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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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A battery charge/state of charge indicator (BC/SCI) system for electric vehicle use was developed. The original and subsequent objectives for the BC/SCI and the rationale for those objectives are described. The requirements generated from the objectives are listed and a description of the BC/SCI is provided. The power section problem, the tests, and the test results are discussed.
Part of an AFGL payload flown on the STS-4 mission consisted of experiments to measure in-situ electric fields, electron densities, and vehicle charging. During this flight some 11 hours of data were acquired ranging from 5 minute snapshots up to continuous half-orbits. These experiments are described and results presented for such vehicle induced events as a main engine burn, thruster firings and water dumps in addition to undisturbed periods. The main characteristic of all the vehicle induced events is shown to be an enhancement in the low frequency noise (less than 2 kHz), in both the electrostatic and electron irregularity (delta N/N) spectra. The non-event results indicate that the electrostatic broadband emissions show a white noise characteristic in the low frequency range up to 2 kHz at an amplitude of 10 db above the shuttle design specification limit, falling below that limit above 10 kHz. The vehicle potential remained within the range of -3 to +1 volt throughout the flight which exhibits normal behavior for a satellite in a low equatorial orbit.
Predicting whether or not vehicle batteries contain sufficient charge to support operations over the remainder of a given flight plan is critical for electric aircraft. This paper describes an approach for identifying upper and lower uncertainty bounds on predictions that aircraft batteries will continue to meet output power and voltage requirements over the remainder of a flight plan. Battery discharge prediction is considered here in terms of the following components; (i) online battery state of charge estimation; (ii) prediction of future battery power demand as a function of an aircraft flight plan; (iii) online estimation of additional parasitic battery loads; and finally, (iv) estimation of flight plan safety. Substantial uncertainty is considered to be an irremovable part of the battery discharge prediction problem. However, high-confidence estimates of flight plan safety or lack of safety are shown to be generated from even highly uncertain prognostic predictions.
Lead-acid and nickel-zinc traction cells were rapidly and efficiently charged using a high rate tapered direct current (HRTDC) charge method which could possibly be used for on-the-road service recharge of electric vehicles. The HRTDC method takes advantage of initial high cell charge acceptance and uses cell gassing rate and temperature as an indicator of charging efficiency. On the average, in these preliminary tests, 300 amp-hour nickel-zinc traction cells were given a HRTDC (initial current 500 amps, final current 100 amps) to 78 percent of rated amp-hour capacity within 53 minutes at an amp-hour efficiency of 92 percent and an energy efficiency of 52 percent. Three hundred amp-hour lead-acid traction cells were charged to 69 percent of rated amp-hour capacity within 46 minutes at an amp-hour efficiency of 91 percent with an energy efficiency of 64 percent. In order to find ways to further decrease the recharge times, the effect of periodically (0 to 400 Hz) pulse discharging cells during a constant current charging process (94% duty cycle) was investigated. Preliminary data indicate no significant effect of this type of pulse discharging during charge on charge acceptance of lead-acid or nickel-zinc cells.
Lead-acid and nickel-zinc traction cells were rapidly and efficiently charged using a high rate taped dc charge (HRTDC) method which could possibly be used for on-the-road service recharge of electric vehicles. The HRTDC method takes advantage of initial high cell charge acceptance and uses cell gassing rate and temperature as an indicator of charging efficiency. On the average, 300 amp-hour nickel-zinc traction cells were given a HRTDC to 78% of rated amp-hour capacity within 53 minutes at an amp-hour efficiency of 92% and an energy efficiency of 52%. Three-hundred amp-hour lead-acid traction cells were charged to 69% of rated amp-hour capacity within 46 minutes at an amp-hour efficiency of 91% with an energy efficiency of 64%.
An innovative modification has been made to a previously patented design for the Phase Change Material (PCM) Thermal Generator, which works in water where ocean temperature alternatively melts wax in canisters, or allows the wax to re-solidify, causing high-pressure oil to flow through a hydraulic generator, thus creating electricity to charge a battery that powers the vehicle. In this modification, a similar thermal PCM device has been created that is heated and cooled by the air and solar radiation instead of using ocean temperature differences to change the PCM from solid to liquid. This innovation allows the device to use thermal energy to generate electricity on land, instead of just in the ocean.
The Vehicle Charging and Potential Experiment formed part of the scientific experiments comprising the NASA Office of Space Sciences-1 pallet flown at an altitude of 250 km and an orbital inclination of 40 deg on the third flight of the Space Shuttle Orbiter in March 1982. Details of the objectives, the instrumentation of the experiment, and some preliminary results from the individual instruments are presented. The experiment studied passive vehicle charging using thermal plasma probes, and it is concluded that apart from moderate v x B (L)-induced electrical potential resulting from the large size and restricted plasma contact area, the Space Shuttle Orbiter behaved similarly to smaller spacecraft orbiting in the same ionospheric region. However, an unexpected passive orbital result was the detection of enhanced plasma density in the vicinity of the Orbiter under restricted orbital conditions, apparently composed of contaminant molecular species and with an electron component having a temperature considerably higher than the ambient ionosphere electrons. The active vehicle charging experiments were performed by using a 100 mA/1 keV electron beam emitted out of the payload bay. Charging measurements made during short electron beam pulse emissions provide an upper limit of about 1 mF for the Orbiter capacitance. Steady-state charging measurements showed that for the vehicle attitudes covered by the mission, the induced electrical potential was normally only a few volts, except during some nighttime conditions when it rose above 40 V.
The International Space Station (ISS) maintains an altitude typically between 300 km and 400 km in low Earth orbit (LEO) which itself is situated in the Earth's ionosphere. The ionosphere is a region of partially ionized gas (plasma) formed by the photoionization of neutral atoms and molecules in the upper atmosphere of Earth. It is important to understand what electron density the spacecraft is/will be operating in because the ionized gas along the ISS orbit interacts with the electrical power system resulting in charging of the vehicle. One instrument that is already operational onboard the ISS with a goal of monitoring electron density, electron temperature, and ISS floating potential is the Floating Potential Measurement Unit (FPMU). Although this tool is a valuable addition to the ISS, there are limitations concerning the data collection periods. The FPMU uses the Ku band communication frequency to transmit data from orbit. Use of this band for FPMU data runs is often terminated due to necessary observation of higher priority Extravehicular Activities (EVAs) and other operations on ISS. Thus, large gaps are present in FPMU data. The purpose of this study is to solve the issue of missing environmental data by implementing a secondary electron density data source, derived from the COSMIC satellite constellation, to create a model of ISS orbital environments. Extrapolating data specific to ISS orbital altitudes, we model the ionospheric electron density along the ISS orbit track to supply a set of data when the FPMU is unavailable. This computer model also provides an additional new source of electron density data that is used to confirm FPMU is operating correctly and supplements the original environmental data taken by FPMU.
This introductory paper describes several concepts that could be used for augmenting the energy state of electric Vertical Take-Off and Landing (eVTOL) vehicles. Advanced Air Mobility (AAM) electric vehicles, just like conventional vehicles, could need additional charge due to depleted batteries (e.g., strong winds along the way) while approaching their destination. There are three indirect charging and five direct charging concepts presented in this paper. The concepts are in preliminary research stage and are being refined. Considering the concepts are for the year 2045 timeframe, there is sufficient time to evolve them, along with the designs of the AAM vehicles. The paper describes more details and discussion on the desirability, viability, and feasibility of these energy augmentation concepts. A discussion of barriers and initial investigation approach for three concepts is presented.
This introductory paper describes several concepts that could be used for augmenting the energy state of electric Vertical Take-Off and Landing (eVTOL) vehicles. Advanced Air Mobility (AAM) electric vehicles, just like conventional vehicles, could need additional charge due to depleted batteries (e.g., strong winds along the way) while approaching their destination. There are three indirect charging and five direct charging concepts presented in this paper. The concepts are in preliminary research stage and are being refined. Considering the concepts are for the year 2045 timeframe, there is sufficient time to evolve them, along with the designs of the AAM vehicles. The paper describes more details and discussion on the desirability, viability, and feasibility of these energy augmentation concepts. A discussion of barriers and initial investigation approach for three concepts is presented.
Gas-emmission and time-integrated-current measurements indicate battery charge state. Tests indicate possibility of monitoring state of charge of lead/acid batteries at any stage in charging cycle by measuring charging current and either gas evolution or electrode potential. Data then processed by microcomputer. Uses include cell voltage, cell pressure, cell temperature and rate of gas recombination on catalyst.
Characterizing the electromagnetic interaction of a satellite in low Earth, high inclination orbit with the space plasma environment and identifying viable charging mitigation strategies is a critical mission design task. High inclination orbits expose the vehicle to auroral charging environments that can potentially charge surfaces to kilovolt potentials and electric thruster propulsion systems will interact with the ambient plasma environment throughout the orbit. NASA is designing the Iodine Satellite (iSAT) cubesat mission to demonstrate operations of an iodine electric thruster system. The spacecraft will be deployed as a secondary payload from a launch vehicle which has not yet been identified so the program must plan for the worst case environments over a range of orbital inclinations. We will first present results from a NASA and Air Force Charging Analyzer Program (Nascap) -2k surface charging calculation used to evaluate the effects of auroral charging on the spacecraft and to provide the charging levels at other locations in orbit for a thruster plume interaction analysis for the iSAT mission. We will then discuss results from the thruster interactions analysis using the Electric Propulsion Interactions Code (EPIC) with inputs from Nascap-2k. The results of these analyses are being used by the iSAT program to better understand how their spacecraft will interact with the space plasma environment in the range of environments that could be encountered when the final mission orbit is selected.
In 1999, the NASA/Human Exploration and Development of Space (HEDS) enterprise selected a number of payloads to fly to the Martian surface in an 03 opportunity (prior to the MPL loss). Part of a proposed experiment, ECHOS, was selected to specifically understand the electrical charging hazards from tribocharged dust in the ambient atmosphere, in dust devils, and in larger storms. It is expected that Martian dust storms become tribocharged much like terrestrial dust devils which can possess almost a million elementary charges per cubic centimeter. The ECHOS package features a set of instruments for measuring electric effects: a radio to detect AC electric fields radiating from discharges in the storm,a DC electric field system for sensing electrostatic fields from concentrations of charged dust grains, and a lander electrometer chain for determining the induced potential on its body and MAV (Mars Ascent Vehicle) during the passages of a charged dust storm. Given that electricity is a systemic process originating from wind-blown dust, we also proposed to correlate the electrical measurements with fundamental fluid/meteorological observations, including wind velocity and vorticity, temperature, and pressure. Triboelectricity will also affect local chemistry, and chemical-sensing devices were also considered a feature of the package. The primary HEDS objectives of the ECHOS sensing suite is to discover and monitor the natural electrical hazards associated with dust devils and storms, and determine their enviro-effectiveness on human systems. However, ECHOS also has a strong footprint in the overarching science objectives of the Mars Surveyor Program.
The Lunar Vehicle Active Charge Control System (LVACCS) is an instrument in development that will measure the electrical potential of a spacecraft or rover under both positive and negative charge conditions, actively discharge negative charge, and charge the vehicle to a safe positive potential. All surfaces immersed in a plasma charge to non-zero potentials. In very low-density plasma (< 100 cm^-3), space vehicle charging can interfere with plasma probe measurements, degrade avionics performance, or cause a total system failure. LVACCS provides real-time in-situ space situational awareness that overcome these challenges in a small package. LVACCS has been awarded a NASA PICASSO grant to bring the instrument to technology readiness level 5. LVACCS consists of two main components: a collimated photoelectron gun (CPEG) and a Spacecraft Charge Detector (SCD).Monitoring charge build up on the lunar surface isa critical first step to understand hazards for crew, rovers, and instrumentation operating there.The LVACCS instrument will be able to fill in this gap providing robust lunar surface potential measurements and active charge dissipation.
A model of the satellite charging at high altitudes (SCATHA P78-2) satellite was used to simulate the charging response of SCATHA at geosynchronous orbit. The model includes a description of the geometry, currents to exposed surface materials, and electrical connections on the spacecraft. The charging response of the vehicle to that predicted by the NASCAP model for the Day 87, 1979 eclipse charging event, in which the spacecraft charged to several kilovolts negative during a magnetospheric substorm are compared. Double Maxwellian representations of the plasma environment reproduce the charging response observed experimentally.
Software-in-the-loop and Hardware-in-the-loop testing of failure prognostics and decision making tools for aircraft systems will facilitate more comprehensive and cost-effective testing than what is practical to conduct with flight tests. A framework is described for the offline recreation of dynamic loads on simulated or physical aircraft powertrain components based on a real-time simulation of airframe dynamics running on a flight simulator, an inner-loop flight control policy executed by either an autopilot routine or a human pilot, and a supervisory fault management control policy. The creation of an offline framework for verifying and validating supervisory failure prognostics and decision making routines is described for the example of battery charge depletion failure scenarios onboard a prototype electric unmanned aerial vehicle.
An electric passenger vehicle was tested to develop data characterizing the state of the art of electric and hybrid vehicles. The test measured vehicle maximum speed, range at constant speed, range over stop-and-go driving schedules, maximum acceleration, gradeability and limit, road energy consumption, road power, indicated energy consumption, braking capability and battery charge efficiency. The data obtained are to serve as a baseline to compare improvements in electric and hybrid vehicle technologies and to assist in establishing performance standards.
Tests are performed on an improved battery charger. The primary purpose of the testing is to develop test methodologies for battery charger evaluation. Tests are developed to characterize the charger in terms of its charge algorithm and to assess the effects of battery initial state of charge and temperature on charger and battery efficiency. Tests show this charger to be a considerable improvement in the state of the art for electric vehicle chargers.