• Journal of Aerospace System Engineering
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• v.15 no.1
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• pp.40-46
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• 2021

In this study, an aerodynamic configuration design and study on performance enhancement of a tilt-rotor UAV were conducted for improving mission capabilities compared to multi-copter type UAV, MC-1/2/3 developed for disaster and policing operation. To improve performance, a new TR5X configuration was developed by modifying the fuselage and tail shape of TR60 UAV and additionally attaching an extended wing to the nacelle. Aerodynamic performances of TR60 and TR5X were compared through computational fluid dynamics (CFD) analysis. In addition, flight performance analysis of full aircraft was conducted. Results showed that main performance requirements of TR5X were satisfied.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.77-77
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• 2002

초소형 공중발사체 설계 시 하이브리드 모터의 적용가능성에 대한 연구를 실시하였다. HTPB/LOX를 추진제로 하여 마차바퀴형 연료 그레인, 산화제 탱크 가압방식을 사용하였고, 성능특성을 계산하기 위하여 하이브리드 연료의 연소율이 일정하다고 가정 하였다. 본 연구에 사용된 임무는 중량 3.5kg의 나노위성을 근지점 고도 200km, 원지점 고도 1,500km의 타원궤도로 진입시키는 것을 목적으로 하는 로켓의 1단 부분에 관한 것으로 1단의 발사속도는 M=1.3, 발사고도는 12km, 연소종료 고도는 40km이다. 1단에 대한 페이로드 중량은 127.5kg이고, 속도증가분($\Delta$V)은 3,330m/s이다. 모선은 F-4E를 사용하였고 모선의 특성상 발사체의 총 중량이 1,000kg이하로 제한되고 길이와 직경이 5m${\times}$5m로 제한되나 1단에 대한 길이의 제한조건은 현재까지 명확히 정립되지 않은 상태이다. 설계과정에서의 변수는 연료 그레인 포트 개수, 초기 산화제 플럭스, 연소실 압력을 사용했고, 설계 제한조건은 추진제 중량, 평균 비추력, 평균 추력, 연소시간, 1단 길이, 직경, 연소시간이고, 이들의 범위는 모선의 특성과 초소형 공중발사체의 임무특성에 맞게 설정하였다.

• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.55-64
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• 2002

A conceptual design of propulsion system for a geosynchronous communication satellite with 12 years design life is presented in this paper. Propellant mass budget for the design life is calculated using total velocity increment (ΔV) flowed-down from mission requirement analysis. Sizes of the fuel and oxidizer tank are derived based on the calculated propellant mass budget, and mass of the pressurant as well as the size and pressure of pressurant tank are calculated too. Thruster positioning, number of rocket engines, and position of tank are determined through Trade-Off Study with Structure & Mechanical Subsystem. Propulsion system configuration and its schematics are presented finally.

• Journal of the Korea Society for Simulation
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• v.20 no.2
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• pp.59-66
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• 2011

In this paper, a method to allocate a submarine search mission to an ASW(Anti-Submarine Warfare) helicopter is proposed. The aim of the proposed method is to increase the submarine detection capability. For this purpose, we modeled the behaviors that the ASW helicopter conduct during the search mission, and the relations between the behaviors are also modeled. To measure quantitatively the effectiveness of ASW search mission, the measure of effectiveness(MOP) is defined. Scenarios are designed to analyze the effectiveness utilizing the ASW mission model. We conducted simulations applying the designed scenarios and some parameters concerned with the friendly ship and the enemy submarine interacting each other in the ASW missions. We analyzed the result of simulation depending on the dipping interval and the pattern of dipping positions in the situation that the helicopter operates for a long time and should resupply several times on the friendly ship. From the analyzed data, we suggested the practical value of ratio between the detectable range of the sonar and the dipping interval to improve the effectiveness of ASW mission.

• Aerospace Engineering and Technology
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• v.9 no.2
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• pp.176-184
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• 2010

The power simulator for a LEO satellite is a useful tool to analyze mission validity and energy balance for various mission scenarios by estimating power generation, power consumption, depth of discharge, bus voltage, charging/discharging current, etc. In this paper, it is described the calculation algorithm of the solar array (SA) power, the satellite load power and the battery modeling method to develop a satellite power simulation. To simulate the SA power generation, three different operation modes (DET, MPPT, CV) of SAR (Solar Array Regulator) are considered with a SA model. The satellite load power is estimated using the satellite unit power database, the unit on/off configuration at some satellite operation modes. The bus voltage and battery charging/discharging current at the specific DoD (Depth of Discharge) are calculated based on the battery characteristics. By this satellite power simulator, it can be conveniently analyzed the energy balance and the validity of a planned mission of a LEO satellite.

• Journal of the military operations research society of Korea
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• v.37 no.1
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• pp.119-128
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• 2011

This paper deals with a study on mission analysis considering the effectiveness measurement of UAV system operations. This mission analysis process is composed of 5 steps; (1) creation of a base model in MANA, (2) design of input parameter set using experiment design, (3) mapping input parameter set to the MANA scenario file, (4) data farming and model run in batch process, and (5) statistical analysis of the simulation result. In the result of this study, the effect of input parameter to the dependent parameter was shown to decrease in the order classification range, sweep width, height, speed, FOV(Field of view), and classification probability. The study also shows that the operational effectiveness of an improved scenario proposed can increase 10.2% from the base scenario.

• Bulletin of the Korean Space Science Society
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• 2009.10a
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• pp.40.1-40.1
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• 2009

미래 한국의 달궤도선 임무에 대비하여 달 근접 궤도 전파기인(orbit propagator) YSPLOP ver. 1(Yonsei Lunar Precise Orbit Propagator version 1)을 개발 하였다. 개발된 궤도 전파기의 성능은 상용 소프트웨어인 STK Astrogator를 이용하여 검증되었다. 개발된 궤도 전파기를 이용, 달 궤도선의 운용에 있어서 다양한 섭동력들이 궤도선의 수명(orbital decay)에 미치는 영향을 분석하였다. YSPLOP ver. 1은 정밀한 달 중심 탐사선의 위치산출을 위하여 M-EME2000 (Moon-Centered, Earth Mean Equator and Equinox of J2000) 좌표계, M-MME2000 (Moon-Centered, Moon Mean Equator and IAU vector of epoch J2000) 좌표계 그리고 M-MEPMD (Moon-Centered, Moon Mean Equator and Prime Meridian) 좌표계를 이용하여 탐사선의 상태(state) 정보를 산출한다. 또한 태양, 지구, 달, 화성, 목성의 중력에 의한 섭동력 및 태양풍에 의한 영향을 포함할 수 있도록 설계되었으며, 달 근접 궤도선의 궤도 운동에 가장 큰 영향을 미칠 수 있는 섭동력인 달의 비대칭 중력장에 의한 영향 또한 고려하도록 하였다. 달의 비대칭 중력장 모델 (Lunipotential model)은 LP165p 모델이 사용되었으며 행성의 정밀한 위치 산출을 위하여 JPL의 DE405 천체력이 사용되었다. 개발된 궤도 전파기를 이용, 달고도 100 km, 궤도 경사각 $90^{\circ}$인 달 중심의 극궤도를 약 30일 동안 전파한 결과, YSPLOP ver. 1의 성능은 STK Astrogator와 비교하여 보았을 때 약 수 m의 오차를 보이는 것으로 확인되었다. 달의 극궤도 탐사선의 궤도 수명을 분석한 결과, 최소한 달의 비대칭 중력장이 70 by 70 이상으로 고려되어야 함을 확인하였으며 이때 달 궤도선의 수명은 약 160일으로 나타났다. 아울러 달 근접 환경에서의 지구 중력에 의한 섭동력은 달 궤도선의 운동에 있어서 무시 할 수 없는 정도의 많은 영향을 끼치고 있음을 확인하였다. 이 연구를 통하여 개발된 궤도 전파기는 미래 한국의 달 궤도선 및 착륙선의 임무 설계시 사용 될 수 있다. 또한 이 연구에서 제시된 달 근접 환경에서의 다양한 섭동력들이 달 궤도선의 운동에 미치는 영향에 대한 해석 결과는 추후 달 근접 임무 설계시 고려되어야 하는 섭동력들의 기본 사양을 제공할 것이다.

• Journal of the Korea Institute of Information and Communication Engineering
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• v.25 no.2
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• pp.320-326
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• 2021

AVTE(Air Vehicle Test Equipment) is an equipment to inspect and check the status of on-board aircraft LRUs(Line Replacement Units) before and after flight for performing successful UAV(Unmanned Aerial Vehicle) missions. This paper suggests utilization of the AVTE as an operation support-equipment by implementing several critical functions for UAV-operation on the AVTE. The AVTE easily sets initialization(default) data and compensates for the installation and position errors of the LRUs which provide critical mission data and situation image with pilots without additional individual operation support-equipment. Major fault list and situation image data could be downloaded after flight using the AVTE in the event of UAV emergency situation or unusual occurrence on duty as well. We anticipate the suggested operational approach of the AVTE could dramatically reduce the cost and man power for design and manufacture of additional operation support equipment and effectively diminish workload of the operator.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.1
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• pp.93-103
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• 2005

SEDT(System Engineering Design Tool) has been developed for small satellite conceptual design with an aim to verifying the nanosatellite HAUSAT-2 design. The program can calculate the mass and power of whole satellite system having specific mission and estimate the system cost based on mission and user requirements. It is containing various analysis data of more than 200 small satellites. The database will provide the trend analysis results of the small satellites which will become important design factors. This tool has also been verified by applying more than 10 small satellite data through case studies.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.629-634
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• 2011

A prestudy on expendable turbine engine for high-speed vehicle was conducted. The two possible mission profiles were established to decide the engine requirements and Design Point, and Design Point analysis was performed with the values of design parameter which were obtained from similar class engines and technical references. The results showed that Specific Net Thrust is 2599.4 ft/s and Specific Fuel Consumption is 1.483 lb/($lb^*h$) at the flight condition of Sea Level, Mach 1.2. It was also found through the performance analysis on the two possible mission profiles that major design parameters for determining Net Thrust were Turbine Inlet Temperature for low supersonic flight speed and Compressor Exit Temperature for high supersonic flight speed. In addition, simple turbojet engine with axial compressor, straight annular combustor, axial turbine and fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost light engine.