Katsuhiko Nakajima

Thermal design verification of a large deployable antenna for a communications satellite

A large deployable antenna for a communication satellite requires sophisticated thermal control to satisfy the temperature requirements for electrical characteristics, and its performance must be confirmed by a thermal balance test. The results of a tradeoff study of the thermal control method for an antenna conducted in an effort to meet temperature requirement demands indicate that the thermal design of an antenna system can be accomplished by using passive thermal control techniques and heaters in spite of the large and complicated structure. Antenna system thermal balance tests are limited by the volume of the space simulation chamber. To overcome this problem, we introduce a two-step thermal design verification method consisting of component level tests and a whole antenna system level test. This paper describes the thermal control method, the thermal design verification method, and the predicted antenna temperatures in a geostationary orbit obtained from the verified thermal analytical model.

Experimental results for capillary looped pipe applied to direct cooling method

The heat transport characteristics of capillary looped pipes are obtained from experimental results. Both indirect-cooling and direct-cooling capillary looped pipes were tested. An indirect-cooling looped pipe was tested to determine the flow direction of the working fluid and the relationship between the liquid charge quantity and the heat transfer characteristics. Tests show that using the direct-cooling configuration is superior to the indirect-cooling configuration.

Deployment mechanism for a large reflector - Thermal-design verification using flight data

The thermal design of the deployment mechanism of a 3.5-m-diam reflector was certified using flight temperatures. The flight temperatures were recorded from Engineering Test Satellite Six, which was launched on Aug. 28, 1994, using an H-II vehicle. The satellite reflectors were successfully deployed on the sixth day after liftoff. The temperatures of the deployment mechanism and the difference in temperatures between the inner and the outer ring of the bearing for the deployment mechanism fell within the allowable temperature range. The minimum heating rate for the deployment mechanism was determined by taking into consideration the minimum allowable bearing temperature. The temperatures were evaluated using a thermal mathematical model of an antenna and a detailed model of the deployment mechanism. The antenna model was used to obtain boundary temperatures in the detailed model. This detailed model included the thermal contact resistance between the inner and the outer rings of the bearings. The predicted temperatures agreed with the flight data within 6%. The correct boundary temperatures are important in determining the exact temperature for the deployment mechanism. The thermal contact resistance between the inner and the outer ring was evaluated, considering the elastic deformation of the bearing due to the temperature difference between them. The thermal resistance in the flight agreed well with the value estimated in the ground tests.

Thermal Design Evaluation of Large Reflector Deployment Mechanism by Using Flight Data.

On the deployment mechanism of a large deployable onboard antenna of ETS-VI (Engineering Test Satellite Type 6), flight data are compared with predicted one. The deployment mechanism is thermally suitably controlled by heaters and by MLI (Multi-Layer Insulation) blankets. Such thermal control permits of deploying the main reflector within allowable temperature limits. That mechanism has thermally been evaluated with a mathematical model including thermal resistances between the inner and outer bearing rings. A good accuracy of the proposed method has also been demonstrated from flight data.

Theoretical Study on the Thermal Contact Resistance of a Space-Use Deep Groove Ball Bearing.

The thermal contact resistance between the balls and the inner and outer rings of a space-use deep groove ball bearing is analyzed assuming that heat transfer between smooth contacting elements occurs through the elastic contact areas. It is also assumed that the stationary bearing sustains axial and/or radial loads under steady-state temperature condition. The shapes and sizes of the contact areas are calculated using the Hertzian theory. The thermal analysis is based on an isolated isothermal elliptic contact area supplying heat to an insulated half-space. The formulation of the resistance is given as a function of a geometric factor of the contact area and the thermal conductivity of the bearing. In particular, an expression for the axial load is derived with careful consideration of changes in contact angle induced by elastic deformation at the contact area.

Experimental Investigation of the Thermal Contact Resistance of a Space-Use Deep Groove Ball Bearing.

An investigation is conducted to experimentally verify an analytical method that determines the thermal contact resistance between the balls and the inner and outer rings of a space-use ball beardng. A single row bearing made from stainless steel 440C is tested in a vacuum environment, and steady-state temperature distributions are measured to evaluate the heat flow across the stationary bearing. Test results are given for the conditions of axial, radial and combined loading. Excellent agreement between the measured and predicted values of the thermal contact resistance is found under all types and magnitudes of the applied loads. It is concluded that the proposed calculation method accurately predicts the thermal contact resistance between the elements of a dry bearing with smooth contact surfaces.

A Design Method of Axially Grooved Heat Pipes Embedded in Equipment Panel for Communication Satellite.

A calculation method of the maximum heat load for an axially grooved heat pipe which is embedded in a honeycomb sandwich panel with multi-point heating is developed by considering the estimation of heat flux rate along the heat pipe. A thermal mathematical model for the panel is also used to estimate the net heat input to the heat pipe. The maximum heat loads predicted for the heat pipe embedded in the panel show good agreement with the data obtained from tests which has been performed in a vacuum chamber. A minimum weight design method for rectangular grooved heat pipes which satisfied heat transport capabilities required are also proposed as a result of this study.

Temperature Calculation of Rectangular Radiative Fins Using a Linearized Method.

An infinite series solution presented for a thin rectangular fin is developed for the steady temperature distribution in a two-dimensional rectangular sandwich panel fin heated within a rectangular footprint region, and losing energy to environment by linearized radiation. The solutions approximate a spacecraft application where a heat dissipating electronic component is mounted to a heat-sink plate or a equipment panel. The comparison of numerical results obtained from the proposed method and the lumped nodal method shows that the formulations will be useful in evaluating heat-sink designs where geometry, heat loads, thermal properties, and environmental parameters change frequently.

Thermal analysis method of high capacity communications satellite with heat pipes

Thermal analysis method for heat pipe embedded communications equipment panel is treated in this paper. The main problem of the thermal analysis is how to construct the mathematical model under the limitation of computer CPU memory size. The mathematical model for the heat pipe embedded panel is first established based on the experiments. The essence of this method is to divide panel area into several small regions and perform thermal analysis independently using the fact of low thermal conductivity of honeycomb sandwich panel. To check the correctness of this method, the experiment using the test panel which thermally simulates the north communications equipment panel of two-ton class high capacity communications satellite has been conducted. The experiment shows the methods works well. The transponder size for the high capacity communications satellite is required to be as small as possible. The small size transponder makes it easy to install transponders into the satellite. But, the smallness of the transponder size means the high generated heat to foot print area size ratio. The generated heat must be dispersed to the wider area than foot print area to keep transponder temperature within permissible range. The weight of thermal control system tends to increase as long as the conventional thermal control technique such as heat sinks and thermal plugs are used. The heat pipe embedded honeycomb sandwich panel with aluminum alloy face sheet? is a promissing solution to ease the situation. Two problems must be solved to make the successful thermal analysis for the heat pipe embedded communications equipment panel. One is how to construct the mathematical model for thermal analysis. The temperature distribution of the panel with the embedded heat pipes is pretty different from that of the panel with heat sinks. The other problem is how to evade the difficulty arising from the increase of the CPU memory size. The communications equipment panel with many transponders requires naturally many nodes to make * Research Engineer ** Senior Research Engineer, Member AIAA accurate temperature prediction. The use of heat pipe increase the number of nodes further. The thermal analysis treating the entire panel at a time requires the large CPU memory size the ordinary CPU memory size restriction. i m w g This paper describes the exact thermal mathematical model of the heat pipe embedded communications equipment panel based on the experiments. Then the method of analysis which circumvent the difficulty of the increase of CPU memory size is described. 2.Thermal Mathematical Model of the Heat Pipe Embedded Panel Thermal Control Using the Embedded Heat Pipes The most important problem for satellite design is how to disperse the heat from transponders efficiently. The transponder size is desired to be as small as possible because the small size makes it easy to accommodate many transponders in the satellite of which size is limited by the rocket fairing dimen-ons. But, the smaller the transponder size is the more difficult the heat dissipation becomes, because the required heat dissipation per unit area increases. The conventional thermal control method for transponders uses heat sinks which makes the temperature distribution uniform using high thermal conductive material. However, the increase in transponder heat density requires a= increase in heat sink weight. On the other hand, the heat pipe can extend the heat along the pipe to practically any length. The required heat radiation becomes possible using the thin face sheet without the heavy heat sinks. The heat pipe, however, must satisfy two conditions to perform the role of heat sinks. It must be light weight, and it must have a form which can give high thermal conductance between transponders and the heat pipe container. The thickness of the heat pipe container must be cut as thin as possible and the unnecessary part of the support section must also be removed. The high conductance can be realized by attaching large fin to the heat pipe section on which a transponder is laid. The cross section of the heat pipe must be as shown in Fig.1. Fig.2 shows the cross section of the heat pipe embedded honycomb panel. The transponder and heat pipe fin are tightened by bolts to get good contact between the transponder and the face sheet, the face sheet and the heat pipe. F i g . 1 C r o s s s e c t i o n o f t h e h e a t p i p e