Jonathan P. Moody

A Microstructure-Based Time-Dependent Crack Growth Model for Life and Reliability Prediction of Turbopropulsion Systems

The objective of this investigation was to develop an innovative methodology for life and reliability prediction of hot-section components in advanced turbopropulsion systems. A set of generic microstructure-based time-dependent crack growth (TDCG) models was developed and used to assess the sources of material variability due to microstructure and material parameters such as grain size, activation energy, and crack growth threshold for TDCG. A comparison of model predictions and experimental data obtained in air and in vacuum suggests that oxidation is responsible for higher crack growth rates at high temperatures, low frequencies, and long dwell times, but oxidation can also induce higher crack growth thresholds (ΔKth or Kth) under certain conditions. Using the enhanced risk analysis tool and material constants calibrated to IN 718 data, the effect of TDCG on the risk of fracture in turboengine components was demonstrated for a generic rotor design and a realistic mission profile using the DARWIN® probabilistic life-prediction code. The results of this investigation confirmed that TDCG and cycle-dependent crack growth in IN 718 can be treated by a simple summation of the crack increments over a mission. For the temperatures considered, TDCG in IN 718 can be considered as a K-controlled or a diffusion-controlled oxidation-induced degradation process. This methodology provides a pathway for evaluating microstructural effects on multiple damage modes in hot-section components.

Probabilistic Fretting Fatigue Assessment of Aircraft Engine Disks

Fretting fatigue is a random process that continues to be a major source of damage associated with the failure of aircraft gas turbine engine components. Fretting fatigue is dominated by the fatigue crack growth phase and is strongly dependent on the magnitude of the stress values in the contact region. These stress values often have the most influence on small cracks where traditional long-crack fracture mechanics may not apply. A number of random variables can be used to model the uncertainty associated with the fatigue crack growth process. However, these variables can often be reduced to a few primary random variables related to the size and location of the initial crack, variability associated with applied stress and crack growth life models, and uncertainty in the quality and frequency of nondeterministic inspections. In this paper, an approach is presented for estimating the risk reduction associated with the nondestructive inspection of aircraft engine components subjected to fretting fatigue. Contact stress values in the blade attachment region are estimated using a fine mesh finite element model coupled with a singular integral equation solver and combined with bulk stress values to obtain the total stress gradient at the edge of contact. This stress gradient is applied to the crack growth life prediction of a mode I fretting fatigue crack. A probabilistic model of the fretting process is formulated and calibrated using failure data from an existing engine fleet. The resulting calibrated model is used to quantify the influence of inspection on the probability of fracture of an actual military engine disk under real life loading conditions. The results can be applied to quantitative risk predictions of gas turbine engine components subjected to fretting fatigue.

Development of a Probabilistic Methodology for Predicting Hot Corrosion Fatigue Crack Growth Life of Gas Turbine Engine Disks

Advanced Ni-based gas turbine disks are expected to operate at higher service temperatures in aggressive environment for longer time durations. Exposures of Ni-based alloys to alkaline-metal salts and sulfur compounds at elevated temperatures can lead to hot corrosion fatigue crack growth in engine disks. Type II hot corrosion involves the formation and growth of corrosion pits in Ni-based alloys at a temperature range of 650°C to 750°C. Once formed, these corrosion pits can serve as stress concentration sites where fatigue cracks can initiate and propagate to failure under subsequent cyclic loading. In this paper, a probabilistic methodology is developed for predicting the corrosion fatigue crack growth life of gas turbine engine disks made from a powder-metallurgy Ni-based superalloy (ME3). Key features of the approach include (1) a pit growth model that describes the depth and width of corrosion pits as a function of exposure time, (2) a cycle-dependent crack growth model for treating fatigue, and (3) a time-dependent crack growth model for treating corrosion. This set of deterministic models is implemented into a probabilistic life-prediction code called DARWIN. Application of this approach is demonstrated for predicting corrosion fatigue crack growth life in a gas turbine disk based on ME3 properties from the literature. The results of this study are used to assess the conditions that control the transition of a corrosion pit to a fatigue crack, and to identify the pertinent material parameters influencing corrosion fatigue life and disk reliability.Copyright

Integration of Manufacturing Process Simulation with Probabilistic Damage Tolerance Analysis of Aircraft Engine Components

Interfaces between manufacturing process simulation software (DEFORM) and probabilistic damage tolerance analysis software (DARWIN) have been developed for bulk residual stresses and average grain size. These interfaces permit full-field results from manufacturing process simulations to be incorporated in predictions of fracture life and reliability. Approaches were presented for modeling the effects of location-specific bulk residual stress and average grain size on crack growth behavior. The interface and the proposed approaches were implemented in prototype software and used to perform demonstration examples for an idealized engine disk. The exercise demonstrates the practical potential for Integrated Computational Materials Engineering (ICME) that directly addresses component integrity.

An Integrated Software Tool for High Fidelity Probabilistic Assessments of Metallic Aero-Engine Components

Most current tools and methodologies to predict the life and reliability of fracture critical gas turbine engine components rely on stress intensity factor solutions that assume highly idealized component and crack geometries, and this can lead to highly conservative results in some cases. This paper describes a new integrated methodology to perform these assessments that combines one software tool for creating high fidelity crack growth simulations (FRANC3D) with another software tool for performing probabilistic fatigue crack growth life assessments of turbine engine components (DARWIN). DARWIN employs finite element models of component stresses, while FRANC3D performs automatic adaptive re-meshing of these models to simulate crack growth. Modifications have been performed to both codes to allow them to share and exchange data and to enhance their shared computational capabilities. Most notably, a new methodology was developed to predict the shape evolution and the fatigue lifetime for cracks that are geometrically complex and not easily parameterized by a small number of degrees of freedom. This paper describes the integrated software system and the typical combined work flow, and it shows the results from a number of analyses that demonstrate the significant features of the system. INTRODUCTION In recent years, the aero-engine industry and its regulators have been placing increasing emphasis on damage tolerance methods for life management of metallic high-energy rotating components. For some applications, probabilistic methods are used to predict the reliability of fracture-critical engine components [1, 2], thereby reducing the potential for overconservatism in deterministic approaches that always assume worst-case conditions. However, the fracture mechanics methods themselves can also introduce excessive conservatism for some problems. This occurs because current engineering tools usually rely on stress intensity factor (SIF) values that assume highly idealized component and crack geometries, such as part-elliptical planar cracks in rectangular prisms. A new integrated methodology has been developed to perform improved deterministic and probabilistic damage tolerance assessments. The methodology combines a software tool for performing fatigue crack growth (FCG) life and fracture reliability assessments of turbine engine components with another software tool for creating high fidelity crack growth simulations. Modifications were performed to both codes to allow them to share and exchange data. A number of enhancements were performed to the shared computational aspects of both codes. Most significantly, a new methodology was developed to predict the shape evolution and the FCG lifetime for cracks that are geometrically complex and not easily parameterized by a small number of degrees of freedom. This paper will first briefly introduce the two software tools, their combined workflow and interfaces, and the most significant enhancements in the methods. The results from a number of demonstration analyses will be shown to illustrate some of the key features and capabilities of the integrated tool. APPROACH The DARWIN® (Design Assessment of Reliability With INspection) computer program integrates a life assessment based on linear elastic fracture mechanics (LEFM) with a finite element (FE) analysis of component stresses, material anomaly data, probability of crack detection, and inspection schedules to determine the probability of fracture as a function of the

An Optimal Automated Zone Allocation Approach for Risk Assessment of Gas Turbine Engine Components

High-energy rotating components of gas turbine engines may contain rare material anomalies that can lead to uncontained engine failures. The Federal Aviation Administration and the aircraft engine industry have been developing enhanced life management methods to address the rare but significant threats posed by these anomalies. One of the outcomes of this effort has been a zone-based risk assessment methodology in which component fracture risk is estimated using groupings of elements called zones that are associated with 2D finite element (FE) stress and temperature models. Previous papers have presented processes for creation of zones either manually or via an automatic algorithm in which zones are assigned to each finite element in a component model. These processes may require significant human time and computer time. The focus of this paper is on the optimal allocation of multiple finite elements to zones that minimizes the total number of zones required to compute the fracture risk of a component. An algorithm is described that uses a relatively coarse response surface method to estimate the conditional risk value at each node in a finite element model. Zones are initially defined for each finite element in the model, and the algorithm identifies and merges zones based on minimizing the influence on component risk. The process continues until all of the zones have been merged into a single zone. The zone sequence is applied in reverse order to identify the minimum number of zones that satisfies component target risk or convergence threshold constraints. This solution provides the optimal allocation of finite elements to zones. The algorithm is demonstrated for a representative gas turbine engine component. The approach significantly improves the computational efficiency of the zone-based risk analysis process.Copyright

A Tool for Probabilistic Damage Tolerance of Hole Features in Turbine Engine Rotors

Over the past two decades, the Federal Aviation Administration (FAA) and the aircraft engine industry (organized through the Rotor Integrity Sub-Committee (RISC) of the Aerospace Industries Association) have been developing enhanced life management methods to address the rare but significant threats posed by undetected material or manufacturing anomalies in high-energy rotating components of gas turbine engines. This collaborative effort has led to the release of several FAA advisory circulars providing guidance for the use of probabilistic damage tolerance methods as a supplement to traditional safe-life methods. The most recent such document is Advisory Circular (AC) 33.70-2 on “Damage Tolerance of Hole Features in High-Energy Turbine Rotors.” In parallel with this effort, the FAA has also been funding research and development activities to develop the technology and tools necessary to implement the new methods, including a series of grants led by Southwest Research Institute® (SwRI®). The most significant outcome of these grants is a probabilistic damage tolerance computer code called DARWIN® (Design Assessment of Reliability With INspection). DARWIN integrates finite element models and stress analysis results, fracture mechanics models, material anomaly data, probability of crack detection, and uncertain inspection schedules with a user-friendly graphical user interface (GUI) to determine the probability of fracture of a rotor disk as a function of operating cycles with and without inspection. This paper provides an overview of new DARWIN models and features that directly support implementation of the new AC on hole features. The paper also simultaneously provides an overview of the AC methodology itself. Component geometry and stresses are addressed through an interface with commercial three-dimensional finite element (FE) models, including management of multiple load steps and multiple missions. Calculations of fatigue crack growth (FCG) life employ a unique interface with the FE models, sophisticated new stress intensity factor solutions for typical crack geometries at holes, shakedown modules, a menu of common FCG equations, and algorithms to address the effects of varying temperatures on crack growth rates. The primary random variables are based on the default anomaly distributions and probability-of-detection (POD) curves provided directly in the AC. Fracture risk is computed on a per-feature basis using one of several available computational methods including importance sampling, response surface, and Monte Carlo simulation. The approach is illustrated for risk prediction of a representative gas turbine engine disk. The results can be used to gain a better understanding of the AC and how the problem is solved using the probabilistic damage tolerance framework provided in DARWIN.Copyright

Integrating Fatigue Crack Growth into Reliability Analysis and Computational Materials Design

Recently a new methodology was developed for automated fatigue crack growth (FCG) life analysis of components based on finite element stress models, weight function stress intensity factor solutions, and algorithms to define idealized fracture geometry models. This paper describes how the new methodology is being used to integrate FCG analysis into highly automated design assessments of component life and reliability. In one application, the FCG model automation is supporting automated calculation of fracture risk due to inherent material anomalies that can occur anywhere in the volume of the component. Automated schemes were developed to divide the component into a computationally optimum number of sub-volumes with similar life and risk values to determine total component reliability accurately and efficiently. In another application, the FCG model automation is supporting integration of FCG life calculations with manufacturing process simulation to perform integrated computational materials engineering. Calculation of full-field, location-specific residual stresses or microstructure is being linked directly with automated life analysis to determine the impact of manufacturing parameters on component reliability.

Influence of Mission Variability on Fracture Risk of Gas Turbine Engine Components

The need for application of probabilistic methods to fatigue life prediction of gas turbine engine components is being increasingly recognized by the U.S. Military. A physics-based probabilistic approach to risk assessment provides improved accuracy compared to a statistical assessment of failure data because it can be used to (1) predict future risk and (2) assess the influences of both deterministic and random variables that are not included in the failure data. Probabilistic risk and fatigue life prediction of gas turbine engine fracture critical components requires estimates of the applied stress and temperature values throughout the life of the component. These values are highly dependent upon the mission type and may vary from flight to flight within the same mission. Currently, standard missions are specified and used during the engine design process, but the associated stresses can differ significantly from stress values that are based on flight data recorder (FDR) information. For this reason, efforts are made to periodically update the standard missions and to assess the impact on component structural integrity and associated risk of fracture. In this paper, the influence of mission type and variability on fracture risk is illustrated for an actual gas turbine engine disk subjected to a number of different mission loadings. Disk stresses associated with each mission were obtained by scaling finite element model results based on RPM values obtained from engine flight recorder data. The variability in stress values throughout the life of the component was modeled using two different approaches to identify the upper and lower bound value influences on the risk of fracture. The remaining variables were based on default values provided in FAA Advisory Circular (AC) 33.14-1 “Damage Tolerance for High Energy Turbine Engine Rotors”. The risk of fracture was computed using a probabilistic damage tolerance computer code called DARWIN® (Design Assessment of Reliability With Inspection) and compared for each mission type to illustrate the maximum influence of mission type on fracture risk. The results can be used to gain insight regarding the influence of mission type and associated variability on the risk of fracture of realistic engine components.© 2012 ASME

Reliability of Fracture Mechanics Approach to Fatigue

Current fatigue design of fracture-critical components, such as tendons and risers, requires dual fatigue life criteria to be satisfied. The S-N approach includes a safety factor (SF) of 10 on the life of the component, while the fracture mechanics (FM) approach includes a safety factor of 5 on the life through-thickness of an acceptable initial flaw. FM provides critical initial flaw sizes such that suitability of the selected NDE methods and weld acceptance criteria can be established. This paper pertains to a comparative fatigue life reliability study between those two approaches. The objective is to develop a rationale for the selection of a safety factor on fatigue life to use in FM calculations. A reliability-based methodology is proposed and implemented. The SFs for FM are obtained by targeting the reliability obtained in fatigue designs based on historically proven S-N damage approach. Random variables entering both approaches were characterized and a number of weld design cases devised to obtain reliabilities. One important variable is the distribution of initial flaw sizes. For this study, flaw distributions were developed from actual inspection records, accounting for the effects of probability of detection and sizing accuracy of the inspection system, as well as the flaw acceptance criteria during fabrication. Comparisons of reliabilities obtained for designs by both approaches for various quality S-N curves, stress spectra, pipe sizes, and initial flaw sizes indicate that there is ample scope to modify downward the current FM safety factor. However, given the limited scope of this study, it is recommended to asses the FM SF using reliability analysis on a project-specific basis.Copyright