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In this work, we conduct a thermal pointing error analysis of the observation satellites considering seasonal and daily temperature variation with interpolated temperature based on prescribed average temperature (PAT) method. Maximum 200 degree temperature excursion is applied to the observation satellites during on-orbit operation, which cause the line of sight (LOS) to deviate from the designated pointing direction due to thermo-elastic deformation. To predict and adjust such deviation, the thermo-elastic deformation analysis with a fine structural finite element model is accomplished with interpolated thermal maps calculated from the results of on-station thermal analysis with a coarse thermal model. After verifying the interpolated temperatures by PAT with two benchmark problems, we evaluate the thermal pointing error.

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In this work, satellite FE (finite element) model updating for the prediction of the effect of micro-vibration is described. In the case of satellites launched in low earth orbit, high agility and more mission accomplishments are required by the customer in order to procure many images from satellites. To achieve the goal, many mechanisms, including high capacity wheels and antennas with multi-axis gimbals have been widely adopted, but they become a source of micro-vibration which could significantly deteriorate the quality of images. To investigate the effect due to the micro-vibration in orbit on the ground, a prediction is conducted through an integrated model coupling the measured jitter sources with FE (finite element) model. Before prediction, the FE model is updated to match simulation results with the modal survey test. Subsequently, the quality of FE model is evaluated in terms of frequency deviation error, the resemblance of mode shapes and FRFs (frequency response functions) between test and analysis.

When developing medium satellites or large satellites, coupled load analysis(CLA) is performed in order to verify satellite design as a final assessment under launch environment. Maximum acceleration, gap between adjacent parts, internal loads obtained from CLA are used to assess the safety of satellite design by comparing them with the allowable loads of every component. To achieve reliable CLA results, satellite FE model have to be properly updated to match with the sine vibration test results. In this paper, the validation procedure of satellite FE model and its results are discussed.

Korean lunar explorer will be launched by korean launcher KSLV-2 in the 2020s in accordance with national space development strategy. Korean lunar explorer is composed of two unmanned orbiter and lander and should be developed as small size and light weight within 550kg of launch mass due to launcher`s loading capability. A structure of lunar explorer is required to have sufficient stiffness and strength under launch and operational environment as well as to accommodate mission equipment. This paper describes the result of a preliminary study on structural model design for korean lunar explorer.

In earth orbit, a great number of orbital debris move around in extremely high velocity, and they become serious threats to satellites. In this study, smoothed particle hydrodynamics(SPH) is used to analyze the damage of a low earth orbit satellite due to the hypervelocity impact with orbital debris. The damage of honeycomb sandwich panel(HC/SP) used for walls of a satellite is analyzed with respect to impact velocities. For the additional analysis to examine the safety of interior components of the satellite, an attached electronic box and an offset electronic box are considered. As a result of the analysis considering the orbital debris having a probability of collision more than 2% at altitude of 685km, it is shown that the HC/SP can be perforated but only small craters are formed on both the attached electronic box and the offset electronic box.

The MFS (Mlti-Functional Structure) concept, which integrates the electronics, thermal control and structure into a single packaging system, has been developed and applied to reduce the volume and weight of the satellite. Therefore, this MFS can eliminate the bulky chassis/frames, cables and connectors of the electronic equipment. The main point of this traditional MFS is the replacement of the electrical chassis/frames with MCMs (Multi-Chip Modules) that require much costs and efforts for developing. This paper shows the new MFS concept that effectively saves the volume and weight. The structure including the thermal control and radiation shielding elements will be designed and manufactured as the rectangular grid-stiffened structure. The rectangular grid-stiffened structure is the modification of the iso-grid structure, and provides the enough spaces for putting the general PCBs without the chassis/frames.