Engineering topics

Pinero, Luis R.

Publications and source records attributed to Pinero, Luis R..

At least 19 records

NASA NTRS2019

Development of a High-Propellant Throughput Small Spacecraft Electric Propulsion System to Enable Lower Cost NASA Science Missions

This paper describes recent progress at the NASA Glenn Research Center (GRC) in the development and demonstration of an integrated high-propellant throughput small spacecraft electric propulsion (HT-SSEP) system based on a Hall-effect thruster. A center-mounted cathode and an innovative magnetic circuit topology were implemented in the design of the Hall-effect thruster to achieve high-propellant throughput, high performance, and efficient packaging. To minimize technical risk, the HT-SSEP development approach sought to limit design features and materials to those with a clear path-to-flight. A propellant throughput capability of greater than 100 kg at a minimum thruster efficiency of 50% was targeted. The proof-of-concept NASA-H64M laboratory model (LM) thruster was designed, fabricated, and tested at GRC in fiscal year 2018. The thruster development leveraged heritage Hall-effect thruster design and manufacturing processes wherever appropriate. Recent NASA advances in Hall-effect thruster technology were also leveraged. A scalable discharge power supply (DPS) capable of powering the H64M-LM was developed, then demonstrated as part of an integrated system test. The DPS uses commercial off-the-shelf components with spaceflight equivalents. A keeper supply with DC ignitor was breadboarded, then demonstrated with a laboratory cathode. Finally, feed system trade studies were performed to ascertain what feed system architecture might be appropriate for an HT-SSEP system. This paper details the motivations for the project, the development approach, the chosen sub-system architectures, design considerations, and test results.

Benavides, Gabriel F.↗

NASA NTRS2019

Development of a High-Propellant Throughput Small Spacecraft Electric Propulsion System to Enable Lower Cost NASA Science Missions

Benavides, Gabriel F.↗

NASA NTRS2017

The Impact of Harness Impedance on Hall Thruster Discharge Oscillations

Hall thrusters exhibit characteristic discharge voltage and current oscillations during steady-state operation. The lower frequency breathing-mode current oscillations are inherent to each thruster and could impact thruster operation and PPU design. The design of the discharge output filter, in particular, the output capacitor is important because it supplies the high peak current oscillations that the thruster demands. However, space-rated, high-voltage capacitors are not readily available and can have significant mass and volume. So, it is important for a PPU designer to know what is the minimum amount of capacitance required to operate a thruster. Through SPICE modeling and electrical measurements on the Hall Effect Rocket with Magnetic Shielding (HERMeS) thruster, it was shown that the harness impedance between the power supply and the thruster is the main contributor towards generating voltage ripple at the thruster. Also, increasing the size of the discharge filter capacitor, as previously implemented during thruster tests, does not reduce the voltage oscillations. The electrical characteristics of the electrical harness between the discharge supply and the thruster is crucial to system performance and could have a negative impact on performance, life and operation.

Pinero, Luis R.↗

NASA NTRS2017

The Impact of Harness Impedance on Hall Thruster Discharge Oscillations

Hall thrusters exhibit characteristic discharge voltage and current oscillations during steady-state operation. The lower frequency breathing-mode current oscillations are inherent to each thruster and could impact thruster operation and power processing unit (PPU) design. The design of the discharge output filter, in particular, the output capacitor is important because it supplies the high peak current oscillations that the thruster demands. However, space-rated, high-voltage capacitors are not readily available and can have significant mass and volume. So, it is important for a PPU designer to know what is the minimum amount of capacitance required to operate a thruster. Through Simulation Program with Integrated Circuit Emphasis modeling and electrical measurements on the Hall Effect Rocket with Magnetic Shielding thruster, it was shown that the harness impedance between the power supply and the thruster is the main contributor towards generating voltage ripple at the thruster. Also, increasing the size of the discharge filter capacitor, as previously implemented during thruster tests, does not reduce the voltage oscillations. The electrical characteristics of the electrical harness between the discharge supply and the thruster is crucial to system performance and could have a negative impact on performance, life and operation.

hollow cathodes↗

NASA NTRS2017

The Impact of Harness Impedance on Hall Thruster Discharge Oscillations

Hall thrusters exhibit characteristic discharge voltage and current oscillations during steady-state operation. The lower frequency breathing-mode current oscillations are inherent to each thruster and could impact thruster operation and power processing unit (PPU) design. The design of the discharge output filter, in particular, the output capacitor is important because it supplies the high peak current oscillations that the thruster demands. However, space-rated, high-voltage capacitors are not readily available and can have significant mass and volume. So, it is important for a PPU designer to know what is the minimum amount of capacitance required to operate a thruster. Through Simulation Program with Integrated Circuit Emphasis modeling and electrical measurements on the Hall Effect Rocket with Magnetic Shielding thruster, it was shown that the harness impedance between the power supply and the thruster is the main contributor towards generating voltage ripple at the thruster. Also, increasing the size of the discharge filter capacitor, as previously implemented during thruster tests, does not reduce the voltage oscillations. The electrical characteristics of the electrical harness between the discharge supply and the thruster is crucial to system performance and could have a negative impact on performance, life and operation.

ion engines↗

NASA NTRS2016

Status of the Development of Flight Power Processing Units for the NASAs Evolutionary Xenon Thruster - Commercial (NEXT-C) Project

A pathfinder prototype unit and two flight power processing units (PPUs) are being developed by the Aerojet Rocketdyne Corporation in Redmond, Washington and ZIN Technologies in Cleveland, Ohio, in support of the NEXT-C Project. This project is being led by the NASA Glenn Research Center in Cleveland, Ohio, and will also yield two flight thrusters. This hardware is being considered to be provided as Government Furnished Equipment for the New Frontiers Program, and is applicable to a variety of planetary science missions and astrophysics science missions. The design of the NEXT-C PPU evolves from the hardware fabricated under the NEXT technology development project. The power processing unit operates from two sources: a wide input 80 to 160 V high-power bus and a nominal 28 V low-power bus. The unit includes six power supplies. Four power supplies (beam, accelerator, discharge, and neutralizer keeper) are needed for steady state operation, while two cathode heater power supplies (neutralizer and discharge) are utilized during thruster startup. The unit in total delivers up to 7 kW of regulated power to a single gridded-ion thruster. Significant modifications to the initial design include: high-power adaptive-delay control, upgrade of design to EEE-INST-002 compliance, telemetry accuracy improvements, incorporation of telemetry to detect plume-mode operation, and simplification of the design in select areas to improve manufacturability and commercialization potential. The project is presently in the prototype phase and preparing for qualification level environmental testing.

NASAaEuro(TM)s Evolutionary Xenon Thruster↗

NASA NTRS2016

Performance of a High-Fidelity 4kW-Class Engineering Model PPU and Integration with HiVHAc System

The High Voltage Hall Accelerator (HiVHAc) propulsion system consists of a thruster, power processing unit (PPU), and propellant feed system. An engineering model PPU was developed by Colorado Power Electronics, Inc. funded by NASA's Small Business Innovative Research Program. This PPU uses an innovative 3-phase resonant converter to deliver 4 kW of discharge power over a wide range of input and output voltage conditions. The PPU includes a digital control interface unit that automatically controls the PPU and a xenon flow control module (XFCM). It interfaces with a control computer to receive highlevel commands and relay telemetry through a MIL-STD-1553B interface. The EM PPU was thoroughly tested at GRC for functionality and performance at temperature limits and demonstrated total efficiencies a high as 95 percent. Integrated testing of the unit was performed with the HiVHAc thruster and the XFCM to demonstrate closed-loop control of discharge current with anode flow. Initiation of the main discharge and power throttling were also successfully demonstrated and discharge oscillations were characterized.

Power Processing Unit↗

NASA NTRS2016

Performance of a High-Fidelity 4kW-Class Engineering Model PPU and Integration with HiVHAc System

The High Voltage Hall Accelerator (HiVHAc) propulsion system consists of a thruster,power processing unit (PPU), and propellant feed system. An engineering model PPU was developed by Colorado Power Electronics, Inc. funded by NASA's Small Business Innovative Research Program. This PPU uses an innovative 3-phase resonant converter to deliver 4 kW of discharge power over a wide range of input and output voltage conditions.The PPU includes a digital control interface unit that automatically controls the PPU and a xenon flow control module (XFCM). It interfaces with a control computer to receive high level commands and relay telemetry through a MIL-STD-1553B interface. The EM PPU was thoroughly tested at GRC for functionality and performance at temperature extremes and demonstrated total efficiencies a high as 95 percent. It was integrated with the HiVHAc thruster and the XFCM to demonstrate closed-loop control of discharge current with anode flow. Initiation of the main discharge and power throttling were also successfully demonstrated and discharge oscillations were characterized.

Electric Propulsion↗

NASA NTRS2015

High Power Silicon Carbide (SiC) Power Processing Unit Development

NASA GRC successfully designed, built and tested a technology-push power processing unit for electric propulsion applications that utilizes high voltage silicon carbide (SiC) technology. The development specifically addresses the need for high power electronics to enable electric propulsion systems in the 100s of kilowatts. This unit demonstrated how high voltage combined with superior semiconductor components resulted in exceptional converter performance.

Power Processing↗

NASA NTRS2015

High Input Voltage, Silicon Carbide Power Processing Unit Performance Demonstration

A silicon carbide brassboard power processing unit has been developed by the NASA Glenn Research Center in Cleveland, Ohio. The power processing unit operates from two sources: a nominal 300 Volt high voltage input bus and a nominal 28 Volt low voltage input bus. The design of the power processing unit includes four low voltage, low power auxiliary supplies, and two parallel 7.5 kilowatt (kW) discharge power supplies that are capable of providing up to 15 kilowatts of total power at 300 to 500 Volts (V) to the thruster. Additionally, the unit contains a housekeeping supply, high voltage input filter, low voltage input filter, and master control board, such that the complete brassboard unit is capable of operating a 12.5 kilowatt Hall effect thruster. The performance of the unit was characterized under both ambient and thermal vacuum test conditions, and the results demonstrate exceptional performance with full power efficiencies exceeding 97%. The unit was also tested with a 12.5kW Hall effect thruster to verify compatibility and output filter specifications. With space-qualified silicon carbide or similar high voltage, high efficiency power devices, this would provide a design solution to address the need for high power electric propulsion systems.

Power↗

NASA NTRS2015

Development of High-Power Hall Thruster Power Processing Units at NASA GRC

NASA GRC successfully designed, built and tested four different power processor concepts for high power Hall thrusters. Each design satisfies unique goals including the evaluation of a novel silicon carbide semiconductor technology, validation of innovative circuits to overcome the problems with high input voltage converter design, development of a direct-drive unit to demonstrate potential benefits, or simply identification of lessonslearned from the development of a PPU using a conventional design approach. Any of these designs could be developed further to satisfy NASA's needs for high power electric propulsion in the near future.

Spacecraft Power Supplies↗

NASA NTRS2015

High Input Voltage, Silicon Carbide Power Processing Unit Performance Demonstration

A silicon carbide brassboard power processing unit has been developed by the NASA Glenn Research Center in Cleveland, Ohio. The power processing unit operates from two sources - a nominal 300-Volt high voltage input bus and a nominal 28-Volt low voltage input bus. The design of the power processing unit includes four low voltage, low power supplies that provide power to the thruster auxiliary supplies, and two parallel 7.5 kilowatt power supplies that are capable of providing up to 15 kilowatts of total power at 300-Volts to 500-Volts to the thruster discharge supply. Additionally, the unit contains a housekeeping supply, high voltage input filter, low voltage input filter, and master control board, such that the complete brassboard unit is capable of operating a 12.5 kilowatt Hall Effect Thruster. The performance of unit was characterized under both ambient and thermal vacuum test conditions, and the results demonstrate the exceptional performance with full power efficiencies exceeding 97. With a space-qualified silicon carbide or similar high voltage, high efficiency power device, this design could evolve into a flight design for future missions that require high power electric propulsion systems.

Technology↗

NASA NTRS2014

High Input Voltage Discharge Supply for High Power Hall Thrusters Using Silicon Carbide Devices

A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because they were the best choice to handle the high voltage stress while delivering high efficiency and low specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also developed and will be integrated with the discharge supply into a vacuum-rated brassboard power processing unit with full flight functionality. This design could be evolved into a flight unit for future missions that requires high power electric propulsion.

Electric Propulsion↗

NASA NTRS2012

NASA's Evolutionary Xenon Thruster (NEXT) Power Processing Unit (PPU) Capacitor Failure Root Cause Analysis

The NASA s Evolutionary Xenon Thruster (NEXT) project is developing an advanced ion propulsion system for future NASA missions for solar system exploration. A critical element of the propulsion system is the Power Processing Unit (PPU) which supplies regulated power to the key components of the thruster. The PPU contains six different power supplies including the beam, discharge, discharge heater, neutralizer, neutralizer heater, and accelerator supplies. The beam supply is the largest and processes up to 93+% of the power. The NEXT PPU had been operated for approximately 200+ hr and has experienced a series of three capacitor failures in the beam supply. The capacitors are in the same, nominally non-critical location-the input filter capacitor to a full wave switching inverter. The three failures occurred after about 20, 30, and 135 hr of operation. This paper provides background on the NEXT PPU and the capacitor failures. It discusses the failure investigation approach, the beam supply power switching topology and its operating modes, capacitor characteristics and circuit testing. Finally, it identifies root cause of the failures to be the unusual confluence of circuit switching frequency, the physical layout of the power circuits, and the characteristics of the capacitor.

Soeder, James F.↗

NASA NTRS2010

Integration Testing of a Modular Discharge Supply for NASA's High Voltage Hall Accelerator Thruster

NASA s In-Space Propulsion Technology Program is developing a high performance Hall thruster that can fulfill the needs of future Discovery-class missions. The result of this effort is the High Voltage Hall Accelerator thruster that can operate over a power range from 0.3 to 3.5 kW and a specific impulse from 1,000 to 2,800 sec, and process 300 kg of xenon propellant. Simultaneously, a 4.0 kW discharge power supply comprised of two parallel modules was developed. These power modules use an innovative three-phase resonant topology that can efficiently supply full power to the thruster at an output voltage range of 200 to 700 V at an input voltage range of 80 to 160 V. Efficiencies as high as 95.9 percent were measured during an integration test with the NASA103M.XL thruster. The accuracy of the master/slave current sharing circuit and various thruster ignition techniques were evaluated.

Pinero, Luis R.↗

NASA NTRS2009

NASA's Evolutionary Xenon Thruster (NEXT) Component Verification Testing

Component testing is a critical facet of the comprehensive thruster life validation strategy devised by the NASA s Evolutionary Xenon Thruster (NEXT) program. Component testing to-date has consisted of long-duration high voltage propellant isolator and high-cycle heater life validation testing. The high voltage propellant isolator, a heritage design, will be operated under different environmental condition in the NEXT ion thruster requiring verification testing. The life test of two NEXT isolators was initiated with comparable voltage and pressure conditions with a higher temperature than measured for the NEXT prototype-model thruster. To date the NEXT isolators have accumulated 18,300 h of operation. Measurements indicate a negligible increase in leakage current over the testing duration to date. NEXT 1/2 in. heaters, whose manufacturing and control processes have heritage, were selected for verification testing based upon the change in physical dimensions resulting in a higher operating voltage as well as potential differences in thermal environment. The heater fabrication processes, developed for the International Space Station (ISS) plasma contactor hollow cathode assembly, were utilized with modification of heater dimensions to accommodate a larger cathode. Cyclic testing of five 1/22 in. diameter heaters was initiated to validate these modified fabrication processes while retaining high reliability heaters. To date two of the heaters have been cycled to 10,000 cycles and suspended to preserve hardware. Three of the heaters have been cycled to failure giving a B10 life of 12,615 cycles, approximately 6,000 more cycles than the established qualification B10 life of the ISS plasma contactor heaters.

Herman, Daniel A.↗

NASA NTRS2007

NEXT Thruster Component Verification Testing

Component testing is a critical part of thruster life validation activities under NASA s Evolutionary Xenon Thruster (NEXT) project testing. The high voltage propellant isolators were selected for design verification testing. Even though they are based on a heritage design, design changes were made because the isolators will be operated under different environmental conditions including temperature, voltage, and pressure. The life test of two NEXT isolators was therefore initiated and has accumulated more than 10,000 hr of operation. Measurements to date indicate only a negligibly small increase in leakage current. The cathode heaters were also selected for verification testing. The technology to fabricate these heaters, developed for the International Space Station plasma contactor hollow cathode assembly, was transferred to Aerojet for the fabrication of the NEXT prototype model ion thrusters. Testing the contractor-fabricated heaters is necessary to validate fabrication processes for high reliability heaters. This paper documents the status of the propellant isolator and cathode heater tests.

Pinero, Luis R.↗

NASA NTRS2007

Performance of the NEXT Engineering Model Power Processing Unit

The NASA s Evolutionary Xenon Thruster (NEXT) project is developing an advanced ion propulsion system for future NASA missions for solar system exploration. An engineering model (EM) power processing unit (PPU) for the NEXT project was designed and fabricated by L-3 Communications under contract with NASA Glenn Research Center (GRC). This modular PPU is capable of processing up from 0.5 to 7.0 kW of output power for the NEXT ion thruster. Its design includes many significant improvements for better performance over the state-of-the-art PPU. The most significant difference is the beam supply which is comprised of six modules and capable of very efficient operation through a wide voltage range because of innovative features like dual controls, module addressing, and a high current mode. The low voltage power supplies are based on elements of the previously validated NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) PPU. The highly modular construction of the PPU resulted in improved manufacturability, simpler scalability, and lower cost. This paper describes the design of the EM PPU and the results of the bench-top performance tests.

Pinero, Luis R.↗