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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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Solar electric power source easily storable and transportable. Source employs blanket of photovoltaic cells rolled up when not in use. The hollow core of canister reel used as storage space for radio equipment or first aid supplies, etc.
This paper suggests a new method for optimizing yaw maneuvers on the International Space Station (ISS). Yaw rotations are the most common large maneuvers on the ISS often used for docking and undocking operations, as well as for other activities. When maneuver optimization is used, large maneuvers, which were performed on thrusters, could be performed either using control moment gyroscopes (CMG), or with significantly reduced thruster firings. Maneuver optimization helps to save expensive propellant and reduce structural loads ‐ an important factor for the ISS service life. In addition, optimized maneuvers reduce contamination of the critical elements of the vehicle structure, such as solar arrays. This paper presents an analytical solution for optimizing yaw attitude maneuvers. Equations describing pitch and roll motion needed to counteract the major torques during a yaw maneuver are obtained. A yaw rate profile is proposed. Also the paper describes the physical basis of the suggested optimization approach. In the obtained optimized case, the torques are significantly reduced. This torque reduction was compared to the existing optimization method which utilizes the computational solution. It was shown that the attitude profiles and the torque reduction have a good match for these two methods of optimization. The simulations using the ISS flight software showed similar propellant consumption for both methods. The analytical solution proposed in this paper has major benefits with respect to computational approach. In contrast to the current computational solution, which only can be calculated on the ground, the analytical solution does not require extensive computational resources, and can be implemented in the onboard software, thus, making the maneuver execution automatic. The automatic maneuver significantly simplifies the operations and, if necessary, allows to perform a maneuver without communication with the ground. It also reduces the probability of command errors. The suggested analytical solution provides a new method of maneuver optimization which is less complicated, automatic and more universal. A maneuver optimization approach, presented in this paper, can be used not only for the ISS, but for other orbiting space vehicles.
High-strain composite deployable structures have been developed for systems such as solar arrays, camera masts or solar sailing propulsion elements. Composite booms in such applications are often flattened and then rolled into a small footprint for low-packaged volume and are then deployed in space. There is a need to obtain deformation for long, slender composite booms on earth through gravity offloading by suspending them vertically and applying distal end (tip) loads. Three-dimensional digital image correlation (3D-DIC), along with other measurement techniques, were used to obtain strain and displacement along the length of a 7.5 m subscale composite Triangular, Rollable, and Collapsible (TRAC) boom in preparation for full scale testing of a 30 m boom. However, incorporating 3D-DIC as a primary measurement tool on long, slender, high-aspect-ratio boom structures presents significant challenges. Challenges include small correlated area due to high aspect ratio, limited standoff distance due to size of test area, coordinate system alignment of multiple camera systems along the length of the boom, nodal mesh extraction for adequate test/analysis correlation, as well as measurement comparison between DIC and other instrumentation used such as fiber optic strain sensing (FOSS) and laser displacement tracking. The contents of the proposed paper will focus on techniques and methods for overcoming the previously mentioned challenges associated with applying 3D-DIC to long, slender boom structures. Results from subscale test along with lessons learned will be discussed.
The X-Ray Telescope (XRT) is an instrument on the National Aeronautics and Space Administration (NASA) SWIFT spacecraft. The thermoelectric cooler (TEC) for the charge coupled device (CCD) of the XRT requires a radiator temperature of -35 C or colder, and a goal of -55 C to minimize the damage by radiation. The waste heat rejected from the TEC to the radiator is in the 8 W to 20 W range. In the Phase A baseline design, the XRT radiator is mounted to the rear end of the XRT telescope tube and is very close to the bottom closeout of the spacecraft bus. The bottom closeout is multi-layer insulation (MLI) blankets. At sun angles between 90 deg and 180 deg, there is direct solar impingement on the bottom closeout. When the rolls +/- 5 deg, the XRT radiator is exposed to direct solar radiation. The radiator also has a view factor to the solar arrays. The results of thermal analysis showed that the flight temperature prediction of the radiator exceeds the temperature requirement of -35 C substantially at sun angles from 110 deg to 180 deg. A new location on the anti-sun side of the spacecraft is proposed for the radiator. It requires a heat pipe to couple the TEC and the radiator thermally. The results of thermal analysis show that the flight temperature prediction of the proposed radiator meets the temperature requirement at all sun angles.
A simulation test of the six-battery nickel-cadmium mission to determine battery life and electrical power system (EPS) performance characteristics for the Hubble Space Telescope (HST) program, is described. The basic HST power system requirements are to provide power generation, energy storage, and EPS control and distribution for 2.5 years with the nickel-cadmium (NiCd) batteries at an end of life solar array of 2 years. Mission simulation life testing begain in April of 1986, and the batteries have completed their 2.5 year mission requirement. Conditions as close as practical to the actual predicted mission profiles were used. These included solar array degradation, load variations, beta angle changes, temperature changes (with excursions to 10 C), battery reconditioning, safemode simulations, and off-normal roll activities. Discounting system cycling interruptions caused by hardware, software, human error, and periodic updates and revisions, the HST EPS breadboard hardware operated continuously for 30 months and demonstrated the power system's ability to meet the HST requirement.
Progress is reported on the design and development of lightweight, flexible, FEP encapsulated solar cell modules. Such modules are intended to form standardized building blocks for larger roll-up or fold-up solar cell blankets having specific performance in the order of 100 W/sq m and 80 W/kg in space near earth. Modules are constructed by heat-laminating interconnected 200-micron thick single crystal silicon solar cells with solderless front and back contacts between two layers of clear FEP. The outer FEP layer is 125 microns thick and serves as the solar cell cover material. The inner layer is 50 microns thick and serves as the adhesive between the solar cells and a 25-micron thick loadbearing Kapton substrate sheet.
This poster highlights the Game Changing Development Program task, Space Act Agreement between Maxar Technologies and NASA to test 5-junction solar cells on roll out compatible blankets through qualification protocols similar to AIAA Standards S-111/112. A description of the planned work and current status is included.
A six-degree-of-freedom (6DOF) simulation was constructed and exercised for a large solar electric propulsion (SEP) vehicle operating in low Earth orbit Nominal power was 500 kWe, with the large array sizes implied. Controllability issues, including gravity gradient, roll maneuvering for Sun tracking, and flexible arrays, and flight control methods, were investigated. Initial findings are that a SEP vehicle of this size is controllable and could be used for orbit raising of heavy payloads.
The present work assesses the increase in flight time and fuel consumption due to introducing attitude constraints on both the thrust vector and the plane of the solar cell arrays on geocentrically orbiting spacecraft. A modified version of the SECKSPOT computer program calculates nearly time-optimal trajectories for the constrained case of zero pitch and roll. Unconstrained cases are generated with the SECKSPOT code. It is concluded that with a pitch constraint but without a roll constraint, power would not be a function of thrust direction, and so the time-optimal thruster direction would be along the projection of the primer vector in the plane normal to the radius vector. The roll constraint would cause power to become a function of thrust angle and sun angle. For certain sun angles the locus of the ratio of power to maximum power is concave and thus there may be jumps in the control angle. Comparisons are made for a SERT-C type mission between constrained and unconstrained cases in an inverse square gravity field.
The Hubble space telescope (HST) solar array was designed to meet specific output power requirements after 2 years in low-Earth orbit, and to remain operational for 5 years. The array, therefore, had to withstand 30,000 thermal cycles between approximately +100 and -100 C. The ability of the array to meet this requirement was evaluated by thermal cycle testing, in vacuum, two 128-cell solar cell modules that exactly duplicated the flight HST solar array design. Also, the ability of the flight array to survive an emergency deployment during the dark (cold) portion of an orbit was evaluated by performing a cold-roll test using one module.
This paper presents the design and developmental testing associated with the bearing, motor, and roll ring module (BMRRM) used for the beta rotation axis on International Space Station Alpha (ISSA). The BMRRM with its controllers located in the electronic control unit (ECU), provides for the solar array pointing and tracking functions as well as power and signal transfer across a rotating interface.
Solar sails provide propulsion from solar photonic pressure without the need for propellant. This advantage can be reduced or eliminated if propellant is required for attitude control. The current state of the art for solar sail attitude control is that of the solar sail mission Near Earth Asteroid Scout, which employs an Adjustable Mass Translator for pitch and yaw momentum management of reaction wheels, but must use a cold gas Reaction Control System for roll axis momentum management. It is highly desirable for solar sails to use an attitude control system that does not include propellant, so here we present a novel method of managing momentum on the roll axis. For the upcoming Solar Cruiser solar sail mission, for roll axis momentum management we plan to use Reflectivity Control Devices (RCDs), thin-film Polyimide Liquid Crystal Devices that are capable of changing optical properties on command, thus affording the opportunity to provide “photonic only” control if placed appropriately on the surface of the sail. RCD design for roll momentum management is multifaceted, and must include performance (degree of change in optical properties), RCD array size, locations of clusters on the sail, orientation with respect to the sail plane, and orientation out of the sail plane. Here we provide an explanation of the basic technology, optimal orientation, sizing, and location of RCDs, results of trade studies for Solar Cruiser, lessons learned from Solar Cruiser, and scalability for future solar sail missions. RCDs will dramatically improve the performance of the Solar Cruiser Momentum Management System compared to that of NEA Scout and will continue to benefit other future solar sail missions.
Solar sails provide propulsion from solar photonic pressure without the need for propellant. This advantage can be reduced or eliminated if propellant is required for attitude control. The current state of the art for solar sail attitude control is that of the solar sail mission Near Earth Asteroid Scout, which employs an Adjustable Mass Translator for pitch and yaw momentum management of reaction wheels, but must use a cold gas Reaction Control System for roll axis momentum management. It is highly desirable for solar sails to use an attitude control system that does not include propellant, so here we present a novel method of managing momentum on the roll axis. For the upcoming Solar Cruiser solar sail mission, for roll axis momentum management we plan to use Reflectivity Control Devices (RCDs), thin-film Polyimide Liquid Crystal Devices that are capable of changing optical properties on command, thus affording the opportunity to provide “photonic only” control if placed appropriately on the surface of the sail. RCD design for roll momentum management is multifaceted, and must include performance (degree of change in optical properties), RCD array size, locations of clusters on the sail, orientation with respect to the sail plane, and orientation out of the sail plane. Here we provide an explanation of the basic technology, optimal orientation, sizing, and location of RCDs, results of trade studies for Solar Cruiser, lessons learned from Solar Cruiser, and scalability for future solar sail missions. RCDs will dramatically improve the performance of the Solar Cruiser Momentum Management System compared to that of NEA Scout and will continue to benefit other future solar sail missions.
The Pioneer Venus Orbiter (PVO) solar panel power output capability has degraded much more severely than has the power output capability of solar panels that have operated in earth-orbiting spacecraft for comparable periods of time. The incidence of solar proton events recorded by the spacecraft's scientific instruments accounts for this phenomenon only in part. It cannot explain two specific forms of anomalous behavior observed: 1) a variation of output per spin with roll angle, and 2) a gradual degradation of the maximum output. Analysis indicates that the most probable cause of the first anomaly is that the solar cells underneath the spacecraft's magnetometer boom have been damaged by a reverse biasing of the cells that occurs during pulsed shadowing of the cells by the boom as the spacecraft rotates. The second anomaly might be caused by the effects on the solar array of substances from the upper atmosphere of Venus.
The ISS utilizes two large rotating mechanisms, the SARJ, as part of the solar arrays alignment system for more efficient power generation. The SARJ is a 10.3m circumference, nitrided 15-5PH steel race ring of triangular cross-section, with 12 sets of trundle bearing assemblies transferring load across the rolling joint. The SARJ mechanism rotates continuously and slowly - once every orbit, or every 90 minutes. In 2008, the starboard SARJ suffered a lubrication failure, resulting in severe damage (spalling) of one of the race ring surfaces. Extensive effort was conducted to prevent the port SARJ from suffering the same failure, and fortunately was ultimately successful in recovering the functionality of the starboard SARJ. The M&P function was key in determining the cause of failure and the means for mechanism recovery. From a M&P lessons-learned perspective, observations are made concerning the original SARJ design parameters (boundary conditions), the perceived need for nitriding the race ring, the test conditions employed during qualification, the environmental controls used for the hardware preflight, and the lubrication robustness necessary for complex kinematic mechanisms expecting high-reliability and long-life.
The International Space Station (ISS) utilizes two large rotating mechanisms, the solar alpha rotary joints (SARJs), as part of the solar arrays' alignment system for more efficient power generation. Each SARJ is a 10.3m circumference, nitrided 15-5PH steel race ring of triangular cross-section, with 12 sets of trundle bearing assemblies transferring load across the rolling joint. The SARJ mechanism rotates continuously and slowly - once every orbit, or every 90 minutes. In 2007, the starboard SARJ suffered a lubrication failure, resulting in severe damage (spalling) to one of the race ring surfaces. Extensive effort was conducted to prevent the port SARJ from suffering the same failure, and fortunately that effort was ultimately successful in also recovering the functionality of the starboard SARJ. The M&P engineering function was key in determining the cause of failure and the means for mechanism recovery. From a M&P lessons-learned perspective, observations are made concerning the original SARJ design parameters (boundary conditions), the perceived need for nitriding the race ring, the test conditions employed during qualification, the environmental controls used for the hardware preflight, and the lubrication robustness necessary for complex kinematic mechanisms expecting high-reliability and long-life.
Substrates survive high deposition temperatures. High-temperature-resistant polymers candidate materials for use as substrates of lightweight, flexible, radiation-resistant solar photovoltaic cells. According to proposal, thin films of copper indium diselenide or cadmium telluride deposited on substrates to serve as active semiconductor layers of cells, parts of photovoltaic power arrays having exceptionally high power-to-weight ratios. Flexibility of cells exploited to make arrays rolled up for storage.
Seven low cost multi-100 kW planar solar array modules were fabricated and tested. Two different designs were used, demonstrating advanced solar array construction practices. Both module types utilized second generation gridded back cells featuring high efficiency and IR transparency. A silicon dioxide AR coating optimized for transmission at gamma = 1.7 microns was applied to the back surface. Two interconnect types, a single sheet printed circuit and a roll type, with alternate approaches to increasing transparency and reducing cost were designed and fabricated. Hinge stress and electrical power optimization were also examined. Two point designs were studied. The first design used a coilage longeron mast and is autonomously deployable. The second design used a Stac Beam for high natural frequency response and required astronaut assistance and assembly on orbit. It was conclusively demonstrated that planar arrays are the most cost effective design for use on the space station or other high power applications.