한국과학기술원 도서관

서지주요정보
Low thrust resonance-orbit-based multiple gravity-assist trajectory = 공명궤도 기반 저추력 다중 중력도움 궤적
서명 / 저자	Low thrust resonance-orbit-based multiple gravity-assist trajectory = 공명궤도 기반 저추력 다중 중력도움 궤적 / Jinsung Lee.
발행사항	[대전 : 한국과학기술원, 2024].
Online Access	원문보기 원문인쇄

소장정보

등록번호

8043060

소장위치/청구기호

학술문화관(도서관)2층 학위논문

DAE 24021

휴대폰 전송

도서상태

이용가능(대출불가)

사유안내

반납예정일

리뷰정보

초록정보

This paper presents a methodology for optimizing a low-thrust gravity assist trajectory to achieve the Earth-Moon L1 periodic orbit, leveraging the resonance-orbital structure as a guiding framework. The formulation employed in this study utilizes the Earth-Moon circular restricted three-body problem to describe the system’s dynamics. The proposed optimization procedure involves the determination of the gravity-assist geometry and the subsequent identification of the gravity-assist linking through the solution of a multiple-point boundary value problem. The periapsis rotation angle is designed by solving a gradient descent optimization problem in the gravity-assist geometry determination step. This process results in trajectories that intentionally deviate from the symmetry observed in resonance orbits. The multiple-point boundary value problem focuses on resolving a minimum-fuel challenge by connecting two intermediate resonance-like orbits characterized by rotated periapses. The optimal control problem unfolds in two steps. Firstly, a relatively straightforward two-point boundary problem, serving as an approximation to the original problem, is established and solved. The solution obtained from this step serves as the initial guess for the subsequent and more intricate multiple-point boundary value problem. To assess the efficacy of the proposed methodology, the low-thrust resonance gravity-assist trajectory is compared against trajectories designed using conventional approaches involving low-thrust propulsion. This comparative analysis aims to validate and highlight the efficiency of the proposed method in optimizing trajectories for space missions within the Earth-Moon system. The proposed methodology converged to a fuel-optimal solution that is 40% more efficient than the previous research. In addition to the LTRGA trajectory, possible extended missions to the Low-Lunar Orbit (LLO), and hyperbolic escape trajectory from Earth is discussed.

이 논문은 지구-달 시스템의 L1 주기궤도에 위성 투입을 위한 공명궤도 기반 저추력 다중 중력 스윙바이 달 탐사 궤적을 소개한다. 지구-달 원주제한삼체문제를 기반하여 궤적을 설계 및 최적화 하였다.저추력 다중 중력 스윙바이는 중력 스윙바이 최적 기하를 계산하는 문제와 다중점 경계값 최적화 문제를 기반으로 설계된다. 최적 기하 결정 단계는 그래디언트 감소 최적화 문제를 활용하여 근지점을 회전시키며 공명궤도의 대칭성을 깨트려 중력 스윙바이 궤적의 성능을 이끌어낸다. 다음, 최적의 중력 스윙바이 기하를 초기와 말기 경계값으로 사용하여 두 공명궤도를 연결하는 최소 연료 문제를 풀게된다. 최적 제어 문제의 첫 번째 단계는 원래 문제를 근사화한 상대적으로 쉬운 두 점 경계값 문제를 사용한다. 저추력 공명궤도 기반 궤도는 다양한 L1 주기궤도 전이방법론과 비교되었다. 현재 알려져있는 최소연료 최저값보다 약 40% 의 추진제를 절약하는 궤도를 찾을 수 있었다.

서지기타정보

서지기타정보
청구기호	{DAE 24021
형태사항	vii, 105 p. : 삽도 ; 30 cm
언어	영어
일반주기	저자명의 한글표기 : 이진성 지도교수의 영문표기 : Jaemyung Ahn 지도교수의 한글표기 : 안재명 Including appendix
학위논문	학위논문(박사) - 한국과학기술원 : 항공우주공학과,
서지주기	References : p. 99-102
주제	지구-달 시스템 공명궤도 다중 중력 스윙바이 저추력 궤도 최적화 Trajectory optimization Earth-Moon system Resonance orbits Multiple gravity assist

QR CODE

책소개

전체보기

이 주제의 인기대출도서

표/그림

모든 표/그림 보기

Trajectory types from Earth to the Moon with various measurement of merits

Pareto optimal diagram of the trajectory types shown in Table 1.1. WSB: Weak Stabil- ity Boundary trajectory. CLT: Conventional Low-Thrust Trajectory. LTRGA: Low-Thrust Resonance Gravity Assist Trajectory.

Three-body problem diagram

Circular restricted three-body problem coordinate frame.

Zero velocity curve opening at Earth-Moon L1 with decrease in C

Four different Hill's Regions with trajectories

Earth-Moon system's Lagrange points plotted above the Sun-Earth Lagrange point

Figure ofEq. 2.46. The roots of the curves show the location of the Lagrange points.

Solar system planet's relative mass to the Sun and their absolute masses Planets Mass (relative to Sun xE - 6) Absolute mass (kg)

Solar system planet's Moon's relative mass to their central planet and their absolute masses Moons Mass (relative to Central planet Absolute mass (kg)

Periodic orbits of the circular-restricted three-body problem environment. A shows the Lyapunov orbit at Lagrange points 1 and 2. B shows the northern and southern Halo orbit of the Lagrange points 1 and 2. C shows the vertical and ecliptic orbit ofthe Lagrange points 4 and 5.

L1 periodic orbit with trajectory line color matching the Jacobi integral level

L2 periodic orbit with trajectory line color matching the Jacobi integral level

Manifold trajectories of Earth-Moon L1 and L2 periodic orbits. Blue and red trajectories show the stable and unstable manifolds. W defines the manifold groups. Superscript (R,L),(S,U) defines right,left,stable, and unstable.

A) Sun-Earth system's manifold-based trajectories ofL1/2and L4/5 (stable (blue), unstable (red), low-thrust arc (green)), B) Magnified view of WS and WU from the periodic orbit ofL1 and L2 C) WU from L2, WS from L5 and low-thrust arc from L5, D) WU from L1, WS from L4, and low-thrust arc from L4,

Interior (red) and exterior (black) resonance orbits. Thefigure shows resonance orbits with multiple levels of Jacobi integral.

Resonance orbit ratio and trajectory image with Jacobi integral of 3.05. The trajectory from top to bottom and from left to right shows resonance orbit with larger resonance orbit ratio, posing larger osculation orbital elements' semi-major axis

External resonance orbit visualization

Differential correction error function visualization

Differential correction visualization for correcting the halo orbit initial conditions

Differential correction visualization for correcting N:M=1:3 resonance orbit's initial condi- tions. Thisfigure shows the initial orbit from the Tisserand parameter in the two-body problem in red (14 = 0) to the Earth-Moon system'S mass ratio in blue, which is tabulated in (table 2.2). Dashed line trajectories are the trajectory solutions during mass parameter continuation.

Time dependent boundary conditions for solving two-point boundary value problems

Poincare section with Manifold Traces

Poincare filtering algorithm

Poincare filtering algorithm applied for transfer from L1 to L2 with 600 rotation from x-axis

Poincare filtering algorithm applied for transfer from L1 to L2 with 90o rotation from x-axis

Poincare filtering algorithm applied for transfer from L1 to L2 with 120o rotation from x-axis

Keplerian map of 1:3 resonance orbit with different Jacobi integral and various other resonance orbit's semi-major axis level

Resonance orbit accessibility chart for four different Jacobi integral levels (C1,02,03, and 04/5);r- reachable, u - unreachable; orange cells used in Case A

Tisserand diagram and keplerian map of the multi-moon and Earth-Moon systems, respec tively.

A) Tisserand-Poincare diagram with reachability surface. B) Keplerian diagram wit reachability surface. C.a) Flyby geometry in rotating reference frame for resonance N:M. C.b) Flyb geometry in rotating reference frame for resonance N:M. D) Flyby orbit change was seen in an inerti frame ofthe central body.

Periapsis Map (central body: Earth) of stable EM-L1 manifold and resonance orbit with various N:M values plotted in Earth-Moon rotating frame

Difference in Jacobi integrals (AC, expressed in color) associated with L1 manifold (gray- scale level curve) and resonance orbits (red dots) plotted in Periapsis Map

AC - rp map generation procedure. The black diamonds are the intersection line 0 the manifold and resonance orbit with the Jacobi integral of 3.12 and 3.15, which are the termina boundary conditions for solving the trajectory optimization from the terminal resonance orbit to the stable manifold. 57

Optimization algorithm for low-thrust resonance gravity-assist trajectory.

N:M = 1:3 (blue) and 2:5 (magenta) resonance orbits with optimal gravity assist trajectory (dashed black); eopt= -4.45 deg (angle enlarged for visualization)

Keplerian map of 1:3 resonance orbit with varying C; semi-major axis is defined with the Earth as the central body

Switching function, optimal steering law, and the throttle level (E= 1,0.3, and 1e-3); Time Unit (TU) defined in Table 2.2

Sequence ofresonance orbit structures (N:M=1:3, 2:5, 1:2) and the stable manifold of the desired L1 periodic orbit.

First low-thrust trajectory optimization problem. The following optimal control problem links the first two gravity assist maneuvers (x1(113 →2: 5) and xb(2:5→1:2))

Second low-thrust trajectory optimization problem. The following optimal control problem links the first two gravity assist maneuvers (x8(2 :5 →1:2) and xb(1:2 → Wii))

Optimization procedure of the first optimal control problem shown in figure 5.6

Parameters, initial and terminal boundary conditions, and canonical units

Results of LTRGA of Earth-Moon transfer solutions

Initial conditions for the test-case C shown in Figure 5.11

Profiles of orbital characteristics for LTRGA trajectory solution A (GTO to EM-L1); all elements are measured with respect to the Earth

Representative Pareto optimal trajectories (solutions A, B, C and D presented in Figure

Time expended trajectory diagram for optimizing the trajectory to Low-Lunar orbit from interior resonance orbits.

Earth-Moon L1 periodic orbit's manifold propagated until the state resembles periapsis with respect to the Moon. Thin lines are the manifold trajectories. Bold lines are the manifold trajectories with an altitude of 100km from the Moon's surface.

Trajectory from multiple resonance orbits to stable-unstable manifold connection atL1,and ultimately to Low-lunar orbit via Q-law guidance. The red trajectory shows the LTRGA region of the trajectory. Blue trajectory shows the apoapsis decrease in the Moon's orbit. The blue circle shows the terminal boundary condition for a stable-unstable connection.

Magnified view of the trajectory from figure 6.3. The red trajectory shows the LTRGA region of the trajectory. Blue trajectory shows the apoapsis decrease in the Moon's orbit. The blue circle shows the terminal boundary condition for a stable-unstable connection.

Full trajectory (figure 6.3) plotted in inertial frame. The red trajectory shows the LTRGA region ofthe trajectory. Blue trajectory shows the apoapsis decrease in the Moon's orbit. The black line shows the orbit ofthe Moon.

Time expanded diagram of the trajectory sequence for Earth escape trajectory from initial resonance orbits. The red trajectory on the left shows the N:M=1:3 resonance orbit, green shows 2:5 and magenta shows a 1:2 resonance orbit. black lines are the manifold of the desired trajectory and heteroclinic orbits in between Earth-Moon L1 and L2. The yellow trajectory shows the unstable manifold from Ea

Keplerian map ofthe external resonance orbits with C3 energy. The colors of the plot show the differences in the Jacobi integral.

Earth escape trajectory plotted in the Earth-Moon rotating frame. The red trajectory or the left shows the N:M=1:3 resonance orbit, green shows 2:5, and magenta shows a 1:2 resonance orbit black lines are the manifold of the desired trajectory and heteroclinic orbits in between Earth-Moon L1 and L2. The yellow trajectory shows the unstable manifold from Earth-Moon L2, the darker green shows the ex

Earth escape trajectory plotted in the Earth-Moon rotating frame, magnified near the Moor region to visualize the complex transfer phase between Earth-Moon L1 and L2. The red trajectory 01 the left shows the N:M=1:3 resonance orbit, green shows 2:5, and magenta shows a. 1:2 resonance orbit black lines are the manifold of the desired trajectory and heteroclinic orbits in between Earth-Moon L1 and L

Earth escape trajectory plotted in the Earth-centered inertial frame. The red trajectory on the left shows the N:M=1:3 resonance orbit, green shows 2:5, and magenta shows a 1:2 resonance orbit. black lines are the manifold of the desired trajectory and heteroclinic orbits in between Earth-Moon L1 and L2. The yellow trajectory shows the unstable manifold from Earth-Moon L2, the darker green shows t