1,721,020 research outputs found
Earth escape from a sun-earth halo orbit using unstable manifold and lunar swingbys
No full textThis paper investigates the Earth escape for spacecraft in a Sun-Earth halo orbit. The escape trajectory consists of first ejecting to the unstable manifold associated with the halo orbit, then coasting along the manifold until encountering the Moon, and finally performing lunar-gravity-assisted escape. The first intersection of the manifold tube and Moon's orbit results in four intersection points. These four manifold-guided encounters have different relative velocities (v∞) to the Moon; therefore, the corresponding lunar swingbys can result in different levels of characteristic energy (C3) with respect to the Earth. To further exploit these manifold-guided lunar encounters, subsequent swingbys utilizing solar perturbation are considered. A graphical method is introduced to reveal the theoretical upper limits of the C3 achieved by double and multiple swingbys. The numerically solved Sun-perturbed Moon-to-Moon transfers indicate that a second lunar swingby can efficiently increase C3. Compared to the direct low-energy escape along the manifold, applying a portion of the lunar swingbys before escape is shown to be more advantageous for deep-space mission design.</p
Phasing Delta-V for transfers from Sun–Earth halo orbits to the Moon
No full textInspired by successful extended missions such as the ISEE-3, an investigation for the extended mission that involves a lunar encounter following a Sun-Earth halo orbit mission is considered valuable. Most previous studies present the orbit-to-orbit transfers where the lunar phase is not considered. Intended for extended missions, the present work aims to solve for the minimum phasing ∆V for various initial lunar phases. Due to the solution multiplicity of the two-point boundary value problem, the general constrained optimization algorithm that does not identify multiple feasible solutions is shown to miss minima. A two-step differential corrector with a two-body Lambert solver is developed for identifying multiple solutions. The minimum ∆V associated with the short-way and long-way approaches can be recovered. It is acquired that the required ∆V to cover all initial lunar phases is around 45 m/s for the halo orbit with out-of-plane amplitude Az greater than 3.5×105 km, and 14 m/s for a small halo orbit with Az=1×105 km. In addition, the paper discusses the phasing planning based on the ∆V result and the shift of lunar phase with halo orbit revolution.</p
Storm-time atmospheric density modeling using neural networks and its application in orbit propagation
No full textUpper atmospheric densities during geomagnetic storms are usually poorly estimated due to a lack of clear understanding of coupling mechanisms between the thermosphere and magnetosphere. Consequently, the orbit determination and propagation for low-Earth-orbit objects during geomagnetic storms have large uncertainties. Artificial neural networks are often used to identify nonlinear systems in the absence of rigorous theory. In the present study, an attempt has been made to model the storm-time atmospheric density using neural networks. Considering the debate over the representative of geomagnetic storm effect, i.e. the geomagnetic indices ap and Dst, three neural network models (NNM) are developed with ap, Dst and a combination of ap and Dst respectively. The density data used for training the NNMs are derived from the measurements of the satellites CHAMP and GRACE. The NNMs are evaluated by looking at: (a) the mean residuals and the standard deviations with respect to the density data that are not used in training process, and (b) the accuracy of reconstructing the orbits of selected objects during storms employing each model. This empirical modeling technique and the comparisons with the models NRLMSIS-00 and Jacchia-Bowman 2008 reveal (1) the capability of neural networks to model the relationship between solar and geomagnetic activities, and density variations; and (2) the merits and demerits of ap and Dst when it comes to characterizing density variations during storms.</p
Dynamics of spacecraft attitude considering high-fidelity solar radiation pressure
No full textSolar radiation pressure (SRP) is a major disturbance for a spacecraft operated in deep space missions, affecting its orbit and attitude, and high accuracy of the SRP calculation is required. The accuracy of SRP calculation is affected by the shape and surface characteristics of the satellite. However, existing computational models do not sufficiently take into account the exact reflective properties of the satellite surface. A detailed reflection model and considering the optical characteristics yield a different SRP and its torque from the conventional ones, resulting in a different attitude motion. In this study, an SRP calculation method that considers the exact reflective properties is applied to investigate the attitude dynamics.</p
How are multiple satellites seen from the ground? Relative apparent motion and formation stabilization
No full textThis paper answers how multiple satellites are seen from the ground. This question is inspired by space-advertising, a public exhibition in the night sky using a dot matrix of satellites that are bright enough to be seen by the naked eye. Thus, it is important for space advertisement that the specific dot matrix is seen. Moreover, the stability of the dot matrix during a visible span is very valuable. To stabilize the dot matrix, this study formulates an apparent position of a dot from a representative dot seen from the ground. The formulation, linear functions of a set of relative orbital elements, reveals the appearance of the dot matrix. The proposed relative variable in the formulation drives the instability of the dot matrix, thereby revealing an initial stable configuration of deputies from a chief. The arbitrary dot matrix designed using the configuration is stable even at low elevations without orbital control during the visible span. (Figure presented.)</p
Attitude motion under full orbit perturbations
No full textThis paper introduces an effort to precisely describe attitude motion under full orbit perturbations. To define the attitude of an Earth-orbiting spacecraft, this study introduces a reference frame being affected by orbit perturbations. Unlike the commonly adopted reference frame, the reference frame introduced here is fully perturbed, so that it rotates about not only the pitch axis, but also the yaw axis. To incorporate the mutual coupling effect between attitude motion and orbit motion, the method introduced considers the spacecraft as consisting of multiple facets and carefully models orbit perturbation-induced torque that varies the attitude. This paper focuses on the attitude dynamics of a small satellite with relatively low moments of inertia under full orbit perturbations, and provides some interesting results and the outcome from the method introduced
Close approach analysis method considering attitude and shape using dual quaternions
No full textThis paper proposes a close approach analysis method considering object attitude change. Implementing a detailed close approach analysis is important for space situational awareness. Conventional conjunction assessment methods assume the shape of the object as a encompassing sphere, which excludes considering the attitude change. However, most of the spacecraft are not spherical, and conventional methods result in overestimation. The proposed method in this paper consists of two steps: 1) calculation of the time of close approach with two objects as point masses and 2) interpolation of the first and last states in the time duration when the close distance is less than a given threshold. Dual quaternions are employed to express the translational and rotational motions, which ensures the uniqueness of the interpolation. Numerical simulations are performed to verify the proposed method and compare it with a conventional method.</p
Estimation of orbital parameters of broken-up object using in-situ debris measurement satellite
No full textCollisions and explosions of satellites generate a large amount of sub-millimeter-size debris, which can cause fatal damage to a spacecraft. However, such tiny debris cannot be tracked from the ground. Therefore, an in-situ debris measurement satellite, which can detect impacts with sub-millimeter-size debris, has been proposed. Based on this concept, previous studies proposed the method to estimate some orbital parameters of the broken-up object. In those studies, it is assumed that the measurement satellite detects impacts at the intersection of the orbital planes of the measurement satellite and the broken-up object. However, simulated measurement data includes fragments contrary to this assumption. Thus, this study introduces a new approach to estimating the orbital parameters of the broken-up object from simulated measurement data. In this study, the inclination, the right ascension of the ascending node, and the nodal precession rate are estimated from the history of geocentric declination using the iteratively reweighted nonlinear least square method.</p
Origin identification of sub-millimeter-sized debris from in-situ debris measurements
No full textThis study introduces a new method to estimate the orbital parameters (inclination, nodal precession rate, and right ascension of the ascending node) of the broken-up object using in-situ debris measurements. To reduce collision probability with tiny debris generated from breakups of satellites, it is required to identify the origin of breakups and predict the orbit of debris. In general, small debris cannot be tracked from the ground, thus an in-situ measurement is an effective method to detect sub-millimeter-sized debris. Previous studies developed a method to identify the location of breakup origin from an in-situ measurement, however, it can be applied to an ideal situation where the detection occurs at the intersection of the orbit planes. This study introduces a new method to estimate the orbital parameters by fitting the model of geocentric declination to the measurement data. This method can be applied to practical situations that cannot be validated in previous studies. This paper also adopts iteratively reweighted least squares to improve estimation accuracy and mitigate initial value dependencies.</p
Stabilization of space-advertisement satellite formation
No full textSpace-advertisement is getting closer with advances in satellite formation flying, which displays an artificial constellation in the night sky. The formation shape composed of pixels of satellites equipping light sources is observed from the ground and different from formation in relative orbits. For realizing the advertisement shape of high quality, there are two requirements, placing the pixels of the advertisement shape properly and keeping the shape during a visible span. It is not enough to satisfy the requirements by the usual coordinate systems and equations. Considering a topocentric coordinate system, the new transforming equations are derived so that an arbitrary advertisement shape is designed at the target time. Also, stable orbits that keep the formation shape without control are found and an example of stable orbits is presented.</p
- …
