FASTRAN is an advanced life-prediction code based on the physics of the growth of micro- and macro-cracks in metallic materials. The code was developed over the past 40 years, at NASA Langley Research Center and Mississippi State University, and it is used worldwide in numerous aircraft companies, research organizations, and other industries.

The basic concept is driven by the experimental observation that cracks close during crack propagation. The model simulates the closure process and predicts the rate of crack propagation under arbitrary load histories. The code uses fatigue-crack growth rate data generated under constant-amplitude loading. The code is able to predict the fatigue life of metallic materials from micro-structural features, such as inclusion-particles or voids. The fatigue-life methodology is based on micro-crack growth and has been substantiated by experimental observations made on a number of engineering materials over the past two decades. This is an alternative to the traditional S-N (stress- or strain-life) approaches. The code is able to predict fatigue-crack growth under variable-amplitude and spectrum load histories, such as those that occur in aircraft structures. The code has what is termed “Rainflow on the fly” methodology, which does not require rainflow counting on variable-amplitude load histories.

The code has application to fatigue-life analyses used by the U. S. Navy and the damage-tolerance (crack growth) analyses used by the U. S. Air Force. The FASTRAN model may be incorporated into the NASGRO life-prediction code, and the model is currently an option in the AFGROW (life-prediction code).

During the past 30 years, the fatigue-crack-growth rate data generated from the ASTM load-shedding method has produced inappropriately high fatigue-crack-growth thresholds (Kth) that have been affected by load-history effects. In the past 15 years, compression precracking methods have been developed to generate fatigue-crack-growth rate data in the so-called threshold region with minimal load-history effects. The new method has been demonstrated on a wide class of materials from aluminum alloys, titanium alloys, steels and superalloys. These new methods are currently being incorporated into the revised ASTM fatigue-crack-growth standard (E-647).

It is known that environmental fatigue crack propagation depends sensitively on stress intensity range as well as loading parameters such as load ratio (R) and frequency. For example, the range of loading frequencies for effective molybdate inhibition of corrosion fatigue of aluminum alloy (AA) 7075 is dependent on the environmental conditions and R. However, the full combined effects of complex airframe load-time spectra on corrosion fatigue is not yet characterized or understood. In particular, (a) the relationship between the presence of an atmospheric environment (rather than full immersion) and loading history has not been examined in detail, and (b) inhibition of this environmental effect has not been investigated.

Additionally, current structural integrity prognosis programs do not explicitly account for the synergistic effect between environment and loading parameters. Specifically, how does the crack growth rate change as a function of cold/dry and hot/moist environments with and without inhibitors coupled with variable amplitude loading where maximum stress, frequency, and stress ratio vary in time. Tools such as AFGROW can effectively model airframe spectrum loading, but the necessary data for environment- spectrum interaction have not been available until recently and will be augmented in this effort. A proper characterization and understanding of the spectrum-environment interaction is critical for the next generation of prognosis programs.

In performing damage tolerance analysis (DTA) for aircraft structures, there are instances in which it is necessary to determine not only the life of a crack from a hole to the edge of a part, but then the growth of a secondary crack on the opposite side of the hole to failure. These types of fracture mechanics models have recently been added to the crack growth software we use for DTA. Southwest Research Institute and the United States Air Force performed testing designed to simulate continuing damage scenarios in which a crack from a hole has failed the primary ligament and a secondary crack is growing from the opposite side of the hole. Test results were then compared analytically to stress intensity values calculated using AFGROW, NASGRO, DSTO-developed equations, and finite element analysis. Because of the nature of the cracking, there is significant in-plane bending, and the different methods available to model the boundary conditions played a large role in the accuracy of the results. When boundary conditions were applied in a manner that closely modeled the testing, stress intensities matched well except near the transition from corner crack to through crack. Agreement with and deviations from test results and the reasons for them will be presented.

Gulfstream' current damage tolerance analysis tools are no longer supported by developers. Afgrow was selected to be the new DTA tool. The benefits show increases in efficiency, accuracy, and consistency in running DTA analyses, combined with an increase in efficiency in generating report-ready DT results. The 2 year review shows the yearly and expected future savings based on average usage within Gulfstream. Gulfstream analyses are supported by historically proven test evidence, correlated to in-house solutions. The plug-in capabilities of Afgrow have been rigorously tested and pushed to their limits. All in-house solutions have been successfully migrated to Afgrow. New COM functions have been added upon request by Gulfstream to increase the functionality of the plug-in module in Afgrow.

An example for continuing damage is used to show the progress of the current development. This example not only shows some of the work-arounds currently needed to run an analysis, but also shows the improvements possible for further development of the plug-in functionality.

A brief introduction into the standardized output based on the XML processing of an analysis run is shown, before concluding with some Gulfstream requests for future developments.

With funding and support from AFRL, Hill Engineering and A-10 ASIP are in the process of developing and implementing a structural modification on the A-10 that uses laser shock peening to enhance fatigue performance. The presentation provides an overview of the proposed modification, and describes the design and optimization of the surface treatment process. Surface treatment design uses computational tools that predict: 1) residual stress from surface treatments, and 2) the effects of residual stress on fatigue performance. Predicted behavior is validated against test data from experiments in element and full-scale articles. The predictive capabilities, and the validating test data, provide for estimates of benefit, including potential cost savings, life extension, or increases in time between structural inspections.

In order to be able to take the benefits of the cold working (CW) life improvement factor (LIF) for components and fleet application, more realistic LIF values, that reflect in-service experience and conditions, need to be determined through testing and modeling. Due to various variabilities, the lab test determined CW LIFs are deemed too high compared to those of full scale tests or in-service situations. The National Research Council of Canada (NRC) is improving methods and tools to simulate hole cold expansion and estimate its effects on fatigue life. The main objective of this project is to develop practical life improvement factors, representative of in-service experience and taking into account the variability observed in the field. In this presentation, 3D finite element simulation results will be compared with the “Residual Stress Database for Coldworked Holes”, measured using contour methods, under several USAF sponsored SBIR and RIF (Rapid Innovation Fund) projects. Key results such as the residual stresses at the entry and exit faces, the residual hole expansion, and the out-of-plane deformation, will be discussed. The preliminary development of an automated StressCheck/AFGROW-based tool for calculating the stress intensity factors resulting from the simulated cold expansion residual stresses will also be presented, and compared with some results from the previous SBIR and RIF projects.

The spectrum filtering capability in AFGROW was added to allow users to modify the spectrum tension and compression loads as a function of crack length. The primary purpose of this was to provide a means to incorporate an existing method to account for the effect of interference fit fasteners. After modeling several interference cases with StressCheck ®, this method did not appear to be the best approach. An effort is currently being conducted to develop K-solutions for cracks at holes with interference fit fasteners. This presentation will review the results of this effort for the axial loading case, and discuss plans for further development and implementation in AFGROW.

LexTech has expanded its crack growth rate database for use in AFGROW. There are now over 30 additional materials available that have been curve fit using the Harter-T method. This presentation will explain how to use these data in AFGROW. Comparison plots of the data and curve fits will be released on the AFGROW website in PDF format. The curve fits will also be released in the tabular look-up format. These files will contain the property values specific to the material and will be ready for use in the AFGROW tabular look-up dialog. This presentation will also give the users a chance to give suggestions on how they would like the data to be presented in AFGROW in later versions.

Depending on the usage of the data, there are many reasons why automated K-decreasing methods might be used in developing crack growth data sets. One reason might be to obtain limited crack growth data during the pre-crack phase of a toughness test, where the parameter, C, could be carefully chosen to validate the test material crack growth response to an expected behavior from previous testing while ending at an acceptable stress intensity prior to toughness testing. Another reason might be to obtain very low delta K values either to support the development of a long crack threshold value for compressive residual stress consideration or simply as a parameter for a specific crack growth model. Alternatively, there may be the need to perform an independent check on the validity of crack growth rate data generated by “non-standard” testing methods. This presentation will explore various motivations and the associated processes used to achieve the desired results while striving to preserve the integrity of the data through application of consistent test procedures.

On the RC-135s, repairs are typically substantiated by following the established standard practices in the TOs or by performing an equivalent strength analysis. In some cases, the standard repair is not preferred or possible due to schedule, cost, part availability, or adjacent existing repairs/modifications. When the damage occurs on the fuselage, we are able to make a few simplifying assumptions that allow us to perform a rough damage tolerance analysis to determine if a crack in the existing damage can remain structurally acceptable until the next depot input when it can be inspected and reevaluated. An example analysis will be presented highlighting the use of AFGROW to substantiate deferred maintenance.

Both A-10 and T-38 ASIP use the checkbox in AFGROW for “Keep a/c constant” for corner cracks. This approach typically does not allow the analytical crack shapes to match crack shapes observed through test. This presentation will discuss a test and analysis program being performed by SwRI, which will investigate the use of L-S material properties for the ‘a’ crack tip and L-T material properties for the ‘c’ crack tip, to identify if that approach will allow a/c to vary naturally in AFGROW and more closely match the corner crack shapes seen in test.

Fatigue crack growth analysis (Damage Tolerance) has been the predominant design tool for aircrafts around the world; but for aging aircraft, it is necessary to evaluate a field event such as a cracked structure to ascertain the remaining life considering variations in loading, material, and geometry during the airplane life. Therefore, a probabilistic damage tolerance analysis (PDTA) program funded by the Federal Aviation Administration (FAA) has been developed such that Government and OEM engineers can conduct a risk assessment of aircraft structural issues in support of policy decisions. A PDTA approach also provides a mechanism whereby inspection and maintenance operations can be included into the simulation, thus providing engineers the opportunity to assess the benefits of different maintenance actions.

A comprehensive probabilistic damage tolerance method requires a combination of deterministic crack growth, probabilistic methods, random variable modeling, and inspection and repair methods to provide a cumulative probability-of-failure and single-flight-probability-of-failure with and without inspection.

This presentation will describe the probabilistic methodology to be utilized in a computer software program (SMART|DT) that performs risk assessment of aircraft fleets allowing an interface with three different crack growth codes (NASGRO®, AFGROW®, and FASTRAN®) as the crack growth engine. The methodology can assess a large range of random variables and calculate the extreme value distribution of maximum load per flight from an internally generated spectrum.

This presentation will focus on the SMART-AFGROW® interphase implementation to solve different risk assessment cases.

Accurate quantification of crack tip stress intensity values is paramount in the analysis of damage tolerant structures. The present analytical investigation seeks to determine the stress intensity solutions for crack geometries outside the existing valid solution space and expand the analyst’s ability to capture representative crack growth behavior. The focus of this investigation is to calculate the stress intensity factors of single quarter-elliptical corner cracks emanating from centrally located holes in finite width plates under various loading conditions (remote tension, bending, and pin loading). Many of the available finite width corrections are singled valued and universally applied to all locations along the crack front. Early investigations into the validity of this application indicated that this correction procedure produces stress intensity values +/- 30% from new solutions. The crack depth to length ratio and depth to thickness ratio can also significantly influence the accuracy of historical finite width solutions and corrections. The analytical investigation utilizes the three dimensional virtual crack closure technique and well- structured, completely hexahedral, element meshes. Stress intensity values are generated for a wide range of ratios for crack depth to crack length, crack depth to sheet thickness, hole radius to sheet thickness, and sheet width to hole diameter. This effort is being executed under a DoD Technical Corrosion Collaboration program.

This talk will capture the improvements of BAMF over that last year to include: detailing the changes within the crack growth routines, model modifications, as well as the program improvements. There will be a discussion on the new feature of multi-point multi-crack BAMF with examples of multiple cracks growing in residual stress fields as well as continuing damage models that exhibit crack interaction.

In life prediction of structural components containing an initial flaw it is particularly important to determine stress intensity factors (SIFs) to a high degree of accuracy. In the particular case of cracks emanating from cold worked holes a very important factor affecting the prediction of fatigue crack propagation is the aleatory variation of experimentally-derived residual stress distributions.

This presentation will discuss a model of fatigue crack propagation and present results of simulations when uncertainties in the input residual stress distribution are considered. Life predictions are compared with experimental fatigue life data for aluminum alloys.

We examine the effect different methods for cold working countersink holes has on life predictions using BAMF. At least three methods are used to engineer residual stress in countersink holes:

1. cold work a straight shank hole then drill the countersink
2. drill the countersink, but cold work only the straight shank and
3. drill the countersink, then use the patented FTI nosecap (CsCx(TM)) to engineer residual stress both in the straight shank and countersink regions.

While the highest stress due to the applied far field tension is at the corner of the countersink and the straight shank, thus is the likely crack formation location, residual stress will generally decrease the effect stress the crack experiences, such that the crack formation location could move from the expected location, depending on the residual stress induced.

Information on the latest research and development efforts and plans beyound AFGROW Release 5.3.