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Article
Publication date: 16 March 2012

Bai Zhengfeng and Zhao Yang

The purpose of this paper is to establish the dynamics model of spacecraft during deployment of oblique solar panel using Auto Dynamic Analysis of Mechanical System…

Abstract

Purpose

The purpose of this paper is to establish the dynamics model of spacecraft during deployment of oblique solar panel using Auto Dynamic Analysis of Mechanical System (ADAMS) and to study the attitude motion of the spacecraft system during the oblique solar panel deployment.

Design/methodology/approach

For the case of an oblique solar panel on spacecraft, the dynamics virtual prototype model of deployment of oblique solar panels on spacecraft is established and the dynamics simulation is carried out using ADAMS. The effects of solar panel deployment on the attitude motion of spacecraft with different oblique angles are studied and the attitude motion regularities of spacecraft system are discussed. First, the effects on attitude motion of spacecraft are compared between the normal solar panel deployment and oblique solar panel deployment on a spacecraft. Then the attitude motion of spacecraft during the deployment of solar panel with different oblique angles on spacecraft is studied.

Findings

The effects of oblique angle of solar panel deployment on the attitude motion of spacecraft are significant in yaw axis. The bigger the oblique angle, the bigger the changes of yaw angle of spacecraft. However, the bigger the oblique angle, the smaller the changes of roll angle of spacecraft. The effects of oblique angle on pitch angle of spacecraft are slight.

Practical implications

Providing a practical method to study the attitude motion of spacecraft system during deployment of solar panel and improving the engineering application of spacecraft system, which put forward up spacecraft system to the practical engineering.

Originality/value

The paper is a useful reference for engineering design of a spacecraft attitude control system and ground text.

Details

Aircraft Engineering and Aerospace Technology, vol. 84 no. 2
Type: Research Article
ISSN: 0002-2667

Keywords

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Article
Publication date: 1 November 2006

Ru Fang, Shijie Zhang and Xibin Cao

Hill equations have definite limitation in the design of multiple spacecraft formation flying in eccentric orbits. To solve the problem, the design method of spacecraft

Abstract

Purpose

Hill equations have definite limitation in the design of multiple spacecraft formation flying in eccentric orbits. To solve the problem, the design method of spacecraft formation flying in a circular reference orbit based on Hill equation can be generalized and applied to spacecraft formation flying in eccentric orbits.

Design/methodology/approach

In this paper, T‐H equation is expressed as the explicit function form of reference orbit true anomaly, and the state transition matrix of relative motion of spacecraft formation flying in eccentric orbits is derived. According to the requirement that relative dynamics equation of spacecraft formation flying in eccentric orbits has periodicity solution, the paper theoretically gives the initial condition needed by the long‐term close‐distance spacecraft formation flying including the relationship between relative position and relative velocity. Without perturbation the spacecraft formation, which satisfies the initial periodicity restriction, can keep long‐term close‐distance flying without the need of active control.

Findings

Based on the theoretical analysis, some numerical simulations are carried out. The results demonstrate that each spacecraft in eccentric orbits can run in a periodic motion surrounding the center spacecraft under some conditions. And spacecraft formation reconfiguration is implementing according to missions.

Originality/value

Combined with the periodicity restriction primary condition a new method about spacecraft formation reconfiguration is put forward. The method given by this paper can be applied to eccentric orbits of arbitrary eccentricity, and provides theoretical reference for orbit design of spacecraft formation flying in eccentric orbits.

Details

Aircraft Engineering and Aerospace Technology, vol. 78 no. 6
Type: Research Article
ISSN: 0002-2667

Keywords

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Article
Publication date: 19 March 2021

Chengxi Zhang, Jin Wu, Yulong Huang, Yu Jiang, Ming-zhe Dai and Mingjiang Wang

Recent spacecraft attitude control systems tend to use wireless communication for cost-saving and distributed mission purposes while encountering limited communication…

Abstract

Purpose

Recent spacecraft attitude control systems tend to use wireless communication for cost-saving and distributed mission purposes while encountering limited communication resources and data exposure issues. This paper aims to study the attitude control problem with low communication frequency under the sampled-data.

Design/methodology/approach

The authors propose constructive control system structures based on quantization and event-triggered methods for intra-spacecraft and multi-spacecraft systems, and they also provide potential solutions to shield the control system's data security. The proposed control architectures can effectively save communication resources for both intra-spacecraft and multi-spacecraft systems.

Findings

The proposed control architectures no longer require sensors with trigger-ing mechanism and can achieve distributed control schemes. This paper also provides proposals of employing the public key encryption to secure the data in control-loop, which is transmitted by the event-triggered control mechanism.

Practical implications

Spacecraft attempts to use wireless communication, yet the attitude control system does not follow up promptly to accommodate these variations. Compared with existing approaches, the proposed control structures can save communication resources of control-loop in multi-sections effectively, and systematically, by rationally configuring the location of quantization and event-triggered mechanisms.

Originality/value

This paper presents several new control schemes and a necessary condition for the employment of encryption algorithms for control systems based on event-based communication.

Details

Aircraft Engineering and Aerospace Technology, vol. 93 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

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Article
Publication date: 1 July 2003

Gianmarco Radice

The purpose of this paper is to present a new approach in the concept and implementation of autonomous micro‐spacecraft. The one true “artificial agent” approach to…

Abstract

The purpose of this paper is to present a new approach in the concept and implementation of autonomous micro‐spacecraft. The one true “artificial agent” approach to autonomy requires the micro‐spacecraft to interact in a direct manner with the environment through the use of sensors and actuators. As such, little computational effort is required to implement such an approach, which is clearly of great benefit for limited micro‐satellites. Rather than using complex world models, which have to be updated, the agent is allowed to exploit the dynamics of its environment for cues as to appropriate actions to achieve mission goals. The particular artificial agent implementation used here has been borrowed from studies of biological systems, where it has been used successfully to provide models of motivation and opportunistic behaviour. The so‐called “cue‐deficit” action selection algorithm considers the micro‐spacecraft to be a non‐linear dynamical system with a number of observable states. Using optimal control theory rules are derived which determine which of a finite repertoire of behaviours the satellite should select and perform. The principal benefits of this approach is that the micro‐spacecraft is endowed with self‐sufficiency, defined here to be the ability to achieve mission goals, while never placing itself in an irrecoverable position.

Details

Kybernetes, vol. 32 no. 5/6
Type: Research Article
ISSN: 0368-492X

Keywords

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Article
Publication date: 26 August 2014

Haoyang Cheng, John Page, John Olsen and Nathan Kinkaid

– This paper aims to investigate the decentralised strategy to coordinate the reconfiguration of multiple spacecraft.

Abstract

Purpose

This paper aims to investigate the decentralised strategy to coordinate the reconfiguration of multiple spacecraft.

Design/methodology/approach

The system of interest consists of multiple spacecraft with independent subsystem dynamics and local constraints, but is linked through their coupling constraints. The proposed method decomposes the centralised problem into smaller subproblems. It minimises the fuel consumption of multiple spacecraft performing a reconfiguration manoeuvre through an iterative computation. In particular, each agent optimises its individual cost function using the most recently available local solution for the other agents.

Findings

The simulation scenarios include spacecraft formation reconfiguration and close manoeuvres around obstacles were conducted. The simulation results showed the fast convergence of the proposed algorithm, while local and inter-vehicle constraints were maintained.

Originality/value

The main advantage of this approach is that it adopts a linear form of the objective function. This allows the local optimisation problem to be formulated as a mixed-integer, linear programming problem, most of which can be quickly solved with resort to commercial software.

Details

Aircraft Engineering and Aerospace Technology: An International Journal, vol. 86 no. 5
Type: Research Article
ISSN: 0002-2667

Keywords

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Article
Publication date: 19 June 2018

Chingiz Hajiyev and Ahmet Sofyali

The purpose of this paper is to present a two-stage approach for estimation of spacecraft’s position and velocity by indirect linear measurements from a single antenna.

Abstract

Purpose

The purpose of this paper is to present a two-stage approach for estimation of spacecraft’s position and velocity by indirect linear measurements from a single antenna.

Design/methodology/approach

In the first stage, direct nonlinear antenna measurements are transformed to linear x-y-z coordinate measurements of spacecraft’s position, and statistical characteristics of orbit determination errors are analyzed. Variances of orbit parameters’ errors are chosen as the accuracy criteria. In the second stage, the outputs of the first stage are improved by the designed Extended Kalman Filter for estimation of the spacecraft’s position and velocity on indirect linear x-y-z measurements.

Findings

The complex content of the measurement matrix in the conventional method causes periodic singularities in simulation results. In addition, the convergence of the filter using direct measurements is highly dependent on the initialization parameters’ values due to the nonlinear partial derivatives in the Jacobian measurement matrix. The comparison of the accuracy of both methods shows that the estimation by using indirect measurements reduces the absolute estimation errors. The simulation results show that the proposed two-stage procedure performs both with better estimation accuracy and better convergence characteristics. The method based on indirect measurements provides an unnoticeably short transient duration.

Practical implications

The proposed method can be recommended for satellite orbit estimation regarding its presented superiorities.

Originality/value

Inputting the single antenna measurements to the filter indirectly results in a quite simpler measurement matrix. As a result, the convergence of the filter is faster and estimation errors are lower.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 5
Type: Research Article
ISSN: 1748-8842

Keywords

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Article
Publication date: 5 March 2018

Xu Kang and Dechang Pi

The purpose of this paper is to detect the occurrence of anomaly and fault in a spacecraft, investigate various tendencies of telemetry parameters and evaluate the…

Abstract

Purpose

The purpose of this paper is to detect the occurrence of anomaly and fault in a spacecraft, investigate various tendencies of telemetry parameters and evaluate the operation state of the spacecraft to monitor the health of the spacecraft.

Design/methodology/approach

This paper proposes a data-driven method (empirical mode decomposition-sample entropy-principal component analysis [EMD-SE-PCA]) for monitoring the health of the spacecraft, where EMD is used to decompose telemetry data and obtain the trend items, SE is utilised to calculate the sample entropies of trend items and extract the characteristic data and squared prediction error and statistic contribution rate are analysed using PCA to monitor the health of the spacecraft.

Findings

Experimental results indicate that the EMD-SE-PCA method could detect characteristic parameters that appear abnormally before the anomaly or fault occurring, could provide an abnormal early warning time before anomaly or fault appearing and summarise the contribution of each parameter more accurately than other fault detection methods.

Practical implications

The proposed EMD-SE-PCA method has high level of accuracy and efficiency. It can be used in monitoring the health of a spacecraft, detecting the anomaly and fault, avoiding them timely and efficiently. Also, the EMD-SE-PCA method could be further applied for monitoring the health of other equipment (e.g. attitude control and orbit control system) in spacecraft and satellites.

Originality/value

The paper provides a data-driven method EMD-SE-PCA to be applied in the field of practical health monitoring, which could discover the occurrence of anomaly or fault timely and efficiently and is very useful for spacecraft health diagnosis.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 2
Type: Research Article
ISSN: 1748-8842

Keywords

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Article
Publication date: 2 January 2018

Yue Wang, Rui Zhong and Shijie Xu

The purpose of this paper is to assess the orbital perturbation caused by the gravitational orbit–attitude coupling of spacecraft in the proximity of asteroids.

Abstract

Purpose

The purpose of this paper is to assess the orbital perturbation caused by the gravitational orbit–attitude coupling of spacecraft in the proximity of asteroids.

Design/methodology/approach

The gravitational orbit–attitude coupling perturbation (GOACP), which has been neglected before in the close-proximity orbital dynamics about asteroids, is investigated and compared with other orbital perturbations. The GOACP has its origin in the fact that the gravity acting on a non-spherical extended body is actually different from that acting on a point mass located at the body’s center of mass, which is the approximated model in the orbital dynamics. Besides, a case study of a tethered satellite system is given by numerical simulations.

Findings

It is found that the ratio of GOACP to the asteroid’s non-spherical gravity is the order of ρ/ae, where ρ is the spacecraft’s characteristic dimension and ae is the asteroid’s mean radius. It can also be seen that as ρ increases, GOACP will also increase but the solar radiation pressure (SRP) will decrease due to the decreasing area-to-mass ratio. The GOACP will be more significant than SRP at small orbital radii for a large-sized spacecraft. Based on the results by analyses and simulations, it can be concluded that GOACP needs to be considered in the orbital dynamics for a large-sized spacecraft in the proximity of a small asteroid.

Practical implications

This study is of great importance for the future asteroids missions for scientific explorations and near-Earth objects mitigation.

Originality/value

The GOACP, which has been neglected before, is revealed and studied.

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Article
Publication date: 2 January 2018

Yong Xie, Pan Liu and Guoping Cai

The purpose of this paper is to present an on-orbit frequency identification method for spacecraft directly using attitude maneuver data. Natural frequency of flexible…

Abstract

Purpose

The purpose of this paper is to present an on-orbit frequency identification method for spacecraft directly using attitude maneuver data. Natural frequency of flexible solar arrays plays an important role in attitude control design of spacecraft with solar arrays, and its precision will directly affect the accuracy of attitude maneuver. However, when the flexibility of the solar arrays is large, because of air damping, gravity effect etc., the frequency obtained by ground test shows great error compared with the on-orbit real value. One solution to this problem is to conduct on-orbit identification during which proper identification methods are used to obtain the parameters of interest based on the real on-orbit data of spacecraft.

Design/methodology/approach

The observer/Kalman filter identification and eigensystem realization algorithm are used as identification methods, and the attitude maneuver controller is designed using the rigid-body dynamics method.

Findings

Two conclusions are drawn in this paper according to results of numerical simulations. The first one is that the attitude controller based on the rigid-body dynamics method is effective in attitude maneuver of the spacecraft. The second one is that the on-orbit parameter identification can be directly achieved by using attitude maneuver data of spacecraft without adding additional missions.

Practical implications

Based on the methods proposed in this paper, it is convenient to obtain the natural frequencies of the spacecraft using the data of the attitude maneuver, which may greatly reduce the cost of on-orbit identification test.

Originality/value

The way of obtaining natural frequencies based on attitude maneuver data of spacecraft provides high originality and value for practical application.

Details

Aircraft Engineering and Aerospace Technology, vol. 90 no. 1
Type: Research Article
ISSN: 1748-8842

Keywords

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Article
Publication date: 6 June 2008

Christie Alisa Maddock and Massimiliano Vasile

The purpose of this paper is to present a methodology and experimental results on using global optimization algorithms to determine the optimal orbit, based on the mission…

Abstract

Purpose

The purpose of this paper is to present a methodology and experimental results on using global optimization algorithms to determine the optimal orbit, based on the mission requirements, for a set of spacecraft flying in formation with an asteroid.

Design/methodology/approach

A behavioral‐based hybrid global optimization approach is used to first characterize the solution space and find families of orbits that are a fixed distance away from the asteroid. The same optimization approach is then used to find the set of Pareto optimal solutions that minimize both the distance from the asteroid and the variation of the Sun‐spacecraft‐asteroid angle. Two sample missions to asteroids, representing constrained single and multi‐objective problems, were selected to test the applicability of using an in‐house hybrid stochastic‐deterministic global optimization algorithm (Evolutionary Programming and Interval Computation (EPIC)) to find optimal orbits for a spacecraft flying in formation with an orbit. The Near Earth Asteroid 99942 Apophis (2004 MN4) is used as the case study due to a fly‐by of Earth in 2029 leading to two potential impacts in 2036 or 2037. Two black‐box optimization problems that model the orbital dynamics of the spacecraft were developed.

Findings

It was found for the two missions under test, that the optimized orbits fall into various distinct families, which can be used to design multi‐spacecraft missions with similar orbital characteristics.

Research limitations/implications

The global optimization software, EPIC, was very effective at finding sets of orbits which met the required mission objectives and constraints for a formation of spacecraft in proximity of an asteroid. The hybridization of the stochastic search with the deterministic domain decomposition can greatly improve the intrinsic stochastic nature of the multi‐agent search process without the excessive computational cost of a full grid search. The stability of the discovered families of formation orbit is subject to the gravity perturbation of the asteroid and to the solar pressure. Their control, therefore, requires further investigation.

Originality/value

This paper contributes to both the field of space mission design for close‐proximity orbits and to the field of global optimization. In particular, suggests a common formulation for single and multi‐objective problems and a robust and effective hybrid search method based on behaviorism. This approach provides an effective way to identify families of optimal formation orbits.

Details

International Journal of Intelligent Computing and Cybernetics, vol. 1 no. 2
Type: Research Article
ISSN: 1756-378X

Keywords

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