Engineering topics

Austin Pennington

Publications and source records attributed to Austin Pennington.

4 records

NASA NTRSDate not supplied

Progressive Damage Failure Analysis of Post-Buckled Composite Single-Stringer Panel with Teflon Inserts

The high strength-to-weight and stiffness-to-weight ratio materials, such as laminated composites, are advantageous for modern aircraft. Laminated composites with initial flaws are susceptible to delamination under buckling loads. PDA tools help enhance the industry’s understanding of the mechanisms for damage initiation and growth in composite structures while assisting in the design, analysis, and sustainment methods of these composite structures. The global-local modeling approach for the single-stringer post-buckled panel was evaluated through this effort, using Teflon inserts to simulate the defect of damage during manufacturing. This understanding is essential for designing the post-buckled structure, reducing weight while predicting damage initiation location, and addressing a potential design review for future aircraft repairs. In this work, the initial damage was captured with Teflon inserts as the starting configuration; and any reference to the damage initiation refers to any damage beyond the “initial unbonded region.” The effort aims to develop, evaluate, and enhance methods to predict damage initiation and progression and the failure of post-buckled hat-stiffened panels using multiple Abaqus FEA Virtual Crack Closure Technique (VCCT) definitions. Validation of the PDA using the VCCT material model was performed on a large single-stringer panel subjected to compressive loading. The compressive loading of the panel caused the skin to buckle before any damage began to occur locally. In addition, comparisons are made for critical aspects of the damage morphology, such as a growth pattern that included delamination from the skin-stiffener interface to the skin and ply interfaces. When compared against the experimental data produced through the NASA Advanced Composites Project (ACP), the present model captured damage migration from one surface to another, and model validations were within 4% of the experimental data.

PDA↗

NASA NTRSDate not supplied

FEA-Based Failure Analysis of Post-Buckled Composite Multi-Hat Stiffened Panel Under Static Compression

The aerospace industry uses composite multi-hat stiffened panels for skins because of their high strength-to-weight and stiffness-to-weight ratios. However, laminated composites with initial flaws are susceptible to delamination under buckling loads. The goal is to ensure damage initiation occurs past the buckled state, only stable damage progression at loads below the ultimate compressive load, and debonding failure past the ultimate compressive load. This approach reduces weight and addresses a potential design review for future aircraft repairs. The need exists to better understand damage initiation and growth mechanisms in composite structures while assisting in the design, analysis, and sustainment methods. Teflon inserts simulate damage defects during manufacturing, while barely visible impact helps capture human error during manufacturing. This work shows the interaction between the skin-stringer and ply-ply separations at the post-buckled state for both defect events using Abaqus Virtual Crack Closure Technique (VCCT). The initial configuration for the Teflon case was to leave the Teflon region unbonded, and the initial impact damage region approximation came from the experimental data and was considered the initial configuration. When compared against the experimental data produced through the NASA Advanced Composites Project (ACP), the present model captured damage transition from skin-hat to ply1-ply2. The model validations were within ~5% of the experimental data.

composite↗

NASA NTRSDate not supplied

Integration of Fatigue R-Curve Effects into VCCT for Durability Predictions, Part 1: Buckled Composite Single-Stringer Stiffened Panels

Composite structures have become popular in modern aircraft because they help reduce weight and increase durability. In addition, hat-stiffened panels provide the stability that the airframe skin needs. However, they can be subject to delamination in the post-buckling regime. Progressive damage analysis (PDA) methods can help predict interlaminar and intralaminar failure events. Many aircraft structures are subject to cyclic loading in the post-buckling regime. Hence, fatigue life prediction becomes essential for design and sustainment purposes. Under the NASA Advanced Composites Project (ACP), composite panels stiffened with single hat-stringers were subject to a cyclic loading sequence from a pre-buckling state to a post-buckling state. This test campaign aimed to provide damage initiation and growth data for three initial damage scenarios: nominally pristine, initial Teflon inserts representing manufacturing defects, and impact-induced damage. This work uses the Abaqus FEA Virtual Crack Closure Technique (VCCT), capabilities enhanced through a novel empirical method integrating fatigue R-curve effects (different from R-ratio) into the Paris Law via user-defined subroutine to simulate the fatigue response of the panels with single hat-stringers. The analysis predictions were within 5% of the test results using this novel method.

single hat↗

NASA NTRSDate not supplied

Integration of Fatigue R-Curve Effects into VCCT for Durability Predictions, Part 2: Buckled Composite Multi-Stringer Stiffened Panels

Composite structures with hat-stiffened panels have gained traction in modern aircraft because they help reduce weight, increase durability, and provide stability to airframe skins. For weight reduction, skins should be designed to operate in the post-buckling regime. This causes possible interlaminar and intralaminar failure. However, composite panels stiffened with multiple hat-stringers are subject to cyclic loading even in the post-buckling regime. Under the NASA Advanced Composites Project (ACP), composite panels with multiple hatstringers were subject to a cyclic loading sequence from a pre-buckling state to a post-buckling state. This effort aimed to predict interlaminar failure for cyclic compression loading with pre-existent impact-induced damage. This work uses the Abaqus FEA Virtual Crack Closure Technique (VCCT), capabilities enhanced through an empirical method to integrate fatigue R-curve effects (different from R-ratio) into the Paris Law via user-defined subroutine to simulate the fatigue response o

ACP↗