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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 199 records · Page 11

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NASA NTRS1974

SEASAT study documentation

The proposed spacecraft consists of a bus module, containing all subsystems required for support of the sensors, and a payload module containing all of the sensor equipment. The two modules are bolted together to form the spacecraft, and electrical interfaces are accomplished via mated connectors at the interface plane. This approach permits independent parallel assembly and test operations on each module up until mating for final spacecraft integration and test operations. Proposed program schedules recognize the need to refine sensor/spacecraft interfaces prior to proceeding with procurement, reflect the lead times estimated by suppliers for delivery of equipment, reflect a comprehensive test program, and provide flexibility for unanticipated problems. The spacecraft systems are described in detail along with aerospace ground equipment, ground handling equipment, the launch vehicle, imaging radar incorporation, and systems tests.

Source record↗

NASA NTRS1985

A swirl flow evaporative cold plate

A forced flow evaporative cold plate is under development for future application to the thermal bus concept being pursued by NASA for Space Station Thermal Control. The vaporizer is a swirl-flow device employing a spiral tube coil geometry sandwiched between conductive metal plates upon which electric components could be mounted. This concept is based on the inherent phase separation that occurs in a two phase stream in curvilinear flow. This is a zero 'g' design with one 'g' all-attitude capability and is capable of high heat transfer coefficients, good isothermality, and the ability to function at heat fluxes approaching 5w/sq cm on the cold plates (10w/sq cm on the tube wall) with Freon 114. The advantages of this design over other two phase evaporator approaches are high heat flux capability, simplified control requirements, insensitivity to micro-gravity oscillations, and inexpensive manufacturability. The program included design, fabrication, and test of such a cold plate utilizing an existing test stand developed for two-phase thermal management system (TPTMS) testing. Test results analysis and conclusions are included.

Niggemann, R. E.↗

NASA NTRS1980

Application of fiber optics in spacecraft

The unique requirements placed on an optical fiber data system in space are briefly reviewed in this paper. These include temperature extremes, particle radiation, and, in some cases, a stringent electrical power budget. An overriding need is for the best possible reliability. Results related to the power consumption of a fiber terminal pair are reported, with the conclusion that a 32-port transmission star bus configuration could be operated error-free at 10 Mbits/s on 1/8 watt per transmitting terminal. The design of a transmitter terminal optimized for low idling power consumption is discussed. Although the electronic elements on a spacecraft are protected, the fiber cable may be subjected to large temperature extremes. Measured effects of temperature on fibers and cables are discussed and related to bus design. Radiation effects are briefly reviewed, with special reference to the long-duration low dose rate environment of a spacecraft.

Johnston, A. R.↗

NASA NTRS2002

Directionally Sensitive Silicon Radiation Sensor (VCELL)

Sensors are a mission critical element in many NASA programs and require some very unique properties such as small size, low power, high reliability, low weight. Low cost sensors offer the possibility of technology transfer to the public domain for commercial applications. One sensor application that is important to many NASA programs is the ability to point at a radiation source, such as the sun. Such sensors may be an integral part of the guidance and control systems in space platforms and in remote exploratory vehicles. Sun/solar pointing is also important for ground-based systems such as solar arrays. These systems are not required to be small and lightweight. However, if a sensor with a sun pointing capability was developed that is very small, rugged, lightweight and at the same time low cost, it certainly could be used in existing and perhaps many new ground based applications. The objective of the VCELL (Directionally Sensitive Silicon Radiation Sensor) research is to develop a new and very unique silicon based directionally sensitive radiation sensor which can be fabricated using conventional monolithic IC technologies and which will meet the above requirements. The proposed sensor is a novel silicon chip that is directionally sensitive to incident radiation, providing azimuth and elevation information on the incident radiation. The resulting sensor chip will be appropriate for integration into a silicon IC or useful in a hybrid structure to be interfaced with a standard IEEE 1451 bus interface IC to create an Intelligent Sensor. It is presently estimated that it will require about three man-years of effort to complete the VCELL research and development. This includes the optical, electrical, mechanical and silicon fabrication and testing as well as computer simulations and theoretical analysis and modeling including testing in simulated space environments. This report summarizes the sensor research completed this summer as part of the Summer Faculty Fellowship Program. The primary effort was focused on activity necessary to fabricate prototype sensor.

Koy B. Cook↗

NASA NTRS2004

Development of Electronic Load Controllers for Free-Piston Stirling Convertors Aided by Stirling Simulation Model

The free-piston Stirling convertor end-to-end modeling effort at the NASA Glenn Research Center has produced a software-based test bed in which free-piston Stirling convertors can be simulated and evaluated. The simulation model includes all the components of the convertor: the Stirling cycle engine, heat source, linear alternator, controller, and load. So far, it has been used in evaluating the performance of electronic controller designs. Three different controller design concepts were simulated using the model: 1) Controllers with parasitic direct current loading. 2) Controllers with parasitic alternating current loading. 3) Controllers that maintain a reference current. The free-piston Stirling convertor is an electromechanical device that operates at resonance. It is the function of the electronic load controller to ensure that the electrical load seen by the machine is always great enough to keep the amplitude of the piston and alternator oscillation at the rated value. This is done by regulating the load on the output bus. The controller monitors the instantaneous voltage, regulating it by switching loads called parasitic loads onto the bus whenever the bus voltage is too high and removing them whenever the voltage is too low. In the first type of controller, the monitor-ing and switching are done on the direct-current (dc) bus. In the second type, the alternating current bus is used. The model allows designers to test a controller concept before investing time in hardware. The simulation code used to develop the model also offers detailed models of digital and analog electronic components so that the resulting designs are realistic enough to translate directly into hardware circuits.

Regan, Timothy F.↗

NASA NTRS1974

The Atmosphere Explorer power subsystem

The present work describes the design and in-flight performance of the power subsystem for the first three Atmosphere Explorer (AE) spacecraft. The subsystem provides all electrical power to the spacecraft loads and consists of a solar array, three batteries, and a power supply electronics unit. Power from the solar array is supplied to the spacecraft loads through an unregulated power bus, which varies between -26 and -38 V. Loads which are normally on only during data collection phases of an orbit are fed from a tightly regulated -24.5 V bus. Power subsystem performance during the first three months in orbit has been nominal, except for an anomally in the shunt limiter protection circuitry. A malfunctioning shunt control amplifier was reconnected via ground command.

Obenschain, A.↗

NASA NTRS1984

Motor/generator and electronic control considerations for energy storage flywheels

A spacecraft electric power supply system is described. Requirements of the system are to accelerate a momentum wheel to a fixed maximum speed when solar energy is available and to maintain a constant voltage on the spacecraft bus under varying loads when solar energy is not available. Candidate motor types, pulse width modulated current control systems, and efficiency considerations are discussed. In addition, the Lunar Roving Vehicle motors are described along with their respective efficiencies.

Nola, F. J.↗

NASA NTRSDate not supplied

SR-1 Freedom Thermal Architecture Challenges

The SR-1 Freedom mission aims to demonstrate nuclear electric propulsion (NEP) and deliver the Skyfall helicopter payload to Mars. Repurposing the existing Gateway Power and Propulsion Element (PPE) spacecraft and combining it with a 20 kWe-class nuclear power module (NPM) presents a variety of thermal architecture challenges. These include unprecedented waste heat rejection requirements, integration with a spacecraft bus originally designed for a different mission profile, and novel packaging constraints. This presentation will describe the simplified concept of operations as it relates to on-orbit thermal environments, early-phase architecture trades and supporting analyses, and planned forward work. We will also discuss our strategy for integrated thermal modeling of the complete spacecraft and examples of interface challenges necessitated by this ambitious effort.

Thermal↗

NASA NTRS2013

30-kW SEP Spacecraft as Secondary Payloads for Low-Cost Deep Space Science Missions

The Solar Array System contracts awarded by NASA's Space Technology Mission Directorate are developing solar arrays in the 30 kW to 50 kW power range (beginning of life at 1 AU) that have significantly higher specific powers (W/kg) and much smaller stowed volumes than conventional rigid-panel arrays. The successful development of these solar array technologies has the potential to enable new types of solar electric propulsion (SEP) vehicles and missions. This paper describes a 30-kW electric propulsion vehicle built into an EELV Secondary Payload Adapter (ESPA) ring. The system uses an ESPA ring as the primary structure and packages two 15-kW Megaflex solar array wings, two 14-kW Hall thrusters, a hydrazine Reaction Control Subsystem (RCS), 220 kg of xenon, 26 kg of hydrazine, and an avionics module that contains all of the rest of the spacecraft bus functions and the instrument suite. Direct-drive is used to maximize the propulsion subsystem efficiency and minimize the resulting waste heat and required radiator area. This is critical for packaging a high-power spacecraft into a very small volume. The fully-margined system dry mass would be approximately 1120 kg. This is not a small dry mass for a Discovery-class spacecraft, for example, the Dawn spacecraft dry mass was only about 750 kg. But the Dawn electric propulsion subsystem could process a maximum input power of 2.5 kW, and this spacecraft would process 28 kW, an increase of more than a factor of ten. With direct-drive the specific impulse would be limited to about 2,000 s assuming a nominal solar array output voltage of 300 V. The resulting spacecraft would have a beginning of life acceleration that is more than an order of magnitude greater than the Dawn spacecraft. Since the spacecraft would be built into an ESPA ring it could be launched as a secondary payload to a geosynchronous transfer orbit significantly reducing the launch costs for a planetary spacecraft. The SEP system would perform the escape from Earth and then the heliocentric transfer to the science target.

Dawn spacecraft↗

NASA NTRS1995

Helmet-Mounted Display Research Capabilities of the NASA/Army Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL)

The Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL) is a UH-60A Black Hawk helicopter that is being modified by the US Army and NASA for flight systems research. The principal systems that are being installed in the aircraft are a Helmet Mounted Display (HMD) and imaging system, and a programmable full authority Research Flight Control System (RFCS). In addition, comprehensive instrumentation of both the rigid body of the helicopter and the rotor system is provided. The paper will describe the capabilities of these systems and their current state of development. A brief description of initial research applications is included. The wide (40 X 60 degree) field-of-view HMD system has been provided by Kaiser Electronics. It can be configured as a monochromatic system for use in bright daylight conditions, a two color system for darker ambients, or a full color system for use in night viewing conditions. Color imagery is achieved using field sequential video and a mechanical color wheel. In addition to the color symbology, high resolution computer-gene rated imagery from an onboard Silicon Graphics Reality Engine Onyx processor is available for research in virtual reality applications. This synthetic imagery can also be merged with real world video from a variety of imaging systems that can be installed easily on the front of the helicopter. These sensors include infrared or tv cameras, or potentially small millimeter wave radars. The Research Flight Control System is being developed for the aircraft by a team of contractors led by Boeing Helicopters. It consists of a full authority high bandwidth fly-by-wire actuators that drive the main rotor swashplate actuators and the tail rotor actuator in parallel. This arrangement allows the basic mechanical flight control system of the Black Hawk to be retained so that the safety pilot can monitor the operation of the system through the action of his own controls. The evaluation pilot will signal the fly-by-wire actuators through the flight computer from electrical sidearm controllers located in the right hand cockpit. The system will have very substantial input/output capacity and impressive computational power. These systems are installed in the aircraft using predominantly a MIL-STD 1553B data bus architecture. Sensor data from the RFCS, the basic aircraft and rotor system instrumentation including navigation information, and the HMD system are easily exchanged among user systems, or are available at the systems operator station located in the cabin for real time monitoring or data recording.

Jacobsen, R. A.↗

NASA NTRS2007

SiC Multi-Chip Power Modules as Power-System Building Blocks

The term "SiC MCPMs" (wherein "MCPM" signifies "multi-chip power module") denotes electronic power-supply modules containing multiple silicon carbide power devices and silicon-on-insulator (SOI) control integrated-circuit chips. SiC MCPMs are being developed as building blocks of advanced expandable, reconfigurable, fault-tolerant power-supply systems. Exploiting the ability of SiC semiconductor devices to operate at temperatures, breakdown voltages, and current densities significantly greater than those of conventional Si devices, the designs of SiC MCPMs and of systems comprising multiple SiC MCPMs are expected to afford a greater degree of miniaturization through stacking of modules with reduced requirements for heat sinking. Moreover, the higher-temperature capabilities of SiC MCPMs could enable operation in environments hotter than Si-based power systems can withstand. The stacked SiC MCPMs in a given system can be electrically connected in series, parallel, or a series/parallel combination to increase the overall power-handling capability of the system. In addition to power connections, the modules have communication connections. The SOI controllers in the modules communicate with each other as nodes of a decentralized control network, in which no single controller exerts overall command of the system. Control functions effected via the network include synchronization of switching of power devices and rapid reconfiguration of power connections to enable the power system to continue to supply power to a load in the event of failure of one of the modules. In addition to serving as building blocks of reliable power-supply systems, SiC MCPMs could be augmented with external control circuitry to make them perform additional power-handling functions as needed for specific applications: typical functions could include regulating voltages, storing energy, and driving motors. Because identical SiC MCPM building blocks could be utilized in a variety of ways, the cost and difficulty of designing new, highly reliable power systems would be reduced considerably. Several prototype DC-to-DC power-converter modules containing SiC power-switching devices were designed and built to demonstrate the feasibility of the SiC MCPM concept. In anticipation of a future need for operation at high temperature, the circuitry in the modules includes high-temperature inductors and capacitors. These modules were designed to be stacked to construct a system of four modules electrically connected in series and/or parallel. The packaging of the modules is designed to satisfy requirements for series and parallel interconnection among modules, high power density, high thermal efficiency, small size, and light weight. Each module includes four output power connectors two for serial and two for parallel output power connections among the modules. Each module also includes two signal connectors, electrically isolated from the power connectors, that afford four zones for signal interconnections among the SOI controllers. Finally, each module includes two input power connectors, through which it receives power from an in-line power bus. This design feature is included in anticipation of a custom-designed power bus incorporating sockets compatible with snap-on type connectors to enable rapid replacement of failed modules.

Lostetter, Alexander↗

NASA NTRS2001

Sub-Kilowatt Radioisotope Electric Propulsion for Outer Solar System Exploration

An activity has begun to perform a multi-mission trajectory/systems analysis study that examines the combined benefit of three key technology development areas. The three key technologies leveraged in the study include sub-kilowatt ion propulsion, Stirling radioisotope power systems and microelectronics/lightweight spacecraft bus technologies. This study is being performed jointly by NASA Glenn Research Center and the Applied Physics Laboratory to leverage their combined areas of expertise in advanced power/propulsion and spacecraft design. Missions examined in this study include missions to outer planets' moons, such as Europa, Titan, and Triton, and a Comet Nucleus Sample Return mission. Additional information is contained in the original extended abstract.

Gefert, L. P.↗

NASA NTRS2004

A 100 kW-Class Technology Demonstrator for Space Solar Power

A first step in the development of solar power from space is the flight demonstration of critical technologies. These fundamental technologies include efficient solar power collection and generation, power management and distribution, and thermal management. In addition, the integration and utilization of these technologies into a viable satellite bus could provide an energy-rich platform for a portfolio of payload experiments such as wireless power transmission (WPT). This paper presents the preliminary design of a concept for a 100 kW-class free-flying platform suitable for flight demonstration of technology experiments. Recent space solar power (SSP) studies by NASA have taken a stepping stones approach that lead to the gigawatt systems necessary to cost-effectively deliver power from space. These steps start with a 100 kW-class satellite, leading to a 500 kW and then a 1 MW-class platform. Later steps develop a 100 MW bus that could eventually lead to a 1-2 GW pilot plant for SSP. Our studies have shown that a modular approach is cost effective. Modular designs include individual laser-power-beaming satellites that fly in constellations or that are autonomously assembled into larger structures at geosynchronous orbit (GEO). Microwave power-beamed approaches are also modularized into large numbers of identical units of solar arrays, power converters, or supporting structures for arrays and microwave transmitting antennas. A cost-effective approach to launching these modular units is to use existing Earth-to-orbit (ETO) launch systems, in which the modules are dropped into low Earth orbit (LEO) and then the modules perform their own orbit transfer to GEO using expendable solar arrays to power solar electric thrusters. At GEO, the modules either rendezvous and are assembled robotically into larger platforms, or are deployed into constellations of identical laser power-beaming satellites. Since solar electric propulsion by the modules is cost-effective for both self-transport of the modules from LEO to GEO, and for on-orbit stationkeeping and repositioning capability during the satellite’s lifetime, this technology is also critical in technology development for SSP. The 100 kW-class technology demonstrator will utilize advanced solar power collection and generation technologies, power management and distribution, advanced thermal management, and solar electric propulsion. State-of-the-art solar concentrators, highly efficient multi-junction solar cells, integrated thermal management on the arrays, and innovative deployable structure design and packaging make the 100 kW satellite feasible for launch on one existing launch vehicle. Early SSP studies showed that a major percentage of the on-orbit mass for power-beaming satellites was from massive power converters at the solar arrays, at the bus, at the power transmitter, or at combinations of these locations. Higher voltage mays and power management and distribution (PMAD) systems reduce or eliminate the need for many of these massive power converters, and could enable direct-drive of high-voltage solar electric thrusters. Lightweight, highly efficient thermal management systems are a critical technology that must be developed and flown for SSP feasibility. Large amounts of power on satellites imply that large amounts of waste heat will need to be managed. In addition, several of the more innovative lightweight configurations proposed for SSP satellites take advantage of solar concentrators that are intractable without advanced thermal management technologies for the solar arrays. These thermal management systems include efficient interfaces with the WPT systems or other high-power technology experiments, lightweight deployable radiators that can be easily integrated into satellite buses, and efficient reliable thermal distribution systems that can pipe heat from the technology experiments to the radiators. In addition to demonstrating the integration and use of these mission-critical technologies, the 100 kw-class satellite will provide a large experiment deck for a portfolio of technology experiments. Current plans for this technology demonstrator allow 2000 kg of payload capability and up to 100 kW of power. The technology experiments could include one or more wireless power transmission demonstrations, either to the Earth’s surface or to a suitable space-based receiver. Technology experiments to quantify the on-orbit performance of critical technologies for SSP or space exploration are welcomed. In addition, the technology experiments provide an opportunity for international cooperation, to advance technology readiness levels of SSP technologies that require flight demonstration. This paper will present the preliminary design for a 100 kW solar-powered satellite and a variety of technology experiments that may be suitable for flight demonstration. In addition, a space-to-Earth-surface WPT experiment will be discussed.

Connie Carrington↗

NASA NTRS1984

Electronic scanning pressure measuring system and transducer package

An electronic scanning pressure system that includes a plurality of pressure transducers is examined. A means obtains an electrical signal indicative of a pressure measurement from each of the plurality of pressure transducers. A multiplexing means is connected for selectivity supplying inputs from the plurality of pressure transducers to the signal obtaining means. A data bus connects the plurality of pressure transducers to the multiplexing means. A latch circuit is connected to supply control inputs to the multiplexing means. An address bus is connected to supply an address signal of a selected one of the plurality of pressure transducers to the latch circuit. In operation, each of the pressure transducers is successively scanned by the multiplexing means in response to address signals supplied on the address bus to the latch circuit.

Coe, C. F.↗

NASA NTRS2000

Solar Power System Options for the Radiation and Technology Demonstration Spacecraft

The Radiation and Technology Demonstration (RTD) Mission has the primary objective of demonstrating high-power (10 kilowatts) electric thruster technologies in Earth orbit. This paper discusses the conceptual design of the RTD spacecraft photovoltaic (PV) power system and mission performance analyses. These power system studies assessed multiple options for PV arrays, battery technologies and bus voltage levels. To quantify performance attributes of these power system options, a dedicated Fortran code was developed to predict power system performance and estimate system mass. The low-thrust mission trajectory was analyzed and important Earth orbital environments were modeled. Baseline power system design options are recommended on the basis of performance, mass and risk/complexity. Important findings from parametric studies are discussed and the resulting impacts to the spacecraft design and cost.

Kerslake, Thomas W.↗

NASA NTRS2022

Electrical Ground Support Equipment for the Sampling Caching System of the Mars 2020 Rover

In this work we describe in detail the architecture, design, testing and operation of the Electrical Ground Support Equipment (EGSE) “Blue Box” used to test and validate the Sampling Caching System (SCS) of the Mars 2020 Perseverance rover. The Blue Box architecture is centered around COTS motor controllers and COTS input-output modules communicating over an EtherCAT bus. A custom, low-level safety subsystem ensures no harm can be done to the flight articles. The modular architecture of the EGSE reduces cost and complexity while expediting assembly time. The Blue Box drives the 19 actuators of the SCS which span the main robotic arm, the corer system, the internal sample handling arm, the sample tube sealing system and the gas dust removal tool; mimicking the Rover Motor Control Assembly (RMCA). Due to the limited availability of RMCA’s, the EGSE enabled and performed the bulk of testing activities for SCS. The majority of the SCS actuators are composed of a 3-phase DC brushless motors, hall sensors for commutation, dual resolvers for output angular measurement, brakes, heaters and platinum thermistors. Additionally, the EGSE read 12 strain gauges forming part of a force torque sensor, and switches used for external positioning references. Over the 3-year span of the V&V campaign for the SCS, over 32 EGSE systems were built, tested and deployed to test venues at JPL and externally. The EGSE tested several families of the SCS subsystem, ranging from engineering units, life test units and two flight units. Test venues that this EGSE supported included lab benches, ultra-clean cleanrooms, ATLO facilities, and thermal vacuum chambers. Together with the test software systems, SSDEV and SSDEV-ECAT, the Blue Box EGSE enabled the team to efficiently test flight hardware and flight software together. We go over the safety features and fault management techniques employed to protect flight hardware. The effects of the long, 50-feet, EGSE harnesses on motor performance, EMI, electrical noise, and motor control performance are explained. Mitigations to these unwanted effects, including shielding strategy and inductance compensation, are summarized. We go over an excerpt of notable anomalies that this EGSE suffered through its operation, along with investigations and resolutions. Lessons learned, areas of improvement as part of future work, and recommendations for future implementations for similar EGSE’s, are shared.

Levine, Dan↗

NASA NTRSDate not supplied

Short Circuiting the Controller – Missteps in Maintenance and Inspection of Process and Wiring in STS-93

The primary objective of the Space Transportation System mission 93 (STS-93) was to deploy the Chandra X-Ray Observatory. Chandra, the world's most powerful X-Ray telescope, allowed scientists from around the world to study some of the most distant and dynamic objects in the universe. Stripped of nearly 7,000 pounds of its own gear to make room for the payload, the orbiter assigned to this mission was Space Shuttle Columbia. Prior to STS-93, Columbia had flown 25 flights and was NASA's oldest and heaviest orbiter. On July 23, 1999, after two prior launch scrubs, Eileen Collins and her crew of four launched from Kennedy Space Center. About five seconds after launch, Mission Control at Johnson Space Center detected a voltage drop on one of Columbia's electrical buses. As a result of this power fluctuation, a primary and back-up Main Engine controller dropped offline. Given design redundancy, the two remaining controllers supported all three engines. If there had been any other AC bus issues, one engine of the three on the Orbiter would have shut down. The redundant set of digital computer units in each engine controller saved Columbia and her crew from a very risky contingency abort. Post-flight inspection revealed soot on a screw head and a hole in an adjacent Kapton insulated wire. The single strand of polyimide wire was located nearly half-way down the payload bay. The Shuttle Independent Assessment Team (SIAT) reported that the wire had rubbed and chaffed against a burred screw head. The burr was later determined to be the result of overtightening of the screw by a technician during a maintenance refurbishment. Alone, the burr may not have been problematic, but later, during another ground processing event, possibly years after, someone inadvertently stepped on the wiring harness. With the pressure and motion of unintended contact, some of the Kapton insulation rubbed off against the burred screw head. The SIAT suspected the wire damage was pre-existing and was caused 4 or 5 years prior to the flight. Finally, the intense vibrations during the launch sequence allowed contact between the exposed conductor and exposed metal area on the burred screw head, resulting in the arcing and shorting of the wire. Due to the quick turnaround times of Space Shuttle Orbiters, wiring issues caused from multiple maintenance events were often overlooked. Additionally, failing to incorporate thorough and early inclusion of human systems integration (HSI) applications as a crucial part of the decision process can result in these types of misshaps. In order to reduce human error associated with integrated manufacturing, maintenance, refurbishment and flight preparations, wire inspection criteria should be refined and standardized, visual inspection processes should be quantified, and technicians should be certified by specially trained instructors. This case, among many others, unveils why human error management and development of safety metrics is a vital piece in the development of complex systems, and why it should be supported aggressively and implemented program wide.

Human Systems Integration↗

NASA NTRS2004

Options Studied for Managing Space Station Solar Array Electrical Hazards for Sequential Shunt Unit Replacement

The U.S. solar array strings on the International Space Station are connected to a sequential shunt unit (SSU). The job of the SSU is to shunt, or short, the excess current from the solar array, such that just enough current is provided downstream to maintain the 160-V bus voltage while meeting the power load demand and recharging the batteries. Should an SSU fail on-orbit, it would be removed and replaced with the on-orbit spare during an astronaut space walk or extravehicular activity (EVA) (see the photograph). However, removing an SSU during an orbit Sun period with input solar array power connectors fully energized could result in substantial hardware damage and/or safety risk to the EVA astronaut. The open-circuit voltage of cold solar-array strings can exceed 320 V, and warm solar-array strings could feed a short circuit with a total current level exceeding 240 A.

Delleur, Ann M.↗