Roger Myers - Academia.edu (original) (raw)
Papers by Roger Myers
Journal of Applied Physics, 2021
Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have b... more Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have been identified as enabling technologies for applications such as lunar tugs, large satellite orbital transfer vehicles, and solar system exploration. These large thrusters introduce significant testing challenges due to the propellant flow rate exceeding the pumping speed available in most laboratories. Even with proposed upgrades in mind, the likelihood that multiple vacuum facilities will exist in the near future to allow long duration testing of high-power Hall thrusters operating at power levels in excess of 100 kW remains extremely low. In this article, we numerically explore the feasibility of testing Hall thrusters in a quasi-steady mode defined by pulsing the mass flow rate between a nominal and a low value. Our simulations indicate that sub-second durations available before the chamber reaches critical pressure are sufficiently long to achieve the steady-state current and flow f...
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017
The next phase of robotic and human deep space exploration missions requires high performance, hi... more The next phase of robotic and human deep space exploration missions requires high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdyne’s Advanced Electric Propulsion System (AEPS) program will complete development and qualification of a 13kW flight EP system to support NASA exploration. The first use of the AEPS is planned for the NASA Power & Propulsion Element, which is the first element of NASA’s cis-lunar Gateway. The flight AEPS system includes a magnetically shielded long-life Hall thruster, power processing unit (PPU), and xenon flow controller (XFC). The Hall thruster, originally developed and demonstrated by NASA’s Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 12.5kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation b...
Large scale cargo transportation to support human missions to the Moon and Mars will require very... more Large scale cargo transportation to support human missions to the Moon and Mars will require very high power Solar Electric Propulsion (SEP) systems operating between 200 and 400 kW. Aerojet Rocketdyne’s NextSTEP program is developing and demonstrating a 100 kW EP system, the XR-100, using a Nested Hall Thruster (NHT) designed for powers up to 200 kW, a modular power processor and a modular flow controller. The three year program objective is to operate the integrated EP system continuously at 100 kW for 100 h, advancing this very high power Electric Propulsion (EP) system to Technology Readiness Level (TRL) 5. With our University of Michigan, Jet Propulsion Laboratory and NASA Glenn Research Center teammates, Aerojet Rocketdyne has completed the initial phase of the program, including operating the thruster at up to 30 kW to validate the thermal models and developing and operating multiple power processor modules in the required series/parallel configuration. The current phase includes completing a TRL 4 integrated system test at reduced power to validate all system operating phases. Design upgrades to demonstrate the TRL 5 capabilities are underway. This paper will present the high power XR-100 capabilities, overall program and design approach and the latest test results for the 100 kW EP system demonstration program. 1 NextSTEP Program Manager, jerry.jackson@rocket.com. 2 NextSTEP Chief Engineer, may.allen@rocket.com. 3 NextSTEP Technical Consultant, roger.myers@rocket.com. 4 NextSTEP Technical Consultant, william.hoskins@rocket.com. 5 NextSTEP PPU Lead, erich.soendker@rocket.com. 6 NextSTEP System and Thruster Lead, benjamin.welander@rocket.com. 7 NextSTEP Feed System Lead, arturo.tolentino@rocket.com. 8 NextSTEP PPU Discharge Supply Lead, sam.hablitzel@rocket.com. 9 University of Michigan, Hall Thruster SME, shall@ umich.edu. 10 University of Michigan, Robert J. Vlassic Dean of Engineering, Richard F. and Eleanor A. Towner Professor, and Arthur F. Thurnau Professor, Department of Aerospace Engineering, and Laboratory Director, PEPL, alec.gallimore@umich.edu. 11 University of Michigan, Assistant Professor and Laboratory Co-Director, PEPL, bjorns@umich.edu. 12 NASA JPL Thruster Lead, richard.r.hofer@jpl.nasa.gov. 13 NASA JPL Hollow Cathode Lead, Dan.M.Goebel@jpl.nasa.gov. 14 NASA NextSTEP technical monitor, eric.pencil@nasa.gov. The 35th International Electric Propulsion Conference, Georgia Institute of Technology, USA October 8 – 12, 2017 Published by the Electric Rocket Propulsion Society with permission https://ntrs.nasa.gov/search.jsp?R=20180004449 2018-11-18T06:06:11+00:00Z
24th Joint Propulsion Conference, 1988
Mass loss and surface temperature measurements, filter photography and SEM surface characterizati... more Mass loss and surface temperature measurements, filter photography and SEM surface characterization are used to study cathode phenomena in a steady state self-field MPD thruster operated at power levels between 15 and 30 kW. The equilibrium cathode temperature is found to be approximately 3200 K over most of its length, for which evaporation and thermionic emission adequately explain measured erosion rates and current levels. The start-up phase is characterized by spot current attachment with associated high erosion rates.
This material may be reproduced by or for the U.S. Government pursuant to the copyright license u... more This material may be reproduced by or for the U.S. Government pursuant to the copyright license under the clause at FAR 52.227-14 [Dec 2007]. vii DST and exploration ground system operations incurred between 2038 and the end of the orbital mission in 2040. Columns may not sum to the total listed due to rounding.
The International Space Station (ISS) program is developing a plasma contactor to mitigate the ha... more The International Space Station (ISS) program is developing a plasma contactor to mitigate the harmful effects of charge collection on the station's large photovoltaic arrays. The purpose of the present test was to examine the effects of charge collection on the solar array electrical circuit and to verify the effectiveness of the plasma contactor. The results showed that the plasma contactor was able to eliminate structure arcing for any array output voltage. However, the current requirements of the plasma contactor were higher than those for prior testing and predicted by analysis. Three possible causes for this excess current demand are discussed. The most likely appeared to be a high local pressure on or very near the surface of the array as a result of vacuum tank conditions. Therefore, in actual space conditions, the plasma contactor should work as predicted.
Flight experience for Electric Propulsion (EP) supplied by Aerojet since the first launch of Aero... more Flight experience for Electric Propulsion (EP) supplied by Aerojet since the first launch of Aerojet's Electrothermal Hydrazine Thruster in 1983 is reviewed. In total, over 210 spacecraft have flown one of four different EP technologies provided by Aerojet: resistojets, arcjet systems, a pulsed plasma thruster system, and Hall thruster systems. The flight systems include over 100 power processing units of seven different designs and over 500 individual thrusters, as well as propellant management hardware. The development history and basic characteristics of each flight system are reviewed. This survey discusses the application of Aerojet's flight systems in the context of the historical use of electric propulsion flight programs worldwide. Roughly two-thirds of all currently operational spacecraft with EP are flying Aerojet electric propulsion. The usage of total spacecraft flying EP and user type by year is traced. Additionally, trends in the major characteristics of EP bea...
Anode power deposition is a dominant power loss mechanism for arcjets and magnetoplasmadynamic (M... more Anode power deposition is a dominant power loss mechanism for arcjets and magnetoplasmadynamic (MPD) thrusters. In this study, a free burning arc experiment was operated at pressures and current densities similar to those in arcjets and MPD thrusters in an attempt to identify the physics controlling this loss mechanism. Use of a free burning arc allowed for the isolation of independent variables controlling anode power deposition and provided a convenient and flexible way to cover a broad range of currents, anode surface pressures, and applied magnetic field strengths and orientations using an argon gas. Test results showed that anode power deposition decreased with increasing anode surface pressure up to 6.7 Pa and then became insensitive to pressure. Anode power increased with increasing arc current, while the electron number density near the anode surface increased linearly. Anode power also increased with increasing applied magnetic field strength due to an increasing anode fall...
Many high-temperature processes comprise large-scale phenomena. Studying spatial and temporal cor... more Many high-temperature processes comprise large-scale phenomena. Studying spatial and temporal correlations of physical processes between several locations within characteristic scales provides desired information on macroscopic physical processes. Achieved with emission spectroscopy by use of multiple optical fibers. Simultaneous coupling of light from these fibers into single available spectrometer and/or monochromator not accomplished without added expense of two-dimensional array and increased complexity of calibration. Quasi-simultaneous coupling, while maintaining optimum alignment and maximum throughput of broadband emission, achieved by use of fiber optic multiscanner. Instrument used successfully in study of frozen-flow losses internal to flow of plasma inside nozzle of arc jet. Instrument includes two hollow disks of different sizes and stepping motor.
Acta Astronautica, 2019
The use of Electric Propulsion (EP) on satellites for commercial, defense, and space science miss... more The use of Electric Propulsion (EP) on satellites for commercial, defense, and space science missions has been increasing in recent decades, from the first successful operation in 1964 aboard the Zond-2 spacecraft to the present day. This paper provides an overview of the technological and commercial development of EP systems that have been deployed. A review of the early years of EP application ends in 1980, when the first geostationary commercial satellite using EP, Intelsat-V, was launched. Beyond 1980, all EP-based spacecraft deployment data through 2018 are presented, divided by spacecraft type: GEOsynchronous satellite, LEO satellites, deep-space missions and small satellites. To date, a total of 587 spacecraft have been launched with some variant of electric propulsion. During the 1960s and 1970s, all 48 spacecraft using EP were government missions, with the US and USSR leading in the development, production, and flight of these systems. These first platforms included a variety of pulsed plasma thrusters, resistojets, arcjets, ion thrusters and Hall thrusters. The number of GEO satellites with electric propulsion systems has increased significantly since 1981, from an average of less than 5 satellites per year during the 1980s to over 15 in recent years. The corresponding annual fraction of EP based GEO satellite launches, compared to all GEO satellite launches, has increased from 20% during the 1980s to over 40% in recent years. For LEO applications, a gradual increase in the utilization of EP has been realized. Of the 167 EP-based LEO platforms deployed, resistojets were the most prolific legacy thruster type (124 S/C) with Hall thrusters gaining traction in recent years (25 S/C), appearing on 19 of 45 satellite missions in the past decade. Of all EPbased LEO missions, approximately half served as testbeds for new technologies. Through 2018, eight deep space spacecraft with EP have been launched, with the US, Japan, and the European Union leading these efforts. Small satellites are also benefiting from this technology, with 24 EP-based small satellites launched to date. Nearly half of these were launched between 2016 and 2018, demonstrating accelerated growth and a large potential for the future of this spacecraft class. George Washington University Formation Flight Orbit Maintenance Pegasus (QB50 AT03) Academic 2 2017 PPT University of Applied Sciences Wiener Neustadt Tech demo AOBA-VELOX 3 Academic 2 2016 PPT Nanyang University Orbit Maintenance AeroCube 8D Technological 3 2016 Electrospray MIT Tech demo AeroCube 8C Technological 3 2016 Electrospray MIT Tech demo Horyu 4 (AEGIS) Academic 10 2016 VAT Kyushu Institute of Technology (KIT) Attitude Control Orbit Maintenance BRICsat-P Academic 1.9 2015 VAT George Washington University Attitude Control Orbit Change AeroCube 8B Technological 3 2015 Electrospray MIT Tech demo AeroCube 8A Technological 3 2015 Electrospray MIT Tech demo WREN Commercial 0.25 2013 PPT StaDoKo Tech demo CUSat top Academic 25 2013 PPT Cornell University Attitude Control Orbit Maintenance CUSat bottom Academic 25 2013 PPT Cornell University Attitude Control Orbit Maintenance STRaND-1 Academic 3.5 2012 PPT SSTL Attitude Control PROITERES Academic 15 2012 PPT Osaka Institute of Technology (OIT) Tech demo FalconSat-III Military 50 2007 PPT Busek Attitude Control Illinois Observing Nanosatellite (ION) Academic 2 2006 VAT University of Illinois Launch Failure
19th International Electric Propulsion Conference, May 1, 1987
The energy deposition and acceleration mechanisms in the electrothermal-electromagnetic hybrid re... more The energy deposition and acceleration mechanisms in the electrothermal-electromagnetic hybrid regime of coaxial plasma thruster operation are examined both theoretically and experimentally. Theoretical results show that the major trade-offs in the hybrid regime are between efficiency and specific impulse: increasing the influence of electromagnetic forces increases I(sp), but within the operating range examined, decreases the efficiency. Experiments conducted in the predominantly electromagnetic regime agree with the predictions. Anode power deposition is the dominant loss process. 26 references.
25th Joint Propulsion Conference, 1989
Journal of Propulsion and Power, 1994
Anode heat flux measurements of a water cooled segmented anode applied-field MPD thruster were ma... more Anode heat flux measurements of a water cooled segmented anode applied-field MPD thruster were made to investigate anode heat transfer phenomena. Pure argon and argon-hydrogen mixtures were used as propellants for a variety of thruster currents, propellant mass flow rates, and axial applied magnetic field strengths. The thruster was operated in two modes; with all four segments active, and with two of the segments floating. In addition, thrust and specific impulse were determined for each operating condition. The results show that the heat flux to the anode increases monotonically with axial magnetic field strength and thruster current. Between 50 and 75 percent of the anode heat flux is transported by the current carrying electrons. Convective and radiative heat transfer account for the remaining portion of the power deposited in the anode. The addition of hydrogen to the argon propellant results in the reduction of the fraction of anode power deposited by the anode fall to a level equivalent to that deposited by convection and radiation.
Journal of Propulsion and Power, 1991
Mass-loss and surface-temperature measurements, filter photography, and scanning electron microsc... more Mass-loss and surface-temperature measurements, filter photography, and scanning electron microscope surface characterization were used to study cathode phenomena in a steady-state, self-field magnetoplasmadynamic thruster operated at power levels between 15 and 30 kW. The steady-state cathode temperature was found to be above 3000 K over most of its length, and evaporation and simple thermionic emission adequately explained the measured erosion rates and current levels. The steady-state cathode power balance was dominated by electron cooling and radiation. The arc ignition phase was found to last approximately 2 s and was characterized by spot current attachment with associated high erosion rates.
Pulsed plasma thrusters (PFTs) offer the combined benefits of extremely low average electric powe... more Pulsed plasma thrusters (PFTs) offer the combined benefits of extremely low average electric power requirements (1 to 150 W), high specific impulse (- lo00 s), and system simplicity derived from the use of an inert solid propellant. Potential applications range from orbit insertion ...
The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) ... more 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 devel...
The nuclear launch approval process has no formal criteria that define a tolerable level of risk.... more The nuclear launch approval process has no formal criteria that define a tolerable level of risk. In principle, the President or his or her designee, the Director of the Office of Science and Technology Policy (OSTP), has the authority to decide what level of risk is tolerable as dictated by National Security Council Memorandum/Presidential Directive No. 25 (NSC/PD-25) and National Space Policy 2010. In practice, relatively few criteria support an approval decision, and the lack of established guidelines leaves safety reviews unbounded by anything other than budget and the launch window.
Journal of Applied Physics, 2021
Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have b... more Hall effect thrusters operating at power levels in excess of several hundreds of kilowatts have been identified as enabling technologies for applications such as lunar tugs, large satellite orbital transfer vehicles, and solar system exploration. These large thrusters introduce significant testing challenges due to the propellant flow rate exceeding the pumping speed available in most laboratories. Even with proposed upgrades in mind, the likelihood that multiple vacuum facilities will exist in the near future to allow long duration testing of high-power Hall thrusters operating at power levels in excess of 100 kW remains extremely low. In this article, we numerically explore the feasibility of testing Hall thrusters in a quasi-steady mode defined by pulsing the mass flow rate between a nominal and a low value. Our simulations indicate that sub-second durations available before the chamber reaches critical pressure are sufficiently long to achieve the steady-state current and flow f...
53rd AIAA/SAE/ASEE Joint Propulsion Conference, 2017
The next phase of robotic and human deep space exploration missions requires high performance, hi... more The next phase of robotic and human deep space exploration missions requires high performance, high power solar electric propulsion systems for large-scale science missions and cargo transportation. Aerojet Rocketdyne’s Advanced Electric Propulsion System (AEPS) program will complete development and qualification of a 13kW flight EP system to support NASA exploration. The first use of the AEPS is planned for the NASA Power & Propulsion Element, which is the first element of NASA’s cis-lunar Gateway. The flight AEPS system includes a magnetically shielded long-life Hall thruster, power processing unit (PPU), and xenon flow controller (XFC). The Hall thruster, originally developed and demonstrated by NASA’s Glenn Research Center and the Jet Propulsion Laboratory, operates at input powers up to 12.5kW while providing a specific impulse over 2600s at an input voltage of 600V. The power processor is designed to accommodate an input voltage range of 95 to 140V, consistent with operation b...
Large scale cargo transportation to support human missions to the Moon and Mars will require very... more Large scale cargo transportation to support human missions to the Moon and Mars will require very high power Solar Electric Propulsion (SEP) systems operating between 200 and 400 kW. Aerojet Rocketdyne’s NextSTEP program is developing and demonstrating a 100 kW EP system, the XR-100, using a Nested Hall Thruster (NHT) designed for powers up to 200 kW, a modular power processor and a modular flow controller. The three year program objective is to operate the integrated EP system continuously at 100 kW for 100 h, advancing this very high power Electric Propulsion (EP) system to Technology Readiness Level (TRL) 5. With our University of Michigan, Jet Propulsion Laboratory and NASA Glenn Research Center teammates, Aerojet Rocketdyne has completed the initial phase of the program, including operating the thruster at up to 30 kW to validate the thermal models and developing and operating multiple power processor modules in the required series/parallel configuration. The current phase includes completing a TRL 4 integrated system test at reduced power to validate all system operating phases. Design upgrades to demonstrate the TRL 5 capabilities are underway. This paper will present the high power XR-100 capabilities, overall program and design approach and the latest test results for the 100 kW EP system demonstration program. 1 NextSTEP Program Manager, jerry.jackson@rocket.com. 2 NextSTEP Chief Engineer, may.allen@rocket.com. 3 NextSTEP Technical Consultant, roger.myers@rocket.com. 4 NextSTEP Technical Consultant, william.hoskins@rocket.com. 5 NextSTEP PPU Lead, erich.soendker@rocket.com. 6 NextSTEP System and Thruster Lead, benjamin.welander@rocket.com. 7 NextSTEP Feed System Lead, arturo.tolentino@rocket.com. 8 NextSTEP PPU Discharge Supply Lead, sam.hablitzel@rocket.com. 9 University of Michigan, Hall Thruster SME, shall@ umich.edu. 10 University of Michigan, Robert J. Vlassic Dean of Engineering, Richard F. and Eleanor A. Towner Professor, and Arthur F. Thurnau Professor, Department of Aerospace Engineering, and Laboratory Director, PEPL, alec.gallimore@umich.edu. 11 University of Michigan, Assistant Professor and Laboratory Co-Director, PEPL, bjorns@umich.edu. 12 NASA JPL Thruster Lead, richard.r.hofer@jpl.nasa.gov. 13 NASA JPL Hollow Cathode Lead, Dan.M.Goebel@jpl.nasa.gov. 14 NASA NextSTEP technical monitor, eric.pencil@nasa.gov. The 35th International Electric Propulsion Conference, Georgia Institute of Technology, USA October 8 – 12, 2017 Published by the Electric Rocket Propulsion Society with permission https://ntrs.nasa.gov/search.jsp?R=20180004449 2018-11-18T06:06:11+00:00Z
24th Joint Propulsion Conference, 1988
Mass loss and surface temperature measurements, filter photography and SEM surface characterizati... more Mass loss and surface temperature measurements, filter photography and SEM surface characterization are used to study cathode phenomena in a steady state self-field MPD thruster operated at power levels between 15 and 30 kW. The equilibrium cathode temperature is found to be approximately 3200 K over most of its length, for which evaporation and thermionic emission adequately explain measured erosion rates and current levels. The start-up phase is characterized by spot current attachment with associated high erosion rates.
This material may be reproduced by or for the U.S. Government pursuant to the copyright license u... more This material may be reproduced by or for the U.S. Government pursuant to the copyright license under the clause at FAR 52.227-14 [Dec 2007]. vii DST and exploration ground system operations incurred between 2038 and the end of the orbital mission in 2040. Columns may not sum to the total listed due to rounding.
The International Space Station (ISS) program is developing a plasma contactor to mitigate the ha... more The International Space Station (ISS) program is developing a plasma contactor to mitigate the harmful effects of charge collection on the station's large photovoltaic arrays. The purpose of the present test was to examine the effects of charge collection on the solar array electrical circuit and to verify the effectiveness of the plasma contactor. The results showed that the plasma contactor was able to eliminate structure arcing for any array output voltage. However, the current requirements of the plasma contactor were higher than those for prior testing and predicted by analysis. Three possible causes for this excess current demand are discussed. The most likely appeared to be a high local pressure on or very near the surface of the array as a result of vacuum tank conditions. Therefore, in actual space conditions, the plasma contactor should work as predicted.
Flight experience for Electric Propulsion (EP) supplied by Aerojet since the first launch of Aero... more Flight experience for Electric Propulsion (EP) supplied by Aerojet since the first launch of Aerojet's Electrothermal Hydrazine Thruster in 1983 is reviewed. In total, over 210 spacecraft have flown one of four different EP technologies provided by Aerojet: resistojets, arcjet systems, a pulsed plasma thruster system, and Hall thruster systems. The flight systems include over 100 power processing units of seven different designs and over 500 individual thrusters, as well as propellant management hardware. The development history and basic characteristics of each flight system are reviewed. This survey discusses the application of Aerojet's flight systems in the context of the historical use of electric propulsion flight programs worldwide. Roughly two-thirds of all currently operational spacecraft with EP are flying Aerojet electric propulsion. The usage of total spacecraft flying EP and user type by year is traced. Additionally, trends in the major characteristics of EP bea...
Anode power deposition is a dominant power loss mechanism for arcjets and magnetoplasmadynamic (M... more Anode power deposition is a dominant power loss mechanism for arcjets and magnetoplasmadynamic (MPD) thrusters. In this study, a free burning arc experiment was operated at pressures and current densities similar to those in arcjets and MPD thrusters in an attempt to identify the physics controlling this loss mechanism. Use of a free burning arc allowed for the isolation of independent variables controlling anode power deposition and provided a convenient and flexible way to cover a broad range of currents, anode surface pressures, and applied magnetic field strengths and orientations using an argon gas. Test results showed that anode power deposition decreased with increasing anode surface pressure up to 6.7 Pa and then became insensitive to pressure. Anode power increased with increasing arc current, while the electron number density near the anode surface increased linearly. Anode power also increased with increasing applied magnetic field strength due to an increasing anode fall...
Many high-temperature processes comprise large-scale phenomena. Studying spatial and temporal cor... more Many high-temperature processes comprise large-scale phenomena. Studying spatial and temporal correlations of physical processes between several locations within characteristic scales provides desired information on macroscopic physical processes. Achieved with emission spectroscopy by use of multiple optical fibers. Simultaneous coupling of light from these fibers into single available spectrometer and/or monochromator not accomplished without added expense of two-dimensional array and increased complexity of calibration. Quasi-simultaneous coupling, while maintaining optimum alignment and maximum throughput of broadband emission, achieved by use of fiber optic multiscanner. Instrument used successfully in study of frozen-flow losses internal to flow of plasma inside nozzle of arc jet. Instrument includes two hollow disks of different sizes and stepping motor.
Acta Astronautica, 2019
The use of Electric Propulsion (EP) on satellites for commercial, defense, and space science miss... more The use of Electric Propulsion (EP) on satellites for commercial, defense, and space science missions has been increasing in recent decades, from the first successful operation in 1964 aboard the Zond-2 spacecraft to the present day. This paper provides an overview of the technological and commercial development of EP systems that have been deployed. A review of the early years of EP application ends in 1980, when the first geostationary commercial satellite using EP, Intelsat-V, was launched. Beyond 1980, all EP-based spacecraft deployment data through 2018 are presented, divided by spacecraft type: GEOsynchronous satellite, LEO satellites, deep-space missions and small satellites. To date, a total of 587 spacecraft have been launched with some variant of electric propulsion. During the 1960s and 1970s, all 48 spacecraft using EP were government missions, with the US and USSR leading in the development, production, and flight of these systems. These first platforms included a variety of pulsed plasma thrusters, resistojets, arcjets, ion thrusters and Hall thrusters. The number of GEO satellites with electric propulsion systems has increased significantly since 1981, from an average of less than 5 satellites per year during the 1980s to over 15 in recent years. The corresponding annual fraction of EP based GEO satellite launches, compared to all GEO satellite launches, has increased from 20% during the 1980s to over 40% in recent years. For LEO applications, a gradual increase in the utilization of EP has been realized. Of the 167 EP-based LEO platforms deployed, resistojets were the most prolific legacy thruster type (124 S/C) with Hall thrusters gaining traction in recent years (25 S/C), appearing on 19 of 45 satellite missions in the past decade. Of all EPbased LEO missions, approximately half served as testbeds for new technologies. Through 2018, eight deep space spacecraft with EP have been launched, with the US, Japan, and the European Union leading these efforts. Small satellites are also benefiting from this technology, with 24 EP-based small satellites launched to date. Nearly half of these were launched between 2016 and 2018, demonstrating accelerated growth and a large potential for the future of this spacecraft class. George Washington University Formation Flight Orbit Maintenance Pegasus (QB50 AT03) Academic 2 2017 PPT University of Applied Sciences Wiener Neustadt Tech demo AOBA-VELOX 3 Academic 2 2016 PPT Nanyang University Orbit Maintenance AeroCube 8D Technological 3 2016 Electrospray MIT Tech demo AeroCube 8C Technological 3 2016 Electrospray MIT Tech demo Horyu 4 (AEGIS) Academic 10 2016 VAT Kyushu Institute of Technology (KIT) Attitude Control Orbit Maintenance BRICsat-P Academic 1.9 2015 VAT George Washington University Attitude Control Orbit Change AeroCube 8B Technological 3 2015 Electrospray MIT Tech demo AeroCube 8A Technological 3 2015 Electrospray MIT Tech demo WREN Commercial 0.25 2013 PPT StaDoKo Tech demo CUSat top Academic 25 2013 PPT Cornell University Attitude Control Orbit Maintenance CUSat bottom Academic 25 2013 PPT Cornell University Attitude Control Orbit Maintenance STRaND-1 Academic 3.5 2012 PPT SSTL Attitude Control PROITERES Academic 15 2012 PPT Osaka Institute of Technology (OIT) Tech demo FalconSat-III Military 50 2007 PPT Busek Attitude Control Illinois Observing Nanosatellite (ION) Academic 2 2006 VAT University of Illinois Launch Failure
19th International Electric Propulsion Conference, May 1, 1987
The energy deposition and acceleration mechanisms in the electrothermal-electromagnetic hybrid re... more The energy deposition and acceleration mechanisms in the electrothermal-electromagnetic hybrid regime of coaxial plasma thruster operation are examined both theoretically and experimentally. Theoretical results show that the major trade-offs in the hybrid regime are between efficiency and specific impulse: increasing the influence of electromagnetic forces increases I(sp), but within the operating range examined, decreases the efficiency. Experiments conducted in the predominantly electromagnetic regime agree with the predictions. Anode power deposition is the dominant loss process. 26 references.
25th Joint Propulsion Conference, 1989
Journal of Propulsion and Power, 1994
Anode heat flux measurements of a water cooled segmented anode applied-field MPD thruster were ma... more Anode heat flux measurements of a water cooled segmented anode applied-field MPD thruster were made to investigate anode heat transfer phenomena. Pure argon and argon-hydrogen mixtures were used as propellants for a variety of thruster currents, propellant mass flow rates, and axial applied magnetic field strengths. The thruster was operated in two modes; with all four segments active, and with two of the segments floating. In addition, thrust and specific impulse were determined for each operating condition. The results show that the heat flux to the anode increases monotonically with axial magnetic field strength and thruster current. Between 50 and 75 percent of the anode heat flux is transported by the current carrying electrons. Convective and radiative heat transfer account for the remaining portion of the power deposited in the anode. The addition of hydrogen to the argon propellant results in the reduction of the fraction of anode power deposited by the anode fall to a level equivalent to that deposited by convection and radiation.
Journal of Propulsion and Power, 1991
Mass-loss and surface-temperature measurements, filter photography, and scanning electron microsc... more Mass-loss and surface-temperature measurements, filter photography, and scanning electron microscope surface characterization were used to study cathode phenomena in a steady-state, self-field magnetoplasmadynamic thruster operated at power levels between 15 and 30 kW. The steady-state cathode temperature was found to be above 3000 K over most of its length, and evaporation and simple thermionic emission adequately explained the measured erosion rates and current levels. The steady-state cathode power balance was dominated by electron cooling and radiation. The arc ignition phase was found to last approximately 2 s and was characterized by spot current attachment with associated high erosion rates.
Pulsed plasma thrusters (PFTs) offer the combined benefits of extremely low average electric powe... more Pulsed plasma thrusters (PFTs) offer the combined benefits of extremely low average electric power requirements (1 to 150 W), high specific impulse (- lo00 s), and system simplicity derived from the use of an inert solid propellant. Potential applications range from orbit insertion ...
The next phase of space exploration missions requires high power Solar Electric Propulsion (SEP) ... more 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 devel...
The nuclear launch approval process has no formal criteria that define a tolerable level of risk.... more The nuclear launch approval process has no formal criteria that define a tolerable level of risk. In principle, the President or his or her designee, the Director of the Office of Science and Technology Policy (OSTP), has the authority to decide what level of risk is tolerable as dictated by National Security Council Memorandum/Presidential Directive No. 25 (NSC/PD-25) and National Space Policy 2010. In practice, relatively few criteria support an approval decision, and the lack of established guidelines leaves safety reviews unbounded by anything other than budget and the launch window.