Earticle

본 논문의 목적은 대함유도무기 비행시험의 업무 효율 증대를 위한 프레임워크 구조 정 의와 프레임워크 개발 결과를 제시하는 것이다. 우선 프레임워크 구조 정의를 위하여 프레 임워크 구성요소 식별, 영역 분류, 단계 구분 및 프레임워크 템플릿 개발을 수행하였다. 프 레임워크의 구성요소는 표준 프로세스, 표준 산출물, 이슈사항 체크리스트로 구성되어 있 다. 프레임워크의 영역 분류는 체계수행, 시험안전, 시험지원으로 나누어져 있다. 프레임워 크의 단계 구분은 비행시험 계획, 준비, 수행, 후속으로 이루어져 있다. 프레임워크 개발을 위한 프레임워크 템플릿과 ID 식별코드 체계를 함께 제시하였다. 프레임워크 구조 정의를 한 후 대함유도무기 비행시험을 위한 총 15개 프레임워크가 개발되었다. 개발된 프레임워 크에 대한 평가 설문을 실시하여 프레임워크의 실용성을 입증하였다. 대함유도무기 비행시 험에 적용코자 개발된 프레임워크는 표준화된 업무 정립을 통해 대함유도무기 체계개발 성 과 창출에 기여할 것으로 판단된다.

The purpose of this paper is to develop the framework for increasing efficiency of the business processes for anti-ship missile flight test. First of all, for establishing the framework structure definition, it is performed framework components identification, region classification, step classification and framework template. The Frame components include standard processes, standard outputs and issue checklists. Framework regions are divided into system work region, test safe region and test support region. Framework steps consist of flight test plan, flight test readiness, flight test execution and flight test follow-up step. Framework template and ID classification schemes are proposed. Since the framework structure definition, the total of 15 frameworks for anti-ship missile flight test are developed. The evaluation of the developed framework is conducted to demonstrate the practicality of the framework. It is believed that the proposed framework will contribute to anti-ship missile development through the standardized processes.

항공탑재시험은 복잡해지는 무기체계 개발과정에서 데이터 획득을 위해 수행하는 중요한 시험 중 하나이다. 본 논문은 무기체계 개발 과정중 수행하는 항공탑재시험용 POD 시스템에 대한 설계 내용 및 시험 결과를 소개한다. 항공탑재 시험용 POD는 좌·우 2개의 POD가 설계 및 제작되었고 각각의 POD는 항공기 연료탱크와 동일한 외형과 질량특성을 갖도 록 하여 감항인증 절차를 생략할 수 있도록 하였다. 또한 무기체계 개발에 필요한 표적 영상 데이터 계측, 항법 데이터 획득, 알고리듬 검증 및 분석에 필요한 기준 데이터를 획득할 수 있도록 구성품들을 적절하게 배치하였다. 항공탑재시험용 POD 시스템은 기계적, 전기적 요소들이 모두 반영된 복합적인 시스템이며 개발된 POD 시스템은 반복적으로 항공탑재시험에 사용되어 무기체계 개발에 필요한 다양한 데이터를 성공적으로 획득하였다.

Captive Flight Test (CFT) is one of the most important tests to acquire data when developing complex weapon systems. In this paper, we introduce the design and test result of our POD system for CFT. POD system uses POD set which consists of left and right POD. The exterior and mass properties of POD set are equal to those of fuel tank for aircraft so that we can omit Airworthiness Certification. Also, we adequately placed inner-equipments in order to acquire data including target image, navigation result and reference data to verify and analyse software algorithm. The POD system for CFT we developed is complex system as both mechanical and electronic factors are applied. As we repeatedly performed CFT, useful and various data for weapon development were acquired.

본 논문에서는 다수 레이더가 발사체를 탐지 및 추적하는 시험 상황 시스템에서 발사체의 예상 위치 추정을 위한 실시간 시험자료 처리 기법을 다루고자 한다. 다수 레이더들은 각각의 고유한 계측특성을 가지고 있으며, 동일한 발사체에 대하여 각기 다른 위치정보를 계측한다. 다수 레이더의 계측값에 대한 신뢰도를 향상시키기 위하여 다수 레이더가 측정한 동일 발사체에 대한 서로 다른 위치정보를 융합하여 통합된 좌표를 공유할 수 있는 기법이 필요하다. 이 논문에서는 다수 레이더에서 계측한 발사체 정보를 수집하는데 필요한 송신 시간 및 계산 지연을 고려하고 특정시간 이후의 발사체에 대한 정확한 예상 좌표 추정을 계산하기 위하여 칼만필터 알고리즘, 선형예측 알고리즘 및 이동평균 알고리즘을 적용한다. 또한, 동일 발사체에 대한 다수 레이더의 위치정보를 융합하기 위하여 각 레이더가 송신한 발사체의 위치 정보 중에서 유효하지 않은 위치 정보를 잡음으로 정의하고, 제거하는 기법을 제안한다. 다수레이더가 계측한 시험자료 중에서 잡음에 해당하는 좌표를 제거하고, 위치좌표를 융합한 후, 일정 시간 이후의 통합좌표 예측 정확도에 대한 성능을 평가하였으며, 세 가지 위치좌표 추정 알고리즘을 사용하여 예상한 위치 정보와 원격측정장비(telemetry) 위치정보 사이의 오차를 비교 분석하였다.

This paper addresses a real-time flight test data processing algorithm for the estimation of the coordinate position of projectiles sensed by multiple radars. Since multiple radars have their own features to search and track targets, they sensor the targets in different results of their estimation. In order to improve the reliability of estimation by multiple radars, multiple radar sites need to share an integrated coordinate position. We thus apply Kalman filter algorithm, linear regression algorithm and moving average algorithm to the estimation of expected future position of a projectile, as considering communication and computation delay. We also define invalidated information, in case that any radar inappropriately detect the coordinate position of the projectile, and suggest a method to exclude the invalidated positions of the projectile. To evaluate algorithms we suggest, in an experiment, we measure and analyze the error between estimated position of the projectile using three algorithms and its position by telemetry.

An airborne radar is an essential aviation electronic system for the aircraft to perform various civil and/or military missions in all weather environments. This paper presents the design, development, and test results of the multi-mode X-band pulsed Doppler radar system test model for helicopter-borne flight test. This radar system consists of antenna unit, transmitter and receiver unit, radar signal & data processing unit and display unit. The developed core technologies include the planar array antenna, TWTA transmitter, coherent I/Q detector, digital pulse compression, MTI, DSP based Doppler FFT filter, adaptive CFAR capability. The design performance of the developed radar system is verified through various ground fixed and moving vehicle test as well as helicopter-borne field tests including MTD (Moving Target Detector) capability for the Doppler compensation due to the moving platform motion.

본 연구의 목적은 파편분산 영역산출을 위한 M&S 방법과 해상 안전을 위한 선박소개 프로세스를 제시하고, 해상 플랫폼 위치 선정 시 고려요소를 도출하는데 있다. 미사일 관련 국방 연구개발 사업에서 비행시험은 필수적인 요소이며, 해상 안전영역을 보장하기 위해서는 우선적으로 M&S 기법이나 수식에 의한 방법으로 정확하게 파편 분산영역 을 예측하는 것이 필요하다. 해상에서 미사일 비행시험 시 민간 선박이나 상선의 피해를 방지하기 위해서는 요격탄이나 표적시스템을 발사하기 전에 사전 선박소개가 이루어져야 한다. 하지만 복잡한 선박소개가 시스템적인 절차에 의한 방법이 아닌 과거 경험을 통해 이루어지는 경우가 연구개발 과정에서 많이 있었다. 그 동안 경험적 판단으로 이루어지고 있었던 선박소개 프로세스를 정립하고, 해상 플랫폼 해상에서 이용할 경우 플랫폼 위치를 선정하기 위해 전문가들의 의견을 Delphi 기법을 이용하여 고려요소들을 도출하였다. 특히, 해상 플랫폼 위치 선정 시 고려사항으로는 인전영역 확보, 계측장비 계측 가능성, 설치 용이성, 시험요원 접근 가능성, 어선 등 민간요소 배제가 중요한 것으로 도출되었다.

This research purpose is to suggest the ship dispersion process for obtaining maritime safety and consideration factors relating with determining the place of maritime launching platform in the missile flight test. A flight test is essential at military R&D stage related with the missile. Prior to securing the maritime safety zone, it is necessary to predict the debris area as accurately as possible through M&S tools or equations. Ship dispersion before launching interceptor or target should be carried out in order to prevent the damage of all vessels including civilian fishing boats. However, the ship dispersion has been based on experience in R&D institutions rather than specific procedures. Both the smooth cooperation with the related agencies and the accurate prediction of objective points and debris area based on M&S are required to achieve successful flight tests including ship dispersion. Consideration factors for selecting the place of the maritime launching platform were derived using the Delphi technique based on the opinions of experts. As the Delphi results, securing the flight safety area, the possibility that measuring equipments work, the ease of installation, the accessibility of the researchers, and the exclusion of civilian factors such as fishing boats were important.

국내외의 여러 경제 분석에서도 드론 시장이 날로 성장할 것이라는 예측을 하고 있으며, 그중에서 완구 및 취미용 드론 시장의 성장이 점진적으로 확대가 될 것으로 예측되고있다. 시장의 성장으로 발생하는 경제적 순기 능으로 드론의 기대가치는 높아져 가고 있으나, 관리 및 운용 미숙으로 발생하는 사고 또한 증가하고 있다. 완구용 드론의 성능 차이는 산업용 드론과 비교할 수 없을 정도로 낮지만, 일반 구매자는 그 성능의 여부를 알 수 없어 사용상에 사고를 유발할 수 있다. 본 연구에 사용된 완구용 드론은 KC와 CE 인증을 받고 국내에서 판매되고 있는 20종을 사용하였으며, 비행시간 및 조종 거리 측정에 사용되었다. 비행이 가능한 시간 범위는 최소 3분∼최대 12 분까지로 측정되었으며, 조종 가능한 거리는 최소 20m∼최대 약 380m까지로 측정되었다. 따라서 완구용 드론의 전파출력에 대한 샘플링 검사를 통한 제품 안전성 확보가 필요하며, 한계 비행거리를 넘었을 때는 자동으로 고도를 낮추거나 호버링을 하는 알고리즘 탑재가 필요하다. 향후 연구 방향으로는 완구용 드론의 실제 고도 테스트와 충격 량 테스트를 통한 완구용 드론의 안전기준 제정을 위한 자료 구축을 할 예정이다.

Economic analysis predicts that the drone market will grow, and the growth of the toy and hobby drone market is expected to gradually expand. Drone expectations are rising due to the net economic function of drone market growth, but accidents due to improper management and operations are also increasing. The difference in toy drone performance is incomparably small compared to industrial drone performance, but the ordinary buyer can not know whether the difference can cause an accident during use. The toy drones used in this study were obtained from KC and CE certification, and 20 kinds of drones were used. The flight time ranged from a minimum of 3 minutes to a maximum of 12 minutes, and the control distance ranged from a minimum of 20m to a maximum of 380m. Therefore, it is necessary to secure product safety through sampling inspection of the radio wave output of toy drones, and it is also necessary to mount an algorithm that automatically lowers the altitude or hover when exceeding the limit flight distance. For future research, we will build data to establish toy drone safety standards through a altitude testing and impact testing of toy drone.

본 연구에서는 모듈형 PAV(Personal Air Vehicle)의 설계 및 Prototype 제작을 통해 기존 PAV의 활용성과 확장성에 대한 가능성을 확인하고자 한다. 모듈형 PAV는 비행을 담당하는 Aerial Module, 결합을 위한 Connector Module, 지상에서 구동되는 Ground Module, 그리고 운송용 Transport Module로 구성된다. Aerial Module 하단에 부착된 Connector Module을 통해 Transport Module이 탈부착 되는 방식이다. 실험 진행은 실제 비행이 가능한 1:8.5 비율의 소형 Prototype을 설계한 후 제작하여 비행 테스트를 했다. 비행 로그 분석 결과 기체는 Roll, Pitch, Yaw 각도에서 안정적인 자세 제어를 보였다. 또한 진동과 가속도 수치도 허용범위 내에 있는 것을 확인했다. 이를 통해 탈부착이 가능한 모듈형 PAV의 실현 가능성을 확인했다.

This study aims to evaluate the usability and scalability potential of modular PAV by designing and prototyping a modular PAV system. The modular PAV consists of an Aerial Module for flight, a Connector Module for combination, a Ground Module that operates on the ground, and a Transport Module for transportation. The Transport Module can be detached and attached via the Connector Module attached to the bottom of the Aerial Module. For the experiment, a small prototype with a scale of 1:8.5 capable of actual flight was designed, manufactured, and flight tested. The flight log analysis results showed that the airframe had stable attitude control in the Roll, Pitch, and Yaw angles. In addition, it was confirmed that the vibration and acceleration values w ere within the acceptable range. Through this, we confirmed the feasibility of a modular PAV that can be detached.

EOTS for drones is showing another aspect of market expansion in detection and recognition areas previously occupied by artificial satellites. The two-axis EOTS for drones controls the vibration or disturbance caused by the drone during the mission so that EOTS can accurately recognize the goal. Vibration generated by drones is transmitted to EOTS. Therefore, it is essential to develop a stabilization controller that attenuates vibrations transmitted from drones so that EOTS can maintain the viewing angle. Therefore, it is necessary to standardize drone disturbance and secure the performance of EOTS disturbance attenuation controller optimized for disturbance level through this. In this paper, a method of standardizing drone disturbance applied to EOTS is studied, through which EOTS controller simulation is performed and stabilization controller shape is selected and designed.

The full results of troubleshooting process related to the flight control system of a tilt-rotor type UAV in the flight tests are described. Flight tests were conducted in helicopter, conversion, and airplane modes. The vehicle was flown using automatic functions, which include speed-hold, altitude-hold, heading-hold, guidance modes, as well as automatic take-off and landing. Many unexpected problems occurred during the envelope expansion tests which were mostly under those automatic functions. The anomalies in helicopter mode include vortex ring state (VRS), long delay in the automatic take-off, and the initial overshoot in the automatic landing. In contrast, the anomalies in conversion mode are untrimmed AOS oscillation and the calibration errors of the air data sensors. The problems of low damping in rotor speed and roll rate responses are found in airplane mode. Once all of the known problems had been solved, the vehicle in airplane mode gradually reached the maximum design speed of 440km/h at the operation altitude of 3km. This paper also presents a comprehensive detailing of the control systems of the tilt-rotor unmanned air vehicle (UAV).

Flight test of flapping micro air vehicles (FMAVs) is carried out using an instrumented measurement system to obtain various engineering parameters and hence to assess the flight performance of the vehicles through the data investigation. An indoor flight test facility equipped with a motion capture system and tracking cameras is used for the work presented in this paper. Maneuvers including straight-level flight, ground flapping, takeoff and landing are tested. Spatial position and orientation data are obtained from the retro-reflective tracking markers attached to the vehicles. Subsequent test analysis is carried out by generating performance parameters from raw data and then assessing the flight performance by comparison of the vehicles. The main findings of this work confirm that the test method and procedures presented here enable the systematic numerical data measurement and assessment of the flying performances of these vehicles, and show the applicability for the test and evaluation of general flapping MAVs.

최근 수년간, GPS와 관련된 연구들이 항공분야를 비롯한 다양한 분야에서 진행되었다. 특히, GPS를 항공기의 착륙 유도 시스템에 이용하려는 연구가 많이 이뤄지고 있다. 이러한 시도들은 GPS의 경제성, 신뢰성, 정확성 등의 장점들을 십분 활용하기 위한 것이라고 할 수 있다. 서울대학교 GPS 실험실에서도 이러한 경향에 보조를 맞추어, GPS를 기반으로 하는 항공기 착륙 시스템을 개발하고, 헬리콥터를 이용한 비행실험을 수행하고 있다. 그 동안 누적된 실시간 DGPS 시스템 개발기술들을 바탕으로 항공기 착륙 시스템을 확장, 보강하여 최근의 비행실험을 실시하였다. 본 논문에서는 새롭게 구성된 항공기 착륙 시스템을 소개하고 이를 이용한 비행실험 결과를 분석하였다. 기존의, 기본적인 실시간 DGPS 시스템에서 추가, 발전된 부분은 세 가지로 분류할 수 있다. 첫 번째는, 단일 GPS 안테나를 이용하여 항공기의 자세를 추정하는 부분이고, 두 번째는, 통합적인 cockpit display이다. 이 display는 가상현실을 이용하여 조종사에게 기존의 ILS 정보와 그 이외의 다양한 정보들을 보여준다. 마지막으로, 공항의 기상상태에 관계없이 조종사가 공항에 접근할 수 있도록, 전자지도를 삽입하여 안전한 착륙을 시도할 수 있도록 시스템을 구성하였다. 이렇게 새롭게 구성된 시스템을 이용하여 김해 국제공항에서 비행실험을 수행하였다. 분석된 결과를 바탕으로, 이 시스템이 정확도 측면에서, CAT-I을 충분히 만족시킴을 확인하였으며, 신뢰도 높은 자세결정이 이뤄지고 있음을 확인하였다.

In recent, there has been remarkable progress in the field of GPS applications. In a few years, an appreciable number of aircraft will adopt GPS as a landing guidance system because GPS is more economic, more reliable and more accurate than any other aviation systems. In this respect, we have performed several helicopter landing flight tests based on the real-time DGPS system made in SNUGL (Seoul National University GPS Laboratory). From the experimental results, we found several problems Which should be fixed to adopt DGPS as a aircraft landing guidance system. In this paper, we will introduce the problems found in tests and also suggest modifications to solve the problems. Our modifications can be classified into three parts. The first is about the attitude determination with single GPS antenna. The second deals with the cockpit display module. The display was devised to integrate the Instrument Landing System(ILS) with tunnel-the-sky using virtual reality. With the display, pilot can achieve more safe landings. The last part is the digital map. We inserted digital map into our system and put direction indicator on the map using position information from GPS. It is very useful for pilot to find airports even in bad weather. Using the newly designed DGPS landing system, we conducted flight test at Kimhae International Airport, Pusan, Korea. It was successful! Our system can also satisfy Category-I criterion for aircraft landing approach and determine attitude angle with a high level of reliability. It is supported by video materials.

Various R&D programs for rocket-based combined cycle (RBCC) engines have progressed worldwide for the space development and the defense applications. As a way toward indigenous domestic RBCC program, a preliminary design of flight test conditions was studied in this study for a sub-scale RBCC engine using a sounding rocket. Launch and flight profiles were calculated for several booster options and compared with that of HyShot II program. The result shows that the Korea Sounding Rocket-II (KSR-II) is a proper candidate to perform the flight test available in Korea. The recommend flight test conditions with KSR-II are Mach 6.0 with a test vehicle of 230 kg and Mach 7.4 with 50 kg. Present study will soon be followed by a design of sub-scale RBCC for a flight test using a sounding rocket.

A numerical prediction of the transient fuel temperature in an aircraft was made and verified with a flight test. The analysis was studied with the finite difference method. Numerical calculation was performed by an explicit method of the modified Dufort-Frankel scheme. Convective heat transfer coefficients were used in calculating heat transfer between the aircraft surface and the ambient air. For an aircraft on the ground, an empirical equation represented as a function of free-stream air velocity was used. And, the heat transfer coefficient for flat plate turbulent flow suggested by Eckert was employed for in-flight phases. The governing equations used in this analysis are the mass and energy conservation equations on fuel and oils. The analysis was verified with the flight test data of a fuel system with additional fuel supplies and return concept. As a result of the verification, the difference of the fuel temperatures obtained by the analysis from those of the flight test data was relatively small with a tendency to increase in the later phases of the flight.

This paper presents a distribution simulation model for dual purpose improved conventional munitions based on flight test data. A systematic procedure for designing a dispersion simulation model is proposed. A new accumulated broken line graph was suggested for designing the distribution shape. In the process of verification and simulation for the distribution simulation model, verification was performed by first comparing data with firing test results, and an application simulation was then conducted. The Monte Carlo method was used in the simulations, which reflected the relationship between ejection conditions and real distribution data. Before establishing the simulation algorithm, the dominant ejection parameter of the submunitions was examined. The relationships between ejection conditions and distribution results were investigated. Five key distribution parameters were analyzed with respect to the ejection conditions. They reflect the characteristics of clustered particle dynamics and aerodynamics.

Nowadays, many kinds of research for airship have being studied withincrease of interests in airship. But these make little progress even now. In flighttest, the data acquisition from the actual flight test has lots of difficulties becausethe airship dynamic response is slow and sensitive to extemal environment. In thispaper, through the actual flight test, appropriateness of the mathematical dynamicmodel was presented by showing the test results in various conditions. Thetuming, the acceleration, and the deceleration motions were tested and analyzed.

High Altitude and Long Endurance (HALE) aircraft are capable of providing intelligence, surveillance and reconnaissance (ISR) capabilities over vast geographic areas when equipped with advanced sensor packages. As their use becomes more widespread, the demand for additional range, endurance and payload capability will increase and designers are exploring non-conventional configurations to meet the increasing demands. One such configuration is the joined-wing concept. A joined-wing aircraft is one that typically connects a front and aft wings in a diamond shaped planform. One such example is the Boeing SensorCraft configuration. While the joined-wing configuration offers potential benefits regarding aerodynamic efficiency, structural weight, and sensing capabilities, structural design requires careful consideration of elastic buckling resulting from the aft wing supporting, in compression, part of the forward wing structural loading. It has been shown already that this is a nonlinear phenomenon, involving geometric nonlinearities and follower forces that tend to flatten the entire configuration, leading to structural overload due to the loss of the aft wing's ability to support the forward wing load. Severe gusts are likely to be the critical design condition, with flight control system interaction in the form of Gust Load Alleviation (GLA) playing a key role in minimizing the structural loads. The University of Victoria Center for Aerospace Research (UVic-CfAR) has built a 3-meter span scaled and flexible wing UAV based on the Boeing SensorCraft design. The goal is to validate the nonlinear structural behavior in flight. The main objective of this research work is to perform Ground Vibration Tests (GVT) to characterize the dynamic properties of the scaled flight vehicle. Results from the experimental tests are used to characterize the modal dynamics of the aircraft, and to validate the numerical models. The GVT results are an important step towards a safe flight test program.

Although computational-fluid-dynamic methods and wind-tunnel testing can provide data about the aerodynamic characteristics of an aircraft, the determination of these and other characteristics from flight data plays and important role. The object of this study is the verification of overall aircraft system performance to improve the stability of vehicle. We have test the Velocity-173, canard-type airplane to obtain the stability data. We adopt the two identifications method, EKF and MLE, for the parameter estimation. The results are compared with those of conventional type airplane.

Real-time monitoring is one of the prime necessities in a weapon flight test that is required for the efficient and timely collection of large amounts of high-rate sampled data acquired by an event-trigger. The wireless sensor network is a good candidate to resolve this requirement, especially considering the inhospitable environment of a weapon flight test. In this paper, we propose a priority based multi-channel MAC protocol with CSMA/CA over a single radio for a real-time monitoring of a weapon flight test. Multi-channel transmissions of nodes can improve the network performance in wireless sensor networks. Our proposed MAC protocol has two operation modes: Normal mode and Priority Mode. In the normal mode, the node exploits the normal CSMA/CA mechanism. In the priority mode, the node has one of three grades - Class A, B, and C. The node uses a different CSMA/CA mechanism according to its grade that is determined by a signal level. High grade nodes can exploit more channels and lower backoff exponents than low ones, which allow high grade nodes to obtain more transmission opportunities. In addition, it can guarantee successful transmission of important data generated by high grade nodes. Simulation results show that the proposed MAC exhibits excellent performance in an event-triggered real-time application.

Since 1990´s, there has been many researches for the development of the Unmanned Aerial Vehicle (UAV). Especially, for the development of digital electronics, the technologies of UAV toward to the miniaturization low-cost, and high reliability. Therefore, recent trends for the development of UAV are focused on the development modern Flight Control System (FCS). In this paper, focusing on the FCS, the development process for Sejong Unmanned Research Vehicle -1 (SURV-1) from design to flight test is presented.