High-Power DC Switchgear Protection for Megawatt Electrified Powertrain Systems
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Nuclear electric propulsion development and qualification facilities
This paper summarizes the findings of a Tri-Agency panel consisting of members from the National Aeronautics and Space Administration (NASA), U.S. Department of Energy (DOE), and U.S. Department of Defense (DOD) that were charged with reviewing the status and availability of facilities to test components and subsystems for megawatt-class nuclear electric propulsion (NEP) systems. The facilities required to support development of NEP are available in NASA centers, DOE laboratories, and industry. However, several key facilities require significant and near-term modification in order to perform the testing required to meet a 2014 launch date. For the higher powered Mars cargo and piloted missions, the priority established for facility preparation is: (1) a thruster developmental testing facility, (2) a thruster lifetime testing facility, (3) a dynamic energy conversion development and demonstration facility, and (4) an advanced reactor testing facility (if required to demonstrate an advanced multiwatt power system). Facilities to support development of the power conditioning and heat rejection subsystems are available in industry, federal laboratories, and universities. In addition to the development facilities, a new preflight qualifications and acceptance testing facility will be required to support the deployment of NEP systems for precursor, cargo, or piloted Mars missions. Because the deployment strategy for NEP involves early demonstration missions, the demonstration of the SP-100 power system is needed by the early 2000's.
Dutt, D. S.↗
High-power converters for space applications
Phase 1 was a concept definition effort to extend space-type dc/dc converter technology to the megawatt level with a weight of less than 0.1 kg/kW (220 lb./MW). Two system designs were evaluated in Phase 1. Each design operates from a 5 kV stacked fuel cell source and provides a voltage step-up to 100 kV at 10 A for charging capacitors (100 pps at a duty cycle of 17 min on, 17 min off). Both designs use an MCT-based, full-bridge inverter, gaseous hydrogen cooling, and crowbar fault protection. The GE-CRD system uses an advanced high-voltage transformer/rectifier filter is series with a resonant tank circuit, driven by an inverter operating at 20 to 50 kHz. Output voltage is controlled through frequency and phase shift control. Fast transient response and stability is ensured via optimal control. Super-resonant operation employing MCTs provides the advantages of lossless snubbing, no turn-on switching loss, use of medium-speed diodes, and intrinsic current limiting under load-fault conditions. Estimated weight of the GE-CRD system is 88 kg (1.5 cu ft.). Efficiency of 94.4 percent and total system loss is 55.711 kW operating at 1 MW load power. The Maxwell system is based on a resonance transformer approach using a cascade of five LC resonant sections at 100 kHz. The 5 kV bus is converted to a square wave, stepped-up to a 100 kV sine wave by the LC sections, rectified, and filtered. Output voltage is controlled with a special series regulator circuit. Estimated weight of the Maxwell system is 83.8 kg (4.0 cu ft.). Efficiency is 87.2 percent and total system loss is 146.411 kW operating at 1 MW load power.
Park, J. N.↗
Megawatt level electric propulsion perspectives
For long range space missions, deliverable payload fraction is an inverse exponential function of the propellant exhaust velocity or specific impulse of the propulsion system. The exhaust velocity of chemical systems are limited by their combustion chemistry and heat transfer to a few km/s. Nuclear rockets may achieve double this range, but are still heat transfer limited and ponderous to develop. Various electric propulsion systems can achieve exhaust velocities in the 10 km/s range, at considerably lower thrust densities, but require an external electrical power source. A general overview is provided of the currently available electric propulsion systems from the perspective of their characteristics as a terminal load for space nuclear systems. A summary of the available electric propulsion options is shown and generally characterized in the power vs. exhaust velocity plot. There are 3 general classes of electric thruster devices: neutral gas heaters, plasma devices, and space charge limited electrostatic or ion thrusters.
Jahn, Robert G.↗
Inert gas thrusters
Inert gas thrusters considered for space propulsion systems were investigated. Electron diffusion across a magnetic field was examined utilizing a basic model. The production of doubly charged ions was correlated using only overall performance parameters. The use of this correlation is therefore possible in the design stage of large gas thrusters, where detailed plasma properties are not available. Argon hollow cathode performance was investigated over a range of emission currents, with the positions of the inert, keeper, and anode varied. A general trend observed was that the maximum ratio of emission to flow rate increased at higher propellant flow rates. It was also found that an enclosed keeper enhances maximum cathode emission at high flow rates. The maximum cathode emission at a given flow rate was associated with a noisy high voltage mode. Although this mode has some similarities to the plume mode found at low flows and emissions, it is encountered by being initially in the spot mode and increasing emission. A detailed analysis of large, inert-gas thruster performance was carried out. For maximum thruster efficiency, the optimum beam diameter increases from less than a meter at under 2000 sec specific impulse to several meters at 10,000 sec. The corresponding range in input power ranges from several kilowatts to megawatts.
Kaufman, H. R.↗
Hybrid Power Management
An engineering discipline denoted as hybrid power management (HPM) has emerged from continuing efforts to increase energy efficiency and reliability of hybrid power systems. HPM is oriented toward integration of diverse electric energy-generating, energy-storing, and energy-consuming devices in optimal configurations for both terrestrial and outer-space applications. The basic concepts of HPM are potentially applicable at power levels ranging from nanowatts to megawatts. Potential applications include terrestrial power-generation, terrestrial transportation, biotechnology, and outer-space power systems. Instances of this discipline at prior stages of development were reported (though not explicitly labeled as HPM) in three prior NASA Tech Briefs articles: "Ultracapacitors Store Energy in a Hybrid Electric Vehicle"(LEW-16876), Vol. 24, No. 4 (April 2000), page 63; "Photovoltaic Power Station With Ultracapacitors for Storage" (LEW-17177), Vol. 27, No. 8 (August 2003), page 38; and "Flasher Powered by Photovoltaic Cells and Ultracapacitors" (LEW-17246), Vol. 24, No. 10 (October 2003), page 37. As the titles of the cited articles indicate, the use of ultracapacitors as energy-storage devices lies at the heart of HPM. An ultracapacitor is an electrochemical energy-storage device, but unlike in a conventional rechargeable electrochemical cell or battery, chemical reactions do not take place during operation. Instead, energy is stored electrostatically at an electrode/electrolyte interface. The capacitance per unit volume of an ultracapacitor is much greater than that of a conventional capacitor because its electrodes have much greater surface area per unit volume and the separation between the electrodes is much smaller. Power-control circuits for ultracapacitors can be simpler than those for batteries, for two reasons: (1) Because of the absence of chemical reactions, charge and discharge currents can be greater than those in batteries, limited only by the electrical resistances of conductors; and (2) whereas the charge level of a battery depends on voltage, temperature, age, and load condition, the charge level of an ultracapacitor, like that of a conventional capacitor, depends only on voltage.
Eichenberg, Dennis↗