• Journal of the Optical Society of Korea
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• v.17 no.6
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• pp.533-537
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• 2013

Space telescope optics is one of the major parts of any space mission used to observe astronomical targets or the Earth. This kind of space mission typically involves bulky and complex opto-mechanics with a long optical tube, but attempts have been made to observe a target with a small satellite. In this paper, we describe the optical design of a reflecting telescope for use in a CubeSat mission. For this design we adopt the off-axis segmented method for astronomical observation techniques based on a Ritchey-Chr$\acute{e}$tien type telescope. The primary mirror shape is a rectangle with dimensions of $8cm{\times}8cm$, and the secondary mirror has dimensions of $2.4cm{\times}4.1cm$. The focal ratio is 3 which can yield a 0.383 degree diagonal angle in a $1280{\times}800$ CMOS color image sensor with a pixel size of $3{\mu}m{\times}3{\mu}m$. This optical design can capture a ${\sim}4km{\times}{\sim}2.3km$ area of the earth's surface at 700 km altitude operation.

• The Bulletin of The Korean Astronomical Society
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• v.46 no.1
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• pp.38.3-38.3
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• 2021

We are developing an optomechanical design of infrared telescope for the CubeSat and Unmanned Aerial Vehicle (UAV) which adapts the Linear Astigmatism Free- Three Mirror System in the confocal off-axis condition. The small entrance pupil (diameter of 40 mm) and the fast telescope (f-number of 1.9) can survey large areas. The telescope structure consists of three mirror modules and a sensor module, which are assembled on the base frame. The mirror structure has duplex layers to minimize a surface deformation and physical size of a mirror mount. All the optomechanical parts and three freeform mirrors are made from the same material, i.e., aluminum 6061-T6. The Coefficient of Thermal Expansion matching single material structure makes the imaging performance to be independent of the thermal expansion. We investigated structural characteristics against external loads through Finite Element Analysis. We confirmed the mirror surface distortion by the gravity and screw tightening, and the overall contraction/expansion following the external temperature environment change (from -30℃ to +30℃).

• Journal of Aerospace System Engineering
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• v.15 no.5
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• pp.81-88
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• 2021

In general, a non-explosive nylon wire cutting-based holding and release mechanism has been used to store and deploy deployable solar panels of CubeSat. However, with this method, accessing the solar panel's access port for charging the cube satellite's battery and electrical inspection and testing of the PCB and payloads while the solar panel is in storage is difficult. Additionally, the mechanism must have a reliable release function in an in-orbit environment, and reusability for stow and deploy of the solar panel, which is a hassle for the operator and difficult to maintain a consistent nylon wire fastening process. In this study, we proposed a pin-puller-based solar panel holding and release mechanism that can easily deploy a solar panel without cutting nylon wires by separating constraining pins. The proposed mechanism's release function and performance were verified through a solar panel deployment test and a maximum separation load measurement test. Through this, we also verified the design feasibility and effectiveness of the pin-puller-based separation device.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.5
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• pp.432-439
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• 2015

MEMS solid thruster module is composed of solid thruster and its control board. It was developed for the purpose of an academic research. Therefore, thermo-mechanical design and verification for space usage were not considered in the design phase. To mount it on a cube satellite without any design modification, technical efforts at the system level structure design is required. In this study, we proposed a structural design concept to mount the MEMS thruster module by using brackets for guaranteeing structure safety under launch loads and easier mating and de-mating of MEMS thruster module during test phase. The effectiveness of the design has been verified through structural analysis and vibration test. In addition, electrical connection method using spring pins between MEMS thruster and control board is effective for guaranteeing the structural safety under launch vibration loads.

• Journal of Astronomy and Space Sciences
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• v.29 no.1
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• pp.33-40
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• 2012

Triplet Ionospheric Observatory (TRIO) CubeSatforIon, Neutral, Electron MAgneticfields (CINEMA) is a CubeSat with the weight 3 kg that will be operated in the orbit conditions of about 800 km altitude and $90^{\circ}$ inclination angle, using the S-band and ultra-high frequency (UHF)-band communication frequencies. Regarding the communication antenna loaded on the satellite, the two patch antennas has the downlink function in the S-band, whereas the two whip antennas has the function to receive the command sent by the ground station to the satellite in the UHF-band. The uplink ground station that communicates through the UHF-band with the CINEMA satellite was established at Kyung Hee University. The system is mainly composed of a terminal node controller, a transceiver, and a helical antenna. The gain of the helical antenna established at the Kyung Hee University ground station was 9.8 dBi. The output of the transceiver was set to be 5 W (6.9 dB) for the communication test. Through the far-field test of the established system, it was verified that the Roman characters, figures and symbols were converted into packets and transmitted to the satellite receiver in the communication speed of 9,600 bps.

• Journal of Aerospace System Engineering
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• v.17 no.4
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• pp.104-109
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• 2023

The Cube Satellite Contest has been held six times as of August 2023, and a total of 21 teams have been selected. Fifteen Cube Satellites selected in previous contests were successfully launched and entered into low-Earth orbit. The six Cube Satellites selected in the sixth contest in 2022 are currently undergoing detailed design, and are scheduled to be launched in 2025 using a Korean launch vehicle. In this study, we analyzed the initial operation reports submitted by the selected teams of the Cube Satellite Contest in 2012, 2013, 2015, 2017, and 2019 to assess mission performance and identify causes of mission failure. Based on the submitted reports, an improvement scheme to enhance mission success for future Cube Satellites is proposed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.8
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• pp.686-696
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• 2017

The major objective of this study is to develop and verify the KAUSAT-5 based on the modular 3U CubeSat standard platform. In the mechanical system design of a 3U standard platform, subsystem and micro equipment functions/performance should be integrated and miniaturized on micro-sized PCBs and electrical capability was maximized to accommodate multiple payloads. KAUSAT-5 is 3U-sized Cubesat which will be operated in Low Earth Orbit(LEO), which implements mainly two scientific missions; one is to observe the Earth through infrared camera and the other is to measure space radiation with a Geiger Muller tube. An additional mission is to verify the equipment(device) such as VSCMG and fuzzy logic-based MPPT internally developed. The results of ETB, qualification and acceptance level environmental tests were shown to verify standard platform and KAUSAT-5 Cubesat.

• Journal of Astronomy and Space Sciences
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• v.32 no.2
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• pp.145-149
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• 2015

A newly designed Tissue Equivalent Proportional Counter (TEPC) has been developed for the CubeSat mission, SIGMA (Scientific cubesat with Instruments for Global Magnetic field and rAdiation) to investigate space radiation. In order to test the performance of the TEPC, we have performed heavy ion beam experiments with the Heavy Ion Medical Accelerator in Chiba (HIMAC), Japan. In space, human cells can be exposed to complex radiation sources, such as X-ray, Gamma ray, energetic electrons, protons, neutrons and heavy charged particles in a huge range of energies. These generate much a larger range of Linear Energy Transfer (LET) than on the ground and cause unexpected effects on human cells. In order to measure a large range of LET, from 0.3 to $1,000keV/{\mu}m$, we developed a compact TEPC which measures ionized particles produced by collisions between radiation sources and tissue equivalent materials in the detector. By measuring LET spectra, we can easily derive the equivalent dose from the complicated space radiation field. In this HIMAC experiment, we successfully obtained the linearity response for the TEPC with Fe 500 MeV/u and C 290 MeV/u beams and demonstrated the performance of the active radiation detector.

• Journal of Aerospace System Engineering
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• v.14 no.3
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• pp.51-59
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• 2020

In the case of cubesat, a PCB-based deployable solar panel advantageous in terms of weight reduction and electrical circuit design is widely used considering the limited weight and volume of satellites. However, because of the low thermal conductivity of PCB, there is a limit relative to heat dissipation. In this paper, the thermal gap pad is applied to the contact between the PCB-based solar panel and the aluminum stiffener mounted on the outside of the panel. Thus, the heat transfer from the solar cell to the rear side of the panel is facilitated. It maximizes the heat dissipation performance while maintaining the merits of PCB panel, and thus, it is possible to improve the power generation efficiency from reducing the temperature of the solar cell. The effectiveness of the thermal design of the 6U cubesat's deployable solar panel using the thermal gap pad has been verified through on-orbit thermal analysis based on the results, compared with the conventional PCB-based solar panel.

• Journal of The Korean Astronomical Society
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• v.54 no.1
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• pp.9-16
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• 2021

The Ozone Dynamics Investigation Nano-Satellite (ODIN) is a CubeSat design proposed by Chungnam National University as contribution to the CubeSat Competition 2019 sponsored by the Korean Aerospace Research Institute (KARI). The main objectives of ODIN are (1) to observe the polar ozone column density (latitude range of 60° to 80° in both hemispheres) and (2) to investigate the chemical dynamics between stratospheric ozone and ozone depleting substances (ODSs) through spectroscopy of the terrestrial atmosphere. For the operation of ODIN, a highly efficient power system designed for the specific orbit is required. We present the conceptual structural design of ODIN and an analysis of power generation in a sun synchronous orbit (SSO) using two different configurations of 3U solar panels (a deployed model and a non-deployed model). The deployed solar panel model generates 189.7 W through one day which consists of 14 orbit cycles, while the non-deployed solar panel model generates 152.6 W. Both models generate enough power for ODIN and the calculation suggests that the deployed solar panel model can generate slightly more power than the non-deployed solar panel model in a single orbit cycle. We eventually selected the non-deployed solar panel model for our design because of its robustness against vibration during the launch sequence and the capability of stable power generation through a whole day cycle.