• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.2
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• pp.169-179
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• 2010

Satisfying the same probability of loss of control and essentially two fail operative performance with a triplex computer architecture requires a lot of modification of the conventional redundancy management design techniques, previously employed in quadruplex digital flight control computer. T-50 FCS for triplex redundancy management design applied an advanced digital flight control architecture with an I/O controller which is functionally independent of the digital computer to achieve the same reliability and special failure analysis and isolation schemes for fail operational goals with a triplex configuration. The analysis results indicated that the triplex flight control system is to satisfy the safety requirement utilizing the advanced flight control techniques and the system performance of the implemented flight control system was verified by failure mode effect test.

• 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.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.33 no.11
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• pp.80-87
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• 2005

T-50 supersonic jet trainer aircraft using digital flight-by-wire flight control system receives aircraft flight conditions such as altitude, VCAS(Calibrated Airspeed) and Angle of Attack from IMFP(Integrated Multi-Function Probe). IMFP sensors information have triplex structure using three IMFP sensors. Air-data selection logic is mid-value selection in three information from three IMFP sensors in order to have more reliability. From supersonic flight test at high altitude, air-data information is dropped simultaneously because of supersonic shock wave effect. This error information may affect to aircraft stability and safety in supersonic area at low altitude. This paper propose that sensitivity analysis and HQS(Handling Quality Simulator) pilot simulation in order to analyze flight stability and controllability in supersonic area at low altitude when these information is applied to flight control law.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.3
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• pp.81-86
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• 2006

This paper deals with a driving current fault monitor design methodology for a DDV actuation system which has a dual hydraulic power supply system, and triplex electric control capability. A fault existing among these redundant channels should be detected accurately and removed timely, and the remaining channels are to be reconfigured in order to compensate the role of a removed faulty channel. An integrated analysis on the aerodynamics, flight control laws, and DDV actuation system is essential for the design of an actuation system fault monitor. A method to define a fault transient boundary which specifies a maximum travel of an actuation system caused by the first faulty operation is proposed based on the top level requirement on the fault effect specified in MIL-F-8785C.