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At least 235 records · Page 13

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NASA NTRS1989

Comparison of sine dwell and broadband methods for modal testing

The objectives of modal tests for large complex spacecraft structural systems are outlined. The comparison criteria for the modal test methods, namely, the broadband excitation and the sine dwell methods, are established. Using the Galileo spacecraft modal test and the Centaur G Prime upper stage vehicle modal test as examples, the relative advantage or disadvantage of each method is examined. The usefulness or shortcomings of the methods are given from a practical engineering viewpoint.

Chen, Jay-Chung↗

NASA NTRS1996

Braided Composites for Aerospace Applications. (Latest citations from the Aerospace Database)

The bibliography contains citations concerning the design, fabrication, and testing of structural composites formed by braiding machines. Topics include computer aided design and associated computer aided manufacture of braided tubular and flat forms. Applications include aircraft and spacecraft structures, where high shear strength and stiffness are required.

COMPOSITE STRUCTURES↗

NASA NTRS2005

Ionizing Radiation Environment on the International Space Station: Performance vs. Expectations for Avionics and Material

The role of structural shielding mass in the design, verification, and in-flight performance of International Space Station (ISS), in both the natural and induced orbital ionizing radiation (IR) environments, is reported. Detailed consideration of the effects of both the natural and induced ionizing radiation environment during ISS design, development, and flight operations has produced a safe, efficient manned space platform that is largely immune to deleterious effects of the LEO ionizing radiation environment. The assumption of a small shielding mass for purposes of design and verification has been shown to be a valid worst-case approximation approach to design for reliability, though predicted dependences of single event effect (SEE) effects on latitude, longitude, SEP events, and spacecraft structural shielding mass are not observed. The Figure of Merit (FOM) method over predicts the rate for median shielding masses of about 10g/cm(exp 2) by only a factor of 3, while the Scott Effective Flux Approach (SEFA) method overestimated by about one order of magnitude as expected. The Integral Rectangular Parallelepiped (IRPP), SEFA, and FOM methods for estimating on-orbit (Single Event Upsets) SEU rates all utilize some version of the CREME-96 treatment of energetic particle interaction with structural shielding, which has been shown to underestimate the production of secondary particles in heavily shielded manned spacecraft. The need for more work directed to development of a practical understanding of secondary particle production in massive structural shielding for SEE design and verification is indicated. In contrast, total dose estimates using CAD based shielding mass distributions functions and the Shieldose Code provided a reasonable accurate estimate of accumulated dose in Grays internal to the ISS pressurized elements, albeit as a result of using worst-on-worst case assumptions (500 km altitude x 2) that compensate for ignoring both GCR and secondary particle production in massive structural shielding.

Koontz, Steven L.↗

NASA NTRS1985

Advances and trends in structures and dynamics; Proceedings of the Symposium, Washington, DC, October 22-25, 1984

Among the topics discussed are developments in structural engineering hardware and software, computation for fracture mechanics, trends in numerical analysis and parallel algorithms, mechanics of materials, advances in finite element methods, composite materials and structures, determinations of random motion and dynamic response, optimization theory, automotive tire modeling methods and contact problems, the damping and control of aircraft structures, and advanced structural applications. Specific topics covered include structural design expert systems, the evaluation of finite element system architectures, systolic arrays for finite element analyses, nonlinear finite element computations, hierarchical boundary elements, adaptive substructuring techniques in elastoplastic finite element analyses, automatic tracking of crack propagation, a theory of rate-dependent plasticity, the torsional stability of nonlinear eccentric structures, a computation method for fluid-structure interaction, the seismic analysis of three-dimensional soil-structure interaction, a stress analysis for a composite sandwich panel, toughness criterion identification for unidirectional composite laminates, the modeling of submerged cable dynamics, and damping synthesis for flexible spacecraft structures.

Noor, A. K.↗

NASA NTRS2006

Crew Launch Vehicle (CLV) Upper Stage Configuration Selection Process

The Crew Launch Vehicle (CLV), a key component of NASA's blueprint for the next generation of spacecraft to take humans back to the moon, is being designed and built by engineers at NASA s Marshall Space Flight Center (MSFC). The vehicle s design is based on the results of NASA's 2005 Exploration Systems Architecture Study (ESAS), which called for development of a crew-launch system to reduce the gap between Shuttle retirement and Crew Exploration Vehicle (CEV) Initial Operating Capability, identification of key technologies required to enable and significantly enhance these reference exploration systems, and a reprioritization of near- and far-term technology investments. The Upper Stage Element (USE) of the CLV is a clean-sheet approach that is being designed and developed in-house, with element management at MSFC. The USE concept is a self-supporting cylindrical structure, approximately 115' long and 216" in diameter, consisting of the following subsystems: Primary Structures (LOX Tank, LH2 Tank, Intertank, Thrust Structure, Spacecraft Payload Adaptor, Interstage, Forward and Aft Skirts), Secondary Structures (Systems Tunnel), Avionics and Software, Main Propulsion System, Reaction Control System, Thrust Vector Control, Auxiliary Power Unit, and Hydraulic Systems. The ESAS originally recommended a CEV to be launched atop a four-segment Space Shuttle Main Engine (SSME) CLV, utilizing an RS-25 engine-powered upper stage. However, Agency decisions to utilize fewer CLV development steps to lunar missions, reduce the overall risk for the lunar program, and provide a more balanced engine production rate requirement prompted engineers to switch to a five-segment design with a single Saturn-derived J-2X engine. This approach provides for single upper stage engine development for the CLV and an Earth Departure Stage, single Reusable Solid Rocket Booster (RSRB) development for the CLV and a Cargo Launch Vehicle, and single core SSME development. While the RSRB design has changed since the CLV Project's inception, the USE design has remained essentially a clean-sheet approach. Although a clean-sheet upper stage design inherently carries more risk than a modified design, it does offer many advantages: a design for increased reliability; built-in extensibility to allow for commonality/growth without major redesign; and incorporation of state-of-the-art materials, hardware, and design, fabrication, and test techniques and processes to facilitate a potentially better, more reliable system. Because consideration was given in the ESAS to both clean-sheet and modified USE designs, this paper will highlight the advantages and disadvantages of both approaches and provide a detailed discussion of trades/selections made that led to the final upper stage configuration.

Davis, Daniel J.↗

NASA NTRS1964

Performance evaluation of ''thermo-lag'' material for entry heat protection of advanced manned spacecraft. vol. 3- structural investigations final report, 1 oct. 1962 - 31 jan. 1964

Mechanical properties and structural analysis of Thermo-lag material for entry heat protection of manned spacecraft

MANNED SPACECRAFT↗

NASA NTRS1983

The measurements of vehicle glow on the Space Shuttle

From the combined data set of glow observations on STS-3, STS-4 and STS-5 some of the properties of the shuttle glow were observed. Comparison of the STS-3 (240 km) and STS-5 (305 km) photographs show that the intensity of the glow is about a factor of 3.5 brighter on the low altitude (STS-3) flight. The orbiter was purposely rotated about the x axis in an experiment on STS-5 to observe the dependence of the intensity on the angle of incidence between the spacecraft surface normal and the velocity vector. For a relatively large angle between the velocity vector and the surface normal there is an appreciable glow, provided the surface is not shadowed by some other spacecraft structure. As the angle becomes less the glow intensifies. The grating experiments (STS-4 photography only, STS-5 image intensifier photography) provided a preliminary low resolution spectra of the spacecraft glow. Accurate wavelength calibrations of the STS-5 instrument permitted measuring of the spectrum and intensity of the Earth's airglow.

Mende, S. B.↗

NASA NTRS1983

Measurement of vehicle glow on the Space Shuttle

From the combined data set of glow observations on STS-3, STS-4 and STS-5 some of the properties of the Shuttle glow were observed. Comparison of the STS-3 (240 km) and STS-5 (305 km) photographs show that the intensity of the glow is about a factor of 3.5 brighter on the low altitude (STS-3) flight. The orbiter was purposely rotated about the x axis in an experiment on STS-5 to observe the dependence of the intensity on the angle of incidence between the spacecraft surface normal and the velocity vector. For a relatively large angle between the velocity vector and the surface normal there is an appreciable glow, provided the surface is not shadowed by some other spacecraft structure. As the angle becomes less the glow intensifies. The grating experiments (STS-4 photography only, STS-5 image intensifier photography) provided a preliminary low resolution spectra of the spacecraft glow. Accurate wavelength calibrations of the STS-5 instrument permitted measuring of the spectrum and intensity of the earth's airglow. Previously cited in STAR as N83-30509

Mende, S. B.↗

NASA NTRS1984

Measurements of vehicle glow on the Space Shuttle

Mende, S. B.↗

NASA NTRS1985

Structural analysis model validation using modal test data

The design of a structure which can survive a prescribed dynamic environment is generally based on the use of an analytical model of the structure. The modal test is employed to verify the analytical model. This paper is concerned with the involved validation process, taking into account modal test results from realistic complex spacecraft structural systems as examples to demonstrate the proposed validation procedure. Attention is given to a load analysis model definition, a pretest analysis and test/analysis model, and the correlation parameters. Coordinate transformation is considered along with mode identification, frequency orthogonality, effective mass, mode shape, local kinetic energy, modal force, strain energy, and modal forces by acceleration measurements. Two examples are used to illustrate the procedure.

Chen, J.-C.↗

NASA NTRS2015

Comparison of Analysis with Test for Static Loading of Two Hypersonic Inflatable Aerodynamic Decelerator Concepts

Acceptance of new spacecraft structural architectures and concepts requires validated design methods to minimize the expense involved with technology demonstration via flight-testing. Hypersonic Inflatable Aerodynamic Decelerator (HIAD) architectures are attractive for spacecraft deceleration because they are lightweight, store compactly, and utilize the atmosphere to decelerate a spacecraft during entry. However, designers are hesitant to include these inflatable approaches for large payloads or spacecraft because of the lack of flight validation. This publication summarizes results comparing analytical results with test data for two concepts subjected to representative entry, static loading. The level of agreement and ability to predict the load distribution is considered sufficient to enable analytical predictions to be used in the design process.

Lyle, Karen H.↗

NASA NTRS1971

Toxicology of spacecraft materials

The procedures for determining the toxicity of products outgassed from spacecraft structures are discussed. The test equipment involved in the tests and the criteria for acceptability are described. The use of animals as the final step in determining toxicity of a spacecraft environment is explained.

Harris, E. S.↗

NASA NTRS1985

Material interactions with the low earth orbital environment Accurate reaction rate measurements

Interactions between spacecraft surfaces and atomic oxygen within the low earth orbital (LEO) environment have been observed and measured during Space Shuttle flights over the past 3 yr. The results of these experiments have demonstrated that interaction rates for many materials proposed for spacecraft applications are high and that protective coatings must be developed to enable long-lived operation of spacecraft structures in the LEO environment. A flight experiment discussed herein uses the Space Shuttle as an orbiting exposure laboratory to obtain accurate reaction rate measurements for materials typically used in spacecraft construction. An ion-neutral mass spectrometer, installed in the Orbiter cargo bay, will measure diurnal ambient oxygen densities while material samples are exposed at low altitude (222 km) to the orbital environment. From in situ atomic oxygen density information and postflight material recession measurements, accurate reaction rates can be derived to update the Space Station materials interaction data base. Additionally, gases evolved from a limited number of material surfaces subjected to direct oxygen impingement will be identified using the mass spectrometer. These measurements will aid in mechanistic definitions of chemical reactions which cause atom-surface interactions and in validating results of upcoming degradation studies conducted in ground-based neutral beam laboratories.

Visentine, J. T.↗

NASA NTRS1982

The Magsat mission

The cesium-vapor scalar and fluxgate vector magnetometers aboard the Magsat spacecraft, which has a twilight, sun-synchronous orbit with a 96.76-deg inclination, have together measured the earth's magnetic field magnitude to accuracies better than 2 nT, and the magnitude of each field component to better than 6 nT. The magnetometers are located at the end of a boom to isolate them from the effects of spacecraft fields, and an optical system was used to measure the attitude of the vector magnetometer and sun sensor at the boom relative to the two star cameras of the main spacecraft structure. The data gathered is available from the National Space Science Data Center in several formats and is currently undergoing analysis. Scalar and vector error budget values are given for the spacecraft's five error sources, which include the instrument, position and time errors, digitization noise, attitude errors, and spacecraft fields.

Langel, R.↗

NASA NTRS2017

Additive Manufacturing: An Enabling Technology for the MoonBEAM 6U CubeSat Missions

The Advanced Concepts Office at the NASA Marshall Space Flight Center completed a mission concept study for the Moon Burst Energetics All-sky Monitor (MoonBEAM). The goal of the concept study was to show the enabling aspects that additive manufacturing can provide to CubeSats. In addition to using the additively manufactured tanks as part of the spacecraft structure, the main propulsion system uses a green propellant, which is denser than hydrazine. Momentum unloading is achieved with electric microthrusters, eliminating much of the propellant plumbing. The science mission, requirements, and spacecraft design are described.

Hopkins, R. C.↗

NASA NTRS1992

Summary of ionizing radiation analysis on the Long Duration Exposure Facility

The ionizing radiation measurements flown on the Long Duration Exposure Facility (LDEF) were contained in 15 experiments which utilized passive detectors to pursue objectives in astrophysics and to measure the radiation environment and dosimetric quantities. The spacecraft structure became sufficiently radioactive to permit additional important studies. The induced activity allows extensive radiation mapping in the structure, and independent comparison with experiment dosimetric techniques, and significant studies of secondary effects. The long exposure time, attitude stability, and number and types of measurements produced a unique and critical set of data for low Earth orbit that will not be duplicated for more than a decade. The data allow an unprecedented test, and improvement if required, of models of the radiation environment and the radiation transport methods that are used to calculate the internal radiation and its effects in spacecraft. Results of measurements in the experiments, as well as from radioactivity in the structure, have clearly shown effects from the directional properties of the radiation environment, and progress was made in the dosimetric mapping of LDEF. These measurements have already influenced some Space Station Freedom design requirements. Preliminary results from experiments, reported at this symposium and in earlier papers, show that the 5.8 years exposure considerably enhanced the scientific return of the radiation measurements. The early results give confidence that the experiments will make significant advances in the knowledge of ultra heavy cosmic rays, anomalous cosmic rays, and heavy ions trapped in the radiation belts. Unexpected phenomena were observed, which require explanation. These include stopping iron group ions between the energy ranges anticipated for anomalous and galactic cosmic rays in the LDEF orbit. A surprising concentration of the Be-7 nuclide was discovered on the 'front' surface of LDEF, apparently transported up from the stratosphere with exceptional efficiency.

Parnell, T. A.↗

NASA NTRS1974

Viking Mars lander 1975 dynamic test model/orbiter developmental test model forced vibration test

The Viking Mars Lander 1975 dynamic test model and orbiter developmental test model were subjected to forced vibration sine tests. Flight acceptance (FA) and type approval (TA) test levels were applied to the spacecraft structure in a longitudinal test configuration using a 133,440-N (30,000-lb) force shaker. Testing in the two lateral axes (X, Y) was performed at lower levels using four 667-N (150-lb) force shakers. Forced vibration qualification (TA) test levels were successfully imposed on the spacecraft at frequencies down to 10 Hz. Measured responses showed the same character as analytical predictions, and correlation was reasonably good. Because of control system test tolerances, orbiter primary structure generally did not reach the design load limits attained in earlier static testing. A post-test examination of critical orbiter structure disclosed no apparent damage to the structure as a result of the test environment.

Fortenberry, J.↗

NASA NTRS1981

A CCD tracker for closed-loop instrument pointing at Halley's comet

An optical tracker based on CCD technology has been developed for the closed-loop pointing of the narrow-angle imaging system of the proposed Halley's comet intercept mission at the comet nucleus. The target tracker, needed to assure the accurate pointing of science instruments as the nucleus rapidly passes the spacecraft, is a component of the pointing control system, which also includes a two-axis articulated instrument platform and high-resolution optical encoders to measure platform orientation relative to spacecraft structure. The device was designed on the basis of a CCD imaging array to calculate target centroid position based on the optical intensity of all pixels containing the image, and a processor to perform spacecraft attitude control, pointing control and target tracker control and data processing functions. The tracker has been configured to operate in both target acquisition and tracking modes. Preliminary simulation results and past experiments on similar trackers indicate that excellent tracking performance can be achieved, and a breadboard version of the tracker is currently being fabricated.

Stanton, R. H.↗