• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.40 no.2
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• pp.165-170
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• 2012

This paper contains the test method to obtain aerodynamic hinge moments acting on the control surface of air vehicle wing. During the flight, hinge moments make difference between actual control surface angle and control angle which is measured by sensor of actuator. The hinge moments can be obtained by using this difference. Static ground load test and calibration test were conducted to obtain torsional stiffness of control surface actuation system. This results are used to calculate hinge moments. In addition, the mechanical errors of actuation system such as slip angle of mounting point and backlash could be estimated. Using flight test results, this experimental measurement method of hinge moment acting on control surface is conducted. The results of this method are similar to those of numerical simulation method, and the validity of this method is proved.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.3
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• pp.93-100
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• 2006

Modern versions of supersonic jet fighter aircraft using a digital flight-by-wire flight control system design utilizes a control surface reconfiguration in order to guarantee the aircraft flight stability when a control surface is failed. The T-50 flight control laws are designed such that the surface reconfiguration mode controls the aircraft using non-failed control surfaces when one of the control surfaces is failed. In this paper, linear analysis and HQS(Handling Quality Simulator) pilot simulations are performed to analyze the flight stability and handling quality when the surface reconfiguration mode is engaged for aircraft landing configuration. It is found that the aircraft flight stability and handling quality is satisfied to level 1 requirements when the T-50 flight control law is changed to the surface reconfiguration mode.

• Aerospace Engineering and Technology
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• v.5 no.2
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• pp.240-247
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• 2006

Tilt rotor aircraft is a multi-configuration airplane which has three independent flight modes; helicopter, conversion, and aiplane. The control surface mixer resign is reqctired to generate and distribute efficient control forces and moments in each flight mode. In the conversion mode, the thrust vector is changed from helicopter mode to airplane, therefore the thrust vector makes undesired forces and moments which affect on pitch, roll and yaw dynamics. This paper describes the design results of control surface mixer design which minimize the undesired forces and moments due to nacelles tilting angle change for 4O% scaled model.

• Aerospace Engineering and Technology
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• v.6 no.2
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• pp.21-28
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• 2007

The goal of this paper is to represent a technique of fault detection for the control surface as a base research of the fault tolerant control system for safety improvement of UAV. The real-time system identification based on the recursive Fourier Transform was implemented for the fault detection of the control surface and verified through the HILS and flight test. The failures of the control surface are detected by comparing the control derivatives in fault condition with the normal condition. As a result from the flight test, we have confirmed that the control derivatives of fault condition less than normal condition.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.11
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• pp.999-1005
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• 2007

In case of an abnormal condition of control surface, the real-time estimation of aerodynamic derivatives are required for the reconfigurable control system to be flight for missions or return to the head office. The goal of this paper is to represent a technique of fault detection to the control surface as a base research to the fault tolerant control system for safety improvement of UAV. The real-time system identification for the fault detection to the control surface was applied with the recursive Fourier Transform and verified through the HILS and flight test. The failures of the control surface are detected by comparing the control derivatives in fault condition with the normal condition. As a result from the flight test, we have confirmed that the control derivatives of fault condition less than about 50% in the normal condition.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.8
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• pp.109-117
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• 2004

A study on gust load alleviation using aircraft control surface was performed. Aeroservoelastic model including control surface controller was formulated and validated by comparing the results of continuous gust response analysis with those of MSC/NASTRAN. Optimal control with output feedback was adopted for designing the control surface controller, and the effects of gust load alleviation was validated by performing the numerical simulation for the controller designed.

• Composites Research
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• v.36 no.3
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• pp.211-216
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• 2023

When designing ships and aircraft structures, it is important to design them to satisfy weight reduction and strength. Currently, studies related to topology optimization using 3D printed composite materials are being actively conducted to satisfy the weight reduction and strength of the structure. In this study, structural analysis was performed to analyze the applicability of 3D printed composite materials to the flight control surface, one of the parts of an aircraft or unmanned aerial vehicle. The optimal topology shape of the flight control surface for the bending load was analyzed by considering three types (hexagonal, rectangular, triangular) of the topology shape of the flight control surface. In addition, the bending strength of the flight control surface was analyzed when four types of reinforcing materials (carbon fiber, glass fiber, high-strength high-temperature glass fiber, and kevlar) of the 3D printed composite material were applied. As a result of comparing the three-point bending test results with the finite element method results, it was confirmed that the flight control surface with hexagonal topology shape made of carbon fiber and Kevlar had excellent performance. And it is judged that the 3D printed composite can be sufficiently applied to the flight control surface.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.12
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• pp.1177-1183
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• 2010

In this paper, a parallel bank of multiple adaptive unknown input observers approach suggested by D.Wang is applied to detect a single fault of control surface actuator and to estimate the actuator position of lock-in-place fault using a small fixed-wing UAV model with eight control surfaces. This paper shows that not only the fault diagnosis algorithm detects and estimates each faults of lock-in-place in 1 second by simulation but also it may be unavailable to isolate among two same-shaped rudders.

• Aerospace Engineering and Technology
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• v.12 no.1
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• pp.54-63
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• 2013

In general, multiple hydraulic servo actuators are installed on one control surface of aero-dynamically highly loaded condition aircraft for redundancy management to satisfy flight control safety requirements. If motions of multiple actuators are not synchronized, control surface is deformed from its free stressed state. In result, force fight conditions are generated on each actuator due to restoration reaction force of deformed control surface. In addition, force fight is induced from severe initial rigging tolerance. Force fight condition of multiple actuators affects control accuracies and reduces operational life of actuators and control surface due to fatigue phenomenon. In this study, we designed controller using force feedback to reduce force fight of duplex servo actuation system.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.9
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• pp.65-71
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• 2004

Redundant control surfaces arc adopted to modern aircraft designs because of high performance and fault tolerant control, so efficient redundancy management is necessary to take advantage of redundant control surfaces. This paper focuses on the control allocation scheme as one of redundancy management methods. A modified pseudo inverse redistribution method is proposed. The existing method sets all saturated controls as their limit values when a pseudo inverse is calculated. But the modified scheme sets only one saturated control as its limit value and redistributes remaining controls. It is shown that the proposed scheme is superior to the existing method by several numerical examples.