Lunar Transfer Research Articles

Mass Driver Design Traveling Earth to the Moon

In this article, the flight of a mass driver was designed for launch from the Earth with Electro Magnetic Space Launching System (EMSLS). Then the orbit exit from the Earth at 185 km and orbit entry the Moon at 100kmwere examined with respect to change of trajectories by using chemical fuel and the engine in the mass driver. Electromagnetically launched mass drivers should orbit with a specified orbital velocity at a designated altitude. In this paper, the energy is transferred externally to a mass driver throughout the flight path the electromagnetic coil system called multistage (EMSLS) designated in order to achieve the specified orbital velocity. The mass driver is synchronously accelerated by a voltage through the capacitors which are used for storing energy. This energy is transferred through a switching inductor to the circuit of the mass driver so that the mass driver is launched into the orbit with a muzzle velocity. However, this fact creates high air drag energy losses due to atmospheric conditions and high velocity obtained in EMSLS. Thus, in the mass driver at 21km altitude an engine starts to increase the velocity of the system to reach orbital velocity. The final aim of this article is to capture the transfer of Δv cost for traveling to the Moon. At any given arrival time in order to guide the system, designers only consider the gravity of the Earth and gravity of the Moon by using a Direct Lunar Transfer Trajectory for the Earth to the Moon approach. Finally, EMSLS was evaluated as a more advantageous and complimentary alternative to chemical propulsion systems for space transportation.

ロケットの近地点引数制約を考慮した実数値遺伝的アルゴリズムによる3インパルス月遷移軌道設計

ロケットの近地点引数制約を考慮した実数値遺伝的アルゴリズムによる3インパルス月遷移軌道設計

Lunar flyby transfers between libration point orbits

Lunar flyby transfers between libration point orbits

Mechanical Limitations of Motorised Tethers for Payload Orbital Transfer

In recent decades much research has focused on the use of space tether systems for payload orbital transfers. Orbit raising through the exchange of momentum using hanging tether systems in addition to enhanced momentum exchange resulting from motorised torque-driven systems have both been considered. In this paper the fo- cus is on the performance of symmetrically laden motorised momentum exchange tethers with the aim of conducting orbital transfers. These consist of two propulsion tethers symmetrically attached to a motor shaft with payloads, to be later trans- ferred, attached to their free ends. Rotation is achieved through reaction against two further tethers and masses attached to the motor stator with an estimated mass of this system, minus the propulsion tethers and payloads, of 1.5 tonnes. By deriv- ing the maximum tension acting on a system orbiting an oblate Earth expressions are obtained for the maximum rotational velocity that the tether systems are capa- ble of withstanding, based upon the material properties of the tethers themselves. Having obtained these expressions the performance of both hanging and motorised tether systems for conducting orbital transfers is analysed in terms of the gain in energy when the orbital geometry and tether length are varied. From this analysis the material-specifc tether length which gives the greatest gain in rotational kinetic energy to the payloads is established for a motorised system. Finally, by utilising this material-specific length, the minimum semi-major axis which is capable of per- forming lunar transfers and Earth escape trajectories is obtained.

Trajectory optimization for a lunar orbiter using a pattern search method

The Republic of Korea plans to launch a lunar orbiter and lander by 2020. There are several ways to enter lunar orbit: direct transfer, phasing loop transfer, weak stability boundary transfer, and spiral transfer trajectory. In this study, trajectory optimization is investigated for a lunar orbiter using a pattern search method that minimizes the required delta-V for direct lunar transfer. This method generates neighborhood points near the initial condition and then determines whether there is a new point that can reduce the value of the objective function. Classical methods require the gradient and acceleration of the objective function, but pattern search does not. Six poll methods and nine search methods are chosen; thus, 54 combinations of poll and search methods are available. The pattern search method can reduce the required delta-V on average by a few meters per second for a time of flight of five days and more than 10 m/s for a time of flight of four or six days, regardless of whether translunar injection is performed at the ascending or descending node.

Orbital design and optimal control of cislunar transport mission using spinning tether system

To transport payload into lunar transfer orbit using spinning tether system, reusable payload transport design were proposed and analysed. When final design was decided, open-loop reference control of payload transport stage was developed using optimal control theories. For the insufficient thrust case in LEO orbit, Hamiltonian on/off control was used to solve the problem. Numerical results indicate that the final transport design is able to transport payload into desired orbit, and the control law could satisfy mission requirement. Compared to conventional rocket transport, much less propellant is required by the spinning system with system reusability.

The effect of the Earth’s oblateness on predicting the shadow conditions of a distant spacecraft: Application to a fictitious lunar explorer

The effect of the Earth’s oblateness on predicting the shadow events of a lunar spacecraft caused by the Earth’s shadow is analyzed in this study. To ensure a reliable analysis, the proven ‘line-of-intersection’ method is modified and directly applied to predict the shadow conditions using a spheroidal model of the Earth and a conical shadow model. Two major lunar mission phases, namely, transfer and orbiting, are considered with corresponding fictitious initial conditions, and eclipse events are predicted and the results are compared using both spherical and spheroidal Earth models. For the lunar transfer phase, for which an Earth-bound highly elliptical orbit is assumed, not only the predicted entry and exit times of an event but also its duration are found to be more strongly shifted as the apogee altitude increases; for perigee and apogee altitudes of 1000 and 380,000km, respectively, the maximum difference in predicted duration is found to be approximately 0.76min for a penumbra event. For the lunar orbiting phase, for which a circular orbit around the Moon at an altitude of 100km is assumed, a prediction difference of approximately half a minute on average and approximately one minute at maximum (e.g., 0.73min for qumbra events, 1.03min for penumbra events and 1.32min for ‘instantaneous’ full sunlight events) can occur. The results of the present analysis highlight the importance of modeling the oblate shape of the Earth when predicting the shadow events of a distant spacecraft, and they are expected to provide numerous insights for any missions involving highly elliptical orbits around the Earth or travel to the Moon.

Analysis on Tracking Schedule and Measurements Characteristics for the Spacecraft on the Phase of Lunar Transfer and Capture

In this work, the preliminary analysis on both the tracking schedule and measurements characteristics for the spacecraft on the phase of lunar transfer and capture is performed. To analyze both the tracking schedule and measurements characteristics, lunar transfer and capture phases' optimized trajectories are directly adapted from former research, and eleven ground tracking facilities (three Deep Space Network sties, seven Near Earth Network sites, one Daejeon site) are assumed to support the mission. Under these conceptual mission scenarios, detailed tracking schedules and expected measurement characteristics during critical maneuvers (Trans Lunar Injection, Lunar Orbit Insertion and Apoapsis Adjustment Maneuver), especially for the Deajeon station, are successfully analyzed. The orders of predicted measurements' variances during lunar capture phase according to critical maneuvers are found to be within the order of mm/s for the range and micro-deg/s for the angular measurements rates which are in good agreement with the recommended values of typical measurement modeling accuracies for Deep Space Networks. Although preliminary navigation accuracy guidelines are provided through this work, it is expected to give more practical insights into preparing the Korea's future lunar mission, especially for developing flight dynamics subsystem.

Error Analysis and Trajectory Correction Maneuvers of Lunar Transfer Orbit

For a returnable lunar probe, this paper studies the characteristics of both the Earth-Moon transfer orbit and the return orbit. On the basis of the error propagation matrix, the linear equation to estimate the first midcourse trajectory correction maneuver (TCM) is figured out. Numerical simulations are performed, and the features of error propagation in lunar transfer orbit are given. The advantages, disadvantages, and applications of two TCM strategies are discussed, and the computation of the second TCM of the return orbit is also simulated under the conditions at the reentry time.

Low-energy lunar transfers using spatial transit orbits

This paper is concerned with natural and artificial low-energy lunar transfers in three-dimensional space. The main contribution of this paper is that the limitations of the planar manifold assumption, which is adopted in previous low-energy orbit design methods, are avoided by describing the transfer orbits with more realistic spatial transit and non-transit orbits. To start, the limitations of the previous design methods for the low-energy trajectories are highlighted, and the boundaries of the spatial transit orbits, which can enter into or escape from the potential well near the Moon through the L1 or L2 bottleneck regions of the zero velocity surface, are defined on a Poincaré section by using the necessary and sufficient condition of transition. Next, by considering the dominant gravity bodies in different orbit segments the motion near the Moon is analyzed in the Earth–Moon circular restricted three-body problem (CR3BP). For natural celestial bodies, the statistical characteristics of the lunar collision trajectories are studied. For the artificial celestial bodies, the investigation is focused on the achievable range of inclination and height of the low lunar orbit (LLO). Then, the motion between the Earth and the Moon is studied in the Earth–Moon based Sun-perturbed bicircular four-body problem (B4BP). For natural and artificial celestial bodies, the Earth-origin trajectories and the trajectories from the low Earth orbits are analyzed. Compared to the current planar manifold based design methods, the technique introduced in this paper can evaluate the lunar transfer orbits more accurately. Also, some lunar transfer trajectories which do not exist in the manifold based models can be found, and the heights and inclinations of the parking orbits around the Earth and the Moon can also be analyzed.

Comparison of low-energy lunar transfer trajectories to invariant manifolds

In this study, transfer trajectories from the Earth to the Moon that encounter the Moon at various flight path angles are examined, and lunar approach trajectories are compared to the invariant manifolds of selected unstable orbits in the circular restricted three-body problem. Previous work focused on lunar impact and landing trajectories encountering the Moon normal to the surface, and this research extends the problem with different flight path angles in three dimensions. The lunar landing geometry for a range of Jacobi constants is computed, and approaches to the Moon via invariant manifolds from unstable orbits are analyzed for different energy levels.

Pre-LOI trajectory maneuvers of the CHANG′E-2 libration point mission

This paper addresses pre-LOI (before Lissajous orbit insertion) trajectory maneuvers for the libration point mission of CHANG′E-2, which is the first Chinese satellite to fly to the Sun-Earth L2 libration point and the first satellite ever to fly to the L2 point from lunar orbit. First, a transfer trajectory for the mission is designed based on InterPlanetary Superhighway theory under the constraint of the remaining propellant, TT&C (telemetry, track and command) and sunlight conditions. The effects of trajectory maneuver errors on the mission are also analyzed. Second, based on the analysis results, the article investigates the trajectory maneuver schemes for the lunar escape and transfer trajectory maneuvers employing mathematical models and operational strategies. The mission status based on our maneuver schemes is then presented, and the results of control operation and corresponding analysis are provided in detail. Finally, the future trajectory maneuver in the Lissajous orbit is discussed. According to the analysis of the CHANG′E-2 TT&C data, the trajectory maneuver schemes proposed in the article work properly and efficiently.

Targeting Low-Energy Ballistic Lunar Transfers

Numerous low-energy ballistic transfers exist between the Earth and Moon that require less fuel than conventional transfers, but require three or more months of transfer time. An entirely ballistic lunar transfer departs the Earth from a particular declination at some time in order to arrive at the Moon at a given time along a desirable approach. Maneuvers may be added to the trajectory in order to adjust the Earth departure geometry or time in order to meet mission requirements. In this article, we develop a method to determine the locations of one or two maneuvers along a transfer such that a spacecraft may depart the Earth from a fixed declination, e.g., 28.5°, at a designated time and follow a low-energy transfer to a particular lunar arrival using a minimal amount of transfer ΔV. One practical application of this study is to characterize the launch period for a mission that aims to launch from a particular launch site, such as Cape Canaveral, Florida, and arrive at a particular orbit at the Moon on a given date using a three-month low-energy transfer. The targeting method developed here is applied to several example transfers with varying geometries. For the purposes of this article, the transfers are designed to target halo orbits about the Earth-Moon L2 point.

Conceptual Design of Korea Aerospace Research Institute Lunar Explorer Dynamic Simulator

In lunar explorer development program, computer simulator is necessary to provide virtual environments that vehicle confronts in lunar transfer, orbit, and landing missions, and to analyze dynamic behavior of the spacecraft under these environments. Objective of simulation differs depending on its application in spacecraft development cycle. Scope of use cases considered in this paper includes simulation of software based, processor and/or hardware in the loop, and support of ground-based flight test of developed vehicle. These use cases represent early phase in development cycle but reusability of modeling results in the next design phase is considered in defining requirements. A simulator architecture in which simulator platform is located in the middle and modules for modeling, analyzing, and three dimensional visualizing are connected to that platform is suggested. Baseline concepts and requirements for simulator development are described. Result of trade study for selecting simulation platform and approaches of defining other simulator components are summarized. Finally, characters of lunar elevation map data which is necessary for lunar terrain generation is described.

Using space manifold dynamics to deploy a small satellite constellation around the Moon

The possibility of communicating with the far side of the Moon is essential for keeping a continuous radio link with lunar orbiting spacecraft, as well as with manned or unmanned surface facilities in locations characterized by poor coverage from Earth. If the exploration and the exploitation of the Moon must be sustainable in the medium/long term, we need to develop the capability to realize and service such facilities at an affordable cost. Minimizing the spacecraft mass and the number of launches is a driving parameter to this end. The aim of this study is to show how Space Manifold Dynamics can be profitably applied in order to launch three small spacecraft onboard the same launch vehicle and send them to different orbits around the Moon with no significant difference in the Delta-V budgets. Internal manifold transfers are considered to minimize also the transfer time. The approach used is the following: we used the linearized solution of the equations of motion in the Circular Restricted Three Body Problem to determine a first–guess state vector associated with the Weak Stability Boundary regions, either around the collinear Lagrangian point L1 or around the Moon. The resulting vector is then used as initial state in a numerical backward-integration sequence that outputs a trajectory on a manifold. The dynamical model used in the numerical integration is four-body and non-circular, i.e. the perturbations of the Sun and the lunar orbital eccentricity are accounted for. The trajectory found in this way is used as the principal segment of the lunar transfer. After separation, with minor maneuvers each satellite is injected into different orbits that lead to ballistic capture around the Moon. Finally, one or more circularization maneuvers are needed in order to achieve the final circular orbits. The whole mission profile, from launch to insertion into the final lunar orbits, is modeled numerically with the commercial software STK.

Autonomous Midcourse Navigation for Lunar Return

Autonomous navigation systems provide the vehicle with estimates of its states without ground support. This work develops an autonomous navigation architecture for lunar transfer using optical sensors and celestial navigation. Measurement and errormodels are developed for two classes of celestial measurements, the elevation of known stars from the Earth’s or moon’s limb, and the apparent radius of the Earth or moon. Monte Carlo methods are used to support the development of measurement error models. The proposed architecture is tested with linear covariance techniques; navigation errors and trajectory dispersions are obtained to confirm the feasibility of the approach. The navigation system is required to provide 0.5 deg flight-path angle accuracy at entry interface for mission safety. The simulation results show that the proposed autonomous navigation systemmeets the reentry safety requirement.

Orbit Maneuver Compensation of KAGUYA for Its Safe and Accurate Lunar Transfer

Reported in this paper are the results of the orbit maneuver compensation in KAGUYA's Lunar transfer. Because of the uncoupled allocation of the attitude control thrusters, extra velocity increment (δv ) is induced whenever KAGUYA performs an orbit maneuver. Since the observed level of δv was unacceptable range from the point of maneuver accuracy requirement, it was compensated by means of deducting estimated δv from the orbit maneuver command. The δv estimation model was updated step-by-step during the Lunar transfer, which leaded to significant improvement of the orbit maneuver accuracy and resulted in the omission of the last trajectory correction maneuver. The method of the compensation and its results are introduced in detail.

Direct lunar halo orbit transfers

This paper provides a robust survey of short-duration, two-burn transfers from the Earth to lunar halo orbits. The results show that several families of such transfers exist, including quick four- to five-day transfers, as well as more efficient transfers that require three or more weeks. The trajectories surveyed in this paper all begin from 185-km circular Earth parking orbits, with varying inclinations; they transfer to many sample halo orbits about the Earth-Moon L1 and L2 points. It has been found that the quickest lunar transfers require between 3.6 and 4.1 km/s, depending on the halo orbit destination. If more time is available, then any lunar halo orbit considered in this paper may be reached for as little as 3.59 to 3.65 km/s. Finally, the paper discusses the costs and benefits of short and longer halo transfers applied to real lunar mission designs.

한국형 달탐사 임무 예비 설계 소프트웨어의 개발

향후 우리나라의 달탐사 임무에 대비하여 순간 추력을 이용한 한국형 달탐사 예비 임무 설계 소프트웨어를 개발하였다. 달 탐사 임무 수행을 위한 지구 출발 단계, 달 천이 단계, 달 도착 및 임무 수행 궤도 단계를 포함한 임무 설계가 이루어 졌다. 이 소프트웨어를 이용하면 순간 추력을 사용한 최적의 달탐사 비행궤적을 설계할 수 있다. 이를 바탕으로 우리나라의 우주 발사체인 KSLV-II를 사용할 때의 발사 가능한 달 탐사선의 최대 질량을 산출하여 보았다. 아울러 심우주 추적망을 이용하여 탐사선의 추적 가능 여부에 대한 해석이 이루어 졌으며 탐사선과의 통신, 태양 전지판의 지향점 해석 그리고 식기간의 분석을 위한 지구-달-탐사선-태양 간의 기하학적 위치에 대한 해석도 함께 이루어졌다. Preparing for future Korean Lunar missions, preliminary Lunar mission design software is developed using a impulsive thrusting method. Developed software is capable of design and analysis every required mission phases to design Lunar mission, including the Earth departure, Lunar transfer, Lunar arrival and mission operation phases. Also, assuming that KSLV-II is selected as a launch vehicle, future Korean Lunar explorer's mass budget is estimated based on driven optimal trajectory characteristics. Tracking analysis is also performed using Deep Space Network including angle geometry analysis between Earth - Moon - Lunar explorer - Sun which are very important for communication, solar panel pointing strategy and eclipse analysis when Lunar missions are under designing phase.

Study on the Characteristics of Two-burn Translunar Trajectory

In this paper, the translunar trajectory of the simple sequence is investigated. The trajectory discussed in this paper is a two-burn ballistic trajectory from the low earth parking orbit to the low lunar orbit. The problem is practical in that the geometric relation from the launch site to the moon is fully considered. The paper includes three subjects. First is the structure of the problem and the solution space. The problems are defined and the solutions are grouped by the properties of the lunar transfer sequence. Second are the characteristics of optimal solutions. The topics discussed are the transition of the required velocity increment based on the launch date, and the difference according to the property of the lunar transfer sequence. Third is the sensitivity analysis of a number of items to the deviation of the parameters from the optimal solution.