Spacecraft Parameters Research Articles

Evaluation of the Efficiency of Propulsion Systems Usage with Energy Accumulation at Spacecraft

The problem of cooperative optimization of trajectory and main design parameters of spacecraft, equipped with propulsion systems with energy accumulation, during low Earth orbit formation and correction in non-central gravitational field is considered. The working principle of these propulsion systems is a periodic cycle of energy accumulation from the energy source at the passive stages of orbital transfer, and its discharge to supply power to the engine during the active stages of orbital transfer. The special characteristics of propulsion systems with energy accumulation is limited operation time per one switching, which determines by the energy, stored in an accumulator. For design-ballistic analysis, the mass model of the spacecraft with an uncontrollable power-limited engine with energy accumulator is used. The general problem is separated into dynamic and parametric parts. Using averaging mathematical models of motion and Pontriagin's maximum principle, the calculation method of the ballistic characteristics of approximately-optimal orbital transfer is obtained. The method of design-ballistic optimization is obtained as an iterative procedure. It was found, that inclusion of an energy accumulator in a propulsion system provides an increasing of orbital transfer efficiency.

Method of predictive studies of the effectiveness of spacecraft modifications with integrated subsystem replacement

The article reviews application features of the method of bilevel coordinated optimization of spacecraft modification parameters during integrated replacement of subsystems. The algorithm for problem solution is presented in the article. It is shown that in the case of implementation of a balanced bilevel schematic of optimization search, the agreement of project solutions for spacecraft modification is provided by integrated replacement of subsystems. Project models are shown for the problem solution of optimization parameters for spacecraft modification with the integrated replacement of subsystems of the payload equipment module and a unified space platform.

Optimization of project-ballistic parameters for low earth orbit spacecraft with propulsion systems with energy storage

Optimization of project-ballistic parameters of the spacecraft with propulsion systems with energy accumulation during the orbital transfer between low Earth orbits in the non-central gravitational field of the Earth is considered. The mass model of the spacecraft with an uncontrollable power-limited engine with the energy accumulator is used. The method of the project-ballistic optimization is obtained as an iterative procedure. Energy storage usage efficiency during the orbital transfer between low Earth orbits is estimated.

Time-varying modal parameters identification of large flexible spacecraft using a recursive algorithm

In existing identification methods for on-orbit spacecraft, such as eigensystem realization algorithm (ERA) and subspace method identification (SMI), singular value decomposition (SVD) is used frequently to estimate the modal parameters. However, these identification methods are often used to process the linear time-invariant system, and there is a lower computation efficiency using the SVD when the system order of spacecraft is high. In this study, to improve the computational efficiency in identifying time-varying modal parameters of large spacecraft, a faster recursive algorithm called fast approximated power iteration (FAPI) is employed. This approach avoids the SVD and can be provided as an alternative spacecraft identification method, and the latest modal parameters obtained can be applied for updating the controller parameters timely (e.g. the self-adaptive control problem). In numerical simulations, two large flexible spacecraft models, the Engineering Test Satellite-VIII (ETS-VIII) and Soil Moisture Active/Passive (SMAP) satellite, are established. The identification results show that this recursive algorithm can obtain the time-varying modal parameters, and the computation time is reduced significantly.

Efficiency of using electric propulsion engines for the task of keeping in a near-circular orbit

The paper deals with the problem of choosing the type and the main design parameters of the correcting propulsion system of spacecraft functioning in low near-circular Earth orbits. The efficiency of using chemical (impulse) and low-thrust electric propulsion engines is analyzed. Analytical relations for calculating spacecraft design parameters and specifying the areas of efficient use of electric propulsion engines are obtained for the criterion of useful load fraction and the assumption of constant power dependent acceleration. The results of calculating the areas of efficient use and design parameters of spacecraft for various operational altitudes are shown.

On-orbit identification of flexible parameters of spacecraft

This paper studies the on-orbit identification of flexible parameters of spacecraft using a typical spacecraft as research object. First, the coupled dynamic equation of the system is established with the Lagrange method and the model order of the solar array is reduced using the inertia complete criterion. Then, the particle swarm optimization is adopted to determine the optimal positions of acceleration sensors on the solar array. Finally, the observer/Kalman filter identification and eigensystem realization algorithm are applied to identify the flexible parameters of the system. In the simulations, three types of external excitations, including white noise, sine and impulse, are considered as system input, respectively. Simulation results indicate that the presented method in this paper is effective in identifying the flexible parameters of the spacecraft.

태양동기궤도 위성군 궤도 최적화에 관한 연구

This paper presents a sun-synchronous orbit design which effectuvely includes the requirements derived from spacecraft to ground station contact and spacecraft to target image accessibility. For this purpose, operation parameters of multiple spacecraft are defined as Contact Overlap, Contact Overlap Gap, Access Overlap, Access Overlap Gap. These parameters are used to form a Figure of Merit that reflects the operational requirements. The Figure of Merit is optimized to increase the efficiency of operating multiple spacecraft in constellation and is used to determine the operational orbit of each spacecraft that constitutes the constellation.

Применение экспериментального алгоритма расчета параметров движения КА после выведения его наорбиту

Применение экспериментального алгоритма расчета параметров движения КА после выведения его наорбиту

Electrodynamic control for spacecraft attitude stability in the geomagnetic field

The electrodynamic method of spacecraft attitude control is studied in an octupole geomagnetic field approximation. An analytical derivation is given for refined laws of control over spacecraft electrodynamic parameters, which ensure the required spacecraft guidance. Spacecraft equilibrium position is proved to be stable under continually acting disturbances. The effectiveness and practical feasibility of the proposed spacecraft attitude control method is confirmed by numerical analysis.

Trajectory correction of the Spektr-R spacecraft motion

The results of refining the parameters of the Spektr-R spacecraft (RadioAstron project) motion after it was launched into the orbit of the Earth’s artificial satellite in July 2011 showed that, at the beginning of 2013, the condition of staying in the Earth’s shadow was violated. The duration of shading of the spacecraft exceeds the acceptable value (about 2 h). At the end of 2013 to the beginning of 2014, the ballistic lifetime of the spacecraft completed. Therefore, the question arose of how to correct the trajectory of the motion of the Spektr-R satellite using its onboard propulsion system. In this paper, the ballistic parameters that define the operation of onboard propulsion system when implementing the correction, and the ballistic characteristics of the orbital spacecraft motion before and after correction are presented.

Dynamics of an Unstabilized Spacecraft During the Deployment of an Elastic Pantograph Structure

An unstabilized spacecraft with a pantograph structure deployed in orbit to carry solar batteries is the subject of study. The objective of the study is to construct a mathematical model of this system taking into account the elastic properties of elements of the pantograph in longitudinal and transverse directions. This model is based on the Lagrangian formalism as applied to a mechanical system with rheonomic constraints. The expressions for the coefficients of the equations of motions are obtained using Mathematica 5 © . A Fortran application software package is used for numerical simulation of dynamic processes. This package can be adapted, if necessary, to study other structures. The behavior of the spacecraft during the deployment of the pantograph is numerically analyzed using different values of spacecraft parameters and the parameters of the deployment process in the gravity field.

Problems in formation and control of a required microacceleration level at spacecraft design, tests, and operation

In this paper, we consider the main aspects of a problem in formation and control of a required microacceleration level as the most important characteristic of the space laboratory intended for realizing the gravitational and sensitive technological processes. The estimation model and methodology to control and estimate the microaccelerations are presented that are necessary for the well-founded choice of the spacecraft parameters and carrying out their flight and space tests, when the new special-purpose space laboratories are designed and constructed.

Probe measurements of plasma parameters in torch plasmatron

An electric probe for a torch plasmatron has been manufactured so that to match physical parameters for a descent spacecraft at the Earth atmosphere. There is no material able to bear temperatures 5000–8000 K for long, so the probe has been made of refractory metals in a ceramic shell and it has been used as a measuring unit for a short time, with its metal part and ceramic shell being collapsed simultaneously. The experimental measurement of plasma electro-conductivity is performed with the double probe inside a torch plasmatron. The oscillograms for electric current and voltage are obtained, the volt-ampere characteristics of various plasmas (nitrogen and argon) are plotted, the plasma electro-conductivity is experimentally determined and numerically calculated.

Combined spacecraft orbit and attitude control through extended Kalman filtering of magnetometer, gyro, and GPS measurements

The main goal of this research is to establish spacecraft orbit and attitude control algorithms based on extended Kalman filter which provides estimates of spacecraft orbital and attitude states. The control and estimation algorithms must be capable of dealing with the spacecraft conditions during the detumbling and attitude acquisition modes of operation. These conditions are characterized by nonlinearities represented by large initial attitude angles, large initial angular velocities, large initial attitude estimation error, and large initial position estimation error. All of the developed estimation and control algorithms are suitable for application to the next Egyptian scientific satellite, EGYPTSAT-2. The parameters of the case-study spacecraft are similar but not identical to the former Egyptian satellite EGYPTSAT-1. This is done because the parameters of EGYPTSAT-2 satellite have not been consolidated yet. The sensors utilized are gyro, magnetometer, and GPS. Gyro and magnetometer are utilized to provide measurements for the estimates of spacecraft attitude state vector where as magnetometer and GPS are utilized to provide measurements for the estimates of spacecraft orbital state vector.

Application of the Modified Method of Transporting Trajectory to Optimize Interplanetary Transfers Combining Low and High Thrust

An approach to optimize trajectories of interplanetary transfers combining low and high trust is proposed. The approach employs the modified method of transporting trajectory to optimize the heliocentric section of the trajectory. An Earth–Mars transfer in 180 astronomical days is used as an example to compare the optimal trajectories and parameters of spacecraft predicted by the classical and modified methods of transporting trajectory

How to Really Measure Low Energy Electrons in Space

There is little argument in the space plasma physics community that in-situ, low energy electron measurements are technically challenging. The primary obstacle has been the effects of spacecraft charging on the measured three-dimensional electron velocity space distribution. A successful spacecraft charging correction algorithm used with the three-dimensional electron instrument aboard the Ulysses spacecraft has clarified the role of spacecraft and instrument parameters in the eventual reconstruction of low energy electron distributions. Suggestions for instrument and spacecraft modifications that can minimize spacecraft charging effects are presented in this paper. The emphasis is on designs that lend themselves to robust correction algorithms.

In-Flight Dynamical Method to Verify Sample Collection for Small-Body Sample Return Mission

An in-flight dynamical method is introduced for use by sample-collection missions to identify the presence and quantity of collected sample material. The method, denoted as G-Sample, implements a maximum-likelihood estimator to identify the collected samplemass, based on onboard force-sensormeasurements, thruster firings, and a dynamicsmodel of the spacecraft. In terms of hardware, G-Sample requires only a force sensor. Hardware complexities are avoided by treating sample mass determination largely as a computational problem in modern estimation, rather than requiring the need for additional specialized equipment or sophisticated sensing architectures. Simulation examples are provided for a spacecraft configuration with a sample-collection device mounted on the end of an extended boom. In one representative example, a 1000 g sample mass is estimated to within 110 g (95% confidence) under realistic assumptions of thruster profile error, spacecraft parameter uncertainty, and sensor noise. For convenience to future mission design, an overall sample-mass estimation error budget is developed to approximate the effect of model uncertainty, sensor noise, data rate, and thrust profile error on the expected estimate of collected sample mass.

Is the 130 GeV line real? A search for systematics in the Fermi-LAT data

Our recent claims of a Galactic center feature in Fermi-LAT data at approximately 130 GeV have motivated a large number of papers proposing explanations ranging from dark matter annihilation to monoenergetic pulsar winds. Because of the importance of such interpretations for physics and astrophysics, a discovery will require not only additional data, but a thorough investigation of possible LAT systematics. While we do not have access to the details of each event reconstruction, we do have information about each event from the public event lists and spacecraft parameter files. These data allow us to search for suspicious trends that could indicate a spurious signal. We consider several hypotheses that might make an instrumental artifact more apparent at the Galactic center, and find them implausible. We also search for an instrumental signature in the Earth limb photons, which provide a smooth reference spectrum for null tests. We find no significant 130 GeV feature in the Earth limb sample. However, we do find a marginally significant 130 GeV feature in Earth limb photons with a limited range of detector incidence angles. This raises concerns about the 130 GeV Galactic center feature, even though we can think of no plausible model of instrumental behavior that connects the two. A modest amount of additional limb data would tell us if the limb feature is a statistical fluke. If the limb feature persists, it would raise doubts about the Pass 7 processing of E > 100 GeV events. At present we find no instrumental systematics that could plausibly explain the excess Galactic center emission at 130 GeV.

Multidisciplinary Collaborative Optimization of General Parameters of Low Orbit Spacecraft

This article takes the optimization of general parameters of low-orbit spacecraft as object. Three subsystems of attitude and orbit control subsystem, power subsystem and structure subsystem were researched. Coupling relations between subsystems and system were described. Analysis models of the subsystems were provided. Meanwhile, optimization models of the researched problem were established based on collaborative optimization (CO). Sequential quadratic programming (NLPQL) was selected as search strategy of CO subsystem optimizer. While NLPQL, adaptive simulated annealing algorithm (ASA) and multi-island genetic algorithm (MIGA) were respectively selected as the search strategy of CO system-level optimizer to search. Furthermore, multidisciplinary feasible method (MDF) was used to optimize the same problem. NLPQL, ASA, MIGA were respectively selected as the search strategy of MDF system-level optimizer. The optimization result of CO was compared to the optimization result of MDF shows that using CO can get the result which close to the optimal result. That proves CO can be effectively applied to multidisciplinary design optimization of the general parameters of the low-orbit spacecraft.

Constant thrust fuel-optimal control for spacecraft rendezvous

In this paper, constant thrust rendezvous is studied and the optimal rendezvous time is calculated by using continuous genetic algorithm. Firstly, the relative position parameters of the target spacecraft are obtained by using the vision measurement and the target maneuver positions are calculated through the isochronous interpolation method. Then, the results of the calculation of constant thrust rendezvous is founded by processing with multivariate linear regression method. Next, a new switching control law is designed based on the thrust acceleration sequence and the on time of thrusters which can be computed by the time series analysis method. The perturbations and fuel consumptions are addressed during the computation of the on time of thrusters.At last, a monte carlo analysis is performed to calculate the relative velocities at which the rendezvous maneuvers can be carried out and simulation examples are given to illustrate the validity of the method presented in this paper. The simulation results show that with the switching control law can ensure the success of the rendezvous.