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Attachment methods for thermal control composite system comprised of optical solar reflectors and multilayer insulation to minimize heating from incident solar energy
Four advanced space radiator concepts that were pursued in an integrated effort to develop multi-mission-use and low cost heat rejection systems which can overcome the limitations of current radiator systems are briefly discussed and described. Also, in order to establish a firm background to compare the advanced space radiator concepts, the Orbiter active thermal control system is also briefly described.
MgF2 was investigated as a phase-change energy-storage material for LEO power systems using solar heat to run thermal cycles. It provides a high heat of fusion per unit mass at a high melting point (1536 K). Theoretical evaluation showed the basic chemical compatibility of liquid MgF2 with refractory metals at 1600 K, though transient high pressures of H2 can occur in a closed container due to reaction with residual moisture. The compatibility was tested in two refractory metal containers for over 2000 h. Some showed no deterioration, while there was evidence that the fluoride reacted with hafnium in others. Corollary tests showed that the MgF2 supercooled by 10-30 K and 50-90 K.
During this U.S. Spacesuit Knowledge Capture Program event, Dr. Eugene Ungar presented the background on the Earth and Moon solar, albedo, and outgoing longwave radiation (OLR) natural thermal environment parameters. He reviewed the interrelationship of the parameters and explore the physically correct method of combining them. Dr. Ungar also presented a sample analysis of a representative case.
At large distances from the Sun (e.g. Jupiter/Saturn), the solar flux is less than 4% that at Earth. This requires very large solar arrays to meet the power demands of a typical new mission concept. Heat loss through Multi-Layer Insulation (MLI) blankets typically constitutes the vast majority of the total heat loss. Hence, improvements to MLI blankets that reduce heat loss are advantageous to the spacecraft design. The overall effective emittance (ε*) of MLI is usually a range dependent on the number of layers, size of blanket, seams, feedthroughs, layer density and operating temperatures of the heat source and heat sink. A concept has been developed at JPL to reduce the ε* by as much as a factor of two, which produces a corresponding reduction in heat losses. This concept utilizes two MLI blankets physically separated by traditional bumpers or spacers used for micrometeoroid protection. The outer surface of the inner blanket and the inner surface of the outer blanket are low emissivity surfaces to further minimize the total ε* of the overall dual MLI system. Analytical predictions of a dual MLI concept have been made using test data based ε* correlations. A development test has been conducted to validate the dual blanket design’s performance. This paper will describe the dual blanket design concept, schemes for its implementation, and the corresponding test results to validate its performance.
Leveraging the speed of reduced-order models (ROMs), thermal analysis teams have previously had access to rapid optimization, sensitivity studies, rapid model correlation, and uncertainty quantification, to name a few. An advantage of the developed approach is its relative robustness, which enabled expansion to include structural models in addition to thermal models leading to both rapid thermal-structural and rapid structural-only analyses. This method was tested for a sample thermal-structural application and showed good performance. It was then applied to a complex NASA mission and showed how ROMs could be used to quickly evaluate system uncertainties. Finally, these methods were applied to a structural-only application and used to better optimize and understand the structural design.
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Thermal protection for spacecraft entering at escape velocity
AIMP-D spacecraft thermal design and mission planning report
Most spacecraft thermal analysis tools assume that the spacecraft is in orbit around a planet and are designed to calculate solar and planetary fluxes, as well as radiation to space. On NASA Constellation projects, thermal analysts are also building models of vehicles in their pre-launch condition on the surface of a planet. This process entails making some modifications in the building and execution of a thermal model such that the radiation from the planet, both reflected albedo and infrared, is calculated correctly. Also important in the calculation of pre-launch vehicle temperatures are the natural environments at the vehicle site, including air and ground temperatures, sky radiative background temperature, solar flux, and optical properties of the ground around the vehicle. A group of Constellation projects have collaborated on developing a cohesive, integrated set of natural environments that accurately capture worst-case thermal scenarios for the pre-launch and launch phases of these vehicles. The paper will discuss the standardization of methods for local planet modeling across Constellation projects, as well as the collection and consolidation of natural environments for launch sites. Methods for Earth as well as lunar sites will be discussed.
An integrated thermal protection system (TPS) for a spacecraft includes a grid that is bonded to skin of the spacecraft, e.g., to support the structural loads of the spacecraft. A plurality of thermally insulative, relatively large panels are positioned on the grid to cover the skin of the spacecraft to which the grid has been bonded. Each panel includes a rounded front edge and a front flange depending downwardly from the front edge. Also, each panel includes a rear edge formed with a rounded socket for receiving the rounded front edge of another panel therein, and a respective rear flange depends downwardly from each rear edge. Pins are formed on the front flanges, and pin receptacles are formed on the rear flanges, such that the pins of a panel mechanically interlock with the receptacles of the immediately forward panel. To reduce the transfer to the skin of heat which happens to leak through the panels to the grid, the grid includes stringers that are chair-shaped in cross-section.
Thermal Fluids and Analysis Workshop, Silver Spring MD NCTS 21070-15. NASA, the Defense Department and commercial interests are actively engaged in developing miniaturized spacecraft systems and scientific instruments to leverage smaller cheaper spacecraft form factors such as CubeSats. This paper outlines research and development efforts among Goddard Space Flight Center personnel and its several partners to develop innovative embedded thermal control subsystems. Embedded thermal control subsystems is a cross cutting enabling technology integrating advanced manufacturing techniques to develop multifunctional intelligent structures to reduce Size, Weight and Power (SWaP) consumption of both the thermal control subsystem and overall spacecraft. Embedded thermal control subsystems permit heat acquisition and rejection at higher temperatures than state of the art systems by employing both advanced heat transfer equipment (integrated heat exchangers) and high heat transfer phenomena. The Goddard Space Flight Center Thermal Engineering Branch has active investigations seeking to characterize advanced thermal control systems for near term spacecraft missions. The embedded thermal control subsystem development effort consists of fundamental research as well as development of breadboard and prototype hardware and spaceflight validation efforts. This paper will outline relevant fundamental investigations of micro-scale heat transfer and electrically driven liquid film boiling. The hardware development efforts focus upon silicon based high heat flux applications (electronic chips, power electronics etc.) and multifunctional structures. Flight validation efforts include variable gravity campaigns and a proposed CubeSat based flight demonstration of a breadboard embedded thermal control system. The CubeSat investigation is technology demonstration will characterize in long-term low earth orbit a breadboard embedded thermal subsystem and its individual components to develop optimized operational schema.
Operational characteristics of actuator for spacecraft thermal control system
This paper details the design developments of the Europa Clipper spacecraft thermal control system and its primary element the Heat Redistribution System (HRS). The Europa Clipper spacecraft will launch in the early 2020s and will have multiple flybys of Jupiter’s icy moon, Europa. The HRS, which utilizes a Mechanically Pumped Fluid Loop (MPFL) to reclaim waste heat from an avionics vault module and redistribute it to the propulsion module and radiator, has become an integral part of the spacecraft architecture both in terms of mechanical configuration and spacecraft operations capability. A walkthrough on the planned implementation of the HRS on the primary spacecraft modules is provided along with a description on thermal control of spacecraft hardware not located on the HRS vault, where thermal isolation and tolerance to extreme thermal environment drives the hardware designs.
Entry corridor thermal entry limits for Apollo spacecraft defined for design of thermal protection system
A simple thermal analysis program for Space Shuttle payloads has been developed to accommodate the user who requires an easily understood but dependable analytical tool. The thermal analysis program includes several thermal subprograms traditionally employed in spacecraft thermal studies, a data management system for data generated by the subprograms, and a master program to coordinate the data files and thermal subprograms. The language and logic used to run the thermal analysis program are designed for the small user. In addition, analytical and storage techniques which conserve computer time and minimize core requirements are incorporated into the program.
The Wide Field/Planetary Camera is an imaging system developed by the Jet Propulsion Laboratory for the NASA Hubble Space Telescope currently scheduled to be launched in December 1989 aboard the space shuttle. The temperature control design of the instrument utilizes multilayered insulation, electric resistance heaters, aluminum/ammonia heat pipes, thermoelectric coolers, temperature control coatings, and space radiators. A feedback control system maintains stable sensor temperatures. Thermal capacitance maintains stable optics and electronics temperatures during transient conditions. Schedule slips and launch delays have allowed extensive thermal testing of the instrument. Six instrument thermal vacuum tests and a spacecraft thermal vacuum test were performed. Several modifications have been made to the instrument to correct icing and contamination problems that have been discovered during thermal vacuum testing. This paper describes the thermal design, last instrument thermal vacuum test, results, and thermal model correlation.