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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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All sourcesNASA NTRSDOE OSTI
NASA NTRS1992

Space reactor power system programs overview

The present development history and current development status evaluation of space reactor power system technologies gives attention to subsystem and component readiness and performance, and assesses the technology data base available in each case. This data base characterization gives attention to the most compatible reactor-power conversion system combinations for prospective DOD and commercial missions, as well as NASA missions. Candidate systems for near, middle, and far term application are selected and prioritized on the basis of technical risk. The programs covered encompass SNAPs 1, 2, 8, and 10A, SNAP 50, and SP-100.

Bloomfield, Harvey S.↗
NASA NTRS2010

NASA Electronic Parts and Packaging (NEPP) Program - Update

This slide presentation reviews the goals and mission of the NASA Electronic Parts and Packaging (NEPP) Program. The NEPP mission is to provide guidance to NASA for the selection and application of microelectronics technologies, to improve understanding of the risks related to the use of these technologies in the space environment and to ensure that appropriate research is performed to meet NASA mission assurance needs. The program has been supporting NASA for over 20 years. The focus is on the reliability aspects of electronic devices. In this work the program also supports the electronics industry. There are several areas that the program is involved in: Memories, systems on a chip (SOCs), data conversion devices, power MOSFETS, power converters, scaled CMOS, capacitors, linear devices, fiber optics, and other electronics such as sensors, cryogenic and SiGe that are used in space systems. Each of these area are reviewed with the work that is being done in reliability and effects of radiation on these technologies.

LaBel, Kenneth A.↗
NASA NTRS1965

Low data rate telemetry.

Noncoherent low power spacecraft communication system with 1 watt transmitted power and omnidirectional spacecraft antenna

SPACECRAFT COMMUNICATION↗
NASA NTRS2024

Validation of Multiphase CFD Predictions of the SLS Launch Environment against Artemis I Flight Data

NASA’s Space Launch System (SLS) is powered by four Aerojet-Rocketdyne RS-25 engines, previously referred to as Space Shuttle Main Engines (SSMEs), with two RSRMV solid rocket boosters (SRBs) derived from the Shuttle booster as seen in Fig. 1. Ignition and the start transient of the SRBs, each of which generate 3.6 million pounds of thrust, create a series of large magnitude pressure waves separate from launch acoustics that have the potential to damage the vehicle if unmitigated. This portion of the SLS launch environment consists of igniter shock (IS), ignition overpressure (IOP), and duct overpressure (DOP).

T. Rivord↗
NASA NTRS2014

Low-Cost, Class D Testing of Spacecraft Photovoltaic Systems Can Reduce Risk

The end-to-end verification of a spacecraft photovoltaic power generation system requires light! Specifically, the standard practice for doing so is the Large Area Pulsed Solar Simulation (LAPSS). A LAPSS test can characterize a photovoltaic system's efficiency via its response to rapidly applied impulses of simulated sunlight. However, a Class D program on a constrained budget and schedule may not have the resources to ship an entire satellite for a LAPSS test alone. Such was the case with the Lunar Atmospheric and Dust Environment Explorer (LADEE) program, which was also averse to the risk of hardware damage during shipment. When the Electrical Power System (EPS) team was denied a spacecraft-level LAPSS test, the lack of an end-to-end power generation test elevated to a project-level technical risk. The team pulled together very limited resources to not only eliminate the risk, but build a process to monitor the health of the system through mission operations. We discuss a process for performing a low-cost, end-to-end test of the LADEE photovoltaic system. The approach combines system-level functional test, panel-level performance results, and periodic inspection (and repair) up until launch. Following launch, mission operations tools are utilized to assess system performance based on a scant amount of data. The process starts in manufacturing at the subcontractor. The panel manufacturer provides functional test and LAPSS data on each individual panel. We apply an initial assumption that the per-panel performance is sufficient to meet the power generation requirements. The manufacturer's data is also carried as the performance allocation for each panel during EPS system modeling and initial mission operations. During integration and test, a high-power, professional theater lamp system provides simulated sunlight to each panel on the spacecraft, thereby permitting a true end-to-end system test. A passing test results in a step response to nearly full-rated current at the appropriate solar array switch in the power system. A metal-halide bulb, infrared imagers, and onboard spacecraft measurements are utilized to minimize risk of thermal damage during test. Data is provided to support test results for both passing and marginal panels. Prior to encapsulation in the launch vehicle, each panel is inspected for damage by the panel manufacturer. Cracked cells or other damage is amended on-site. Because the photovoltaic test system is inexpensive and portable, each repaired panel can be re-verified immediately. Post-launch, the photovoltaic system is again characterized for per-panel deviations from the manufacturer's performance test. This proved especially tricky as the LADEE spacecraft performs only one current measurement on the entire array. The algorithm for Matlab tools to assess panel performance based on spacecraft attitude is discussed. While not as precise and comprehensive as LAPSS, the LADEE approach leverages minimal resources into an ongoing assessment program that can be applied through numerous stages of the mission. The project takes a true Class D approach in assessing the technical value of a spacecraft level performance test versus the programmatic risk of shipping the spacecraft to another facility. The resources required are a fraction of that for a LAPSS test, and is easy to repeat. Further, the test equipment can be handed down to future projects without building an on-site facility.

photovoltaic↗
NASA NTRS1994

Flight experience with lightweight, low-power miniaturized instrumentation systems

Engineers at the NASA Dryden Flight Research Facility (NASA-Dryden) have conducted two flight research programs with lightweight, low-power miniaturized instrumentation systems built around commercial data loggers. One program quantified the performance of a radio-controlled model airplane. The other program was a laminar boundary-layer transition experiment on a manned sailplane. The purpose of this article is to report NASA-Dryden personnel's flight experience with the miniaturized instrumentation systems used on these two programs. This article will describe the data loggers, the sensors, and the hardware and software developed to complete the systems. It also describes how the systems were used and covers the challenges encountered to make them work. Examples of raw data and derived results will be shown as well. For some flight research applications where miniaturized instrumentation is a requirement, the authors conclude that commercially available data loggers and sensors are viable alternatives. In fact, the data loggers and sensors make it possible to gather research-quality data in a timely and cost-effective manner.

LOW POWER MINIATURIZED INSTRUM↗
NASA NTRS1995

Micro Weather Stations for Mars

A global network of weather stations will be needed to characterize the near-surface environment on Mars. Here, we review the scientific and measurement objectives of this network. We also show how these objectives can be met within the cost-constrained Mars Surveyor Program by augmenting the Mars Pathfinder-derived landers with large numbers of very small (less than 5 liter), low-mass (less than 5 kg), low-power, low-cost Mini-meteorological stations. Each station would include instruments for measuring atmospheric. pressures, temperatures, wind velocities, humidity, and airborne dust abundance. They would also include a data handling, telemetry, power, atmospheric entry, and deployment systems in a rugged package capable of direct entry and a high-impact landing. In this paper, we describe these systems and summarize the data-taking strategies and data volumes needed to achieve the surface meteorology objectives for Mars.

Crisp, David↗
NASA NTRS1994

Simulation of Shuttle launch G forces and acoustic loads using the NASA Ames Research Center 20G centrifuge

The high cost and long times required to develop research packages for space flight can often be offset by using ground test techniques. This paper describes a space shuttle launch and reentry simulating using the NASA Ames Research Center's 20G centrifuge facility. The combined G-forces and acoustic environment during shuttle launch and landing were simulated to evaluate the effect on a payload of laboratory rates. The launch G force and acoustic profiles are matched to actual shuttle launch data to produce the required G-forces and acoustic spectrum in the centrifuge test cab where the rats were caged on a free-swinging platform. For reentry, only G force is simulated as the aero-acoustic noise is insignificant compared to that during launch. The shuttle G-force profiles of launch and landing are achieved by programming the centrifuge drive computer to continuously adjust centrifuge rotational speed to obtain the correct launch and landing G forces. The shuttle launch acoustic environment is simulated using a high-power, low-frequency audio system. Accelerometer data from STS-56 and microphone data from STS-1 through STS-5 are used as baselines for the simulations. This paper provides a description of the test setup and the results of the simulation with recommendations for follow-on simulations.

Shaw, T. L.↗
NASA NTRS2005

Health Management Applications for International Space Station

Traditional mission and vehicle management involves teams of highly trained specialists monitoring vehicle status and crew activities, responding rapidly to any anomalies encountered during operations. These teams work from the Mission Control Center and have access to engineering support teams with specialized expertise in International Space Station (ISS) subsystems. Integrated System Health Management (ISHM) applications can significantly augment these capabilities by providing enhanced monitoring, prognostic and diagnostic tools for critical decision support and mission management. The Intelligent Systems Division of NASA Ames Research Center is developing many prototype applications using model-based reasoning, data mining and simulation, working with Mission Control through the ISHM Testbed and Prototypes Project. This paper will briefly describe information technology that supports current mission management practice, and will extend this to a vision for future mission control workflow incorporating new ISHM applications. It will describe ISHM applications currently under development at NASA and will define technical approaches for implementing our vision of future human exploration mission management incorporating artificial intelligence and distributed web service architectures using specific examples. Several prototypes are under development, each highlighting a different computational approach. The ISStrider application allows in-depth analysis of Caution and Warning (C&W) events by correlating real-time telemetry with the logical fault trees used to define off-nominal events. The application uses live telemetry data and the Livingstone diagnostic inference engine to display the specific parameters and fault trees that generated the C&W event, allowing a flight controller to identify the root cause of the event from thousands of possibilities by simply navigating animated fault tree models on their workstation. SimStation models the functional power flow for the ISS Electrical Power System and can predict power balance for nominal and off-nominal conditions. SimStation uses realtime telemetry data to keep detailed computational physics models synchronized with actual ISS power system state. In the event of failure, the application can then rapidly diagnose root cause, predict future resource levels and even correlate technical documents relevant to the specific failure. These advanced computational models will allow better insight and more precise control of ISS subsystems, increasing safety margins by speeding up anomaly resolution and reducing,engineering team effort and cost. This technology will make operating ISS more efficient and is directly applicable to next-generation exploration missions and Crew Exploration Vehicles.

Alena, Richard↗
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 NTRS1986

Data communication requirements for the advanced NAS network

The goal of the Numerical Aerodynamic Simulation (NAS) Program is to provide a powerful computational environment for advanced research and development in aeronautics and related disciplines. The present NAS system consists of a Cray 2 supercomputer connected by a data network to a large mass storage system, to sophisticated local graphics workstations, and by remote communications to researchers throughout the United States. The program plan is to continue acquiring the most powerful supercomputers as they become available. In the 1987/1988 time period it is anticipated that a computer with 4 times the processing speed of a Cray 2 will be obtained and by 1990 an additional supercomputer with 16 times the speed of the Cray 2. The implications of this 20-fold increase in processing power on the data communications requirements are described. The analysis was based on models of the projected workload and system architecture. The results are presented together with the estimates of their sensitivity to assumptions inherent in the models.

Levin, Eugene↗
NASA NTRS1982

The tethered retrievable platform concept and utilization

Main components of the current configuration of a proposed tethered satellite system (TSS) for deploying instrumentation up to 100 km upward or downward from the Orbiter in LEO are described. The TSS consists of the Deployer Module, the Tether, and the Satellite Module, and remains connected to the equipment in the Orbiter bay for the entire mission. The Deployer Module is mounted on a pallet, and has a boom, a tether control mechanism, and a tether control capability, in addition to data acquisition and control systems. The Satellite Module features a thermal control, electrical power and distribution, data management and communication, attitude determination, propulsion, and payload capability. Deployment is unpowered, taking advantage instead of gravity gradients to move the payload from the Orbiter. Safety features include a cutter mechanism to disengage the tether from the Orbiter if retrieval of the payload is judged to endanger the crew and the Orbiter.

Laue, J. H.↗
NASA NTRS2013

Overview of the Habitat Demonstration Unit Power System Integration and Operation at Desert RATS 2010

A habitat demonstration unit (HDU) was constructed at NASA Johnson Space Center (JSC) and designed by a multicenter NASA team led out of NASA Kennedy Space Center (KSC). The HDU was subsequently utilized at the 2010 Desert Research and Technology Studies (RATS) program held at the Black Point Lava Flow in Arizona. This report describes the power system design, installation and operation for the HDU. The requirements for the power system were to provide 120 VAC, 28 VDC, and 120 VDC power to the various loads within the HDU. It also needed to be capable of providing power control and real-time operational data on the load's power consumption. The power system had to be capable of operating off of a 3 phase 480 VAC generator as well as 2 solar photovoltaic (PV) power systems. The system operated well during the 2 week Desert RATS campaign and met all of the main goals of the system. The power system is being further developed to meet the future needs of the HDU and options for this further development are discussed.

Colozza, Anthony J.↗
NASA NTRS2012

Radar Waveform Pulse Analysis Measurement System for High-Power GaN Amplifiers

This work presents a measurement system to characterize the pulsed response of high-power GaN amplifiers for use in space-based SAR platforms that require very strict amplitude and phase stability. The measurement system is able to record and analyze data on three different time scales: fast, slow, and long, which allows for greater detail of the mechanisms that impact amplitude and phase stability. The system is fully automated through MATLAB, which offers both instrument control capability and in-situ data processing. To validate this system, a high-power GaN HEMT amplifier operated in saturation was characterized. The fast time results show that variations to the amplitude and phase are correlated to DC supply transients, while long time characteristics are correlated to temperature changes.

high-power amplifier (HPA)↗
NASA NTRS1980

Spacelab's interface to payload experiments

The Spacelab Payload Standard Modular Electronics (SPSME) program has been designed to provide a standardized set of space-qualified low-power electronic modules from which an experimenter may economically implement a command and data management system. SPSME is based on the Computer Automated Measurement and Control (CAMAC) interface standards. To the system of CAMAC, SPSME adds Spacelab interface modules, a flight qualified system crate, and an efficient light-weight power supply to produce a system compatible with the Spacelab environments, the Spacelab Command and Data Management System interface, and any experiment.

Golemon, W. F.↗
NASA NTRS2013

Application of Spectroscopic Doppler Velocimetry for Measurement of Streamwise Vorticity

A spectroscopic Doppler velocimetry technique has been developed for measuring two transverse components of velocity and hence streamwise vorticity in free jet flows. The nonintrusive optical measurement system uses Mie scattering from a 200 mW green continuous-wave laser interacting with dust and other tracer particulates naturally present in the air flow to measure the velocities. Scattered light is collected in two opposing directions to provide measurements of two orthogonal velocity components. An air-spaced Fabry-Perot interferometer is used for spectral analysis to determine the optical frequency shift between the incident laser light and the Mie scattered light. This frequency shift is directly proportional to the velocity component in the direction of the bisector of the incident and scattered light wave propagation vectors. Data were acquired for jet Mach numbers of 1.73 and 0.99 using a convergent 1.27-cm diameter round nozzle fitted with a single triangular "delta-tab". The velocity components and the streamwise vorticity calculated from the measurements are presented. The results demonstrate the ability of this novel optical system to obtain velocity and vorticity data without any artificial seeding and using a low power laser system.

Fagan, Amy↗