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A parallel battery system for the SMM spacecraft is described. The battery system performance as a function of lifetime over orbit was evaluated. The following equipment performance specifications were examined during a typical orbit: battery current and discharges, voltage limitations, battery temperature variations, and current sensor performance. Tabulated battery performance data is also included.
Nine groups of cells were studied. The first group was the control. The basic design was cadmium in the positive plate, and no negative treatment. The plate loading level was the IUE. The separator was pellon. The test regime for the remaining cells was a 90 minute cycle with 40% depth of discharge and a temperature of 20 C. The charge rate was 9.6 amps to a voltage limit. The discharge rate was 9.6 amps.
Three batteries were operated in parallel from a common bus during charge and discharge. SMM utilized NASA Standard 20AH cells and batteries, and LANDSAT-D NASA 50AH cells and batteries of a similar design. Each battery consisted of 22 series connected cells providing the nominal 28V bus. The three batteries were charged in parallel using the voltage limit/current taper mode wherein the voltage limit was temperature compensated. Discharge occurred on the demand of the spacecraft instruments and electronics. Both flights were planned for three to five year missions. The series/parallel configuration of cells and batteries for the 3-5 yr mission required a well controlled product with built-in reliability and uniformity. Examples of how component, cell and battery selection methods affect the uniformity of the series/parallel operation of the batteries both in testing and in flight are given.
The solar cells in a photovoltaic module which must be electrically isolated from module exterior surfaces to satisfy a variety of safety and operating considerations are discussed. The performance and reliability of the insulation system are examined. Technical requirements involve the capability of withstanding the differential voltage from the solar cells to the module frame. The maximum system voltage includes consideration of maximum open circuit array voltages achieved under low-temperature, high-irradiance conditions, and transient overvoltages due to system feedback of lightning transients. The latter is bounded by the characteristics of incorporated voltage limiting devices such as MOVs.
The design and testing of charge control concepts for several power systems is presented. The design of the CRAM charge control system for the SAS satellites and the design of the battery charge regulator for the AMPTE satellites is reviewed. Another charge control concept examined is for the OPEN program and its use of a microprocessor. The use of the voltage limiter is also discussed.
JPL is developing ambient temperature secondary lithium cells for future spacecraft applications. Prior studies on experimental laboratory type Li-TiS2 cells yielded promising results in terms of cycle life and rate capability. To further assess the performance of this cell, 5 Ah engineering model cells were developed. Initially baseline cells were designed and fabricated. Each cell had 15 cathodes and 16 anodes and the ratio of anode to cathode capacity is 6:1. A solution of 1.5 molar LiAsF6 in 2Me-THF was used as the electrolyte. Cells were evaluated for their cycle life at C/1 and C/5 discharge rates and 100 percent depth of discharge. The cells were cycled between voltage limits 1.7 and 2.8 volts. The rate of charge in all cases is C/10. The results obtained indicate that cells can operate at C/10 to C/2 discharge rates and have an initial energy density of 70 Wh/kg. Cells delivered more than 100 cycles at C/2 discharge rate. The details of cell design, the test program, and the results obtained are described.
JPL is developing ambient temperature secondary lithium cells for future spacecraft applications. Prior studies on experimental laboratory type Li-TiS2 cells yielded promising results in terms of cycle life and rate capability. To further assess the performance of this cell, 5 Ah engineering model cells were developed. Initially baseline cells were designed and fabricated. Each cell had 15 cathodes and 16 anodes and the ratio of anode to cathode capacity is 6:1. A solution of 1.5 molar LiAsF6 in 2Me-THF was used as the electrolyte. Cells were evaluated for their cycle life at C/1 and C/5 discharge rates and 100 percent depth of discharge. The cells were cycled between voltage limits 1.7 and 2.8 volts. The rate of charge in all cases is C/10. The results obtained indicate that cells can operate at C/10 to C/2 discharge rates and have an initial energy density of 70 Wh/kg. Cells delivered more than 100 cycles at C/2 discharge rate. The details of cell design, the test program, and the results obtained are described.
JPL is involved in a NASA sponsored program to develop ambient temperature secondary cells for future space missions. After several years of research on various cathode materials, Titanium Disulfide (TiS2) was selected in view of its intrinsic reversibility and high faradaic utilization. In the last two years, the efforts were focused on improving the cycle life of the system and developing 1 Ah cells. Several approaches including the use of mixed solvent electrolytes, the operation of cells at low temperature, and the cycling of cells under different voltage limits, were initially examined to improve the cycle life performance of the Li-TiS2 system. Spiral wound 1 Ah cells fabricated incorporating the improvements from the above studies have delivered more than 600 cycles at 50% DOD. Work is in progress to identify alternate anode materials that can improve the cycle life of the cells to 1000 cycles at 50% DOD. This paper summarizes the advances made in the Li-TiS technology at JPL since 1989.
Considerable controversy has arisen during the recent discussions over a new version of the RTCA DO160C/ED 14C Section 22 document at the European Committee for Aviation Electronics. Section 22 is concerned with lightning waveform tests to equipment. Investigations of some of these controversies with circuit analysis and measurements indicate the impedance characteristics required of the transient generators and the possibility of testing to a voltage limit even for current waveforms.
Most magnetic bearing control schemes use a bias current with a superimposed control current to linearize the relationship between the control current and the force it delivers. For most operating conditions, the existence of the bias current requires more power than alternative methods that do not use conventional bias. Two such methods are examined which diminish or eliminate bias current. In the typical bias control scheme it is found that for a harmonic control force command into a voltage limited transconductance amplifier, the desired force output is obtained only up to certain combinations of force amplitude and frequency. Above these values, the force amplitude is reduced and a phase lag occurs. The power saving alternative control schemes typically exhibit such deficiencies at even lower command frequencies and amplitudes. To assess the severity of these effects, a time history analysis of the force output is performed for the bias method and the alternative methods. Results of the analysis show that the alternative approaches may be viable. The various control methods examined were mathematically modeled using nondimensionalized variables to facilitate comparison of the various methods.
Nickel-Hydrogen cells are being cycled under a LEO (low earth orbit) test regime to examine the benefits of operating the cells at lower States of Charge (SOC) than typically used. A group of four cells are cycled using a voltage limiting charge regime that limits the State of Charge that the cells are allowed to reach. The test cells are then compared to identical cells being cycled at or near 100% State of Charge using a constant current charge regime.
Results of tests characterizing an 8-cm diameter ion source are presented. The tests were conducted in three separate vacuum test facilities at the University of Alabama-Huntsville, Colorado State University, and L3 Communications' ETI division. Standard ion optics tests describing electron backstreaming and total-voltage-limited impingement current behavior as a function of beam current were used as guidelines for selecting operating conditions where more detailed ion beam measurements were performed. The ion beam was profiled using an in-vacuum actuating probe system to determine the total ion current density and the ion charge state distribution variation across the face of the ion source. Both current density and ExB probes were utilized. The ion current density data were used to obtain integrated beam current, beam flatness parameters, and general beam profile shapes. The ExB probe data were used to determine the ratio of doubly to singly charged ion current. The ion beam profile tests were performed at over six different operating points that spanned the expected operating range of the DAWN thrusters being developed at L3. The characterization tests described herein reveal that the 8-cm ion source is suitable for use in (a) validating plasma diagnostic equipment, (b) xenon ion sputtering and etching studies of spacecraft materials, (c) plasma physics research, and (d) the study of ion thruster optics at varying conditions.
The performance testing results of the first prototype model NEXT ion engine, PM1, are presented. The NEXT program has developed the next generation ion propulsion system to enhance and enable Discovery, New Frontiers, and Flagship-type NASA missions. The PM1 thruster exhibits operational behavior consistent with its predecessors, the engineering model thrusters, with substantial mass savings, enhanced thermal margins, and design improvements for environmental testing compliance. The dry mass of PM1 is 12.7 kg. Modifications made in the thruster design have resulted in improved performance and operating margins, as anticipated. PM1 beginning-of-life performance satisfies all of the electric propulsion thruster mission-derived technical requirements. It demonstrates a wide range of throttleability by processing input power levels from 0.5 to 6.9 kW. At 6.9 kW, the PM1 thruster demonstrates specific impulse of 4190 s, 237 mN of thrust, and a thrust efficiency of 0.71. The flat beam profile, flatness parameters vary from 0.66 at low-power to 0.88 at full-power, and advanced ion optics reduce localized accelerator grid erosion and increases margins for electron backstreaming, impingement-limited voltage, and screen grid ion transparency. The thruster throughput capability is predicted to exceed 750 kg of xenon, an equivalent of 36,500 hr of continuous operation at the full-power operating condition.
The paper introduces the midpoint monitoring method for detecting cell voltage imbalance in a series string of 8 electrochemical cells. Monte Carlo analysis is performed for normal cell End of Discharge open and closed circuit voltage behavior to assess the extent to which derived guard band voltage limits will detect over discharged cells, provide false positive indication, or fail to detect over discharged cells. Forward work is proposed to refine the battery monitoring method for operational use in 28V lithium ion batteries.
The United States National Aeronautics and Space Administration (NASA) Near Space Network’s Space Relay (SR) System provides communication relay services to a number of scientific and manned space missions with its Tracking and Data Relay Satellite (TDRS) constellation. NASA’s eighth Tracking and Data Relay Satellite (TDRS) has been experiencing a decline in the health of its power subsystem, and recent reliability analyses indicate that the end of its mission serving the Space Relay is imminent. Launched in 2000, TDRS-8 is the oldest TDRS built upon a Boeing 601 platform, and as such it is the first TDRS to experience these age-related failures. The most prevalent elements of the power subsystem to decline are the solar array circuitry and the Bus Voltage Limiters (BVLs), which prevent too much power from the solar arrays being transferred to the bus by shunting excess current. This means that while the loss of solar array circuits introduces concerns that the solar arrays will continue providing the spacecraft with enough current to remain power positive, the loss of BVLs introduces a concern that the bus may not be adequately protected from overvoltage events. Spacecraft engineers at the TDRS primary ground terminal, the White Sands Complex (WSC) in Las Cruces, New Mexico, and at NASA's Goddard Space Flight Center (GSFC) have also been working in collaboration with Boeing to develop innovative techniques to extend the serviceable life of TDRS-8.
The results of performance tests with two titanium optics sets are presented and compared to those of molybdenum optics. All tests were conducted on a 30 cm ion thruster that was nearly identical to the NASA Solar Electric Propulsion Technology Application Readiness (NSTAR) thruster design. Optics performance tests were conducted over a thruster input power range of 0.5 to 4.6 kW. Optics performance including impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, near-field beam current density profiles, beam divergence angles, and beam divergence thrust correction factors were determined throughout this power range. The impingement-limited total voltages for titanium optics were within 10 to 55 V of those for molybdenum optics. Electron backstreaming limit magnitude as a function of peak beam current density for both molybdenum and titanium optics were within a few volts of each other, indicating similar hot grid gaps for these two grid materials during steady-state operation. Beam divergence half-angles at 90 percent of the total beam current and thrust correction factors for both titanium optics sets were within one degree and one percent, respectively, of those for molybdenum optics. When thruster power was increased to 2.3 kW immediately following discharge ignition, the titanium screen grid came into contact with the accelerator grid within five minutes of ignition. Relative to molybdenum, titanium's larger thermal expansion and smaller thermal conductivity likely caused the screen grid to thermally expand more relative to the accelerator grid during startup.
The results of performance tests with two 40 cm ion optics sets are presented and compared to those of 30 cm ion optics with similar aperture geometries. The 40 cm ion optics utilized both NSTAR and TAG (Thick-Accelerator-Grid) aperture geometries. All 40 cm ion optics tests were conducted on a NEXT (NASA's Evolutionary Xenon Thruster) laboratory model ion engine. Ion optics performance tests were conducted over a beam current range of 1.20 to 3.52 A and an engine input power range of 1.1 to 6.9 kW. Measured ion optics' performance parameters included near-field radial beam current density profiles, impingement-limited total voltages, electron backstreaming limits, screen grid ion transparencies, beam divergence angles, and start-up transients. Impingement-limited total voltages for 40 cm ion optics with the NSTAR aperture geometry were 60 to 90 V lower than those with the TAG aperture geometry. This difference was speculated to be due to an incomplete burn-in of the TAG ion optics. Electron backstreaming limits for the 40 cm ion optics with the TAG aperture geometry were 8 to 19 V higher than those with the NSTAR aperture geometry due to the thicker accelerator grid of the TAG geometry. Because the NEXT ion engine provided beam flatness parameters that were 40 to 63 percent higher than those of the NSTAR ion engine, the 40 cm ion optics outperformed the 30 cm ion optics.
The effect of collector, guard-ring potential imbalance on the observed collector-current-density J, collector-to-emitter voltage V characteristic was evaluated in a planar, fixed-space, guard-ringed thermionic converter. The J,V characteristic was swept in a period of 15 msec by a variable load. A computerized data acquisition system recorded test parameters. The results indicate minimal distortion of the J,V curve in the power output quadrant for the nominal guard-ring circuit configuration. Considerable distortion, along with a lowering of the ignited-mode striking voltage, was observed for the configuration with the emitter shorted to the guard ring. A limited-range performance map of an etched-rhenium, niobium, planar converter was obtained by using an improved computer program for the data acquisition system.