Tein-Min Tan

Use of Statical Indentation Laws in the Impact Analysis of Laminated Composite Plates

The low-velocity impact response of graphite/epoxy laminates was investigated theoretically and experimentally. A nine-node isoparametric plate finite element in conjunction with an empirical contact law was used for the theoretical investigation. Theoretical results are in good agreement with strain-gage experimental data. The results of the investigation indicate that the present theoretical procedure describes the impact response of laminate for low-impact velocities.

Fatigue Testing of a Stiffened Lap Joint Curved Fuselage Structure

In April 1988, Aloha Airlines flight 243 experienced an explosive midair decompression that resulted in the separation of an 18-foot section of the fuselage crown of the Boeing 737 airplane. Investigations revealed that the linkup of small cracks emanating from multiple rivet holes in a debonded lap joint contributed to the catastrophic failure. This cracking scenario, known as multiple-site damage, is one of two sources of widespread fatigue damage; a type of structural degradation characterized by the simultaneous presence of fatigue cracks at multiple structural elements that are of sufficient size and density whereby the structure will no longer meet its damage tolerance requirement This study, sponsored by the National Aging Aircraft Research Program initiated by the Federal Aviation Administration in response to the Aloha accident, investigates multiple-site damage initiation and growth behavior in a pristine narrow-body fuselage panel. The test panel, a curved 6 x 10 ft stiffened structure containing six frames, seven stringers, and a longitudinal lap joint, was tested at the Federal Aviation Administration Full-Scale Aircraft Structural Test Evaluation and Research facility. The panel was subjected to a fatigue test with constant-amplitude cyclic loading, simulating the major modes of load associated with fuselage pressurization. Nondestructive inspections were conducted during the fatigue test to detect and monitor crack formation and growth. Multiple-site damage cracks were visually detected after about 80% of the fatigue life. Cracks developed and linked in the upper rivet row of the lap joint in the outer skin layer and formed a long fatigue crack before the termination of the fatigue test A residual strength test was then conducted by subjecting the panel to quasi-static loads until catastrophic failure. Fractographic examinations were conducted to reconstruct crack growth history. Preliminary results show multiple crack origins and significant subsurface crack growth.

Computational methodology to predict damage growth in unidirectional composites—I. Theoretical formulation and numerical implementation

The theoretical formulation and numerical implementation of a computational methodology for predicting both the initiation and growth of damage in a unidirectional composite monolayer is presented. The methodology has been implemented into a finite element program to form the Micromechanics Analysis and Damage Growth In Composites (MADGIC) code. A node splitting and nodal force relaxation algorithm that is capable of generating new crack surfaces has been incorporated to simulate damage initiation and growth. One of the unique features of this code is that the instantaneous direction of damage progression is dictated by the local mechanics and failure criteria. Thus, the crack path need not be preselected. Common modes of damage that take place in composites, including fiber breakage, matrix cracking and fiber-matrix debonding, are simulated using the node splitting mechanisms in conjunction with mechanistic failure criteria. An incremental elastic-plastic algorithm with J2 flow theory and isotropic hardening has also been incorporated to account for matrix plastic deformation when analyzing damage growth in metal matrix composites. In order to efficiently model standard laboratory size composite specimens, a hybrid micromechanical-anisotropic continuum model has been used consisting of a heterogeneous region enclosing the micromechanical damage processing zone, and an outer homogeneous region to which the far-field load is applied.

Engineering Design of Tough Ceramic Matrix Composites for Turbine Components

This paper provides a review of the material design concepts for the toughening of ceramic matrix composites by three-dimensional fiber architecture. To establish a communication link between the structural and the materials engineers, an integrated design methodology is presented with an example. Through a Fabric Geometry Model (FGM), the contribution of three-dimensional fiber architecture is translated into a stiffness matrix for finite element structural analysis. With the feedback from the structural analysis, this design methodology provides an effective means to screen reinforcement materials systems for three-dimensional fabric-reinforced composite components

A modified finite element-transfer matrix for control design of space structures

The Finite Element-Transfer Matrix (FETM) method was developed for reducing the computational efforts involved in structural analysis. While being widely used by structural analysts, this method does, however, have certain limitations, particularly when used for the control design of large flexible structures. In this paper, a new formulation based on the FETM method is presented. The new method effectively overcomes the limitations in the original FETM method, and also allows an easy construction of reduced models that are tailored for the control design. Other advantages of this new method include the ability to extract open loop frequencies and mode shapes with less computation, and simplification of the design procedures for output feedback, constrained compensation, and decentralized control. The development of this new method and the procedures for generating reduced models using this method are described in detail and the role of the reduced models in control design is discussed through an illustrative example.

Control of Flexible Structures by Finite Element-Transfer Matrix Approach

Abstract For a flexible structure to meet its mission objectives successfully, a close interaction between structural analysis and control group is needed at both the modelling and control design phases. For such an interaction to be beneficial, an integration of suitable methods from both the groups is essential. The finite element-transfer matrix (FETM) approach and the method of decentralized control using optimal control theory are ideally suited for this integration: the FETM method has been developed by structural analysts to overcome the ‘dimension-problem’, which problem is also encountered by control engineers; the decentralized control strategy is being adopted by control engineers to control large flexible structures. This paper presents the development of this method and demonstrates its applicability to the control of distributed parameter systems. Specifically, as an illustration, the control of a simply-supported beam is investigated

Reduced modeling of flexible structures for decentralized control

Based upon the modified finite element - transfer matrix method [1], this paper presents a technique for reduced modeling of flexible structures for decentralized control. The modeling decisions are carried out at (finite-) elemental level, and are dictated by control objectives. A simply supported beam with two sets of actuators and sensors (linear force actuator and linear position and velocity sensors) is considered for illustration. In this case, it is conjectured that the decentrally controlled closed loop system is guaranteed to be at least marginally stable.

Full-scale testing of fuselage panels

A unique state-of-the-art facility to assess the structural integrity of aircraft fuselage structure was established at the Federal Aviation Administration (FAA) William J. Hughes Technical Center. The Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility is capable of testing full-scale fuselage panel specimens under conditions representative of those seen by an aircraft in actual operation. The test fixture features a novel adaptation of mechanical, fluid, and electronic components and is capable of applying pressurization, longitudinal, hoop, frame, and shear loads to a fuselage panel. A high-precision, Remote Controlled Crack Monitoring (RCCM) system was developed to inspect and record crack initiation and progression over the entire fuselage panel test surface. A detailed description of the FASTER facility along with representative results from a variety of experimental test programs will be presented.

Computational methodology to predict damage growth in unidirectional composites. II : case studies

Abstract A computational methodology was used to simulate the damage growth processes in center notched unidirectional Boron-Epoxy, Boron-Aluminum, and Silicon Carbide-Titanium composites subjected to quasi-static tensile loading. This methodology uses a specially developed finite element program, Micromechanics Analysis and Damage Growth In Composites (MADGIC) code and a hybrid micromechanical-anisotropic continuum model. The unique feature of the approach is that multiple modes of damage can be simulated simultaneously and the crack path need not be preselected. The direction and path of damage growth are dictated by the local mechanics in conjunction with the failure criteria. This paper reports on the applications of this computational methodology to several case studies. Predictions of the damage growth process in unidirectional aluminum and titanium matrix composites were correlated with experimental observations, which have shown to be highly dependent on the properties of the constituents. The computational simulations captured the salient features of the observed notch-dip damage evolution in each of the materials evaluated. The predictions agreed qualitatively quite well with the experimental observations of the failure process.