• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.11
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• pp.80-87
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• 2005

T-50 supersonic jet trainer aircraft using digital flight-by-wire flight control system receives aircraft flight conditions such as altitude, VCAS(Calibrated Airspeed) and Angle of Attack from IMFP(Integrated Multi-Function Probe). IMFP sensors information have triplex structure using three IMFP sensors. Air-data selection logic is mid-value selection in three information from three IMFP sensors in order to have more reliability. From supersonic flight test at high altitude, air-data information is dropped simultaneously because of supersonic shock wave effect. This error information may affect to aircraft stability and safety in supersonic area at low altitude. This paper propose that sensitivity analysis and HQS(Handling Quality Simulator) pilot simulation in order to analyze flight stability and controllability in supersonic area at low altitude when these information is applied to flight control law.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.2
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• pp.195-199
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• 2010

(255,223) RS(Reed-Solomon) code which is the CCSDS(Consultative Committee for Space Data Systems) standard was used in the STSAT-3 to correct errors during the downlink of payload data. The RS encoder developed by VHDL was implemented in MMU(Mass Memory Unit). Moreover, the RS decoder developed by C-language was implemented in the DRS(Data Receiving System) of ground station. In this paper, we reported the design and analysis results of RS(255,223) for STSAT-3. The BER(Bit Error Rate) performance from MMU to DRS was confirmed through the downlink test at 16 Mbps. Also, the error correction performance and capability of RS(255,223) was tested by the manual attenuation of the RF(Radio Frequency) signal in the X-band transmitter resulting in putting some errors in the communication line.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.9
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• pp.802-808
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• 2014

MEMS based solid propellant thruster researched for the purpose of an academic research will be verified at space environment through CubeSat program. For this, the temperature of the MEMS thruster should be within allowable operating temperature range by proper thermal control to prevent the ignition failure caused by ignition time delay and to guarantee the structural safety of the MEMS thruster in the low temperature. In this study, we proposed an effective thermal control strategy, that is to use micro-igniter as a heater and temperature sensor for active thermal control instead of using additional heater. The effectiveness of the strategy has been verified through on-orbit thermal analysis of CubeSats with MEMS thruster.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.4
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• pp.309-318
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• 2019

The thermal design of the Lunar Terrain Imager (LUTI) on the Korean Pathfinder Lunar Orbiter (KPLO) was performed and the soundness of the thermal design was verified by thermal analysis. The thermal environment of the lunar mission orbit should be reflected in the thermal design because the IR radiation of the lunar surface is important, unlike the earth orbit. The components or modules exposed to the outside of the satellite are insulated with MLI as much as possible, but the camera tube and the radiator are functionally exposed, so the thermal shield using the concept of radiation shape factor is mounted on the front to mitigate IR radiation. The IR emissivity is important in the front side of the radiator that receives little solar radiation, and components that are susceptible to thermal deformation such as the tube use a radiation heater to minimize the temperature gradient. Through the investigation of computational results, it was confirmed that the thermal design of LUTI is stable in various situations.

• The Journal of Korean Institute of Electromagnetic Engineering and Science
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• v.30 no.5
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• pp.339-347
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• 2019

The power circuit design of an on-board S-band transmitter for KSLV-II with derating(operation of a component at less than its maximum rated specification to enhance reliability) is investigated. The power circuit of the transmitter consists of linear voltage regulators, DC/DC converters for regulating the DC supply, and diodes for reverse voltage protection. After analyzing the load current of the components, derating requirements are explored. Furthermore, power dissipation and junction temperature rise are considered with respect to the load current. The analysis is compared to the results from an engineering model of the transmitter. The temperature of the components is derated by >$40^{\circ}C$ in an environment where the ambient temperature is $+60^{\circ}C$, which is the acceptance test specification of high temperature.

• 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.42 no.7
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• pp.586-593
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• 2014

The CNUSAIL-1 project aims to develop and operate a 3U-sized cube satellite with solar sail mechanism. The primary mission is to successfully deploy the solar sail in a low earth orbit, and the secondary mission is to collect the scientific data for the effect of the solar sail deployment and operation on orbit maneuver and attitude change of the cube satellite. For this, the bus system will collect and transmit the dynamic data of the satellite and the visual images of the solar sail operation. This paper describes solar sail mission and conceptual design of CNUSAIL-1. The actuation/operation of the solar sail and the bus system are preliminarily designed in terms of attitude control system, communication system, electrical power system, command and data handling system, structure and thermal control system is designed.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.171-181
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• 2003

The nose fairings of KSR-III are designed to be separated from the rocket by explosive force at the mission altitude to expose the payload. Adequate amount of separation force should be imposed to allow safe separation without collision between the fairings and the rocket, and the separation device was designed for the separation at very high altitude where almost no air load was expected. As the development of KSR-III goes on, several design changes have made and lower separation altitude of 45km is expected as a result. Under these circumstances, it is required to determine if the nose fairings can be separated without collision with much severer air load than for the design condition. In this study, the 6-DOF motion analysis program, PASEM, which was developed to predict the strap-on booster separation, is modified to simulate the pivotal motion of the fairings at early stages of separation. The accuracy of pivot motion simulation is validated by comparison with the results of ground test and the accurate separation conditions are deduced from it. Trajectory simulations are performed to see if separation without collision is possible with varying angle of attack, direction of gravity, and the effect of gust. It is also found that reducing the separation angle of the clamshell hinge from 60 degrees to 40 degrees can enhance separation safety and separation at lower altitude of 40km can be done without collision.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.9
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• pp.827-832
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• 2012

In this paper, we will design the speed measurement board of LEO Satellite's reaction wheel which has two speed measuring methods as M-Method type and T-Method type. therefore we can use the advantage of two methods. and we will verify the availability of design on the on-board computer at the real LEO Satellite(KOMPSAT-3). In the reaction wheels satellite that can change the satellite's attitude is one of the leading drivers by the rotational inertia of the motor will perform attitude control. Reaction methods for detecting wheel rotation speed generated during a certain period T internal reaction wheel tacho pulse counting M-Method to detect wheel speed and wheel tacho pulses are generated by measuring the time between the detection rate can be divided into T-Method. M-method is simple to implement and benefit measurement time is constant, but slow fall in the velocity measurement accuracy is a disadvantage. In contrast, the time between tacho pulses to measure the T-Method to measure the precise speed at low speed and to measure the time delay is small, has the advantage. However, this method also in the actual implementation and the complexity of the operation at different speeds depending on the speed of operation has the disadvantage.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.36 no.11
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• pp.1399-1404
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• 2012

Because satellites operate in space, it is impossible to repair them when they malfunction. Therefore, to ensure the normal function of the payload used in the satellites, accurate assembly and installation of parts are crucial. To prevent abnormal functioning in the extreme environments during launch and in space, it is essential to test changes at the parts and system levels by performing alignment measurement before and after the launch environment test and the space environment test. Recently, noncontact three-dimensional precision machinery for medium- and large-sized parts has been developed. One of these is the theodolite measurement system, which is widely used in the aerospace industry. This study measures the angle of the dual thruster module that is used to control the attitude of KOMPSAT by using a theodolite, and alignment measurement and a reliability analysis are performed.