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Hazards of interaction of meteoroids with thin metallic shields of spacecraft structures and computations of impact damage for definition of penetration
Glass fiber reinforced plastic material for spacecraft structural components
Structural design of spacecraft subjected to thermal sterilization prior to planetary landing
Conceptual mechanization for horizon definition spacecraft structures and thermal subsystem - spin-stabilized, hexagonal cylinder for launch of two-stage Improved Delta /DSV-3N/
In the design of spacecraft structures, fine tuning the structure to achieve minimum weight with natural frequency constraints is a time consuming process. Here, a novel implementation of the method of optimality criterion (OC) is developed. In this new implementation of OC, the free vibration analysis results are used to compute the eigenvalue sensitivity data required for the formulation. Specifically, the modal elemental strain and kinetic energies are used. Additionally, normalized design parameters are introduced as a second level linking that allows design variables of different values to be linked together. With the use of this novel formulation, synthesis of structures with natural frequency constraint can be carried out manually using modal analysis results. Design examples are presented to illustrate this novel implementation of the optimality criterion method.
Fatigue and response to noise of aircraft and spacecraft viscoelastic panels, shell structure, and payloads
Simulation of meteoroid interaction and impact damage to spacecraft structures
Space simulation experimental studies of penetration mechanics of meteoroids into typical spacecraft structures
An attempt is made to indicate the research studies which should be vigorously supported in order to provide necessary information for the solution of structural dynamic problems of missile and space vehicles. The problem areas are discussed in terms of the disciplines or functions required in their solution. Among the latter are: (1) interactions of the complete system and the environment, (2) criteria for design conditions and performance, (3) interactions of aerodynamic forces with flexible structures, (4) motion of liquids, (5) vibration, (6) impulsive loading and transient responses, (7) guidance and control, (8) testing, and (9) materials considerations. In addition, some correlation is provided for identifying the problems in terms of the environmental conditions in which they occur.
Elastomeric self-sealing structural materials for spacecraft operating in ameteoroid environment
Cassini is a large robotic spacecraft currently under development at the Jet Propulsion Laboratory (JPL) whose interplanetary scientific mission is to explore Saturn, its ring, and its moons. Cassini is scheduled to launch on a Titan IV rocket with a Centaur upper stage booster, and will be protected during ascent through the atmosphere by a lightweight aluminum payload fairing (PLF). As a result of the extreme noise levels generated by the powerful Titan IV at liftoff, and the acoustic transparency of the PLF, Cassini is predicted to experience sever acoustic levels. Furthermore, the high acoustic levels, coupled with the size and configuration of the spacecraft, will induce unacceptable random vibration levels on the structure and critical spacecraft components. Studies showed that the use of Tuned Vibration Absorbers (TVAs) would be effective in reducing vibration. A series of reverberant acoustic tests were performed on a partial development test model (DTM) of Cassini to evaluate the effectiveness of TVAs in reducing the structural vibration. The test results showed that significant vibration attenuation was achieved.
New ideas and techniques recently put to use at the Jet Propulsion Laboratory for structural dynamics of spacecraft are presented. This paper deals with practical problems rather than elaborate mathematical theories and is concerned with the system approach for structural dynamics, an approach which has received attention in the recent past owing to the use of the fast Fourier transform algorithm which permits an economical use of digital computers. Concept of dynamics mass in the frequency domain is introduced. Reaction forces and moments at the base of a spacecraft in boosted flight configuration are determined. A combination of digital and analog techniques for special problems is presented. The examples reported are on actual spacecraft.
The paper discusses a potential application of electric propulsion to perform orbit transfer of a large spacecraft structure to geosynchronous orbit (GEO) from LEO, utilizing a nuclear reactor space power source in the spacecraft on a shared basis. The discussions include spacecraft, thrust system, and nuclear reactor space power system concepts. Emphasis is placed on orbiter payload arrangements, spacecraft launch constraints, and spacecraft LEO assembly and deployment sequences.
Design constraints, design description, handling and fabricating, and shock and vibration testing of SERT 2 spacecraft structures
Spacecraft design for minimization of structure interaction with control system, noting susceptibility of extendible booms to solar environment
This document is the final report for the project entitled, "Multi-Scale Sizing of Lightweight Multifunctional Spacecraft Structural Components," funded under the NRA entitled "Cross-Enterprise Technology Development Program" issued by the NASA Office of Space Science in 2000. The project was funded in 2001, and spanned a four year period from March, 2001 to February, 2005. Through enhancements to and synthesis of unique, state of the art structural mechanics and micromechanics analysis software, a new multi-scale tool has been developed that enables design, analysis, and sizing of advance lightweight composite and smart materials and structures from the full vehicle, to the stiffened structure, to the micro (fiber and matrix) scales. The new software tool has broad, cross-cutting value to current and future NASA missions that will rely on advanced composite and smart materials and structures.
A program implemented by NASA Langley Research Center to develop and evaluate new polymers and polymer matrix composites for spacecraft structural applications is examined. Various polymeric films, moldings, and adhesives are evaluated for resistance to atomic oxygen and high energy electron and UV radiation. Thin films from the poly(arylene ether)s containing phenylphosphine oxide groups and the siloxane-epoxies exhibited minor weight loss compared to Kapton polyimide after exposure. Large doses (greater than 10 exp 9 rads) of electron radiation, simulating 30 yr of exposure in GEO, are found to alter the chemical structure of epoxies by both chain scission and cross-linking. The thermal cycling representative of both LEO and GEO environments can cause microcracking in composites which can in turn affect the dimensional stability and produce mechanical property reductions. The processing and fabrication issues associated with precision composite spacecraft components are also addressed.
The calculated results of a semiempirical model for electron-caused electromagnetic pulse (ECEMP) are compared to the experimental data for three spacecraft geometries. The appropriateness of certain model assumptions which have been employed in the absence of a microscopic theory for dielectric breakdown and associated electron blowoff is discussed. Results are limited to the exterior response of spacecraft structures, although neither the model nor the experiments were limited to the outside problem. Rationales for model assumptions are provided.