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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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NASA NTRSDate not supplied

GOES-T Goddard Magnetometer Sensor Unit Flight Thermal Predictions and On-Orbit Thermal Performance

The GOES-T Goddard Magnetometer (GMAG) sensor unit (SU) thermal design was validated by SU level thermal balance (TB) testing. The correlated SU thermal model was integrated to the observatory thermal model to predict the outboard SU temperatures and operating heater duty cycles in the worst cold case and worst hot case orbits. An orbit with the highest solar beta angle (23.5°) and an orbit with the lowest solar beta angle (0°) were run for each case. A comparison in the SU temperature predictions and heater duty cycles were made. The Autumnal Equinox (AE) cold case orbit is slightly colder than the Winter Solstice (WS) cold case orbit. The WS hot case orbit is slightly warmer than the March Equinox (VE) hot case orbit. Additionally, the lunar eclipse cold case orbit was run. It is the coldest environment. Using a 23°C set point for the SU cover heater controller, there is positive heater control for all the orbits. The heater duty cycle predictions for the AE worst cold case orbit and WS worst hot case orbit were 44.4% and 33% respectively. During the lunar eclipse, the heater cycle increases to 68.8% at the end of the eclipse. The temperature predictions for the bobbins are in the 21.9°C to 22.8°C range in the AE worst cold case orbit and in the 21.8°C to 23.0°C range in the WS worst hot case orbit. The SU temperatures during nominal operation on orbit were expected to be within the design goal, despite seasonal variations in solar beta angle and solar irradiance, and no solar flux in a lunar eclipse. The maximum operating SU cover flight heater duty cycle was expected to comply with GOLD Rule 4.25. The SU thermal design is now flight proven based on its long-term (>1 year) thermal performance on orbit.

GOES-T↗
NASA NTRS1991

Solar array orientations for a space station in low earth orbit

A large portion of the drag of a space station in LEO is generated by its solar array; for a baseline 25-kW solar array in 334-km orbit, 1800 kg of reboost propellant/year is needed to counteract solar array drag. A study is conducted of the drag reduction potential of three possible solar array orientations: sun-pointing, sun-pointing during illumination/edge-on during eclipse, and edge-on during entire orbit. An 18.5-percent drag-makeup propellant reduction is found to be obtainable with the sun-pointing/edge-on eclipse orientation technique.

Landis, Geoffrey A.↗
NASA NTRS1992

Laser beamed power: Satellite demonstration applications

It is possible to use a ground-based laser to beam light to the solar arrays of orbiting satellites, to a level sufficient to provide all or some of the operating power required. Near-term applications of this technology for providing supplemental power to existing satellites are discussed. Two missions with significant commercial pay-off are supplementing solar power for radiation-degraded arrays and providing satellite power during eclipse for satellites with failed batteries.

Landis, Geoffrey A.↗
NASA NTRS1993

Improved treatment of GPS force parameters in precise orbit determination applications

Data collected from the International Global Positioning System (GPS) Service (IGS) have been processed at JPL to determine presise orbits for the satellites of the GPS. We have tested a filtering technique to improve modeling of GPS solar radiation pressure force parameters. The new approach improves orbit quality for eclipsing satellites by a factor of two, with typical results in the 25 - 50 cm range.

Vigue, Yvonne↗
NASA NTRS2014

The Orbit of Transneptunian Binary Manwe and Thorondor and Their Upcoming Mutual Events

A new Hubble Space Telescope observation of the 7:4 resonant transneptunian binary system (385446) Manwe has shown that, of two previously reported solutions for the orbit of its satellite Thorondor, the prograde one is correct. The orbit has a period of 110.18 +/- 0.02 days, semimajor axis of 6670 +/- 40 km, and an eccentricity of 0.563 +/- 0.007. It will be viewable edge-on from the inner Solar System during 2015- 2017, presenting opportunities to observe mutual occultation and eclipse events. However, the number of observable events will be small, owing to the long orbital period and expected small sizes of the bodies relative to their separation. This paper presents predictions for events observable from Earth-based telescopes and discusses the associated uncertainties and challenges.

Hubble Space Telescope observations↗
NASA NTRS2021

Cluster after 20 Years of Operations: Science Highlights and Technical Challenges

The Cluster mission was the first constellation using four identical spacecraft to study Sun-Earth connection plasma processes. Using four spacecraft in a tetrahedron shape, it could measure, for the first time, 3D quantities such as electrical currents, plasma gradients or divergence of the electron pressure tensor and 3D structures such as boundaries, surface waves or vortices. Launched in pairs in July and August 2000, on two Soyuz rockets from Baikonur, the four spacecraft have been collecting data continuously for more than 20 years. The mission faced many challenges during the years of operations as some spacecraft subsystems had a lifetime of a few years beyond the initial two-year mission. The major one was to operate without functioning batteries and to successfully pass short and long eclipses, up to 3 h long, without damaging the on-board computers and transmitters and without freezing the fuel. More than 1,000 eclipses have been successfully passed since 2010 using a specially made procedure which switches off the complete spacecraft before entering into eclipse and switches it on when the Sun is again illuminating the solar panels. During 20 years, many discoveries and science results have been published in more than 2,700 scientific papers. A few highlights are presented here, focusing on how varying the spacecraft separation was essential to achieve the science goals of the mission. The Cluster Science Data System and the Cluster archive allows public access to all science data as well as spacecraft ancillary data.

C P Escoubet↗
NASA NTRSDate not supplied

Analysis of the Artemis I Orion Spacecraft Power System Performance

NASA successfully completed an uncrewed test flight of the Orion spacecraft during the 26-day Artemis I mission in November and December 2022. The Artemis I mission profile included several potentially stressing features for Orion electrical power system (EPS) performance, including eclipse duration, multiple propulsive or navigational maneuvers which constrained positioning of the solar array wings (SAWs), and the proximity and phasing of various events together. All of these features vary significantly with Earth-sun-moon geometry, providing a unique challenge for predicting EPS performance before an exact launch date is known. This presentation will provide a brief mission overview, discuss the different computer models with varying levels of fidelity used to analyze Orion EPS performance, as well as the screening process developed to incorporate EPS performance as a consideration for launch epoch selection. Final preflight model predictions of EPS performance will be compared to in-flight telemetry measurements, and several EPS anomalies observed will be briefly discussed.

Orion↗
NASA NTRS1977

Helios-1 Faraday rotation experiment - Results and interpretations of the solar occultations in 1975

The first of two solar occultations of the satellite Helios-1 in 1975 occurred in April when the satellite's ray path approached the west limb of the sun to a minimum distance of 1.63 solar radii. The second occultation took place in late August/early September when Helios-1 was totally eclipsed by the photosphere. Measurements of the polarization angle of the linearly polarized telemetry signal were performed with automatic tracking polarimeters at the 64 m Goldstone Tracking Station in California and also at the 100 m radio telescope in Effelsberg, West Germany. The coronal Faraday rotation as a function of the solar offset for both occultations is shown in graphs. The theoretical significance of the observations is investigated.

Volland, H.↗
NASA NTRSDate not supplied

A Notional Artemis Lunar Surface Exploration Package (ArLSEP) based on the Gandalf Staff Platform

Introduction: The Artemis program is planning to deliver crew and cargo to the lunar surface, but there is no current package for supporting lunar in-struments and experiments similar to the Apollo Lunar Surface Exploration Package (ALSEP). This abstract provides a possible concept for such a package using the Gandalf Staff Platform as a common core. Gandalf Staff: The Gandalf Staff is an early prototype system developed over FY’21/FY’22 using NASA Science Technology Mission Directorate (STMD) Center Information Fund (CIF) grants to de-sign, build and test “proof-of-concept” components. These components include a 24v battery powered monopole that powers a suite of subsystems, including a Graphical User Interface (GUI) for crew, surface voice and data communications, Lunar Search and Rescue (LunaSAR) navigation and communications, LiDAR, field site external lighting, 360-degree camera, and a geothermal instrument for measuring sub-surface temperature gradient. The staff can be carried independently by an Extra-Vehicular Activity (EVA) astronaut, or can be mounted into a tripod for “hands free” support at a surface site being investigated. The staff can be attached to an external solar array and power storage system for long-duration operations. [1,2] ALSEP: An ASLEP flew on each mission Apollo 12 to Apollo 17. For Apollo 11, a simplified packaged called the Early Apollo Scientific Experiments Pack-age (EASEP) was flown. Each package included a “Central Station” that provided the power and communications connected to a variety of instruments and sensors. The power was provided by a Radioisotope Thermoelectric Generator (RTG) fueled by Plutoni-um-238 generating 70 watts of power (initially, decayed over time) [3]. The communications system provide for direct to Earth data transfer from the lunar surface. Each pack-age was stowed externally in the Lunar Module (LM) Scientific Equipment (SEQ) bay with a mass up to 163 kg (Apollo 17). The crew unloaded the ALSEP from the LM and deployed the instruments on the lunar surface. Although designed to operate for only 1 year, many sites operated for up to 8 years successfully [4]. The Active Seismic Experiment (ASE) included 3 geophones for detecting seismic waves created by mortars and thumpers deployed by the crew. Other active experiments measured the lunar atmosphere, the heat flow in the subsurface, the lunar gravity and potential gravity waves, the lunar magnetic field, the solar wind and plasma interactions in cislunar space. Passive experiments included collectors for dust and cosmic rays, and retroreflectors for precise measurements of distance using a laser from Earth. The ALSEP program continues to generate insights into lunar formation and evolution. ArLSEP Concepts: The lunar surface science package for the Artemis program will hopefully exceed the capability of the ALSEP. There are multiple issues for discussion leading to the design of a new ArLSEP, needing requirements definition from the science community, NASA mission architecture, and NASA budget planners. 1. Delivery Mechanism Two possible projects currently provide capability to deliver scientific cargo to the lunar surface: 1) the Commercial Lunar Payload Services (CLPS) [5] and the Human Landing System (HLS) [6, 7]. Each project is controlled by a different organization within NASA and budgeted with different criteria although both support lunar exploration. The HLS system delivers crew (and potentially cargo) to human landing sites. If an ArLSEP is “predeployed” to such a site, the design must include power (either from the vehicle or independently) to keep the electronics functioning until deployed by the crew. If an ArLSEP is delivered on a vehicle after the crew is present on the lunar surface, safety protocols require adequate distance from the humans for impact from descent propelled sur-face regolith ejecta. This distance can not exceed the capability of the crew to walk (if no rover) to the vehicle for ArLSEP deployment. 2. Overall Guidelines The general design of ArLSEP will likely follow the ALSEP with a common system for communications and power; however, significant architecture differences between Apollo and Artemis exist. Power: The RTG will not be available for early Artemis missions nor likely follow-on Lunar Exploration Transportation Services (LETS) missions [8]. Thus, ArLSEP power must be supplied by solar arrays with sufficient battery capability to “keep alive” necessary electronics during any lunar surface eclipse period. Communication: The Artemis program is developing a series of communications satellites for lunar orbit to provide surface transmission of data and voice to Earth. Called “LunaNET”, this network is component useful for ArLSEP since south polar locations may not always have direct “line-of-sight” to Earth [9]. 3. Concept of Operations (ConOps) The general ConOps for ArLSEP is to deliver the package to lunar surface before the crew arrives, and then have the crew deploy the package after some period of time. This requires coordinated design (for power systems) and launch window (for schedule) on both the cargo and crew missions. Once the ArLSEP is deployed, it will operate autonomously for a number of years. It should be designed to be EVA compatible for crew maintenance and upgrade. 4. Notional Design (for discussion purpose only) The landing site near the South Pole is expected to have no eclipse cycle exceeding 5 days, so the “keep alive” power is 144 hours (6 days to include margin). A 12v ArLSEP will use rechargeable LiFePO4 cells, which are common in the Electric Vehicle (EV) industry. With a current of 5 amps and a 125 watt system, the mass is about 90kg. The comm. system and structure adds another 10kg, thus the “Central Station” is approximately 100kg. The solar power is collected on four arrays (each 2m above the surface), and the entire ArLSEP is designed to stow in a 2m x 1m x 1m volume. The experiment and instrument design will vary for each installation and add mass to the total (although they are expected to fit within the 2m3 volume). Seismic wave generation will likely not be provided with mortars, thus an electric “thumper” will be required. Active instruments such as imaging systems and sensing instruments will benefit from the additional power and communication capability provided by ArLSEP. Passive systems such as retroreflectors, witness plates, and cosmic dust collectors can be added to either the landing vehicle and/or the ArLSEP. With repeated HLS missions to the same human site, the ArLSEP can be expanded and easily maintained for long duration science collection on the lunar surface.

ALSEP↗
NASA NTRS1999

Solution for Direct Solar Impingement Problem on Landsat-7 ETM+ Cooler Door During Cooler Outgas in Flight

There was a thermal anomaly of the Landsat-7 Enhanced Thematic Mapper Plus (ETM+) radiative cooler cold stage during the cooler outgas phase in flight. With the cooler door in the outgas position and the outgas heaters enabled, the cold stage temperature increased to a maximum of 323 K when the spacecraft was in the sunlight, which was warmer than the 316.3 K upper set point of the outgas heater controller on the cold stage. Also, the outgas heater cycled off when the cold stage was warming up to 323 K. A corrective action was taken before the attitude of the spacecraft was changed during the first week in flight. One orbit before the attitude was changed, the outgas heaters were disabled to cool off the cold stage. The cold stage temperature increase was strongly dependent on the spacecraft roll and yaw. It provided evidence that direct solar radiation entered the gap between the cooler door and cooler shroud. There was a concern that the direct solar radiation could cause polymerization of hydrocarbons, which could contaminate the cooler and lead to a thermal short. After outgas with the cooler door in the outgas position for seven days, the cooler door was changed to the fully open position. With the cooler door fully open, the maximum cold stage temperature was 316.3 K when the spacecraft was in the sunlight, and the duty cycle of the outgas heater in the eclipse was the same as that in the sunlight. It provided more evidence that direct solar radiation had entered the gap between the cooler door and cooler shroud. Cooler outgas continued for seven more days, with the cooler door fully open. The corrective actions had prevented overheating of the cold stage and cold focal plane array (CFPA), which could damage these two components. They also minimized the risk of contamination on the cold stage, which could lead to a thermal short.

Choi, Michael K.↗
NASA NTRS2010

Discovery of Eclipses from the Accreting Millisecond X-Ray Pulsar Swift J1749.4-2807

We report the discovery of X-ray eclipses in the recently discovered accreting millisecond X-ray pulsar SWIFT J1749.4-2807. This is the first detection of X-ray eclipses in a system of this type and should enable a precise neutron star mass measurement once the companion star is identified and studied. We present a combined pulse and eclipse timing solution that enables tight constraints on the orbital parameters and inclination and shows that the companion mass is in the range 0.6-0.8 solar mass for a likely range of neutron star masses, and that it is larger than a main-sequence star of the same mass. We observed two individual eclipse egresses and a single ingress. Our timing model shows that the eclipse features are symmetric about the time of 90 longitude from the ascending node, as expected. Our eclipse timing solution gives an eclipse duration (from the mid-points of ingress to egress) of 2172+/-13 s. This represents 6.85% of the 8.82 hr orbital period. This system also presents a potential measurement of "Shapiro" delay due to general relativity; through this technique alone, we set an upper limit to the companion mass of 2.2 Solar mass .

Markwardt, C. B.↗
NASA NTRS1986

An IBM PC-based math model for space station solar array simulation

This report discusses and documents the design, development, and verification of a microcomputer-based solar cell math model for simulating the Space Station's solar array Initial Operational Capability (IOC) reference configuration. The array model is developed utilizing a linear solar cell dc math model requiring only five input parameters: short circuit current, open circuit voltage, maximum power voltage, maximum power current, and orbit inclination. The accuracy of this model is investigated using actual solar array on orbit electrical data derived from the Solar Array Flight Experiment/Dynamic Augmentation Experiment (SAFE/DAE), conducted during the STS-41D mission. This simulator provides real-time simulated performance data during the steady state portion of the Space Station orbit (i.e., array fully exposed to sunlight). Eclipse to sunlight transients and shadowing effects are not included in the analysis, but are discussed briefly. Integrating the Solar Array Simulator (SAS) into the Power Management and Distribution (PMAD) subsystem is also discussed.

Emanuel, E. M.↗
NASA NTRS1983

Discovery of a massive unseen star in LMC X-3

Spectroscopic observations of the optical counterpart of LMC X-3 show it to be a spectroscopic binary in the Large Magellanic Cloud with an orbital period of 1.70 days. The B3 main-sequence primary has a large radial velocity amplitude indicating a mass function of 2.3 solar masses. LMC X-3 is an extremely luminous and variable X-ray source, but no 1.7 day X-ray periodicity has been detected. The lack of optical and X-ray eclipses limits the inclination to less than 70 degrees and implies a mass for the unseen star of more than 9.0 solar masses, if the B star has a normal mass. The system now appears to be the strongest evidence for the existence of a stellar mass black hole and is the first extragalactic example.

Remillard, R.↗
NASA NTRS1988

Thermal analysis of heat storage canisters for a solar dynamic, space power system

A thermal analysis was performed of a thermal energy storage canister of a type suggested for use in a solar receiver for an orbiting Brayton cycle power system. Energy storage for the eclipse portion of the cycle is provided by the latent heat of a eutectic mixture of LiF and CaF2 contained in the canister. The chief motivation for the study is the prediction of vapor void effects on temperature profiles and the identification of possible differences between ground test data and projected behavior in microgravity. The first phase of this study is based on a two-dimensional, cylindrical coordinates model using an interim procedure for describing void behavor in 1-g and microgravity. The thermal analysis includes the effects of solidification front behavior, conduction in liquid/solid salt and canister materials, void growth and shrinkage, radiant heat transfer across the void, and convection in the melt due to Marangoni-induced flow and, in 1-g, flow due to density gradients. A number of significant differences between 1-g and o-g behavior were found. This resulted from differences in void location relative to the maximum heat flux and a significantly smaller effective conductance in 0-g due to the absence of gravity-induced convection.

Wichner, R. P.↗
NASA NTRS1991

Design and fabrication of an autonomous rendezvous and docking sensor using off-the-shelf hardware

NASA Marshall Space Flight Center (MSFC) has developed and tested an engineering model of an automated rendezvous and docking sensor system composed of a video camera ringed with laser diodes at two wavelengths and a standard remote manipulator system target that has been modified with retro-reflective tape and 830 and 780 mm optical filters. TRW has provided additional engineering analysis, design, and manufacturing support, resulting in a robust, low cost, automated rendezvous and docking sensor design. We have addressed the issue of space qualification using off-the-shelf hardware components. We have also addressed the performance problems of increased signal to noise ratio, increased range, increased frame rate, graceful degradation through component redundancy, and improved range calibration. Next year, we will build a breadboard of this sensor. The phenomenology of the background scene of a target vehicle as viewed against earth and space backgrounds under various lighting conditions will be simulated using the TRW Dynamic Scene Generator Facility (DSGF). Solar illumination angles of the target vehicle and candidate docking target ranging from eclipse to full sun will be explored. The sensor will be transportable for testing at the MSFC Flight Robotics Laboratory (EB24) using the Dynamic Overhead Telerobotic Simulator (DOTS).

Grimm, Gary E.↗
NASA NTRS1998

Closed Cycle Engine Program Used in Solar Dynamic Power Testing Effort

NASA Lewis Research Center is testing the world's first integrated solar dynamic power system in a simulated space environment. This system converts solar thermal energy into electrical energy by using a closed-cycle gas turbine and alternator. A NASA-developed analysis code called the Closed Cycle Engine Program (CCEP) has been used for both pretest predictions and post-test analysis of system performance. The solar dynamic power system has a reflective concentrator that focuses solar thermal energy into a cavity receiver. The receiver is a heat exchanger that transfers the thermal power to a working fluid, an inert gas mixture of helium and xenon. The receiver also uses a phase-change material to store the thermal energy so that the system can continue producing power when there is no solar input power, such as when an Earth-orbiting satellite is in eclipse. The system uses a recuperated closed Brayton cycle to convert thermal power to mechanical power. Heated gas from the receiver expands through a turbine that turns an alternator and a compressor. The system also includes a gas cooler and a radiator, which reject waste cycle heat, and a recuperator, a gas-to-gas heat exchanger that improves cycle efficiency by recovering thermal energy.

Ensworth, Clint B., III↗
NASA NTRS2020

On Orbit Flight Testing of the Roll-Out Solar Array

The Roll-Out Solar Array (ROSA) is an innovative, lightweight solar array with a flexible substrate that makes use of the stored strain energy in its composite structural members to provide deployment without the use of motors while still packaging efficiently. The ROSA flight experiment was launched to the International Space Station (ISS)on June 3rd, 2017 as a part of the eleventh SpaceX commercial resupply mission. After two weeks of storage, ROSA was extracted from the Dragon spacecraft, deployed, and put through a week of tests to verify its structural characteristics as well as the performance of photovoltaics. The structural dynamics experiments included active excitation of out-of-plane bending modes, and a study of the thermal-structural interaction of the structure during day-night transitions. Data was gathered from a small number of accelerometers located at key points on the solar array and from cameras observing numerous photogrammetry targets spread out over the structure and the photovoltaic blanket. The goal in this effort was to better understand the performance of ROSA and to improve modeling efforts for future designs of similar solar arrays. Of particular interest are the first few system modes and mode shapes of the array, the amount of structural damping present, and degree of structural-thermal interaction seen during eclipse exit.

ROSA↗
NASA NTRS2000

Earth Observing System (EOS) Terra Spacecraft 120 Volt Power Subsystem: Requirements, Development and Implementation

Built by the Lockheed-Martin Corporation, the Earth Observing System (EOS) TERRA spacecraft represents the first orbiting application of a 120 Vdc high voltage spacecraft electrical power system implemented by the National Aeronautics and Space Administration (NASA) Goddard Space Flight Center (GSFC). The EOS TERRA spacecraft's launch provided a major contribution to the NASA Mission to Planet Earth program while incorporating many state of the art electrical power system technologies to achieve its mission goals. The EOS TERRA spacecraft was designed around five state-of-the-art scientific instrument packages designed to monitor key parameters associated with the earth's climate. The development focus of the TERRA electrical power system (EPS) resulted from a need for high power distribution to the EOS TERRA spacecraft subsystems and instruments and minimizing mass and parasitic losses. Also important as a design goal of the EPS was maintaining tight regulation on voltage and achieving low conducted bus noise characteristics. This paper outlines the major requirements for the EPS as well as the resulting hardware implementation approach adopted to meet the demands of the EOS TERRA low earth orbit mission. The selected orbit, based on scientific needs, to achieve the EOS TERRA mission goals is a sun-synchronous circular 98.2degree inclination Low Earth Orbit (LEO) with a near circular average altitude of 705 kilometers. The nominal spacecraft orbit is approximately 99 minutes with an average eclipse period of about 34 minutes. The scientific goal of the selected orbit is to maintain a repeated 10:30 a.m. +/- 15 minute descending equatorial crossing which provides a fairly clear view of the earth's surface and relatively low cloud interference for the instrument observation measurements. The major EOS TERRA EPS design requirements are single fault tolerant, average orbit power delivery of 2, 530 watts with a defined minimum lifetime of five years (EOL). To meet these mission requirements, while minimizing mass and parasitic power losses, the EOS TERRA project relies on 36, 096 high efficiency Gallium Arsenide (GaAs) on Germanium solar cells adhered to a deployable flexible solar array designed to provide over 5,000 watts of power at EOL. To meet the eclipse power demands of the spacecraft, EOS TERRA selected an application of two 54-cell series connected Individual Pressure Vessel (IPV) Nickel-Hydrogen (NiH2) 50 Ampere-Hour batteries. All of the spacecraft observatory electrical power is controlled via the TERRA Power Distribution Unit (PDU) which is designed to provide main bus regulation of 120 Vdc +/- -4% at all load interfaces through the implementation of majority voter control of both the spacecraft's solar array sequential shunt unit (SSU) and the two battery bi-directional charge and discharge regulators. This paper will review the major electrical power system requirement drivers for the EOS TERRA mission as well as some of the challenges encountered during the development, testing, and implementation of the power system. In addition, spacecraft test and early on orbit performance results will also be covered.

Keys, Denney J.↗