Results for “charging demand” — Page 2

NASA NTRS1986

A double layer model for solar X-ray and microwave pulsations

The wide range of wavelengths over which quasi-periodic pulsations have been observed suggests that the mechanism causing them acts upon the supply of high energy electrons driving the emission processes. A model is described which is based upon the radial shrinkage of a magnetic flux tube. The concentration of the current, along with the reduction in the number of available charge carriers, can rise to a condition where the current demand exceeds the capacity of the thermal electrons. Driven by the large inductance of the external current circuit, an instability takes place in the tube throat, resulting in the formation of a potential double layer, which then accelerates electrons and ions to MeV energies. The double layer can be unstable, collapsing and reforming repeatedly. The resulting pulsed particle beams give rise to pulsating emission which are observed at radio and X-ray wavelengths.

Tapping, K. F.↗

NASA NTRSDate not supplied

NASA Space Communications and Navigation Approach to Cloud Capability Implementation

The integration of cloud services and capabilities across government and commercial sectors has continued to increase over the last several years. Cloud architectures provide on-demand process, storage, and transfer of data as a service without the need to purchase expensive hardware, infrastructure, and their associated maintenance costs. Most cloud environments charge the user for only the resources they use, making cloud implementations highly scalable and easily adjustable to the user’s changing demands.

cloud computing↗

NASA NTRS1981

Advances in space power research and technology at the National Aeronautics and Space Administration

Progress and plans in various areas of the NASA Space Power Program are discussed. Solar cell research is narrowed to GaAs, multibandgap, and thin Si cells for arrays in planar and concentrator configurations, with further work to increase cell efficiency, radiation hardness, develop flexible encapsulants, and reduce cost. Electrochemical research is concentrating on increasing energy and power density, cycle and wet stand life, reliability and cost reduction of batteries. Further development of the Ni-H2 battery and O2-H2 fuel cell to multihundred kW with a 5 year life and 30,000 cycles is noted. Basic research is ongoing for alkali metal anodes for high energy density secondary cells. Nuclear thermoelectric propulsion is being developed for outer planets exploration propulsion systems, using Si-Ge generators, and studies with rare earth chalcogenides and sulfides are mentioned. Power Systems Management seeks to harmonize increasing power supply levels with inner and outer spacecraft environments, circuits, demands, and automatic monitoring. Concomitant development of bipolar transistors, an infrared rectenna, spacecraft charging measurement, and larger heat pipe transport capacity are noted.

Mullin, J. P.↗

NASA NTRS1984

LEC GaAs for integrated circuit applications

Recent developments in liquid encapsulated Czochralski techniques for the growth of semiinsulating GaAs for integrated circuit applications have resulted in significant improvements in the quality and quantity of GaAs material suitable for device processing. The emergence of high performance GaAs integrated circuit technologies has accelerated the demand for high quality, large diameter semiinsulating GaAs substrates. The new device technologies, including digital integrated circuits, monolithic microwave integrated circuits and charge coupled devices have largely adopted direct ion implantation for the formation of doped layers. Ion implantation lends itself to good uniformity and reproducibility, high yield and low cost; however, this technique also places stringent demands on the quality of the semiinsulating GaAs substrates. Although significant progress was made in developing a viable planar ion implantation technology, the variability and poor quality of GaAs substrates have hindered progress in process development.

Kirkpatrick, C. G.↗

NASA NTRS2002

In-situ, On-demand Lubrication System for Space Mechanisms

Many of today's spacecraft have long mission lifetimes. Whatever the lubrication method selected, the initial lubricant charge is required to last the entire mission. Fluid lubricant losses are mainly due to evaporation, tribo-degradation, and oil creep out of the tribological regions. In the past, several techniques were developed to maintain the appropriate amount of oil in the system. They were based on oil reservoirs (cartridges, impregnated porous parts), barrier films, and labyrinth seals. Nevertheless, all these systems have had limited success or have not established a proven record for space missions. The system reported here provides to the ball-race contact fresh lubricant in-situ and on demand. The lubricant is stored in a porous cartridge attached to the inner or the outer ring of a ball bearing. The oil is released by heating the cartridge to eject oil, taking advantage of the greater thermal expansion of the oil compared to the porous network. The heating may be activated by torque increases that signal the depletion of oil in the contact. The low surface tension of the oil compared to the ball bearing material is utilized and the close proximity of the cartridge to the moving balls allows the lubricant to reach the ball-race contacts. This oil resupply system can be used to avoid a mechanism failure or reduce torque to an acceptable level and extend the life of the component.

Marchetti, Mario↗

NASA NTRS2005

Using chargeback to meet customer demand and drive continuous improvement in IT services

Explore the source record for details and available documents.

unified charging system (UCS)↗

NASA NTRS2014

Spacecraft Charging Analysis of a CubeSat

Spacecraft charging occurs when charged particles from the surrounding space plasma environment contact a spacecraft and unequal charging currents result in a net charge density accumulation on or in spacecraft materials. Charging becomes a threat when differential potentials between two points on the spacecraft or between the spacecraft and the ambient space environment build to the level that electric fields associated with the potentials exceed the electric breakdown strength of the spacecraft materials and electrostatic discharge arcs are generated. Electrostatic discharges resulting from spacecraft charging can adversely affect telemetry and cause irreparable damage to electronics. Other spacecraft charging effects include damage of solar arrays and thermal protection, enhancement of contamination of surfaces, and degradation of optics. Typically, the large government and commercial space programs include spacecraft charging analysis as part of the design process. CubeSat projects, however, usually do not have the time or funding to include a spacecraft charging analysis due to their low budget and quick-turnaround requirements. CubeSat projects also tend to rely heavily on commercial "off-the-shelf" products, many of which are not qualified for use in space, and are particularly vulnerable to the effects of the space environment. As the demand for longer and more complex CubeSat missions increases, it is becoming more and more important to consider the effects of spacecraft charging in the design process. Results of surface charging analysis using Nascap-2k on a typical CubeSat design for a polar orbit scenario are illustrated. These results show that for a polar orbiting CubeSat, spacecraft charging could be an issue and steps should be taken to mitigate the effects for these small satellites.

Willis, Emily M.↗

NASA NTRS2023

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Spacecraft↗

NASA NTRSDate not supplied

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Spacecraft↗

NASA NTRSDate not supplied

Development of Two High-Energy Bus ‘Cores’

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Avery Brock↗

NASA NTRS1989

Stratified charge rotary engine combustion studies

Analytical and experimental studies of the combustion process in a stratified charge rotary engine (SCRE) continue to be the subject of active research in recent years. Specifically to meet the demand for more sophisticated products, a detailed understanding of the engine system of interest is warranted. With this in mind the objective of this work is to develop an understanding of the controlling factors that affect the SCRE combustion process so that an efficient power dense rotary engine can be designed. The influence of the induction-exhaust systems and the rotor geometry are believed to have a significant effect on combustion chamber flow characteristics. In this report, emphasis is centered on Laser Doppler Velocimetry (LDV) measurements and on qualitative flow visualizations in the combustion chamber of the motored rotary engine assembly. This will provide a basic understanding of the flow process in the RCE and serve as a data base for verification of numerical simulations. Understanding fuel injection provisions is also important to the successful operation of the stratified charge rotary engine. Toward this end, flow visualizations depicting the development of high speed, high pressure fuel jets are described. Friction is an important consideration in an engine from the standpoint of lost work, durability and reliability. MSU Engine Research Laboratory efforts in accessing the frictional losses associated with the rotary engine are described. This includes work which describes losses in bearing, seal and auxillary components. Finally, a computer controlled mapping system under development is described. This system can be used to map shapes such as combustion chamber, intake manifolds or turbine blades accurately.

Shock, H.↗

NASA NTRS1986

The Voyager 2 encounter with the Uranian system

Discoveries which have occurred due to the success of the Voyager 2 flyby of Uranus and its moons are summarized, along with techniques used to maximize the data collected and minimize the demands on the spacecraft. The 11 spacecraft science packages, including photopolarimeters, spectrometers, a magnetometer, radios and a charged particle detector gathered information on the ring system, the scattering, emitting and reflecting characteristics of the Uranian atmosphere, and the five major satellites. The data revealed 10 minor satellites between Miranda and Uranus and the similarities between the Uranus, Saturn and Jupiter atmospheres. Extensive tabular data are provided for the physical dimensions of the nine rings and the 15 satellites. Finally, data on the magnetosphere and upper atmosphere environment of Uranus and imagery of the major moons are discussed.

Stone, E. C.↗

NASA NTRS2003

Independent Testing of JWST Detector Prototypes

The Independent Detector Testing Laboratory (IDTL) is jointly operated by the Space Telescope Science Institute (STScI) and the Johns Hopkins University (MU), and is assisting the James Webb Space Telescope (JWST) mission in choosing and operating the best near-infrared detectors under a NASA Grant. The JWST is the centerpiece of the NASA Office of Space Science theme, the Astronomical Search for Origins, and the highest priority astronomy project for the next decade, according to the National Academy of Science. JWST will need to have the sensitivity to see the first light in the Universe to determine how galaxies formed in the web of dark matter that existed when the Universe was in its infancy (z approx. 10 - 20). To achieve this goal, the JWST Project must pursue an aggressive technology program and advance infrared detectors to performance levels beyond what is now possible. As part of this program, NASA has selected the IDTL to verify comparative performance between prototype JWST detectors developed by Rockwell Scientific (HgCdTe) and Raytheon (InSb). The IDTL is charged with obtaining an independent assessment of the ability of these two competing technologies to achieve the demanding specifications of the JWST program within the 0.6 - 5 approx. mum bandpass and in an ultra-low background (less than 0.01 e'/s/pixel) environment. We describe results from the JWST Detector Characterization Project that is being performed in the IDTL. In this project, we are measuring first-order detector parameters, i.e. dark current, read noise, QE, intra-pixel sensitivity, linearity, as functions of temperature, well size, and operational mode.

Figer, D. F.↗

NASA NTRS2004

Independent Testing of JWST Detector Prototypes

The Independent Detector Testing Laboratory (IDTL) is jointly operated by the Space Telescope Science Institute (STScI) and the Johns Hopkins University (JHU), and is assisting the James Webb Space Telescope (JWST) mission in choosing and operating the best near-infrared detectors. The JWST is the centerpiece of the NASA Office of Space Science theme, the Astronomical Search for Origins, and the highest priority astronomy project for the next decade, according to the National Academy of Science. JWST will need to have the sensitivity to see the first light in the Universe to determine how galaxies formed in the web of dark matter that existed when the Universe was in its infancy (z is approximately 10-20). To achieve this goal, the JWST Project must pursue an aggressive technology program and advance infrared detectors to performance levels beyond what is now possible. As part of this program, NASA has selected the IDTL to verify comparative performance between prototype JWST detectors developed by Rockwell Scientific (HgCdTe) and Raytheon (InSb). The IDTL is charged with obtaining an independent assessment of the ability of these two competing technologies to achieve the demanding specifications of the JWST program within the 0.6-5 micron bandpass and in an ultra-low background (less than 0.01 e(-)/s/pixel) environment. We describe results from the JWST Detector Characterization Project that is being performed in the LDTL. In this project, we are measuring first-order detector parameters, i.e. dark current, read noise, QE, intra-pixel sensitivity, linearity, as functions of temperature, well size, and operational mode.

Figer, Donald F.↗

NASA NTRS2004

Arc-Second Pointer for Balloon-Borne Astronomical Instrument

A control system has been designed to keep a balloon-borne scientific instrument pointed toward a celestial object within an angular error of the order of an arc second. The design is intended to be adaptable to a large range of instrument payloads. The initial payload to which the design nominally applies is considered to be a telescope, modeled as a simple thin-walled cylinder 24 ft (approx.= 7.3 m) long, 3 ft (approx.= 0.91 m) in diameter, weighing 1,500 lb (having a mass of .680 kg). The instrument would be mounted on a set of motor-driven gimbals in pitch-yaw configuration. The motors on the gimbals would apply the control torques needed for fine adjustments of the instrument in pitch and yaw. The pitch-yaw mount would, in turn, be suspended from a motor mount at the lower end of a pair of cables hanging down from the balloon (see figure). The motor in this mount would be used to effect coarse azimuth control of the pitch-yaw mount. A notable innovation incorporated in the design is a provision for keeping the gimbal bearings in constant motion. This innovation would eliminate the deleterious effects of static friction . something that must be done in order to achieve the desired arc-second precision. Another notable innovation is the use of linear accelerometers to provide feedback that would facilitate the early detection and counteraction of disturbance torques before they could integrate into significant angular-velocity and angular-position errors. The control software processing the sensor data would be capable of distinguishing between translational and rotational accelerations. The output of the accelerometers is combined with that of angular position and angular-velocity sensors into a proportional + integral + derivative + acceleration control law for the pitch and yaw torque motors. Preliminary calculations have shown that with appropriate gains, the power demand of the control system would be low enough to be satisfiable by means of storage batteries charged by solar batteries during the day.

Ward, Philip R.↗

NASA NTRS2015

Verification of a Remaining Flying Time Prediction System for Small Electric Aircraft

This paper addresses the problem of building trust in online predictions of a battery powered aircraft's remaining available flying time. A set of ground tests is described that make use of a small unmanned aerial vehicle to verify the performance of remaining flying time predictions. The algorithm verification procedure described here uses a fully functional vehicle that is restrained to a platform for repeated run-to-functional-failure experiments. The vehicle under test is commanded to follow a predefined propeller RPM profile in order to create battery demand profiles similar to those expected in flight. The fully integrated aircraft is repeatedly operated until the charge stored in powertrain batteries falls below a specified lower-limit. The time at which the lower-limit on battery charge is crossed is then used to measure the accuracy of remaining flying time predictions. Accuracy requirements are considered in this paper for an alarm that warns operators when remaining flying time is estimated to fall below a specified threshold.

Hogge, Edward F.↗

NASA NTRS2010

Human Research Program Space Radiation Standing Review Panel (SRP)

The Space Radiation Standing Review Panel (SRP) met at the NASA Johnson Space Center (JSC) on December 9-11, 2009 to discuss the areas of current and future research targeted by the Space Radiation Program Element (SRPE) of the Human Research Program (HRP). Using evidence-based knowledge as a background for identified risks to astronaut health and performance, NASA had identified gaps in knowledge to address those risks. Ongoing and proposed tasks were presented to address the gaps. The charge to the Space Radiation SRP was to review the gaps, evaluate whether the tasks addressed these gaps and to make recommendations to NASA s HRP Science Management Office regarding the SRP's review. The SRP was requested to evaluate the practicality of the proposed efforts in light of the demands placed on the HRP. Several presentations were made to the SRP during the site visit and the SRP spent sufficient time to address the SRP charge. The SRP made a final debriefing to the HRP Program Scientist, Dr. John B. Charles, on December 11, 2009. The SRP noted that current SRPE strategy is properly science-based and views this as the best assurance of the likelihood that answers to the questions posed as gaps in knowledge can be found, that the uncertainty in risk estimates can be reduced, and that a solid, cost-effective approach to risk reduction solutions is being developed. The current approach of the SRPE, based on the use of carefully focused research solicitations, requiring thorough peer-review and approaches demonstrated to be on the path to answering the NASA strategic questions, addressed to a broad extramural community of qualified scientists, optimally positioned to take advantage of serendipitous discoveries and to leverage scientific advances made elsewhere, is sound and appropriate. The SRP viewed with concern statements by HRP implying that the only science legitimately deserving support should be "applied" or, in some instances that the very term "research" might be frowned upon. We understand the desire of management to ensure that research stay focused on mission objectives, but the terms used are code words fraught with different meaning for scientists. Such expressions, taken at face value, convey a profoundly flawed view of science, can easily lead down counterproductive paths, and have the potential to irretrievably corrupt NASA requirements. The SRP understands and endorses the mandate to keep research efforts focused on the mission needs. However, thoughtful application of knowledge gained by understanding the mechanisms and pathways of biological effects cannot be replaced.

Woloschak, Gayle↗

NASA NTRS1983

The effect of cell design and test criteria on the series/parallel performance of nickel cadmium cells and batteries

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.

Halpert, G.↗