• Journal of the Korea Institute of Military Science and Technology
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• v.21 no.3
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• pp.349-360
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• 2018

This paper proposes various methods for constructing a HWIL simulation system including Inertial Navigation System(INS) and Guidance Control Unit(GCU) under the assumption that the INS identifies the initial attitude of an aviation body through its own alignment and that it is a package consisting of an inertial sensor and a navigation computation module. This paper also presents a real-time computing technology and a way to calculate the command of the Flight Motion System(FMS) analogous to the acutal flight environment. The proposed HWIL simulation system is constructed by applying the above-mentioned methods and the results of running a series of simulations confirm high effectiveness and usefulness of the system. Finally, minor error factors that could be acquired only in HWIL simulation Environment are analyzed.

• Proceedings of the Korean Institute of Navigation and Port Research Conference
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• v.1
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• pp.109-113
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• 2006

This paper presents an in-flight alignment method for strapdown inertial navigation systems based on the line-of-sight information. Unlike the existing methods, the proposed method utilizes only the 2-axis angle measurements of the onboard image sensor and does not require any explicit range measurements between the vehicle and landmarks. To improve the accuracy of all the position, velocity, and attitude estimates through the in-flight alignment, an error model of the image-sensor-aided SDINS is derived. A simulation study demonstrates that the accuracy of SDINS can be improved by the line-of-sight information only.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.6
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• pp.72-80
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• 2004

In this paper, the design procedure of a guidance algorithm in the boosting phase of missiles with free-flight after thrust cut-off is introduced. The purpose of the guidance is to achieve a required velocity vector at the thrust cut-off. Trajectory optimizations for four cost functions are performed to investigate implementable trajectories in the pitch plane. It is observed from the optimization results that high angle of attack maneuver in the beginning of the flight are required to satisfy the constraints. The proposed guidance algorithm consists of the pitch program to produce open-loop pitch attitude command and the yaw attitude command generator to nullify the velocity to go. The pitch program utilizes the pitch attitude histories obtained from the trajectory optimization.

• 제어로봇시스템학회:학술대회논문집
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• 2005.06a
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• pp.265-270
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• 2005

This paper presents a new attitude stabilization and control of an unmanned helicopter based on neural network compensation. A systematic derivation on the dynamics of an unmanned small-scale helicopter is performed. Combined rotor-fuselage-tail dynamics is derived in body-fixed reference frame with its origin at the C.G. of the helicopter. And the resulting nonlinear equation of motion consists of 6-DOF air vehicle dynamics as well as the rotor flapping and engine torque equations. A simulation model was modified using the existing simulator for an unmanned helicopter dynamic model, which reflects the unmanned test helicopter(CNUHELI). The dynamic response of the refined model was compared with the flight test data. It can be shown that a good coincidence was accomplished between the real unmanned helicopter system and the mathematical model. This dynamic model was linearized for classical controller design using small perturbation method. A Neuro-PD control system was designed for both longitudinal and lateral flight modes, and the results were compared with the PD-only control response. Simulation results show that the proposed Neuro-PD control system demonstrates better performance.

• 제어로봇시스템학회:학술대회논문집
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• 2003.10a
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• pp.519-524
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• 2003

In this report, a longitudinal adaptive flight control law is presented for the automatic landing system of a Japanese automatic landing flight experiment vehicle (ALFLEX). The longitudinal adaptive flight control law is designed to track an output of the vehicle to a guidance signal from the guidance portion of the automatic landing system. The proposed adaptive control law in the attitude control portion adjusts the controller gains continuously online as flight conditions change, in spite of the existence of unmodeled dynamics. The number of the controller gains to be adjusted is decreased to 1/2 from the previous studies. Computer simulation involving six-degree-of-freedom (DOF) nonlinear flight dynamics is performed to examine the effectiveness of the proposed adaptive control law. In order to verify the influence of the dispersion of the initial conditions, the Monte Carlo simulation is also applied. The initial conditions are more widely dispersed than the previous studies. As a result, except under the unsuitable initial conditions, the ALFLEX successfully landed on the runway.

• Journal of Biosystems Engineering
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• v.36 no.6
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• pp.476-483
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• 2011

The aerial application using an unmanned helicopter has been already utilized and an attitude controller would be developed to enhance the operational convenience and safety of the operator. For a preliminary study of designing flight controller, a state space model for an RC helicopter would be identified. Frequency sweep flight tests were performed and time history data were acquired in the previous study. In this study, frequency response of the flight test data of a small unmanned helicopter was analyzed by using the CIFER software. The time history flight data consisted of three replications each for collective pitch, aileron, elevator and rudder sweep inputs. A total of 36 frequency responses were obtained for the four control stick inputs and nine outputs including linear velocities and accelerations and angular velocities in 3-axis. The results showed coherence values higher than 0.6 for every primary control inputs and corresponding on-axis outputs for the frequency range from 0.07 to 4 Hz. Also the analysis of conditioned frequency response showed its effectiveness in evaluating cross coupling effects. Based on the results, the dynamic characteristics of the model helicopter can further be analyzed in terms of transfer functions and the undamped natural frequency and damping ratio of each critical mode.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.2
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• pp.173-180
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• 2014

In this paper, the instrument flight certification process and flight test results of Korean Utility Helicopter (KUH) are presented. For the instrument flight certification, the suitability of installed equipments and instruments have been reviewed and verified by ground and flight tests. Next, static and dynamic stability test are conducted in accordance with FAR-29 Appendix B. The static stability is determined by the change of speed and attitude according to control inputs. The dynamic stability is evaluated by how quickly the response of the helicopter due to long and short period control inputs are decreased. The pilot workload evaluation are also carried out by simulated IMC flight tests. This paper presents the workload assessment results when some failures are occurred at cockpit instruments, engine or flight control systems as well as the normal situation. After the simulated IMC flight test is completed, actual instrument flight test are conducted in a real IMC environment according to the air traffic controls.

• Journal of Electrical Engineering and Technology
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• v.7 no.4
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• pp.615-625
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• 2012

GPS attitude outputs or carrier phase observables can be effectively utilized to compensate the attitude error of the strapdown inertial navigation system. However, when the integer ambiguity is not correctly resolved and/or a cycle slip occurs, an erroneous GPS output can be obtained. If the erroneous GPS output is applied to the attitude determination GPS/INS (ADGPS/INS) integrated navigation system, the performance of the system can be degraded. This paper proposes an ADGPS/INS integration system using the triple difference carrier phase observables. The proposed integration system contains a cycle slip detection algorithm, in which the inertial information is combined. Computer simulations and flight test were performed to verify effectiveness of the proposed navigation system. Results show that the proposed system gives an accurate and reliable navigation solution even when the integer ambiguity is not correctly resolved and the cycle slip occurs.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.21 no.4
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• pp.34-40
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• 2013

Attitude command response type required for enhanced handling qualities of helicopter can be implemented by mechanical automatic flight control system with SAS actuators which have limited authorities. However, the early saturation of SAS actuator hinders the helicopter from following the attitude command for large stick command. Auto-trim controller can delay SAS actuator's saturation by utilizing trim actuators and allows the attitude command response type for larger stick command. This paper describes the control law for limited authority system of helicopter with auto-trim. Limited authority system is applied to BO-105 linear dynamic model and simulation is performed along with handling quality analysis.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.6
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• pp.455-462
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• 2017

In recent decades, many researchers have been struggling to developing the compound aircraft that is capable high speed and VTOL flight. And in recent years, multi-copters are very popular because of having advantages of VTOL and easy handling, but they are lack of doing long-range mission. Therefore, we presents simple aircraft architecture which is equipped fixed wing, multi propellers and no control surfaces. In this paper, we designed the attitude control for the compound aircraft prototype and measured the attitude control performance with flight test for validating prototype's performance. We analysed the attitude control test result comparing with similar size of a fixed wing aircraft. The performance was almost same as fixed wing aircraft.