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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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The present spacecraft solar array simulator addresses both dc and ac characteristics as well as changes in illumination and temperature and performance degradation over the course of array service life. The computerized control system used allows simulation of a complete orbit cycle, in addition to automated diagnostics. The simulator is currently interfaced with the NASA EOS testbed.
The Mir Cooperative Solar Array (MCSA) was developed jointly by the United States (US) and Russia to provide approximately 6 kW of photovoltaic power to the Russian space station Mir. The MCSA was launched to Mir in November 1995 and installed on the Kvant-1 module in May 1996. Since the MCSA photovoltaic panel modules (PPMs) are nearly identical to those of the International Space Station (ISS) photovoltaic arrays, MCSA operation offered an opportunity to gather multi-year performance data on this technology prior to its implementation on ISS. Two specially designed test sequences were executed in June and December 1996 to measure MCSA performance. Each test period encompassed 3 orbital revolutions whereby the current produced by the MCSA channels was measured. The temperature of MCSA PPMs was also measured. To better interpret the MCSA flight data, a dedicated FORTRAN computer code was developed to predict the detailed thermal-electrical performance of the MCSA. Flight data compared very favorably with computational performance predictions. This indicated that the MCSA electrical performance was fully meeting pre-flight expectations. There were no measurable indications of unexpected or precipitous MCSA performance degradation due to contamination or other causes after 7 months of operation on orbit. Power delivered to the Mir bus was lower than desired as a consequence of the retrofitted power distribution cabling. The strong correlation of experimental and computational results further bolsters the confidence level of performance codes used in critical ISS electric power forecasting. In this paper, MCSA flight performance tests are described as well as the computational modeling behind the performance predictions.
This final report presents conclusions/recommendations concerning the TRMM Solar Array; deliverable list and schedule summary; waivers and deviations; as-shipped performance data, including flight panel verification matrix, panel output detail, shadow test summary, humidity test summary, reverse bias test panel; and finally, quality assurance summary.
The MESUR solar array power model initially uses assumed values for many input variables.
Settling of atmospheric dust onto the surface of the solar array is potentially a lifetime-limiting factor for a power system on any Mars mission. For long-term operation of arrays on Mars, it will be necessary to develop techniques to remove deposited dust. Dust is expected to adhere to the array by Van der Waals adhesive forces. These forces are quite strong at the dust particle sizes expected. Dust removal methods must overcome this force. Removal methods can be categorized briefly into four categories: natural, mechanical, electromechanical, and electrostatic. The environment of Mars is expected to be an ideal one for use of electrostatic dust-removal techniques.
The Integrated Solar Array and Reflectarray Antenna (ISARA) mission will demonstrate a reflectarray antenna that increases downlink data rates for CubeSats from the existing baseline rate of 9.6 kilobits per second (kbps) to more than100 megabits per second (Mbps). A secondary payload called the CubeSat Multispectral Observation System (CUMULOS), is an experimental remote sensing payload also being demonstrated on this mission. A launch date for the ISARA spacecraft is currently pending.
Carbon composite solar array panel coupons similar to those considered for use on the planned NASA Europa Mission were exposed to energetic electrons to observe the size and rate of electrostatic discharge production. The discharge size and rate were found to be primarily dependent on the quantity of dielectric binder pooled on the surface of the sample. Cooling the samples to cryogenic temperatures increased the size of discharges but an end of mission radiation dose had only minimal effects. Samples were composed of M55J carbon composite impregnated with an RS3 binder. Additional samples tested included a co-cured Kapton surface.
Pointing the solar array to the Sun is critical for the power management in all so-lar-powered space missions. This is particularly true for NASA’s Psyche mis-sion, which will explore the large asteroid (16) Psyche orbiting the Sun at around 3 AU. Due to the large distance to the Sun, use of electric propulsion, and sophisticated scientific orbits at Psyche, the definition of this pointing accu-racy became a challenging task for the spacecraft design, which involves the GNC, Power, and Thermal subsystems as well as Mission Design. This paper describes the story of how this complicated requirement development evolved for Psyche mission.
With the retirement of Space Shuttle cargo delivery capability and the ten year life extension of the International Space Station (ISS) more emphasis is being put on preservation of the service life of ISS critical components. Current restrictions on the United States Orbital Segment (USOS) Solar Array (SA) positioning during Russian Vehicle (RV) departure from ISS nadir and zenith ports cause SA to be positioned in the plume field of Service Module thrusters and lead to degradation of SAs as well as potential damage to Sun tracking Beta Gimbal Assemblies (BGA). These restrictions are imposed because of the single fault tolerant RV Motion Control System (MCS), which does not meet ISS Safety requirements for catastrophic hazards and dictates 16 degree Solar Array Rotary Joint position, which ensures that ISS and RV relative motion post separation, does lead to collision. The purpose of this paper is to describe a methodology and the analysis that was performed to determine relative motion trajectories of the ISS and separating RV for nominal and contingency cases. Analysis was performed in three phases that included ISS free drift prior to Visiting Vehicle separation, ISS and Visiting Vehicle relative motion analysis and clearance analysis. First, the ISS free drift analysis determined the worst case attitude and attitude rate excursions prior to RV separation based on a series of different configurations and mass properties. Next, the relative motion analysis calculated the separation trajectories while varying the initial conditions, such as docking mechanism performance, Visiting Vehicle MCS failure, departure port location, ISS attitude and attitude rates at the time of separation, etc. The analysis employed both orbital mechanics and rigid body rotation calculations while accounting for various atmospheric conditions and gravity gradient effects. The resulting relative motion trajectories were then used to determine the worst case separation envelopes during the clearance analysis. Analytical models were developed individually for each stage and the results were used to build initial conditions for the following stages. In addition to the analysis approach, this paper also discusses the analysis results, showing worst case relative motion envelopes, the recommendations for ISS appendage positioning and the suggested approach for future analyses.
Four-paddle solar array orientation change effect on spin rate of rotating satellite
The development of a 100 kW planar low cost solar array is presented. Tasks are defined, objectives are stated, and test results are given. A study of alternate contact configuration concluded that gridded back contact cells sharply reduced solar absorptance, that copper contacts will require a developmental program before interface can be achieved, and that gridded back contact cells using Ti-Pd-Ag as the contact materials appear to be the most cost-effective design.
The investigation of the vibrational disturbances of the Hubble Space Telescope that were discovered soon after deployment in orbit is described in detail. It was found that the disturbances were particularly evident during orbital day-night crossings, and that the magnitude of the disturbances was considerably larger than the design jitter requirement. This paper describes the process by which the vibrations were characterized and isolated to a particular mechanism. The analysis of the flight data and comparisons with computer simulation results showed that the source of the disturbances was the thermally driven deformation of the solar arrays in conjunction with frictional effects in the array mechanisms. The control system was successfully modified to attenuate the disturbances to tolerable levels pending mechanical and thermal redesign of the solar arrays. The new arrays were installed during the first Space Telescope servicing mission, and in combination with the enhanced control system algorithm reduced the disturbances to satisfactory levels.
The solar array module plasma interactions experiment (SAMPIE) is an approved NASA flight experiment manifested for Shuttle deployment in early 1994. The SAMPIE experiment is designed to investigate the interaction of high voltage space power systems with ionospheric plasma. To study the behavior of solar cells, a number of solar cell coupons (representing design technologies of current interest) will be biased to high voltages to measure both arcing and current collection. Various theories of arc suppression will be tested by including several specially modified cell coupons. Finally, SAMPIE will include experiments to study the basic nature of arcing and current collection. This paper describes the rationale for a space flight experiment, the measurements to be made, and the significance of the expected results. A future paper will present a detailed discussion of the engineering design.
The Solar Array Module Plasma Interactions Experiment (SAMPIE) is an approved NASA Space Shuttle space flight experiment to be launched in Jul. 1993. The SAMPIE experiment is designed to investigate the interaction of high voltage space power systems with ionospheric plasma. To study the behavior of solar cells, a number of cell coupons, representing technologies of current interest, will be biased to high voltages to characterize both negative potential arcing and positive potential current collection. Additionally, various theories of arc suppression will be tested by including several specially modified cell coupons. Finally, SAMPIE will include experiments to study the basic nature of these interactions. The rationale for a space flight experiment, the measurements to be made, the significance of the expected results, and the current design status of the flight hardware are described.
Current estimates of solar-electric-propulsion spacecraft performance are based upon a solar-array output power which is degraded by approximately 10-13% to account for possible losses caused by proton, electron and micrometeorite damage. Past studies have used a worst case analysis in which the maximum degradation was taken to occur at the beginning of the mission. This paper presents a comparison of mission studies using a hypothetical exponential decrease in power with time, with those using a sudden degradation of solar power. These comparisons indicate that the performance gain by using a time varying degradation during the mission is quite small for outbound missions in the solar system. In addition an indication of the power allocation strategy to be followed during a mission is presented.
The National Aeronautics and Space Administration's (NASA's) Space Technology 5 (ST5) is designed to flight-test the concept of miniaturized 'small size" satellites and innovative technologies in Earth's magnetosphere. Three satellites will map the intensity and direction of the magnetic fields within the inner magnetosphere. Due to the small area available for the solar arrays, and to meet the mission power requirements, very high-efficiency multijunction solar cells were selected to power the spacecraft built by NASA Goddard Space Flight Center (GSFC). This was done in partnership with the Air Force Research Lab (AFRL) through the Dual-Use Science and Technology (DUS&T) program. Emcore's InGaP/lnGaAs/Ge Advanced triple-junction (ATJ) solar cells, exhibiting an average air mass zero (AMO) efficiency of 28.0% (one-sun, 28 C), were used to populate the arrays. Each spacecraft employs 8 identical solar panels (total area of about 0.3 square meters), with 15 large-area solar cells per panel. The requirement for power is to support on-orbit average load of 13.5 W at 8.4 V, with plus or minus 5% off pointing. The details of the solar array design, development and qualification considerations, as well as ground electrical performance & shadowing analysis results are presented.
Postflight analysis data on Nimbus 1 solar array drive motor failure
Analysis of operating problems of high voltage solar array in earth orbit