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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 469 records · Page 26

All sources NASA NTRS DOE OSTI

NASA NTRS1982

Failure characteristics of graphite-epoxy structural components loaded in compression

Failure characteristics of compressively loaded graphite-epoxy components are described. Experimental results for both strength-critical laminates and structural components with postbuckling strength are presented. Effects of low speed impact damage and circular holes on compressive strength are discussed. Delamination and shear crippling failure mechanisms that limit the performance of strength-critical laminates are described. Transverse shear and skin-stiffener separation failure mechanisms that limit the performance of components with postbuckling strength are also described. The influence of matrix properties on compressive strength improvements for impact damaged laminates is discussed. Experimental data and results from a failure analysis for strength-critical laminates with cutouts are discussed and compared with impact damage results. Typical postbuckling test results are compared with analytical predictions.

Starnes, J. H., Jr.↗

NASA NTRS1983

Failure characteristics of graphite-epoxy structural components loaded in compression

Starnes, J. H., Jr.↗

NASA NTRS2002

Characterization of Ceramic Matrix Composite Fasteners Exposed in a Combustion Liner Rig Test

Combustion tests on SiC/SiC CMC (Ceramic Matrix Composite) components were performed in an aircraft combustion environment using the Rich-burn, Quick-quench, Lean-burn (RQL) sector rig. SiC/SiC fasteners were used to attach several of these components to the metallic rig structure. The effect of combustion exposure on the fastener material was characterized via microstructural examination. Fasteners were also destructively tested, after combustion exposure, and the failure loads of fasteners exposed in the sector rig were compared to those of as-manufactured fasteners. Combustion exposure reduced the fastener failure load by 50% relative to the as-manufactured fasteners for exposure times ranging from 50 to 260 hours. The fasteners exposed in the combustion environment demonstrated failure loads that varied with failure mode. Fasteners that had the highest average failure load, failed in the same manner as the unexposed fasteners.

Verrilli, Michael J.↗

NASA NTRS2019

Adaptive Control Law for PID

This paper is the final report for my Spring 2019 internship at the Kennedy Space Center in Cape Canaveral, Florida. The official title of my internship is 'Launch Vehicle Control Study internship,' and I spent the spring working in GMRO- Granular Mechanics and Regolith Operations. There are two components to my project involving developing an adaptive control law for a PID controller. The first component involves performing a deep study and analysis of adaptive control laws, with the goal of developing stability proofs about the control law and the gain and phase margins. The second component involves analyzing the data received after implementing the adaptive control law. Unfortunately, as of the time of writing this paper, the data is not considered â€˜clean' enough to analyze. Therefore, in this paper I will give an overview of adaptive control law stability proofs and will write about the data analysis separately. Section II includes several general definitions, and the later sections have additional definitions at the end of each section.A PID (proportional integral derivative) controller is a common control algorithm which isimplemented in NASA’s rocket launch system. The idea behind the PID controller is to calculatethe error (current position – desired position) and drive it to zero by using proportional, integral,and derivative influences on the controller. For example, when a sailor steers a ship which isheading towards location 𝑥 but would like to turn the ship to reach location 𝑥∗, he/she wouldinitially turn the wheel significantly, and as the ship proceeds towards 𝑥∗ the sailor would slowlyshift the wheel back to its original position, and thereby drive the error, 𝑥∗ − 𝑥, to zero.

Eisenberg, Yael↗

DOE OSTI2020

Modernization of Technical Requirements for Licensing of Advanced Non-Light Water Reactors: Safety Classification and Performance Criteria for Structures, Systems, and Components

This report, “Modernization of Technical Requirements for Licensing of Advanced Non-Light Water Reactors: Safety Classification and Performance Criteria for Structures, Systems and Components,” represents a key element in the development of a framework for the efficient licensing of advanced non-Light Water Reactors (non-LWRs). It is the result of a Licensing Modernization Project (LMP) led by Southern Company and cost-shared by the United States Department of Energy (DOE).

42 ENGINEERING↗

NASA NTRS1983

An assessment of the use of antimisting fuel in turbofan engines

An evaluation was made on the effects of using antimisting kerosene (AMK) on the performance of the components from the fuel system and the combustor of current in service JT8D aircraft engines. The objectives were to identify if there were any problems associated with using antimisting kerosene and to determine the extent of shearing or degradation required to allow the engine components to achieve satisfactory operation. The program consisted of a literature survey and a test program which evaluated the antimisting kerosene fuel in laboratory and bench component testing, and assessed the performance of the combustor in a high pressure facility and in an altitude relight/cold ignition facility.

Fiorentino, A. J.↗

NASA NTRS1989

Electric propulsion options for 10 kW class earth space missions

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment have been evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA II 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10(7) to 2.1x10(7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA II 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10(6) to 3.6x10(6) N-s, and approximately 662 to 1072 m/s, respectively.

Patterson, M. J.↗

NASA NTRS1989

Electric Propulsion Options for 10 kW Class Earth-Space Missions

Five and 10 kW ion and arcjet propulsion system options for a near-term space demonstration experiment were evaluated. Analyses were conducted to determine first-order propulsion system performance and system component mass estimates. Overall mission performance of the electric propulsion systems was quantified in terms of the maximum thrusting time, total impulse, and velocity increment capability available when integrated onto a generic spacecraft under fixed mission model assumptions. Maximum available thrusting times for the ion-propelled spacecraft options, launched on a DELTA 2 6920 vehicle, range from approximately 8,600 hours for a 4-engine 10 kW system to more than 29,600 hours for a single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 1.2x10 (exp 7) to 2.1x10 (exp 7) N-s, and 3550 to 6200 m/s, respectively. Maximum available thrusting times for the arcjet propelled spacecraft launched on the DELTA 2 6920 vehicle range from approximately 528 hours for the 6-engine 10 kW hydrazine system to 2328 hours for the single-engine 5 kW system. Maximum total impulse values and maximum delta-v's range from 2.2x10 (exp 6) to 3.6x10 (exp 6) N-s, and approximately 662 to 1072 m/s, respectively.

Patterson, M. J.↗

NASA NTRS1999

Microgravity Vibration Output Testing of Space Station Rotary Joints

The mission of the International Space Station is to provide a working laboratory in orbit for research in engineering, life sciences, and microgravity. Ensuring that the mechanical equipment on Space Station does not unduly disturb the microgravity environment is of paramount importance in meeting the Station's mission. The large inertia being moved by the Space Station's solar array and thermal radiator rotary joints make them one of the largest potential disturbance sources. The present paper describes the mechanical and control system design of these joints, their disturbance producing characteristics, and analytical predictions of some key performance indicators. The component and system functional tests performed to measure the actual vibration output of the joints are detailed and the test results discussed. Results of the rotary joint test program presented here show that the joints do meet the Space Station microgravity requirements, ensuring that this unique laboratory for microgravity research will be unaffected by the operation of the largest moving machinery on board.

Boucher, Robert L.↗

NASA NTRS2012

Progress in the Fabrication and Testing of Telescope Mirrors for the James Webb Space Telescope

The telescope of the James Webb Space Telescope (JWST) is an f/20, three mirror anastigmat design, passively cooled (40K) in an L2 orbit. The design provides diffraction limited performance (Strehl > or = 0.8) at .=2 m. To fit within the launch vehicle envelope (Arianne V), the 6.6 meter primary mirror and the secondary mirror support structure are folded for launch, then deployed and aligned in space. The primary mirror is composed of 18 individual, 1.3 meter (flat flat) hexagonal segments, each adjustable in seven degrees of freedom (six rigid body + radius of curvature) provided by a set of high precision actuators. The actuated secondary mirror (approx.0.74m) is similarly positioned in six degrees of rigid body motion. The approx..70x.51m, fixed tertiary and approx. 0.17m, flat fine steering mirror complete the telescope mirror complement. The telescope is supported by a composite structure optimized for performance at cryogenic temperatures. All telescope mirrors are made of Be with substantial light-weighting (21 kg for each 1.3M primary segment). Additional Be mounting and supporting structure for the high precision (approx.10nm steps) actuators are attached to the primary segments and secondary mirror. All mirrors undergo a process of thermal stabilization to reduce stress. An extensive series of interferometric measurements guide each step of the polishing process. Final polishing must account for any deformation between the ambient temperature of polishing and the cryogenic, operational temperature. This is accomplished by producing highly precise, cryo deformation target maps of each surface which are incorporated into the final polishing cycle. All flight mirrors have now completed polishing, coating with protected Au and final cryo testing, and the telescope is on track to meet all system requirements. We here review the measured performance of the component mirrors and the predicted performance of the flight telescope.

Bowers, Charles↗

DOE OSTI2025

Advanced Diagnosis and Accelerated Testing of Balance of System Components for Utility Scale PV Installations: October 1, 2022-September 30, 2024

A study of the durability of PV Balance of System components was performed. Specifically, wire cable jackets and cable connectors were examined within the direct current (DC) PV Power Transmission Chain (PTC). Degraded and failed samples have been obtained from utility PV installations to provide feedback on the degradation modes and the related damage-enabling considerations in today's PV systems. An industry interface group (including system owners, system inspectors, component manufacturers, and test labs) was used to help identify and obtain field-failed samples, for feedback (including samples and experimental design), and to facilitate the subsequent dissemination of the results of this study. Samples were empirically studied using accelerated stress testing with steady-state conditions (cable jackets) in addition to combined-accelerated stress testing (cable jackets, connectors and uncapped connectors). Steady-state accelerated testing has been performed using at least one applied stressor (e.g. UV light) to aid understanding of jacket durability relative to its application. Component- and material-focused failure analysis was conducted to develop an understanding and advise the PV industry. In-depth characterization will be applied selectively to field- and artificially aged-samples, to gain scientific understanding of the structural, chemical, electrical, mechanical, and thermal properties enabling degradation.

24 POWER TRANSMISSION AND DISTRIBUTION↗

NASA NTRS1973

Injector design guidelines for gas/liquid propellant systems

Injector design guidelines are provided for gas/liquid propellant systems. Information was obtained from a 30-month applied research program encompassing an analytical, design, and experimental effort to relate injector design parameters to simultaneous attainment of high performance and component (injector/thrust chamber) compatibility for gas/liquid space storable propellants. The gas/liquid propellant combination studied was FLOX (82.6% F2)/ ambient temperature gaseous methane. Design criteria that provide for simultaneous attainment of high performance and chamber compatibility are presented for both injector types. Parametric data are presented that are applicable for the design of circular coaxial and like-doublet injectors that operate with design parameters similar to those employed. However, caution should be exercised when applying these data to propellant combinations whose elements operate in ranges considerably different from those employed in this study.

Falk, A. Y.↗

NASA NTRSDate not supplied

Signal and Power Integrity Design Methodology for High-Performance Flight Computing Systems

Computing capabilities of space systems have in-creased onboard performance by orders of magnitude with the use of radiation-tolerant field-programmable gate arrays (FPGA)and processors. The incorporation of signal and power integrity analysis with printed circuit board (PCB) design in reliable computing architectures for space systems has become critical to enable future mission capabilities. Developers launch high-performance processors into a breadth of orbits and missions, running varying applications that create challenges for designing reliable computing hardware. Specifically, for these designs, academic and industry research has focused on component radiation performance, fault mitigation, and reliable architectures. How-ever, other design parameters including electromagnetic interference (EMI), PCB stackup, signal integrity (SI), voltage regulator module (VRM) design, and power distribution network (PDN)are often deprioritized or disregarded as the design matures. Since these characteristics are becoming more significant in high-performance processor designs, this research presents a hardware design and analysis methodology for high-performance, space-computing systems that focuses on a holistic design approach and PDN reliability. While these challenges exist across all space hardware, the reduced PCB dimensions imposed by SmallSats and CubeSats introduce additional hurdles, specifically to VRM and decoupling design. By examining the relationship between the PDN and radiation performance, an analytical relationship is developed that incorporates Total Ionizing Dose and Single-Event Transients to ensure reliability throughout the mission duration. The presented design methodology is applied to the SpaceCube v3.0 Mini, an FPGA-based on-board science data processing system developed at NASA Goddard Space Flight Center.

Advanced avionics↗

NASA NTRSDate not supplied

High Risk Spacecraft Materials Offgassing

NASA-STD-6001B, Determination of Offgassed Products (Test 7), provides the offgassing characteristics under standardized conditions for materials and assembled articles to be located within habitable spacecraft environments. Experience with Test 7 has found certain material types to be of higher risk for offgassing undesirable compounds aboard spacecraft than others. Formaldehyde and acrolein are historically high T value offgassed components of the offgassed compound target list because they have low spacecraft maximum allowable concentration (SMAC) values assigned by the JSC Toxicology Group. Carbon disulfide, benzene, acrylonitrile, and furan are additional target compounds of concern due to their lower thresholds of toxicity as determined by the JSC Toxicology Group. Materials offgassing siloxanes are also of concern due to their degradation effects on environmental control and life support system (ECLSS) components and performance. Spacecraft materials and articles defined in this manuscript as high risk were identified after examining and condensing data for these compounds of concern from approximately 3000 tests performed over 30 years. Summaries of high risk material and article types based on highest Multi-Purpose Crew Vehicle (MPCV) T values are also presented. Historical analysis shows high risk components are produced largely from test materials and articles in the general categories of electronic/powered components, foams, paints/coatings/films, adhesives/tapes, epoxy/resins, liquids/gels, Nomex® with surface treatments, markers/pens/inks, dry film lubricants, thermoplastics, hygiene items (deodorants, lip balms), and silicone rubber. These data are intended to be a resource for spacecraft materials and processes managers, designers, and toxicologists. High risk materials and articles intended for use aboard spacecraft should be tested in accordance with NASA-STD-6001B Test 7.

Vanessa D Buchanan↗

NASA NTRS1972

An H2-O2 auxiliary power unit for space shuttle

Auxiliary power units operating on hydrogen and oxygen have the potential for significantly improved performance and payload for space shuttle vehicles. A recuperated open cycle hydrogen-oxygen turbine driven power system concept with pressure modulated reactant control is considered. The reference system, the control mode, component arrangement, performance potential, and the technology program for this system are discussed.

Beremand, D. G.↗

NASA NTRS1973

Simplified procedures for correlation of experimentally measured and predicted thrust chamber performance

Thrust chamber performance is evaluated in terms of an analytical model incorporating all the loss processes that occur in a real rocket motor. The important loss processes in the real thrust chamber were identified, and a methodology and recommended procedure for predicting real thrust chamber vacuum specific impulse were developed. Simplified equations for the calculation of vacuum specific impulse are developed to relate the delivered performance (both vacuum specific impulse and characteristic velocity) to the ideal performance as degraded by the losses corresponding to a specified list of loss processes. These simplified equations enable the various performance loss components, and the corresponding efficiencies, to be quantified separately (except that interaction effects are arbitrarily assigned in the process). The loss and efficiency expressions presented can be used to evaluate experimentally measured thrust chamber performance, to direct development effort into the areas most likely to yield improvements in performance, and as a basis to predict performance of related thrust chamber configurations.

Powell, W. B.↗

NASA NTRS2004

Results From a Parametric Acoustic Liner Experiment Using P and W GEN1 HSR Mixer/Ejector Model

This report documents the results of an acoustic liner test performed using a Gen 1 HSR mixer/ejector model installed on the Jet Exit Rig in the Nozzle Acoustic Test Rig in the Aeroacoustic Propulsion Laboratory or NASA Glenn Research Center. Acoustic liner effectiveness and single-component thrust performance results are discussed. Results from 26 different types of single-degree-of-freedom and bulk material liners are compared with each other and against a hardwall baseline. Design parameters involving all aspects of the facesheet, the backing cavity, and the type of bulk material were varied in order to study the effects of these design features on the acoustic impedance, acoustic effectiveness and on nozzle thrust performance. Overall, the bulk absorber liners are more effective at reducing the jet noise than the single-degree-of-freedom liners. Many of the design parameters had little effect on acoustic effectiveness, such as facesheeet hole diameter and honeycomb cell size. A relatively large variation in the impedance of the bulk absorber in a bulk liner is required to have a significant impact on the noise reduction. The thrust results exhibit a number of consistent trends, supporting the validity of this new addition to the facility. In general, the thrust results indicate that thrust performance benefits from increased facesheet thickness and decreased facesheet porosity.

Boyd, Kathleen C.↗

NASA NTRS2013

NASA-USRP Summer 2013 Internship Final Report

Three major projects were undertaken during the Summer 2013 USRP Internship: (A) assisting the cTAPS group with component and pressure vessel system analyses and documentation, (B) designing a hoisting fixture for a solid rocket motor, (C) finding an alternative to removing the DOT rated gaseous nitrogen tank from the roof for hydrostatic testing. Hypergolic Material Assessments (HMAs) and pressure calculations were performed on components of pressure systems. Additionally, component information was logged in the Standard Parts Database to provide a location where system designers can find information regarding components, including their specifications and compatibility with fluids. A hoisting fixture was designed to hoist a solid rocket motor and meets the specifications related to stress and size. However, there are issues with the fixtures bolt head allotment, the bolt spacing, and the complexity of the part. Finally, calculations were performed on an expiring DOT rated gaseous nitrogen tank in an attempt to re-rate it per ASME standards. This was unsuccessful so other options are being explored for the tank. While much progress was made on all three projects, there is still work to be performed on each project to achieve the desired results.

Gurganus, S. Christine↗