Dong-Hyun Cho

Optimal Perilune Altitude of Lunar Landing Trajectory

In general, the lunar landing stage can be divided into two distinct phases: de-orbit and descent, and the descent phase usually comprises two sub-phases: braking and approach. And many optimization problems of minimal energy are usually focused on descent phases. In these approaches, the energy of de-orbit burning is not considered. Therefore, a possible low perilune altitude can be chosen to save fuel for the descent phase. Perilune altitude is typically specified between 10 and 15km because of the mountainous lunar terrain and possible guidance errors. However, it requires more de-orbit burning energy for the lower perilune altitude. Therefore, in this paper, the perilune altitude of the intermediate orbit is also considered with optimal thrust programming for minimal energy. Furthermore, the perilune altitude and optimal thrust programming can be expressed by a function of the radius of a parking orbit by using continuation method and co-state estimator.

A PIV Study of Flow Patterns Over Stationary and Pitch-Oscillating Airfoils with Blowing Jet

A particle image velocimetry (PIV) technique was employed to investigate the effects of blowing jet on the flow characteristics over stationary and pitch-oscillating airfoils. The Reynolds number was 7.84×10? based on the chord length. It was found that for stationary airfoil cases, continuous and pulsating blowing jets successfully reduced separated wake region at high angles of attack. A comparison study of two different types of jet blowing indicated that pulsating jet is more effective than continuous jet for flow separation control. Pulsating leading-edge blowing postpones flow separation and increased stall angle of attack by 2°-3°. For pitch-oscillating airfoil cases, the PIV results showed that blowing jet efficiently delays the separation onset point during pitch-up stroke, whereas it does not prevent flow separation during pitch-down stroke, even at angles of attack smaller than static ones.

Satellite Trajectory Correction Maneuver for Lunar Mission based on Three-Body Dynamics

ABSTRACT During the lunar mission, spacecraft are subject to various unexpected disturbance sources such as third body attraction, solar pressure and operating impulsive maneuver error. Therefore, efficient trajectory correction maneuver (TCM) strategy must be required to follow the designed mission trajectory. In the early days of space exploration, the mission trajectory has been designed by using patched conic approach based on two-body dynamics for the lunar mission. Thus the TCM based on two-body dynamics has been usually adopted. However, with the advanced in computing power, the mission trajectory based on three-body dynamics is attempted recently. Thus, these approaches based on two-body dynamics are essentially different from real environment and large amount of energy for the TCM is required. In this work, we study the trajectory correction maneuver based on three-body dynamics. 초 록 달탐사 임무를 수행하는 인공위성의 경우 임무수행을 하는 과정에서 3체에 의한 인력, 태양풍 그리고 추력시스템의 추력오차 등의 많은 예기치 못한 외부 섭동력에 영향을 받게 된다. 따라서 주어진 임무궤도를 따라서 인공위성이 운영되기 위해서 궤적 보정 기동이 필요하다. 우주 탐사의 초창기 시절에는 이러한 임무궤도는 주로 2체 운동방정식에 기반을 한 패치 코닉(Patched Conic)기법으로 생성을 하였으며, 이로 인해 2체 운동방정식에 기반을 한 궤적 보정 기동이 많이 사용되어져 왔다. 하지만 최근 컴퓨터 연산능력의 향상에 기인하여 이러한 임무궤도를 지구-인공위성-달의 3체 운동방정식에 기반하여 설계하고 있는 추세이다. 따라서 기존의 2체 운동방식 기반의 궤적 보정 기동으로는 실제 우주환경과 많은 차이를 보이기 때문에 달의 작용권구(Sphere of Influence)에 접근할수록 많은 궤도오차를 보이며, 이를 보정하기 위해서 많은 에너지가 필요하게 된다. 따라서 본 논문에서는 3체 운동방정식에 기인한 궤적 보정 기동에 대하여 기술하고자 한다.Key Words : Trajectory Correction Maneuver(궤적 보정 기동), Lunar Mission(달탐사 임무), Three Body Dynamics(3체 운동방정식)

Experimental Study for the Safety Analysis of an External Store Separation from Fighter Aircraft

The prediction of the separation trajectories of external stores carried by military aircraft is an important task in the area of aircraft design having the objective to define the operational, release envelopes. This paper presents the results obtained for safe store separation from a fighter aircraft by experimental methods in the subsonic wind tunnel. The problems associated with separation of external stores can be studied by the use of several wind-tunnel test techniques. Attention is given the two most useful techniques: 1) dynamically scaled drop-model testing, 2) grid testing. A description of each method is given and data obtained are shown to validate the similarity within acceptable limits.

Flow Separation Control Effects of Blowing Jet on an Airfoil

An experimental study has been conducted to investigate the flow separation control effects of a blowing jet on an elliptic airfoil at a Reynolds number of 7.84×105 based on the chord length. A blowing jet was obtained by pressing a plenum inside the airfoil and ejecting flow out of a thin jet slot that located in leading edge or trailing edge. The experimental results have shown that the blowing jet had an effect of suppressing the flow separation, resulting in the higher suction pressure distribution and higher normal force. The increase in Cn was more pronounced at higher incidence, whereas the effectiveness of the blowing jet reduced at lower incidences. The leading edge pulsating blowing with 90° was the most effective in controlling the flow separation than other types of blowing jet configuration tested in this research. Moreover, when the pulsating blowing was applied, the stall angle was postponed about 2°-3°. The continuous and pulsating blowing jet is a direct and effective flow separation control for improving the aerodynamic characteristics and performances of airfoil.An experimental study has been conducted to investigate the flow separation control effects of a blowing jet on an elliptic airfoil at a Reynolds number of 7.84×10? based on the chord length. A blowing jet was obtained by pressing a plenum inside the airfoil and ejecting flow out of a thin jet slot that located in leading edge or trailing edge. The experimental results have shown that the blowing jet had an effect of suppressing the flow separation, resulting in the higher suction pressure distribution and higher normal force. The increase in C n was more pronounced at higher incidence, whereas the effectiveness of the blowing jet reduced at lower incidences. The leading edge pulsating blowing with 90° was the most effective in controlling the flow separation than other types of blowing jet configuration tested in this research. Moreover, when the pulsating blowing was applied, the stall angle was postponed about 2°~3°. The continuous and pulsating blowing jet is a direct and effective flow separation control for improving the aerodynamic characteristics and performances of airfoil.

Orbit Transfer Trajectory Optimization With Electric Engine

In this paper, it is studied optimal coplanar orbit transfer of a satellite from a low Earth orbit to a low Earth orbit using electric propulsion system. In this case the satellite is passing through the shadow area of the Earth. During this time, supplying electric power to the satellite is going to be difficult due to the solar array which is under insufficient sun power. Henceforth, the Earths shadow effect should be taken into account for obtaining optimal trajectory for orbit transfer. The optimization method used in this paper is indirect method under the assumption of the initially circular orbit model. And the fact of turning off the electric propulsion system in the earth shadow area mean that it is possible to ON/OFF control of electric propulsion system. Therefore, in this study, the variable thrust system is also considered with direction control of thrust vector.

Optimal Phase Angle Design for the Lunar Lander

To design the optimal lunar landing trajectory, the periapsis of de-orbit burn phase is usually used as a starting point of the powered descent phase. And the optimal problem is constructed based on the initial states at this point. For this optimal problem, some kinds of result trajectories have the increase in their own altitude at the early stage to earn the enough time to reduce their huge horizontal velocity for the low thrust, and this phenomenon increase the consumption of the propellant mass. For this reason, it is possible to exist another phase angle to reduce this phenomenon instead of zero for periapsis of de-orbit burn phase because the vertical velocity is not zero at this point. To find out this optimal phase angle, in this paper, the initial free state optimal problem method is applied without any additional assumption and change of hardware for the traditional optimal lunar landing problem. Using this approach, it shows that the optimal phase angle does not always equal to the periapsis for some cases, and it is possible to reduce the fuel consumption of the lunar lander.

Analysis of Space debris collision risk using KARISMA for KOMPSAT Satellite series

Korea has successfully launched multipurpose satellites such as KOMPSAT (KOrea Multi Purpose SATellite) in low Earth orbit (LEO) since 1999. Thus far, KOMPSAT-2, KOMPSAT-3, and KOMPSAT-5 are in operation, and KOMPSAT-3A will be launched this year. Korea also successfully launched COMS (Communication, Ocean, and Meteorological Satellite), its first geostationary satellite, in 2006. The Korea Aerospace Research Institute (KARI) is therefore now concerned about the risk posed by space debris to these satellites. In fact, a preliminary analysis of the collision risk for KOMPSAT-1 (which was left out owing to a communication system malfunction in 2008) and KOMPSAT-2 was performed in light of Chinese’s Anti Satellite Test (ASAT) in 2007. The development of a full-scale analysis system for collision risk at KARI has been further encouraged by the satellite collision accident between the Iridium 33 satellite and the defunct Cosmos 2251 satellite at an altitude of 770 km, which significantly increased the amount of space debris in the LEO region of space. With this background in mind, the KARI Space debris collision risk MAnagement system (KARISMA) has been in development since 2011. KARISMA was finalized in July 2013 and then tested for the KOMPSAT satellite series. This paper presents the analysis results for the space collision risk of the KOMPSAT satellites using KARISMA as well as the collision avoidance (COLA) maneuver planning with various maneuver strategies for several conjunction events. The characteristics and architecture of KARISMA are also discussed with detailed operational views. Among the various features of KARISMA, the efficient process for both the analysis and the management of the collision risk and a user-friendly UI including various 2D and 3D displays for the results and conjunction geometry are the most remarkable. However, KARISMA has been developed as standalone software. Therefore, we plan to expand and modify it for various users in Korea by considering a server-client concept as a further work.

Comparison of Global Optimization Methods for Insertion Maneuver into Earth-Moon L2 Quasi-Halo Orbit Considering Collision Avoidance

A spacecraft placed in an Earth-Moon L2 quasi-halo orbit can maintain constant communication between the Earth and the far side of the Moon. This quasi-halo orbit could be used to establish a lunar space station and serve as a gateway to explore the solar system. For a mission in an Earth-Moon L2 quasi-halo orbit, a spacecraft would have to be transferred from the Earth to the vicinity of the Earth-Moon L2 point, then inserted into the Earth-Moon L2 quasi-halo orbit. Unlike the near Earth case, this orbit is essentially very unstable due to mutually perturbing gravitational attractions by the Earth, the Moon and the Sun. In this paper, an insertion maneuver of a spacecraft into an Earth-Moon L2 quasi-halo orbit was investigated using the global optimization algorithm, including simulated annealing, genetic algorithm and pattern search method with collision avoidance taken into consideration. The result shows that the spacecraft can maintain its own position in the Earth-Moon L2 quasi-halo orbit and avoid collisions with threatening objects.

Trajectory correction maneuver design based on B-plane targeting for low-thruster

For the interplanetary mission, the spacecraft must require the trajectory correction maneuver (TCM) to reduce the trajectory error due to the influence of unexpected external perturbation forces. There are many kinds of researches focused on these TCM techniques for the impulsive thrust. Among them, the B-plane targeting method is mainly applied to the interplanetary mission as TCM method. However, there are brilliant developments in the propulsion systems; especially the electrical propulsion system is adopted in some spacecraft. Thus, it is necessary to design the TCM algorithm for such low-thruster for future interplanetary missions. Therefore, in this paper, the TCM algorithm for the low thrust such like electrical propulsion system is suggested to correct the target B-vector and target time to closest approach (TCA). To design the low thrust guidance scheme, the Lyapunov feedback control is used in this paper.