• Title, Summary, Keyword: flight control

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Lateral-Directional Dynamic Inversion Control Applied to Supersonic Trainer (초음속 고등훈련기 가로-방향축 모델역변환 비행제어법칙 설계)

  • Kim, Chongsup;Ji, Changho;Cho, In-Je
    • Journal of Aerospace System Engineering
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    • v.8 no.4
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    • pp.24-31
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    • 2014
  • The modern version of aircrafts is allowed to guarantee the superior handing qualities within the entire flight envelope by imposing the adequate stability and flying qualities on a target aircraft through the various techniques of flight control law design. Generally, the flight control law of the aircraft in service applies the various techniques of the verified control algorithm, such as dynamic inversion and eigenstructure assignment. The supersonic trainer employs the RSS(Relaxed Static Stability) concept in order to improve the aerodynamic performance in longitudinal axis and the longitudinal control laws employ the dynamic inversion with proportional-plus-integral control method. And, lateral-directional control laws employ the blended roll system of both beta-betadot feedback and simple roll rate feedback with proportional control method in order to guarantee aircraft stability. In this paper, the lateral-directional flight control law is designed by applying dynamic inversion control technique as a different method from the current supersonic trainer control technique, where the roll rate command system is designed at the lateral axis for the rapid response characteristics, and the sideslip command system is adopted at the directional axis for stability augmentation. The dynamic inversion of a simple 1st order model is applied. And this designed flight control law is confirmed to satisfy the requirement presented from the military specification. This study is expected to contribute to design the flight control law of KF-X(Korean Fighter eXperimental) which will proceed into the full-scale development in the near future.

Moving Mass Actuated Reentry Vehicle Control Based on Trajectory Linearization

  • Su, Xiao-Long;Yu, Jian-Qiao;Wang, Ya-Fei;Wang, Lin-lin
    • International Journal of Aeronautical and Space Sciences
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    • v.14 no.3
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    • pp.247-255
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    • 2013
  • The flight control of re-entry vehicles poses a challenge to conventional gain-scheduled flight controllers due to the widely spread aerodynamic coefficients. In addition, a wide range of uncertainties in disturbances must be accommodated by the control system. This paper presents the design of a roll channel controller for a non-axisymmetric reentry vehicle model using the trajectory linearization control (TLC) method. The dynamic equations of a moving mass system and roll control model are established using the Lagrange method. Nonlinear tracking and decoupling control by trajectory linearization can be viewed as the ideal gain-scheduling controller designed at every point along the flight trajectory. It provides robust stability and performance at all stages of the flight without adjusting controller gains. It is this "plug-and-play" feature that is highly preferred for developing, testing and routine operating of the re-entry vehicles. Although the controller is designed only for nominal aerodynamic coefficients, excellent performance is verified by simulation for wind disturbances and variations from -30% to +30% of the aerodynamic coefficients.

Comparison Study of Nonlinear CSAS Flight Control Law Design Using Dynamic Model Inversion and Classical Gain Scheduling (항공기 CSAS 설계를 위한 고전적 Gain Scheduling 기법과 Dynamic Model Inversion 비선형 기법의 비교 연구)

  • Ha, Cheol-Geun;Im, Sang-Su;Kim, Byeong-Su
    • Journal of Institute of Control, Robotics and Systems
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    • v.7 no.7
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    • pp.574-581
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    • 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.

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Reconfigurable Flight Control Design for the Complex Damaged Blended Wing Body Aircraft

  • Ahn, Jongmin;Kim, Kijoon;Kim, Seungkeun;Suk, Jinyoung
    • International Journal of Aeronautical and Space Sciences
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    • v.18 no.2
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    • pp.290-299
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    • 2017
  • Reconfigurable flight control using various kinds of adaptive control methods has been studied since the 1970s to enhance the survivability of aircraft in case of severe in-flight failure. Early studies were mainly focused on the failure of actuators. Recently, studies of reconfigurable flight controls that can accommodate complex damage (partial wing and tail loss) in conventional aircraft were reported. However, the partial wing loss effects on the aerodynamics of conventional type aircraft are quite different to those of BWB(blended wing body) aircraft. In this paper, a reconfigurable flight control algorithm was designed using a direct model reference adaptive method to overcome the instability caused by a complex damage of a BWB aircraft. A model reference adaptive control was incorporated into the inner loop rate control system enhancing the performance of the baseline control to cope with abrupt loss of stability. Gains of the model reference adaptive control were polled out using the Liapunov's stability theorem. Outer loop attitude autopilot was designed to manage roll and pitch of the BWB UAV as well. A 6-DOF dynamic model was built-up, where the normal flight can be made to switch to the damaged state abruptly reflecting the possible real flight situation. 22% of right wing loss as well as 25% loss for both vertical tail and rudder control surface were considered in this study. Static aerodynamic coefficients were obtained via wind tunnel test. Numerical simulations were conducted to demonstrate the performance of the reconfigurable flight control system.

Flight Test of Pitch Control Force for an Airplane (항공기 피치 조종력 비행시험)

  • Lee, Jung-hoon
    • Journal of Aerospace System Engineering
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    • v.8 no.3
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    • pp.20-26
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    • 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.

A Study on Aircraft Flight Stability of T-50 Air Data Reconfiguration Mode (T-50 형상 재구성 모드의 항공기 비행 안정성에 관한 연구)

  • Kim, Chong-Sup;Hwang, Byung-Moom;Hwang, Min-Hwan;Bae, Myung-Whan
    • Journal of the Korean Society for Aeronautical & Space Sciences
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    • v.33 no.5
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    • pp.57-64
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    • 2005
  • Modern versions of supersonic jet fighter aircraft using digital flight-by-wire flight control system receive aircraft flight condition such as altitude, airspeed and AoA(angle of attack) from IMFP(Integrated Multi-Function Probe). IMFP sensors data have triplex structure using three IMFP sensors. An air data reconfiguration mode is applied to a T-50 flight control law to guarantee the aircraft flight stability when 2 or 3 IMFP sensors data are invalided. In this study, linear analysis and HQS(Handling Quality Simulator) pilot simulation are performed to analyze flight stability when the air data reconfiguration mode is applied to the control law. And we propose an example that the air data reconfiguration mode is applied to the control law due to second failure of IMFP during T-50 flight. It is found that the aircraft flight stability is not affected when the T-50 flight control law is changed to the air data reconfiguration mode.

A Study on Parameter Estimation for General Aviation Canard Aircraft

  • Kim, Eung Tai;Seong, Kie-Jeong;Kim, Yeong-Cheol
    • International Journal of Aeronautical and Space Sciences
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    • v.16 no.3
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    • pp.425-436
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    • 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.

System Identification and Stability Evaluation of an Unmanned Aerial Vehicle From Automated Flight Tests

  • Jinyoung Suk;Lee, Younsaeng;Kim, Seungjoo;Hueonjoon Koo;Kim, Jongseong
    • Journal of Mechanical Science and Technology
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    • v.17 no.5
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    • pp.654-667
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    • 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.

Design of a Digital Adaptive Flight Control Law for the ALFLEX

  • Ito, Hideya;Shimada, Yuzo;Uchiyama, Kenji
    • 제어로봇시스템학회:학술대회논문집
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    • pp.519-524
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    • 2003
  • In this report, a longitudinal adaptive flight control law is presented for the automatic landing system of a Japanese automatic landing flight experiment vehicle (ALFLEX). The longitudinal adaptive flight control law is designed to track an output of the vehicle to a guidance signal from the guidance portion of the automatic landing system. The proposed adaptive control law in the attitude control portion adjusts the controller gains continuously online as flight conditions change, in spite of the existence of unmodeled dynamics. The number of the controller gains to be adjusted is decreased to 1/2 from the previous studies. Computer simulation involving six-degree-of-freedom (DOF) nonlinear flight dynamics is performed to examine the effectiveness of the proposed adaptive control law. In order to verify the influence of the dispersion of the initial conditions, the Monte Carlo simulation is also applied. The initial conditions are more widely dispersed than the previous studies. As a result, except under the unsuitable initial conditions, the ALFLEX successfully landed on the runway.

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Simulation of the control force of the light aircraft using flight test data (비행시험 자료를 이용한 경항공기의 조종력 시뮬레이션)

  • 김정환;황명신;이정훈
    • 제어로봇시스템학회:학술대회논문집
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    • pp.203-206
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    • 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.

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