• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.8-27
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• 2002

In this study, the load analysis of a composite canard aircraft is performed numerically. Excel visual basic program for PC is used to calculate aerodynamic coefficients, loads and moments etc.. The basic data required for the load analysis such as aircraft configuration and dimension, parts and its weight and coordinate etc. are obtained from Catia modeling, measurement or material density. Aircraft weight, center of gravity, inertia moment, structural design speeds, wing load distribution, forces and moments are evaluated by using these data. V-n diagram is also represented for selecting critical loads applied to the wing and fuselage. The V-n diagram is investigated to decide the flight envelope of canard aircraft for design speed VA, VC, VD and load factor +3.8G, -1.52G at maximum weight of 2,573 lbs and sea level. In the future, the results of the wing and fuselage load analysis is to represented by using selected critical loads.

• 한국항공운항학회:학술대회논문집
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• 2007.06a
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• pp.109-112
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• 2007

• Aerospace Engineering and Technology
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• v.6 no.1
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• pp.35-44
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• 2007

In this study was made the flutter analysis for the export model of Firefly(Bandi-ho), the small canard aircraft. Stiffness model based on internal load generation finite element model was generated. Mass model based on the weight DB for weight control was generated. Aerodynamic model based on Doublet Lattice Method was generated. Preliminary flutter analysis was made. Based on it, major vibration modes are identified and experimentally obtained via the ground vibration test. The obtained normal mode frequencies were used to correlate the finite element model. Flutter analysis was made again and major flutter mechanisms were summarized. The most important flutter root was identified as a coupled root between rigid body roll mode and anti-symmetric wing pitching mode.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.22-28
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• 2003

This paper presents a method for estimating static aerodynamic derivatives by analyzing data obtained from the flying quality evaluation test of a small canard aircraft. The aerodynamic derivatives extracted from maximum likelihood estimation method and from the proposed method in this paper are compared in the same polt. Reliable static aerodynamic derivatives were extracted from a limited number of the flight tests by the proposed method. The parameter data obtained from this method can be used as reference for the conventional parameter identification methods such as maximum likelihood estimation method.

• Journal of the Korea Institute of Military Science and Technology
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• v.20 no.3
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• pp.376-383
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• 2017

This paper describes the gyroscopic stability and the drag characteristics of the smart munition with a course correction fuze(CCF). A ballistic analysis was conducted to figure out the effect of the canards on the gyroscopic stability of the projectile. The analysis used the commercial ammunition performance evaluation software: Projectile Design and Analysis System(PRODAS). In particular, we compared the PRODAS analysis results to real field test results to investigate the influence of the CCF mounted projectile. In addition, some ballistic simulations were carried out to provide the conditions suitable for wind tunnel tests. Experimental results show that the added drag force by the canards is almost uniform regardless of the Mach number when the projectile is at the normal position where the angle of rotation and the angle of attack are both 0 degrees. However, as the angle of attack of the projectile increases, the additional drag force depends on the deflection of the canards.

• Journal of the Korean Society of Marine Environment & Safety
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• v.20 no.4
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• pp.443-450
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• 2014

In this study, we have discussed about the effect of hydrofoil arrangement and longitudinal moment characteristic on longitudinal motion stability of fully-submerged hydrofoil by the experiment of tandem hydrofoil model. First of all, tandem hydrofoil model that has canard wing arrangement has been made and characteristics of lift force and drag force by performing the lift force and drag force measuring experiment has also been estimated. Besides, tandem hydrofoil model's wing arrangement which has the initial stability and self stability of longitudinal motion has also been determined. In longitudinal stability experiment of tandem hydrofoil model, the motion characteristic of pitch and heave and the longitudinal stability of foil borne condition by variation of self stability of longitudinal moment and longitudinal distance are estimated. The result from the experiment and it's important conclusion can be described as below; Increase the self stability for longitudinal moment, the higher self stability for pitch motions in a constant pitch angles. By increasing the self stability for longitudinal moment, the range of fluctuation of pitch motion and heave motion for pitch angle also will change relatively small and longitudinal stability is excellent. Lastly, when the lift force of hydrofoil is remain constants, we can conclude that securing the enough self stability for longitudinal moment is essential for stable foil borne condition of tandem hydrofoil.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.219-224
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• 2005

In order to investigate transient behaviour of the CRW(Canard Rotor Wing) type UAV(Uninhabited Aerial Vehicle) propulsion system during flight mode transition considering flow control valve operation, the propulsion system was modelled using SIMULINK commercial program. For transient simulation of the main engine system, the ICV(Inter-Component Volume) method was applied. The valve system is to control the gas flow of the rotary duct system and the main duct system, and the analysis was performed with an assumption that the total gas mass flow of the main engine is the same as summation of the rotary duct flow and the main duct flow, and with consideration of valve loss, flow rate and effective area in valve angle variation. The performance analysis was carried out during flight mode transitions from the rotary flight mode to the fixed wing flight mode and vice versa mode at altitude of 1Km, flight Mach number 0.1 and maximum engine rpm.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.36 no.11
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• pp.1327-1334
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• 2012

This study shows the conformance-based design results of a fourth-order dynamic actuator showing a performance variation caused by variation in the components as well as aero-induced disturbances. The actuator comprises a BLDC motor, spur gear and worm gear assembly, and canard. The actuator performance was evaluated by using time-variant angle information of the canard. Based on the response surface models, critical system variables were screened using F-tests, and then, the performance was approximated as a function of the variables because it is difficult to determine the performance of a high-order dynamic system as a function of system variables through analytical approaches. In this study, the conformance of uncertain performance to the specification was defined as a probability measure. The design variables obtained by optimizing the measure can provide actuator performance conforming to the specifications considered, even though there is a variation in both the components and the aero-induced disturbances.

• 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 computational fluids engineering
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• v.13 no.2
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• pp.1-7
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• 2008

For Aerodynamic analysis of vehicle at altitude, 100km, the validity of governing equations based on continuum model, was reviewed. Also, as the preliminary study for the sub-orbital space plane development, a candidate geometry was suggested and computational fluid dynamic(CFD) analysis was performed for various angles of attack in subsonic and supersonic flow regimes to analyze the aerodynamic characteristics and performance. The inviscid flow analyses showed that the stall starts at angle of attack above $20^{\circ}$, the maximum drag is generated at angle of attack, $87^{\circ}$ and the maximum lift to drag ratio is about 8 in subsonic flow. In supersonic, the stall angle is about $40^{\circ}$ and the maximum drag is generated at angle of attack, $90^{\circ}$. Also, mach number distribution of re-entry vehicle was computed versus altitudes.