Finite element analysis of a fuselage

We conclude that the single-plane with ten actuators scheme is feasible for the shape control, and the actuators do not damage the fuselage. This condition is not valid for thick plates in which there are considerable stresses developed normal to the surface of the plate.

The stress distribution on the plate is altered by the existence of the stress concentration factor due to the existence of the discontinuity — circular hole. It provides the space for power plant, cargo, controls, accessories, passengers and other requirements depending on the purpose i.

The discovered flaws or cracks are then analyzed for their criticality. The long and short length approximations are matching with the experimental results on their respective domains.

Radial stresses at the root of the notch Figure 1.

The symmetry boundary condition is applied on the lines on left of the plate and the line on the Finite element analysis of a fuselage of the plate and ahead of the crack tip.

Firstly, we develop a finite element model with detailed material property, ply design, fixture structure, and actuators installation considered. This is essential to accurately capture the high gradients of stress variations near the crack tip. The direction of the crack is defined by selecting the three adjacent left nodes from the crack tip.

Fracture mechanics, combined with the conventional fatigue design modals now became an integral part of mechanical engineering design. In order to improve the dimensional quality and increase the productivity, a new shape control system has been proposed to conduct dimensional shape adjustment before the assembly process.

For the initial analysis in the section 1. Thus, plane stress assumption is valid in this case. Substituting these values at the above equation we get,?? Suppose a given plate with a crack, is loaded in tension. There are many commercial tools available for solving the real life problems using FE methods.

By using finite element analysis, we conduct the feasibility analysis of this new shape control system. But, the long length approximated equation fail to achieve satisfactory results in this case. The state of stress existing in the portion of the fuselage between the rivet holes is considered to be similar to that of a wide thin plate subjected to uniform stress containing a circular hole in the middle Figure 1.

The finite element model is then validated and calibrated by physical experimental data. A new composite fuselage shape control system is proposed. For the verification of its results, the analysis is repeated by varying the length of singular elements.

Thus, it is dangerous to use this equation for design purposes. The condition prevailing in this case is known as plane strain case. Airbus Corporation Due to their criticality, the aircraft fuselages are periodically inspected using NonDestructive Testing NDT methods for any potential flaws.

Simplified Model of Crack prevalent in relatively thin plates in which there is no considerable variation in displacements along the direction of thickness. The second set is calculated using the approximated equation for a short length crack from section 1.

From the theoretical calculations in section 1. Thus, it can be said that the FE solution is valid. However for a thick plate, material away from the free surfaces is not free to deform laterally as it is constrained by the surrounding material.

The tangential stress at any point on the plate is given by,?? The FE analysis can be repeated for different lengths of cracks in order to find out the relationship between the stress intensity factor and the crack length.

The condition is mainly Figure 1. From the figure, it is clear that the FEA results are most closely matching the empirical relation based on experimental results. But, the short length equation fails in this case.Finite Element Analysis of a Fuselage Crack. Topics: Finite element Now-a-days a number of finite element analysis packages are available commercially and number of users is increasing.

A user without a basic course on finite element analysis may produce dangerous results. Fracture in metals Objective Fracture parameters Verification study PowerPoint Presentation Finite Element Analysis Verification study: Unstiffened curved panel Bulging of crack in the fuselage FEA of Aircraft Fuselage Components of fuselage Finite Element Model Boundary Conditions and loading Von Mises Stress Stress distribution.

Finite Element Modeling for Fracture Mechanics Analysis of Aircraft Fuselage Structure Sandeepkumar Gowda1, 3Lakshminarayana H. V.2, Kiran Kumar N. 1Post Graduate Student, 2Professor,3Asst.

Professor Dept. of Mechanical Engineering, Dayananda Sagar College of Engineering, Bangalore. Based on finite element analyses, the present paper also presents a weight comparison between a composite fuselage and an aluminium alloy comparison assess the weight reduction obtained with the use of composite materials for designing the fuselage.


Finite Element Analysis of a Fuselage Crack

II. Finite Element Analysis The aforementioned global FE model of composite panels, metallic fittings, mechanical fasteners, and the COLTS test fixture developed to support the design effort through linear structural strength and stability analyses was adapted to perform nonlinear analyses.

Finite element analysis of a fuselage
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