• 제어로봇시스템학회논문지
• /
• 제7권7호
• /
• pp.574-581
• /
• 2001

In this paper we design and evaluate the longitudinal nonlinear N(aub)z-CSAS(Command and Stability Augmentation System) flight control law in \"DMI(Dynamic Model Inversion)-method\" and classical \"Gain Scheduling-method\", respectively, to meet the handling quality requirements associated with push-over pull-up maneuver. It is told that the flight control law designed in \"DM-method\" is adequate to the full flight regime without gain scheduling and is efficient to produce the time response shape desired to the handling quality requirements. On the contrary, the flight control law designed in \"Gain Scheduling-method\" is easy to be implemented in flight control computer and insensitive to variation of the actuator model characteristics.n of the actuator model characteristics.

• 항공우주시스템공학회지
• /
• 제8권3호
• /
• pp.20-26
• /
• 2014

This paper presents the procedures and the results of the pitch control force via flight test for a light airplane in order to make out the stability of the aircraft and the compliance with concerned regulation. The flight test procedures were determined in order to obtain the aircraft type certification. The instrumentation equipments including airspeed indicator, accelerometer, and pitch control force measurement tools are used to perform the flight test. For the flight test, the airspeed and the pitch control force with related normal acceleration are measured sustaining turn flight with bank angle derived from trim speed. The flight test results showed that the handling qualities of the airplane are complied with the KAS-23, the regulation of the Korean government for the light airplane type certification.

• International Journal of Aeronautical and Space Sciences
• /
• 제16권3호
• /
• pp.425-436
• /
• 2015

This paper presents the procedures used for estimating the stability and control derivatives of a general aviation canard aircraft from flight data. The maximum likelihood estimation method which accounts for both process and measurement noise was used for the flight data analysis of a four seat canard aircraft, the Firefly. Without relying on the parameter estimation method, several aerodynamic derivatives were obtained by analyzing the steady state flight data. A wind tunnel test, a flight test of a 1/4 scaled remotely controlled model aircraft, and the prediction of aerodynamic coefficients using the USAF Stability and Control Digital Data Compendium (DATCOM), Advanced Aircraft Analysis (AAA), and Computer Fluid Dynamics (CFD) were performed during the development phase of the Firefly and the results were compared with flight determined derivatives of a full scaled flight prototype. A correlation between the results from each method could be used for the design of the canard aircraft as well as for building the aerodynamic database.

• 한국항공우주학회지
• /
• 제33권5호
• /
• pp.57-64
• /
• 2005

현대의 고성능 전투기에 탑재되어 있는 전기식 비행제어계통(Digital Fly-By-Wire Flight Control System)은 통합 다기능 감지기(IMFP : Integrated Multi-Function Probe)에 의해 항공기의 고도/속도/받음각 정보를 얻는다. T-50에 적용되어 있는 3개의 IMFP는 3중 결함 및 분리되지 않는 2중 결함에 대해서 비행 안정성(Flight Stability)을 확보하기 위해 형상 재구성 모드(Air Data Reconfiguration Mode)를 제어법칙에 적용했다. 본 논문에서는 항공기 운용 시 발생할 수 있는 IMFP 결함으로 인한 형상 재구성 모드 제어법칙에 대해, 비행 안정성을 해석하기 위하여 선형해석(Linear Analysis) 및 HQS( Handling Quality Simulator) 조종사 시뮬레이션을 수행하였고, T-50 비행시험 시, 발생했던 IMFP 결함으로 인해 제어법칙이 형상 재구성 모드로 적용되었던 사례를 제시했다. 그 결과, T-50 훈련기의 제어법칙이 형상 재구성 모드로 전환될 경우, 항공기 안정성에는 영향이 없다는 것을 알았다.

• Journal of Positioning, Navigation, and Timing
• /
• 제11권2호
• /
• pp.109-117
• /
• 2022

This paper proposes a method for detecting flight anomalies of drones through the difference between the command of flight controller (FC) and the navigation solution. If the drones make a flight normally, control errors generated by the difference between the desired control command of FC and the navigation solution should converge to zero. However, there is a risk of sudden change or divergence of control errors when the FC control feedback loop preset for the normal flight encounters interferences such as strong winds or navigation sensor abnormalities. In this paper, we propose the method with a deep neural network model that predicts the control error in the normal flight so that the abnormal flight state can be detected. The performance of proposed method was evaluated using the real-world flight data. The results showed that the method effectively detects anomalies in various situation.

• 제어로봇시스템학회논문지
• /
• 제10권2호
• /
• pp.139-144
• /
• 2004

The flight control system designed for an unmanned airship, which is under development by KARI, is in reduced. First, the dynamic characteristics of the airship are addressed, which are fairly different from those of the nominal aircraft. In order to implement autonomous flight for the unmanned airship, flight control logic is designed including autopilot and guidance law. The autopilot is designed under consideration of the velocity region of the unmanned airship. The guidance laws are implemented in main operational modes such as point navigation, station keeping and spiral up/down for emergency return. Their simulation results are also presented in order to validate performances of the flight control system.

• 한국항공운항학회지
• /
• 제18권4호
• /
• pp.1-8
• /
• 2010

In the aerospace field, the study on a morphing-wing is in progress to improve flight performance and perform multi flight mission. There are many concepts of morphing-wing such as camber-change, wing-twist, variable-span, and so on. In this study, the aerodynamic characteristics and flight control of an air vehicle with a variable-span morphing wing (VSMW) have been investigated. VSMW with symmetric span control(SSC) can increase cruising range of aircraft by reducing drag in various flight condition. VSMW with anti-symmetric span control(ASSC) can be used in the roll control of an aircraft. The flight control about pure rolling dynamic system and full dynamic system have been performed about the cruise missile.

• 제어로봇시스템학회:학술대회논문집
• /
• 제어로봇시스템학회 1996년도 한국자동제어학술회의논문집(국내학술편); 포항공과대학교, 포항; 24-26 Oct. 1996
• /
• pp.203-206
• /
• 1996

The purpose of this paper is to find how to determine the parameters of the basic control system design such as hinge moment coefficients and to display the controllability of the ChangCong-91. Since the estimation from the flight test of real aircraft is the most reliable, we performed the flight test of ChangGong-91 to get the various parameters such as velocity, height, control force, control surface deflection, 3 axis acceleration, 3 axis angular rate, pitch angle, angle of attack temperature and so on. We recorded the flight test data in VHS tapes and stored them to personal computer using A/D(analog to digital) converter. Flight test was done in various conditions, and the acquired data was processed with parameter identification method such as least square method. These data will be utilized for the development of Autopilot System design and Control Loading System design.

• Journal of Mechanical Science and Technology
• /
• 제17권5호
• /
• pp.654-667
• /
• 2003

This paper presents a consequence of the systematic approach to identify the aerodynamic parameters of an unmanned aerial vehicle (UAV) equipped with the automatic flight control system. A 3-2-1-1 excitation is applied for the longitudinal mode while a multi-step input is applied for lateral/directional excitation. Optimal time step for excitation is sought to provide the broad input bandwidth. A fully automated programmed flight test method provides high-quality flight data for system identification using the flight control computer with longitudinal and lateral/directional autopilots, which enable the separation of each motion during the flight test. The accuracy of the longitudinal system identification is improved by an additional use of the closed-loop flight test data. A constrained optimization scheme is applied to estimate the aerodynamic coefficients that best describe the time response of the vehicle. An appropriate weighting function is introduced to balance the flight modes. As a result, concurrent system models are obtained for a wide envelope of both longitudinal and lateral/directional flight maneuvers while maintaining the physical meanings of each parameter.

• 한국항공우주학회지
• /
• 제37권7호
• /
• pp.701-708
• /
• 2009

개발된 비행제어컴퓨터(DFLCC: Developed Flight Control Computer)의 비행제어법칙은 고등훈련기급 항공기의 시제 최종 형상의 비행제어소프트웨어(OFP: Operation Flight Program)를 기반으로 개발되었다. 비행제어법칙은 상용 개발 툴을 이용하여 GUI(Graphic User Interface) 환경에서 설계되며, C 코드로 변환되어 OFP 에 반영된다. 그리고 OFP는 정형화된 검증절차를 통하여 검증되는데, 검증과정을 거치기 전에 비실시간 환경에서 C코드로 변환된 비행제어법칙과 기반이 되는 비행제어법칙의 상사성(similarity)을 검증하고, 구성된 HILS(Hardware In-the-Loop Simulator) 환경의 신뢰성(reliability)을 사전에 검증하는 절차가 필요하다. 비행제어법칙의 상사성은 비실시간 환경에서 고등훈련기급 항공기의 시제 최종 버전의 비행제어법칙과 개발된 비행제어법칙의 응답특성을 상호 비교하여 검증된다. 또한, 구성된 HILS 환경의 신뢰성은 비실시간 시뮬레이션 툴을 기반으로 HILS 결과와 항공기 응답특성을 비교하여 검증된다. 본 논문에서는 항공기 응답을 직접 비교함으로써 개발된 비행제어법칙의 상사성과 HILS 환경의 신뢰성을 검증하였다.