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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS2015

High-Temperature, Wirebondless, Ultracompact Wide Bandgap Power Semiconductor Modules

Silicon carbide (SiC) and other wide bandgap semiconductors offer great promise of high power rating, high operating temperature, simple thermal management, and ultrahigh power density for both space and commercial power electronic systems. However, this great potential is seriously limited by the lack of reliable high-temperature device packaging technology. This Phase II project developed an ultracompact hybrid power module packaging technology based on the use of double lead frames and direct lead frame-to-chip transient liquid phase (TLP) bonding that allows device operation up to 450 degC. The new power module will have a very small form factor with 3-5X reduction in size and weight from the prior art, and it will be capable of operating from 450 degC to -125 degC. This technology will have a profound impact on power electronics and energy conversion technologies and help to conserve energy and the environment as well as reduce the nation's dependence on fossil fuels.

Elmes, John↗
NASA NTRSDate not supplied

The Next Challenge: High Power Electric Aircraft Propulsion

Introduction of the Aeronautic Research Mission Directorate Technical Challenge 10.2 “High Power Electric Aircraft Propulsion” to the power electronics community. The presentation was given as part of a panel discussion “Challenges of Power Electronics Systems for Sustainable Aviation” during the 2024 Sustainable Aviation Workshop held at The Ohio State University.

power electroics↗
NASA NTRSDate not supplied

MUSTANG: A Workhorse for NASA Spaceflight Avionics

The Modular Unified Space Technology Avionics for Next Generation (MUSTANG) is a small integrated Avionics system including Command and Data Handling (C&DH), Power System Electronics (PSE), Attitude Control System Interfaces (ACS), and Propulsion Electronics. The MUSTANG Avionics Architecture is built upon many years of knowledge capture and lessons learned at the Goddard Space Flight Center. With a motivation towards modularity and keeping board redesign costs to a minimum, MUSTANG offers flexibility in features with a backplane-less design and allows the user to choose the options (cards) needed for their system. It incorporates a distributed power system that provides secondary power to all its subcomponents reducing the number of primary services needed for an Avionics. MUSTANG can be integrated into one system or divided into several smaller components. MUSTANG supports redundancy and cross-strap ability for a more robust and reliable Avionics system. A variation of MUSTANG exists for Instrument Electronics called iMUSTANG and allows the user to select functionality applicable to the instrument electronics. MUSTANG is not meant to replace Avionics for all spacecraft. There are limitations due to its relatively compact size, but the MUSTANG design has proven broadly applicable on many spacecraft and instrument bus avionics architectures.

MUSTANG↗
NASA NTRS2004

The Space Technology 5 Avionics System

The Space Technology 5 (ST5) mission is a NASA New Millennium Program project that will validate new technologies for future space science missions and demonstrate the feasibility of building launching and operating multiple, miniature spacecraft that can collect research-quality in-situ science measurements. The three satellites in the ST5 constellation will be launched into a sun-synchronous Earth orbit in early 2006. ST5 fits into the 25-kilogram and 24-watt class of very small but fully capable spacecraft. The new technologies and design concepts for a compact power and command and data handling (C&DH) avionics system are presented. The 2-card ST5 avionics design incorporates new technology components while being tightly constrained in mass, power and volume. In order to hold down the mass and volume, and quali& new technologies for fUture use in space, high efficiency triple-junction solar cells and a lithium-ion battery were baselined into the power system design. The flight computer is co-located with the power system electronics in an integral spacecraft structural enclosure called the card cage assembly. The flight computer has a full set of uplink, downlink and solid-state recording capabilities, and it implements a new CMOS Ultra-Low Power Radiation Tolerant logic technology. There were a number of challenges imposed by the ST5 mission. Specifically, designing a micro-sat class spacecraft demanded that minimizing mass, volume and power dissipation would drive the overall design. The result is a very streamlined approach, while striving to maintain a high level of capability, The mission's radiation requirements, along with the low voltage DC power distribution, limited the selection of analog parts that can operate within these constraints. The challenge of qualifying new technology components for the space environment within a short development schedule was another hurdle. The mission requirements also demanded magnetic cleanliness in order to reduce the effect of stray (spacecraft-generated) magnetic fields on the science-grade magnetometer.

Speer, Dave↗
NASA NTRS2014

Controllable Multi-Functional Heat Exchanger for Space Applications

Team Spacering will design SpaceRing, a cooling and waste heat recycling system that addresses NASA’s interest for a cooling and power-generating unit in its prototype deep space habitat[1]. The merits of this system are as follows: - Controllable cooling is critical in maintaining a habitable environment and in the functionality of life support systems, powering electronics and other devices in a spacecraft. - Power is expensive to produce in a space environment. Converting wasted thermal energy into electricity will save energy and money for the space station. Space habitation is a highly complex endeavor. Consequently, the SpaceRing must adhere to additional constraints, which include: - Lack of natural fluid or gas convection due to micro-gravity environment - Limited range for heat source and cold sink temperature - Subject to high fluctuations of incoming thermal energy - Minimal weight due to the high cost of transporting items into space After consulting with NASA and experts in thermal studies, several design objectives were identified and are listed below in order of importance: - High cooling capacity - Reliable safety control and alert system - Maximization of energy generation from wasted heat - Optimization of cooling capability and power generation - Long operational life and minimization of maintenance In space, conventional cooling systems utilize passive heat pipe systems for small-scale heat rejection and pumped pipes with heat exchangers for large-scale heat rejection. Thermal power generation in space is mainly achieved by using thermoelectric material that produces electric energy from temperature differentials. We propose a design that combines cooling and thermoelectric power generation. Using a hot pipe connected to a heat source, a cold pipe connected to a radiator, a heat exchanger in contact with both pipes, and a pump in each pipe to move the fluid, we make an active cooling system. The thermoelectric material is inserted in the heat exchanger between the hot and cold pipes to extract energy from the maximum temperature difference coming from the two pipes. It also provides controllable cooling capacity through the use of the thermoelectric material as cooling modules. We have tested our system under target heat source and sink conditions and measured the temperatures and flow rates. Our preliminary results are promising, demonstrating the output of 25 Watts of power on average. We are in the process of running control tests to compare the heat transfer performance without the thermoelectric generators, which is vital in determining whether or not the generators introduce an unacceptable amount of thermal resistance. While some control tests have been run already, the results have been noisy. We have modified our testing environment and data acquisition methods to allow for simple comparison between the control test results and thermoelectric performance results. Lastly, we have modified the heat exchanger design slightly to include acrylic baffles. This will allow the heat to flow through the heat exchanger more efficiently. We have presented our preliminary results to NASA and will deliver the final product by the end of May 2014.

Lee Xiong↗
NASA NTRS1972

Delta-92 Telesat-A operations summary

Telesat-A, which is the first of the Canadian Telesat satellites to be launched for the domestic satellite communications system is described. The launch vehicle, designated Delta-92 consists of a DSV 3p-11 extended long tank first stage with an MB-3 engine, augmented by low-drag Castor 2 solid motors. The spacecraft has a spin-stabilized electronic system powered by 23,000 solar cells, with sufficient on-board battery capability to provide full capacity power during eclipse of the solar cells. A 60-inch wide circular directional antenna which remains constantly aimed at Canada is included.

Source record↗
NASA NTRS1989

NASA standard GAS Can satellite

The Get-Away Special canister (GAS Can) satellite is a small, (150 lb) low-cost satellite making it possible for commercial and scientific institutions to conduct experiments in space on an economical and short-term basis. The current model is called Xsat (Exceptional Satellite) and is designed to be launched from a GAS canister on the STS Orbiter; also provided is a low-cost automated PC-operated ground station for commercial, scientific, and government users. The Xsat structure is diagrammed, and details such as payload interface, weight restrictions, and structural loads are described in detail, pointing out that Xsat has a maximum payload weight of 50 lbs, and has a natural vibration frequency of around 45 Hz, with a minimum requiremet of 35 Hz. Thermal designs, power system, electronics, computer design and bus system, and satellite operations are all outlined.

Cudmore, Patrick H.↗
NASA NTRS1991

Long Duration Exposure Facility (LDEF) low-temperature Heat Pipe Experiment Package (HEPP) power system results

An overview of a self contained Direct Energy Transfer (DET) Power System which was developed to provide power to the Long Duration Exposure Facility (LDEF) Heat Pipe Experiment Package (HEPP) Experiment. A brief synopsis of the flight characteristics of the power system are presented to illustrate the system's capability to function successfully in a low Earth orbit (LEO) space environment. The successful space flight operations of the battery, solar arrays, and power system electronics during the flight period of almost six years (approximately 32,000 LEO cycles) are discussed.

Tiller, Smith E.↗
NASA NTRS2001

Performance of High-Speed PWM Control Chips at Cryogenic Temperatures

The operation of power electronic systems at cryogenic temperatures is anticipated in many NASA space missions such as planetary exploration and deep space probes. In addition to surviving the space hostile environment, electronics capable of low temperature operation would contribute to improving circuit performance, increasing system efficiency, and reducing development and launch costs. As part of the NASA Glenn Low Temperature Electronics Program, several commercial high-speed Pulse Width Modulation (PWM) chips have been characterized in terms of their performance as a function of temperature in the range of 25 to -196 C (liquid nitrogen). These chips ranged in their electrical characteristics, modes of control, packaging options, and applications. The experimental procedures along with the experimental data obtained on the investigated chips are presented and discussed.

Elbuluk, Malik E.↗
NASA NTRS2001

Flight Computer Design for the Space Technology 5 (ST-5) Mission

As part of NASA's New Millennium Program, the Space Technology 5 mission will validate a variety of technologies for nano-satellite and constellation mission applications. Included are: a miniaturized and low power X-band transponder, a constellation communication and navigation transceiver, a cold gas micro-thruster, two different variable emittance (thermal) controllers, flex cables for solar array power collection, autonomous groundbased constellation management tools, and a new CMOS ultra low-power, radiation-tolerant, +0.5 volt logic technology. The ST-5 focus is on small and low-power. A single-processor, multi-function flight computer will implement direct digital and analog interfaces to all of the other spacecraft subsystems and components. There will not be a distributed data system that uses a standardized serial bus such as MIL-STD-1553 or MIL-STD-1773. The flight software running on the single processor will be responsible for all real-time processing associated with: guidance, navigation and control, command and data handling (C&DH) including uplink/downlink, power switching and battery charge management, science data analysis and storage, intra-constellation communications, and housekeeping data collection and logging. As a nanosatellite trail-blazer for future constellations of up to 100 separate space vehicles, ST-5 will demonstrate a compact (single board), low power (5.5 watts) solution to the data acquisition, control, communications, processing and storage requirements that have traditionally required an entire network of separate circuit boards and/or avionics boxes. In addition to the New Millennium technologies, other major spacecraft subsystems include the power system electronics, a lithium-ion battery, triple-junction solar cell arrays, a science-grade magnetometer, a miniature spinning sun sensor, and a propulsion system.

Speer, David↗
NASA NTRS2004

Space Technology 5: Pathfinder for Future Micro-Sat Constellations

The Space Technology 5 (ST-5) Project, currently in the implementation phase, is part of the National Aeronautics and Space Administration (NASA) s New Millennium Program (NMP). ST-5 will consist of a constellation of three miniature satellites, each with mass less than 25 kg and size approximately 60 cm by 30 cm. ST-5 addresses technology challenges, as well as fabrication, assembly, test and operations strategies for future micro-satellite missions. ST-5 will be deployed into a highly eccentric, geo-transfer orbit (GTO). This will expose the spacecraft to a high radiation environment as well as provide a low level magnetic background. A three-month flight demonstration phase is planned to validate the technologies and demonstrate concepts for future missions. Each ST-5 spacecraft incorporates NMP competitively-selected breakthrough technologies. These include Cold Gas Micro-Thrusters for propulsion and attitude control, miniature X-band transponder for space-ground communications, Variable Emittance Coatings for dynamic thermal control, and CULPRiT ultra low power logic chip used for Reed-Solomon encoding. The ST-5 spacecraft itself is a technology that can be infused into future missions. It is a fully functional micro-spacecraft built within tight volume and mass constraints. It is built to withstand a high radiation environment, large thermal variations, and high launch loads. The spacecraft power system is low-power and low-voltage, and is designed to turn on after separation &om the launch vehicle. Some of the innovations that are included in the ST-5 design are a custom spacecraft deployment structure, magnetometer deployment boom, nutation damper, X-band antenna, miniature spinning sun sensor, solar array with triple junction solar cells, integral card cage assembly containing single card Command and Data Handling and Power System Electronics, miniature magnetometer, and lithium ion battery. ST-5 will demonstrate the ability of a micro satellite to perform research-quality science. Each ST-5 spacecraft will deploy a precision magnetometer to be used both for attitude determination and as a representative science instrument. The spacecraft has been developed with a low magnetic signature to avoid interference with the magnetometer. The spacecraft will be able to detect and respond autonomously to science events, i.e. significant changes in the magnetic field measurements. The three spacecraft will be a pathfinder for future constellation missions. They will be deployed to demonstrate an appropriate geometry for scientific measurements as a constellation. They will be operationally managed as a constellation, demonstrating automation and communication strategies that will be useful for future missions. The technologies and future mission concepts will be validated both on the ground and in space. Technologies will be validated on the ground by a combination of component level and system level testing of the flight hardware in a thermal vacuum environment. In flight, specific validation runs are planned for each of the technologies. Each validation run consists of one or more orbits with a specific validation objective. This paper will describe the ST-5 mission, and the applicability of the NMP technologies, spacecraft, and mission concepts to future missions. It will also discuss the validation approach for the ST-5 technologies and mission concepts.

Carlisle, Candace↗
NASA NTRS2004

Robotic Exploration of the Surface and Atmosphere of Venus

Venus, the "greenhouse planet", is a scientifically fascinating place. In many ways it can be considered "Earth's evil twin." A huge number of important scientific questions remain to be answered: 1) Before the runaway greenhouse effect, was early Venus temperate? 2) Did Venus once have an ocean? 3) What causes the geological resurfacing of the planet? 4) Is Venus still geologically active? 5) What is the "snow" on Venus mountaintops? 6) Can we learn about Earth's climate from Venus? 7) Is the atmosphere of Venus suitable for life? To address these and other scientific questions, a robotic mission to study the surface and atmosphere of Venus has been designed. The mission includes both surface robots, designed with an operational lifetime of 90 days on the surface of Venus, and also solar-powered airplanes to probe the middle atmosphere. At 450 Celsius, and with 90 atmospheres of pressure of carbon-dioxide atmosphere, the surface of Venus is a hostile place for operation of a probe. This paper will present the mission design, discuss the technology options for materials, power systems, electronics, and instruments, and present a short summary of the mission.

Landis, Geoffrey A.↗
NASA NTRS2016

Global Precipitation Measurement (GPM) Spacecraft Lithium Ion Battery Micro-Cycling Investigation

The Global Precipitation Measurement (GPM) spacecraft was jointly developed by NASA and JAXA. It is a Low Earth Orbit (LEO) spacecraft launched on February 27, 2014. The power system is a Direct Energy Transfer (DET) system designed to support 1950 watts orbit average power. The batteries use SONY 18650HC cells and consist of three 8s by 84p batteries operated in parallel as a single battery. During instrument integration with the spacecraft, large current transients were observed in the battery. Investigation into the matter traced the cause to the Dual-Frequency Precipitation Radar (DPR) phased array radar which generates cyclical high rate current transients on the spacecraft power bus. The power system electronics interaction with these transients resulted in the current transients in the battery. An accelerated test program was developed to bound the effect, and to assess the impact to the mission.

Lithium Ion Battery Performance↗
NASA NTRS1994

Alarm- And Power-Monitoring System

Electronic central monitoring system, called Remote Monitor Alarm System, RMAS, used to monitor malfunction alarms and power supplies of remotely located equipment modules of transmitting and receiving equipment in fiber-optic communication network at Kennedy Space Center. Includes central monitoring unit at location convenient for technicians, plus remote terminal unit at each remote site containing equipment to be monitored.

Stute, Rob↗
NASA NTRS1980

Electronic system for high power load control

Parallel current paths are divided into two groups, with control devices in the current paths of one group each having a current limiting resistor, and the control devices in the other group each having no limiting resistor, so that when the control devices of the second group are turned fully on, a short circuit is achieved by the arrangement of parallel current paths. Separate but coordinated control signals are provided to turn on the control devices of the first group and increase their conduction toward saturation as a function of control input, and when fully on, or shortly before, to turn on the control devices of the second group and increase their conduction toward saturation as a function of the control input as that input continues to increase. Electronic means may be used to generate signals. The system may be used for 1-V characteristic measurements of solar arrays as well as for other load control purposes.

Miller, E. L.↗
NASA NTRS2020

Advanced Modular Power System Fully Populated Electronics Enclosure Thermal Testing

A thermal test of an enclosure housing the AMPS electronics cards was performed. The enclosure was fully populated with 17 nonfunctioning AMPS electronics cards. Heaters were applied to the cards to simulate their operation. A cold plate was installed on the top of the case. The case was then insulated to simulate its installation within an electronics rack with the face and rear of the enclosure left open. The cold plate was connected to a chiller which provided cooling water at a fixed input temperature to the cold plate. Flow rates from the chiller of 0.0 L/min, 1.4 L/min and 1.8 L/min were tested. The test was operated at each flow rate over a range of heater power levels until the steady state temperature of the electronics cards was reached. Heater power levels from 2.5 W/card to 45 W/card were tested. The data collected was used to determine the steady state operating temperature of the cards, the temperature distribution between the cards, the time to reach steady-state temperature, the change in steady state temperature with change in heater power, the temperature distribution between the cards at the steady state condition and the heat lost to the surrounds by the enclosure. One of the main goals of the testing was to determine how much waste heat the cards could generate and still maintain a touch temperature at or below 40°C. Based on this requirement, for the 3 cooling rates tested, the maximum card waste heat that could be generated by each individual card was 3.9 W, 16.3 W and 17.6 W respectively. An analysis was also set up to provide an analytical model to estimate the card temperature under different cold plate flow rates and card heater power levels. The results off the analysis was compared to that of the experiment of the range of flow rates and power levels tested.

Anthony Colozza↗
NASA NTRS1993

Electrical Actuation Technology Bridging

This document contains the proceedings of the NASA Electrical Actuation Technology Bridging (ELA-TB) Workshop held in Huntsville, Alabama, September 29-October 1, 1992. The workshop was sponsored by the NASA Office of Space Systems Development and Marshall Space Flight Center (MSFC). The workshop addressed key technologies bridging the entire field of electrical actuation including systems methodology, control electronics, power source systems, reliability, maintainability, and vehicle health management with special emphasis on thrust vector control (TVC) applications on NASA launch vehicles. Speakers were drawn primarily from industry with participation from universities and government. In addition, prototype hardware demonstrations were held at the MSFC Propulsion Laboratory each afternoon. Splinter sessions held on the final day afforded the opportunity to discuss key issues and to provide overall recommendations. Presentations are included in this document.

Hammond, Monica↗
NASA NTRS1980

Large space systems technology electronics: Data and power distribution

The development of hardware technology and manufacturing techniques required to meet space platform and antenna system needs in the 1980s is discussed. Preliminary designs for manned and automatically assembled space power system cables, connectors, and grounding and bonding materials and techniques are reviewed. Connector concepts, grounding design requirements, and bonding requirements are discussed. The problem of particulate debris contamination for large structure spacecraft is addressed.

Dunbar, W. G.↗