• Journal of Institute of Control, Robotics and Systems
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• v.14 no.7
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• pp.711-719
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• 2008

Although non-real time simulation and pilot based evaluations are available for the development of flight control computer prior to real flight tests, there are still many risky factors. The control law designed for prototype aircraft often leads to degraded performance from the initial design objectives, therefore, the proper evaluation methods should be applied such that flight control law designed can be verified in real flight environment. The one proposed in this paper is IFS(In-Flight Simulator). Currently, this system has been implemented into the F-18 HARV(High Angle of Attack Research Vehicle), SU-27 and F-16 VISTA(Variable stability. In flight Simulation Test Aircraft) programs. This paper addresses the concept of switching mechanism for FLCC(Flight Control Computer)-SWMC(Switching Mechanism Computer) using 1553B communication based on flight control law of advanced supersonic trainer. And, the fader logic of TFS(Transient Free Switch) and stand-by mode of reset '0' type are designed to reduce abrupt transient and minimize the integrator effect in pitch axis control law. It hans been turned out from the pilot evaluation in real time that the aircraft is controllable during the inter-conversion process through the flight control computer, and level 1 handling qualities are guaranteed. In addition, flight safety is maintained with an acceptable transient response during aggressive maneuver performed in severe flight conditions.

• 제어로봇시스템학회:학술대회논문집
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• 2005.06a
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• pp.1871-1876
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• 2005

In this paper, We present a new nonlinear adaptive control law using a disturbance accommodating control (DAC) observer for a Japanese automatic landing flight experiment vehicle called ALFLEX. A future spaceplane must have ability to deal with greater fluctuations in the stability and control derivatives of flight dynamics, because its flight region is much wider than that of conventional aircraft. In our previous studies, digital adaptive flight control systems have been developed based on a linear-parameter-varying (LPV) model depending on dynamic pressure, and obtained good simulation results. However, under previous control laws, it is difficult to accommodate uncertainties represented by disturbance and nonlinearity, and to design a stable flight control system. Therefore, in this study, we attempted to design a nonlinear adaptive control law using the DAC Observer and inverse dynamic methods. A good tracking property of the obtained system was confirmed in numerical simulation.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.7
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• pp.712-718
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• 2008

This paper deals with a development of flight control law switching system which can be used for flight test of the research control law by switching control law during flight. Through this research program, fader logic and integrator stabilization design has been introduced to minimize the transient response of aircraft caused by flight control law switching and to prevent the divergence of the integrator included in the control law in standby mode. MIL-STD-1553B communication was applied to transfer the data between the two control laws. This paper introduce the control law switching system architecture and major design concept and include the system verification and validation result performed on the flying quality simulator of the advanced trainer.

• Journal of Institute of Control, Robotics and Systems
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• v.9 no.6
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• pp.478-488
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• 2003

A guidance scheme that is suitable for controlling the aircraft flight path is proposed. The concept of miss distance which is commonly used in the missile guidance laws, and Lyapunov stability theorem are effectively combined to obtain the aircraft's trajectory-tracking guidance law. Guidance commands are given in terms of speed and flight path angles, but they perfectly reflect any position and velocity errors between real aircraft trajectory and reference one. The proposed guidance law is easily integrated into the existing flight control system. The new guidance law was extensively tested with various mission scenarios and the fully nonlinear 6-DOF aircraft model. Furthermore, the new guidance law was compared with previous guidance schemes in nonlinear simulation. Results from the numerical simulation show that the proposed guidance law yields better performance than previous ones.

• 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.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.10
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• pp.79-88
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• 2003

The paper presented here covers the work associated with the flight control law design, ground based and in flight simulation and handling qualities assessment of the Fly-by-Wire type Aircraft (FBWA). The FBWA configurations are of the same generic form of the Korean advanced trainer. The normal acceleration (Nz) and pitch rate (q) feedback control system is employed for longitudinal axis and roll rate (p) and lateral acceleration (Ny) feedback flight control law is developed in lateral/ directional axis. The flight tests for the FBW A dynamics evaluation were executed for the target aircraft (FBWA) on the IFS (In-Flight-Simulator) aircraft . The test results showed that Level 1 handling qualities for the most unstable flight regime and Level 1/2 for the landing approach flight regime were achieved. And the designed FBWA flight control law has revealed acceptable CHR (Cooper-Harper handling qualities Ratings).

• Journal of Institute of Control, Robotics and Systems
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• v.13 no.5
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• pp.422-428
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• 2007

This paper presents a neural network based adaptive control approach to a reconfigurable flight control law that keeps handling qualities in the presence of faults or failures to the control surfaces of an aircraft. This approach removes the need for system identification for control reallocation after a failure and the need for an accurate aerodynamic database for flight control design, thereby reducing the cost and time required to develope a reconfigurable flight controller. Neural networks address the problem caused by uncertainties in modeling an aircraft and pseudo control hedging deals with the nonlinearity in actuators and the reconfiguration of a flight controller. The effect of the reconfigurable flight control law is illustrated in results of a nonlinear simulation of an unmanned aerial vehicle Durumi-II.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.11
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• pp.1104-1111
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• 2008

The supersonic advanced trainer based on digital flight-by-wire flight control system uses aircraft flight information such as altitude, calibrated airspeed and angle of attack to calculate flight control law, and this information is measured by IMFP(Integrated Multi-Function Probe) equipment. The information has triplex structure using three IMFP sensors. Final value of informations is selected by mid-value selection logic to have more flight data reliability. As the result of supersonic flight test, pitch oscillation is occurred due to IMFP noise when altitude hold autopilot mode is engaged. This tendency may affect stability and handling quality of an aircraft during autopilot mode. This paper addresses autopilot control law design to remove pitch oscillation and these control laws are verified by non-real time simulation and flight test. Also, pitch response characteristics of pitch attitude hold autopilot mode is improved by upgrading the control law structure and feedback gain tuning during bank turn.

• Journal of Institute of Control, Robotics and Systems
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• v.13 no.7
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• pp.694-702
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• 2007

The SSWM(Software Switching Mechanism) of I-processor concept using non-real time in-house software simulation program is an effective method in order to develop the flight control law in desktop or HQS environment. And, this system has some advantages compare to HSWM(Hardware Switching Mechanism) such as remove the time delay effectiveness and reduce the costs of development. But, if this system loading to the OFP(Operational Flight Program), the OFP guarantee the enough throughput in order to calculate the two control law at once. Therefore, the HSWM(Hardware Switching Mechanism) of 2-processor concept is necessary. This paper addresses the concept of HSWM of the HQS-PC interface using TCP/IP(Transmission Control Protocol/Internet Protocol) communication based on flight control law of advanced supersonic trainer. And, the fader logic of TFS(Transient Free Switch) and stand-by mode of reset '0' type are designed in order to reduce the abrupt transient response and minimize the integrator effect in pitch axis. The result of the analysis based on HQS pilot simulation using HSWM reveals that the flight control systems are switching between two computers without any problem.

• Journal of the Korean Society for Industrial and Applied Mathematics
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• v.19 no.3
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• pp.235-252
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• 2015

Optimal impact angle control guidance law and its variants for intercepting a maneuvering target are introduced in this paper. The linear quadratic(LQ) optimal control theory is reviewed first to setup framework of guidance law derivation, called the sweep method. As an example, the inversely weighted time-to-go energy optimal control problem to obtain the optimal impact angle control guidance law for a fixed target is solved via the sweep method. Since this optimal guidance law is not applicable for a moving target due to the angle mismatch at the impact instant, the law is modified to three different biased proportional navigation(PN) laws: the flight path angle control law, the line-of-sight(LOS) angle control law, and the relative flight path angle control law. Effectiveness of the guidance laws are verified via numerical simulations.