Cheryl A. Rose

Failure criteria for FRP laminates

A new set of six phenomenological failure criteria for fiber-reinforced polymer laminates denoted LaRC03 is described. These criteria can predict matrix and fiber failure accurately, without the curve-fitting parameters. For matrix failure under transverse compression, the angle of the fracture plane is solved by maximizing the Mohr-Coulomb effective stresses. A criterion for fiber kinking is obtained by calculating the fiber misalignment under load and applying the matrix failure criterion in the coordinate frame of the misalignment. Fracture mechanics models of matrix cracks are used to develop a criterion for matrix failure in tension and to calculate the associated in situ strengths. The LaRC03 criteria are applied to a few examples to predict failure load envelopes and to predict the failure mode for each region of the envelope. The analysis results are compared to the predictions using other available failure criteria and with experimental results.

Continuum Damage Mechanics Models for the Analysis of Progressive Failure in Open-Hole Tension Laminates

The performance of a state-of-the-art continuum damage mechanics model for interlaminar damage, coupled with a cohesive zone model for delamination is examined for failure prediction of quasi-isotropic open-hole tension laminates. Limitations of continuum representations of intra-ply damage and the effect of mesh orientation on the analysis predictions are discussed. It is shown that accurate prediction of matrix crack paths and stress redistribution after cracking requires a mesh aligned with the fiber orientation. Based on these results, an aligned mesh is proposed for analysis of the open-hole tension specimens consisting of different meshes within the individual plies, such that the element edges are aligned with the ply fiber direction. The modeling approach is assessed by comparison of analysis predictions to experimental data for specimen configurations in which failure is dominated by complex interactions between matrix cracks and delaminations. It is shown that the different failure mechanisms observed in the tests are well predicted. In addition, the modeling approach is demonstrated to predict proper trends in the effect of scaling on strength and failure mechanisms of quasi-isotropic open-hole tension laminates.

Compression Response of Fluted-Core Composite Panels

In recent years, fiber-reinforced composites have becomemore accepted for aerospace applications. For example, duringNASA’s recent efforts to develop new launch vehicles, compositematerials were considered and baselined for a number of structures, including dry barrel sections, which are primarily loaded in longitudinal compression. Because of mass and stiffness requirements, sandwich composites are often selected for these applications. However, there are a number of manufacturing and in-service concerns associated with traditional honeycomb-core sandwich composites that in certain instances may be alleviated through the use of other core materials or construction methods. A fluted core, which consists of integral angled webmembers with structural radius fillers spaced between laminate facesheets, is one such construction alternative. In this paper, two different fluted-core composite designs were considered: a subscale design and a full-scale design sized for a heavy-lift-launch-vehicle interstage. In particular, the longitudinal compression behavior of fluted-core composites was evaluated with experiments and finite-element analyses. Detailed branched-shell finite-element models were developed, and geometrically nonlinear analyses were conducted to predict both buckling andmaterial failures. Good agreement was obtained between test data and analysis predictions for both failure types. Though the local buckling events are not catastrophic, the resulting deformations contribute to material failures. Consequently, neither the local buckling behavior nor the material failure loads and modes can be predicted by either linear analyses or nonlinear smeared-shell analyses. Compression-after-impact performance of fluted-core composites was also investigated experimentally. Nondestructive inspection of the damage zones indicated that the detectable damage was limited to no more than one flute on either side of any given impact. More study is needed, but this may indicate that an inherent damagearrest capability of fluted core could provide benefits over traditional sandwich designs in certain weight-critical applications.

Effect of Buckling Modes on the Fatigue Life and Damage Tolerance of Stiffened Structures

The postbuckling response and the collapse of composite specimens with a co-cured hat stringer are investigated experimentally and numerically. These specimens are designed to evaluate the postbuckling response and the effect of an embedded defect on the collapse load and the mode of failure. Tests performed using controlled conditions and detailed instrumentation demonstrate that the damage tolerance, fatigue life, and collapse loads are closely tied with the mode of the postbuckling deformation, which can be different between two nominally identical specimens. Modes that tend to open skin/stringer defects are the most damaging to the structure. However, skin/stringer bond defects can also propagate under shearing modes. In the proposed paper, the effects of initial shape imperfections on the postbuckling modes and the interaction between different postbuckling deformations and the propagation of skin/stringer bond defects under quasi-static or fatigue loads will be examined.

Compression Behavior of Fluted-Core Composite Panels

In recent years, fiber-reinforced composites have become more accepted for aerospace applications. Specifically, during NASA s recent efforts to develop new launch vehicles, composite materials were considered and baselined for a number of structures. Because of mass and stiffness requirements, sandwich composites are often selected for many applications. However, there are a number of manufacturing and in-service concerns associated with traditional honeycomb-core sandwich composites that in certain instances may be alleviated through the use of other core materials or construction methods. Fluted-core, which consists of integral angled web members with structural radius fillers spaced between laminate face sheets, is one such construction alternative and is considered herein. Two different fluted-core designs were considered: a subscale design and a full-scale design sized for a heavy-lift-launch-vehicle interstage. In particular, axial compression of fluted-core composites was evaluated with experiments and finite-element analyses (FEA); axial compression is the primary loading condition in dry launch-vehicle barrel sections. Detailed finite-element models were developed to represent all components of the fluted-core construction, and geometrically nonlinear analyses were conducted to predict both buckling and material failures. Good agreement was obtained between test data and analyses, for both local buckling and ultimate material failure. Though the local buckling events are not catastrophic, the resulting deformations contribute to material failures. Consequently, an important observation is that the material failure loads and modes would not be captured by either linear analyses or nonlinear smeared-shell analyses. Compression-after-impact (CAI) performance of fluted core composites was also investigated by experimentally testing samples impacted with 6 ft.-lb. impact energies. It was found that such impacts reduced the ultimate load carrying capability by approximately 40% on the subscale test articles and by less than 20% on the full-scale test articles. Nondestructive inspection of the damage zones indicated that the detectable damage was limited to no more than one flute on either side of any given impact. More study is needed, but this may indicate that an inherent damage-arrest capability of fluted core could provide benefits over traditional sandwich designs in certain weight-critical applications.

Nonlinear Analysis and Scaling Laws for Noncircular Composite Structures Subjected to Combined Loads

Results from an analytical study of the response of a built-up, multi-cell noncircular composite structure subjected to combined internal pressure and mechanical loads are presented. Nondimensional parameters and scaling laws based on a first-order shear-deformation plate theory are derived for this noncircular composite structure. The scaling laws are used to design sub-scale structural models for predicting the structural response of a full-scale structure representative of a portion of a blended-wing-body transport aircraft. Because of the complexity of the full-scale structure, some of the similitude conditions are relaxed for the sub-scale structural models. Results from a systematic parametric study are used to determine the effects of relaxing selected similitude conditions on the sensitivity of the effectiveness of using the sub-scale structural model response characteristics for predicting the full-scale structure response characteristics.

The Nonlinear Response of Cracked Aluminum Shells Subjected to Combined Loads

Abstract The results of a numerical study of the nonlinear re-sponse of thin unstiffened aluminum cylindrical shellswith a longitudinal crack are presented. The shells areanalyzed with a nonlinear shell analysis code that accu-rately accounts for global and structural response phe-nomena. The effects of initial crack length on theprebuckling, buckling and postbuckling responses of atypical shell subjected to axial compression loads, andsubjected to combined internal pressure and axial com-pression loads are described. Both elastic and elastic-plastic analyses are conducted. Numerical results for afixed initial crack length indicate that the buckling loaddecreases as the crack length increases for a given pres-sure load, and that the buckling load increases as the in-ternal pressure load increases for a given crack length.Furthermore, results indicate that predictions from anelastic analysis for the initial buckling load of a crackedshell subjected to combined axial compression and inter-nal pressure loads can be unconservative. In addition,the effect of crack extension on the initial buckling loadis presented.IntroductionThe fail-safe design philosophy, when applied totransport aircraft fuselage structure, requires that thesestructures retain adequate structural integrity in the pres-ence of discrete-source damage or fatigue cracks. Onetype of damage frequently associated with the structuralintegrity of fuselage shell structures is a longitudinalcrack in the fuselage skin that is subjected to circumfer-ential stresses resulting from the internal pressure loads,and to axial stresses resulting from the vertical bendingand shearing of the fuselage that are induced by normalflight loads. The structural response of a transport fuse-lage structure with a crack is influenced by the localstress and displacement gradients near the crack and bythe internal load distribution in the shell. Local fuselageout-of-plane skin displacements near a crack can be large

STAGS Developments for Residual Strength Analysis Methods for Metallic Fuselage Structures

A summary of advances in the Structural Analysis of General Shells (STAGS) finite element code for the residual strength analysis of metallic fuselage structures, that were realized through collaboration between the structures group at NASA Langley, and Dr. Charles Rankin is presented. The majority of the advancements described were made in the 1990s under the NASA Airframe Structural Integrity Program (NASIP). Example results from studies that were conducted using the STAGS code to develop improved understanding of the nonlinear response of cracked fuselage structures subjected to combined loads are presented. An integrated residual strength analysis methodology for metallic structure that models crack growth to predict the effect of cracks on structural integrity is demonstrated

Jim Starnes' Contributions to Residual Strength Analysis Methods for Metallic Structures

A summary of advances in residual strength analyses methods for metallic structures that were realized under the leadership of Dr. James H. Starnes, Jr., is presented. The majority of research led by Dr. Starnes in this area was conducted in the 1990s under the NASA Airframe Structural Integrity Program (NASIP). Dr. Starnes, respectfully referred to herein as Jim, had a passion for studying complex response phenomena and dedicated a significant amount of research effort toward advancing damage tolerance and residual strength analysis methods for metallic structures. Jims efforts were focused on understanding damage propagation in built-up fuselage structure with widespread fatigue damage, with the goal of ensuring safety in the aging international commercial transport fleet. Jims major contributions in this research area were in identifying the effects of combined internal pressure and mechanical loads, and geometric nonlinearity, on the response of built-up structures with damage. Analytical and experimental technical results are presented to demonstrate the breadth and rigor of the research conducted in this technical area. Technical results presented herein are drawn exclusively from papers where Jim was a co-author.