Results for “power electronics control” — Page 7

NASA NTRSDate not supplied

Development and Qualification Status of the Electric Propulsion Systems for the NASA PPE Mission and Gateway Program

NASA is charged with landing the first American woman and next American man on the South Pole of the Moon and establishing sustainable lunar exploration by the end of the decade. To meet this challenge, NASA’s Gateway will develop and deploy critical infrastructure required for operations on the lunar surface and that enables a sustained presence on and around the moon. NASA’s Power and Propulsion Element (PPE), the first planned element of NASA’s cis-lunar Gateway, leverages prior and ongoing NASA and U.S. industry investments in high-power, long-life solar electric propulsion technology investments. NASA awarded a PPE contract to Maxar Technologies to provide a 50 kW-class SEP spacecraft that meets Gateway’s needs, aligns with industry’s heritage spacecraft buses, and allows extensibility for NASA’s Mars exploration goals. Maxar’s PPE concept design, is based on their high heritage, modular, and highly reliable 1300-series bus architecture. The electric propulsion system features three 12 kW Advanced Electric Propulsion (AEPS) thrusters from Aerojet Rocketdyne and four BHT-6000 thrusters from Busek. Maxar-provided power electronics and xenon flow controllers from Moog are utilized in both the 12kW and 6kW electric propulsion strings on the spacecraft. The paper will present overviews of NASA’s Gateway and the PPE Project, status of the development and qualification activities for the two electric propulsion system, and the planned implementation of PPE electric propulsion system as keystone of NASA’s Gateway. The PPE spacecraft is currently heading into the Critical Design Review, with the qualification and flight electric propulsion hardware fabrication already initiated and significant progress being made toward planned qualifications in support of the planned PPE spacecraft co-manifest launch in 2024.

Electric Propulsion↗

NASA NTRSDate not supplied

Development and Qualification Status of the Electric Propulsion Systems for the NASA PPE Mission and Gateway Program

NASA is charged with landing the first American woman and next American man on the South Pole of the Moon and establishing sustainable lunar exploration by the end of the decade. To meet this challenge, NASA’s Gateway will develop and deploy critical infrastructure required for operations on the lunar surface and that enables a sustained presence on and around the moon. NASA’s Power and Propulsion Element (PPE), the first planned element of NASA’s cis-lunar Gateway, leverages prior and ongoing NASA and U.S. industry investments in high-power, long-life solar electric propulsion technology investments. NASA awarded a PPE contract to Maxar Technologies to provide a 50 kW-class SEP spacecraft that meets Gateway’s needs, aligns with industry’s heritage spacecraft buses, and allows extensibility for NASA’s Mars exploration goals. Maxar’s PPE concept design, is based on their high heritage, modular, and highly reliable 1300-series bus architecture. The electric propulsion system features three 12 kW Advanced Electric Propulsion (AEPS) thrusters from Aerojet Rocketdyne and four BHT-6000 thrusters from Busek. Maxar-provided power electronics and xenon flow controllers from Moog are utilized in both the 12kW and 6kW electric propulsion strings on the spacecraft. The paper will present overviews of NASA’s Gateway and the PPE Project, status of the development and qualification activities for the two electric propulsion system, and the planned implementation of PPE electric propulsion system as keystone of NASA’s Gateway. The PPE spacecraft is currently heading into the Critical Design Review, with the qualification and flight electric propulsion hardware fabrication already initiated and significant progress being made toward planned qualifications in support of the planned PPE spacecraft co-manifest launch in 2024.

Electric Propulsion↗

NASA NTRS1978

Electromechanical flight control actuator, volume 2

Schematic diagrams are given for both the four-channel electromechanical actuator and the single-channel power electronics breadboard. Detailed design data is also given on the gears used in the differential gearbox and a copy of the operations manual for the system is included. Performance test results are given for the EMA motor and its current source indicator, the drive control electronics, and the overall system. The power converter waveform test results are also summarized.

Source record↗

NASA NTRS2005

Evaluation of Power Electronic Components and Systems at Cryogenic Temperatures For Space Missions

Power electronic circuits and systems designed for deep space applications and outer planetary exploration are required to operate reliably and efficiently under extreme temperature conditions. This requirement is dictated by the fact that the operational environments associated with some of the space missions would encompass temperatures as low as -183 C. The development and utilization of electronics capable of low temperature operation would not only fulfill the advanced technology requirements, but also would contribute to improving circuit performance, increasing system efficiency, and reducing development and launch costs. These benefits are generally achieved by the improved intrinsic properties of some of the electronic materials at low temperature, reduced device losses, and the elimination of heating elements used in conventional systems at low temperatures. Power electronic circuits are widely used in space power systems in the areas of power management, conditioning, and control. In this work, the performance of certain power electronic components and systems was investigated under low temperature. These include inductors, capacitors, pulse-width-modulation (PWM) controllers, and advanced commercial DC/DC converter modules. Different properties were determined as a function of temperature in the range of 20 C to -140 C, at various current and voltages levels. The experimental procedures along with the experimental data obtained are presented and discussed in this paper.

Elbuluk, Malik E.↗

NASA NTRS2007

Thin, nearly wireless adaptive optical device

A thin, nearly wireless adaptive optical device capable of dynamically modulating the shape of a mirror in real time to compensate for atmospheric distortions and/or variations along an optical material is provided. The device includes an optical layer, a substrate, at least one electronic circuit layer with nearly wireless architecture, an array of actuators, power electronic switches, a reactive force element, and a digital controller. Actuators are aligned so that each axis of expansion and contraction intersects both substrate and reactive force element. Electronics layer with nearly wireless architecture, power electronic switches, and digital controller are provided within a thin-film substrate. The size and weight of the adaptive optical device is solely dominated by the size of the actuator elements rather than by the power distribution system.

Knowles, Gareth↗

NASA NTRS2008

Thin, nearly wireless adaptive optical device

A thin, nearly wireless adaptive optical device capable of dynamically modulating the shape of a mirror in real time to compensate for atmospheric distortions and/or variations along an optical material is provided. The device includes an optical layer, a substrate, at least one electronic circuit layer with nearly wireless architecture, an array of actuators, power electronic switches, a reactive force element, and a digital controller. Actuators are aligned so that each axis of expansion and contraction intersects both substrate and reactive force element. Electronics layer with nearly wireless architecture, power electronic switches, and digital controller are provided within a thin-film substrate. The size and weight of the adaptive optical device is solely dominated by the size of the actuator elements rather than by the power distribution system.

Knowles, Gareth↗

NASA NTRS2009

Thin nearly wireless adaptive optical device

A thin nearly wireless adaptive optical device capable of dynamically modulating the shape of a mirror in real time to compensate for atmospheric distortions and/or variations along an optical material is provided. The device includes an optical layer, a substrate, at least one electronic circuit layer with nearly wireless architecture, an array of actuators, power electronic switches, a reactive force element, and a digital controller. Actuators are aligned so that each axis of expansion and contraction intersects both substrate and reactive force element. Electronics layer with nearly wireless architecture, power electronic switches, and digital controller are provided within a thin-film substrate. The size and weight of the adaptive optical device is solely dominated by the size of the actuator elements rather than by the power distribution system.

Knowles, Gareth J.↗

NASA NTRS1988

Successful completion of a cyclic ground test of a mercury ion auxiliary propulsion system

An engineering model Ion Auxiliary Propulsion System (IAPS) 8-cm thruster (S/N 905) has completed a life test at NASA Lewis Research Center. The mercury ion thruster successfully completed and exceeded the test goals of 2557 on/off cycles and 7057 hr of operation at full thrust. The final 1200 cycles and 3600 hr of the life test were conducted using an engineering model of the IAPS power electronics unit (PEU) and breadboard digital controller and interface unit (DCIU). This portion of the test is described in this paper with a charted history of thruster operating parameters and off-normal events. Performance and operating characteristics were constant throughout the test with only minor variations. The engineering model power electronics unit operated without malfunction; the flight software in the digital controller and interface unit was exercised and verified. Post-test inspection of the thruster revealed facility enhanced accelerator grid erosion but overall the thruster was in good condition. It was concluded that the thruster performance was not drastically degraded by time or cycles. Additional cyclic testing is currently under consideration.

Francisco, David R.↗

NASA NTRS1988

Successful completion of a cyclic ground test of a mercury Ion Auxiliary Propulsion System

An engineering model Ion Auxiliary Propulsion System (IAPS) 8-cm thruster (S/N 905) has completed a life test at NASA Lewis Research Center. The mercury ion thruster successfully completed and exceeded the test goals of 2557 on/off cycles and 7057 hr of operation at full thrust. The final 1200 cycles and 3600 hr of the life test were conducted using an engineering model of the IAPS power electronics unit (PEU) and breadboard digital controller and interface unit (DCIU). This portion of the test is described in this paper with a charted history of thruster operating parameters and off-normal events. Performance and operating characteristics were constant throughout the test with only minor variations. The engineering model power electronics unit operated without malfunction; the flight software in the digital controller and interface unit was exercised and verified. Post-test inspection of the thruster revealed facility enhanced accelerator grid erosion but overall the thruster was in good condition. It was concluded that the thruster performance was not drastically degraded by time or cycles. Additional cyclic testing is currently under consideration.

Francisco, David R.↗

NASA NTRS1995

Control Electronics For Reaction Wheel

Bidirectional operation achieved with single-polarity main power supply. Control circuitry generates pulse-width-modulated 800-Hz waveforms to drive two-phase ac motor and reaction wheel. Operates partly in response to digital magnitude-and-direction torque command generated by external control subsystem and partly in response to tachometric feedback in form of two once-per-revolution sinusoids with amplitudes proportional to speed. Operation in either of two modes called "normal" and "safehold." In normal mode, drive pulses timed so that, on average over one or few cycles, motor applies commanded torque. In safehold mode, pulses timed to keep motor running at set speed in one direction.

Chamberlin, Keith↗

NASA NTRS1976

Investigation of the use of an electronic multifunction display and an electromechanical horizontal situation indicator for guidance and control of powered-lift short-haul aircraft

The use which pilots make of a moving map display from en route through the terminal area and including the approach and go-around flight phases was investigated. The content and function of each of three primary STOLAND displays are reviewed from an operational point of view. The primary displays are the electronic attitude director indicator (EADI), the horizontal situation indicator (HSI), and the multifunction display (MFD). Manually controlled flight with both flight director guidance and raw situation data is examined in detail in a simulated flight experiment with emphasis on tracking reference flight plans and maintaining geographic orientation after missed approaches. Eye-point-of-regard and workload measurements, coupled with task performance measurements, pilot opinion ratings, and pilot comments are presented. The experimental program was designed to offer a systematic objective and subjective comparison of pilots' use of the moving map MFD in conjunction with the other displays.

Clement, W. F.↗

NASA NTRS2012

Compact Autonomous Hemispheric Vision System

Solar System Exploration camera implementations to date have involved either single cameras with wide field-of-view (FOV) and consequently coarser spatial resolution, cameras on a movable mast, or single cameras necessitating rotation of the host vehicle to afford visibility outside a relatively narrow FOV. These cameras require detailed commanding from the ground or separate onboard computers to operate properly, and are incapable of making decisions based on image content that control pointing and downlink strategy. For color, a filter wheel having selectable positions was often added, which added moving parts, size, mass, power, and reduced reliability. A system was developed based on a general-purpose miniature visible-light camera using advanced CMOS (complementary metal oxide semiconductor) imager technology. The baseline camera has a 92 FOV and six cameras are arranged in an angled-up carousel fashion, with FOV overlaps such that the system has a 360 FOV (azimuth). A seventh camera, also with a FOV of 92 , is installed normal to the plane of the other 6 cameras giving the system a > 90 FOV in elevation and completing the hemispheric vision system. A central unit houses the common electronics box (CEB) controlling the system (power conversion, data processing, memory, and control software). Stereo is achieved by adding a second system on a baseline, and color is achieved by stacking two more systems (for a total of three, each system equipped with its own filter.) Two connectors on the bottom of the CEB provide a connection to a carrier (rover, spacecraft, balloon, etc.) for telemetry, commands, and power. This system has no moving parts. The system's onboard software (SW) supports autonomous operations such as pattern recognition and tracking.

Pingree, Paula J.↗

NASA NTRS1991

Large space structures and systems in the space station era: A bibliography with indexes (supplement 03)

Bibliographies and abstracts are listed for 1221 reports, articles, and other documents introduced into the NASA scientific and technical information system between January 1, 1991 and June 30, 1991. Topics covered include large space structures and systems, space stations, extravehicular activity, thermal environments and control, tethering, spacecraft power supplies, structural concepts and control systems, electronics, advanced materials, propulsion, policies and international cooperation, vibration and dynamic controls, robotics and remote operations, data and communication systems, electric power generation, space commercialization, orbital transfer, and human factors engineering.

Source record↗

NASA NTRS2001

Integrated Power Source Grant

Traditional spacecraft power systems incorporate a solar array energy source, an energy storage element (battery), and battery charge control and bus voltage regulation electronics to provide continuous electrical power for spacecraft systems and instruments. Dedicated power conditioning components provide limited fault isolation between systems and instruments, while a centralized power-switching unit provides spacecraft load control. Battery undervoltage conditions are detected by the spacecraft processor, which removes fault conditions and non-critical loads before permanent battery damage can occur. Cost effective operation of a micro-sat constellation requires a fault tolerant spacecraft architecture that minimizes on-orbit operational costs by permitting autonomous reconfiguration in response to unexpected fault conditions. A new micro-sat power system architecture that enhances spacecraft fault tolerance and improves power system survivability by continuously managing the battery charge and discharge processes on a cell-by-cell basis has been developed. This architecture is based on the Integrated Power Source (US patent 5644207), which integrates dual junction solar cells, Lithium Ion battery cells, and processor based charge control electronics into a structural panel that can be deployed or used to form a portion of the outer shell of a micro-spacecraft. The first generation Integrated Power Source is configured as a one inch thick panel in which prismatic Lithium Ion battery cells are arranged in a 3x7 matrix (26VDC) and a 3x1 matrix (3.7VDC) to provide the required output voltages and load currents. A multi-layer structure holds the battery cells, as well as the thermal insulators that are necessary to protect the Lithium Ion battery cells from the extreme temperatures of the solar cell layer. Independent thermal radiators, located on the back of the panel, are dedicated to the solar cell array, the electronics, and the battery cell array. In deployed panel applications, these radiators maintain the battery cells in an appropriate operational temperature range.

Source record↗

NASA NTRS2011

Electro-Mechanical Systems for Extreme Space Environments

Exploration beyond low earth orbit presents challenges for hardware that must operate in extreme environments. The current state of the art is to isolate and provide heating for sensitive hardware in order to survive. However, this protection results in penalties of weight and power for the spacecraft. This is particularly true for electro-mechanical based technology such as electronics, actuators and sensors. Especially when considering distributed electronics, many electro-mechanical systems need to be located in appendage type locations, making it much harder to protect from the extreme environments. The purpose of this paper to describe the advances made in the area of developing electro-mechanical technology to survive these environments with minimal protection. The Jet Propulsion Lab (JPL), the Glenn Research Center (GRC), the Langley Research Center (LaRC), and Aeroflex, Inc. over the last few years have worked to develop and test electro-mechanical hardware that will meet the stringent environmental demands of the moon, and which can also be leveraged for other challenging space exploration missions. Prototype actuators and electronics have been built and tested. Brushless DC actuators designed by Aeroflex, Inc have been tested with interface temperatures as low as 14 degrees Kelvin. Testing of the Aeroflex design has shown that a brushless DC motor with a single stage planetary gearbox can operate in low temperature environments for at least 120 million cycles (measured at motor) if long life is considered as part of the design. A motor control distributed electronics concept developed by JPL was built and operated at temperatures as low as -160 C, with many components still operational down to -245 C. Testing identified the components not capable of meeting the low temperature goal of -230 C. This distributed controller is universal in design with the ability to control different types of motors and read many different types of sensors. The controller form factor was designed to surround or be at the actuator. Communication with the slave controllers is accomplished by a bus, thus limiting the number of wires that must be routed to the extremity locations. Efforts have also been made to increase the power capability of these electronics for the ability to power and control actuators up to 2.5KW and still meet the environmental challenges. For commutation and control of the actuator, a resolver was integrated and tested with the actuator. Testing of this resolver demonstrated temperature limitations. Subsequent failure analysis isolated the low temperature failure mechanism and a design solution was negotiated with the manufacturer. Several years of work have resulted in specialized electro-mechanical hardware to meet extreme space exploration environments, a test history that verifies and finds limitations of the designs and a growing knowledge base that can be leveraged by future space exploration missions.

Mojarradi, Mohammad M.↗

NASA NTRS1986

Power electronics for a 1-kilowatt arc jet thruster

After more than two decades, new space mission requirements have revived interest in arcjet systems. The preliminary development and demonstration of new, high efficiency, power electronic concepts for start up and steady state control of dc arcjets is reported. The design comprises a pulse width modulated power converter which is closed loop configured to give fast current control. An inductor, in series with the arcjet, serves the dual role of providing instantaneous current control, as well as a high voltage arc ignition pulse. Benchmark efficiency, transient response, regulation, and ripple data are presented. Tests with arcjets demonstrate that the power electronics breadboard can start thrusters consistently with no apparent damage and transfer reliably to the nondestructive high voltage arc mode in less than a second.

Gruber, R. P.↗

NASA NTRS1986

Power electronics for a 1-kilowatt arcjet thruster

After more than two decades, new space mission requirements have revived interest in arcjet systems. The preliminary development and demonstration of new, high efficiency, power electronic concepts for start up and steady state control of dc arcjets is reported. The design comprises a pulse width modulated power converter which is closed loop configured to give fast current control. An inductor, in series with the arcjet, serves the dual role of providing instantaneous current control, as well as a high voltage arc ignition pulse. Benchmark efficiency, transient response, regulation, and ripple data are presented. Tests with arcjets demonstrate that the power electronics breadboard can start thrusters consistently with no apparent damage and transfer reliably to the nondestructive high voltage arc mode in less than a second.

Gruber, R. P.↗

NASA NTRS1978

The 30-cm ion thruster power processor

A power processor unit for powering and controlling the 30 cm Mercury Electron-Bombardment Ion Thruster was designed, fabricated, and tested. The unit uses a unique and highly efficient transistor bridge inverter power stage in its implementation. The system operated from a 200 to 400 V dc input power bus, provides 12 independently controllable and closely regulated dc power outputs, and has an overall power conditioning capacity of 3.5 kW. Protective circuitry was incorporated as an integral part of the design to assure failure-free operation during transient and steady-state load faults. The implemented unit demonstrated an electrical efficiency between 91.5 and 91.9 at its nominal rated load over the 200 to 400 V dc input bus range.

Herron, B. G.↗