• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.7
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• pp.701-708
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• 2009

The flight control law of developed flight control computer(DFLCC) is developed based on operation flight program of advanced trainer aircraft full scale development final configuration. The flight control law design is used common use development tool in GUI(Graphic User Interface) environment. The flight control law transformed to C-Code is reflected in OFP. The OFP is verified by the standardized verification process. But, before standardized verification process, we need preliminary verification process such as similarity of flight control law and reliability of developed HILS. Similarity of flight control law is verified by comparing the aircraft response of advanced trainer aircraft and those of the developed control law. Also, reliability of developed HILS is verified by comparing the aircraft response of HILS and Non-real time simulation result. This paper verifies similarity of developed control law and reliability of HILS environment as comparing aircraft response.

• 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 the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.633-639
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• 2007

This paper presents a new homing guidance law based on well-known BPN to achieve an impact time constraint as well as an impact angle constraint. The guidance commands are synthesized by introducing an additional command to control impact-time. The structure of the additional command has a BPN-based loop multiplied by time-varying gains being proportional to the time difference between the required time-to-go and the estimated time-to-go by BPN. Moreover, the proposed homing loop converges to BPN as the time-to-go error is reduced. The performance of the proposed guidance law is evaluated by the computer simulations.

• 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 the Korean Society for Aeronautical & Space Sciences
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• v.43 no.8
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• pp.731-738
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• 2015

The rapid prototyping process and development tool which enable the control law evaluation efficiently are needed to minimize the development cycle, cost and risk of aircraft flight control system. This paper describes a development process that integrates the designed control law into HETLAS to evaluate simulation effectively using nonlinear mathematical models. The desktop engineering simulator was developed using HETLAS for the piloted simulation evaluation of a various control modes and the procedure was developed, which quickly integrates the HETLAS into HQS(Handling Quality Simulator) and HILS(Hardware In the Loop Simulation) environments. This paper presents a rapid prototyping process using HETLAS that significantly shortens the integration process of the control law into the nonlinear math model, HETLAS, and allows the control law designs to be quickly tested in the piloted simulation and HILS environments.

• Journal of Advanced Navigation Technology
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• v.14 no.5
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• pp.596-603
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• 2010

This paper presents a homing guidance law of anti-ship missiles using flight path angle to achieve an impact time constraint as well as an impact angle constraint. the independent variable in the nonlinear engagement model is change d from the flight time to the heading angle of the missile. The proposed guidance law can home a missile to the target with zero miss distance as well as satisfying both of the impact angle and time constraints. The performance of the proposed guidance law is evaluated by the computer simulations.

• 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).

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.1-7
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• 2006

This paper presents a new point navigation guidance law which is useful for unmanned helicopters. Predicting the future position, the guidance law generates velocity and heading commands, which are used as input to autopilot. This method differs from conventional guidance law in that it reorients the direction of flight velocity vector directly, not by bank angle indirectly. For flight tests, we have developed a flight control system for a R/C helicopters. The system consists of a flight control computer, navigation sensors, and a ground station The results of the test show that the proposed law guides a unmanned helicopter along a line path within a given area. In the future, we are planning to extend the guidance law to the mission of path following. i.e., waypoint navigation.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.8
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• pp.558-565
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• 2019

A lot of studies on the survivability of missiles have been widely studied, since the technology of modern warships equipped with state-of-the-art defense systems has been improved. The survivability of missiles can be improved by attacking a target simultaneously using multiple missiles. For this reason, impact time control guidance (ITCG) laws have been widely studied. This paper deals with the effect of time-to-go estimate to ITCG laws. In this paper, two kinds of time-to-go estimate are first introduced in two-dimensional and three-dimensional environment and then ITCG laws are derived using the time-to-go estimate. Numerical simulations are performed to analyze the performance of the designed ITCG laws and the effect of time-to-go estimate is discussed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.3
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• pp.213-219
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• 2021

In this paper, an incremental twisting compensator is proposed for improving the performance of helicopter control and tested on an in-house full-scale helicopter simulator. The proposed compensator has a merit in that an incremental control input (a second-order sliding mode control input or so-called twisting control input) is simply added to improve the performance of helicopter control, while the original flight control structure remains untouched. The proposed control technique has been shown to improve the transient and steady-state response of the in-house helicopter simulator.