Advanced Spacecraft Research Articles

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Thermal Aging Properties of Graphite Fiber Reinforced Peek and Graphite Fiber Reinforced Bmi Composites

Abstract Graphite fiber reinforced poly(ether ether ketone) (PEEK) and graphite fiber reinforced bismaleimide (BMI) composite materials are two kinds of advanced fiber-reinforced polymer matrix composites with good thermal stability and excellent mechanical properties at high temperature. They are currently receiving considerable attention. the main limitation on their application is the lack of knowledge regarding their behaviors during extended use at high temperature. Thermal aging properties are the main parameters for new polymer matrix composites that will be used in advanced spacecraft structural components. From the results of thermal aging effects on the properties—including interlaminar shear strength, drop-weight impact strength, and impact energy—of graphite/PEEK and graphite/BMI composites, it is found that unidirectional graphite fiber reinforced composites retain higher strength compared to multidirectional, and that multidirectional graphite/PEEK composites keep higher property retentions than multidirectional graphite/BMI composites after thermal aging at 190°C. From scanning electron photomicrographs, it is also found that graphite/PEEK composites have better fiber/resin adhesion, intraply adhesion, and microcrack resistance compared to graphite/BMI composites after thermal aging.

A multi-shock concept for spacecraft shielding

The results of an advanced spacecraft shielding program conducted at the NASA Johnson Space Center Hypervelocity Impact Research Laboratory (HIRL) are presented. The results include two new aspects of shielding design: the geometrical configuration and the type of material used for the shield. The geometrical configuration of the shield will be the prime focus of this paper due to its application over a large range of materials. The uniqueness of this concept is in the utilization of a multi-shock (MS) shielding technique where ultra-thin (t s) spaced (ΔS), shield elements are used to repeatedly shock the impacting projectile (diameter d p) to a high enough energy state to cause melting and vaporization at velocities which normally would not produce these results. Although the concept of multi-sheet shields has been proposed and tested many times (Christiansen, 1987; Gehring, 1970; Rajendra and Elfer, 1989; Richardson, 1970), the t s/d p ratio has always been large enough that the shield material has provided a large percentage of the debris plume mass which the back sheet must withstand. This concept does not produce the same results. The low t s/d p adds very little shield material to the debris plume allowing a substantial decrease in the thickness (strength) of the backsheet and the proper spacing between sheets prevents the debris plume from destroying successive sheets prior to the particulates reaching the sheet. The present concept, using aluminum as an analog for comparison to a dual sheet (aluminum) “Whipple shield” results in a 30% reduction in weight. The use of other materials with this concept can result in even greater weight savings. The concept was tested at normal impact, oblique impact, and low velocity impact (2.7 km/s) and performed as well as an equivalent dual sheet shield. The scaling characteristics of the new cincept were tested and verified for impacting projectiles of mass 45 milligrams and 1.27 grams at velocities of 6.7 km/s. The new concept provides a shield which can be tailored to meet many design requirements, produce minimal secondary debris particles, provide a means for designing an augmentable shielding system, and most important reduce the weight of debris shielding.

Assessment of the Tribological Requirements of Advanced Spacecraft Mechanisms

AbstractA survey was conducted of existing technologies for moving mechanical assemblies used in spacecraft applications. The purpose was to identify areas where future requirements for lifetimes in excess of ten years with anticipated speeds, loads, and temperatures might not be satisfied. Some specific mechanisms, such as momentum/reaction wheels, high-speed turbines, pointing and tracking mechanisms, despin mechanisms, and gimbal mechanisms, were identified as areas for potential application of existing but unused technologies. Two major problem areas identified involve boundary-regime lubrication and lubricant supply (active or passive) for long life. Areas where substantial, near-term improvements appear practical include the use of hybrid bearings, new synthetic fluid lubricants, new bearing retainer materials, and properly designed solid-film lubricants.

Performance Trends in Spacecraft Auxiliary Propulsion Systems

A trade study was conducted to compare the performance of two advanced spacecraft auxiliary propulsion system designs. These are a low-thrust (2-5 Ibf) bipropellant system for all on-orbit propulsion functions vs an electrical power augmented hydrazine thruster for north-south stationkeeping in conjunction with a conventional hydrazine system. Both approaches offer a weight savings of approximately 25% over all-catalytic systems for a spacecraft requiring north-south stationkeeping for 7 to 10 years. However, neither system offers any particular performance advantage over the other (i.e., less than 1% difference in total net weight of both configurations). Therefore, system selection criteria must be expanded to include cost, risk, and reliability tradeoffs for upcoming spacecraft missions in the Space Shuttle era. This increased performance issue is of particular importance, because of the potential cost savings associated with reducing the size of auxiliary propulsion systems for spacecraft launched by the Space Shuttle.

An Advanced Spacecraft Attitude Control Prototype: Closed Loop Tests and Architecture Investigations

The future geostationnary communication satellites payloads evolve continuously towards increasing transmission power and pointing requirements to be performed all along long-life (7 to 10 years) missions with a high degree of availability.These evolution trends have to take in account the efforts to cut down the development and recurring costs.The above mentioned items apply for all the satellite ancillary functions and first of all to the attitude control functions; the features required for attitude control systems lead then naturally to a versatile, modular system where the attitude control tasks are performed by digital programmable electronics.The conference paper describes the conception prototyping and tests of an advanced 3 axis attitude control system for geostationnary satellites using as controller a programmable digital electronics called also ACE. The study of ACE integration in a modular, standard environment building the attitude control system architecture have been undertaken.This architecture is organized around a system bus which distributes power and data to all the equipments.

Voyager mission description

The Voyager Project, managed by the Jet Propulsion Laboratory, involves the lauching of two advanced spacecraft to explore the Jovian and Saturnian systems, as well as interplanetary space. The one-month lauch period opens on August 20, 1977, with arrivals at Jupiter in March and July of 1979, and at Saturn in November of 1980 and August of 1981. Gravity-assist swingbys of Jupiter are utilized in order to reduce the lauch energy demands needed to reach Saturn. In addition, a gravity-assist targeting option at Saturn will be maintained on the second-arriving Voyager for a possible continuation on to Uranus, with arrival in January of 1986. Flight through the Jovian and Saturnian systems will achieve close to moderate flyby encounters with several of the natural satellites, including special flyby geometry conditions for Io and Titan, as well as an Earth occultation of the spacecraft's radio signal by the rings of Saturn. The purpose of this paper is to describe the Voyager mission characteristics in order to establish a framework upon which to better understand the objectives and goals of the eleven scientific investigations which are described in subsequent papers.

Development of Dual Bumper Wall Construction for Advanced Spacecraft

This paper describes research and development of the dual bumper wall construction for application to advanced spacecraft. Theoretical penetration mechanics were applied to size single and dual bumper test articles containing various amounts of multilayer insulation. Hypervelocity particle impact testing of candidate wall constructions was completed and the effectiveness established. The dual bumper wall was found to be over twice as efficient as the single bumper wall in resisting complete perforation, for impact velocities where projectile fragmentation occurs and where rear sheet penetration by these fragments is the major source of damage. Quantitative data are presented on the increased penetration resistance afforded by multilayer thermal insulation.

Advanced Data Systems for the Grand Tour

A data system technology is described which relies on standby redundancy, automated flexibility, and extensive use of LSI devices to achieve the desired mission results. The STAR computer is the heart of the system and is truly an advanced next-generation spacecraft computer. Extensive use of large-scale integrated circuits has made it possible to implement the needed complexity throughout the advanced data system for the Grand Tour mission.

Advanced spacecraft stabilization and control techniques.

Covers advancements in spacecraft and tactical and strategic missile systems, including subsystem design and application, mission design and analysis, materials and structures, developments in space sciences, space processing and manufacturing, space operations, and applications of space technologies to other fields.

Malfunction Detection System for Advanced Spacecraft

A design approach for a Malfunction Detection System (MDS) to beused on board a manned spacecraft is considered. This design is directed toward orbiting laboratory type vehicles, in which man will be subjected to hostile environments for extended periods of time. The prime purpose of the MDS will be to monitor and evaluate critical signals and display out-of-tolerance conditions. This paper initially discusses a mathematical model approach to a monitoring-system design. Another topic covered is the feasibility of using some value less than the specified tolerance as a signal "test" limit. Finally, various design decisions and trade-offs involved in choosing a monitoring subsystem configuration are discussed, and the conclusion is that the approach is feasible and could be expanded to include other capabilities.

Fluid Mechanics of Zero-G Propellant Transfer in Spacecraft Propulsion Systems

The basic requirement of a transfer system for storable propellants is to provide propellant in sufficient quantity to assure proper engine operation. The main propulsion systems of advanced spacecraft will make many restarts after long coast periods during which the propellant is in a weightless state. Various vehicle motions, such as rotational maneuvers, may result in dislocation of propellant from the region of the outlet of un-baffled tanks. In this paper a promising method of propellant relocation, settling by means of an induced gravitational field, is considered. Small axially directed rockets are used to generate a low g-field for settling the propellant. The value of the acceleration that must be exceeded to assure propellant flow from top to bottom of the tanks (the critical acceleration) has been determined by a numerical solution of the interface equation. This acceleration is exceeded by several orders of magnitude for practical settling rocket systems. The fluid mechanics of the ensuing “settling” flow and the subsequent rebound or “geysering” is discussed and a criterion for settling time is suggested. These results were obtained through scaled experiments conducted at 1-g conditions.

Sonic-Induced Vibration and Control for Advanced Spacecraft

Advanced spacecraft and space stations are designed to remain in space for a prolonged period of time. In some cases, the structure of these vehicles can be excited by modulation of the internally pressurized media and externally exposed to space vacuum. In other cases, the structure is immersed in vacuum and excited externally by the impingement of the control jets. In either case, atmospheric damping is absent from the side of the structure exposed to vacuum. This problem is investigated theoretically and experimentally in terms of three types of specimen structural panels: aluminum, viscoelastic, and honeycomb. With one side of the panels in near vacuum, sonic excitation was applied to the other side. This excitation consisted of (1) constant-amplitude sinusoid, (2) random-amplitude sinusoid, (3) constant amplitude but random frequency, and (4) random amplitude and frequency. These forcing functions simulate the sonic excitation induced by various sources, such as equipment noise, multiple spacecraft dockings, micrometeor shower, space-station repair work, control jets, etc. The methods considered for controlling these sonic-induced vibrations include the altering of resonances response due to mass coupling, and the attenuation of response due to viscoelastic and aquaplus damping. The relative damping efficiency of treatments embracing different areas of the aquaplus coating was also investigated. The effects of hysteresis and air-damping on a panel of arbitrary curvature can be expressed in a generalized mathematical equation.