Spacecraft Radiator Research Articles

Investigation on Thermal Analysis of Spacecraft Radiators

In this research, the main objective is to compare the thermal properties of the rectangular and honeycomb structured space radiators used in spacecraft. The comparison of radiators was done by conducting steady state thermal analysis on these radiators by applying varying heat loads over it to analyse the thermal properties acting over them. The selection of appropriate material for the radiators was based on tensile strength, resistance to high temperature, and the type of working fluid considered. From the simulation results, the honeycomb structures of spacecraft radiator could withstand to a higher temperature gradient, and the rate of heat dissipation was much higher in comparison with rectangular fin type radiator, and hence it can be considered as a great replacement in near future.

Multiobjective and Multicase Optimization of a Spacecraft Radiator

Optimization is now an established tool for obtaining better solutions in engineering design. Nevertheless, the multiplicity of possible applications and the continuing search for better computational performance has motivated researchers in continuing to propose new methods, as well as improvements on existing ones. Not only optimization has allowed a better exploration of the design space, it also can significantly reduce the time of the design process. In this paper the advantages of using a fully multiobjective and multicase approach for the design of space radiators is presented. The optimization of Amazonia-1 satellite batteries panel thermal design is investigated. Instead of conventional space radiators, in this work an innovative solution considering the fin as heat absorber is analyzed. The results have shown that there are several viable configurations regarding to fin coating material, heater power consumption and additional mass, and the choice of the appropriate radiator configuration shall be evaluated based on system impact. The possibility of testing automatically different coatings and sizes for the fin, allowed by the optimization approach, increased significantly the probability of finding viable solutions on the design space.

Shaded Spacecraft Radiators to Be Used on the Daytime Surface of the Mercury Planet, the Moon, and Asteroids of the Solar System Inner Part

During the daytime a surface of the Moon, Mercury planet, and asteroids of the Solar system inner part, significantly heats up, and infrared radiation of the local soil becomes essential. At the same time direct solar radiation and reflected from the surface solar radiation reach the maximum too. These radiation fluxes can significantly decrease the efficiency of spacecraft radiators in the daytime. This effect is especially strong on the Mercury surface where direct solar radiation is 10 times stronger than solar radiation near the Earth. As a result, on the daytime surface of the Mercury the conventional low-temperature radiators become completely disabled. The article describes the development of the special shaded spacecraft radiators to be used in daytime on the Mercury and other atmosphereless bodies of the Solar system inner part. To solve this task are used mirror shades. The shape of these shades is developed to improve operation conditions of the spacecraft radiator through the appropriate scheme of radiation reflection. The task is discussed in 2D and 3D cases. A new design of shaded spacecraft radiators is proposed, and reasonable proportions of radiators are determined. The performance capability of proposed radiators for environments of the Mercury and the Moon is estimated using the zonal method in view of partial mirror reflection. The calculations showed that the developed shaded spacecraft radiators are capable to work on the Mercury surface as the low-temperature radiators even during the daytime. New radiators provide minimum accepted operating temperature of 241К (-32°С), meanwhile radiators of common design have minimum operating temperature of 479К (206°С). Using such radiators on the Moon enables us to increase effectiveness of spacecraft radiators and to decrease their minimum operating temperature from 270К (-3°С) to 137К (-136°С).

Optimizing the Design of Space Radiators for Thermal Performance and Mass Reduction

AbstractIn this paper, a spacecraft radiator formed in a honeycomb structure is designed to enhance the thermal performance while reducing its mass. Examples of design guidelines for radiator configurations, such as the distance between heat pipes, facesheet thickness, and honeycomb core density, are suggested. To derive the analytic solution of the governing equation, a linear approximation is used and the accuracies of the solutions are verified with a fourth-order finite-difference method. There exist optimal combinations of design parameters that minimize the radiator mass while maintaining its heat rejection capacity. The heat rejection rate that minimizes the mass per unit heat rejection and the pertinent radiator shape also is presented. The combinations of optimal design are different among the three surface treatments and their characteristics are investigated.

Development of a Rotary Joint for a Steerable Spacecraft Radiator

The technology for a steerable spacecraft radiator with a fully functional rotary joint capable of operating with anhydrous ammonia would be very beneficial for thermal control systems onboard future spacecraft. Different slip-ring rotary joints capable of continuous 360 deg rotation and a flexible pressure line joint or “twist capsule” with oscillatory motion were investigated. A rotary joint with commercial-off-the-shelf ethylene propylene diene monomer O rings and one with ethylene propylene diene monomer O rings that were preconditioned to be compatible with ammonia were designed and tested, and they were determined to not be viable candidates for a rotary seal with ammonia working fluid. A rotary joint with ultrahigh-molecular-weight polyethylene lip seals was also explored and has demonstrated at least an equivalent 15 year life in low Earth orbit. The ultrahigh-molecular-weight polyethylene lip seal rotary joint has a dynamic sealing capability from at least to 70°C. A flightlike unit is being built with enhanced fidelity to be tested and qualified for flight. In parallel to the 360 deg rotary joints, a twist capsule rotary joint was also designed and tested. The twist capsule oscillates between 0 and 270 deg, and the first-generation model demonstrated a life capability equivalent to 27 years in low Earth orbit. An improved flightlike twist capsule rotary joint is in development to be tested and qualified for flight.

Node-based spacecraft radiator design optimization

Spacecraft radiators can be designed through thermal analysis of a thermal model divided into nodes. The radiator is thus designed discretely in unit nodes; then, the radiator node distribution over the node division of a candidate radiator region indicates the radiator design as radiator node combinations. In particular, the radiator sizes and the topologies of shapes and locations are key design factors for efficient thermal design. These design factors can be considered, to some extent, at the typical resolution levels for the node division of the thermal model. This paper proposes two radiator design optimization methods that are based on node division of the thermal model; these methods find the optimal solutions of radiator node combinations in different ways. The first method uses an integrated optimization analysis that combines an optimization algorithm with thermal analysis; binary design variables are assigned to each node division to represent the radiator node combination and serve as the interface between two sets of software. The optimization problem is formulated as a multi-objective problem and improved through thermal design heuristics. The second method is a stepwise approach that adds a radiator node with the highest temperature sensitivity until the temperature limits are satisfied. A thermal model of a small spacecraft was developed to verify the proposed optimization methods, and test problems of various versions of the thermal model were appropriately defined. The numerical optimal solutions for the test problems using these methods showed good agreement with the analytic solutions. Therefore, the applicability and feasibility of the present methods for practical radiator design were confirmed.

Multi-Objective Optimization of Node-Based Spacecraft Radiator Design

An efficient multi-objective optimization method is proposed for node-based spacecraft radiator design. In contrast to a conventional manual approach to node-based radiator design with a thermal model, the proposed approach finds an optimum design solution for radiator node combinations through an optimization algorithm. The important parameters of radiator design are radiator size and topology, represented by discrete binary design variables allocated to each node division in the candidate radiator region. The optimization problem is formulated as a multi-objective problem with two or more objectives to minimize the number of radiator nodes and temperature margins of unit boxes, and a genetic algorithm suitable for multi-objective optimization is used. A small thermal model (verification thermal model) was developed to verify the proposed multi-objective optimization of the node-based spacecraft radiator design method, and test problems for this thermal model were defined. The numerical optimal solutions for test problems show good agreement with the analytic optimal solutions. Therefore, the applicability and feasibility of the multi-objective optimization of node-based spacecraft radiator design method to practical radiator design were confirmed.

Flexible Heat Pipes for a Lightweight Spacecraft Radiator

Flexible heat pipes utilizing ultra-high-molecular-weight (UHMW) Ž lm wicks and formed UHMW wicks with Kapton® and aluminized Mylar® containers were designed, fabricated, and tested. Wick characterization was performed on both wicks with acetone, methanol, freon, and ethanol. The thin-Ž lm UHMW wick had a larger permeability and capillary radius. Burst tests were performed on the Mylar and Kapton heat-pipe containers indicating sufŽ cient strength for the heat-pipe application. The heat pipes were charged with methanol as the working  uid and successfully tested at steady-state conditions. Relatively uniform temperatures were obtained during several days of testing.

Thermal management for multifunctional structures

Multifunctional structures (MFS) are an innovative concept that offer a new methodology for spacecraft design, eliminating chassis, cables and connectors, and integrating the electronics into the walls of the spacecraft. The MFS design consists of multilayer flexible circuit patches bonded onto a structural composite panel, and multichip modules (MCMs) performing specific functions are bonded onto the circuit patches which are interconnected via flexible circuit jumpers. Incorporation of the high power density two-dimensional (2-D) and three-dimensional (3-D) MCM's into smaller and more efficient packaging designs still has the fundamental requirement to maintain component temperatures within design limits. Higher component qualification temperatures, such as 393 K, can result in smaller spacecraft radiator areas that are consistent with efficient packaging schemes. During the MFS development effort, a structural radiator panel was fabricated using high thermal conductivity (Hi-K) composite facesheets, and several thermal management designs using combinations of Hi-K doublers (150-1500 W/m-K), Hi-K (150-700 W/m-K) corefill, and deployable radiators to maximize MCM's heat rejection. Results of the thermal vacuum tests and details of the thermal design methodology are presented in this paper.

Electroless nickel plating on magnesium alloy

a traveling wave tube (TWT) for amplification of signals. In a TWT the collector is the main source of heat dissipation. The electrodes are heated by the residual kinetic energy of the collected electrons; however, the elec- trodes must be maintained at a rela- tively low temperature. There are two ways to dissipate this heat into cold space (see Fig. 1). 1. Heat conduction from the elec- trodes to the spacecraft radiator. This is called conduction cooling and is suit-