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All sourcesNASA NTRSDOE OSTI
NASA NTRS2011

Circuit for Communication Over Power Lines

Many distributed systems share common sensors and instruments along with a common power line supplying current to the system. A communication technique and circuit has been developed that allows for the simple inclusion of an instrument, sensor, or actuator node within any system containing a common power bus. Wherever power is available, a node can be added, which can then draw power for itself, its associated sensors, and actuators from the power bus all while communicating with other nodes on the power bus. The technique modulates a DC power bus through capacitive coupling using on-off keying (OOK), and receives and demodulates the signal from the DC power bus through the same capacitive coupling. The circuit acts as serial modem for the physical power line communication. The circuit and technique can be made of commercially available components or included in an application specific integrated circuit (ASIC) design, which allows for the circuit to be included in current designs with additional circuitry or embedded into new designs. This device and technique moves computational, sensing, and actuation abilities closer to the source, and allows for the networking of multiple similar nodes to each other and to a central processor. This technique also allows for reconfigurable systems by adding or removing nodes at any time. It can do so using nothing more than the in situ power wiring of the system.

Krasowski, Michael J.↗
NASA NTRS2012

Optimal Dynamic Sub-Threshold Technique for Extreme Low Power Consumption for VLSI

For miniaturization of electronics systems, power consumption plays a key role in the realm of constraints. Considering the very large scale integration (VLSI) design aspect, as transistor feature size is decreased to 50 nm and below, there is sizable increase in the number of transistors as more functional building blocks are embedded in the same chip. However, the consequent increase in power consumption (dynamic and leakage) will serve as a key constraint to inhibit the advantages of transistor feature size reduction. Power consumption can be reduced by minimizing the voltage supply (for dynamic power consumption) and/or increasing threshold voltage (V(sub th), for reducing leakage power). When the feature size of the transistor is reduced, supply voltage (V(sub dd)) and threshold voltage (V(sub th)) are also reduced accordingly; then, the leakage current becomes a bigger factor of the total power consumption. To maintain low power consumption, operation of electronics at sub-threshold levels can be a potentially strong contender; however, there are two obstacles to be faced: more leakage current per transistor will cause more leakage power consumption, and slow response time when the transistor is operated in weak inversion region. To enable low power consumption and yet obtain high performance, the CMOS (complementary metal oxide semiconductor) transistor as a basic element is viewed and controlled as a four-terminal device: source, drain, gate, and body, as differentiated from the traditional approach with three terminals: i.e., source and body, drain, and gate. This technique features multiple voltage sources to supply the dynamic control, and uses dynamic control to enable low-threshold voltage when the channel (N or P) is active, for speed response enhancement and high threshold voltage, and when the transistor channel (N or P) is inactive, to reduce the leakage current for low-leakage power consumption.

Duong, Tuan A.↗
NASA NTRS2012

Multi-Mission Power Analysis Tool (MMPAT) Version 3

The Multi-Mission Power Analysis Tool (MMPAT) simulates a spacecraft power subsystem including the power source (solar array and/or radioisotope thermoelectric generator), bus-voltage control, secondary battery (lithium-ion or nickel-hydrogen), thermostatic heaters, and power-consuming equipment. It handles multiple mission types including heliocentric orbiters, planetary orbiters, and surface operations. Being parametrically driven along with its user-programmable features can reduce or even eliminate any need for software modifications when configuring it for a particular spacecraft. It provides multiple levels of fidelity, thereby fulfilling the vast majority of a project s power simulation needs throughout the lifecycle. It can operate in a stand-alone mode with a graphical user interface, in batch mode, or as a library linked with other tools. This software can simulate all major aspects of a spacecraft power subsystem. It is parametrically driven to reduce or eliminate the need for a programmer. Added flexibility is provided through user-designed state models and table-driven parameters. MMPAT is designed to be used by a variety of users, such as power subsystem engineers for sizing power subsystem components; mission planners for adjusting mission scenarios using power profiles generated by the model; system engineers for performing system- level trade studies using the results of the model during the early design phases of a spacecraft; and operations personnel for high-fidelity modeling of the essential power aspect of the planning picture.

Wood, Eric G.↗
NASA NTRS2009

Power Sources for Micro-Autonomous Vehicles- Challenges and Prospects

Micro-autonomous vehicle systems are expected to have expanded role in military missions by providing full spectrum intelligence, surveillance and reconnaissance support on the battlefield, suppression of enemy defenses, and enabling co-operative (swarm-like) configurations. Of the numerous demanding requirements of autonomy, sensing, navigation, mobility, etc., meeting the requirement of mission duration or endurance is a very challenging one. This requirement is demanding because of the constraints of mass and volume that limit the quantity of energy that can be stored on-board. Energy is required for mobility, payload operation, information processing, and communication. Mobility requirements typically place an extraordinary demand on the specific energy (Wh/kg) and specific power (W/kg) of the power source; the actual distribution of the energy between mobility and other system functions could vary substantially with the mission type. The power requirements for continuous mobility can vary from 100-1000 W/kg depending on the terrain, ground speed and flight speed. Even with the power source accounting for 30% of the mass of the vehicle, the best of rechargeable batteries can provide only up to 1-2 hours of run-time for a continuous power demand at 100W/kg. In the case of micro-aerial vehicles with flight speed requirements in the range of 5-15 m s-1, the mission times rarely exceed 20 minutes [2]. Further, the power required during take-off and hover can be twice or thrice that needed for steady level flight, and thus the number and sequence of such events is also limited by the mass and size of the power source. For operations such as "perch and stare" or "silent watch" the power demand is often only a tenth of that required during continuous flight. Thus, variation in power demand during various phases of the mission importantly affects the power source selection.

fuel cells↗
NASA NTRS2015

Lightweight Damage Tolerant, High-Temperature Radiators for Nuclear Power and Propulsion

NASA is increasingly emphasizing exploration to bodies beyond near-Earth orbit. New propulsion systems and new spacecraft are being built for these missions. As the target bodies get further out from Earth, high energy density systems, e.g., nuclear fusion, for propulsion and power will be advantageous. The mass and size of these systems, including supporting systems such as the heat exchange system, including thermal radiators, will need to be as small as possible. Conventional heat exchange systems are a significant portion of the total thermal management mass and size. Nuclear electric propulsion (NEP) is a promising option for high-speed, in-space travel due to the high energy density of nuclear fission power sources and efficient electric thrusters. Heat from the reactor is converted to power for use in propulsion or for system power. The heat not used in the power conversion is then radiated to space as shown in figure 1. Advanced power conversion technologies will require high operating temperatures and would benefit from lightweight radiator materials. Radiator performance dictates power output for nuclear electric propulsion systems. Pitch-based carbon fiber materials have the potential to offer significant improvements in operating temperature, thermal conductivity, and mass. These properties combine to allow significant decreases in the total mass of the radiators and significant increases in the operating temperature of the fins. A Center-funded project at NASA Marshall Space Flight Center has shown that high thermal conductivity, woven carbon fiber fins with no matrix material, can be used to dissipate waste heat from NEP systems and because of high specific power (kW/kg), will require less mass and possibly less total area than standard metal and composite radiator fins for radiating the same amount of heat. This project uses an innovative approach to reduce the mass and size required for the thermal radiators to the point that in-space NEP and power is enabled. High thermal conductivity carbon fibers are lightweight, damage tolerant, and can be heated to high temperature. Areal densities in the NASA set target range of 2 to 4 kg/m2 (for enabling NEP) are achieved and with specific powers (kW/kg) a factor of about 7 greater than conventional metal fins and about 1.5 greater than carbon composite fins. Figure 2 shows one fin under test. All tests were done under vacuum conditions.

Craven, Paul D.↗
NASA NTRS2017

The Lightweight Integrated Solar Array and anTenna (LISA-T) Big Power for Small Spacecraft

NASA is developing a space power system using lightweight, flexible photovoltaic devices originally developed for use here on Earth to provide low cost power for spacecraft. The Lightweight Integrated Solar Array and anTenna (LISA-T) is a launch stowed, orbit deployed array on which thin-film photovoltaic and antenna elements are embedded. The LISA-T system is deployable, building upon NASA's expertise in developing thin-film deployable solar sails such the one being developed for the Near Earth Asteroid Scout project which will fly in 2018. One of the biggest challenges for the NEA Scout, and most other spacecraft, is power. There simply isn't enough of it available, thus limiting the range of operation of the spacecraft from the Sun (due to the small surface area available for using solar cells), the range of operation from the Earth (low available power with inherently small antenna sizes tightly constrain the bandwidth for communication), and the science (you can only power so many instruments with limited power). The LISA-T has the potential to mitigate each of these limitations, especially for small spacecraft. Inherently, small satellites are limited in surface area, volume, and mass allocation; driving competition between their need for power and robust communications with the requirements of the science or engineering payload they are developed to fly. LISA-T is addressing this issue, deploying large-area arrays from a reduced volume and mass envelope - greatly enhancing power generation and communications capabilities of small spacecraft and CubeSats. The problem is that these CubeSats can usually only generate between 7W and 50W of power. The power that can be generated by the LISA-T ranges from tens of watts to several hundred watts, at a much higher mass and stowage efficiency. A matrix of options are in development, including planar (pointed) and omnidirectional (non-pointed) arrays. The former is seeking the highest performance possible while the latter is seeking GN&C simplicity. Options for leveraging both high performance, 'typical cost' triple junction thin-film solar cells as well as moderate performance, low cost cells are being developed. Alongside, UHF (ultrahigh frequency), S-band, and X-band antennas are being integrated into the array to move their space claim away from the spacecraft and open the door for more capable multi-element antenna designs such as those needed for spherical coverage and electronically steered phase arrays.

Space Exploration↗
NASA NTRS2017

Low and High-Power Inductive Pulsed Plasma Thruster Development Testing at NASA-MSFC

The inductive pulsed plasma thruster (IPPT) is an electromagnetic plasma accelerator that has been identified in NASA roadmaps as an enabling propulsion technology for some niche low-power missions and for high-power in-space propulsion needs. The IPPT is an electrodeless space propulsion device where a capacitor is charged to an initial voltage and then discharged producing a high current pulse through a coil. The field produced by this pulse ionizes propellant, inductively driving current in a plasma located near the face of the coil. Once the plasma is formed it can be accelerated and expelled at a high exhaust velocity by the electromagnetic Lorentz body force arising from the interaction of the induced plasma current and the magnetic field produced by the current in the coil. Thrusters of this type possess many demonstrated and potential benefits that make them worthy of continued investigation. The electrodeless nature of these thrusters eliminates the lifetime and contamination issues associated with electrode erosion in conventional electric thrusters. Also, a wider variety of propellants are accessible when compatibility with metallic electrodes in no longer an issue. IPPTs have been successfully operated using propellants like ammonia, hydrazine, and CO2, and there is no fundamental reason why they would not operate on other in situ propellants like H2O. It is well-known that pulsed accelerators can maintain constant specific impulse (I(sub sp)) and thrust efficiency (eta(sub t)) over a wide range of input power levels by adjusting the pulse rate to hold the discharge energy per pulse constant. It has also been demonstrated that an inductive pulsed plasma thruster can operate in a regime where eta(sub t) is relatively constant over a wide range of I(sub sp) values (3000-8000 s). Finally, thrusters in this class have operated in single-pulse mode at high energy per pulse, and by increasing the pulse rate they offer the potential to process very high levels of power using a single thruster. There has been significant previous research on IPPTs designed around a planar-coil (flat-plate) geometry. The most notable of these was the Pulsed Inductive Thruster (PIT), with the PIT MkV presently representing the state-of- the-art in pulsed high-power IPPT technological development. In this paper, we focus on two planar-geometry devices that operate at significantly different power levels. Most work performed at NASA-Marshall Space Flight Center (MSFC) has, to date, focused on lower power thruster operation (approx. = 10s to 100s of J/pulse, up to 2-2.5 kW average power throughput) and previously described. The most recent work aimed to assemble a device that could be tested in cyclic mode on a thrust-stand, and which could augment the existing data set for IPPTs. In addition, the thruster was designed to serve as a test-bed for solid state switching circuitry and pulsed gas valves, with the modular design of the device allowing for variation in or upgrades to test configuration. Recently, MSFC obtained on loan from the Georgia Institute of Technology (Atlanta, GA) the PIT MkVI, successor to the PIT MkV. The MkV and MkVI are similar in design with much of the hardware from the former, specifically the capacitors and spark-gap switches, being reused in the latter. The coil is similar in geometry but has bent copper rods used in the latest iteration in place of the Litz wire windings found in the MkV. The MkVI master switch for the spark gaps is located in the vacuum chamber contained within a sealed, pressurized vessel fastened to the back of the thruster. This is different from the MkV where many capacitor charging lines and spark gap-triggering delay lines ran to the thruster from a master trigger located outside the vacuum chamber. The MkVI was damaged during testing soon after its fabrication was completed. The thruster arrived at MSFC still-damaged and mostly disassembled into many individual pieces. The device has been repaired, with a few additional design changes implemented after discussions with the late Prof. Lovberg regarding the initial testing results and issues encountered. In the present work, we present results from testing of both the small IPPT and the larger MkVI thruster. The smaller device (Fig. 1) is tested on a thrust stand on multiple gases to demonstrate its capability to operate in a repetition-rate mode and serve as a IPPT technology-development testbed. The larger MkVI (Fig. 2) is operated for the first time in its newly reconstituted state, demonstrating full-power pulsed operation and, for the first time, repetition-rate operation of a high-power IPPT. The additional upgrades required for synchronous operation of all the pulsed systems in single-pulse and repetition-rate mode are described in detail.

Polzin, Kurt A.↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100hrs to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne (AR) is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2018

Development of High Power Hall Thruster Systems to Enable the NASA Exploration Vision

The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) systems for large-scale science missions and cargo transportation. Development is underway at Aerojet Rocketdyne on Hall thruster systems that are intended to bracket the needs of future NASA SEP missions in support of space exploration. The Advanced Electric Propulsion System (AEPS) program is developing and qualifying a 13.3kW Hall thruster system to be demonstrated on the Power and Propulsion Element (PPE), which is intended to be the first element of a Lunar Outpost Platform - Gateway (LOP-G). The NextSTEP program is integrating a nested Hall thruster into a 100 kW system and testing it for 100 hours. These two programs will provide a path to efficient in-space propulsion that will allow NASA to transfer the large amounts of cargo that is needed to support human missions - first to the moon and then on to Mars. The Advanced Electric Propulsion System (AEPS) program is completing development, qualification and delivery of five flight 13.3kW EP systems to NASA. The flight AEPS system includes a magnetically shielded long-life Hall thruster, Power Processing Unit (PPU) and a Xenon Flow Controller (XFC). The Hall thruster, developed and demonstrated by NASA, operates at input powers up to 12.5 kW while providing a specific impulse over an estimated 2800s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95-140V, consistent with operation beyond the orbit of Mars. The integrated system input power is continuously throttleable between 3 and 13.3kW. Component level testing of the EP String has begun with prototype hardware. The NextSTEP program is developing a 100kW Electric Propulsion (EP) system using a nested Hall thruster designed for powers up to 250kW, a modular power processor and a modular mass flow controller. While the program objective is to operate the integrated EP system continuously at 100kW for 100 hours to demonstrate thermal stability and support the development of system life time models, it builds on decades of experience with long-life Hall thrusters and the design is evolvable to a capability of 250kW. Design upgrades that demonstrate the 100kW EP system have been completed and tested. Aerojet Rocketdyne is excited to support NASA as it extends human reach into deep space and believes that these programs will provide the propulsion to make such missions affordable and sustainable. These systems provide NASA with a range of options to power its deep space transport vehicles. This paper presents the mission requirements for supporting the NASA exploration vision, as well as the status for the high power Hall thruster systems in development.

Jackson, Jerry↗
NASA NTRS2018

A Comparison of Energy Conversion Technologies for Space Nuclear Power Systems

A key element of space nuclear power systems is the energy conversion subsystem that converts the nuclear heat into electrical power. Nuclear systems provide a favorable option for missions that require long-duration power in hostile space environments where sunlight for solar power is absent or limited. There are two primary nuclear power technology options. Radioisotope Power System (RPS) utilize the natural decay heat from Pu238 to generate electric power levels up to about one kilowatt. Fission Power System (FPS) rely on a sustained fission reaction of U235 and offer the potential to supply electric power from kilowatts to megawatts. Example missions for nuclear power include Mars science rovers (e.g. Curiosity, Mars 2020), lunar and Mars surface landers ? including crewed missions, deep space planetary orbiters, Ocean World science landers, and robotic space probes that utilize nuclear electric propulsion. This paper examines the energy conversion technology options that can be used with RPS and FPS, and provides an assessment of their relative performance and technology readiness.

Mason, Lee S.↗
NASA NTRS2019

A Study on the Use of Solid-Oxide Fuel Cells for Increased Power Generation on Small Aircraft

A growing demand for higher-power electrical equipment on aircraft is driving the exploration of alternative electrical power generation devices. “Hybrid-electric” aircraft architectures that leverage multiple or alternative power sources are also becoming more popular with rising concerns over efficiency. Fuel cells are particularly interesting as auxiliary or secondary electrical power generation devices, due to their relative efficiency and longevity, and could be used as a “power pod” for vehicles without enough on-board electrical generation capability. This paper examines the feasibility and potential value of using an externally-mounted, hybridelectric, solid-oxide fuel cell (SOFC) power system on a Class IV unmanned aerial vehicle (UAV) for the purpose of power generation in excess of the baseline electrical load. Moreover, the paper will demonstrate that a SOFC auxiliary power unit (APU) provides significant value to an aircraft by producing a relatively large amount of additional electrical power for a relatively low initial investment when compared to producing additional power by scaling the engine.

Litherland, Brandon L.↗
NASA NTRS2019

Comparison of Energy Conversion Technologies for Space Nuclear Power Systems

A key element of space nuclear power systems is the energy conversion subsystem that converts the nuclear heat into electrical power. Nuclear systems provide a favorable option for missions that require long-duration power in hostile space environments where sunlight for solar power is absent or limited. There are two primary nuclear power technology options: (1) radioisotope power systems (RPSs) utilize the natural decay heat from 238Pu to generate electric power levels up to about 1 kW and (2) fission power systems (FPSs) rely on a sustained fission reaction of 235U and offer the potential to supply electric power from kilowatts to megawatts. Example missions utilizing nuclear power include Mars science rovers (e.g., Curiosity, Mars 2020), lunar and Mars surface landers, crewed surface outposts, deep space planetary orbiters, Ocean World science landers, and robotic space probes that utilize nuclear electric propulsion. This report examines the energy conversion technology options that can be used with RPSs and FPSs, and provides an assessment of their relative performance.

Lee S Mason↗
NASA NTRSDate not supplied

Modeling and Analysis of Stirling Power Convertors

Modeling and Analysis of Stirling Power Convertors Luis A. Rodriguez1 Steven M. Geng, Terry V. Reid, Scott D. Wilson NASA Glenn Research Center, Cleveland, OH, 44135, USA NASA Glenn’s Thermal Energy Conversion Branch is supporting the development of the next generation free-piston Stirling power convertors. American Superconductor (AMSC) and Sunpower Inc. are the two firms under contract to develop the Flexure Isotope Stirling Convertor (FISC) and the Sunpower Robust Stirling Convertor (SRSC), respectively. To comprehend and forecast convertor performance, Sage, ANSYS® Maxwell, and ANSYS® Fluent were used to model the Stirling thermodynamic cycle, alternator electromagnetics, and piston and displacer dynamics. I. Introduction Stirling convertors are being developed by NASA as a potential steady source of electrical power for NASA’s future scientific space missions. Currently, NASA Glenn Research Center has two corporations under contract, American Superconductor (AMSC) and Sunpower Inc., for the development of the next generation of free-piston Stirling convertors for dynamic radioisotope power systems. AMSC is developing the Flexure Isotope Stirling Convertor (FISC), which uses flexures to prevent side motion and rubbing of the piston. Similarly, Sunpower Inc, is developing the Sunpower Robust Stirling Convertor (SRSC). The SRSC uses gas bearings to prevent radial contact of the moving piston. As convertor development continues, it is increasingly important to understand and predict the interactions of components in the system, how they respond to one another, and how they perform as a response to changes in operating conditions. A suitable and enlightening way to demonstrate and foresee these interactions is with the use of accurate modeling software. Sage, ANSYS® Maxwell, and ANSYS® Fluent are the current modeling tools used by NASA to analytically determine convertor performance. Sage is a one-dimensional object-oriented commercial software package used for modeling and optimizing Stirling convertors for Dynamic Radioisotope Power Systems (DRPS) and it is one of the most accurate Stirling convertor codes in use by NASA. This code is the successor to GLIMPS (Globally-Implicit Stirling Cycle Simulation) and GLOP (GLIMPS Optimization) software created by Gedeon Associates [1]. Model input parameters are typically material/gas type, component physical dimensions, temperatures, frequency, charge pressure, and number of time/space nodes. Sage is used to model both the FISC’s and SRSC’s Stirling cycle thermodynamics and piston/displacer dynamics. Performance maps were created and analyzed for both power systems to better understand the relationship between the following conditions: cold-end temperature, hot-end temperature, piston/displacer amplitudes, pressure drop, and thermal input power. The synergy between these conditions will help determine parameter sensitivity. ANSYS® Maxwell was used to create a three-dimensional (3-D) axisymmetric model for both FISC and SRSC alternators. The significant physical components included in each model are the magnets, magnet carrier, outer/inner laminations, and the coil. Inputs to the model are piston amplitude, piston frequency, alternator load, coil resistance, tuning capacitance, and specific material properties. The alternator models calculate terminal voltage, current, piston/current phase, voltage/current phase, coil inductance, terminal power and efficiency. The RI2 losses, core (hysteresis and eddy) losses, and magnet/can eddy losses are also a part of the final results. ANSYS® Fluent is used to build 3-D computational fluid dynamic (CFD) models to examine the Stirling cycle thermodynamics for both the FISC and SRSC systems. Three-dimensional Computer Aided Design (CAD) models were used to create the physical components of each convertor. Steady-state simulations were conducted for hardware testing, prediction of environmental losses during testing, and generation of radiation look-up tables. The model inputs to the aforementioned analysis are the material properties and boundary thermal conditions. The steady-state model calculates temperature and heat flow distributions. Transient 3-D calculations were also part of the CFD analysis. In this study a physically reduced version of the FISC is used to obtain a prediction of available engine power. For the gas bearing SRSC, the transient effort is used to obtain a prediction of bearing pad performance and its sensitivity to micro-channel geometric variation. The model inputs to the transient simulations are the piston amplitude, displacer amplitude, frequency, displacer/piston phase angle, dynamic deforming CFD grid, temperature boundary conditions, and user defined files describing motion profile of piston/displacer. The results of the model are temperature distributions, heat distributions, and PV power produced at pre-determined conditions.

Luis A Rodriguez↗
NASA NTRS2021

Thermal Systems Modeling of Chemical Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such as system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90-100 Wth thermal power and 30-40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗
NASA NTRS2021

Thermal Systems Modeling of Chemical Heat Integrated Power Source (CHIPS) to Survive Lunar Night Environments

This paper presents the results of the systems level thermal modeling for a conceptual Chemical Heat Integrated Power Source (CHIPS). This proposed system offers a combined thermal and electrical power source to support survival of spacecraft operating in extreme low temperature lunar environments, without the use of radioisotope-based sources. A conceptual design study has been completed for this system, that uses heat generated by an exothermic chemical reaction in place of radioisotope or electrical heat sources. The goal of the study was to evaluate the feasibility of such a system through thermodynamic and chemical analysis and thermal modeling, and to identify technology gaps to inform a technology development and maturation plan. The specific technical objectives focused on delivery of 90 to 100 Wth thermal power and 30 to 40 We electrical power for 336 hours to a representative Commercial Lunar Payload Services (CLPS) lander, to support lunar surface survival and limited operations through a lunar night. The total system mass was targeted at ≤50 kg. A highly exothermic chemical reaction system is used to generate on-board electrical power for spacecraft systems, and thermal power to maintain critical spacecraft/lander systems within their allowable flight temperature (AFTs). Based on the very high energy content of the chemical reaction system, a much higher amount of heat per unit mass can be delivered to the spacecraft, relative to a rechargeable lithium-ion battery and electrical heater(s). Since radioisotope heaters or generators are not used, the system will be orders of magnitude lower in cost than a radioisotope heating/power unit, without the attendant regulatory complexities. As part of the CHIPS concept, a fraction of the thermal power generated is converted to electrical power via an appropriate thermal-to-electric converter technology (such as a free piston Stirling converter or a thermoelectric generator module), to provide power to critical loads. This approach is ideally suited to support operation of commercial landers (e.g., via the CLPS program). In most cases, these landers are only designed to operate during a portion of the lunar day, with no provision for survival through the lunar night. By supporting lunar night survival, a mission can be extended through the lunar night and at least into another lunar day, thus turning a nominal eight-day mission into a 36-day mission. Therefore, the current system demonstration will focus on scalability to support at least 336 hours (one lunar night) of continuous thermal and electrical power generation. Although initially targeted to support lunar equatorial landings, the technology is extensible to missions at the lunar poles, other extreme environments in the Solar System or even air-independent applications on Earth (e.g., ocean exploration).

West, William↗
NASA NTRSDate not supplied

Thermoradiative Conversion for Space Power Systems

The thermoradiative cell is a new method for converting heat energy to electrical power, first detailed by Strandberg in 2015. The cell is structurally similar to a photovoltaic cell, in that it is a p-n junction semiconductor device, but thermodynamically operates in the reverse direction, converting the thermal dark current into electrical power by utilizing the recombination radiation from thermally-generated electron hole pairs to radiate waste heat to space. This technology may have application for space missions in converting thermal energy produced by a radioisotope source or from a nuclear reactor into power. The power and efficiency can be calculated as a function of bandgap in the detailed-balance case (the Shockley-Queisser limit), in which all of the thermal emissivity of the cell is due to the recombination of thermally generated electron-hole pairs, and all other recombination losses are ignored. The current produced is directly proportional to the recombination radiation, and thus the more thermally generated pairs, the higher the current. The voltage is proportional to the external bias. These two constraints allow optimization of the optimum bias point for maximum power, and allow calculation of the efficiency at maximum power point. Unlike photovoltaic cells, the maximum power operating point is not the same as the maximum efficiency point, and higher efficiency can be achieved at a higher (negative) bias in the ideal case. Incorporating non-ideal losses, however, shifts the maximum efficiency point toward lower bias. Since a thermoradiative cell operates by radiating directly to space, the current produced by a themoradiative cells will increase with the Stefan-Boltzman radiative efficiency; roughly the fourth power of the temperature. Thus, in contrast to a photovoltaic converter, the power produced is highest at high operating temperatures. Likewise, in contrast to conventional thermal conversion, high radiator temperature increases, rather than decreases the efficiency. Thus, the thermoradiative conversion may fill a mission niche in which small radiator size is required. The basic operation will be summarized, applications to space power discussed, and the requirements for further research outlined.

Photovoltaic Cell↗
NASA NTRSDate not supplied

NASA POWER: Providing Analysis-Ready, Cloud-Optimized Data for AI /ML Training and Applications in Earth Science

As global demand for sustainable development grows, the integration of Earth Observation (EO) data into decision making frameworks has become a primary objective for the scientific community. The NASA Prediction of Worldwide Energy Resources (POWER) project serves as a bridge between NASA EO data and the specialized needs of the renewable energy, sustainable infrastructure and agroclimatology communities. In this poster presentation we will present an overview of POWER data products and services along with its use in diverse research to decision-making workflows. By providing over 40 years of high-resolution historical, hourly and daily solar and meteorological data, POWER transforms satellite observations and global model reanalysis into actionable, Analysis-Ready Dataset (ARD). Currently, the project delivers over 250 industry-friendly parameters to the users from different NASA datasets like CERES SYN1Deg, MERRA-2, and IMERG alongside downscaled CMIP6 climate model data, fulfilling over 16 million requests from 50,000 unique users monthly. To ensure data quality and traceability, these parameters are rigorously validated against the ground-based observations from the Baseline Surface Radiation Network (BSRN) and the Global Surface Summary of the Day (GSOD) – these results will be discussed in the presentation. A newly introduced web-based PaRameter Uncertainty ViEwer (PRUVE) tool will be presented that provides an online validation platform to the users that benchmarks satellite-based and assimilation data products against these surface measurements. To reduce technical barriers to data adoption, POWER data is accessible through RESTful APIs, ESRI ArcGIS Image Services, a web-based Data Access Viewer tool, allowing users to visualize, validate and apply the dataset. For efficient data delivery POWER data is cloud-optimized into Zarr datastore accessible through NASA managed Amazon S3 ensures high-performance allowing users to integrate EO directly into operational pipelines. These customized services will be presented. Use cases from application will be presented from the energy sector - such as for design of generation systems, performance monitoring of solar power plants, in infrastructure sector- optimizing building energy efficiency and thermal comfort, in agriculture – such as driving crop simulation and yield forecasting models to enable climate resilient farming. Furthermore, the shift toward machine learning (ML) in EO research that has positioned POWER as a key provider for training datasets which will be discussed. Use-cases will be presented to showcase how NASA data is enabling the development of predictive tools for climate variability and resource management. The poster will present POWER’s future plans including technology development to enhance data traceability and reproducibility and improving I/O performance to support the rapid integration of new EO products, ensuring that POWER remains a robust scalable backend for the evolving landscape of AI-driven Earth Science. Additionally, POWER is developing an AI Agent and an MCP-Server to enable industry AI-Agentic workflows.

Neha Khadka↗
DOE OSTI2022

Technologies for providing secure emergency power control of high voltage direct current transmission system

Technologies for providing secure emergency power control of a high voltage direct current transmission (HVDC) system include a controller. The controller includes circuitry configured to receive status data indicative of a present physical status of a power system. The circuitry is also configured to obtain an emergency power control command triggered by a remote source. The emergency power control command is to be executed by an HVDC transmission system of the power system. Further, the circuitry is configured to determine, as a function of the status data, whether the emergency power control command is consistent with the present physical status of the power system and block, in response to a determination that the emergency power control command is not consistent with the present physical status of the power system, execution of the emergency power control command by the HVDC transmission system.

Pan, Jiuping↗