Browse Topic: Fatigue
Traditional safe-life methodologies for rotorcraft structural components rely on deterministic safety factors to account for uncertainty in loads, material properties, and operational usage. While effective for ensuring safety, these approaches lead to early retirement lives and reduced aircraft availability. This paper presents an updated digital twin-based probabilistic framework for rotorcraft component fatigue life assessment that integrates a probabilistic stress–life (S-N) material model, machine learning-based load estimation from flight data, and Monte Carlo uncertainty propagation. The approach is demonstrated for a critical location on the CH-146 Griffon main rotor yoke. Compared with earlier work, the present study advances the framework through independent validation of the load-estimation model and application to available in-service flight data from multiple mission categories. A probabilistic sensitivity analysis is used to examine the separate and combined effects of material variability and load-estimation uncertainty on fatigue life, cumulative probability of failure, and hazard rate. For the CH-146 demonstration case, the results indicate that the material fatigue strength uncertainty has a major impact on the lower tail of the life distribution and the corresponding reliability-based life, whereas load-estimation accuracy uncertainty has a secondary influence on risk metrics. The application of the digital twin framework to operational, search and rescue, and training mission data further shows that mission-specific usage variability plays an important role in the evolution of fatigue damage accumulation and structural risk. Overall, the proposed framework provides a more informative basis for risk-based rotorcraft life assessment by explicitly quantifying uncertainty and incorporating aircraft-specific operational data. The study is intended as a step toward validation of the framework rather than a completed operational deployment.
Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work develops structural fatigue life prediction capability in state-of-the-art emerging progressive damage failure analysis tool CDMat developed at the University of Texas Arlington Advanced Materials and Structures Lab. While JIntegral, implemented in CDMat, appears as the most objective and rigorous approach to predict delamination growth-based fatigue life of composite structures, the key material properties of the J-Integral fatigue model have not been measured with the adequate accuracy. This work addressees a fundamental challenge of eliminating the established and routine assumptions and developed a methodology to determine the key material properties meeting the material input data requirements for the JIntegral based structural fatigue life prediction models. This work generated input data for fatigue crack growth propagation, including optimized input data parameters for the in-situ cohesive law and modeling as-manufactured specimen conditions. In addition, a methodology to account for and characterize the effects of fiber bridging in static tests is presented. It uses a simple standard unidirectional panel for testing, before a component is manufactured, fabricated from the same batch of prepreg to qualify the effects of fiber bridging. Also, this work attempted to determine a minimum conservative initial crack size to streamline the fatigue crack propagation prediction. Fatigue predictions have been demonstrated on a representative composite skin–hat stiffener sub-component section and compared with tests.
A typical helicopter drive system consists of a multi-stage gearbox with highly loaded dynamic components such as gears, shafts, and bearings, crucial for safe flight and landing. Planetary reduction stages are commonly used in the final reduction stage of rotorcraft main gearboxes due to their ability to handle high torques at high gear ratios within a compact envelope. The planet gear, a critical component in this arrangement, is subjected to significant loads on both flanks of its teeth and must meet stringent weight and assembly requirements, leading to a thin rim design with integrated bearing races. This design makes the planet gear susceptible to relevant reduction of its fatigue life. This paper explores analysis methods to evaluate the damage resistance of the planetary stage assembly, focusing on the planet gear. The study aims to assess the "growth" or "no growth" condition of the planet gear against defined flaw defects. An iterative calculation loop determines the critical length and position of a crack that may lead to full crack propagation and, in worst cases, to system jamming. Initial crack propagation simulations use NASGRO software, with stress fields derived from a non-linear FEM of the planet gear availing of detailed Transmission3D model for teeth meshing forces evaluation. Further additional analysis can involve a dedicated FE model of the crack, iteratively updating its geometry. The impact of crack propagation on the remaining components of the planetary stage assembly is also addressed, considering the unbalanced load conditions caused by stiffness loss in the planet gear. The dissertation object of this paper aims to outline a comprehensive, effective and efficient procedure to determine the maximum allowable defect size for "no growth" condition and the operational hours until failure, providing a robust approach to support the strength substantiation of the involved components.
The oil cooling fan of a Main Gearbox (MGB) is a mechanically-driven component whose purpose is to force an air flow through an air cooled oil cooler; its performance is crucial in ensuring that the MGB oil temperature does not exceed a predefined threshold, set to alert the crew in case of an abnormal situation. The design and the certification of a cooling fan is a process involving several steps and multiple disciplines; mechanical design, aerodynamic analysis, dedicated tests carried out both on rigs and at aircraft level need to be exploited as complementary tools to assess the correct aero-mechanical behavior of the system. The aerodynamic assessment is associated to performance, measured in terms of MGB oil temperature: considering a comparison between two cooling fans, one outperforms the other if the resultant MGB oil temperature is lower, keeping the same boundary conditions (engine torque, wind speed, ambient temperature, etc.). The correct mechanical behavior is instead associated to multiple requirements that need to be satisfied: target fatigue life, high and low temperature limits, blade loss containment, maximum speed, etc. The process connecting the very first blank page to the certification of the part is usually linear and well defined, and the experience gathered by Leonardo Helicopters over the years has allowed to follow this path easily in most cases. This paper discusses an intriguing situation: the goal of the activity herein discussed was to replace an obsolete oil cooling fan with a new one. The latter was expected to outperform the first both from the aerodynamic point of view (i.e. lower MGB oil temperature) and from the mechanical point of view (i.e. longer life). The preliminary CFD analyses and the rig tests proved that the new fan was able to outperform the legacy one in regards of the aerodynamic requirement; nevertheless, during the last steps of verification carried out at aircraft level, the installation of the new fan resulted in a higher MGB oil temperature. This triggered dedicated troubleshooting: a sequence of investigative tests performed by means of dedicated rigs, targeted at properly identifying the aerodynamic operative point of the cooling fan and at understanding the reason behind the unexpected outcomes obtained. The relevance of this contribution does not lie in the description of the certification process per se, rather in the nonlinear dynamics characterizing the test sequence once the flight activities pointed out an unpredicted behavior.
Traditional safe-life methodologies for rotorcraft structural components often result in overly conservative life estimates, increasing maintenance costs and reducing aircraft availability. This study explores the integration of digital twin concepts with probabilistic modeling and machine learning to enhance structural life assessment, demonstrated through a practical case involving the Royal Canadian Air Force CH-146 Griffon helicopter. A probabilistic fatigue model determines a fatigue life distribution by incorporating material variability and uncertain operational loads inferred directly from flight data. Unlike conventional approaches, this method dynamically estimates load spectra, including uncertainty instead of relying on conservative assumptions. Monte Carlo simulations are used to quantify structural risk and assess the impact of load and material uncertainties. Sensitivity analyses highlight these uncertainties’ contributions to failure probability. The proposed approach provides probabilistic life predictions, supporting risk-based maintenance strategies to potentially optimize operational efficiency. The long-term goal is to develop an adaptive digital twin model that continuously updates with new operational flight data, enhancing predictive accuracy for helicopter fleet management.
Rotorcraft dynamic component fatigue lives and corresponding reliability have long been derived from three major contributors: material strength, loads, and usage. This paper provides a historical perspective of the contribution of aircraft usage to overall U.S. Army rotorcraft dynamic component reliability. A quick background of how we got to a six-nines reliability requirement is first provided. Different types of usage spectra and the nuances and trade-offs of two specific usage gathering methods, pilot surveys and usage monitoring, are discussed. Finally, I describe where usage spectrum fits into fatigue life calculations and the existing reliability policy and requirements. Each OEM (e.g., Bell Helicopter, Boeing, Sikorsky) has been free to develop their own fatigue methods over the years. These differences in method can lead to vastly different results, even with the same input parameters as evidenced by a now well-known round robin problem. There is notable variability between OEM methodologies, each with viable solutions to this trivariate problem. In the interest of normalizing independent U.S. Government (USG) assessments across multiple OEM paradigms, the Army is investigating a USG method to assess the reliability contribution from usage. No new methods are presented herein, only findings of previous work. Uncited opinions herein are those of the author based on literature review, peer discussions, and experience with U.S. Army and U.S. Air Force (USAF) airworthiness processes. Reliability values in this paper are approximate, as there are elements of statistical distribution and non-statistical estimation that contribute.
Previous work documented the use of IVHMS data on the U.S. Army's fleet of UH-60 Black Hawk helicopters to update the fatigue lives of six specific components on the A/L and M models. This paper documents a significant expansion of the level of data applied to the usage spectrum, as well as applying it to all components on the aircraft. As a design spectrum for the yet to be fielded Improved Turbine Engine (ITE) equipped UH-60M, changes due to new engine capability needed to be addressed. The new spectrum has been developed and is being used for planning of flight testing. The spectrum along with flight test loads will be used to generate fatigue lives for the new aircraft. Once deployed for several years the spectrum will be reviewed to determine if any changes are needed. This work highlights what the Army considers to be the most significant issues when applying monitored usage to critical fatigue components, and rationale for dealing with issues such as insufficient data for various purposes.
Shot peened components present a challenge for the structural analyst when nicks, scratches and gouges are discovered. A common repair scheme calls for blending away of the defect with an appropriate grit abrasive. Though the blending operation removes the defect, it also takes away a portion the beneficial compressive layer as well as the cold-worked material. Large repair facilities may have touch-up shot peen capability but technicians in a field repair setting typically do not. If the shot peen cannot be restored, the structural analyst must have a method to quantify the effect on fatigue life of the repaired part. The purpose of this technical paper is to substantiate analytical techniques for evaluating the fatigue life of a shot peened part after a blend operation. In addition to practical methods to estimate the magnitude of the residual stresses, a numerical method is introduced using finite element modeling of shot peen impacts with non-linear finite element code and validation by a simulated Almen strip.
This paper presents a framework with associated concepts to define a method of compliance for the failure rate requirements of the Army Military Airworthiness Certification Criteria (AMACC), Chapter 5, for fleet qualification and first flight. The fleet failure rate requirement is paraphrased as less than one structural failure in 20 million flight hours at 95% confidence and applies specifically to fatigue failure of primary structural elements (PSEs). This method of compliance assumes a reliability model sufficient to support an analytical failure rate that bounds the uncertainty arising from practical constraints of an aircraft qualification program and aims to optimize the competing objectives of safety of flight and operational capability. The requirements as defined verify that the aircraft system will perform as intended for a specified fleet life and flight test program duration and serve as the analytical basis for the assessment of emergent issues identified throughout the product life cycle.
A state-of-the-art emerging progressive damage failure analysis tool CDMat has been successfully applied to multiple material systems on open-hole tension and compression, and double shear bearing laminate coupons under static and fatigue loading including simulation to ultimate failure. CDMat also successfully demonstrated component-level strength/fatigue analysis under the Air Force Composite Airframe Life Extension (CALE) and the Fail-Safe Technologies for Bonded and Unitized Composite Structures (FASTBUCs) Programs. Building on the success of CDMat an integrated software solution for certification and sustainment of rotorcraft primary composite structures is being developed. A method and an algorithm for fatigue crack growth simulation in laminated structures are proposed to improve the accuracy of CDMat fatigue predictions. The method is based on using cohesive material model, tracking material points at the crack front, and calculating the pointwise energy release rate employing the J-integral. The algorithm was implemented as a set of user material subroutines developed within the framework of explicit finite element formulation for ABAQUS. The effectiveness of the method is demonstrated on several examples of Mode I and II fatigue crack growth.
This paper documents the re-evaluation and updates to the previous Partial Regime Recognition Spectrum effort for the MH-47G using Structural Usage Monitoring System (SUMS). Further validation of the SUMS algorithm allowed for additions to the spectrum. These additions include more refined categorization of turn and partial power descent regimes based on angle of bank and descent rates, respectively; high load prorates for turns, partial power descents, level flight, and climbs based on the Cruise Guide Indicator; exceedances of maximum density altitude; and use of occurrences for Landing and Run-On Landing regimes. Additional years of flight data from 2013 to 2019 were included in this effort. The updated usage spectrum for the Army MH-47G aircraft has been delivered to the OEM (Original Equipment Manufacturer). The OEM calculated new fatigue lives and updated the "Fatigue Substantiation Report", which will soon be fielded.
This paper details an analysis methodology for a primary structure component on a tandem rotor helicopter that has been shown to experience fatigue damage in operation. The primary structure component is a web in the aft pylon of the helicopter. The web carries a complex set of loads in flight with various forces and moments applied along its boundaries and rotor torque reacted around a large rectangular cutout in the web. Due to the complex loading applied to the web, there is no clear location that can be considered to carry a "gross" stress, rendering traditional hand calculation methods (such as the use of stress concentration factors applied to a gross stress) impractical. The analysis detailed in this paper considers the application of flight loads to Finite Element Models to determine stresses in the web, which are used to evaluate fatigue life based on various flight conditions.
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Carbon fiber reinforced polymer composites (CFRP) are extensively used as structural components in rotorcraft applications. Here, we report considerable improvement in the fatigue life of CFRP through the infiltration of nanoscale silica particles into the epoxy resin matrix (nanoCFRP). Fumed silica nanoparticles were initially added to the epoxy resin to prepare epoxy-silica nanocomposites, which were demonstrated to have superior fracture and fatigue properties. Fractographic analysis indicated presence of various key toughening mechanisms including crack deflection, plastic void growth as well as a hitherto unreported heterogeneity induced mesoscale toughening effect. The epoxy-silica nanocomposite resin was then used as the matrix material to fabricate nanoCFRP. Cyclic flexural bending tests indicate significant fatigue life enhancement for the nanoCFRP. The enhancement is especially pronounced in the high cycle fatigue regime. This enhancement in high cycle fatigue is indicative of transfer of small-scale toughening mechanisms from the silica-epoxy nanocomposite resin to the nanoCFRP system. Such nanoCFRP show promise to improve the fatigue life and reduce the operational/maintenance cost for next generation rotorcraft.
Presently the fatigue lives of MH-60R dynamic components and airframe are based on a usage spectrum developed using pilot surveys. In order to better define the usage spectrum and to extend component and airframe fatigue life, the Health & Usage Spectrum (HUMS) System was installed on the U.S. Navy MH- 60R Rotorcraft. So far 207 aircraft are equipped with the HUMS systems and 121,334 flight hours of good data have been recorded. The regime recognition programs recognize 315 maneuvers, but are consolidated to 94 maneuvers of MH-60R usage spectrum, for which the component measured loads are available. To better define usage spectrum in detail and compute realistic component fatigue life, an additional maneuver of low Angle Of Bank (AOB) from 10 to 25 degrees was added, but the measured component loads were not available at this AOB to implement HUMS. Thus, measured flight loads data of level flight and AOB turns at 30, 45, and 60 degrees were utilized to derive component loads at 20 degrees by spline cubic interpolation technique. The cubic interpolation technique was applied to measured minimum, average, and maximum loads of variation at 10, 30, 45, and 60 degrees to interpolate load at 20 degree. This technique was applied to interpolate loads for pitch control rod, swash-plate, drag damper, shaft bending moments, blade cuff stresses, and flap deflections. The spline interpolation loads correlated with measured available loads of pitch control rod and blade stresses. The probabilistic fleet usage spectrum of various severities was developed using the HUMS recorded data of 121, 334 hours from 179 rotorcraft with and without low AOB usage. It is evident that fatigue life with 20 AOB split is significantly higher for all dynamic components. Thus, to implement HUMS successfully, it is necessary to compute loads that are not available in the original component fatigue life calculations. Further prorates of gross weigh (GW), velocity and altitude based on the HUMS fleet usage should be implemented to extend component fatigue lives.
Gear design changes impact on gear crack propagation trajectory is investigated through numerical study. General purpose linear elastic fracture mechanics software, FRANC2D and FRANC3D, are used to simulate 2D and 3D gear crack propagation. FRANC can model non-planner, arbitrary shape crack surface for crack tip stress distributions, stress intensity factors, and crack propagation analyses. Maximum tensile stress and NASGRO4 fatigue crack growth models are employed to predict crack propagation direction and life. Three-dimensional idler gear crack propagation simulation shows the predicted crack trajectory is close to the field observation. Various 2D models are simulated to investigate the crack trajectory impact factors and design strategies to prevent gear rim failure. As shown in previous studies, the initial crack position and orientation play pivot role to control gear failure mode - tooth or rim. For a fixed crack position, this study shows the ratio between bend stress and centrifugal stress dominates gear fracture mode. The less centrifugal stress, the crack more likely to break tooth, while lower bend stress more likely lead to break rim. To prevent rim failure through increasing the rim thickness results in a significant weight penalty. The larger the gear, more the weight penalty. Based on the simulation results, the recommended design strategy is to evaluate gear rim failure risk during the gear train layout phase. It is difficult to be improved at the individual gear design phase.
The aim of this study is to evaluate the effect of small defects on the fatigue life of the 18CrNiMo7-6 material. The evaluations are based on comparisons of S/N curves, which are obtained through coupon-level fatigue testing of pristine and intentionally flawed specimen sets. Fractographic examinations are performed on the failed specimens from both sets in order to investigate the fracture characteristics and discover any potential differences in failure mechanisms of two sets. Experimental findings regarding endurance limit of the flawed set are also compared with analytical evaluations. In these evaluations, the small defect effect is also considered on top of the linear fracture mechanics approach. Finally, the differences between the outcomes of fatigue life evaluations, when different methods are used, are discussed.
Advanced structural analysis methods, known as progressive damage and failure analysis (PDFA) tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work shows the ability of the PDFA implemented in CDMat software developed at the University of Texas Arlington Advanced Materials and Structures Lab (AMSL) to predict strength and fatigue failure of a Common Feature Test Component (CFTC) - representative of flight-critical structural attributes and failure modes - without a priori knowledge of the test result. CFTC advanced structural features include a composite skin made of unidirectional tape, a fabric hat stiffener, and a mechanically fastened aluminum rib. CFTC, developed by Boeing under the Air Force Research Laboratory (AFRL) Composite Airframe Life Extension (CALE) Program, Assessing the Durability and Damage Tolerance of Advanced Composite Structural Features, has been the most complex PDFA validation article to date.
A key component for implementing the digital twin approach is to apply a validated high fidelity simulation tool to generate a mapping between the virtual test and structural performance. Due to the high computational intensity of physical simulation tools, their application for a complex system along with its error estimation can be time consuming. In addition, given the limited data gathered from sensors, onsite inspection, and tests at different configurations, it is imperative to create a high fidelity and efficient model based on the previously gathered information and enhance the model when more data points are gathered. Motivated by this, we develop a machine learning based digital twin simulation framework to predict fatigue life of a structural component from available information gathered. Different from a conventional physical simulation approach, the prediction error from the physical simulation and machine learning are explicitly obtained, in addition to the improvement of the computational efficiency at the prediction stage via the trained machine learning model. To illustrate the idea of this modeling strategy, we applied our developed 3D extended finite element toolkit for Abaqus (XFA3D) as a virtual testing tool for fatigue crack path and life prediction of a welded metallic component in conjunction with the observed testing data. Using machine learning techniques, we first estimate prediction error from the physical model based on previous cross validation results, and then predict the fatigue life in the presence of uncertainties associated with fabrication induced imperfection, welding induced residual stress, and the machine learning errors. It is found that the inclusion of as-manufactured characteristics and uncertainties are essential for the application of a digital twin approach for the total life management of aging structures.
Icing of the fuselage and blades may occur when the helicopter is flying in the icing area. If ice accretion occurs in the ADS(Air Data System) of the fuselage, normal speed and altitude information are lost, making it difficult to flight. When windshield icing occurs, the view of pilot is limited and flight is difficult. Also, the ice accretion of the blades deforms the outer shape of the blades (Ref. 1) and makes the dynamic characteristics unstable due to an abnormal weight increase, resulting in deterioration of performance, deterioration of maneuverability, and structural instability. To avoid this, an anti-icing or de-icing system is required. Therefore, if the aircraft is not fitted with a proper anti-icing system, it is not possible to operate under icing conditions. However, it is difficult to design a proper anti-icing system considering the position of anti-icing protection area and icing phenomenon due to limitation of electric power, weight, thermal damage temperature limit, shape and so on. It is essential of understanding of anti-icing mechanism for selection and design of appropriate anti-ice system considering configuration (impingement limit, collection efficiency), material (thermal fatigue limit, heating source) and icing amount (collection efficiency, stagnation point). This paper introduces the preparation procedure (analysis and simulation, design, scaled model test) of icing certification test and the artificial/natural icing flight test method for proving helicopter icing through KUH case, and describes the characteristics of each system for the de/anti-icing system design of the helicopter, the computational analysis method, comparison with the test result, and the improvement method of the de/anti-icing system.
We present an analysis of the influence of various cycle counting approaches on the estimated fatigue life for critical rotorcraft components. The entire study is demonstrated using the pitch link of a utility helicopter similar to the UH-60 Black Hawk helicopter as a representative component. Different counting methods predict fatigue life differently due to the varied fundamental parameters or approaches they take to characterize the stress profile. Hence, it becomes vital to choose an appropriate method for a particular application so that it does not over or under predict fatigue life during the design process. Based on this, we investigated four methods namely, Rainflow, Peak, Level and Range counting
This paper presents three strategies of integrating structural loads analysis with structural flight tests with a goal of increasing confidence without increasing flight test costs. The strategies considered in this paper are direct substitution, normalized load spectra, and loads interpolation. Discussion includes impacts on the fatigue methodology and analysis validation associated with each strategy. The paper also explores potential benefits and challenges for each strategy. While the direct substitution strategy requires minimal changes to the fatigue methodology to apply load factors for greater confidence, this strategy also requires an extensive validation effort. Each of the other two strategies reduces the extent of the validation effort, but also requires changes in the fatigue methodology and industry consensus. Finally, for normalized load spectra and loads interpolation strategies, the paper discusses the relationship of each strategy to loads monitoring and estimation.
Rotorcraft components, which are often made with reinforced fiber composites, are subjected to severe fatigue loadings due to increased performance demands. Therefore, considerable research interest exists in improving fatigue life of conventional fiber reinforced composites. Nanocomposites are a new class of materials which seek to improve mechanical performance of materials by creating nanoscale crack-nanofiller interactions. In this study we demonstrate the fatigue life improvement of conventional composites by addition of SiO2 nanofillers. The epoxy resin was initially modified with nanofillers to test the static fracture toughness. Once the improvement in static facture toughness was confirmed, three phase modified fiber reinforced composites were made using the modified resin. Cyclic tests were performed at various stress level which demonstrate that three phase nanocomposites perform better than conventional fiber reinforced composites. Fractographic analysis suggests that nanofiller de-bonding from the matrix as well as crack deflection around nanofiller clusters contributes to the improved fracture toughness and fatigue life.
This paper presents research addressing the technological gap in the predictive capabilities of modern computational fluid-structures interaction (FSI) in the context of the US Navy's requirement for accurate loads prediction for both critical rotorhead and fuselage components in the context of a real aircraft’s fatigue life tracking program, with the ultimate objective of loads accuracy and performance robust enough to support near real-time lifing assessments across the full flight regime. This research, funded via US Navy STTR N17A-T009, touches on a broad range of innovative research areas, starting with application of the Load Confluence Algorithm (LCA) to a coupled main rotor-fuselage/tail rotor model for the UH-60A. This paper will document research to date, extending beyond traditional approaches such as the main rotor analyzed in isolation or a fuselage model affected only by the inflow generated via an actuator disk model of the main rotor. This includes the development of use of modal super- elements (MSEs) to model more complex rotorcraft components, such as the fuselage for use in main rotor loads simulations, as well as the stationary and rotating swashplates, and the rotating scissors. This paper will also document advances made in geometry modeling and mesh generation, computational mechanics of fluids and structures, modeling of coupled multi-physics phenomena, and high-performance computing, with improved representation of complex geometry and multi-physics phenomena. Lastly, these enhanced whole aircraft loads prediction methods will be examined within the context of future application to a structural fatigue life tracking program, in terms of accuracy, usability, and application performance.
Low Pressure Carburization (LPC) is widely used in the aerospace industry for hardening components made of steels with high alloy content and high heat resistant properties. The traditional gas carburizing process often generates Intergranular Oxidation (IGO) near the surface due to the existence of oxygen in the furnace atmosphere, which needs to be removed by grinding after hardening to restore bending and contact fatigue performance. LPC processing is done in a low pressure chamber without the existence of oxygen, so the surface microstructure is improved by eliminating IGO. High temperature resistant steels require high alloy element contents, and some elements are strong carbide formers, such as Cr, Mn, Mo, and V, etc. During LPC processing, both iron and alloy carbides can be formed, which significantly affect the carburization time required to reach a specified case depth and surface carbon. The carbides formed during the LPC process may not decompose completely prior to quench hardening, and these primary carbides will end up in the final processed parts. If the size of these primary carbides is not controlled, both bending and contact fatigue performance may be decreased. In order to control carbide formation during LPC, the carbon diffusivity of a material must be characterized. This characterization was recently performed under a program between DANTE Solutions and the Combat Capabilities Development Command Aviation and Missile Center (CCDC AvMC). In this research, a specifically designed coupon was used to characterize the carbon diffusivity and carbide forming properties during LPC processes. Using the characterized material properties, LPC process recipes can be designed by using modeling to achieve specific case depth and surface carbon content. The work was demonstrated using Pyrowear 675 steel and DANTE commercial heat treatment modeling software.
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