• International Journal of Control, Automation, and Systems
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• v.6 no.6
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• pp.859-872
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• 2008

In this paper, the problem of partial asymptotic stabilization of nonlinear control cascaded systems with integrators is considered. Unfortunately, many controllable control systems present an anomaly, which is the non complete stabilization via continuous pure-state feedback. This is due to Brockett necessary condition. In order to cope with this difficulty we propose in this work the partial asymptotic stabilization. For a given motion of a dynamical system, say x(t,$x_0,t_0$)=(y(t,$y_0,t_0$),z(t,$z_0,t_0$)), the partial stabilization is the qualitative behavior of the y-component of the motion(i.e., the asymptotic stabilization of the motion with respect to y) and the z-component converges, relative to the initial vector x($t_0$)=$x_0$=($y_0,z_0$). In this work we present new results for the adding integrators for partial asymptotic stabilization. Two applications are given to illustrate our theoretical result. The first problem treated is the partial attitude control of the rigid spacecraft with two controls. The second problem treated is the partial orientation of the underactuated ship.

• Coupled systems mechanics
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• v.2 no.4
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• pp.303-327
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• 2013

The coupled free vibration of flexible structures and on-board liquid in zero gravity space was analyzed, considering the spacecraft main body as a rigid mass, the flexible appendages as two elastic beams, and the on-board liquid as a "spring-mass" system. Using the Lagrangians of a rigid mass (spacecraft main body), "spring-mass" (liquid), and two beams (flexible appendages), as well as assuming symmetric motion of the system, we obtained the frequency equations of the coupled system by applying Rayleigh-Ritz method. Solving these frequency equations, which are governed by three system parameters, as an eigenvalue problem, we obtained the coupled natural frequencies and vibration modes. We define the parameter for evaluating the magnitudes of coupled motions of the added mass (liquid) and beam (appendages). It was found that when varying one system parameter, the frequency curves veer, vibration modes exchange, and the significant coupling occurs not in the region closest to the two frequency curves but in the two regions separate from that region.

• 제어로봇시스템학회:학술대회논문집
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• 1994.10a
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• pp.44-48
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• 1994

A new method of quaternion feedback control for the attitude acquisition of spacecraft is suggested to limit the angular rates of rigid body which are not desirable and make a control algorithm complicate. New attitude acquisition control algorithm is evaluated and compared with the existing quaternion feedback control method for the large slewing maneuvers through simulations. The simulation results reveal that a new method is effective on limiting the angular rates of spacecraft.

• Journal of Mechanical Science and Technology
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• v.18 no.2
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• pp.203-210
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• 2004

Attitude control law design for spacecraft large angle maneuvers is investigated in this paper. The feedback linearization technique is applied to the design of a nonlinear tracking control law. The output function to be tracked is the quaternion attitude parameter. The designed control law turns out to be a combination of attitude and attitude rate tracking commands. The attitude-only output function, therefore, leads to a stable closed-loop system following the given reference trajectory. The principal advantage of the proposed method is that it is relatively easy to produce reference trajectories and associated controller.

• Journal of Institute of Control, Robotics and Systems
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• v.7 no.7
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• pp.589-596
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• 2001

In this paper, a new design methodology for the optimal control of nonlinear systems described by the TS(Takagi-Sugeno) fuzzy model is proposed. First, a new theorem concerning the optimal stabilizing control of a general nonlinear dynamic system is proposed. Next, based on the proposed theorem and the inverse optimal approach, an optimal controller synthesis procedure for a TS fuzzy system is given, Also, it is shown that the optimal controller can be found by solving a linear matrix inequality problem. Finally, the proposed method is applied to the attitude control of a rigid spacecraft to demonstrate its validity.

• Advances in aircraft and spacecraft science
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• v.10 no.5
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• pp.419-437
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• 2023

This paper presents the study of the effects of rigid-body motion simultaneously with the presence of the effects of temporal variation due to the existence of morphing speed on the aeroelastic stability of the two-stage telescopic wings, and hence this is the main novelty of this study. To this aim, Euler-Bernoulli beam theory is used to model the bending-torsional dynamics of the wing. The aerodynamic loads on the wing in an incompressible flow regime are determined by using Peters' unsteady aerodynamic model. The governing aeroelastic equations are discretized employing a finite element method based on the beam-rod model. The effects of rigid-body motion on the length-based stability of the wing are determined by checking the eigenvalues of system. The obtained results are compared with those available in the literature, and a good agreement is observed. Furthermore, the effects of different parameters of rigid-body such as the mass, radius of gyration, fuselage center of gravity distance from wing elastic axis on the aeroelastic stability are discussed. It is found that some parameters can cause unpredictable changes in the critical length and frequency. Also, paying attention to the fuselage parameters and how they affect stability is very important and will play a significant role in the design.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.7
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• pp.26-32
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• 2005

In this paper, the vibration response of the spacecraft solar array is investigated. The solar array model consists of composite thin walled beam and solar blanket, spreader bar. The composite thin walled beam incorporates a number of nonclassical effects of transverse shear, primary and secondary warping, rotary inertia and anisotropy of constituent materials. The solar blanket is a membrane subjected to uniform tension in the z direction. The spreader bar is a rigid member. A coupled thermal structure analysis that includes the effects of structural deformations on heating and temperature gradient is investigated. A stability criterion given in parameters for establishes the conditions for thermal flutter.

• Journal of Astronomy and Space Sciences
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• v.14 no.2
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• pp.367-374
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• 1997

Slewing maneuver and vibration suppression control of flexible spacecraft model by Lyapunov stability theory are considered. The specific model considered in this paper consists of a rigid hub with an elastic appendage attached to the central hub and tip mass. Attitude control to point and stabilize single axis using reaction wheel type device is tested. To control all flexible modes is so critical to designing an active control law. We therefore considered an direct output feeback control design by using Lyapunov stability theory. It is shown that the ouput feedback control law design with proposed configuration gives satisfactory result in slewing performance and vibration suppression control.

• Advances in aircraft and spacecraft science
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• v.5 no.6
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• pp.653-669
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• 2018

This work has the objective to analyze multibody mechanisms of inflatable structures for manned space applications. The focus is on the evaluation of the main characteristics of MaxFlex, a new module of MSC Adams including the effect of nonlinear flexible bodies. MaxFlex integrates the nonlinear Finite Element Analysis (FEA) of Nastran-SOL400-and the Adams multibody capabilities in one unique solver, providing an improvement concerning the concept and technology based on the co-simulation among solvers. MaxFlex converts the equations of motion of the nonlinear FEA into phase-space form and discretizes them according to the multibody system integrator framework. The numerical results deal with an inflatable manned space module having rigid components and a flexible coating made of Kevlar. This paper is a preliminary assessment of the computational capabilities of the software and does not provide realistic guidelines for the actual design of the structure. The analysis leads to some recommendations related to the main issues to consider in a nonlinear simulation including both rigid and flexible components. The results underline the importance of realistic deployment times and applied forces. Also, a proper structural modeling is necessary, but can lead to excessive computational overheads.

• Advances in aircraft and spacecraft science
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• v.9 no.3
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• pp.263-278
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• 2022

The time-varying structural reliability of an aeroelastic launch vehicle subjected to stochastic parameters is investigated. The launch vehicle structure is under the combined action of several stochastic loads that include aerodynamics, thrust as well as internal combustion pressure. The launch vehicle's main body structural flexibility is modeled via the normal mode shapes of a free-free Euler beam, where the aerodynamic loadings on the vehicle are due to force on each incremental section of the vehicle. The rigid and elastic coupled nonlinear equations of motion are derived following the Lagrangian approach that results in a complete aeroelastic simulation for the prediction of the instantaneous launch vehicle rigid-body motion as well as the body elastic deformations. Reliability analysis has been performed based on two distinct limit state functions, defined as the maximum launch vehicle tip elastic deformation and also the maximum allowable stress occurring along the launch vehicle total length. In this fashion, the time-dependent reliability problem can be converted into an equivalent time-invariant reliability problem. Subsequently, the first-order reliability method, as well as the Monte Carlo simulation schemes, are employed to determine and verify the aeroelastic launch vehicle dynamic failure probability for a given flight time.