John G. Bakuckas

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.

Boundary Correction Factors for Elliptical Surface Cracks Emanating from Countersunk Rivet Holes

To predict the crack growth and residual strength of riveted joints subjected to widespread fatigue damage, accurate stress and fracture analyses of corner and surface cracks at a rivet hole are needed. The results presented focus on the calculation of stress-intensity factor (SIF) solutions for cracks at countersunk rivet holes for tension, bending, and wedge load conditions. A wide range of configuration parameters were varied, including the crack size, crack shape, crack location, and length of the straight shank hole. A finite-element-based global-intermediatelocal hierarchical approach was used. The results are expressed as boundary correction factors (BCFs), which are nondimensional representations of the SIF. The BCFs were determined along the crack front in terms of the physical angle, which was measured from the inner surface of the plate to a point on the hole boundary or on the outer surface of the plate. In general, the values of BCFs increased along the crack front, moving from the inner surface of the plate toward the hole boundary or the outer surface. The values of the BCFs were highest for the crack fronts closest to the hole boundary. The trends in the solutions were the same for the three loading conditions.

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.

On the Applicability of Acoustic Emission for Monitoring Damage Progression in Metal Matrix Composites

Damage initiation and accumulation has been monitored through acoustic emission (AE) in several center-notched unidirectional metal matrix composites (MMC) and multidirectional boron/aluminum (B/Al) laminates subjected touniaxial quasi-static tensile loading. Emphasis has been placed on establishing the correlation between the actual damage initiation and growth and AE results. The correspondence between failure processes in the different material systems and the event intensities was investigated, and the results on the event amplitudes are reported. Effect of heat treatment on the failure process and on the associated emission in unidirectional and quasi-isotropic B/Al laminates has also been studied. Notch-tip damage growth was monitored in real-time through optical observations, and effective notch-tip damage extension was determined by applying the compliance matching procedure. Fracture surface morphology was examined via the scanning electron microscope. It is shown that the failure processes in the different material systems can be correlated with the AE event amplitudes. The results indicate that an excellent correlation can be established between the rate of damage growth and the AE results. Thus, it is concluded that AE technique can monitor damage formation and accumulation in MMC in real-time.

Nonlinear Analysis and Post-Test Correlation for a Curved PRSEUS Panel

The Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) concept, developed by The Boeing Company, has been extensively studied as part of the National Aeronautics and Space Administrations (NASA s) Environmentally Responsible Aviation (ERA) Program. The PRSEUS concept provides a light-weight alternative to aluminum or traditional composite design concepts and is applicable to traditional-shaped fuselage barrels and wings, as well as advanced configurations such as a hybrid wing body or truss braced wings. Therefore, NASA, the Federal Aviation Administration (FAA) and The Boeing Company partnered in an effort to assess the performance and damage arrestments capabilities of a PRSEUS concept panel using a full-scale curved panel in the FAA Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility. Testing was conducted in the FASTER facility by subjecting the panel to axial tension loads applied to the ends of the panel, internal pressure, and combined axial tension and internal pressure loadings. Additionally, reactive hoop loads were applied to the skin and frames of the panel along its edges. The panel successfully supported the required design loads in the pristine condition and with a severed stiffener. The panel also demonstrated that the PRSEUS concept could arrest the progression of damage including crack arrestment and crack turning. This paper presents the nonlinear post-test analysis and correlation with test results for the curved PRSEUS panel. It is shown that nonlinear analysis can accurately calculate the behavior of a PRSEUS panel under tension, pressure and combined loading conditions.

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.

Damage accumulation in titanium matrix composites under generic hypersonic vehicle flight simulation and sustained loads

Titanium matrix composites (TMC), such as Ti-15V-3Cr-3Al-3Sn (Ti-15-3) reinforced with continuous silicon-carbide fibers (SCS-6), are being evaluated for use in hypersonic vehicles and advanced gas turbine engines where high strength-to-weight and high stiffness-to-weight ratios at elevated temperatures are critical. Such applications expose the composite to mechanical fatigue loading as well as thermally induced cycles. The damage accumulation behavior of a [0/90]2s laminate made of Ti-15V-3Cr-3Al-3Sn (Ti-15-3) reinforced with continuous silicon-carbide fibers (SCS-6) subjected to a simulated generic hypersonic flight profile, portions of the flight profile, and sustained loads was evaluated experimentally. Portions of the flight profile were used separately to isolate combinations of load and time at temperature that influenced the fatigue behavior of the composite. Sustained load tests were also conducted and the results were compared with the fatigue results under the flight profile and its portions. The test results indicated that the fatigue strength of this materials system is considerably reduced by a combination of load and time at temperature.

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.

Detecting damage in full-scale honeycomb sandwich composite curved fuselage panels through frequency response

Preliminary tests were conducted using frequency response (FR) characteristics to determine damage initiation and growth in a honeycomb sandwich graphite/epoxy curved panel. This investigation was part of a more general study investigating the damage tolerance characteristics of several such panels subjected to quasi-static internal pressurization combined with hoop and axial loading. The panels were tested at the Full-Scale Aircraft Structural Test Evaluation and Research (FASTER) facility located at the Federal Aviation Administration William J. Hughes Technical Center in Atlantic City, NJ. The overall program objective was to investigate the damage tolerance characteristics of full-scale composite curved aircraft fuselage panels and the evolution of damage under quasi-static loading up to failure. This paper focuses on one aspect of this comprehensive investigation: the effect of state-of-damage on the characteristics of the frequency response of the subject material. The results presented herein show that recording the frequency response could be used for real-time monitoring of damage growth and in determining damage severity in full-scale composites fuselage aircraft structures.