• Journal of the Korean Society for Aviation and Aeronautics
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• v.23 no.4
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• pp.22-29
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• 2015

Quadrotor is widely used in variable application nowadays. Due to its inherent unstable characteristics, control system to augment the stability is essential for quadrotor operation. To design control system and verify its performance through simulation, accurate dynamic model is required. Quadrotor dynamic model is simply compared with conventional rotorcraft such as helicopter. However, the accurate dynamic model of quadrotor is not easy to develop because of the highly correlated aerodynamic effect of each rotor. In this paper, quadrotor dynamic model is identified from the flight data using frequency domain approach. Flight test of quadrotor is performed in closed loop configuration with stability augmentation system included. Frequency sweep input is applied in each of lateral, longitudinal, yaw and heave axis separately. The bare dynamic model is identified from the flight data of quadrotor responses and thrust measurement through Pulse Width Modulation(PWM) data. The frequency responses of identified model match well with those of flight data, and time responses of identified model for doublet input in each axis are also shown to agree with flight data.

• Structural Engineering and Mechanics
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• v.87 no.5
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• pp.419-429
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• 2023

This study conducts numerical analyses of a thin-walled composite cylinder under axial compression and internal pressure of 10 kPa. Numerical vibration correlation technique and nonlinear postbuckling analyses are conducted using the nonlinear finite element analysis program, ABAQUS. The single perturbation load approach and measured imperfection data are used to represent the geometric initial imperfection of thin-walled composite cylinder. The buckling knockdown factors are derived using present initial imperfection and analysis methods under axial compression without and with the internal pressure. Furthermore, the buckling knockdown factors are compared with the buckling test and computation time are calculated. In this study, derived buckling knockdown factors in present study have difference within 10% as compared with the buckling test. It is shown that nonlinear postbuckling analysis can derive relatively accurate buckling knockdown factor of present thin-walled cylinders, however, numerical vibration correlation technique derives reasonable buckling knockdown factors compared with buckling test. Therefore, this study shows that numerical vibration correlation technique can also be considered as an effective numerical method with 21~91% reduced computation time than nonlinear postbuckling analysis for the derivation of buckling knockdown factors of present composite cylinders.

• Journal of Aerospace System Engineering
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• v.13 no.6
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• pp.52-59
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• 2019

The rotor blade is a key component that generates the lift, thrust, and control forces required for helicopter flight by the torque transmitted through the hub and the blade pitch angle control, and should be designed to factor vibration characteristics so that there is no risk of resonance with structural safety. In this study, the structural design of the main rotor blade for MPUH(Multi-Purpose Unmanned Helicopter) was conducted and the sectional stiffness measurement of the fabricated blade was performed. The evaluation of the vibration characteristics of the main rotor system was then conducted factoring the measured stiffness distribution. The interior of the rotor blade comprised of the skin, spar, and torsion box, and carbon and glass fiber composites were applied. The Ksec2D program was applied to predict the stiffness of blade, and the results were compared to the measured data. CAMRADII, a comprehensive rotorcraft analysis program, was applied to investigate the natural frequency trends and resonance risks due to the rotor rotation.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.34 no.11
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• pp.1691-1696
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• 2010

The efforts of reflecting the system's uncertainties in design step have been made and robust optimization or reliabilitybased design optimization are examples of the most famous methodologies. The statistical moments of a performance function and the constraints corresponding to probability conditions are involved in the formulation of these methodologies. Therefore, it is essential to effectively and accurately calculate them. The sensitivities of these methodologies have to be determined when nonlinear programming is utilized during the optimization process. The sensitivity of statistical moments and probability constraints is expressed in the integral form and limited to the normal random variable; we aim to expand the sensitivity formulation to nonnormal variables. Additional functional calculation will not be required when statistical moments and failure or satisfaction probabilities are already obtained at a design point. On the other hand, the accuracy of the sensitivity results could be worse than that of the moments because the target function is expressed as a product of the performance function and the explicit functions derived from probability density functions.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.48 no.12
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• pp.997-1004
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• 2020

A synchropter is a type of rotorcraft in which a pair of blades inclined with each other rotates in synchronization. Removing the tail rotor enables an efficient and compact configuration similar to a coaxial-rotor helicopter. This paper describes the design and flight test results of a small synchropter to examine the suitability of a drone system for the army. The synchropter in this paper is a small vehicle with a rotor diameter of 1.4m and a weight of 7kg and was assembled based on commercial parts to examine flight characteristics effectively. The flight control system adopted Pixhawk, which is designed based on an open-architecture. The model-based design technique is applied to develop the control law of the synchropter and a new firmware embedded on the Pixhawk. Through qualitative flight tests, we analyzed the flight characteristics. As a result of the analysis, we confirmed the possibility of application as a drone system of the synchropter.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.1
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• pp.1-7
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• 2013

An aerodynamic rig test of a radial turbine for an auxiliary power unit (APU) was performed at a high-temperature turbine test facility at the Korea Aerospace Research Institute. The pressure ratio, Mach number, and flow coefficient in the rig test are the same as those under normal engine operation conditions. The design pressure ratio is 3.096, design test speed is 34909 rpm, and turbine inlet temperature is $160^{\circ}C$. The turbine has airfoil-type nozzles, and the diameter of the turbine wheel is 175.74 mm. The turbine map is experimentally measured, and the detailed flow at the turbine inlet is measured. The pressure distribution in the nozzle at both the hub and the shroud sides and the pressure distribution along the shroud casing of the turbine wheel were measured, and this confirmed that the expansion process in the turbine wheel is acceptable.

• Journal of Aerospace System Engineering
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• v.15 no.5
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• pp.106-112
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• 2021

In this study, we performed a design and structural analysis of a blade-shaped antenna installation on the rear fuselage of a rotary wing aircraft operated by the military. When the structure is damaged while the aircraft is in operation, it is separated from the aircraft main structure and may collide with the rotor or blades to cause the aircraft to crash. Therefore, structural safety for the modified structure must be secured. The design requirement for the newly installed modified part were established, and the load condition was constructed by applying the load that may occur in the aircraft after the modification. Structure safety for the modified structure was secured by performing structure analysis. To analysis stress and deformation of aircraft structure, we developed finite element model and verified it by using hand calculation method. We confirmed the safety of the modified structure through the final structural integrity analysis.

• Journal of Aerospace System Engineering
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• v.13 no.3
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• pp.56-63
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• 2019

Performance analyses were performed on a propeller developed for use in a PAV (Personal Air Vehicle) under 600 kg Maximum Take-Off Weight (MTOW). The actuator disc theory and CFD analyses were used to estimate the hovering time with regards to MTOW variation for a given battery weight. The interference induced power factor kint was introduced to account for the effect of flow interference between the propellers and to estimate the performance of counter-rotating propellers. The Maximum Figure of Merit (FM) value of the propeller pitch was determined and the design RPM range for the required power inversely obtained from the CFD results. Previous research indicate that the flight time of large multi-copter is limited by the available battery energy density. Similarly, the propeller pitch settings and spacing are important factors in reducing the kint value.

• Journal of Advanced Navigation Technology
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• v.16 no.6
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• pp.1092-1100
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• 2012

In this paper, it has designed and developed the integrated aural cueing system(ACS) system of the reconfigurable tactical Flight Training System(RTT) for the 6 rotorcraft such as UH-1H, UH-60, AH-1H, 500MD, BO-100, and CH-47. RTT is an evolving alternative instructional training system to provide the ability to rehearsal and collectively train, through networked simulators in a unit-collective and combined arms simulated battlefield environment. ACS handles the volume, pitch and repetition of the digitally stored sounds based on commands it receives from the Host server. This paper explained and implemented the conceptual and detail design the ACS system. In order to evaluating the performance of the ACS system, we made the monitoring system for interworking the virtual Host and the ACS system. As the result, it was confirmed the good performance.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.10
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• pp.842-850
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• 2014

This paper presents a trajectory tracking controller for rotorcraft UAVs to improve the tracking performances in the presence of various uncertainties. The proposed tracking method consists of a velocity guidance law based on the relative distance and L1 adaptive augmentation loop for tracking the velocity commands. In the proposed structure, the desired velocity generated by the guidance law is the reference value of the adaptive controller for accurate path tracking. In the guidance law, the desired acceleration is generated based on the relative distance and its derivatives, and then the velocity command of the inner control loop is calculated by integrating the accelerations. $L_1$ augmentation loop supplements the linear controller to guarantee the flight performances such as a tracking accuracy in the presence of the uncertainties. The proposed controller was validated in actual flight tests to successfully demonstrate its capability using a quadrotor UAV.