• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.187-194
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• 2008

Because the types and severities of most engine faults are various and complex, it is not easy that the conventional model based fault detection approach like the GPA(Gas Path Analysis) method can monitor all engine fault conditions. Therefore this study proposed newly a diagnostic algorithm for isolating and diagnosing effectively the faulted components of the smart UAV propulsion system, which has been developed by KARI(Korea Aerospace Research Institute), using the fuzzy logic and the neural network algorithms. A precise performance model should be needed to perform the model-based diagnostics. The based engine performance model was developed using SIMULINK. For the work and mass flow matching between components of the steady-state simulation, the state-flow library was applied. The proposed steady-state performance model can simulate off-design point performance at various flight conditions and part loads, and in order to evaluate the steady-state performance model their simulation results were compared with manufacturer's performance deck data. According to comparison results, it was confirm that the steady-state model well agreed with the deck data within 3% in all flight envelop. The diagnosis procedure of the proposed diagnostic system has the following steps. Firstly after obtaining database of fault patterns through performance simulation, then secondly the diagnostic system was trained by the FFBP networks. Thirdly after analyzing the trend of the measuring parameters due to fault patterns, then fourthly faulted components were isolated using the fuzzy logic. Finally magnitudes of the detected faults were obtained by the trained neural networks. Because the detected faults have almost same as degradation values of the implanted fault pattern, it was confirmed that the proposed diagnostic system can detect well the engine faults.

• Journal of Aerospace System Engineering
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• v.17 no.1
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• pp.79-87
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• 2023

To design a control law of an aircraft, an accurate aircraft dynamic model is required. To obtain an aerodynamic database (DB) to build a dynamic model, a large number of wind tunnel tests are typically required. However, when flight test data of target aircraft exist such as in the process of unmanned conversion of a manned aircraft, an aircraft dynamic model can be obtained through a parameter estimation method and a DB tuning procedure. This paper describes a nonlinear model construction process and a verification method for KC-100 OPV aircraft. Flight data compatibility analysis was performed to determine suitability of the estimation method application. Linear model estimation was performed using the maximum likelihood estimation method. Results of aerodynamic DB tuning process and verification applying the FFS standard to the nonlinear model constructed are presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.4
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• pp.99-104
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• 2003

This paper described the general dynamic point for rotor design and the design procedure of low vibration blade. Generally, rotor rotating natural frequencies are determined to minimize hub loads, blade vibration and to suppress ground resonance at rotor design stage. First, through rotor frequency diagram, natural frequencies must be far away from resonance point and rotating loads generated from blade can be transformed to non-rotating load to predict fuselage vibration. Vibration level was predicted at each forward flight condition by calculating cockpit's vertical acceleration transferred from non-rotating hub load assuming a fuselage as a rigid body. This design method is applied to design current Next-generation Rotor System Blade(NRSB) and will be applied to New Rotor which will be developed Further.

• Journal of Advanced Navigation Technology
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• v.22 no.2
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• pp.84-89
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• 2018

Due to the development of domestic technology, a variety of aircraft launch weapons have been developed, and the importance of aircraft-store integration certification is increasing. The aircraft-Store integration certification is to certify compliance with the armaments/stores integration criteria set out in the Standard ACC and to prove that there is no problem with the safety flight. Therefore, it is necessary to reflect the requirements of the aircraft in the store development process to reduce the design change requirement in the compatibility verification stage. In this paper, the relationship between the Standards ACC, aircraft-store compatibility reference document MIL-HDBK-1763, and the development process of cruise guided weapons have been analyzed. As a result of the analysis, it was concluded that the design changes in the aircraft-store integration certification stage could be reduced if the aerodynamic and structural design requirements were reflected from the conceptual design stage.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.7
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• pp.539-547
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• 2021

Helicopter rotor blade is required to be designed by considering the interacting effects among aerodynamics, flexibility, and controllability. The reverse design allows the structural components to have common characteristics by using the configuration numerics and experimental results. This paper aims to design the composite rotor blade which will feature common characteristics with that of BO-105. The present engineering design procedure is done by dividing the rotor blade into a few sections and composite laminates across the cross section. For each section, variational asymptotic beam sectional analysis (VABS) program is used to evaluate its flapwise, lagwise, and torsion stiffnesses to have discrepancy smaller than certain tolerance. Finally, CAMRAD II is used to predict the stress acting on the rotor blade during the specific flight condition and to check whether the present deign is structurally valid.

• Journal of IKEEE
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• v.23 no.3
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• pp.1108-1111
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• 2019

This paper presents the comprehensive transformer design methodology with the winding method of automatic type suitable for LLC resonant converter. The problem with conventional LLC resonant transformer is that it is difficult to implement the winding method of automatic type because all windings are wound in one bobbin. Therefore, the LLC resonant transformer, which is capable of the winding method of automatic type to multiple section bobbin without insulation tape and barrier tape, is proposed in this paper and the design procedure of the proposed LLC resonant transformer is also shown. Modeling and analysis of the proposed LLC resonant transformer using Maxwell 3D simulation tool are described in detail and a protype 150W LLC resonant converter using the proposed LLC resonant transformer manufactured by 3D modelling is also verified through the experimental test.

• Journal of Institute of Control, Robotics and Systems
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• v.21 no.7
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• pp.691-698
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• 2015

Inertial Navigation Systems (INS) are widely used as the main navigation device for aircraft. To get the initial attitude, the INS requires the initial alignment before navigation starts. An aircraft also needs an engine test procedure that causes some vibrations before flight. An INS can't be aligned in a vibration environment so the initial alignment is performed before the aircraft engine test. Therefore, the initial alignment time of an INS has been a major factor in limiting an aircraft's takeoff response time. In this paper, we designed an initial alignment algorithm that can be executed even in disturbances such as aircraft run-up. We demonstrated verification of the algorithm that is embedded on the real INS and testing methods to evaluate the alignment of the INS. We also analyzed the test results of the proposed initial alignment algorithm that is performed during a real aircraft run-up.

• Journal of Advanced Navigation Technology
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• v.26 no.2
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• pp.47-53
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• 2022

With the recent development of defense technology, various aircraft armaments/equipments are being developed. Among them, cruise guided weapons(missile) for launching aircraft can be installed on aircraft aftersecuring reliability and safety through varioustests. In order to avoid collision with aircraft, missile separated from aircraft must release restraint of wing at the specified time, control wing in the specified direction. In order to confirm this, a test system that can control MIL-STD-1760 according to safety separation procedure and verify release performance is required. test system needs a function to check circuit for release of restraints and a function to check driving performance by controlling MIL-STD-1760. In addition, test system should be simulate environment separated from aircraft. This paper presented a testsystemthat can verify circuit and driving performance mounted on missile according to the safety separation procedure, and it was confirmed that it was normally separated from aircraft through flight tests.

• Journal of Advanced Navigation Technology
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• v.27 no.5
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• pp.621-628
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• 2023

This study derived a reverse engineering procedure for verifying the design identity of original parts and developed parts for PMA(Parts Manufacturer Approval) and STC(Supplemental Type Certificate) of metal brake pads for transport aircraft, which are critical parts among aircraft parts and equipment. In Korea, the regulations for reverse engineering procedures are regulated by the Parts Manufacturer Approval Guidelines, and in the United States, AC No. 21. 303-4. In the reverse engineering for the brake pad, the detailed procedures for each component were determined by selecting verification test items to confirm identity based on sample quantity, dimensional tolerance, mechanical property measurement, material, weight and volume characteristics for each component. In addition, as a result of analyzing the regulation of Korea and United States, in the case of Korea, it is necessary to establish technical standards for braking systems for transport aircraft and regulations related to flight tests.

• Journal of Aerospace System Engineering
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• v.15 no.5
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• pp.8-15
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• 2021

This paper describes a design method using an optimization technique to satisfy the longitudinal handling quality of high maneuverable jet aircraft. The dynamic inversion technique was applied to the target aircraft, and the control gain optimization satisfied the longitudinal short-period handling quality, however, the stability margin was not considered. If the stability margin is not satisfied, it is necessary to directly readjust the gains through trial and error methods for improvement. To improve this, an additional compensator and an optimization constraint were added to the control gain optimization procedure. In addition, the degree of handling quality satisfaction with the optimization result was reevaluated, and additional control evaluation criteria for the convergence of the time response and the steady state error that the flight performance requirement set as the optimization constraint cannot be reflected, and the results are described.