Search results

This paper aims to present the control laws to be used in the control of pendular motion on tethered satellite systems in space during orbiting by using a nonlinear design technique.

This work presents both physical and mathematical models represented in a circular orbit. Euler equation of the rigid body is applied under reasonable assumption so as to form the equations of pendular motion. These equations are then used to develop the control laws using a nonlinear design technique. The control laws are required to drive the in-plane angles and out-of-plane angles of the pendular motion to the required trajectories. Simulations are then conducted to study the control results.

Simulation results show that the control laws in both plane angles of motions considered are able to move the pendular motion to the required trajectory. It was also eminent that a lot of effort is required in the case of the reference trajectory that corresponds to the constant inside-plane. To control the pendular motion of the plane, one requires an extended period of time and it should be controlled within a reasonable range. In the outside-of-plane pendular motion, minimal or no effort is required for the control. The reason is that the trajectory is natural planar.

This research is expected to provide a dynamic control strategy for all tethered satellite space systems.

The research proposes a combined dynamic method for the purpose of improving the control of all types of tether satellites.

The purpose of this paper is to study the dynamic vibrations of the tethered satellite system (TSS).

The energy principle and the variational approach are used to establish the dynamic equations of the TSS. By introducing new generalized coordinates, the equations are transformed into the Hamiltonian system. Then, the symplectic Runge-Kutta (SRK) method is used to solve the canonical equations.

The influence of the tether length on the dynamic behavior of the TSS is very important.

The dynamic responses of the TSS are obtained by using the SRK method.

The purpose of this paper is to develop analytic solutions for a tethered satellite system (TSS) subjected to internal tether tension moment and external aerodynamic torque for spin-up and spin-down manoeuvres.

Analytic solutions for TSS based on the approximation of Euler’s equations of motion via Fresnel integrals and sine and cosine integrals. Test simulation was performed for two cases (spin-up and spin-down manoeuvres). The conclusion is based on graphical interpretation.

The effects of angular velocities on X, Y and Z axes of the TSS under the influence of combined torques from internal tether tension and external aerodynamic drag influenced during spinning manoeuvres are shown graphically.

This research focuses only on a circular orbit, which is one of the simplest orbits without many variables taken into account such as flight path angle and true anomaly. It could get quite complex for other orbit types like elliptic and parabolic orbits.

Practical implications include observing the stability rotational motion of TSS so as to perform a two-way payload exchange via momentum transfer.

In this paper, analytic solutions for a torque motion of a TSS comprising non-linear Euler’s equations of motion are established for spin-up and spin-down manoeuvres.

This paper aims to present the impact analysis of payload rendezvous with tethered satellite system and the design of an adaptive sliding mode controller which can deal with mass parameter uncertainty of targeted payload, so that the proposed cislunar transportation scheme with spinning tether system could be extended to a wider and more practical range.

In this work, dynamical model is first derived based on Langrangian equations to describe the motion of a spinning tether system in an arbitrary Keplerian orbit, which takes the mass of spacecraft, tether and payload into account. Orbital design and optimal open-loop control for the payload tossed by the spinning tether system are then presented. The real payload rendezvous impact around docking point is also analyzed. Based on reference acceleration trajectory given by optimal theories, a sliding mode controller with saturation functions is designed in the close-loop control of payload tossing stage under initial disturbance caused by actual rendezvous error. To alleviate the influence of inaccurate/unknown payload mass parameters, the adaptive law is designed and integrated into sliding mode controller. Finally, the performance of the proposed controller is evaluated using simulations. Simulation results validate that proposed controller is found effective in driving the spinning tether system to carry payload into desired cislunar transfer orbit and in dealing with payload mass parameter uncertainty in a relatively large range.

The results show that unideal rendezvous manoeuvres have significant impact on in-plane motion of spinning tether system, and the proposed adaptive sliding mode controller with saturation functions not only guarantees the stability but also provides good performance and robustness against the parameter and unstructured uncertainties.

This work addresses the analysis of actual impact on spinning tether system motion when payload is docking with system within tolerated docking window, rather than at the particular ideal docking point, and the robust tracking control of deep-space payload tossing missions with the spinning tether system using the adaptive sliding mode controller dealing with parameter uncertainties. This combination has not been proposed before for tracking control of multivariable spinning tether systems.

The purpose of this paper is to present kinetic equations for tether-net system during deorbiting using a novel method differing from the traditional method.

The work presents kinetic equations for tether-net system in which the tether exhibits tensional and tensionless states alternately during deorbiting. Orbital position coordinates of net-capture and abandoned spacecrafts are adopted as generalized coordinates above-mentioned instead of librations and the length of the tether. Geostationary orbit (GEO) and the orbit whose apogee is 300 km above GEO are chosen as the initial and target orbit, respectively. Simulations are conducted to study the deorbiting results considering a variety of parameters and initial conditions.

The distinctive dynamic characteristics of tether-net system can be seen by kinetic equations based on the proposed dynamic modeling strategies. Moreover, the deorbiting results are deeply affected by the initial tension force and librations showed by simulations. The initial tension force and librations should be controlled within a reasonable range.

This is expected to provide dynamic modeling strategies for space tether-net system during deorbiting. Moreover, the preliminary principle of choosing initial conditions and parameters to meet the requirements for deorbiting can be achieved.

The research proposes a novel dynamic modeling method for space tether-net system that differs from traditional tethered system, and also proposes a superior librations expression based on orbital position coordinates of net-capture and abandoned spacecrafts.

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.

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.

It is found that the ratio of GOACP to the asteroid’s non-spherical gravity is the order of ρ/a_e, where ρ is the spacecraft’s characteristic dimension and a_e 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.

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

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

The purpose of this paper is to study the effect of J₃ perturbation of the Earth’s oblateness on satellite orbit compared with J₂ perturbation.

Based on the parametric variation method in the time domain, considering more accurate Earth potential function by considering J₃-perturbation effect, the perturbation equations about satellite’s six orbital elements (including semi-major axis, orbit inclination, right ascension of the ascending node, true anomaly, eccentricity and argument of perigee) has been deduced theoretically. The disturbance effects of J₂ and J₃ perturbations on the satellite orbit with different orbit inclinations have been studied numerically.

With the inclination increasing, the maximum of the semi-major axis increases weakly. The difference of inclination disturbed by the J₂ and J₃ perturbation is relative to orbit inclinations. J₃ perturbation has weak effect on the right ascension and argument of perigee. The critical angle of the right ascension and argument of perigee which decides the precession direction is 90° and 63.43°, respectively. The disturbance effects of J₂ and J₃ perturbations on the argument of perigee, right ascension and eccentricity are weakened when the eccentricity increases, simultaneously, the difference of J₂ and J₃ perturbations on argument of perigee, right ascension and argument of perigee decreases with eccentricity increasing, respectively.

In the future, satellites need to orbit the Earth much more precisely for a long period. The J₃ perturbation effect and the weight compared to J₂ perturbation in LEO can provide a theoretical reference for researchers who want to improve the control accuracy of satellite. On the other hand, the theoretical analysis and simulation results can help people to design the satellite orbit to avoid or diminish the disturbance effect of the Earth’s oblateness.

The J₃ perturbation equations of satellite orbit elements are deduced theoretically by using parametric variation method in this paper. Additionally, the comparison studies of J₂ perturbation and J₃ perturbation of the Earth’s oblateness on the satellite orbit with different initial conditions are presented.

Giant fan blades weighing approximately 500lbs each are being built by Permali Gloucester Ltd for use in a NASA wind tunnel in the USA.

An aerospace worker at Gulf‐stream Aerospace Corp in Savannah, Georgia, instals a lightweight graphite/epoxy composite spoiler manufactured by Lockheed‐Georgia Company on a new Gulfstream IV coporate jet.

The micro‐satellite clusters have been discussed for several years, however, there is not a common framework about its software, and various researches distributed at different domains. In order to conduct the future work well, the purpose of this paper is to systematically describe micro‐satellite clusters' characteristics, clusters software model, and present a distributed testbed to shorten test process, and minimize the development cost.

The cluster characteristics and model is summarized through analyzing the past satellite cluster programs. Then the ground test system is designed to shorten micro‐satellite's development period, improve its reliability.

The clusters' characteristics are discussed, such as coverage, scalability, fault tolerance, low cost, etc. The clusters' data flow and on‐board software architecture are presented according to properties of clusters. Finally, the distributed testbed that focuses on future on‐board software and hardware technologies that aim to rapid design, build, integration, test, deployment, and operation of the future micro‐satellite is designed.

The presentation of software architecture of cluster member can improve the micro‐satellite's development, and the distributed testbed can improve the ground test efficiency, especially, when the micro‐satellite quantity is big.