• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.6
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• pp.473-480
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• 2013

Researches for various lunar landing technologies are in progress for the lunar exploration program planned for early 2020s in Korea. This paper shows the integrated simulation for safe lunar landing guidance/control system in powered descent phase. Generally, the lunar lander uses on/off(bang-bang) controller to control the RCS jet thrusters instead of proportional controller. In this paper, the on/off controller using phase-plane switching function, and thruster selection algorithm to control sixteen thrusters are applied. Also additional guidance commands are calculated by a proposed fuzzy logic guidance algorithm. The simulation results show that lunar lander can follow a reference trajectory which is generated by optimization method, then land on the surface safely.

• Advances in aircraft and spacecraft science
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• v.4 no.5
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• pp.503-514
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• 2017

The attitude control of an aircraft is usually fulfilled by means of thrusters at high altitudes. Therefore, the possibility of using also aerodynamic surfaces would produce the advantage of reducing the amount of fuel for the thrusters to be loaded on board. For this purpose, Zuppardi already considered some aerodynamic problems linked to the use of a wing flap in a previous paper. A NACA 0010 airfoil with a trailing edge flap of 35% of the chord, in the range of angle of attack 0-40 deg and flap deflections up to 30 deg was investigated. Computer tests were carried out in hypersonic, rarefied flow by a direct simulation Monte Carlo code at the altitudes of 65 and 85 km of Earth Atmosphere. The present work continues this subject, considering the same airfoil and free stream conditions but two flap extensions of 45% and 25% of the chord and two flap deflections of 15 and 30 deg. The main purpose is to compare the influence of the flap dimension with that of the flap deflection. The present analysis is carried out in terms of: 1) percentage variation of the global aerodynamic coefficients with respect to the no-flap configuration, 2) increment of pressure and heat flux on the airfoil lower surface due to the Shock Wave-Shock Wave Interaction (SWSWI) with respect to the same quantities with no SWSWI or in no-flap configuration, 3) flap hinge moment. Issues 2) and 3) are important for the design of the mechanical and thermal protection system and of the flap actuator, respectively. Under the above mentioned test and geometrical conditions, the flap deflection is aerodynamically more effective than the flap extension, because it involves higher variation of the aerodynamic coefficients. However, tests verify that a smaller deflection angle involves the advantage of a smaller increment of pressure and heat flux on the airfoil lower surface, due to SWSWI, as well as a smaller hinge moment.

• Journal of Astronomy and Space Sciences
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• v.26 no.1
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• pp.31-46
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• 2009

An Unscented Kalman Filter (UKF) for estimation of the attitude and rate of a spacecraft using only magnetometer vector measurement is developed. The attitude dynamics used in the estimation is the nonlinear Euler's rotational equation which is augmented with the quaternion kinematics to construct a process model. The filter is designed for small satellite in low Earth orbit, so the disturbance torques include gravity-gradient torque, magnetic disturbance torque, and aerodynamic drag torque. The magnetometer measurements are simulated based on time-varying position of the spacecraft. The filter has been tested not only in the standby mode but also in the detumbling mode. Two types of actuators have been modeled and applied in the simulation. The PD controller is used for the two types of actuators (reaction wheels and thrusters) to detumble the spacecraft. The estimation error converged to within 5 deg for attitude and 0.1 deg/s for rate respectively when the two types of actuators were used. A joint state parameter estimation has been tested and the effect of the process noise covariance on the parameter estimation has been indicated. Also, Monte-Carlo simulations have been performed to test the capability of the filter to converge with the initial conditions sampled from a uniform distribution. Finally, the UKF performance has been compared to that of the EKF and it demonstrates that UKF slightly outperforms EKF. The developed algorithm can be applied to any type of small satellites that are actuated by magnetic torquers, reaction wheels or thrusters with a capability of magnetometer vector measurements for attitude and rate estimation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.262-270
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• 2004

Basic experiments were carried out using the THT-IV low-power Hall thruster to examine the influences of magnetic field shape and strength, and acceleration channel length on thruster performance and to establish guidelines for design of high-performance Hall thrusters. Thrusts were measured with varying magnetic field and channel structure. Exhaust plasma diagnostic measurement was also made to evaluate plume divergent angles and voltage utilization efficiencies. Ion current spatial profiles were measured with a Faraday cup, and ion energy distribution functions were estimated from data with a retarding potential analyzer. The thruster was stably operated with a highest performance under an optimum acceleration channel length of 20 mm and an optimum magnetic field with a maximum strength of about 150 Gauss near the channel exit and with some shape considering ion acceleration directions. Accordingly, an optimum magnetic field and channel structure is considered to exist under an operational condition, related to inner physical phenomena of plasma production, ion acceleration and exhaust plasma feature. A new Hall thruster was designed with basic research data of the THT-IV thruster. With the thruster with many considerations, long stable operations were achieved. In all experiments at 200-400 V with 1.5-3 mg/s, the thrust and the specific impulse ranged from 15 to 70 mN and from 1100 to 2300 see, respectively, in a low electric power range of 300~1300 W. The thrust efficiency reached 55 %. Hence, a large map of the thruster performance was successfully made. The thermal characteristics were also examined with data of both measured and calculated temperatures in the thruster body. Thermally safe conditions were achieved with all input powers.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.6
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• pp.581-586
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• 2008

The MEMS solid propellant thrusters have very low thrust level for applying to the propulsion system of micro/nano satellites or the side jet thruster of smart bombs. In this research, the fabrication possibility of planar type MEMS solid propellant thrusters that have enlarged burning surface area was examined and the safety of the structure of thruster during the firing test was confirmed. The performance of a micro igniter which is the key component of the MEMS solid propellant thruster was estimated by the ANSYS Icepak and evaluated by the experiment. Finally, the thrust was measured by the micro force sensor. The levels of thrust were 300, 600 mN in the case of K=15, 20.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.90-100
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• 2019

Hypergolic thrusters have been extensively researched and applied to spacecraft propulsion based on their simplicity and high reliability of ignition. Research on the impingement characteristics of $N_2O_4$/amine has been profoundly carried out since the 1960s in advanced countries, especially the United States. Recently, enhancements to advanced hypergolic thrusters using MON/MMH have been planned by NASA to improve compactness and high performance. In this work, technical studies were investigated on the mixing of hypergolic propellant and its combustion instabilities such as reactive separation flow and popping.

• Journal of Space Technology and Applications
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• v.4 no.1
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• pp.62-73
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• 2024

This study analyzed the effectiveness of orbital change through attitude change in nano-satellites operating in low Earth orbit (LEO) without thrusters, focusing on collision avoidance maneuvers. The results revealed that changes in the satellite's cross-sectional area significantly impact its in-track direction, influenced by the aspect ratio of cross-sectional area change and mission altitude. Notably, satellites at lower altitudes demonstrated significant reduction in collision risks with a small amount of attitude change. Through this study, it is judged that the changing the cross-sectional area through attitude maneuver is a sufficiently suitable method in the operation of nano-satellites without thrusters, and is expected to contribute to improving the safety of satellite operations in the New Space era.

• Journal of Ship and Ocean Technology
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• v.11 no.2
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• pp.1-9
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• 2007

This paper describes the design and development of a hovering AUV constructed at Cheju National University and analyses the dynamic performance of the vehicle using simulation programs. The main purpose of this AUV is to carry out fundamental tests in its station keeping, attitude control, and desired position tracking. Its configuration is similar to the general ROV appearance for underwater works and its dimensions are 0.75m*0.5m*0.5m. It has 4 thrusters of 450 watts for longitudinal/lateral/vertical propulsion and is equipped with a pressure sensor for measuring water depth and a magnetic compass for measuring heading angle. The navigation of the vehicle is controlled by an on-board Pentium III-class computer, which runs with the help of the Windows XP operating system. These give us an appropriate environment for developing various algorithms needed for developing and advancing Hovering AUV.

• IEMEK Journal of Embedded Systems and Applications
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• v.12 no.5
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• pp.277-286
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• 2017

This paper introduces mathematical modelling and control algorithm of expendable mine disposal vehicle. This vehicle has two longitudinal thrusters, one vertical thruster and internal mass moving system which can control pitch rate. Also, the vehicle has an optical camera and forward looking sonar for underwater mine detection and classification. The vehicle is controlled via an optical cable connected with operating console on the mother ship. We describe the vehicle's 6DOF dynamic model and controller which can track the desired trajectory for the way-point tracking. These simulation results shows guidance and maneuvering performance which has other sensor data or not.

• International Journal of Aeronautical and Space Sciences
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• v.17 no.3
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• pp.378-390
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• 2016

A concept of the perch landing assisted by thruster (PLAT) for a fixed wind aircraft is proposed in this paper. The proposed concept is applicable to relatively large unmanned aerial vehicles (UAV), hence can overcome the limitation of existing perch landing technologies. A planar rigid body motion of an aircraft with aerodynamic and thruster forces and moments is modeled. An optimal control problem to minimize the fuel consumption by determining the histories of thruster and elevator deflection angle with specified terminal landing condition is formulated and solved. A parametric study for various initial conditions and thruster parameters is conducted to demonstrate the practicability of the proposed concept.