Several new and updated advanced models are coming with AFGROW version 5.5. This presentation will demonstrate the improved accuracy and range of updated models, as well as preview of the improvements to the AFGROW's advanced model interface.

For the past few decades T-38 has been using legacy spectra to analyze damage tolerant analysis, durability and create maintenance schedules. Periodically the T-38 SPO evaluates the curriculums of the currently flying aircraft to see if changes need to be implemented . The individual aircraft tracking program (IATP) on each aircraft is used to create spectra from current flying conditions. This is a presentation on the study of the current usages used for analysis compared to proposed new usages, the process of this study will be shown. The exceedance data between the usages will be shown which will flow into the test data, the SOLR correlation and the damage tolerance life changes. Southwest Research Institute has developed recommendations to the Air Force from the results of this study.

The B-52 ASIP team is pushing to start using AFGROW for their fleetwide DTA support, as the OEM legacy crack growth analysis software is no longer being supported.

Work is already underway to facilitate the transition process, including the development and use of an A-10 based Excel COM file to run fatigue crack growth analyses in AFGROW for many B-52 details. Comparisons and sensitivity studies are being performed to ensure a smooth, sensible transition occurs and that the models, parameters, and settings used in AFGROW provide the best fit for B-52 ASIP needs.

This work is part of a larger effort being supported by the A-10 ASIP team and contractors, and includes other tasks such as material testing, spectrum testing, residual stress measurement, spectra development verification, detailed damage tolerance analysis incorporating residual stresses, and organic analysis capability development.

A study was performed to evaluate the influence of the Newman-Raju finite width correction and the Harter finite width correction on A-10 critical locations. A comparison of the total fatigue life influence is provided as well as additional insights and observations, with validation to finite element results. The Harter correction was developed to correct for observed unconservatism in some scenarios, so as expected, predicted fatigue life was reduced for a number of critical locations. However, a large number of locations saw an increase in fatigue life as well, though the percent increase was not as significant for many locations as those that had a reduced life.

The A-10 ASIP team has performed multiple studies to help improve their DTA ground rules. In a 2021 AFGROW workshop Mr. Luciano Smith presented a summary of the studies the A-10 team investigated. Additionally, in a 2022 AFGROW workshop Mr. Kaylon Anderson presented the approaches for correlating SOLR values. The A-10 ASIP team updated crack growth rate fits. Resulting in a re-correlation of SOLR values. The updates to the A-10 DTA ground rules and SOLR values required an update to A-10 Damage Tolerance Analysis report. This presentation summarizes the results of the A-10 DTA update, due to the changes of DTA ground rules and SOLR values, specifically the Enhance Wing Assembly (EWA).

As aircraft require sustainment beyond their original service life, sound structural repairs, that are also fatigue and damage tolerant, are common. This presentation will show the process taken utilizing design of experiments to evaluate and attempt to optimize common repair doublers with typical A-10 structural components.

It is well known that as the load on the fastener increases, the fatigue life of the joint decreases. Therefore, repairs are designed so that the applied load will gradually develop; thus, minimizing the load at the first fastener row. Doing so generally results in a stepped thickness of the repair, along with other geometric features, that will reduce the stiffness at the first rows of fasteners. This approach leads to fatigue and damage tolerant friendly repairs.

Stiffness of the repair increases the load at the first fastener; therefore, it is desired to understand if general rules of thumb could be developed for how thick the repair at the first fastener row must be before a second repair needs to be added to step in the load more gradually. It is also desired to understand how sensitive increasing thickness is with respect to the resulting fatigue life at the joint and the corresponding durability & damage tolerance life.

This presentation will show numerous iterations of experimental coupon geometries. Varying thicknesses were analyzed while changing other parameters such as edge margin, load bearing transfer, and first/last fastener orientations. Conclusions, lessons learned and trends from these investigations will be discussed. This work is anticipated to serve as a catalyst to develop guidelines for efficient and confident sustainment doubler-style structural repairs.

Information on the latest research and development efforts and plans beyond AFGROW Release 5.5

Recent fatigue crack growth experiment results call into question established stress intensity factor (SIF) solutions for SEN(T) coupon geometries. This has motivated accuracy assessment and validation of legacy SIF solutions that are commonly associated with aircraft structural analyses. To that end, nonlinear finite element analysis has been observed to significantly improve SIF accuracy by correcting engineering assumptions made during the legacy formulation. Further, modern machine learning approaches to the generation of handbook SIF solutions has proven effective and efficient but requires diligence in generating training data along with selection of appropriate methods and hyperparameters. These potentially negative effects on SIF accuracy must be developed and documented in a best-practices guideline, as the USAF damage tolerance methodology relies on SIF solutions for analysis of aircraft structure and in prescribing inspection intervals. We present recent findings on the effect of nonlinear geometry in SIF computations and recent machine learning attempts at SIF solution development.

Design requirements for aerospace fracture-critical components typically include minimum (limiting) fatigue life and damage tolerance. For the first requirement, estimation of minimum fatigue life is based on statistical extrapolations of extensive S-N fatigue databases. Air Force Research Laboratory (AFRL) has demonstrated the applicability of a mechanism-based approach for predicting the limiting fatigue life of several alloys, without using the fatigue data. The essential inputs for this approach are the microstructural-scale (small) crack growth and life-limiting microstructural features. The application of this approach for predicting the minimum fatigue life of wrought and additively-manufactured (AM) Ti-6Al-4V will be discussed. For the second requirement, damage tolerance based life management of USAF’s critical structural components requires the knowledge of Equivalent Initial Damage Size (EIDS) distribution. Additively-manufactured metals are under consideration for multiple USAF applications, but EIDS data are not readily available in the open literature for AM metals. This presentation will discuss the results from a AFRL study to estimate the EIDS distributions in AM Ti-6Al-4V under both constant amplitude and simulated spectrum fatigue loading.

The A-10 System Program Office (SPO) engineering team is continually working to modernize their analysis methods for fleet management. One aspect of this modernization effort is to account for the benefit of residual stress at a cold worked hole during fatigue life predictions. Currently this benefit is captured using 3D finite element solutions for stress intensity—utilizing StressCheck within the framework of the multi-point crack growth program known as BAMpF. As part of validating these BAMpF models, the A-10 SPO compares baseline AFGROW and BAMpF models (no residual stress) to ensure that the solutions between models match, or are correlated in a predictable way. A recent observation during analysis noted that non-bearing loaded models usually compared well, but bearing loaded models had larger differences than expected. A study was conducted that compared solutions between AFGROW and BAMpF across a variety of plate widths and hole offsets to better understand where AFGROW and BAMpF bearing loaded models align or diverge. Additional investigations were made to propose underlying reasons for any observed divergence in solutions.

In the mid-1970’s, Paris and others developed a load-reduction test method to generate ΔK-rate data at very low crack-growth rates (less than 1e-8 m/cycle) to establish a ΔKth value at 1e-10 m/cycle for various stress ratios (R = P_min/P_max). Hudak et al (1978) developed the load-reduction test procedures for the ASTM Standard E-647. In 1984, the current load-reduction test method was approved by ASTM Committee E-24 and have been maintained by Committee E-08.

The objective of the current presentation is to review the history of load-reduction test methods. Ohji et al (1978) and Minikawa and McEvily (1981) experimentally measured a rise in the crack-opening loads using similar load-reduction test methods. The reasons for the rise in crack-opening loads were not known, but could have been caused by plasticity-, debris- and/or roughness-induced crack-closure mechanisms. Newman (1983), using a crack-closure model, numerically calculated a rise in the crack opening loads using the ASTM load-reduction test procedures. The rise in crack-opening loads were shown to be caused by residual plastic deformations from remote closure. Thus, measurements and calculations have demonstrated that the load-reduction test method causes a load-history effect and does not generate steady-state constant amplitude ΔK-rate data.

Pippan et al (1987, 1994) and Newman (1998-present) developed the compression pre-cracking (CP) test methods to generate load-history free ΔK-rate test data for a wide range of materials. However, these methods have not yet been adopted by the current fatigue-crack-growth testing community and ASTM has not included the CP test methods in ASTM E-647 nor developed a new standard.

Handbook stress intensity factor solutions (K_I) are commonly used for assessing fatigue crack growth rate (FCGR) for a metallic material of interest. Usage of the handbook solutions introduces a source of error since the models used to generate these closed-form solutions do not exactly represent the specimen geometry and the constraints imposed by the testing rig.

In this presentation, single edge notch (SEN) handbook K_I solutions are compared to 3D FEA-based solutions for a K_b specimen. Depending on the boundary conditions imposed to the 3D model, the K_I solution difference is quantified using the handbook solution as a reference. A comparison between a fatigue crack growth rate assessment using conditions more representative of test conditions and, similar to the model used in the closed-form development is provided.

The A-10 System Program Office (SPO) engineering team performed a failure assessment and damage tolerance analysis (DTA) of the hinge fitting for the nose landing gear (NLG) aft door. This was done in response to an NLG door departure in flight during a routine training sortie operating inside the normal operational envelope of the aircraft. An analysis team was tasked to determine the root cause of failure and subsequent inspection requirements for the hinge fitting. The presentation demonstrates how organic Air Force capabilities are leveraged to perform failure analyses, finite element modeling (FEM), and multi-point crack growth analyses to validate results and help define inspection criteria for fleet management. The fracture surface revealed multiple crack origins and heavy corrosion which obscured striation data in the region believed to be the primary crack. The materials lab at Hill AFB was able to perform striation counting at multiple points along the secondary fracture surface to correlate with analysis. Finite element analysis (FEA) was used to model stress fields in the part at the crack location(s), leveraging legacy stress analysis reports to help determine appropriate loads. Because of the complicated geometry and loading of the hinge, a custom multi-point fracture mechanics model was utilized to correlate the crack growth rates in the analysis to the available measured rates on the failed hinge. Then a second multi-point fracture mechanics model was ultimately utilized to generate crack growth curves from a crack originating at the higher-stress primary crack location. The resulting crack growth curves, along with the applicable detectable flaw size, may be used to manage fleet inspection requirements thus reducing the risk of future in-flight failures. Additionally, this evaluation informs the SPO as to how service usage might align or misalign with legacy design assumptions.