• Advances in aircraft and spacecraft science
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• v.9 no.5
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• pp.433-447
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• 2022

This work deals with an explicit guidance and control architecture for autonomous lunar ascent and orbit injection, i.e., the locally-flat near-optimal guidance, accompanied by nonlinear reduced-attitude control. This is a new explicit guidance scheme, based on the local projection of the position and velocity variables, in conjunction with the real-time solution of the associated minimum-time problem. A recently-introduced quaternion-based reduced-attitude control algorithm, which enjoys quasi-global stability properties, is employed to drive the longitudinal axis of the ascent vehicle toward the desired direction. Actuation, based on thrust vectoring, is modeled as well. Extensive Monte Carlo simulations prove the effectiveness of the guidance, control, and actuation architecture proposed in this study for precise lunar orbit insertion, in the presence of nonnominal flight conditions.

• 제어로봇시스템학회:학술대회논문집
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• 2004.08a
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• pp.969-974
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• 2004

For a maneuvering unmanned autonomous helicopter, it is necessary to design a proper controller of each flight mode. In this paper, overall helicopter dynamics is derived and hovering model is linearized and transformed into a state equation form. However, since it is difficult to obtain parameters of stability derivatives in the state equation directly, a linear control model is derived by time-domain parametric system identification method with real flight data of the model helicopter. Then, two different controllers - a linear feedback controller with proportional gains and a robust controller - are designed and their performance is compared. Both proposed controllers show outstanding results by computer simulation. These validated controllers can be used to autonomous flight controller of a real unmanned model helicopter.

• Journal of Institute of Control, Robotics and Systems
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• v.11 no.3
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• pp.246-253
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• 2005

An experimenal study on flight control of high-speed AEV(Aero-levitation Electric Vehicle) scale model in wind-tunnel is conducted. The AEV is to fly at very low altitude in predesigned track so that it is always under the wing-in-ground effect. The experiment is intended to fly the scale model to follow the predesigned altitude schedule while holding its attitude (pitch, roll, and yaw). Especially, the altitude changes for climb, cruise, and descent with constant pitch angle are most important maneuvers. The experiment shows that the required mission flights can be performed with appropriate sensors, processors, and actuators.

• Journal of Aerospace System Engineering
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• v.3 no.3
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• pp.13-16
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• 2009

Aircraft flight control surfaces which are one of the most important elements of safety allow a pilot to adjust and control the aircraft's flight attitude. This paper is described of the control surface deflection angle measuring device. Data analysis through ground test and flight test can provide reliability of this device using the present system. It is also shown that measuring system is capable of detecting failure of control surfaces.

• Journal of Engineering Education Research
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• v.17 no.5
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• pp.17-22
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• 2014

Unmanned aerial vehicles (UAVs) that have been recently commercialized can largely be divided into fixed-wing aircraft and rotor aircraft by their styles and flight characteristics. Although the fixed-wing aircraft represents higher power efficiency, higher speed, longer flight distance and larger loading weight than the rotor aircraft, they have a disadvantage of requiring a space for take-off and landing. On the other hand, the rotor aircraft can implement vertical take-off and landing (VTOL) and represents various flight modes (hovering, steep bank turns and low-speed flights). But they require both precision take-off control and attitude control. In this study, we used a quad-tilt rotor UAV to combine advantages in both the fixed-wing aircraft and the rotor aircraft. The quad-tilt rotor (QTR) system was designed and constructed by adding a tilt device with a servo motor to a general quad-rotor vehicle.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.2
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• pp.290-299
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• 2017

Reconfigurable flight control using various kinds of adaptive control methods has been studied since the 1970s to enhance the survivability of aircraft in case of severe in-flight failure. Early studies were mainly focused on the failure of actuators. Recently, studies of reconfigurable flight controls that can accommodate complex damage (partial wing and tail loss) in conventional aircraft were reported. However, the partial wing loss effects on the aerodynamics of conventional type aircraft are quite different to those of BWB(blended wing body) aircraft. In this paper, a reconfigurable flight control algorithm was designed using a direct model reference adaptive method to overcome the instability caused by a complex damage of a BWB aircraft. A model reference adaptive control was incorporated into the inner loop rate control system enhancing the performance of the baseline control to cope with abrupt loss of stability. Gains of the model reference adaptive control were polled out using the Liapunov's stability theorem. Outer loop attitude autopilot was designed to manage roll and pitch of the BWB UAV as well. A 6-DOF dynamic model was built-up, where the normal flight can be made to switch to the damaged state abruptly reflecting the possible real flight situation. 22% of right wing loss as well as 25% loss for both vertical tail and rudder control surface were considered in this study. Static aerodynamic coefficients were obtained via wind tunnel test. Numerical simulations were conducted to demonstrate the performance of the reconfigurable flight control system.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.2
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• pp.162-169
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• 2011

The flight control system utilize RSS(Relaxed Static Stability) criteria in both longitudinal axes to achieve performance enhancements and improve stability. The aircraft using digital flight-by-wire flight control system receives aircraft flight conditions such as pitch, roll and yaw rate, normal acceleration from RSA(Rate Sensor Assembly) and ASA(Acceleration Sensor Assembly). These sensors has permissible measurement error related to system safety of an aircraft but, unexpected flight motions are happened by sensing errors such as offset, noise and etc. The unexpected pitch down tendency occurred by ASA sensor bias in 1g level flight with pilot hands-off. This paper addresses the design and verification of flight control law to improve of pitch down or up tendency caused by ASA sensor bias. The result of analysis and flight test reveals that pitch down tendency can be improved by pitch attitude feedback system.

• Journal of Aerospace System Engineering
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• v.12 no.3
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• pp.45-52
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• 2018

This paper deals with the analysis of the inner-loop algorithm of the Pixhawk which is a representative multi-copter open source. The algorithm is based on flight control-law structure. The inner-loop algorithm of the Pixhawk can be divided into a position controller and an attitude controller. The position controller generates the attitude of the multi-copter to move to the destination The position controller also generates the demand force and moment acting on each actuator. We confirm that the position controller saturates the desired acceleration and speed by using a proper relational expression. The expression can be used in order to prevent the sudden change in the attitude of a multi-copter.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.6
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• pp.537-544
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• 2009

This paper concerns analysis technique on determining of attitude control gain in the low frequency region using stability area. The stability area is defined by the D-Decomposition method, which was designed by Neimark. In this paper, it is introduced D-Decomposition method from reference paper and design attitude control gain of generic launch vehicle during first stage flight phase. For selecting PD control gain, it is considered the system parameter uncertainty about whole first-stage flight phase, represented the stability area boundary on each case. After deciding the PD control gain using stability area method, it is applied to launch vehicle linear model, and checking the stability margin requirement, frequency response characteristics.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.8
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• pp.63-70
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• 2006

Due to the small size and weight of micro aerial vehicle (MAV), only miniaturized MEMS type sensors are applicable for MAV autonomous flight system. In this paper, we propose a accelerometer and gyro mixing algorithm to improve an attitude performance of MEMS type sensors. The performance of the proposed mixing algorithm is compared with the performance of fuzzy-based mixing algorithm through simulation. The simulation results show that the attitude compensation method through the attitude compensation has better performance than the fuzzy-based mixing method for MAV attitude estimation.