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Earth observation satellite (EOS) scheduling is critical for improving the performance of Earth observation system. Most of the existing studies on satellite observation scheduling problems focus on static tasks and real-time dynamic tasks also known as emergency tasks while few attempts have been made for common dynamic tasks yet. The common dynamic tasks are characterized by uncertain arrival times and delayable observations while the emergency tasks require immediate observations. The number of common dynamic tasks is rapidly increasing with the expansion of satellite observation applications, which poses great challenges to EOS observation scheduling. To address this issue, this paper investigates the satellite observation scheduling problem with common dynamic tasks. Firstly, a centralized onboard dynamic scheduling framework based on a periodic-triggered rolling horizontal optimization (RHO) strategy is proposed and a novel two-stage mathematical model is established for the periodic rescheduling problem. Secondly, we propose a low-complexity onboard periodic rescheduling algorithm (OPR), which consists of a greedy-based task allocation algorithm, a pointer network (Ptr-network) based task scheduling algorithm and an iterative local search algorithm. In the greedy-based task allocation algorithm, we define a Task-EOS Fitness Indicator (TEFI) and each task is greedily allocated to the EOS with maximum TEFI. In the Ptr-network based task scheduling algorithm, the allocated task sets of all EOSs are fed into the Ptr-network in parallel to generate the observation scheduling result. Afterward, an iterative local search algorithm is proposed to further improve the quality of the observation scheduling result. Finally, extensive experiments are conducted to demonstrate the superiority of the OPR algorithm for the satellite observation scheduling problem with common dynamic tasks. [ABSTRACT FROM AUTHOR]

The paper considers a process of a controlled attitude motion of a dual-spin nanosatellite. The main goal of control is to achieve a gravitationally stabilized position in a circular orbit in the shortest possible time. The dual-spin nanosatellite contains a gravitational damper and an internal rotor inside a movable unit. The movable unit can fulfill small controlled inclination relative to the main body of the nanosatellite. The mathematical model of the attitude dynamics of the nanosatellite is built and appropriate control laws are developed. Simulation results confirm the performance of the gravitational damper and efficiency of the control algorithm. [ABSTRACT FROM AUTHOR]

• The mathematical model of the reachable domain of the contingency return trajectories. • Reachable domain calculation method of the contingency return trajectories based on interval analysis. • Parameter characteristic analysis of reachable domain. The investigation on reachable domain based on interval analysis is performed, which is aimed for the contingency return trajectories during the circumlunar flight stage in manned lunar landing engineering. First, the issue of contingency return trajectory, which adopts a non-coplanar maneuver during the circumlunar flight phase, is expounded upon. The mathematical model of the reachable domain of contingency return trajectories is established. Second, the contingency return trajectory is designed. An improved interval branch and bound algorithm is introduced and a reachable domain calculation method of the contingency return trajectories is proposed. Third, in numerical simulations, by comparing the reachable domain obtained by the traditional brute force method, the accuracy of the method presented in this paper is verified. It also indicates that this method has a higher calculation efficiency. Finally, numerous simulations are conducted to discuss the parameter characteristic of reachable domain by adopting this method. The study conclusions can serve as an important reference for the selection and decision of contingency return trajectory plans in the future manned lunar landing engineering. [ABSTRACT FROM AUTHOR]

In recent times, origami-inspired deployable structures have received significant interest from the research community as a major area of research in the fields of science and technology. This paper deals with the geometry and the deformation behavior of deployable cones based on the four-fold line origami concept. First, a novel geometric design algorithm is presented to develop axially foldable conical structures based on generalized Miura-ori pattern, which is then utilized to investigate the influence of four design parameters on folding geometry. The foldability of thus obtained origami patterns is verified with the help of physical models. The geometric incompatibilities during the deployment are highlighted using a simple experiment, showing the unfolding sequence of a paper model. The deformations due to incompatibilities are analyzed by adopting the concept of pin-jointed truss modeling of the origami structures. A mathematical model of the deployment problem is formulated to capture the deployment deformations of the structure, on the basis of which an algorithm for deformation analyses is developed. One example is considered to characterize the deformation behavior of the conical origami structures as the design parameters of the folding pattern are varied. The analytical findings demonstrate that the fold line strains of the origami cone are very sensitive to the changes in three design parameters, number of origami units (m), fold angle (α 0), and angle θ 2 ; except the length ratio (l r 0). The variation in design parameters may induce bi-stability in the deployment path of the conical structure. The findings of the analytical investigation are verified numerically, demonstrating that the strain responses obtained from the numerical analyses are completely consistent with the analytical results. [ABSTRACT FROM AUTHOR]

This paper proposes a procedure for managing the risk of reentering space objects and risk assessment methodologies used for the process. The proposed procedure comprises three phases encompassing the whole reentry stages of space objects. Mathematical models for assessing the impact risk of the reentering space objects by utilizing the information available during different risk management phases and the recommended risk analysis results for public communication are presented. The concept of the conditional casualty expectation is proposed as the metric representing the reentry risk and the method to compute its profile is introduced. A case study on the risk management procedure with the dataset on an actual reentry event is conducted to demonstrate the efficacy of the proposed procedure. [ABSTRACT FROM AUTHOR]

In this paper, a fundamental approach to analyzing the maneuverability of spacecraft with mounted control moment gyros is addressed by searching feasible angular momentum. The geometrical array of control moment gyros considered is a roof array also known as a 2-SPEED system . Due to the simplicity of the singularity envelope of the array considered, it is reasonably practical to conduct rotational performance analysis of the spacecraft. This maneuverability analysis technique uses a unique chart developed in this work, which allows a guaranteed maximum torque output and angular momentum at any time without concern about the singularity problem of the control moment gyros. Therefore, the purpose of this paper is to provide a conservative method for maneuverability analysis of a spacecraft with installed control moment gyros by searching allowable maximum angular momentum. This method is demonstrated using an illustrative example. [ABSTRACT FROM AUTHOR]

Due to high relative velocities, collisions of spacecraft in orbit with Space Debris (SD) or Micrometeoroids (MM) can lead to payload degradation, anomalies as well as failures in spacecraft operation, or even loss of mission. Flux models and impact risk assessment tools, such as MASTER (Meteoroid and Space Debris Terrestrial Environment Reference) or ORDEM (Orbital Debris Engineering Model), and ESABASE2 or BUMPER II are used to analyse mission risk associated with these hazards. Validation of flux models is based on measured data. Currently, as most of the SD and MM objects are too small (millimeter down to micron sized) for ground-based observations (e.g. radar, optical), the only available data for model validation is based upon retrieved hardware investigations e.g. Long Duration Exposure Facility (LDEF), Hubble Space Telescope (HST), European Retrievable Carrier (EURECA). Since existing data sets are insufficient, further in-situ experimental investigation of the SD and MM populations are required. This paper provides an overview and assessment of existing and planned SD and MM impact detectors. The detection area of the described detectors is too small to adequately provide the missing data sets. Therefore an innovative detection concept is proposed that utilises existing spacecraft components for detection purposes. In general, solar panels of a spacecraft provide a large area that can be utilised for in-situ impact detection. By using this method on several spacecraft in different orbits the detection area can be increased significantly and allow the detection of SD and MM objects with diameters as low as 100 μm. The design of the detector is based on damage equations from HST and EURECA solar panels. An extensive investigation of those panels was performed by ESA and is summarized within this paper. Furthermore, an estimate of the expected sensitivity of the patented detector concept as well as examples for its implementation into large and small spacecraft are presented. [ABSTRACT FROM AUTHOR]

WAM is realized making use of the plasma density data collected via the retarding potential analyser on board the Dynamics Explorer 2 spacecraft, capable to model the scintillation climatology over the northern hemisphere high latitude ionosphere. More recently, WAM has been tuned to model the ionospheric scintillations also over the equatorial latitudes. The effort has been done to support the CIGALA (Concept for Ionospheric Scintillation Mitigation for Professional GNSS in Latin America) project in the assessment of the scintillations climatology over Latin America. The concept of the new release of WAM is the same already adopted for the high latitudes: the in situ measurements, supplemented with an ionospheric model and with the irregularity anisotropy model, are treated to describe the morphology of scintillation, provided a suitable propagation model is used. Significant differences have been included in the low latitudes release to account for the anisotropy of the irregularities and for strong scattering regime. The paper describes the new WAM formulation and presents comparisons of the model predictions with the actual measurements collected in Brazil. [ABSTRACT FROM AUTHOR]

In a geodetic radio frequency observing system the phase center offsets and phase center variations of ground antennae are a fundamental component of mathematical models of the system observables. In this paper we describe work aimed at improving the DORIS Starec ground antenna phase center definition model. Seven antennas were analyzed in the Compact Antenna Test Range (CATR), a dedicated CNES facility. With respect to the manufacturer specified phase center offset, the measured antennae varied between −6 mm and +4 mm due to manufacturing variations. To solve this problem, discussions were held with the manufacturer, leading to an improvement of the manufacturing process. This work results in a reduction in the scatter to ±1 mm. The phase center position has been kept unchanged and associated phase law has been updated and provided to users of the International DORIS Service (IDS). This phase law is applicable to all Starec antennas (before and after manufacturing process consolidation) and is azimuth independent. An error budget taking into account these updated characteristics has been established for the antenna alone: ±2 mm on the horizontal plane and ±3 mm on the up component, maximum error values for antennas named type C (Saunier et al., 2016) produced with consolidated manufacturing process. Finally the impact of this updated characterization on positioning results has been analyzed and shows a scale offset only of the order of +12 mm for the Terrestrial Reference Frame. [ABSTRACT FROM AUTHOR]

With the impact of the Global Navigation Satellite System (GNSS), dual-GNSS constellations are playing an increasingly significant role in positioning, navigation and timing (PNT) applications. Aiming at improving from the existing method, i.e., linearization, of solving the single-point positioning problem under a dual-GNSS, this paper develops a closed-form method for solving PNT problems in the case of six satellites. This method reduces the positioning problem to a simple mathematical problem of finding solutions to a quadratic equation, thereby needing only one receiver clock bias (RCB) as variable. By solving the RCB, the positioning information in three dimensions is obtained by utilizing a linear equation. Compared with the existing method, the closed-form method requires no initial position or iterations. This method thus provides a direct solution to single-point positioning. Further, how to check the uniqueness and the validity of the solutions is also derived. Experimental results verify the validity, applicability and efficiency of the proposed method. [ABSTRACT FROM AUTHOR]

Using total electron content (TEC) data over one and a half solar cycles (1999–2015) provided by the Center for Orbit Determination in Europe (CODE), this paper proposes a new empirical TEC model for northeast China (40–50N, 120–130E). The model, called TECM-NEC, involves the multiplication of four separable components, including diurnal variation, seasonal variation, geomagnetic field dependency, and solar dependency. Diurnal variation is composed of three parts: the typical daily variation of TEC; corrections of Mid-latitude Summer Nighttime Anomaly (MSNA) that depend on geographic location, season, and local time; and corrections of day-to-night ratio under different seasons and solar activities. Four sub-harmonics of the year with annual, semiannual, four-, and three-month periods are used to describe seasonal variations. For geomagnetic variation, geomagnetic latitude is based on the latest International Geomagnetic Reference Field (IGRF12) model. Compared with similar empirical models, the solar proxy index F10.7P = (F10.7 + F10.7A)/2, where F10.7A is the 81-day running mean of daily F10.7, is chosen as having linear relationship with TEC for the model. This model has 43 coefficients, which are determined by nonlinear least squares fitting (NLSF) technique. The TECM-NEC model fits with the TEC/CODE input data with a bias of 0.03TECU and a RMS deviation of 2.76TECU. The proposed TECM-NEC model can reproduce the MSNA and nighttime TEC enhancements phenomenon over northeast China. [ABSTRACT FROM AUTHOR]

The time-consuming experimental method for handling qualities assessment cannot meet the increasing fast design requirements for the manned space flight. As a tool for the aircraft handling qualities research, the model-predictive-control structured inverse simulation (MPC-IS) has potential applications in the aerospace field to guide the astronauts’ operations and evaluate the handling qualities more effectively. Therefore, this paper establishes MPC-IS for the manual-controlled rendezvous and docking (RVD) and proposes a novel artificial neural network inverse simulation system (ANN-IS) to further decrease the computational cost. The novel system was obtained by replacing the inverse model of MPC-IS with the artificial neural network. The optimal neural network was trained by the genetic Levenberg–Marquardt algorithm, and finally determined by the Levenberg–Marquardt algorithm. In order to validate MPC-IS and ANN-IS, the manual-controlled RVD experiments on the simulator were carried out. The comparisons between simulation results and experimental data demonstrated the validity of two systems and the high computational efficiency of ANN-IS. [ABSTRACT FROM AUTHOR]

The transient dynamic behaviour of supersonic disk-gap-band parachutes in a Mars entry environment involving fluid structure interactions is studied. Based on the multi-material Arbitrary Lagrange–Euler method, the coupling dynamic model between a viscous compressible fluid and a flexible large deformation structure of the parachute is solved. The inflation performance of a parachute with a fixed forebody under different flow conditions is analysed. The decelerating parameters of the parachute, including drag area, opening loads, and coefficients, are obtained from the supersonic wind tunnel test data from NASA. Meanwhile, the evolution of the three-dimensional shape of the disk-gap-band parachute during supersonic inflation is presented, and the structural dynamic behaviour of the parachute is predicted. Then, the influence of the presence of the capsule on the flow field of the parachute is investigated, and the wake of unsteady fluid and the distribution of shock wave around the supersonic parachute are presented. Finally, the structural dynamic response of the canopy fabric under high-pressure conditions is comparatively analysed. The results show that the disk-gap-band parachute is well inflated without serious collapse. As the Mach numbers increase from 2.0 to 2.5, the drag coefficients gradually decrease, along with a small decrease in inflation time, which corresponds with test results, and proves the validity of the method proposed in this paper. [ABSTRACT FROM AUTHOR]

The thermal subsystem of spacecrafts and payloads is always designed with the help of Thermal Mathematical Models. In the case of the Thermal Lumped Parameter (TLP) method, the non-linear system of equations that is created is solved to calculate the temperature distribution and the heat power that goes between nodes. The accuracy of the results depends largely on the appropriate calculation of the conductive and radiative conductances. Several established methods for the determination of conductive conductances exist but they present some limitations for complex geometries. Two new methods are proposed in this paper to calculate accurately these conductive conductances: The Extended Far Field method and the Mid-Section method. Both are based on a finite element calculation but while the Extended Far Field method uses the calculation of node mean temperatures, the Mid-Section method is based on assuming specific temperature values. They are compared with traditionally used methods showing the advantages of these two new methods. [ABSTRACT FROM AUTHOR]

This paper uses surrogate models to reduce the computational cost associated with spacecraft mission design in three-body dynamical systems. Sampling-based least squares regression is used to project the system response onto a set of orthogonal bases, providing a representation of the Δ V required for rendezvous as a reduced-order surrogate model. Models are presented for mid-field rendezvous of spacecraft in orbits in the Earth–Moon circular restricted three-body problem, including a halo orbit about the Earth–Moon L 2 libration point (EML-2) and a distant retrograde orbit (DRO) about the Moon. In each case, the initial position of the spacecraft, the time of flight, and the separation between the chaser and the target vehicles are all considered as design inputs. The results show that sample sizes on the order of 10 2 are sufficient to produce accurate surrogates, with RMS errors reaching 0.2 m/s for the halo orbit and falling below 0.01 m/s for the DRO. A single function call to the resulting surrogate is up to two orders of magnitude faster than computing the same solution using full fidelity propagators. The expansion coefficients solved for in the surrogates are then used to conduct a global sensitivity analysis of the Δ V on each of the input parameters, which identifies the separation between the spacecraft as the primary contributor to the Δ V cost. Finally, the models are demonstrated to be useful for cheap evaluation of the cost function in constrained optimization problems seeking to minimize the Δ V required for rendezvous. These surrogate models show significant advantages for mission design in three-body systems, in terms of both computational cost and capabilities, over traditional Monte Carlo methods. [ABSTRACT FROM AUTHOR]

The base of the dynamics characteristic research on the parachute and vehicle system is to establish a dynamics model, during the parachute descent phase, which can accurately display the relationship among the velocity, altitude and attitude angles as well as the variation of time. This paper starts with a new tracking law – ADRC in Mars entry guidance, which affects the initial states of the parachute deployment point and determines precision landing capability. Then, the influence of unsteady resistance to the parachute in Martian air is considered as the added mass, and a 6DOF nonlinear mathematical model of the parachute and vehicle system is established. [ABSTRACT FROM AUTHOR]

In this paper, the Global Ionospheric Maps (GIMs) during 1998–2012 have been adopted as the primary source of data to generate the Galileo-like ionospheric correction parameters. Taking the previous day GIMs as ingested dataset and the current day GIMs as reference, we have computed the cumulative distribution function (CDF) of TEC (Total Electron Content) error less than 5 TECU (TEC Unit, 1TECU = 10 16 electron/m 2 ) or 20% (denoted as zi 20 ) to study the performance of NeQuick 2 in providing global ionospheric correction. The statistical analysis is carried out to study how the performance changes with seasons, solar activities and geographic latitudes. Our studies showed that: (1) in general, the NeQuick 2 model driven by the previous day GIMs agrees well with the current day GIMs, with monthly averaged zi 20 exceeds 60%. (2) The performance of GIMs-driven-NeQuick model varies with geographic latitudes. The model behaves better in mid-latitudes than in low-latitudes, which is more prominent in northern hemisphere. (3) There is a nonlinear relationship between zi 20 and F10.7. When the solar activity level is low or medium, zi 20 tends to decrease with increasing solar activities. However, when F10.7 increases further to exceed a certain threshold, zi 20 generally shows either saturation or an increase for high solar epoch. The zi 20 decreases with F10.7 more sharply in mid-latitudes than in low-latitudes, in the southern hemisphere than in the northern hemisphere. (4) The seasonal behavior of zi 20 in different latitudinal belts are quite different. In the southern hemisphere, zi 20 mainly presents annual variation, and it is largest in winter and smallest in summer. In the northern hemisphere, the semi-annual variation is dominant, and zi 20 is larger in solstice than in equinox. During the High Solar Activity years, the zi 20 is much larger in summer than in winter in the low-latitude regions. [ABSTRACT FROM AUTHOR]