Earth Escape Trajectory Research Articles

Hayabusa 2 extension plan: Asteroid selection and trajectory design

The Hayabusa 2 mission is targeted to explore the asteroid (162173) 1999 JU3 and return surface as well as sub-surface samples through a novel impactor. Upon its return, at the end of 2020, the spacecraft will release the capsule for Earth re-entry and drift away from the planet. Based on the current mission profile, the spacecraft is expected to retain 30 kg of xenon propellant for trajectory maneuvers after the capsule is released. This remaining fuel can be used to extend the mission and improve its scientific return by exploring a new target. Work herein outlines an extension plan for Hayabusa 2, detailing the target selection process and its subsequent trajectory design. Due to final Earth escape trajectory, considering the excess velocity and orbital geometry, the only available extension option is an asteroid flyby. One of the most important trajectory characteristic is to maximize the spacecraft's optical detection capabilities. As a result the asteroid 2001 WR1 is identified as the most promising target candidate. The resulting trajectory uses all the available xenon with 100% duty cycle. Furthermore, the extension lasts for 932 days and offers 1.57 days of optical navigation time for a flyby on June 27, 2023.

An analytical description of three-dimensional heliocentric solar sail orbits

This paper introduces a simple analytical approximation to three-dimensional heliocentric solar sail orbits where the only forces considered are solar gravity and solar radiation. The approximation is based upon the previously studied hodograph transformation and provides a description of the inclination, longitude of ascending node and true latitude for a specific set of initial conditions. It is shown that the rotational symmetry of a heliocentric orbit allows this specific solution to be mapped onto a solution with arbitrary initial conditions. The approximation is then compared to the numerical results for a solar sail on an Earth escape trajectory with an area to mass ratio up to twice as high as current technology allows.

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.

저추력기를 이용한 연료 최적의 지구탈출 궤적 설계 연구

A Discrete continuation Method/homotopy approaches are studied for energy/fuel optimal low thrust Earth escape trajectory by solving a two point boundary value problem(TPBVP). Recently, maneuvers using low thrust propulsion system have been identified as emerging technologies. The low thruster is considered as the main actuator for orbit maneuvers. The cost function consists of a energy/fuel consumption function, and constraints are position and velocity vectors at the terminal escape point. Solving the minimum energy/fuel problem directly is not an easy task, so we adopt the homotopy analysis. Using a solution of the minimum energy, which is solved by discrete continuation method, we obtain the solution of the minimum fuel problem.က€

Inverse Solar Sail Trajectory Problem

SOLAR sailing has long been considered for a diverse range of future mission applications. Although low-performance solar sails can be utilized for interplanetary transfer using heliocentric spiral trajectories, high-performance solar sails can enable exotic applications using non-Keplerian orbits. A simple example of such an exotic applicationis levitation,with the solar radiationpressure accelerationexperiencedby the sail exactly balancing solar gravity. Such a static equilibrium allows the solar sail to remain stationary with respect to the sun, or indeed if the sail is turned edgewise to the sun it will fall sunwards on a rectilinear trajectory. Although this static equilibrium is simple to identify, the question of transfer to it from an Earth escape trajectory remains open. This Note will derive an analytic sail steering law that allows the solar sail is be maneuvered from a circular heliocentric orbit, to a static equilibrium location at the same heliocentric distance. The required trajectory will be deŽ ned a priori with the resulting sail steering law derived from the equations of motion. An inverse trajectory problem is, therefore, being solved.

LISA — a study of the ESA cornerstone mission for observing gravitational waves

The primary objective of the Laser Interferometer Space Antenna (LISA) mission is to detect and observe gravitational waves from massive black holes and galactic binaries in the frequency range 10 −4 to 10 −1 Hz. This low-frequency range is inaccessible to ground-based interferometers because of the unshieldable background of local gravitational noise and because ground-based interferometers are limited in length to a few km. LISA is an ESA cornerstone mission and recently had a system study (Ref. 1) carried out by a consortium led by Astrium, which confirmed the basic configuration for the payload with only minor changes, and provided detailed concepts for the spacecraft and mission design. The study confirmed the need for a drag-free technology demonstration mission to develop the inertial sensors for LISA, before embarking on the build of the flight sensors. With a technology demonstration flight in 2005, it would be possible to carry out LISA as a joint ESA-NASA mission with a launch by 2010 subject to the funding programmatics. The baseline for LISA is three disc-like spacecraft each of which consist of a science module which carries the laser interferometer payload (two in each science module) and a propulsion module containing an ion drive and the hydrazine thrusters of the AOCS. The propulsion module is used for the transfer from earth escape trajectory provided by the Delta II launch to the operational orbit. Once there the propulsion module is jettisoned to reduce disturbances on the payload. Detailed analysis of thermal and gravitational disturbances, a model of the drag-free control and of the interferometer operation confirm that the strain sensitivity of the interferometer will be achieved.

Mercury sun-synchronous polar orbiter with a solar sail

An innovative mission MESSAGE (MErcury Solar Sailing Advanced Geoscience Exploration) is proposed where a solar sail spacecraft of about 240kg total mass delivers a limited but highly valuable science payload to a sun-synchronous Mercury orbit in 312 years. Corresponding operation near the terminator offers favorable thermal and remote sensing conditions for the considered instruments. Since Mercury is still the poorest explored and least known planet of the inner solar system, there is a high level of scientific interest for further exploration. Science objectives and a potential payload package of about 20 kg for the proposed solar sail mission considering the special orientation and geometry of a sun-synchronous orbit about Mercury will be discussed. Determination of the elemental composition of the Hermian surface with respect to major rock forming elements and volatiles could be carried out by a gamma-ray spectrometer. In the current mission architecture the spacecraft will be launched into an Earth escape trajectory with a relatively low-cost launch option such as TAURUS or ROCKOT with upper stages. The sail is used as a transportation means for the interplanetary transfer, for orbit capture upon Mercury arrival, and subsequent adjustment of a sun-synchronous orbit about the planet near the terminator. Due to Mercury's 3:2 spin-orbit coupling the eccentric polar orbit of 200 km × 6350 km altitude with periapsis above the north pole will allow complete coverage of the planet's surface within two Mercury years. Optimized interplanetary trajectories as well as simulations of the orbit about Mercury will be shown. A conceptual design of the spacecraft including the deployment concept of the square sail of an estimated size of about 86m × 86m and its support structure will be presented. The lightweight booms are proposed to be manufactured of carbon fibre reinforced plastic profiles that can be stored on a central hub prior to deployment. Possible sail film materials, coatings and folding concepts will be discussed. In a more advanced scenario two identical solar sails could be launched in a stacked configuration by a Med-Lite to realize a dual sun-synchronous Mercury orbiter mission.