated during all SDGTD CDRs. The reader should remember that the SDGTD program is being conducted in an open communication forum, and program participants are encouraged to ask questions or request information. Team members are allowed and encouraged to participate in the reviews on an equal basis. No request for information, as long as it is within the work scope, is refused, so many action items are generated.

Alexander, Dennis↗
NASA NTRS1987

Orbit transfer rocket engine technology program. Phase 2: Advanced engine study

In Phase 2 of the Advanced Engine Study, the Failure Modes and Effects Analysis (FMEA) maintenance-driven engine design, preliminary maintenance plan, and concept for space operable disconnects generated in Phase 1 were further developed. Based on the results of the vehicle contractors Orbit Transfer Vehicle (OTV) Concept Definition and System Analysis Phase A studies, minor revisions to the engine design were made. Additional refinements in the engine design were identified through further engine concept studies. These included an updated engine balance incorporating experimental heat transfer data from the Enhanced Heat Load Thrust Chamber Study and a Rao optimum nozzle contour. The preliminary maintenance plan of Phase 1 was further developed through additional studies. These included a compilation of critical component lives and life limiters and a review of the Space Shuttle Main Engine (SSME) operations and maintenance manual in order to begin outlining the overall maintenance procedures for the Orbit Transfer Vehicle Engine and identifying technology requirements for streamlining space-based operations. Phase 2 efforts also provided further definition to the advanced fluid coupling devices including the selection and preliminary design of a preferred concept and a preliminary test plan for its further development.

Erickson, C.↗
NASA NTRS2012

Performance Assessment of Baseline Cells for the High Efficiency Space Power Systems Project

The Enabling Technology Development and Demonstration (ETDD) Program High Efficiency Space Power Systems (HESPS) Project, formerly the Exploration Technology Development Program (ETDP) Energy Storage Project is tasked with developing advanced lithium-ion cells for future NASA Exploration missions. Under this project, components under development via various in-house and contracted efforts are delivered to Saft America for scale-up and integration into cells. Progress toward meeting project goals will be measured by comparing the performance to these cells with cells of a similar format with Saft s state-of-the-art aerospace chemistry. This report discusses the results of testing performed on the first set of baseline cells delivered by Saft to the NASA Glenn Research Center. This build is a cylindrical "DD" geometry with a 10 Ah nameplate capacity. Testing is being performed to establish baseline cell performance at conditions relevant to ETDD HESPS Battery Key Performance Parameter (KPP) goals including various temperatures, rates, and cycle life conditions. Data obtained from these cells will serve as a performance baseline for future cell builds containing optimized ETDD HESPSdeveloped materials. A test plan for these cells was developed to measure cell performance against the high energy cell KPP goals. The goal for cell-level specific energy of the high energy technology is 180 Wh/kg at a C/10 discharge rate and 0 C. The cells should operate for at least 2000 cycles at 100 percent DOD with 80 percent capacity retention. Baseline DD cells delivered 152 Wh/kg at 20 C. This number decreased to 143.9 Wh/kg with a 0 C discharge. This report provides performance data and summarizes results of the testing performed on the DD cells.

Schneidegger, Brianne T.↗
NASA NTRS2004

Development of an Experiment High Performance Nozzle Research Program

As proposed in the above OAI/NASA Glenn Research Center (GRC) Co-Operative Agreement the objective of the work was to provide consultation and assistance to the NASA GRC GTX Rocket Based Combined Cycle (RBCC) Program Team in planning and developing requirements, scale model concepts, and plans for an experimental nozzle research program. The GTX was one of the launch vehicle concepts being studied as a possible future replacement for the aging NASA Space Shuttle, and was one RBCC element in the ongoing NASA Access to Space R&D Program (Reference 1). The ultimate program objective was the development of an appropriate experimental research program to evaluate and validate proposed nozzle concepts, and thereby result in the optimization of a high performance nozzle for the GTX launch vehicle. Included in this task were the identification of appropriate existing test facilities, development of requirements for new non-existent test rigs and fixtures, develop scale nozzle model concepts, and propose corresponding test plans. Also included were the evaluation of originally proposed and alternate nozzle designs (in-house and contractor), evaluation of Computational Fluid Dynamics (CFD) study results, and make recommendations for geometric changes to result in improved nozzle thrust coefficient performance (Cfg).

Source record↗
NASA NTRS2018

Estimating Software Reliability for Space Launch Vehicles in Probabilistic Risk Assessment (PRA)

It is acutely recognized in the Probabilistic Risk assessment (PRA) field that software plays a defining role in overall system reliability for all modern systems across a wide variety of industries. Regardless if the software is embedded firmware for working components or elements, part of a Human-Machine-Interface, or automated command and control logic, the success of the software to fulfill its function under nominal and off-nominal environments will be a dominant contributor to system reliability. It is also recognized that software reliability prediction and estimation is one of the more challenging and questionable aspects of any PRA or system analyses due to the nature of software and its integration with physics based systems. Irrespective of this dichotomy, any incorporation of software reliability methods requires that the contributions are accountable, quantitative, and tractable. This paper provides a brief overview of software reliability methods, establishes some minimum requirements that the methods should incorporate for completeness, and provides a logic structure for applying software reliability. Model resolution will be discussed that supports current testing plans and trade studies. We will provide initial recommendations for use in the NASA PRA and present a future dynamic option for software and PRA. Space Launch Vehicle Software is recognized to be reliable in static conditions, yet relatively vulnerable to a set of failure modes in changing environments/flight phases. Two quantitative methods were chosen to incorporate software reliability into a Space Launch Vehicle PRA accounting for phase adjustments. One method predicts latent software failure using statistical methods, and the second provides estimates of coding errors and software operating system failures based on test and historical data, respectively. Software uncertainty will also be discussed. We determined that recommendations for PRA software reliability should be modeled at the software module level where multiple software components compose a module and combinations of the software architecture can lead to a functional failure.

Novack, Steven↗
NASA NTRS2013

Flight Test Results of an Axisymmetric Channeled Center Body Supersonic Inlet at Off-Design Conditions

Flight-testing of a channeled center-body axisymmetric supersonic inlet design concept was conducted at the National Aeronautics and Space Administration (NASA) Dryden Flight Research Center in collaboration with the NASA Glenn Research Center (Cleveland, Ohio) and TechLand Research, Inc. (North Olmsted, Ohio). This testing utilized the Propulsion Flight Test Fixture, flown on the NASA F-15B research test bed airplane (NASA tail number 836) at local experiment Mach numbers up to 1.50. The translating channeled center-body inlet was designed by TechLand Research, Inc. (U.S. Patent No. 6,276,632 B1) to allow for a novel method of off-design flow matching, with original test planning conducted under a NASA Small Business Innovative Research study. Data were collected in flight at various off-design Mach numbers for fixed-geometry representations of both the channeled center-body design and an equivalent area smooth center-body design for direct comparison of total pressure recovery and limited distortion measurements.

inlet↗
NASA NTRS2018

Estimating Software Reliability for Space Launch Vehicles in Probabilistic Risk Assessment (PRA)

It is acutely recognized in the Probabilistic Risk Assessment (PRA) field that software plays a defining role in overall system reliability for all modern systems across a wide variety of industries. Regardless of whether the software is embedded firmware for working components or elements, part of a Human-Machine-Interface, or automated command and control logic, the success of the software to fulfill its function under nominal and off-nominal environments will be a dominant contributor to system reliability. It is also recognized that software reliability prediction and estimation is one of the more challenging and questionable aspects of any PRA or system analyses due to the nature of software and its integration with physics based systems. Irrespective of this dichotomy, any incorporation of software reliability methods requires that the contributions are accountable, quantitative, and tractable. This paper provides a brief overview of software reliability methods, establishes some minimum requirements that the methods should incorporate for completeness, and provides a logic structure for applying software reliability. Model resolution will be discussed that supports current testing plans and trade studies. We will provide initial recommendations for use in the National Aeronautics and Space Administration (NASA) PRA and present a future dynamic option for software and PRA. Space Launch Vehicle software is recognized to be reliable in static conditions, yet relatively vulnerable to a set of failure modes in changing environments/flight phases. Two quantitative methods were chosen to incorporate software reliability into a Space Launch Vehicle PRA accounting for phase adjustments. One method predicts latent software failure using statistical methods, and the second provides estimates of coding errors and software operating system failures based on test and historical data. Software uncertainty will also be discussed. It is determined that recommendations for PRA software reliability should be modeled at the software module level where multiple software components compose a module and combinations of the software architecture can lead to a functional failure.

Steven D. Novack↗
NASA NTRS1991

Subscale LOX/hydrogen testing with a modular chamber and a swirl coaxial injector

To support basic injector and chamber technology, the response of a modular calorimeter chamber to an injector with swirl coaxial elements was investigated. The tests supported both the Advanced Launch System (ALS) and the Space Shuttle Main Engine programs. Original test plans included chamber pressures up to 2250 psia, the current ALS engine criteria, but hardware limitations forced conditions to be reduced. With four different chamber configurations, a variety of data was obtained with chamber pressures ranging from 1483 psia to 1679 psia and mixture ratios from 5.24 to 6.90. The swirl coaxial injector showed good chamber wall compatibility, despite misaligned elements, and its high performance was independent of fuel. Heating rates and wall temperatures were acceptable and close to predict profiles, and after scaling to a chamber pressure of 2250 psia, heating rates remained acceptable. In addition, a chamber spool made with the vacuum plasma spray process survived hot-fire testing with no detrimental effects.

Elam, Sandra K.↗
NASA NTRS1994

Ni-MH storage test and cycle life test

Gates Aerospace Batteries is conducting two long term test programs to fully characterize the NiMH cell technology for aerospace applications. The first program analyzes the effects of long term storage upon cell performance. The second program analyzes cycle life testing and preliminary production lot testing. This paper summarizes these approaches to testing the NiMH couple and culminates with initial storage and testing recommendations. Long term storage presents challenges to deter the adverse condition of capacity fade in NiMH cells. Elevated but stabilized pressures and elevated but stabilized end-of-charge voltages also appear to be a characteristic phenomenon of long term storage modes. However, the performance degradation is dependent upon specific characteristics of the metal-hydride alloy. To date, there is no objective evidence with which to recommend the proper method for storage and handling of NiMH cells upon shipment. This is particularly critical due to limited data points that indicate open circuit storage at room temperature for 60 to 90 days will result in irrecoverable capacity loss. Accordingly a test plan was developed to determine what method of mid-term to long-term storage will prevent irrecoverable capacity loss. The explicit assumption is that trickle charging at some rate above the self-discharge rate will prevent the irreversible chemical changes to the negative electrode that result in the irrecoverable capacity loss. Another premise is that lower storage temperatures, typically 0 C for aerospace customers, will impede any negative chemical reactions. Three different trickle charge rates are expected to yield a fairly flat response with respect to recoverable capacity versus baseline cells in two different modes of open circuit. Specific attributes monitored include: end-of-charge voltage, end-of-charge pressure, mid-point discharge voltage, capacity, and end-of-discharge pressure. Cycle life testing and preliminary production lot testing continue to dominate the overall technology development effort at GAB. The cell life test program reflects continuing improvements in baseline cell designs. Performance improvements include lower and more stable charge voltages and pressures. The continuing review of production lot testing assures conformance to the design criteria and expectations. This is especially critical during this period of transferring technology from research and development status to production.

Dell, R. Dan↗
NASA NTRS2007

Experimental Design for the Evaluation of Detection Techniques of Hidden Corrosion Beneath the Thermal Protective System of the Space Shuttle Orbiter

The detection of corrosion beneath Space Shuttle Orbiter thermal protective system is traditionally accomplished by removing the Reusable Surface Insulation tiles and performing a visual inspection of the aluminum substrate and corrosion protection system. This process is time consuming and has the potential to damage high cost tiles. To evaluate non-intrusive NDE methods, a Proof of Concept (PoC) experiment was designed and test panels were manufactured. The objective of the test plan was three-fold: establish the ability to detect corrosion hidden from view by tiles; determine the key factor affecting detectability; roughly quantify the detection threshold. The plan consisted of artificially inducing dimensionally controlled corrosion spots in two panels and rebonding tile over the spots to model the thermal protective system of the orbiter. The corrosion spot diameter ranged from 0.100" to 0.600" inches and the depth ranged from 0.003" to 0.020". One panel consisted of a complete factorial array of corrosion spots with and without tile coverage. The second panel consisted of randomized factorial points replicated and hidden by tile. Conventional methods such as ultrasonics, infrared, eddy current and microwave methods have shortcomings. Ultrasonics and IR cannot sufficiently penetrate the tiles, while eddy current and microwaves have inadequate resolution. As such, the panels were interrogated using Backscatter Radiography and Terahertz Imaging. The terahertz system successfully detected artificially induced corrosion spots under orbiter tile and functional testing is in-work in preparation for implementation.

Kemmerer, Catherine C.↗
NASA NTRS2003

HUMEX, a study on the survivability and adaptation of humans to long-duration exploratory missions, part I: lunar missions

The European Space Agency has recently initiated a study of the human responses, limits and needs with regard to the stress environments of interplanetary and planetary missions. Emphasis has been laid on human health and performance care as well as advanced life support developments including bioregenerative life support systems and environmental monitoring. The overall study goals were as follows: (i) to define reference scenarios for a European participation in human exploration and to estimate their influence on the life sciences and life support requirements; (ii) for selected mission scenarios, to critically assess the limiting factors for human health, wellbeing, and performance and to recommend relevant countermeasures; (iii) for selected mission scenarios, to critically assess the potential of advanced life support developments and to propose a European strategy including terrestrial applications; (iv) to critically assess the feasibility of existing facilities and technologies on ground and in space as testbeds in preparation for human exploratory missions and to develop a test plan for ground and space campaigns; (v) to develop a roadmap for a future European strategy towards human exploratory missions, including preparatory activities and terrestrial applications and benefits. This paper covers the part of the HUMEX study dealing with lunar missions. A lunar base at the south pole where long-time sunlight and potential water ice deposits could be assumed was selected as the Moon reference scenario. The impact on human health, performance and well being has been investigated from the view point of the effects of microgravity (during space travel), reduced gravity (on the Moon) and abrupt gravity changes (during launch and landing), of the effects of cosmic radiation including solar particle events, of psychological issues as well as general health care. Countermeasures as well as necessary research using ground-based test beds and/or the International Space Station have been defined. Likewise advanced life support systems with a high degree of autonomy and regenerative capacity and synergy effects were considered where bioregenerative life support systems and biodiagnostic systems become essential. Finally, a European strategy leading to a potential European participation in future human exploratory missions has been recommended. c2003 COSPAR. Published by Elsevier Ltd. All rights reserved.

Moon↗
NASA NTRS2007

1393 Ring Bus at JPL: Description and Status

Completed Ring Bus IC V&V Phase - Ring Bus Test Plan Completed for SIM Project - Applicable to Other Projects Implemented a Avionics Bus Based upon the IEEE 1393 Standard - Excellent Starting Point for a General Purpose High-Speed Spacecraft Bus - Designed to Meet SIM Requirements for - Real-time deterministic, distributed systems. - Control system requirements - Fault detection and recovery Other JPL Projects Considering Implementation F'light Software Ring Bus Driver Module Began in 2006, Continues Participating in Standard Revision. Search for Earth-like planets orbiting nearby stars and measure the masses and orbits of the planets it finds. Survey 2000 nearby stars for planetary systems to learn whether our Solar System is unusual, or typical. Make a new catalog of star position 100 times more accurate than current measurements. Learn how our galaxy formed and will evolve by studying the dynamics of its stars. Critically test models of exactly how stars shine, including exotic objects like black holes, neutron stars and white dwarfs.

time synchronization↗
NASA NTRS1974

In-harbor and at-sea electromagnetic compatibility survey for maritime satellite L-band shipboard terminal

Geostationary maritime satellites, one over the Pacific and one over the Atlantic Ocean, are planned to make available high-speed communications and navigation (position determination) services to ships at sea. A shipboard satellite terminal, operating within the authorized maritime L-band, 1636.5 to 1645.0 MHz, will allow ships to pass voice, teletype, facsimile, and data messages to shore communication facilities with a high degree of reliability. The shore-to-ship link will also operate in the maritime L-band from 1535.0 to 1543.5 MHz. A significant number or maritime/commercial ships are expected to be equipped with an L-band satellite terminal by the year 1980, and so consequently, there is an interest in determining electromagnetic compatibility between the proposed L-band shipboard terminal and existing, on-board, shipboard communications/electronics and electrical systems, as well as determining the influence of shore-based interference sources. The shipboard electromagnetic interference survey described was conducted on-board the United States Line's American Leader class (15,690 tons) commercial container ship, the "American Alliance" from June 16 to 20, 1974. Details of the test plan and measurements are given.

Source record↗
NASA NTRSDate not supplied

Simplified Applied Loads of a Transonic Truss-Braced Wing Derived from Computational Fluid Dynamics

The transonic truss-braced wing (TTBW) concept is under consideration by the NASA Advanced Air Transport Technology (AATT) project due to its potential to reduce fuel consumption of transport aircraft. To properly evaluate the TTBW concept, detailed structural modeling and subsequent test planning is underway, which requires a simplified set of discrete applied loads that simulate the distributed aerodynamic pressures. Given the unique architecture of the TTBW and its intended operation in the transonic regime, computational fluid dynamics (CFD) is the ideal approach to employ to ensure the aerodynamics of this configuration are sufficiently captured. To convert aerodynamic results into input for structural models, application of a data transfer tool, Discrete Data Transfer Between Dissimilar Meshes (DDTBDM), is presented to resolve aerodynamic pressures from CFD into single load vectors at hardpoints along both the wing and the strut. Evaluation of mapping the aerodynamic forces to hardpoints at rib and spar intersections in a structural model is performed, and the results are also compared with a linear static aeroelastic analysis. The structural responses indicate that distributing forces along only the wing and strut forward spars is insufficient for simulating the distributed aerodynamic loads. Adding moments along the forward spars improved the global response of the wing, but the point moments created unrealistic local rotations within the structure. When distributing loads along both the forward and aft spars of the wing and strut, the global response was less sensitive to the inclusion of point moments, which indicated the arrangement of point forces alone was sufficient input for structural design work and testing and provided a more continuous loading profile.

Finite Element↗
NASA NTRSDate not supplied

Simplified Applied Loads of a Transonic Truss-Braced Wing Derived from Computational Fluid Dynamic

The transonic truss-braced wing (TTBW) concept is under consideration by the NASA Advanced Air Transport Technology (AATT) project due to its potential to reduce fuel consumption of transport aircraft. To properly evaluate the TTBW concept, detailed structural modeling and subsequent test planning is underway, which requires a simplified set of discrete applied loads that simulate the distributed aerodynamic pressures. Given the unique architecture of the TTBW and its intended operation in the transonic regime, computational fluid dynamics (CFD) is the ideal approach to employ to ensure the aerodynamics of this configuration are sufficiently captured. To convert aerodynamic results into input for structural models, application of a data transfer tool, Discrete Data Transfer Between Dissimilar Meshes (DDTBDM), is presented to resolve aerodynamic pressures from CFD into single load vectors at hardpoints along both the wing and the strut. Evaluation of mapping the aerodynamic forces to hardpoints at rib and spar intersections in a structural model is performed, and the results are also compared with a linear static aeroelastic analysis. The structural responses indicate that distributing forces along only the wing and strut forward spars is insufficient for simulating the distributed aerodynamic loads. Adding moments along the forward spars improved the global response of the wing, but the point moments created unrealistic local rotations within the structure. When distributing loads along both the forward and aft spars of the wing and strut, the global response was less sensitive to the inclusion of point moments, which indicated the arrangement of point forces alone was sufficient input for structural design work and testing and provided a more continuous loading profile.

Finite Element↗
NASA NTRS2002

Magnetohydrodynamic Augmented Propulsion Experiment

A fundamental obstacle to routine space access is the specific energy limitations associated with chemical fuels. In the case of vertical take-off, the high thrust needed for vertical liftoff and acceleration to orbit translates into power levels in the 10 GW range. Furthermore, useful payload mass fractions are possible only if the exhaust particle energy (i.e., exhaust velocity) is much greater than that available with traditional chemical propulsion. The electronic binding energy released by the best chemical reactions (e.g., LOX/LH2 for example, is less than 2 eV per product molecule (approx. 1.8 eV per H2O molecule), which translates into particle velocities less than 5 km/s. Useful payload fractions, however, will require exhaust velocities exceeding 15 km/s (i.e., particle energies greater than 20 eV). As an added challenge, the envisioned hypothetical RLV (reusable launch vehicle) should accomplish these amazing performance feats while providing relatively low acceleration levels to orbit (2-3g maximum). From such fundamental considerations, it is painfully obvious that planned and current RLV solutions based on chemical fuels alone represent only a temporary solution and can only result in minor gains, at best. What is truly needed is a revolutionary approach that will dramatically reduce the amount of fuel and size of the launch vehicle. This implies the need for new compact high-power energy sources as well as advanced accelerator technologies for increasing engine exhaust velocity. Electromagnetic acceleration techniques are of immense interest since they can be used to circumvent the thermal limits associated with conventional propulsion systems. This paper describes the Magnetohydrodynamic Augmented Propulsion Experiment (MAPX) being undertaken at NASA Marshall Space Flight Center (MSFC). In this experiment, a 1-MW arc heater is being used as a feeder for a 1-MW magnetohydrodynamic (MHD) accelerator. The purpose of the experiment is to demonstrate that an MHD accelerator can be an effective augmentation system for increasing engine exhaust velocity. More specifically, the experiment is intended to show that electromagnetic effects are effective at producing flow acceleration whereas electrothermal effects do not cause unacceptable heating of the working fluid. The MHD accelerator was designed as an externally diagonalized segmented Faraday channel, which will be inserted into an existing 2-tesla electromagnet. This allows the external power to be connected through two terminals thereby minimizing the complexity and cost associated with powering each segment independently. The design of the accelerator and other components in the flow path has been completed and fabrication activities are underway. This paper provides a full description of MAPX including performance analysis, design, and test plans, and current status.

Litchford, Ron J.↗