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This paper aims to present a deflection model to improve positional accuracy of industrial robots. Earlier studies have demonstrated the lack of accuracy of heavy-duty robots when exposed to high external forces. One application where the robot is pushed to its limits in terms of forces is friction stir welding (FSW). This process requires the robot to deliver forces of several kilonewtons causing deflections in the robot joints. Especially for robots with serial kinematics, these deflections will result in significant tool deviations, leading to inferior weld quality.

This paper presents a kinematic deflection model, assuming a rigid link and flexible joint serial kinematics robot. As robotic FSW is a process which involves high external loads and a constant welding speed of usually below 50 mm/s, many of the dynamic effects are negligible. The model uses force feedback from a force sensor, embedded on the robot, and predicts the tool deviation, based on the measured external forces. The deviation is fed back to the robot controller and used for online path compensation.

The model is verified by subjecting an FSW tool to an external load and moving it along a path, with and without deviation compensation. The measured tool deviation with compensation was within the allowable tolerance for FSW.

The model can be applied to other robots with a force sensor.

The presented deflection model is based on force feedback and can predict and compensate tool deviations online.

The purpose of this paper is to analyze the inelastic buckling of a rectangular thin flat isotropic plate subjected to uniform uniaxial in-plane compression using a work principle, a deformation plasticity theory and Taylor–Maclaurin series formulation.

The non-loaded longitudinal edges of the rectangular plate are clamped, whereas the loaded edges are simply supported (CSCS). Total work error function is applied to Stowell’s plasticity theory in the derivation of the inelastic buckling equation. Mathematical formulation of the Taylor–Maclaurin series deflection function satisfied the boundary conditions of the CSCS rectangular plate. The critical inelastic load of the plate is then derived by applying variational principles.

Values of the plate buckling coefficient are calculated using various values of moduli ratio for aspect ratios ranging from 0.1 to 1.0, in intervals of 0.1. The accuracy of the proposed technique is validated by comparing the results obtained in the present study with solutions from a previous investigation. The percentage differences in the values of the buckling coefficient ranged from −0.122 to −4.685 per cent.

The results indicate that the work principle approach can be used as an alternative approximate method for analyzing inelastic buckling of rectangular thin flat isotropic plates under uniform in-plane compressive loads.

THE purpose of the present report is to present tables and to summarize formulas which are required for the analysis of circular rings, under forces in the plane of the ring or out of the plane. Usually, rings which are redundant structures are calculated by considering strain due to bending only. It is shown that for the stress distribution in a circular ring under given loads, the solution introducing strains due to shearing and axial loads is the same as the solution conventionally used.

Extreme loads are generated in aircraft flight manoeuvres. Among different manoeuvres considered in this paper are motions following a sudden deflection of elevator and response to a vertical gust. Airplane was assumed to be a rigid body of three degrees of freedom in symmetrical motion. Elevator deflection was either of the step change type, or of the sinusoidal type, gust was assumed to be either of the step change type or harmonic, with a gust cycle time corresponding to the time to travel a distance equal to 25 Mean Aerodynamic Chord. In all cases a jump type elevator deflection was assumed to last for 1 second, whilst the airplane response was observed for 3 seconds. The airplane motion, its velocities, accelerations and load acting on the tailplane were calculated by means of numerical integration of the ordinary differential /of motion.

Nonlinear model predictive control (NMPC) is emerging as a way to control unmanned aircraft with flight control constraints and nonlinear and unsteady aerodynamics. However, these predictive controllers do not perform robustly in the presence of physics-based model mismatches and uncertainties. Unmodeled dynamics and external disturbances are unpredictable and unsteady, which can dramatically degrade predictive controllers’ performance. To address this limitation, the purpose of this paper is to propose a new systematic approach using frequency-dependent weighting matrices.

In this framework, frequency-dependent weighting matrices jointly minimize closed-loop sensitivity functions. This work presents the first practical implementation where the frequency content information of uncertainty and disturbances is used to provide a significant degree of robustness for a time-domain nonlinear predictive controller. The merit of the proposed method is successfully verified through the design, coding, and numerical implementation of a robust nonlinear model predictive controller.

The proposed controller commanded and controlled a large unmanned aerial system (UAS) with unsteady and nonlinear dynamics in the presence of environmental disturbances, measurement bias or noise, and model uncertainties; the proposed controller robustly performed disturbance rejection and accurate trajectory tracking. Stability, performance, and robustness are attained in the NMPC framework for a complex system.

The theoretical results are supported by the numerical simulations that illustrate the success of the presented technique. It is expected to offer a feasible robust nonlinear control design technique for any type of systems, as long as computational power is available, allowing a much larger operational range while keeping a helpful level of robustness. Robust control design can be more easily expanded from the usual linear framework, allowing meaningful new experimentation with better control systems.

Such algorithms allows unstable and unsteady UASs to perform reliably in the presence of disturbances and modeling mismatches.

Under this heading are published regularly abstracts of all Reports and Memoranda of the Aeronautical Research Council, Reports and Technical Memoranda of the United States National Advisory Committee for Aeronautics and publications of other similar Research Bodies as issued.

The present study is based on finding the structural response of a tensile membrane structure (TMS) through deformation. The intention of the present research is to develop a basic understanding of reliability analysis and deflection behavior of a pre-tensioned TMS. The mean value first-order second-moment method (MVFOSM) method is used here to evaluate stochastic moments of a performance function with random input variables. Results suggest the influence of modulus of elasticity, the thickness of the membrane, and edge span length are significant for reliability based TMS design.

A simple TMS is designed and simulated by applying external forces (along with prestress), as a manifestation of wind and snow load. A nonlinear analysis is executed to evaluate TMS deflection, followed by calculating the reliability index. Parametric study is done to consider the effect of membrane material, thickness and load location. First-order second moment (FOSM) is used to evaluative the reliability. A comparison of reliability index is done and deflection variations from μ − 3s to μ + 3s are accounted for in this approach.

The effectiveness of deflection is highlighted for the reliability assessment of TMS. Reliability and parametric study collectively examine the proposed geometry and material to facilitate infield design requirements. The estimated β value indicates that most suitable fabric material for a simple TMS should possess an elasticity modulus in the range of 1,000–1,500 MPa, the thickness may be considered to be around 1.00 mm, and additional adjustment of around 5–10 mm is suggested for edge length. The loading position in case of TMS structures can be a sensitive aspect where the rigidity of the surface is dependent on the pre-tensioning of the membrane.

The significance of the parametric study on material and loading for deflection of TMS is emphasized. Due to the lack of consolidated literature in the field combining reliability with deflection limits of a TMS, this work can be very useful for researchers.

The present work outcome may facilitate practitioners in determining effective design methodology and material selection for TMS construction.

The significance of parametric study for serviceability criteria is emphasized. Parameters like pre-stress can be included in future parametric studies to witness in depth behavior of TMS. Due to lack of consolidated literature in the field combining reliability with deflection limits of a TMS, this work can be very useful for the researchers. The present work outcome may facilitate practitioners in determining effective design methodology and material selection for TMS construction.

Presents a procedure for obtaining an improved finite element solution of boundary problems by estimating the principle of exact displacement method in the finite element technique. The displacement field is approximated by two types of functions: the shape functions satisfying the homogeneous differential equilibrium equation and the full clamping element functions as a particular solution of the differential equation between the nodes. The full clamping functions represent the solution of the full clamping state on finite elements. An improved numerical solution of displacements, strains, stresses and internal forces, not only at nodes but over the whole finite element, is obtained without an increase of the global basis, because the shape functions are orthogonal with the full clamping functions. This principle is generally applicable to different finite elements. The contribution of introducing two types of functions based on the principle of the exact displacement method is demonstrated in the solution procedure of frame structures and thin plates.

The purpose of this paper is to develop and apply an exact finite strip (F‐a FSM) for the buckling and initial post‐buckling analyses of box section struts.

The Von‐Karman's equilibrium equation is solved exactly to obtain the buckling loads and deflection modes for the struts. The investigation is then extended to an initial post‐buckling study with the assumption that the deflected form immediately after the buckling is the same as that obtained for the buckling. Through the solution of the Von‐Karman's compatibility equation, the in‐plane displacement functions are developed in terms of the unknown coefficient. These in‐plane and out‐of‐plane deflected functions are then substituted in the total strain energy expressions and the theorem of minimum total potential energy is applied to solve for the unknown coefficient.

The F‐a FSM is applied to analyze the buckling and initial post‐buckling behavior of some representative box sections for which the results were also obtained through the application of a semi‐energy finite strip method (S‐e FSM). For a given degree of accuracy in the results, however, the F‐a FSM analysis requires less computational effort.

In the present F‐a FSM, only one of the calculated deflection modes is used for the initial post‐buckling study.

A very useful and computationally economical methodology is developed for the initial design of struts which encounter post‐buckling.

The originality of the paper is the fact that by incorporating a rigorous buckling solution into the Von‐Karman's compatibility equation, and solving it, a fairly efficient method for post‐buckling stiffness calculation is achieved.

This paper presents first sight on the longitudinal control strategy for an aircraft in the tandem wing configuration. It is an aerodynamic strongly coupled configuration that needs a lot of detailed aerodynamic analysis which describes the mutual impact of the main parts of the aircraft. The purpose of this paper is to build the numerical model that allows to make an analysis of necessary flaps (front and rear) deflection and prepare the control strategy for this kind of aircraft.

Aircrafts’ aerodynamic characteristics were obtained using the MGAERO software which is a commercial computing fluid dynamics tool created by Analytical Methods, Inc. This software uses the Euler flow model. Results from this software were used in the static stability evaluation and trim condition analysis. The trim conditions are the outcome of the optimisation process whose goal was to find the best front and rear flap deflection to achieve the best lift to drag (L/D) ratio.

The main outcome of this investigation is the proposal of strategy for the front and rear flap deflection which ensured the maximum L/D ratio and satisfied the trim condition. Moreover, the analysis of the mutual impact of the front and rear wings and the analysis of the control surface impact on the aerodynamic characteristic of the aircraft are presented.

In terms of aerodynamic computation, MGAERO software uses an inviscid flow model. However, this research is for the conceptual stage of the design and the MGAERO software grantee satisfied accurate respect to relatively low time of computations.

The ultimate goal is to build an aircraft in a tandem wing configuration and to conduct flying tests or wind tunnel tests. The presented result is one of the milestones to achieve this goal.

The aircraft in the tandem wing configuration is an aerodynamic-coupled configuration that needs detailed analysis to find the mutual interaction between the front and rear wings. Moreover, the mutual impact of the front and rear flaps is necessary too. Obtaining these results allowed this study to build the numerical model of the aircraft in the tandem wing configuration. It allows to find the best strategy of flap deflection, which allows to obtain the maximum L/D ratio and satisfy the trim condition.