Consequence of Continued Growth in the GEO and GEO Disposal Orbital Regimes

To date more than 800 spacecraft, upper stages, and apogee kick motors are known to reside in geosynchronous and nearby orbits, including geosynchronous disposal orbits. An even larger number of debris greater than 10 em in diameter have been detected by U.S. and European groundbased sensors. Using projections of geosynchronous deployment characteristics and disposal rates, NASA and Kyushu University models of the geosynchronous and super-geosynchronous orbital regimes have examined the sensitivity of the long-term satellite population to various scenarios. Emphasis has been placed on the rate of collisions in the geosynchronous orbit and in the higher disposal orbits and on the significance of cross-regime contamination. The sensitivity of the long-term environment on low velocity (0-1 km/s) collision breakup model parameters and on the minimum height of disposal orbits has also been explored. Results are presented in terms of both satellite population and spatial density

Using NASA Standard Breakup Model to Describe Low-Velocity Impacts on Spacecraft

The applicability is examined of the hypervelocity collision model included in the NASA standard breakup model 2000 revision to low-velocity collisions possible in space, especially in the geosynchronous regime. The analytic method used in the standard breakup model is applied to experimental data from low-velocity impact experiments previously performed at Kyushu University at a velocity range less than 300 m/s. The projectiles and target specimens used were stainless steel balls and aluminum honeycomb sandwich panels with face sheets of carbon fiber reinforced plastic, respectively. It is concluded that the hypervelocity collision model in the standard breakup model can be applied to low-velocity collisions with some simple modifications

Earth escape from a sun-earth halo orbit using unstable manifold and lunar swingbys

This 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

Storm-time atmospheric density modeling using neural networks and its application in orbit propagation

Upper 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

The phasing problem for sun-earth halo orbit to lunar encounter transfers

Halo orbit missions are of many applications and become popular. An investigation on the extended mission following halo orbit missions would be worthwhile. In a previous study, the strategy of using the unstable manifolds associated with the Sun-Earth L-1/L-2 halo orbit and lunar gravity assists for Earth escape was analyzed to be advantageous for extending the mission. However, in an extension mission where the halo orbit mission is not pre-phased for a lunar swingby, the fuel cost for phasing the halo-to-Moon transfers should be investigated. The current paper aims to give the insight of the minimum phasing AV to encounter the Moon for various lunar phases with respect to the halo orbit. Efforts are made to tackle the problem of multiple optimization directions. The phasing planning is briefly discussed as well

Phasing Delta-V for transfers from Sun–Earth halo orbits to the Moon

Inspired 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

How are multiple satellites seen from the ground? Relative apparent motion and formation stabilization

This 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