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At least 271 records · Page 15

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NASA NTRS2012

Non-intrusive Shock Measurements Using Laser Doppler Vibrometers

Stud mount accelerometers are widely used by the aerospace industry to measure shock environments during hardware qualification. The commonly used contact-based sensors, however, interfere with the shock waves and distort the acquired signature, which is a concern not actively discussed in the community. To alleviate these interference issues, engineers at the Jet Propulsion Laboratory are investigating the use of non-intrusive sensors, specifically Laser Doppler Vibrometers, as alternatives to the stud mounted accelerometers. This paper will describe shock simulation tests completed at the Jet Propulsion Laboratory, compare the measurements from stud mounted accelerometers and Laser Doppler Vibrometers, and discuss the advantages and disadvantages of introducing Laser Doppler Vibrometers as alternative sensors for measuring shock environments.

shock testing↗

NASA NTRS2009

Performance Characteristics of the NEXT Long-Duration Test After 16,550 h and 337 kg of Xenon Processed

The NASA s Evolutionary Xenon Thruster (NEXT) program is developing the next-generation ion propulsion system with significant enhancements beyond the state-of-the-art in ion propulsion to provide future NASA science missions with enhanced mission capabilities at a low total development cost. As part of a comprehensive thruster service life assessment utilizing both testing and analyses, a Long-Duration Test (LDT) was initiated to verify the NEXT propellant throughput capability to a qualification-level of 450 kg, 1.5 times the anticipated throughput requirement of 300 kg from mission analyses conducted utilizing the NEXT propulsion system. The LDT is being conducted with a modified, flight-representative NEXT engineering model ion thruster, designated EM3. As of June 25, 2008, the thruster has accumulated 16,550 h of operation: the first 13,042 h at the thruster full-input-power of 6.9 kW with 3.52 A beam current and 1800 V beam power supply voltage. Operation since 13,042 h, i.e., the most recent 3,508 h, has been at an input power of 4.7 kW with 3.52 A beam current and 1180 V beam power supply voltage. The thruster has processed 337 kg of xenon (Xe) surpassing the NSTAR propellant throughput demonstrated during the extended life testing of the Deep Space 1 flight spare ion thruster. The NEXT LDT has demonstrated a total impulse of 13.3 106 N s; the highest total impulse ever demonstrated by an ion thruster. Thruster performance tests are conducted periodically over the entire NEXT throttle table with input power ranging 0.5 to 6.9 kW. Thruster performance parameters including thrust, input power, specific impulse, and thruster efficiency have been nominal with little variation to date. This paper presents the performance of the NEXT LDT to date with emphasis on performance variations following throttling of the thruster to the new operating condition and comparison of performance to the NSTAR extended life test.

Soulas, George C.↗

NASA NTRS2018

Natural Frequency Testing and Model Correlation of Rocket Engine Structures in Liquid Hydrogen - Phase I, Cantilever Beam

Many structures in the launch vehicle industry operate in liquid hydrogen (LH2), from the hydrogen fuel tanks through the ducts and valves and into the pump sides of the turbopumps. Calculating the structural dynamic response of these structures is critical for successful qualification of this hardware, but accurate knowledge of the natural frequencies is based entirely on numerical or analytical predictions of frequency reduction due to the added-fluid-mass effect because testing in LH2 has always been considered too difficult and dangerous. This fluid effect is predicted to be approximately 4-5% using analytical formulations for simple cantilever beams. As part of a comprehensive test/analysis program to more accurately assess pump inducers operating in LH2, a series of frequency tests in LH2 were performed at NASA/Marshall Space Flight Center's unique cryogenic test facility. These frequency tests are coupled with modal tests in air and water to provide critical information not only on the mass effect of LH2, but also the cryogenic temperature effect on Young's Modulus for which the data is not extensive. The authors are unaware of any other reported natural frequency testing in this media. In addition to the inducer, a simple cantilever beam was also tested in the tank to provide a more easily modeled geometry as well as one that has an analytical solution for the mass effect. This data will prove critical for accurate structural dynamic analysis of these structures, which operate in a highly-dynamic environment.

Brown, Andrew M.↗

NASA NTRS1996

Improved Aircraft Inlet Total-Pressure Distortion Rakes

Innovative rake of pressure transducers for measuring distortion of total pressure at inlet of jet engine developed. Rake features one-piece wagon-wheel design, reducing cost and installation time relative to traditional designs. Design adaptable to inlets of other aircraft. Rake scaled to smaller and larger inlets with minimal qualification requirements.

Ray, Ronald J.↗

NASA NTRS2005

Structural Analysis Made 'NESSUSary'

Everywhere you look, chances are something that was designed and tested by a computer will be in plain view. Computers are now utilized to design and test just about everything imaginable, from automobiles and airplanes to bridges and boats, and elevators and escalators to streets and skyscrapers. Computer-design engineering first emerged in the 1970s, in the automobile and aerospace industries. Since computers were in their infancy, however, architects and engineers during the time were limited to producing only designs similar to hand-drafted drawings. (At the end of 1970s, a typical computer-aided design system was a 16-bit minicomputer with a price tag of $125,000.) Eventually, computers became more affordable and related software became more sophisticated, offering designers the "bells and whistles" to go beyond the limits of basic drafting and rendering, and venture into more skillful applications. One of the major advancements was the ability to test the objects being designed for the probability of failure. This advancement was especially important for the aerospace industry, where complicated and expensive structures are designed. The ability to perform reliability and risk assessment without using extensive hardware testing is critical to design and certification. In 1984, NASA initiated the Probabilistic Structural Analysis Methods (PSAM) project at Glenn Research Center to develop analysis methods and computer programs for the probabilistic structural analysis of select engine components for current Space Shuttle and future space propulsion systems. NASA envisioned that these methods and computational tools would play a critical role in establishing increased system performance and durability, and assist in structural system qualification and certification. Not only was the PSAM project beneficial to aerospace, it paved the way for a commercial risk- probability tool that is evaluating risks in diverse, down- to-Earth application

Source record↗

NASA NTRS2003

Mercury Conditions for the MESSENGER Mission Simulated in High- Solar-Radiation Vacuum Tests

The MESSENGER (Mercury Surface, Space Environment, Geochemistry, and Ranging) spacecraft, planned for launch in March 2004, will perform two flybys of Mercury before entering a year-long orbit of the planet in September 2009. The mission will provide opportunities for detailed characterization of the surface, interior, atmosphere, and magnetosphere of the closest planet to the Sun. The NASA Glenn Research Center and the MESSENGER spacecraft integrator, the Johns Hopkins University Applied Physics Laboratory, have partnered under a Space Act Agreement to characterize a variety of critical components and materials under simulated conditions expected near Mercury. Glenn's Vacuum Facility 6, which is equipped with a solar simulator, can simulate the vacuum and high solar radiation anticipated in Mercury orbit. The MESSENGER test hardware includes a variety of materials and components that are being characterized during the Tank 6 vacuum tests, where the hardware will be exposed to up to 11 suns insolation, simulating conditions expected in Mercury orbit. In 2002, ten solar vacuum tests were conducted, including beginning of life, end of life, backside exposure, and solar panel thermal shock cycling tests. Components tested include candidate solar array panels, sensors, thermal shielding materials, and communication devices. As an example, for the solar panel thermal shock cycling test, two candidate solar array panels were suspended on a lift mechanism that lowered the panels into a liquid-nitrogen-cooled box. After reaching -140 C, the panels were then lifted out of the box and exposed to the equivalent of 6 suns (8.1 kilowatts per square meters). After five cold soak/heating cycles were completed successfully, there was no apparent degradation in panel performance. An anticipated 100-hr thermal shield life test is planned for autumn, followed by solar panel flight qualification tests in winter. Glenn's ongoing support to the MESSENGER program has been instrumental in identifying design solutions and validating thermal performance models under a very aggressive development schedule. The test data have assisted Johns Hopkins engineers in selecting a flight solar array vendor and a thermal shield design. MESSENGER is one in a series of missions in NASA's Discovery Program. Infrared thermography provides data on the thermal gradients in the MESSENGER components during high solar insolation vacuum testing.

Wong, Wayne A.↗

NASA NTRS2021

Computing Risk of Pyrotechnic Devices Using Lot Acceptance Testing

This paper presents an application of statistical engineering to solve the problem of determining the effects of reducing test sample sizes and environment levels in lot acceptance testing (LAT) of ordnance devices used by NASA and DoD space systems. Unlike environmental testing of dedicated qualification hardware, ordnance devices have a test-like-you-fly exception and use lot acceptance sampling or LAT. LAT is performed by randomly selecting a predefined number of devices from an individual lot and exposing them to a more severe dynamic shock and vibration testing environments than flight (sometimes as high as 6 decibels over flight environments) and then performing a destructive test on the device. Recent trends have been suggested to reduce LAT costs by reducing environment test levels or the number of devices per LAT, or both. A method is given here to compute flight risk, so the proposed methods can be compared by subject matter experts to the baseline LAT methods. In addition, this method can be used to determine the best methods to include in future LAT standards for pyrotechnic devices.

Computing Risk↗

NASA NTRS2019

13kW Advanced Electric Propulsion Flight System Development and Qualification

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

in-space propulsion↗

NASA NTRS2019

13kW Advanced Electric Propulsion Flight System Development and Qualification

in-space propulsion↗

NASA NTRS1973

Velocity control propulsion subsystem of the Radio Astronomy Explorer satellite for Goddard Space Flight Center

The Velocity Control Propulsion Subsystem (VCPS) was designed the propulsion required for trajectory and lunar orbit corrections of the spacecraft. A GFE clamp assembly physically attaches the VCPS to the spacecraft and the unit is ejected after completing the required corrections. The VCPS is physically and functionally separated from the spacecraft except for the electrical and telemetry interfaces. A GFE transtage provides the superstructure on which the VCPS is assembled. The subsystem consists of two 5 foot pound rocket engine assemblies, 4 propellant tanks, 2 latching valves, 2 fill and drain valves, a system filter, pressure transducer, gas and propellant manifolds and electrical heaters and thermostats. The RAE-B VCPS program covered the design, manufacture and qualification of one subsystem. This subsystem was to be manufactured, subjected to qualification tests; and refurbished, if necessary, prior to flight. The VCPS design and test program precluded the need for refurbishing the subsystem and the unit was delivered to GSFC at the conclusion of the program.

Braught, W.↗

NASA NTRS1975

Limb radiance inversion radiometer

Engineering and scientific objectives of the LRIR experiment are described along with system requirements, subassemblies, and experiment operation. The mechanical, electrical, and thermal interfaces between the LRIR experiment and the Nimbus F spacecraft are defined. The protoflight model qualification and acceptance test program is summarized. Test data is presented in tables to give an overall view of each test parameter and possible trends of the performance of the LRIR experiment. Conclusions and recommendations are included.

Drozewski, R. W.↗

NASA NTRS2011

Base Flow Investigation of the Apollo AS-202 Command Module

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

Walpot, Louis M. G.↗

NASA NTRS1990

Summary of JAYGO mixing and FSM-1 application of castable inhibitor and liner

Two JAYGO planetary mixers (12 and 42 gallon) are being qualified to mix STW5-3224 liner and STW5-3223 castable inhibitor. These mixers are an integral part of a mix process which allows for safe addition of the asbestos component. An essential part of the engineering evaluation (ETP-0347) of these mixers is the generation of static test fire data. Ultimately, these results will help confirm the adequacy of these mixers for production mixing of liner and inhibitor. (These data are not required for qualification of the Certification Test Plan CTP-0125). The details on the mixing, inhibiting, and sling-lining of JAYGO-mixed castable inhibitor and liner which were applied to the FSM-1 segments are presented. The objectives are the following: (1) to document processing events surrounding the JAYO mixing of castable inhibitor and liner, and the subsequent inhibiting and sling lining onto the FSM-1 segments; and (2) to substantiate the measured properties of these JAYGO-mixed materials (rheological and mechanical) and compare these properties to existing production database.

Evans, Kurt B.↗

NASA NTRS1992

Materials And Processes Technical Information System (MAPTIS) LDEF materials database

The Materials and Processes Technical Information System (MAPTIS) is a collection of materials data which was computerized and is available to engineers in the aerospace community involved in the design and development of spacecraft and related hardware. Consisting of various database segments, MAPTIS provides the user with information such as material properties, test data derived from tests specifically conducted for qualification of materials for use in space, verification and control, project management, material information, and various administrative requirements. A recent addition to the project management segment consists of materials data derived from the LDEF flight. This tremendous quantity of data consists of both pre-flight and post-flight data in such diverse areas as optical/thermal, mechanical and electrical properties, atomic concentration surface analysis data, as well as general data such as sample placement on the satellite, A-O flux, equivalent sun hours, etc. Each data point is referenced to the primary investigator(s) and the published paper from which the data was taken. The MAPTIS system is envisioned to become the central location for all LDEF materials data. This paper consists of multiple parts, comprising a general overview of the MAPTIS System and the types of data contained within, and the specific LDEF data element and the data contained in that segment.

Davis, John M.↗

NASA NTRS2022

Recommendations on the Use of Commercial-Off-The-Shelf (COTS) Electrical, Electronic, and Electromechanical (EEE) Parts for NASA Missions - Phase II

This assessment had two Phases. Phase I captured NASA Centers’ current practices for commercial-off-the-shelf (COTS) Electrical, Electronic, and Electromechanical (EEE) parts 1 used in spaceflight systems and ground support equipment (available at https://ntrs.nasa.gov/citations/20205011579) [ref. 1]. The Phase II report provides guidance for selecting and using COTS parts in NASA missions. The approaches proposed in this report differ from current agency practices. This top-level executive summary touches on these new approaches for using COTS parts but does not provide the detailed information that is critical in understanding the rationale behind these new approaches. Readers will need to read the entire report to gain full understanding and effectively use the recommendations herein. NASA’s historical approach to selecting and applying parts has been to define certain parts, primarily specific classes of military specification (MIL-SPEC) parts, as “standard”, leaving all others, including COTS parts, as nonstandard. Standard parts typically are used without further testing (“use-as-is”). Nonstandard parts are subjected to initial screening and subsequent lot acceptance testing of representative samples from each procured lot per MIL-SPEC or similar requirements. Decades later, top-tier commercial part manufacturers have evolved significant manufacturing, statistical control, and technological improvements that can now provide parts as reliable or more reliable than MIL-SPEC parts, when used within their datasheet limits. Concurrently, the space science and exploration community’s needs demand technological advances unavailable with MIL-SPEC parts. This ongoing change necessitates using COTS parts for space missions. Properly selected COTS parts in appropriate applications can offer performance and supply availability advantages compared to MIL-SPEC parts. Their utility and demonstrated reliability result from large volumes and automated production and testing processes. However, careful review and a thorough understanding of their specifications (i.e., datasheet limitations) is needed, and verifying that manufacturer specifications and reliability meet space hardware application needs are necessary. This report recommends MIL-SPEC screening and non-radiation-related lot acceptance testing be reduced or eliminated in cases where evidence of sufficient quality and reliability exists for COTS parts. The extent of NASA's insight into COTS manufacturers and the amount and nature of the needed evidence will differ by mission and will likely be driven by a mission's resources and associated risk posture. To facilitate this goal, two new terminologies have been defined and described: “Industry Leading Parts Manufacturer (ILPM)” and “Established COTS parts.” An ILPM is a COTS manufacturer that produces high quality and reliable parts. Some parts produced by ILPMs, defined as Established COTS parts, do not need any additional MIL-SPEC or NASA screening and lot acceptance testing to be used in space applications. This report provides guidance for selecting, procuring, and applying COTS parts and for performing part-, board-, and system-level COTS parts verification. The recommendation to select Established COTS parts from ILPMs will assure those COTS parts will have comparable quality to corresponding MIL-SPEC parts. Selecting, applying, and verifying Established COTS parts from ILPMs requires a holistic team approach, engaging parts engineers, circuit designers, quality, reliability, and systems engineers, procurement specialists, radiation specialists, avionics leads, and program/project managers. A mission-specific approach tailored to a project’s Mission, Environment, Applications and Lifetime (MEAL) [ref. 2] requirements should be developed and approved by program/project managers. Any associated risks should be clearly identified, quantified, mitigated, and/or accepted. Different approaches are recommended according to program/project Risk Classes A, B, C, and D [ref. 3] and human-rated missions [ref. 4]: 1. Recommend Classes A and B and human-rated missions consider a “MIL-SPEC parts- based design” approach. ”MIL-SPEC parts-based design” approach is one in which most parts are MIL-SPEC parts and Established COTS parts from ILPMs are used only when an equivalent MIL-SPEC part does not meet functional or size, weight, and power (SWaP) or performance requirements, or is not available. 2. Recommend Classes D and Sub-D missions consider a “System of COTS” approach. “System of COTS” approach is one which most parts are Established COTS parts from ILPMs. 3. Recommend Class C missions determine which approach is the best for their projects; that is, use either a “MIL-SPEC parts-based design” approach, “System of COTS” approach, or a combined approach utilizing elements of both. This report intends to provide guidance in using COTS parts for NASA missions with risk classifications of A through D and human-rated missions; but it does not address the costs of using COTS parts. Costs of using COTS parts in different NASA mission classes can vary significantly even if the same parts are used in different risk postures, due to differing verification levels needed. The guidance does not distinguish between critical or non-critical systems, and a given project will need to apply the appropriate guidance based on their risk posture. The intended audience of this report are NASA personnel and commercial practitioners who support NASA’s spaceflight missions, including spaceflight program or project managers, parts engineers, parts manufacturers, radiation engineers, avionics engineers, system engineers, circuit design engineers, reliability engineers, safety and mission assurance (SMA) personnel, and parts procurement specialists. The NEPP Program will perform a pathfinder study to explore implementing the guidance in this NESC report. An ILPM verification process is not the same as conventional vendor qualification processes performed according to military standards and specifications. This NESC report intends to provide guidance in utilizing available parts data from ILPM manufacturers for parts assurance assessments needed for NASA missions. The report also captured the current practices from DoD and Federal Aviation Administration (FAA) in Section 10. Note each DoD and FAA report was provided by the corresponding agencies regarding their practices, which are independent from the NESC recommendations in the report.

Commercial-Off-The-Shelf↗

NASA NTRSDate not supplied

Process Qualification Concepts in NASA-STD-6030

- Terminology - Definition and role of QMP in the NASA-STD-6030 framework - Qualification checklist for evaluating a candidate AM material process - Types of data needed to evaluate the metrics for a QMP - Advocacy for a singular, contained data entity that enables QMP evaluation and other aspects of the AM ecosystem - Material Engineering Equivalence Baseline - Discussion of Material Engineering Equivalence concepts - Summary

Doug Wells↗

NASA NTRSDate not supplied

High Temperature Devices for Aerospace Applications

This presentation will overview important new capabilities and benefits that next-generation higher temperature electronics promise for a variety of aerospace systems and missions. Initial aerospace infusions of higher temperature electronics to be discussed include jet engines and the prolonged robotic exploration of the Venus surface. Complete redesign of systems compared to prior conventional electronics practices would yield the most revolutionary benefits. However, the inherent higher risk and expense associated with new technology adoption will likely drive a more evolutionary technology infusion process, especially since flight qualification demands large safety margins and testing that ensures reliable operation for far longer than actual mission duration. SiC pn-junction based devices are inherently capable of the most durable operation at the highest application temperatures compared other transistor technologies, but such integrated circuits (ICs) will never achieve integration levels near today’s silicon ICs. Extreme temperature high-power devices will also be de-rated compared to room-temperature performance. While the high temperature market will always be small compared to the conventional temperature mass-market, compatibility and leveraging of existing foundry manufacturing is vital to the establishment of a sustainable high temperature electronics ecosystem sufficient to significantly improve aerospace systems.

High Temperature↗

NASA NTRS2000

Parts Engineering Experiences, Philosophies and Trends

This document is the presentation viewgraphs of the NASA presentations to NASDA, outlining the philosophy and trends of the experiences with engineering parts. Included in the presentations: are (1) the assurance of COTS boards for Space flight, and (2) Peer Review for Ice, Cloud and land Elevation Satellite (ICESat) GPS flight receivers EEE parts. The emphasis is on the methods for qualification of available parts for space flight.

Shaw, Harry↗