• Journal of the Korean Society for Aviation and Aeronautics
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• v.12 no.3
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• pp.1-12
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• 2004

When a straight-in landing from an instrument approach using ILS or VOR/DME is not possible or desirable because of topographical reason or bad weather, a circling approach maneuver is initiated by the pilot to align the aircraft with a runway for landing. Visual contact with the runway is necessary while conducting a circle to land maneuver. This research is to develop a new warning system based on a conventional marker system which alerts pilots to watch out for exceeding the circling approach area. The airborne system also uses the same receiver unit without any new installations. The objective of this research is to design and develop a Yagi antenna in a special form. The research includes computer simulations to determine the size of antenna radiation pattern and to compute an expected flight path in case of alarm to validate effectiveness of the system.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.4
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• pp.76-83
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• 2004

A Propulsion System of the CRW(Canard Rotor Wing) type UAV(Unmanned Aerial Vehicle) was composed of the turbojet engine to generate the propulsive exhaust gas, and the duct system including main and rotary ducts, the nozzle subsystem including main and tip jet nozzle for three flight modes such as lift/landing mode, low speed transition flight mode and high speed forward flight mode. Transient simulation performance utilized the ICV (Inter-component volume) method and simulated using the SIMULINK. Transient performance analysis was performed on 3 cases. Fuel flow schedules to accelerate from Idle to maximum rotational speed were divided into the step increase of the most severe case and ramp increase cases to avoid the overshoot of turbine inlet temperature, and variations of thrust and the turbine inlet temperature were investigated in some transient analysis cases.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.513-520
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• 2004

A steady-state/transient performance simulation model was newly developed for the propulsion system of the CRW (Canard Rotor Wing) type UAV (Unmanned Aerial Vehicle) during flight mode transition. The CRW type UAV has a new concept RPV (Remotely Piloted Vehicle) which can fly at two flight modes such as the take-off/landing and low speed forward flight mode using the rotary wing driven by engine bypass exhaust gas and the high speed forward flight mode using the stopped wing and main engine thrust. The propulsion system of the CRW type UAV consists of the main engine system and the duct system. The flight vehicle may generally select a proper type and specific engine with acceptable thrust level to meet the flight mission in the propulsion system design phase. In this study, a turbojet engine with one spool was selected by decision of the vehicle system designer, and the duct system is composed of main duct, rotor duct, master valve, rotor tip-jet nozzles, and variable area main nozzle. In order to establish the safe flight mode transition region of the propulsion system, steady-state and transient performance simulation should be needed. Using this simulation model, the optimal fuel flow schedules were obtained to keep the proper surge margin and the turbine inlet temperature limitation through steady-state and transient performance estimation. Furthermore, these analysis results will be used to the control optimization of the propulsion system, later. In the transient performance model, ICV (Inter-Component Volume) model was used. The performance analysis using the developed models was performed at various flight conditions and fuel flow schedules, and these results could set the safe flight mode transition region to satisfy the turbine inlet temperature overshoot limitation as well as the compressor surge margin. Because the engine performance simulation results without the duct system were well agreed with the engine manufacturer's data and the analysis results using a commercial program, it was confirmed that the validity of the proposed performance model was verified. However, the propulsion system performance model including the duct system will be compared with experimental measuring data, later.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.2
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• pp.84-91
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• 2005

For long time, the takeoff and landing control of airship was worked by human handling. With the development of the autonomous control system, the exact controls during the takeoff and landing were required and lots of methods and algorithms were suggested. This paper presents the result of airship take-off and landing by buoyancy control using air ballonet volume change and performance control of pitch angle for stable flight within the desired altitude. For the complexity of airship's dynamics, firstly, simple PID controller was applied. Due to the various atmospheric conditions, this controller didn't give satisfactory results. Therefore, new control method was designed to reduce rapidly the error between designed trajectory and actual trajectory by learning algorithm using an artificial neural network. Generally, ANN has various weaknesses such as large training time, selection of neuron and hidden layer numbers required to deal with complex problem. To overcome these drawbacks, in this paper, the RBFN (radial basis function network) controller developed. The weight value of RBFN is acquired by learning which to reduce the error between desired input output through and airship dynamics to impress the disturbance. As a result of simulation, the controller using the RBFN is superior to PID controller which maximum error is 15M.

• Journal of the Korean Society of Safety
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• v.36 no.2
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• pp.49-57
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• 2021

In the case of military supplies, any potential failure and causes of failures must be considered. This study is aimed at examining the failure modes of a rotorcraft landing system to identify the priority items. Failure mode and effects analysis (FMEA) is applied to the rotorcraft landing system. In general, the FMEA is used to evaluate the reliability in engineering fields. Three elements, specifically, the severity, occurrence, and detectability are used to evaluate the failure modes. The risk priority number (RPN) can be obtained by multiplying the scores or the risk levels pertaining to severity, occurrence, and detectability. In this study, different weights of the three elements are considered for the RPN assessment to implement the FMEA. Furthermore, the FMEA is implemented using a fuzzy rule base, similarity aggregation model (SAM), and grey theory model (GTM) to perform a comparative analysis. The same input data are used for all models to enable a fair comparison. The FMEA is applied to military supplies by considering methodological issues. In general, the fuzzy theory is based on a hypothesis regarding the likelihood of the conversion of the crisp value to the fuzzy input. Fuzzy FMEA is the basic method to obtain the fuzzy RPN. The three elements of the FMEA are used as five linguistic terms. The membership functions as triangular fuzzy sets are the simplest models defined by the three elements. In addition, a fuzzy set is described using a membership function mapping the elements to the intervals 0 and 1. The fuzzy rule base is designed to identify the failure modes according to the expert knowledge. The IF-THEN criterion of the fuzzy rule base is formulated to convert a fuzzy input into a fuzzy output. The total number of rules is 125 in the fuzzy rule base. The SAM expresses the judgment corresponding to the individual experiences of the experts performing FMEA as weights. Implementing the SAM is of significance when operating fuzzy sets regarding the expert opinion and can confirm the concurrence of expert opinion. The GTM can perform defuzzification to obtain a crisp value from a fuzzy membership function and determine the priorities by considering the degree of relation and the form of a matrix and weights for the severity, occurrence, and detectability. The proposed models prioritize the failure modes of the rotorcraft landing system. The conventional FMEA and fuzzy rule base can set the same priorities. SAM and GTM can set different priorities with objectivity through weight setting.

• Aerospace Engineering and Technology
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• v.12 no.1
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• pp.22-31
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• 2013

Flight tests on flight control performance of helicopter, conversion and airplane mode for the Smart UAV were completed. Automatic take-off and landing, automatic return home as well as automatic approach to hover were performed in helicopter mode. Climb/descent, left/right turn using speed and altitude hold mode were performed in each $10^{\circ}$ tilt angle in conversion mode. The rotor speed in airplane mode was reduced to 82% from 98% RPM in order to increase rotor efficiency with reducing Mach number at tip of rotors. It reached to the designed maximum speed, $V_{TAS}$=440 km/h at 3 km altitude. This paper presents the flight test result on full envelopment of Smart UAV. Detailed test plan and test data on control performance were also presented to prove that all data meets the flying qualities requirement.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.24 no.8
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• pp.1056-1061
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• 2000

Particles existing in a hard disk drive are known as a major source of TA(thermal asperity). Researchers have investigated how particles induce the TA phenomena, but have not verified yet the reason why and how particles are generated in a HDD. The objective of this study is to investigate why and how particles are generated, and in what condition, the largest number of particles is generated. The number of particles generated in a HDD was measured over the landing zone after various rest times of slider and during various motions and positions of slider. It is found that the large number of particles was generated when the HDD was turned on after a long rest time of slider and that a few of particles were continuously generated when the slider flied over the disk surface. It is thought that the number of particles generated in a HDD was related to the rest time of slider because the rest time of slider increased stiction, and that there were intermittent contacts between the slider and the disk surface when the slider flied over the disk surface.

• Journal of the Korean Society of Visualization
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• v.9 no.2
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• pp.40-47
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• 2011

Tilt-rotor aircraft can be used in various fields because they have the capabilities of the vertical take-off and landing and the high-speed cruise flight. In the present study, the flow analysis of a tilt-rotor aircraft is conducted at various tilt angles. The lift and drag forces of the tilt-rotor aircraft are obtained and the wakes by the rotor-blade are visualized. The result shows that the rotor-blade affects the lift force in a hovering mode and the main wing has an influence on the lift force in a cruise mode. Additional thrust is required at the tilt angle of around 40 degree due to the least lift force. The drag force is dependent on the rotor-blade at overall tilt angles. The minus drag force appears between the tilt angles of 90 degree and 55 degree. Also, the drag force is dramatically increased at the other tilt angles. The wake by rotor-blade affects the flow around the fuselage of the tilt-rotor aircraft at the tilt angles of 75 degree and 60 degree.

• Advances in aircraft and spacecraft science
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• v.2 no.4
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• pp.391-401
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• 2015

Crashworthiness design and certification have been and will continue to be the main concern in aviation safety. The effects of roll angles on fuselage section crashworthiness for typical civil transport category aircrafts were investigated. A fuselage section with waved-plates under cargo floor is suggested, and the finite element model of fuselage section is developed to simulate drop test subjected to 7 m/s impact velocity under conditions of 0-deg, 5-deg, 10-deg and 15-deg roll angles, respectively. A comparative analysis of failure modes, acceleration responses, and energy absorption of fuselage section under various conditions are given. The results show that the change of roll angles will significantly affect fuselage deformation, seat peak overloads, and energy absorption. The crashworthiness capability of aircraft can be effectively improved by choosing appropriate landing way.