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This paper describes the design, implementation, and performance test results of an engineering model of the Position Multiplexer (MUX)-Analog Input Processor (AIP) System for the transmission and continuous measurements of Reflector Control Drive position in SP-100. The specially tailored MUX-AIP combination multiplexes the sensor signals and provides an increase in immunity from low frequency interference by translating the signals up to a higher frequency band. The modulated multiplexed signals are transmitted over a single twisted shielded cable pair from the reflector drives located near reactor to the AIP located at the power conditioning/system controller end of the space craft boom. There the signals are demultiplexed and processed by the AIP, eliminating the need for individual cables for each of the twelve position sensors across the boom.

Recent Generic Flight System (GFS) updates have necessitated revisions in the initial startup and restart control strategies. The design changes that have had the most impact on the control strategies are the addition of the Auxiliary Cooling and Thaw (ACT) system for preheating the lithium filled components, changes in the reactivity worths of the reflectors and safety-rods such that initial cold criticality is achieved with only a small amount of reflector movement following the withdrawal of the safety-rods, and the removal of the scram function from the reflectors. Revised control and operating strategies have been developed and tested using the SP-100 dynamic simulation model, ARIES-GFS. The change in the total reactivity worths of the reflectors and safety-rods has eliminated the need for the use of fast and slow reflector drive speeds during the initial on-orbit approach to criticality.

As part of the SP-100 Space Reactor Power System Program being undertaken for the U. S. Department of Energy, GE is developing a thermoelectric (T/E) power converter which utilizes reactor delivered heat and transforms it into usable electric power by purely static means. This converter is based to GE's product line of successful thermoelectric space power systems. The SP-100 power converter embodies the next generation improvement over the type of T/E converter successfully flown on the six U. S. space missions. That is, conduction coupling of T/E cell to both the heat source and the heat rejection elements. The current technology utilizes radiation coupling in these areas. The conduction coupling technique offers significant improvements in system specific power since it avoids the losses associated with parasitic ΔT's across the radiation gap between the heat source and the hot junction of the thermoelectric (T/E) cell.

The SP-100 Space Reactor Power System is being developed by GE, under contract to the U.S. Department of Energy, to provide electrical power in the range of 10's to 100's of kW. The system represents an enabling technology for a wide variety of earth orbital and interplanetary science missions, nuclear electric propulsion (NEP) stages, and lunar/Mars surface power for the Space Exploration Initiative (SEI). An effective infrastructure of Industry, National Laboratories and Government agencies has made substantial progress since the 1988 System Design Review. Hardware development and testing has progressed to the point of resolving all key technical feasibility issues. The technology and design is now at a state of readiness to support the definition of early flight demonstration missions. Of particular importance is that SP-100 meets the demanding U.S. safety, performance, reliability and life requirements.

The term “SP-100” is synonymous with a set of technologies that can be utilized to provide long lifetime, reliable, safe space power over the range of kilowatts to megawatts [1] using a nuclear reactor as the heat source. This paper describes recent development progress in a number of technology areas such as fuel, materials, reactivity control mechanisms and sensors. Without exception, excellent technical progress is being accomplished in all areas under development to optimize spacecraft performance characteristics.

This paper describes how the SP-100 Space Reactor Power System (SRPS) can be tailored to meet the specific requirements for a lunar surface power system to meet the needs of the consolidation and utilization phases outlined in the 90-day NASA SEI study report. This same basic power system can also be configured to obtain the low specific masses needed to enable robotic interplanetary science missions employing Nuclear Electric Propulsion (NEP). In both cases it is shown that the SP-100 SRPS can meet the specific requirements. For interplanetary NEP missions, performance upgrades currently being developed in the area of light weight radiators and improved thermoelectric material are assumed to be technology ready in the year 2000 time frame. For lunar applications, some system rearrangement and enclosure of critical components are necessary modifications to the present baseline design.

Researchers now have the means to evolve complex manned and unmanned space missions using all of their complex support systems in a fully adaptive visual environment. The expected interactive nature of space missions requires powerful, flexible and comprehensive simulation hardware and software to develop and verify concepts, systems, and procedures. Correlation of visual, sensor, and radar imagery is essential due to new sensor blending and fusion techniques that characterize complex systems and missions. Only through total visual, non-visual and mission environment simulation, combined with analytical tools, can reliable systems and missions be developed. The same can be said of the simulation-based training programs that must be developed for ground and flight mission crews. If maximum situational awareness cannot be trained through simulation, it may be too risky, too expensive or even too late to acquire during a mission.

A new ultrathin-backed semipermeable membrane has been developed which shows considerable promise as a gas separator for engine bleed air to provide nitrogen-rich air for aircraft fuel tank inerting. The membrane is a silicone, polycarbonate copolymer of 1500 Å effective thickness, deposited on a reinforced porous backing. The selective removal of oxygen provides oxygen concentrations of less than 9% in the inerting gas. Small-scale testing demonstrated that the backed membranes are suitable in the aircraft environment. A system using such membranes avoids the logistic and service requirements of tanked liquid nitrogen.

This paper concerns a new technique designed to provide high performance reaction control systems for sounding rockets. Proportional control of differential thrust and simple adaptive control of thrust magnitude (based on the level of demanded thrust) is utilized. The control is being implemented with a combination of electronic and fluidic components for an Aerobee 150 sounding rocket payload whose goal is a pointing stability of 0.1 arc second.

Reliability can be improved by the careful definition of the mechanical properties of engineering materials. Methods to define these properties for design functions by the use of statistics and probability concepts are presented. In addition, methods will be presented for quantitatively measuring the effects of specification screening on the improved properties of the acceptable materials. By selection of the proper design allowables based on required failure rate, reliability can be designed into components using the techniques discussed and illustrated.