A commercial spacecraft should survive on orbit for more than 10 years under the severe circumstances without any maintenance. To realize this subject, not only performance but also other design factors such as reliability, mass, robustness, cost, etc. should be taken into consideration. From point of the thermal design, it is very important to obtain the robust thermal control subsystem with matrix heat pipe layout while minimizing the mass (weight). A new thermal optimization method without compromising the thermal robustness and the mass of thermal subsystem is highly anticipated. This paper proposes a robust thermal design approach for optimizing the heat pipe shape to minimize the mass of the spacecraft panel. We apply a combination of Design of Experiments (DOE), Response Surface Methodology (RSM) and Monte Carlo Simulation to determine the robust design parameters that minimize the mass of the heat pipe structure. Dimensions of the heat pipe design parameters were determined with rationally in a short time and practical robust optimization method was established.
A commercial spacecraft should survive on orbit for more than 10 years under the severe circumstances without any maintenance. To realize this subject, not only performance but also other design factors such as reliability, mass, robustness, cost, etc. should be taken into consideration. From point of the thermal design, it is very important to obtain the robust thermal control subsystem with matrix heat pipe layout while minimizing the mass (weight). A new thermal optimization method without compromising the thermal robustness and the mass of thermal subsystem is highly anticipated. This paper proposes a robust thermal design approach for optimizing the heat pipe shape to minimize the mass of the spacecraft panel. We apply a combination of Design of Experiments (DOE), Response Surface Methodology (RSM) and Monte Carlo Simulation to determine the robust design parameters that minimize the mass of the heat pipe structure. Dimensions of the heat pipe design parameters were determined with rationally in a short time and practical robust optimization method was established.
A deployable radiator (DPR) on the Engineering Test Satellite-VIII (ETS-VIII) was launched to a geo stationary orbit using H-IIA rocket on 18 Dec. 2006, carrying a deployable radiator (DPR). Deployable radiators are anticipated for extension of the radiator area on satellites to mitigate increased heat generation. The DPR is an experimental apparatus on the ETSVIII; it has a reservoir-embedded loop heat pipe (RELHP) for heat transfer from the experimental heat load to the DPR radiator. The RELHP reservoir is embedded in the evaporator for reliable function of the Loop Heat Pipe (LHP). This paper describes the DPR and the RELHP characteristics on the ETS-VIII from its launch operation to the beginning of the test under an orbital environment.
A deployable radiator (DPR) on the Engineering Test Satellite Kiku 8 (ETS-VIII) was launched to a geo stationary orbit by H-IIA rocket on 18Dec. 2006. Deployable radiators are anticipated for extension of the radiator area on satellites to mitigate increased heat generation. The DPR is an experimental apparatus on the Kiku 8; it has a Reservoir Enbedded Loop Heat Pipe (RELHP) for heat transfer from the experimental heat load to the DPR radiator. The RELHP reservoir is embedded in the evaporator for reliable function of the Loop Heat Pipe (LHP). This paper describes the characteristics of the DPR and the RELHP on ETS-VIII from the launch operation time to the first DPR test in an orbital environment.
High-powered satellites need larger heat rejection area. Deployable radiator is one of key technologies for high-powered satellite bus. Reservoir Embedded Loop Heat Pipe (RELHP) is a two-phase heat transfer device and constitutes the deployable radiator. RELHP has an evaporator core which is used as a liquid reservoir to enhance operational reliability. For use on satellites, RELHP is required over 10 years lifetime. In case of conventional heat pipes, it is generally known non condensable gas (NCG) make worse heat transport characteristics. On the other hand, influence of NCG on RELHP is not still obvious. This paper presents the heat transport characteristics of RELHP in case of changing NCG volume by experiment and calculation. It was found that NCG increases temperature rise at evaporator. NCG volume in RELHP has great influence on heat transport characteristics due to pressure increase in reservoir by NCG.
A deployable radiator (DPR) on the Engineering Test Satellite-VIII (ETS-VIII) was launched to a geo stationary orbit by the H-IIA rocket on 18 Dec. 2006. The DPR, an experimental apparatus on ETS- VIII; it has a Reservoir-Embedded Loop Heat Pipe (RELHP) for heat transfer from the experimental heat load to the DPR radiator. The RELHP reservoir is embedded in the evaporator for reliable function of the Loop Heat Pipe (LHP). This paper describes the characteristics of the DPR and the RELHP under an orbital environment over three years.
Heat pipes have capillary pumping limit when they operate under micro-gravity condition or axially top heating mode under normal-gravity. A theoretical model was developed to predict temperature rise in evaporator with asymmetrical heating and the capillary pumping limit. In this model, puddling effect, liquid recession into groove bottom and 3-dimensional heat conduction in the evaporator were taken into consideration. The prediction under micro-and normal-gravity condition agreed well with experimental data. It is found by calculation that the temperature at the evaporator end begin to increase just after all gooves dry out at the evaporator end in case of both symmetrical and asymmetrical heating and the capillary pumping limit is not much affected by heating condition when the thermal conductivity of heat pipe envelope is large.
