Search results

The purpose of this paper is to propose a modified fuel-balanced formation keeping strategy based on actively rotating satellites in the formation in the J2 perturbed environment.

Based on the relative orbital elements theory, the J2 perturbed relative motions between different satellites in the formation are analyzed, and then, the method to estimate fuel required to keep the in-plane and out-of-plane relative motions is presented, based on which a modified fuel-balanced formation keeping strategy is derived by considering both in-plane and out-of-plane J2 perturbations.

Numerical simulations demonstrate that the modified fuel-balanced formation keeping strategy is valid, and the modified fuel-balanced formation keeping strategy requires less total fuel consumption than original Vadali and Alfriend’s method.

The modified fuel-balanced formation keeping strategy is valid for formation flying mission whose member satellite is in circular or near-circular orbit.

The modified fuel-balanced formation keeping strategy can be used to solve formation flying keeping problem, which involves multiple satellites in the formation.

The modified fuel-balanced formation keeping strategy is proposed by considering both in-plane and out-of-plane J2 perturbations, which further reduce the fuel consumption than the original Vadali and Alfriend’s method.

Current procedures of buckling load estimation for thin-walled structures may provide very conservative estimates. Their refinement offers the potential to use structure and material properties more efficiently. Due to the large variety of design variables, for example laminate layup in composite structures, a prohibitively large number of tests would be required for experimental assessment, and thus reliable numerical techniques are of particular interest. The purpose of this paper is to analyze different methods of numerical buckling load estimation, formulate simulation procedures suitable for commercial software and give recommendations regarding their application. All investigations have been carried out for cylindrical composite shells; however similar approaches are feasible for other structures as well.

The authors develop a concept to apply artificial load imperfections with the aim to estimate as good as possible lower bounds for the buckling loads of shells for which the actual physical imperfections are not known. Single and triple perturbation load approach, global and local dynamic perturbation approach and path following techniques are applied to the analysis of a cylindrical composite shell with known buckling characteristics. Results of simulations are compared with published experimental data.

A single perturbation load approach is reproduced and modified. Buckling behavior for negative values of the perturbation load is examined and a pattern similar to a positive perturbation load is observed. Simulations with three perturbation forces show a decreased (i. e. more critical) value of the buckling load compared to the single perturbation load approach. Global and local dynamic perturbation approaches exhibit a behavior suitable for lower bound estimation for structures with arbitrary geometries.

Various load imperfection approaches to buckling load estimation are validated and compared. All investigated methods do not require knowledge of the real geometrical imperfections of the structure. Simulations were performed using a commercial finite element code. Investigations of sensitivity with respect to a single perturbation load are extended to the negative range of the perturbation load amplitude. A specific pattern for a global perturbation approach was developed, and based on it a novel simulation procedure is proposed.

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.

The measurement of geomagnetic field can provide a reliable and economical basis for attitude and orbit information of low earth orbiting satellite. Because the earth's magnetic field is a function of position, and its measurement on the orbit are fully observable, orbit estimation can be obtained using extend Kalman filter (EKF) algorithm. With the assistant of angle velocity information from gyro measurement, attitude estimation can also be obtained. At the same time, gyro drift rate estimation is a part of the filter output. Although orbit and attitude determination are independent of each other, the filter can give the orbit and attitude estimation at the same time. The results of the numerical test show that a signal EKF can estimate both orbit and attitude by using magnetometer and gyro measurement only. The accuracy, usually is sufficient for low earth orbiting satellites.

The purpose of this paper is to conduct a reliable remote manipulation with good contact perception of the remote site. The long-term experience of the authors’ repeatedly confirm that the highest relevance lies in monitoring the wrench acting at a structurally weak point of the work piece rather than monitoring the wrench experienced by the robot end-effector.

The approach followed here is to sense the wrench at the interface of the robot end-effector and the environment. Position and orientation data and environment model are used to arrive at the contact point in real time. The intent of remote contact procedure is understood based on the knowledge of motion trajectory. All the above information is used to develop a wrench transformation to obtain the force diagrams.

The haptic solutions greatly suffer from objectivity, and therefore may result in inconsistency in an operator’s role. Intermediary telepresence through the visual communication of the wrench at the remote site in the form of force diagram provides excellent consistency across the operators and operations. Observing six components of the wrench in separate graphs does not provide on-line error estimate. Force diagrams suggested in the paper are found to be highly effective in perceiving the wrench.

The contact mode operations like assembly, surgery, docking, etc. still suffer due to the lack of easily perceivable wrench visualization. This paper provides solution to such practical issues.

The concept is original, and has evolved steadily over a period of time.

The purpose of this paper is to present a method to conduct small satellite rendezvous mission by using the differential aerodynamic forces under J2 perturbation in low earth orbit (LEO).

Each spacecraft is assumed to be equipped with two large flat plates, which can be controlled for generating differential accelerations in all three directions. Based on the kinetic theory, the aerodynamic lift and drag generated by a flat plate are calculated. To describe the relative dynamics under J2 perturbation, a modified model is derived from the high-fidelity linearized J2 equations proposed by Schweighart and Sedwick.

Simulation results demonstrate that the proposed method is valid and efficient to solve satellite rendezvous problem, and the modified model considering J2 effect shows better accuracy than the Horsley’s Clohessy–Wiltshire-based model.

Because aerodynamic force will reduce drastically as orbital altitude rises, the rendezvous control strategy for small satellites presented in this paper should be limited to the scenarios when satellites are in LEO.

The rendezvous control method in this paper can be applied to solve satellite rendezvous maneuver problem for small satellites in LEO.

This paper proposes a modified differential aerodynamic control model by considering J2 perturbation, and simulation results show that it can achieve higher rendezvous control accuracy.

The equation of motion of a free‐free moving column with a prescribed axial acceleration is formulated based on Hamilton’s principle and the assumed mode method. The slender column is used as a simple model for a missile or a launched vehicle. The column is under the action of aerodynamic forces which are modelled as axial external loads at the ends of the column. The effects of axial sinusoidal perturbations in respect of axial acceleration and the external loads are then examined using Bolotin’s method. The respective regions of instability are determined by converting the resulting equations of boundary frequencies to the standard form of a generalized eigenvalue problem. Instability regions for various combinations of external loads and axial acceleration of the column are presented.

This paper aims to investigate the feasibility of using the combination of Lorentz force and aerodynamic force as a propellantless control method for spacecraft formation.

It is assumed that each spacecraft is equipped with several large flat plates, which can rotate to produce aerodynamic force. Lorentz force can be achieved by modulating spacecraft’s electrostatic charge. An adaptive output feedback controller is designed based on a sliding mode observer to account for unknown uncertainties and the absence of relative velocity measurements. Aiming at distributing the control input, an optimal control allocation method is proposed to calculate the electrostatic charge of the Lorentz spacecraft and control commands for the atmospheric-based actuators.

Numerical examples are provided to demonstrate the effectiveness of the proposed control strategy in the presence of J₂ perturbations. Simulation results show that relative motion in a formation can be precisely controlled by the proposed propellantless control method under uncertainties and unavailability of velocity measurements.

The controllability of the system is not theoretically investigated in the current work.

The proposed control method introduced in this paper can be applied for small satellites formation in low Earth orbit.

The main contribution of the paper is the proposal of the propellantless control approach for satellite formation using the combination of Lorentz force and aerodynamic force, which can eliminate the requirement of the propulsion system.

The purpose of this paper is to establish a model‐based framework allowing the simulation, analysis and optimization of friction stir welding (FSW) processes of metallic structures using industrial robots, with a particular emphasis on the assembly of aircraft components made of aerospace aluminum alloys.

After a first part of the work dedicated to the kinetostatic and dynamical identification of the robotic mechanical system, a complete analytical model of the robotized process is developed, incorporating a dynamic model of the industrial robot, a multi‐axes macroscopic visco‐elastic model of the FSW process and a force/position control unit of the system. These different modules are subsequently implemented in a high‐fidelity multi‐rate dynamical simulation.

The developed simulation infrastructure allowed the research team to analyze and understand the dynamic interaction between the industrial robot, the control architecture and the manufacturing process involving heavy load cases in different process configurations. Several critical process‐induced perturbations such as tool oscillations and lateral/rotational deviations are observed, analyzed, and quantified during the simulated operations.

The presented simulation platform will constitute one of the key technology enablers in the major research initiative carried out by NRC Aerospace in their endeavor to develop a robust robotic FSW platform, allowing both the development of optimal workcell layouts/process parameters and the validation of advanced real‐time control laws for robust handling of critical process‐induced perturbations. These deliverables will be incorporated in the resulting robotic FSW technology packaged for deployment in production environments.

The paper establishes the first model‐based framework allowing the high‐fidelity simulation, analysis and optimization of FSW processes using serial industrial robots.

The purpose of this paper is to analyze the inclusion effects of the linearized time‐varying J₂‐perturbed terms for relative accuracy increase significantly over Melton's problem.

The methodology is based on the previous studies provided by Ross's paper. He gives a set of equations based on the C‐W equations which incorporates the J₂ gravitational perturbations and states in his introduction that this method can be expanded for the elliptical reference orbits as described by Melton.

One challenge in implementing the relative motions is maintaining the relations as it experiences gravitational perturbations, most notably due to non‐spherical Earth. Simulation results show that the inclusion of time‐varying J₂ perturbation terms in the derived linear equations increased the accuracy of the solution significantly in the out‐of‐orbit‐plane direction, while the accuracy within the orbit plane remained roughly unchanged.

By reason of replacing approximate terms (e, M) in this solution, for continues accuracy increase of time‐varying parameters containing θ(t) and R_O(t), this solution could be useful in the element‐errors evaluation and analysis of orbital multiple rendezvous missions, that are involved to the limited orbit periods.

The originality of this paper is to develop Melton's researches. He provided a method for generalizing the linear equations of motion to an elliptical orbit which enabled the determination of a time‐explicit, approximate solution. In this regard, there is no paper which has evaluated the inclusion effects of the linearized time‐varying J₂ perturbation terms over Melton's equations with such an approach.