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NASA NTRS1977

Electric prototype power processor for a 30cm ion thruster

An electrical prototype power processor unit was designed, fabricated and tested with a 30 cm mercury ion engine for primary space propulsion. The power processor unit used the thyristor series resonant inverter as the basic power stage for the high power beam and discharge supplies. A transistorized series resonant inverter processed the remaining power for the low power outputs. The power processor included a digital interface unit to process all input commands and internal telemetry signals so that electric propulsion systems could be operated with a central computer system. The electrical prototype unit included design improvement in the power components such as thyristors, transistors, filters and resonant capacitors, and power transformers and inductors in order to reduce component weight, to minimize losses, and to control the component temperature rise. A design analysis for the electrical prototype is also presented on the component weight, losses, part count and reliability estimate. The electrical prototype was tested in a thermal vacuum environment. Integration tests were performed with a 30 cm ion engine and demonstrated operational compatibility. Electromagnetic interference data was also recorded on the design to provide information for spacecraft integration.

Biess, J. J.↗

NASA NTRS1992

Beamed microwave power transmission and its application to space

The general principles and special components of beamed microwave power transmission systems are outlined and their application to the space program are discussed. The beamed system is defined as starting with a dc source of power at the transmitting end, converting it to a microwave beam for transmission through space, and ending with the dc power output at the receiving end. An experimentally measured and certified dc-to-dc efficiency of 54 percent has been achieved, using this definition. The application discussed is that of a LEO to GEO transportation system that depends upon vehicles propelled by electric thrusters whose power is supplied by a microwave beam originating at the earth's surface. The advantages of the all-electronic system over a chemically propelled system are enumerated. The principles of space propulsion, particularly as they relate to electric propulsion, are outlined. Key components of the system and environmental considerations are discussed.

Brown, William C.↗

NASA NTRSDate not supplied

NASA Spacecraft Trade Modeling System (NSTRDMS)

A rapid mission analysis tool is developed to support the ongoing design of the Lunar Transit trajectory of the Power and Propulsion Element (PPE). A 50-kW class electric propulsion system is envisioned to transit a massive vehicle be-tween a Medium Earth Orbit (MEO) parking orbit and a lunar L2 southern Near Rectilinear Halo Orbit (NRHO). A parameterization is developed by which the Lunar Transit can be analyzed in the context of varying vehicle mass, solar elec-tric propulsion (SEP) configurations, and solar array power output. A rapid and novel mission analysis tool enables a wide array of these trade analyses to be completed without the need for extensive computing resources or time. This tool is shown to be useful in the analysis of a reference trajectory, where changes to the baseline vehicle architecture or off-nominal operational scenarios (such as electric thruster failures) can be rapidly assessed by the mission designer.

Low thrust↗

NASA NTRS2013

Standard Practices for Usage of Inductive Magnetic Field Probes with Application to Electric Propulsion Testing

Inductive magnetic field probes (also known as B-dot probes and sometimes as B-probes or magnetic probes) are useful for performing measurements in electric space thrusters and various plasma accelerator applications where a time-varying magnetic field is present. Magnetic field probes have proven to be a mainstay in diagnosing plasma thrusters where changes occur rapidly with respect to time, providing the means to measure the magnetic fields produced by time-varying currents and even an indirect measure of the plasma current density through the application of Ampère's law. Examples of applications where this measurement technique has been employed include pulsed plasma thrusters and quasi-steady magnetoplasmadynamic thrusters. The Electric Propulsion Technical Committee (EPTC) of the American Institute of Aeronautics and Astronautics (AIAA) was asked to assemble a Committee on Standards (CoS) for Electric Propulsion Testing. The assembled CoS was tasked with developing Standards and Recommended Practices for various diagnostic techniques used in the evaluation of plasma thrusters. These include measurements that can yield either global information related to a thruster and its performance or detailed, local data related to the specific physical processes occurring in the plasma. This paper presents a summary of the standard, describing the preferred methods for fabrication, calibration, and usage of inductive magnetic field probes for use in diagnosing plasma thrusters. Inductive magnetic field probes (also called B-dot probes throughout this document) are commonly used in electric propulsion (EP) research and testing to measure unsteady magnetic fields produced by time-varying currents. The B-dot probe is relatively simple in construction, and requires minimal cost, making it a low-cost technique that is readily accessible to most researchers. While relatively simple, the design of a B-dot probe is not trivial and there are many opportunities for errors in probe construction, calibration, and usage, and in the post-processing of data that is produced by the probe. There are typically several ways in which each of these steps can be approached, and different applications may require more or less vigorous attention to various issues.

Polzin, Kurt A.↗

NASA NTRS1995

Direct drive options for electric propulsion systems

Power processing units (PPU's) in an electric propulsion system provide many challenging integration issues. The PPU must provide power to the electric thruster while maintaining compatibility with all of the spacecraft power and data systems. Inefficiencies in the power processor produce heat, which must be radiated to the environment in order to ensure reliable operation. Although PPU efficiencies are generally greater than 0.9, heat loads are often substantial. This heat must be rejected by thermal control systems which generally have specific masses of 15-30 kg/kW. PPU's also represent a large fraction of the electric propulsion system dry mass. Simplification or elimination of power processing in a propulsion system would reduce the electric propulsion system specific mass and improve the overall reliability and performance. A direct drive system would eliminate all or some of the power supplies required to operate a thruster by directly connecting the various thruster loads to the solar array. The development of concentrator solar arrays has enabled power bus voltages in excess of 300 V which is high enough for direct drive applications for Hall thrusters such as the Stationary Plasma Thruster (SPT). The option of solar array direct drive for SPT's is explored to provide a comparison between conventional and direct drive system mass.

Hamley, John A.↗

NASA NTRS2023

Overview of the NASA Spacecraft Trade Modeling System (NSTRDMS), A Rapid Mission Analysis Tool

Low thrust↗

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 NTRS2003

The Electric Propulsion Interactions Code (EPIC): A Member of the NASA Space Environment and Effects Program (SEE) Toolset

Science Applications International Corporation is currently developing the Electric Propulsion Interactions Code, EPIC, as part of a project sponsored by the Space Environments and Effects Program at NASA Marshall Space Flight Center. Now in its second year of development, EPIC is an interactive computer toolset that allows the construction of a 3-D spacecraft model, and the assessment of a variety of interactions between its subsystems and the plume from an electric thruster. This paper reports on the progress of EPZC including the recently added ability to exchange results the NASA Charging Analyzer Program, Nascap-2k. The capability greatly enhances EPIC's range of applicability. Expansion of the toolset's various physics models proceeds in parallel with the overall development of the software. Also presented are recent upgrades of the elastic scattering algorithm in the electric propulsion Plume Tool. These upgrades are motivated by the need to assess the effects of elastically scattered ions on the SIC for ion beam energies that exceed loo0 eV. Such energy levels are expected in future high-power (>10 kW) ion propulsion systems empowered by nuclear sources.

Mikellides, Ioannis G.↗

NASA NTRS1979

Sputtering in mercury ion thrusters

Ground-based tests of Hg ion thrusters have identified sputter erosion of thruster components as one of the main life limiting phenomena. Subsequent measurements have revealed that sputtering rates can be affected by background gases at pressures as low as 10 to the -10th torr. With the recent interest in thin film technology, sputtering in the presence of reactive gases has been studied in great detail. This paper presents the results of many of those studies and applies them to the sputtering of electric thrusters. A model, which assumes that chemisorption is the dominant mechanism, is applied to the sputtering rate measurements of the screen grid of a 30-cm thruster in the presence of nitrogen. The model utilizes inputs from a variety of experimental and analytical sources. The model of environmental effects on sputtering was applied to thruster conditions of low discharge voltage and a discussion of the comparison of theory and experiment is presented.

Mantenieks, M. A.↗

NASA NTRS2005

Reusable Hybrid Propellant Modules for Outer-Space Transport

A report summarizes the concept of reusable hybrid propellant modules (HPMs), which would be used in outer space for long-term cryogenic storage of liquefied spacecraft-propellant gases, including for example, oxygen and hydrogen for combustion-based chemical rocket engines and xenon for electric thrusters. The HPM concept would provide the fundamental building block for an efficient, reusable in-space transportation system for both crewed and uncrewed missions. Each HPM would be equipped to implement an advanced zero-boil-off method of managing cryogenic fluids, and would include a fluid-transfer interface comprising standardized fittings that would be compatible with fittings on all supply facilities and on spacecraft to be supplied. The HPM, combined with a chemical or electric orbital transfer spacecraft, would provide an integrated propulsion system. HPMs would supply chemical propellant for time-critical transfers such as crewed missions, and utilize the more efficient electric-propulsion transfer vehicles to transport filled HPMs to the destinations and to return empty HPMs back to near-Earth orbits or other intermediate locations for replenishment and reuse. The HPM prepositioned using electric propulsion would provide the chemical propellant for the crew s return trip in a much more efficient manner than a chemical-only approach. The propellants to fill the HPMs would be delivered from the Earth or other initial supply locations to the intermediate locations by use of automated, compatible spacecraft designed specifically for that purpose. Additionally, multiple HPMs could be aggregated and positioned in orbits and on planets, moons, and asteroids to supply fluids to orbiting and interplanetary spacecraft.

Mazanek, Daniel D.↗

NASA NTRS2019

Non-Intrusive Characterization of the Wear of the HERMeS Thruster Using Optical Emission Spectroscopy

The Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5 kW Hall thruster electric propulsion string that has been in development by NASA Glenn Research Center(GRC) and NASA JPL since 2012. Due to the magnetically shielded design, service life-limiting erosion of the boron nitride discharge has been virtually eliminated. The inner front pole cover (IFPC) has now been identified as the component defining erosion-based service life. Optical emission spectroscopy (OES) is used as an in-situ diagnostic to measure relative erosion trends during operation of the HERMeS thruster during a series of short duration wear tests. Erosion trends obtained from the OES data will be compared to traditional erosion data measured with a non-contact profilometer.

Gray, Timothy G.↗

NASA NTRS2007

Electric Propulsion Interactions Code (EPIC): Recent Enhancements and Goals for Future Capabilities

The Electric Propulsion Interactions Code (EPIC) is the leading interactive computer tool for assessing the effects of electric thruster plumes on spacecraft subsystems. EPIC, developed by SAIC under the sponsorship of the Space Environments and Effects (SEE) Program at the NASA Marshall Space Flight Center, has three primary modules. One is PlumeTool, which calculates plumes of electrostatic thrusters and Hall-effect thrusters by modeling the primary ion beam as well as elastic scattering and charge-exchange of beam ions with thruster-generated neutrals. ObjectToolkit is a 3-D object definition and spacecraft surface modeling tool developed for use with several SEE Program codes. The main EPIC interface integrates the thruster plume into the 3-D geometry of the spacecraft and calculates interactions and effects of the plume with the spacecraft. Effects modeled include erosion of surfaces due to sputtering, re-deposition of sputtered materials, surface heating, torque on the spacecraft, and changes in surface properties due to erosion and deposition. In support of Prometheus I (JIMO), a number of new capabilities and enhancements were made to existing EPIC models. Enhancements to EPIC include adding the ability to scale and view individual plume components, to import a neutral plume associated with a thruster (to model a grid erosion plume, for example), and to calculate the plume from new initial beam conditions. Unfortunately, changes in program direction have left a number of desired enhancements undone. Variable gridding over a surface and resputtering of deposited materials, including multiple bounces and sticking coefficients, would significantly enhance the erosion/deposition model. Other modifications such as improving the heating model and the PlumeTool neutral plume model, enabling time dependent surface interactions, and including EM1 and optical effects would enable EPIC to better serve the aerospace engineer and electric propulsion systems integrator. We review EPIC S overall capabilities and recent modifications, and discuss directions for future enhancements.

Gardner, Barbara M.↗

NASA NTRS2014

5.8kV SiC PiN Diode for Switching of High-Efficiency Inductive Pulsed Plasma Thruster Circuits

Inductive Pulsed Plasma Thruster (IPPT) pulse circuits, such as those needed to operate the Pulsed Inductive Thruster (PIT), are required to quickly switch capacitor banks operating at a period of μs while conducting current at levels on the order of at least 10 kA. [1,2] For all iterations of the PIT to date, spark gaps have been used to discharge the capacitor bank through an inductive coil. Recent availability of fast, high‐power solid state switching devices makes it possible to consider the use of semiconductor switches in modern IPPTs. In addition, novel pre‐ionization schemes have led to a reduction in discharge energy per pulse for electric thrusters of this type, relaxing the switching requirements for these thrusters. [3,4] Solid state switches offer the advantage of greater controllability and reliability, as well as decreased drive circuit dimensions and mass relative to spark gap switches. The use of solid state devices such as Integrated Gate Bipolar Transistors (IGBTs), Gate Turn‐off Thyristors (GTOs) and Silicon‐Controlled Rectifiers (SCRs) often involves the use of power diodes. These semiconductor devices may be connected antiparallel to the switch for protection from reverse current, or used to reduce power loss in a circuit by clamping off current ringing. In each case, higher circuit efficiency may be achieved by using a diode that is able to transition, or 'switch,' from the forward conducting state ('on' state) to the reverse blocking state ('off' state) in the shortest amount of time, thereby minimizing current ringing and switching losses. Silicon Carbide (SiC) PiN diodes offer significant advantages to conventional fast‐switching Silicon (Si) diodes for high power and fast switching applications. A wider band gap results in a breakdown voltage 10 times that of Si, so that a SiC device may have a thinner drift region for a given blocking voltage. [5] This leads to smaller, lighter devices for high voltage applications, as well as reduced forward conduction losses, faster reverse recovery time (faster turn‐off), and lower‐magnitude reverse recovery current. In addition, SiC devices have lower leakage current as compared to their Si counterparts, and a high thermal conductivity, potentially allowing the former to operate at higher temperatures with a smaller, lighter heatsink (or no heatsink at all).

Toftul, Alexandra↗

NASA NTRS2007

Bismuth Propellant Feed System Development at NASA-MSFC

NASA-MSFC has been developing liquid metal propellant feed systems capable of delivering molten bismuth at a prescribed mass flow rate to the vaporizer of an electric thruster. The first such system was delivered to NASA-JPL as part of the Very High Isp Thruster with Anode Layer (VHITAL) program. In this system, the components pictured were placed in a vacuum chamber and heated while the control electronics were located outside the chamber. The system was successfully operated at JPL in conjunction with a propellant vaporizer, and data was obtained demonstrating a new liquid bismuth flow sensing technique developed at MSFC. The present effort is aimed at producing a feed-system for use in conjunction with a bismuth-fed Hall thruster developed by Busek Co. Developing this system is more ambitious, however, in that it is designed to self-contain all the control electronics inside the same vacuum chamber as an operating bismuth-fed thruster. Consequently, the entire system, including an on-board computer, DC-output power supplies, and a gas-pressurization electro-pneumatic regulator, must be designed to survive a vacuum environment and shielded to keep bismuth plasma from intruding on the electronics and causing a shortcircuit. In addition, the hot portions of the feed system must be thermally isolated from the electronics to avoid failure due to high heat loads. This is accomplished using a thermal protection system (TPS) consisting of multiple layers of aluminum foil. The only penetrations into the vacuum chamber are an electrically isolated (floating) 48 VDC line and a fiberoptic line. The 48 VDC provides power for operation of the power supplies and electronics co-located with the system in the vacuum chamber. The fiberoptic Ethernet connection is used to communicate user-input control commands to the on-board computer and transmit real-time data back to the external computer. The partially assembled second-generation system is shown. Before testing at Busek, a more detailed flow sensor calibration will be performed to accurately quantify the flow monitoring capabilities. This effort is funded under a Technology Innovation Program (TIP) award from NASA-MSFC's Technology Transfer office and performed under SAA8-061060.

Polzin, Kurt A.↗

NASA NTRS1991

Dual-purpose self-deliverable lunar surface PV electrical power system

A safe haven and work supported PV power systems on the lunar surface will likely be required by NASA in support of the manned outpost scheduled for the post-2000 lunar/Mars exploration and colonization initiative. Initial system modeling and computer analysis shows that the concept is workable and contains no major high risk technology issues which cannot be resolved in the circa 2000 to 2025 timeframe. A specific selection of the best suited type of electric thruster has not been done; the initial modeling was done using an ion thruster, but Rocketdyne must also evaluate arc and resisto-jets before a final design can be formulated. As a general observation, it appears that such a system can deliver itself to the Moon using many system elements that must be transported as dead payload mass in more conventional delivery modes. It further appears that a larger power system providing a much higher safe haven power level is feasible if this delivery system is implemented, perhaps even sufficient to permit resource prospecting and/or lab experimentation. The concept permits growth and can be expanded to include cargo transport such as habitat and working modules. In short, the combined payload could be manned soon after landing and checkout. NASA has expended substantial resources in the development of electric propulsion concepts and hardware that can be applied to a lunar transport system such as described herein. In short, the paper may represent a viable mission on which previous investments play an invaluable role. A more comprehensive technical paper which embodies second generation analysis and system size will be prepared for near-term presentation.

Arnold, Jack H.↗

NASA NTRS2006

Modular, Reconfigurable, High-Energy Technology Development

The Modular, Reconfigurable High-Energy (MRHE) Technology Demonstrator project was to have been a series of ground-based demonstrations to mature critical technologies needed for in-space assembly of a highpower high-voltage modular spacecraft in low Earth orbit, enabling the development of future modular solar-powered exploration cargo-transport vehicles and infrastructure. MRHE was a project in the High Energy Space Systems (HESS) Program, within NASA's Exploration Systems Research and Technology (ESR&T) Program. NASA participants included Marshall Space Flight Center (MSFC), the Jet Propulsion Laboratory (JPL), and Glenn Research Center (GRC). Contractor participants were the Boeing Phantom Works in Huntsville, AL, Lockheed Martin Advanced Technology Center in Palo Alto, CA, ENTECH, Inc. in Keller, TX, and the University of AL Huntsville (UAH). MRHE's technical objectives were to mature: (a) lightweight, efficient, high-voltage, radiation-resistant solar power generation (SPG) technologies; (b) innovative, lightweight, efficient thermal management systems; (c) efficient, 100kW-class, high-voltage power delivery systems from an SPG to an electric thruster system; (d) autonomous rendezvous and docking technology for in-space assembly of modular, reconfigurable spacecraft; (e) robotic assembly of modular space systems; and (f) modular, reconfigurable distributed avionics technologies. Maturation of these technologies was to be implemented through a series of increasingly-inclusive laboratory demonstrations that would have integrated and demonstrated two systems-of-systems: (a) the autonomous rendezvous and docking of modular spacecraft with deployable structures, robotic assembly, reconfiguration both during assembly and (b) the development and integration of an advanced thermal heat pipe and a high-voltage power delivery system with a representative lightweight high-voltage SPG array. In addition, an integrated simulation testbed would have been developed containing software models representing the technologies being matured in the laboratory demos. The testbed would have also included models for non-MRHE developed subsystems such as electric propulsion, so that end-to-end performance could have been assessed. This paper presents an overview of the MRHE Phase I activities at MSFC and its contractor partners. One of the major Phase I accomplishments is the assembly demonstration in the Lockheed Martin Advanced Technology Center (LMATC) Robot-Satellite facility, in which three robot-satellites successfully demonstrated rendezvous & docking, self-assembly, reconfiguration, adaptable GN&C, deployment, and interfaces between modules. Phase I technology maturation results from ENTECH include material recommendations for radiation hardened Stretched Lens Array (SLA) concentrator lenses, and a design concept and test results for a hi-voltage PV receiver. UAH's accomplishments include Supertube heatpipe test results, which support estimates of thermal conductivities at 30,000 times that of an equivalent silver rod. MSFC performed systems trades and developed a preliminary concept design for a 100kW-class modular reconfigurable solar electric propulsion transport vehicle, and Boeing Phantom Works in Huntsville performed assembly and rendezvous and docking trades. A concept animation video was produced by SAIC, wllich showed rendezvous and docking and SLA-square-rigger deployment in LEO.

Carrington, Connie↗

NASA NTRSDate not supplied

Overview of High Temperature Material Needs for Space Nuclear Propulsion Reactors

Space nuclear propulsion systems are capable of enabling future crewed missions to Mars. Two primary options of these systems exist: nuclear thermal propulsion (NTP) and nuclear electric propulsion (NEP). NTP systems use a reactor as a heat exchanger to heat a hydrogen propellant to produce high specific impulse (≥ 900 s) and thrust (10 - 100 klbf). A NTP materials challenge is development of nuclear fuel capable of operating for short durations (hours) under high power densities (≥ 5 MW/L) and temperatures (> 2700 K) in a corrosive hydrogen environment. In NEP systems, the reactor heats a working fluid which transfers heat to a power conversion system to power electric thrusters. NEP is capable of higher specific impulse (≥ 2000 s) but limited to low thrust (10s lbf). Light weight vessel and in-core structural materials that are radiation-resistant, creep-resistant, and dimensionally stable while operating at ~1500 K over mission times up to several years are needed for NEP. Both systems can benefit from high temperature neutron moderators. Through the Space Nuclear Propulsion project, NASA, supported by the DOE, is developing NTP and NEP systems. Technology maturation planning and risk reduction activities related to fabrication and testing of reactor material candidates is ongoing. This presentation overviews NTP and NEP materials needs and current ceramic and composite development activities.

nuclear thermal propulsion↗

NASA NTRSDate not supplied

Characterizing Electrical Arc Behavior in Ion Engines

Gridded ion engines represent a class of electric thrusters characterized by high propulsive efficiency and high total impulse capabilities. These engines operate by applying high voltage to multi-aperture electrodes, thereby electrostatically accelerating plasma to generate thrust. Within the course of normal operations, periodic occurrences of high voltage breakdowns manifest within the thruster. These breakdowns constitute an inherent aspect of operation, with observations indicating a diminishing frequency and total count over the thruster's operational lifespan. To safeguard the propulsion system during such occurrences, the power processing unit (PPU) employs an automated arc extinction sequence, termed "recycle," which is triggered upon detection of an over-current condition. The initial phase of this sequence involves momentarily deactivating the high voltage and reducing the discharge current to a predetermined level, minimizing plasma impingement as the ion beam is defocused. The high voltage is subsequently reapplied, and the discharge current is returned to its nominal value. The entire sequence typically spans approximately 800 milliseconds. Throughout a recycle event, the propulsion system experiences various current and voltage transients as the arc current traverses the system and different power supplies are toggled on and off.

electric propulsion↗