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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 523 records · Page 29
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Space transportation system payload interface verification

The paper considers STS payload-interface verification requirements and the capability provided by STS to support verification. The intent is to standardize as many interfaces as possible, not only through the design, development, test and evaluation (DDT and E) phase of the major payload carriers but also into the operational phase. The verification process is discussed in terms of its various elements, such as the Space Shuttle DDT and E (including the orbital flight test program) and the major payload carriers DDT and E (including the first flights). Five tools derived from the Space Shuttle DDT and E are available to support the verification process: mathematical (structural and thermal) models, the Shuttle Avionics Integration Laboratory, the Shuttle Manipulator Development Facility, and interface-verification equipment (cargo-integration test equipment).

Everline, R. T.↗
NASA NTRSDate not supplied

Flight and Direct to Earth/Space Relay Communication System Architecture for GSFC CubeSat Missions

The CubeSat platform is finding increasing use in space science applications due to its low cost and comparative ease of launch. It is becoming a key scientific discovery tool in low Earth orbit (LEO) and beyond, including geosynchronous equatorial orbit (GEO), the Lagrange Points, Lunar missions, and more. The increasing complexity of these missions and their scientific goals must be supported by equal advancements in communications technology. Higher data rates and greater reliability are required every year. However, the reduced Size, Weight, and Power (SWaP) constraints of CubeSat platforms introduce unique challenges in the area of satellite communications. There is currently a lack of communication equipment tailored specifically to the CubeSat platform. This lack of standardized, tested equipment extends development time and reduces mission confidence. Furthermore, missions utilizing the CubeSat platform are often subject to more difficult design constraints. Antenna placement, size, and pointing are often subordinate to the requirements of the payload instruments and mission goals. Traditional link margin estimation techniques are insufficient in these cases, as they emphasize worst case scenarios. In reality the actual link parameters may vary widely even during a single pass. This presents new challenges in predicting communications performance and scheduling ground station contacts, but also new opportunities for improving efficiency. This paper presents the integration, testing, and validation process for a new software defined radio (SDR) designed for the CubeSat platform in conjunction with Vulcan Wireless, Inc. The SDR is planned for use on 5 upcoming CubeSat missions at NASAs Goddard Space Flight Center (GSFC) including a Geosynchronous Transfer Orbit (GTO) mission and it may also serve as a standard and well-tested option for future missions by enabling a standardized, rapid and low cost CubeSat communication system network integration process. Detailed simulations have been developed to estimate the communication performance of these missions, taking the unique antenna placements and attitude behavior of each satellite into account. These simulations allow a much more accurate analysis of the expected link margin, which varies considerably during each pass for the NASA Space Relay (SR) and Direct to Earth (DTE) network. The modelling procedures are outlined, and the results are used to predict communications performance of the missions.

Space Networks↗
NASA NTRS1994

Multi-lab comparison on R-curve methodologies: Alloy 2024-T3

In an effort to determine an optimum method for ranking the fracture toughness of developmental aluminum alloys for High Speed Civil Transport applications, five labs performed K and/or J based fracture tests on aluminum alloy 2024-T3. Two material thicknesses were examined: 0.063 in. and 0.125 in. Middle crack tension and compact tension specimens were excised from 60 in. wide middle crack tension panels which had been previously tested at Boeing. The crack resistance curves generated were compared to the R-curves from 60 in. wide specimens. The experimental program indicated that effective stress intensity from secant compliance based crack length and stress intensity calculated from J-integral testing were equivalent. In addition, comparison of different specimen sizes and configurations indicated that standard validity requirements for compact tension specimens may be overly restrictive.

Reynolds, Anthony P.↗
NASA NTRS2008

Feasibility of Conducting J-2X Engine Testing at the Glenn Research Center Plum Brook Station B-2 Facility

A trade study of the feasibility of conducting J-2X testing in the Glenn Research Center (GRC) Plum Brook Station (PBS) B-2 facility was initiated in May 2006 with results available in October 2006. The Propulsion Test Integration Group (PTIG) led the study with support from Marshall Space Flight Center (MSFC) and Jacobs Sverdrup Engineering. The primary focus of the trade study was on facility design concepts and their capability to satisfy the J-2X altitude simulation test requirements. The propulsion systems tested in the B-2 facility were in the 30,000-pound (30K) thrust class. The J-2X thrust is approximately 10 times larger. Therefore, concepts significantly different from the current configuration are necessary for the diffuser, spray chamber subsystems, and cooling water. Steam exhaust condensation in the spray chamber is judged to be the key risk consideration relative to acceptable spray chamber pressure. Further assessment via computational fluid dynamics (CFD) and other simulation capabilities (e.g. methodology for anchoring predictions with actual test data and subscale testing to support investigation.

Schafer, Charles F.↗
NASA NTRS1982

Refurbishment of SRB aluminum components by walnut hull blast removal of protective coatings

A test program was conducted to develop, optimize, and scale up an abrasive blasting procedure was developed for refurbishment of specific SRB components: aft skirt, forward skirt, frustrum, and painted piece parts. Test specimens utilizing 2219 T87 aluminum substrate of varying thicknesses were prepared and blasted at progressively increasing pressures with selected abrasives. Specimens were analyzed for material response. The optimum blasting parameters were determined on panel specimens and verified on a large cylindrical integrated test bed.

Colberg, W. R.↗
NASA NTRS2018

Simulations of the NASA Langley 14- by 22-Foot Subsonic Tunnel for the Juncture Flow Experiment

NASA’s Transformational Tools and Technologies Program’s Juncture Flow experiment aims to provide data to improve Computational Fluid Dynamics (CFD) modeling in the juncture flow region. The experiment is planned to provide validation-quality data for CFD that focuses on the separation bubble near the wing-body juncture trailing edge region. Because wind tunnel tests associated with the Juncture Flow project have been designed for the purpose of CFD validation, considerable effort is going into modeling and simulating the wind tunnel. This is not only important because wind tunnel wall effects can play a role in integrated testing uncertainties, but also because the better the boundary conditions are known, the better CFD can accurately represent the experiment. This paper builds on the recent CFD efforts to model the NASA Langley 14- by 22-Foot Subsonic Tunnel. Current best practices in simulating wind tunnels are evaluated. The features of each method, as well as some of their pros and cons, are highlighted. Boundary conditions and modeling techniques currently used by CFD for empty-tunnel simulations are also described. Preliminary CFD studies associated with modeling the Juncture Flow model are summarized, with the intention to determine sensitivities of the flow near the wing-body juncture region of the model to a variety of modeling decisions.

Subsonic↗
NASA NTRSDate not supplied

The Marshall Grazing Incidence X-ray Spectrometer (MaGIXS) Solar Sounding Rocket Campaign – Calibration and Performance

The Marshall Grazing Incidence X-ray Spectrometer (MaGIXS) is a sounding rocket experiment that is designed to observe, for the first time, soft X-ray spectra of high-temperature, low-emission plasma of coronal structures spatially resolved along a narrow slit.MaGIXS observation involves a set of high temperature spectral lines in soft X-rays from 0.5 - 2.0 keV from an active region core,which will extend the DEM coverage from 3MK to 10MK constraining the slope of the DEM fall-off. The novel instrument designincludes a Wolter-I type telescope and a 3-optic grazing-incidence spectrometer. The spectrometer consists of a finite conjugatemirror pair and a blazed planar, varied line spaced grating, which disperses the rays on to a CCD and provides a high spatial andspectral resolution. Component level instrument testing, integration of the instrument and end-to-end X-ray calibration are carriedout using the X-ray and Cryogenic Facility (XRCF) at NASA Marshall Space Flight Center. MaGIXS is scheduled for launch in2021. We will present the results of X-ray calibration tests for MaGIXS and discuss the expected inflight performance throughdifferent solar observation scenarios.

P. S. Athiray↗
NASA NTRS1976

Solid electrolyte oxygen regeneration system

A program to design, develop, fabricate and assemble a one-man, self-contained, solid electrolyte oxygen regeneration system (SX-1) incorporating solid electrolyte electrolyzer drums was completed. The SX-1 is a preprototype engineering model designed to produce 0.952 kg (2.1 lb)/day of breathable oxygen (O2) from the electrolysis of metabolic carbon dioxide (CO2) and water vapor. The CO2 supply rate was established based on the metabolic CO2 generation rate for one man of 0.998 kg (2.2 lb)/day. The water supply rate (0.254 kg (0.56 lb)/day) was designed to be sufficient to make up the difference between the 0.952 kg (2.1 lb)/day O2 generation specification and the O2 available through CO2 electrolysis, 0.726 kg (1.6 lb)/day. The SX-1 was successfully designed, fabricated and assembled. Design verification tests (DVT) or the CO Disproportionators, H2 separators, control instrumentation, monitor instrumentation, water feed mechanism were successfully completed. The erratic occurrence of electrolyzer drum leakage prevented the completion of the CO2 electrolyzer module and water electrolyzer module DVT's and also prevented the performance of SX-1 integrated testing. Further development work is required to improve the solid electrolyte cell high temperature seals.

Shumar, J. W.↗
NASA NTRS1996

Parallelization of Rocket Engine System Software (Press)

The main goal is to assess parallelization requirements for the Rocket Engine Numeric Simulator (RENS) project which, aside from gathering information on liquid-propelled rocket engines and setting forth requirements, involve a large FORTRAN based package at NASA Lewis Research Center and TDK software developed by SUBR/UWF. The ultimate aim is to develop, test, integrate, and suitably deploy a family of software packages on various aspects and facets of rocket engines using liquid-propellants. At present, all project efforts by the funding agency, NASA Lewis Research Center, and the HBCU participants are disseminated over the internet using world wide web home pages. Considering obviously expensive methods of actual field trails, the benefits of software simulators are potentially enormous. When realized, these benefits will be analogous to those provided by numerous CAD/CAM packages and flight-training simulators. According to the overall task assignments, Hampton University's role is to collect all available software, place them in a common format, assess and evaluate, define interfaces, and provide integration. Most importantly, the HU's mission is to see to it that the real-time performance is assured. This involves source code translations, porting, and distribution. The porting will be done in two phases: First, place all software on Cray XMP platform using FORTRAN. After testing and evaluation on the Cray X-MP, the code will be translated to C + + and ported to the parallel nCUBE platform. At present, we are evaluating another option of distributed processing over local area networks using Sun NFS, Ethernet, TCP/IP. Considering the heterogeneous nature of the present software (e.g., first started as an expert system using LISP machines) which now involve FORTRAN code, the effort is expected to be quite challenging.

Cezzar, Ruknet↗
NASA NTRS2017

Boundary Condition Study for the Juncture Flow Experiment in the NASA Langley 14x22-Foot Subsonic Wind Tunnel

Because future wind tunnel tests associated with the NASA Juncture Flow project are being designed for the purpose of CFD validation, considerable effort is going into the characterization of the wind tunnel boundary conditions, particularly at inflow. This is important not only because wind tunnel flowfield nonuniformities can play a role in integrated testing uncertainties, but also because the better the boundary conditions are known, the better CFD can accurately represent the experiment. This paper describes recent investigative wind tunnel tests involving two methods to measure and characterize the oncoming flow in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The features of each method, as well as some of their pros and cons, are highlighted. Boundary conditions and modeling tactics currently used by CFD for empty-tunnel simulations are also described, and some results using three different CFD codes are shown. Preliminary CFD parametric studies associated with the Juncture Flow model are summarized, to determine sensitivities of the flow near the wing-body juncture region of the model to a variety of modeling decisions.

Rumsey, C. L.↗
NASA NTRS2017

Development of an IVE/EVA Compatible Prototype Cold-Gas Cubesat Propulsion System at NASA/JSC

Cold-gas propulsion systems are well suited for some applications because they are simple to design and build, have low operating costs, and are non-toxic. The inherent tradeoff, however, is their relatively low impulse density. Nevertheless, a modest propulsion system, sized for Cubesats and designed for affordability, presents an attractive system solution for some missions, such as an on-orbit inspection free-flyer. NASA has a long-standing effort to develop propulsion systems appropriate for very high delta-V cubesat missions, such as geo transfer orbits, and there are commercially available Cubesat propulsion systems with considerably more impulse capability, but, these are both prohibitively expensive for some development customers and face compatibility constraints for crewed applications, such as operation within ISS. A relatively conventional cold-gas system has been developed at NASA/JSC taking advantage of existing miniature industrial components, additive manufacturing techniques and in-house qualification of the system. The result is a nearly modular system with a 1U form factor. Compressed nitrogen is stored in a small high-pressure tank, then regulated and distributed to 12 thrusters. Maneuvering thrust can be adjusted, with a typical value of 40 mN, and the delta-V delivered to a 3U Cubesat would be approximately 7 m/s. These values correspond to the performance parameters for an inspection mission previously established at JSC for inspection of the orbiter prior to reentry. Environmental testing was performed to meet ISS launch and workmanship standards, along with the expected thermal environment for an inspection mission. Functionality has been demonstrated, and performance in both vacuum and relevant blow down scenarios was completed. Several avenues for further improvement are also explored. Details of the system, components, integration, tests, and test data are presented in this paper.

Radke, Christopher↗
NASA NTRS1995

Development Status of the NSTAR Ion Propulsion System Power Processor

A 0.5-2.3 kW xenon ion propulsion system is presently being developed under the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) program. This propulsion system includes a 30 cm diameter xenon ion thruster, a Digital Control Interface Unit, a xenon feed system, and a power processing unit (PPU). The PPU consists of the power supply assemblies which operate the thruster neutralizer, main discharge chamber, and ion optics. Also included are recycle logic and a digital microcontroller. The neutralizer and discharge power supplies employ a dual use configuration which combines the functions of two power supplies into one, significantly simplifying the PPU. Further simplification was realized by implementing a single thruster control loop which regulates the beam current via the discharge current. Continuous throttling is possible over a 0.5-2.3 kW output power range. All three power supplies have been fabricated and tested with resistive loads, and have been combined into a single breadboard unit with the recycle logic and microcontroller. All line and load regulation test results show the power supplies to be within the NSTAR flight PPU specified power output of 1.98 kW. The overall efficiency of the PPU, calculated as the combined efficiencies of the power supplies and controller, at 2.3 kW delivered to resistive loads was 0.90. The component was 6.16 kg. Integration testing of the neutralizer and discharge power supplies with a functional model thruster revealed no issues with discharge ignition or steady state operation.

Hamley, John A.↗
NASA NTRS2013

Methane Post-Processor Development to Increase Oxygen Recovery beyond State-of-the-Art Carbon Dioxide Reduction Technology

State‐of‐the‐art life support carbon dioxide (CO2) reduction technology, based on the Sabatier reaction, is theoretically capable of 50% recovery of oxygen from metabolic CO2. This recovery is constrained by the limited availability of reactant hydrogen. Post‐processing of the methane byproduct from the Sabatier reactor results in hydrogen recycle and a subsequent increase in oxygen recovery. For this purpose, a Methane Post‐Processor Assembly containing three sub‐systems has been developed and tested. The assembly includes a Methane Purification Assembly (MePA) to remove residual CO2 and water vapor from the Sabatier product stream, a Plasma Pyrolysis Assembly (PPA) to partially pyrolyze methane into hydrogen and acetylene, and an Acetylene Separation Assembly (ASepA) to purify the hydrogen product for recycle. The results of partially integrated testing of the sub‐systems are reported.

Abney, Morgan↗
NASA NTRS2013

Methane Post-Processor Development to Increase Oxygen Recovery beyond State-of-the-Art Carbon Dioxide Reduction Technology

State-of‐the‐art life support carbon dioxide (CO2) reduction technology, based on the Sabatier reaction, is theoretically capable of 50% recovery of oxygen from metabolic CO2. This recovery is constrained by the limited availability of reactant hydrogen. Post‐processing of the methane byproduct from the Sabatier reactor results in hydrogen recycle and a subsequent increase in oxygen recovery. For this purpose, a Methane Post‐Processor Assembly containing three sub‐systems has been developed and tested. The assembly includes a Methane Purification Assembly (MePA) to remove residual CO2 and water vapor from the Sabatier product stream, a Plasma Pyrolysis Assembly (PPA) to partially pyrolyze methane into hydrogen and acetylene, and an Acetylene Separation Assembly (ASepA) to purify the hydrogen product for recycle. The results of partially integrated testing of the sub‐systems are reported

Abney, Morgan B.↗
NASA NTRS1999

Summary of HEAT 1 Aeroacoustics Installation Effects

A critical part of the NASA High-Speed Research (HSR) program is the demonstration of satisfactory suppression of the jet noise present at low airspeeds. One scheme for reducing jet exhaust noise generated by a future High-Speed Civil Transport (HSCT) is the use of a mixer/ ejector system which would entrain large quantities of ambient air into the exhaust flow from the powerplant in order to cool and slow the jet exhaust before it leaves the tailpipe. Of the variety of factors which can affect the noise suppression characteristics of the mixer/ejector system, the influence of the wing flow field and high-lift devices is not well understood. The effectiveness of the noise suppression device must be evaluated in the presence of the wing/high-lift system before definitive assessments can be made concerning HSCT noise. Of nearly equal importance is the evaluation of the performance of the high-lift system(s) in the presence of realistic propulsion units which feature high ambient flow entrainment rates and jet thrust coefficients. These noise suppressors must provide the required acoustic attenuation while not overly degrading the thrust efficiency of the propulsion system or the lift enhancement of the high-lift devices on the wing. The overall objective of the NASA High-lift Engine Aeroacoustics Technology program is to demonstrate satisfactory interaction between the jet noise suppressor and the high-lift system at airspeeds and angles of attack consistent with takeoff, climb, approach, and landing. In support of this program, an isolated aeroacoustic test of a 13.5%-scale, candidate mixer/ejector nozzle was performed in the Ames' Research Center 40- by 80-Foot Wind Tunnel. The purpose of the test was to measure the baseline aeroacoustic performance characteristics of this nozzle in isolation from the aerodynamic flowfield induced by an HSCT airframe. The test documented the acoustic signature of the nozzles with treated and hardwall ejector surfaces and with changes in the ratio of ejector-duct-to-jet-area over a wide range range of nozzle pressure ratios and freestream Mach numbers. The test also measured the thrust performance, ambient-flow aspiration ratio, and internal and external static pressures on the nozzles. The isolated aeroacoustic performance data has been compared with results obtained with this nozzle installed on a 13.5% Boeing Reference H HSCT configuration, semi-span model. The semi-span, aeroacoustics integration test documented the first-order effects of the airframe flowfield on the acoustic performance of the nozzle and the effect of the nozzle secondary inlet flows on the aerodynamic performance of the wing high-lift systems. This investigation is critical to understanding the mutual installation effects of mixer/ejector nozzles and wing high-lift systems.

Smith, Brian E.↗
NASA NTRS2022

Testing Mars 2020 Flight Software and Hardware in the Surface System Development Environment

The Mars 2020 (M2020) Perseverance Rover is NASA's most advanced planetary rover mission to date. It includes a novel Sample Caching Subsystem (SCS) which will collect rock cores for possible future return to Earth, as well as an improved mobility system with enhanced autonomous navigation which will enable it to traverse faster and farther than prior rovers. The development of both systems required extensive flight software and flight hardware testing. To support this testing, we developed the Surface System Development Environment (SSDEV) and used it for a wide variety of testing. SSDEV is a bundled subset of M2020 Flight Software which runs on commercially available Linux computers and can be combined with multiple backend options for simulation and hardware control. The SSDEV architecture enabled our teams to perform much more testing of flight software and flight hardware than would have otherwise been possible. As a secondary benefit, the SSDEV-based test campaigns also helped our teams enter the operations phase of the mission with greater readiness of operations products and tools. In this paper, we summarize the motivation for SSDEV, provide an overview of the SSDEV architecture, list several examples of how SSDEV was used, and summarize lessons learned. SSDEV is not a substitute for integrated testing with flight-like avionics, but it enabled substantially more testing than would have otherwise been possible and also provided some unique benefits. We recommend architectures like SSDEV to future projects that need to perform extensive hardware and software testing using a limited set of flight-like avionics.

Wai, Dennis↗
NASA NTRS1991

Isothermal Dendritic Growth Experiment - Science, engineering, and hardware development for USMP space flights

The Isothermal Dendritic Growth Experiment (IDGE) has been designed to provide microgravity data on dendritic growth for a critical test of theory. This paper updates progress on constructing a crystal growth chamber suitable for space flight. The IDGE chamber is constructed from glass and stainless steel and is hermetically sealed by electron beam welds and glass-metal seals. Initial tests of the chambers sample's melting point plateau show that the new chamber design is capable of preserving the 99.9995 percent purity of succinonitrile. Dendrite growth can be initiated in the center of the IDGE chamber by means of thermo-electric coolers and a capillary injector tube (stinger). The new IDGE chamber is ready for fully integrated tests with the prototype IDGE engineering hardware at NASA's Lewis Research Center.

Glicksman, M. E.↗
NASA NTRS2012

Extensible Adaptable Simulation Systems: Supporting Multiple Fidelity Simulations in a Common Environment

Common practice in the development of simulation systems is meeting all user requirements within a single instantiation. The Joint Polar Satellite System (JPSS) presents a unique challenge to establish a simulation environment that meets the needs of a diverse user community while also spanning a multi-mission environment over decades of operation. In response, the JPSS Flight Vehicle Test Suite (FVTS) is architected with an extensible infrastructure that supports the operation of multiple observatory simulations for a single mission and multiple mission within a common system perimeter. For the JPSS-1 satellite, multiple fidelity flight observatory simulations are necessary to support the distinct user communities consisting of the Common Ground System development team, the Common Ground System Integration & Test team, and the Mission Rehearsal Team/Mission Operations Team. These key requirements present several challenges to FVTS development. First, the FVTS must ensure all critical user requirements are satisfied by at least one fidelity instance of the observatory simulation. Second, the FVTS must allow for tailoring of the system instances to function in diverse operational environments from the High-security operations environment at NOAA Satellite Operations Facility (NSOF) to the ground system factory floor. Finally, the FVTS must provide the ability to execute sustaining engineering activities on a subset of the system without impacting system availability to parallel users. The FVTS approach of allowing for multiple fidelity copies of observatory simulations represents a unique concept in simulator capability development and corresponds to the JPSS Ground System goals of establishing a capability that is flexible, extensible, and adaptable.

McLaughlin, Brian J.↗