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The cassini Power Subsystem provides power generation, management and distribution functions for the Saturn Orbiter Spacecraft. Power is derived from three parallel Radioisotope Thermoelectric Generators.
Performance and reliability studies of solar conversion power supply subsystem for Nimbus-B satellite
Nimbus-B satellite solar conversion power supply subsystem operation
Solar conversion power supply subsystem for Nimbus D satellite, and engineering tests
The power requirement of vehciles desgined to transport cargo supporting a piloted expedition to Mars is in the range of megawatts. Therefore, it is imperative that the megawatt - class power processing unit designed for high-power nuclear electric propulsion vehicles using turboalternators and advanced magnetoplasmadynamic thrusters shouls be such that the overall system efficienty is as high as possible with minimum system specific mass.
In the near future, Air Force mission payloads will require significant increases in power. Sophisticated sensing systems such as infrared focal plane detector arrays and radar will be employed by the Air Force to fulfill its strategic objectives. These payloads will demand that the power subsystem provide up to 50 kW at the end of mission life, more than an order of magnitude greater than is currently required. Some of these payloads must be flown in low-Earth polar orbits to satisfy mission objectives, and it is likely that large (500 to 600 sq m) solar photovoltaic arrays will operate in the low-Earth polar environment. The standard 28 volt power subsystem is not weight efficient for the array power levels being considered. The impact of the solar array operating voltage on the total weight of the array and the subsystem power conditioning and distribution components is illustrated. In the interest of reducing power subsystem weight, higher array operating voltages are considered. The problems which the higher array voltage present to the array designer are discussed. In order to provide a maximum return on the tremendous investment of resources required to develop and place these assets in orbit, they must be designed to operate effectively for extended periods of time. To achieve this, the system must be able to function in the threat-induced and natural space environment.
Photovoltaic systems, the state of the art of power conditioning subsystem components, and the design and operational interaction between photovoltaic systems and host utilities are detailed in this document. Major technical issues relating to the design and development of power conditioning systems for photovoltaic application are considered; these include: (1) standards, guidelines, and specifications; (2) cost effective hardware design; (3) impact of advanced components on power conditioning development; (4) protection and safety; (5) quality of power; (6) system efficiency; and (7) system integration with the host utility. Theories of harmonic distortion and reactive power flow are discussed, and information about power conditioner hardware and manufacturers is provided.
The advanced Television Infrared Observation satellite program is a cooperative effort between the National Aeronautics and Space Administration (NASA), the National Oceanic and Atmospheric Administration (NOAA), the United Kingdom, Canada and France, for providing day and night global environmental and associated data. NASA is responsible for procurement launch, and checkout of these spacecraft before transferring them over to NOAA, who operates the spacecraft to support weather forecasting, severe storm tracking, and 'meteorological research by the National Weather Service. These spacecraft with all weather monitoring instruments imposed challenging requirements for the onboard electrical power subsystem (EPS). This paper provides first a brief overview of the overall power subsystem, followed by a description of batteries. A unique power subsystem design which provides 'tender-loving-care' to these batteries is highlighted. This is followed by the on-orbit maintenance and performance data of the batteries since launch.
Manufacturing and qualification testing of solar conversion power supply subsystem for Nimbus 4
Autonomous power management has been proposed as a method to perform optimization of power subsystem performance in connection with the management of multikilowatt space platforms. A concept for a 250-kW utility-type power subsystem was developed. A Cassegrain concentrator solar array primary source is conditioned by a solar array switching unit which supplies seventeen 220 +20 Vdc power channels. A power management subsystem provides the monitoring and control of the overall electrical power subsystem. The discussed system concept for autonomous management of high power space platforms utilizes on-board microprocessors in a decentralized data management architecture. A data bus protocol and a data bus contention resolution scheme were selected in conjunction with the dencentralized management architecture.
The requirements, capabilities, and design of the Space Shuttle Orbiter electrical power generation (EPG) subsystem are presented. The Orbiter EPG subsystem is designed to have the flexibility to supply the basic Orbiter electrical loads and the power requirements of the payloads which, in some cases, are completely dependent on the Orbiter for both power and heat rejection. These needs are supplied by three hydrogen/oxygen fuel cell powerplants (FCPs), having the capability of providing a total of 14 kW average and up to 24 kW peak in the basic Orbiter configuration. Kits permit dedication of one FCP to the payload by providing an additional reactant tank pair for a seven-day mission and additional heat rejection capability.
NASA and military applications for high power spacecraft are presently being planned. A perspective of autonomous power management for low earth orbit large space platforms is presented. A multichannel 250 kWe utility-type power subsystem is used as a baseline system. The need for automation is reviewed, based on power subsystem complexity, survivability requirements, and cost benefits. On-board versus ground management is discussed with respect to these needs. In addition, the utilization of autonomous power management to enable technology readiness of large complex power subsystems is described. Recommendations are made for further technology thrusts.
Concepts for the 1990's space station envision an initial operational capability with electrical power output requirements of approximately 75 kW and growth power requirements in the range of 300 kW over a period of a few years. Photovoltaic and solar dynamic power generation techniques are contenders for supplying this power to the space station. A study was performed to identify growth power subsystem impacts on other space station subsystems. Subsystem interactions that might suggest early design changes for the space station were emphasized. Quantitative analyses of the effects of power subsystem mass and projected area on space station controllability and reboost requirements were conducted for a range of growth station configurations. Impacts on space station structural dynamics as a function of power subsystem growth were also considered.
The results of the Independent Orbiter Assessment (IOA) of the Failure Modes and Effects Analysis (FMEA) and Critical Items List (CIL) are presented. The IOA approach features a top-down analysis of the hardware to determine failure modes, criticality, and potential critical items. To preserve independence, this analysis was accomplished without reliance upon the results contained within the NASA FMEA/CIL documentation. This report documents the independent analysis results corresponding to the Orbiter Electrical Power Distribution and Control (EPD and C)/Electrical Power Generation (EPG) hardware. The EPD and C/EPG hardware is required for performing critical functions of cryogenic reactant storage, electrical power generation and product water distribution in the Orbiter. Specifically, the EPD and C/EPG hardware consists of the following components: Power Section Assembly (PSA); Reactant Control Subsystem (RCS); Thermal Control Subsystem (TCS); Water Removal Subsystem (WRS); and Power Reactant Storage and Distribution System (PRSDS). The IOA analysis process utilized available EPD and C/EPG hardware drawings and schematics for defining hardware assemblies, components, and hardware items. Each level of hardware was evaluated and analyzed for possible failure modes and effects. Criticality was assigned based upon the severity of the effect for each failure mode.
A potassium-Rankine power conversion system model was developed under Contract No. NAS3-25808 for the NASA-LeRC. This model predicts potassium-Rankine performance for turbine inlet temperatures (TIT) from 1200 - 1600 K, TIT to condenser temperature ratios from 1.25-1.6, power levels from 100 to 10,000 kWe, and lifetimes from 2-10 years. The model is for a Rankine cycle with reheat for turbine stage moisture control. The model assumes heat is supplied from a lithium heat transport loop. The model does not include a heat source or a condenser/heat rejection system model. These must be supplied by the user.
The launches of the International Sun-Earth Explorers in October 1977 (ISEE-A) and August 1978 (ISEE-C) marked the first successful implementation of an electrostatically clean spacecraft design on a US-built satellite. The power subsystem design selected was required to operate without induced or coupled electromagnetic interference while meeting the criteria of low cost, low weight (with the resulting removal of almost all redundancy), modular construction techniques, long life (more than 3 years), and maximum utilization of previously qualified/flown designs. To save money, both the ISEE-A and -C power subsystem designs had to be identical even though the two missions are flown in vastly different orbits. Additionally, the requirement for a three year mission utilizing a single silver-cadmium battery had never been imposed before. A power subsystem configuration which met all of the specified requirements was developed. Excellent correlation between preflight and actual flight performance is demonstrated.
Solar auxiliary power supply unit using nickel cadmium storage batteries for Nimbus 4 satellite
A breadboard power system incorporating autonomous functions of monitoring, fault detection and recovery, command and control was developed, tested and evaluated to demonstrate technology feasibility. Autonomous functions including switching of redundant power processing elements, individual load fault removal, and battery charge/discharge control were implemented by means of a distributed microcomputer system within the power subsystem. Three local microcomputers provide the monitoring, control and command function interfaces between the central power subsystem microcomputer and the power sources, power processing and power distribution elements. The central microcomputer is the interface between the local microcomputers and the spacecraft central computer or ground test equipment.