• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.12
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• pp.1216-1221
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• 2008

A low earth orbit satellite with a fixed solar array always has a sun-pointing attitude during daylight, and changes into a nadir-pointing attitude for a imaging mission. Since external heating sources to the satellite panels are Earth　irradiation and Albedo during most of daylight in a sun-pointing attitude, the thermal environment condition is relatively stable. However, direct sunlight which is the greatest environmental heating has an affect on the satellite panels during a mission period (10% of one orbit) in a nadir-pointing attitude. In satellite thermal design, thermal effects of a nadir-pointing mission attitude due to this thermal environment change need to be evaluated although the duration of a nadir-pointing attitude is short. Therefore, a nadir-pointing attitude during a mission is incorporated into thermal model and by the thermal analysis result, thermal effects on the satellite are investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.9
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• pp.795-801
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• 2014

STEP Cube Lab. classified as a pico-satellite has been being developed by SSTL(Space Technology Synthesis Laboratory) in Chosun University. Its main mission objective is to perform the on-orbit verification of core space technologies, which will be the potential candidates for future space missions. In this paper, to guarantee successful mission operation of the cube satellite under extremely severe space thermal environment condition, the system level thermal design and analysis has been performed. The effectiveness of the design has been verified through on-orbit thermal analysis of cube satellite.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.17-23
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• 2020

Upper stage of launch vehicle mainly injects a lunar explorer from low earth orbit to the moon at a distance of 380,000 km. In foreign lunar explorer, the upper stage is separated from the explorer after the explorer is injected into the earth-moon transfer trajectory, and the lunar explorer then uses on-board propellant to carry out mid-course correction maneuvers and lunar orbit insertion maneuvers. This study describes a newly presented small liquid upper stage. Using a small liquid upper stage with a wet mass of 2.9 tonnes, the lunar explorer not only can be injected earth-moon transfer trajectory but also can be performed lunar orbit insertion. This study provides acceptable mass range of the lunar explorer and the scope of acceptable mission range also describes based on the launch from Naro Space Center.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.1
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• pp.31-38
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• 2022

TR module using in satellite must consider discriminative electrical and mechanical requirements compare to the one using in ground and aircraft system since not only the environment level of vibration and shock during the launch stage but also the level of radiation, vacuum and thermal variation from orbit environment are more severe than atmosphere condition. This paper describes the environmental conditions of launch and the orbit and, suggests design method of TR module applying GaN to satisfy the unique environmental requirements of satellite systems by especially focusing on parts selection, derating design, RF budget design, manufacturing process design, and thermal design of TR module.

• Journal of Space Technology and Applications
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• v.4 no.1
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• pp.27-47
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• 2024

In the New space era, satellites are being developed to perform on-orbit service (OOS) missions. Various missions for orbital service include failure repair, refueling, towing, component replacement, and space construction, and in order to do so, a robot arm payload must be mounted. Unlike conventional satellite payloads, the robot arm payload is not move in a fixed state, but is a payload that must move continuously to perform the mission. It is also characterized by the need to perform the mission while being directly exposed to outer space, rather than existing inside the structure of the satellite. Due to the characteristics of these payloads, thermal design and interpretation that can be operated smoothly in an extreme space thermal environment is essential, but there are not many papers on thermal design and interpretation of the robot arm. This paper introduces and summarizes cases of thermal design and interpretation of robot arm payloads developed so far, and finally, it intends to suggest directions for thermal design and interpretation of robot arm payloads to be developed in the future.

• Journal of Astronomy and Space Sciences
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• v.27 no.3
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• pp.195-204
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• 2010

The optimal Earth-Moon transfer trajectory considering spacecraft's visibility from the Daejeon ground station visibility at both the trans lunar injection (TLI) and lunar orbit insertion (LOI) maneuvers is designed. Both the TLI and LOI maneuvers are assumed to be impulsive thrust. As the successful execution of the TLI and LOI maneuvers are crucial factors among the various lunar mission parameters, it is necessary to design an optimal lunar transfer trajectory which guarantees the visibility from a specified ground station while executing these maneuvers. The optimal Earth-Moon transfer trajectory is simulated by modifying the Korean Lunar Mission Design Software using Impulsive high Thrust Engine (KLMDS-ITE) which is developed in previous studies. Four different mission scenarios are established and simulated to analyze the effects of the spacecraft's visibility considerations at the TLI and LOI maneuvers. As a result, it is found that the optimal Earth-Moon transfer trajectory, guaranteeing the spacecraft's visibility from Daejeon ground station at both the TLI and LOI maneuvers, can be designed with slight changes in total amount of delta-Vs. About 1% difference is observed with the optimal trajectory when none of the visibility condition is guaranteed, and about 0.04% with the visibility condition is only guaranteed at the time of TLI maneuver. The spacecraft's mass which can delivered to the Moon, when both visibility conditions are secured is shown to be about 534 kg with assumptions of KSLV-2's on-orbit mass about 2.6 tons. To minimize total mission delta-Vs, it is strongly recommended that visibility conditions at both the TLI and LOI maneuvers should be simultaneously implemented to the trajectory optimization algorithm.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.9
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• pp.752-759
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• 2012

This paper presents a mission orbit design method for spacecraft which use the sun-synchronous and ground repeat orbits. In this work, we have proposed a new design procedure, "Nonlinear simulation-based numerical optimization technique" using the commercial S/W's such as STK (Satellite Tool kit) and Matlab, which are widely adopted S/W's in the area of orbital mechanics and engineering analysis. Inclusion of all the perturbation effects on the spacecraft not only can more precisely satisfy the mission requirements for sun-synchronicity and repeated ground track, and also operational requirements such as minimum change in the S/C local time, maximization of the contact time with a specified ground station, etc. can be appropriately considered. Design examples for LEO sun-synchronous mission are presented to demonstrate the usefulness of the proposed method in this paper.

• Aerospace Engineering and Technology
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• v.1 no.2
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• pp.57-65
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• 2002

The Electrical Power System (EPS) shall supply required power to maintain spacecraft and payload during the mission. The EPS sizing are based on space environment, satellite mission and lifetime, and allocated budgets. The type of the primary and secondary power is determined according to satellite design-level and allocated subsystem budgets. The design of EPS has closely related to system and others' subsystems design. To supply the sufficient power to the satellite, the implementation of the larger power source and energy storage is impossible actually. And there will be some problems of the attitude control of the satellite, the handling power capability of the electronic boxes, and launch vehicle selection caused by EPS oversizing. Also, the thermal control is not easy in the space by extra power. And the maintenance of the satellite within the specific orbit from orbit-drag is a big design burden of the thruster. So the various technologies have been developed to optimize the EPS sizing and to operate the power system efficiently.

• Journal of Astronomy and Space Sciences
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• v.38 no.4
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• pp.217-227
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• 2021

In this study, we describe an analytical process for designing a low Earth orbit constellation for discontinuous regional coverage, to be used for a surveillance and reconnaissance space mission. The objective of this study was to configure a satellite constellation that targeted multiple areas near the Korean Peninsula. The constellation design forms part of a discontinuous regional coverage problem with a minimum revisit time. We first introduced an optimal inclination search algorithm to calculate the orbital inclination that maximizes the geometrical coverage of single or multiple ground targets. The common ground track (CGT) constellation pattern with a repeating period of one nodal day was then used to construct the rest of the orbital elements of the constellation. Combining these results, we present an analytical design process that users can directly apply to their own situation. For Seoul, for example, 39.0° was determined as the optimal orbital inclination, and the maximum and average revisit times were 58.1 min and 27.9 min for a 20-satellite constellation, and 42.5 min and 19.7 min for a 30-satellite CGT constellation, respectively. This study also compares the revisit times of the proposed method with those of a traditional Walker-Delta constellation under three inclination conditions: optimal inclination, restricted inclination by launch trajectories from the Korean Peninsula, and inclination for the sun-synchronous orbit. A comparison showed that the CGT constellation had the shortest revisit times with a non-optimal inclination condition. The results of this analysis can serve as a reference for determining the appropriate constellation pattern for a given inclination condition.

• Journal of Aerospace System Engineering
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• v.16 no.5
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• pp.86-93
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• 2022

Generally, a deployable solar panel is used primarily to achieve sufficient power output to perform the mission. However, temperature distribution on the antenna reflector may increase due to the shading effect induced by the presence of the deployable solar panels. Appropriate thermal design is critical to minimize the thermal deformation of the mesh antenna reflector in harsh on-orbit thermal environments to ensure remote frequency (RF) performance. In this paper, we proposed a dual-surface primary reflector consisting of a mesh antenna and a flexible fabric membrane sheet. This design strategy can contribute to thermal stabilization by using a flexible solar panel on the rear side of membrane sheet to reduce the temperature distribution caused by the deployable solar panel. The effectiveness of the mesh antenna design strategy investigates through on-orbit thermal analysis.