The Experts below are selected from a list of 177 Experts worldwide ranked by ideXlab platform

Michael R Hill - One of the best experts on this subject based on the ideXlab platform.

  • the effects of laser peening and shot peening on fretting fatigue in ti 6al 4v coupons
    Tribology International, 2009
    Co-Authors: Michael R Hill
    Abstract:

    Abstract This paper describes testing of Ti–6Al–4V coupons in fretting fatigue and compares the effects of mechanical surface treatments on performance. Fretting fatigue tests were performed using a proving ring for fretting load, bridge-type fretting pads, and applied tension–tension cyclic fatigue stress. As-machined (AM), shot peened (SP), and laser peened (LP) coupons were evaluated, and data generated to compare residual stress, surface condition, lifetime, and fractographic detail encountered for each. Near-surface residual stress in SP and LP coupons was similar. The layer of compressive residual stress was far deeper in LP coupons than in SP coupons and, consequently, subsurface tensile residual stress was significantly greater in LP coupons than in SP coupons. SP coupons exhibited a rough surface and had the greatest volume of fretting-induced wear. LP coupons exhibited a wavy surface and had a small volume of wear localized at wave peaks. SP coupons had the greatest fretting fatigue lifetime, with significant improvement over AM coupons. Lifetimes of LP coupons were similar to those for SP coupons at high fatigue stress, but fell between AM and SP coupons at lower fatigue stress. Fractographic evaluation showed that fractures of AM samples were preceded by initiation of fretting-induced Cracks, transition of a Lead fretting Crack to mode-I fatigue Crack growth, and Crack growth to failure. SP and LP samples exhibited behavior similar to AM samples at high fatigue stress, but in coupons tested at low stress the Lead Crack initiated subsurface, near the measured depth of maximum tensile residual stress, despite the presence of fretting-induced Cracks. The level of fatigue stress above which Lead Cracks were initiated by fretting was higher for LP than for SP, and was predicted with good accuracy using an analysis based on linear elastic fracture mechanics, the fatigue Crack growth threshold stress intensity factor range, and superposition of measured residual stress and applied fatigue stress.

  • The effects of laser peening and shot peening on fretting fatigue in Ti–6Al–4V coupons
    Tribology International, 2009
    Co-Authors: Kevin K. Liu, Michael R Hill
    Abstract:

    Abstract This paper describes testing of Ti–6Al–4V coupons in fretting fatigue and compares the effects of mechanical surface treatments on performance. Fretting fatigue tests were performed using a proving ring for fretting load, bridge-type fretting pads, and applied tension–tension cyclic fatigue stress. As-machined (AM), shot peened (SP), and laser peened (LP) coupons were evaluated, and data generated to compare residual stress, surface condition, lifetime, and fractographic detail encountered for each. Near-surface residual stress in SP and LP coupons was similar. The layer of compressive residual stress was far deeper in LP coupons than in SP coupons and, consequently, subsurface tensile residual stress was significantly greater in LP coupons than in SP coupons. SP coupons exhibited a rough surface and had the greatest volume of fretting-induced wear. LP coupons exhibited a wavy surface and had a small volume of wear localized at wave peaks. SP coupons had the greatest fretting fatigue lifetime, with significant improvement over AM coupons. Lifetimes of LP coupons were similar to those for SP coupons at high fatigue stress, but fell between AM and SP coupons at lower fatigue stress. Fractographic evaluation showed that fractures of AM samples were preceded by initiation of fretting-induced Cracks, transition of a Lead fretting Crack to mode-I fatigue Crack growth, and Crack growth to failure. SP and LP samples exhibited behavior similar to AM samples at high fatigue stress, but in coupons tested at low stress the Lead Crack initiated subsurface, near the measured depth of maximum tensile residual stress, despite the presence of fretting-induced Cracks. The level of fatigue stress above which Lead Cracks were initiated by fretting was higher for LP than for SP, and was predicted with good accuracy using an analysis based on linear elastic fracture mechanics, the fatigue Crack growth threshold stress intensity factor range, and superposition of measured residual stress and applied fatigue stress.

Lorrie Molent - One of the best experts on this subject based on the ideXlab platform.

  • The Lead Crack concept applied to defect growth in aircraft composite structures
    Composite Structures, 2017
    Co-Authors: Lorrie Molent, C. Forrester
    Abstract:

    Currently, the cyclic fatigue growth and residual strength of damaged aircraft composite structures under operational loads is not fully understood. This Leads to structures generally being designed to a no damage growth criterion with many knock down factors included to cover unknown/untested effects. Thus, full optimisation of composite aircraft structures is unlikely to be achieved under the no damage growth criterion. In 2009 the US Federal Aviation Administration (FAA) introduced a slow growth approach to certifying composite, adhesively bonded structures and bonded repairs which could improve the situation and is worthy of further investigation. In this paper the growth of some (limited) damage types available in the literature are reviewed and a framework proposed to address the damage tolerance assessment of these structures.

  • Fatigue Cracking from a corrosion pit in an aircraft bulkhead
    Engineering Failure Analysis, 2014
    Co-Authors: Simon Barter, Lorrie Molent
    Abstract:

    Abstract For aircraft, the repair and management of corrosion can be a major through life cost driver as well as an availability degrader. A common form of corrosion in aircraft structural alloys is pitting corrosion. This paper describes fatigue Cracking in a laboratory environment that had initiated from a significant corrosion pit in a locally shot-peened aluminium alloy 7050-T7451 bulkhead of an F/A-18 Hornet aircraft. Extensive study of this material under Hornet type loading has shown that if the initial equivalent pre-Crack size is determined, then the fatigue Crack growth follows the Lead Crack form, as was the case for this corrosion-pit initiated Crack in the absence of retardation affects. This is a valuable observation which should allow the growth of Cracks from corrosion pits to be determined through fatigue Crack modelling without the need to account for the potential of corrosion-assisted fatigue effects, which are postulated here to be insignificant for combat-type aircraft. The main outcome of the investigation was that whilst the physical size of the corrosion pit was large (442 μm deep) its effect as a fatigue Crack initiator i.e. its effective pre-Crack size (EPS), was significantly smaller (i.e. approximately 10 μm deep).

  • Typical fatigue-initiating discontinuities in metallic aircraft structures
    International Journal of Fatigue, 2012
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept, based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and the fatigue Cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

  • the Lead Crack fatigue lifing framework
    International Journal of Fatigue, 2011
    Co-Authors: Lorrie Molent, Simon Barter, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. Furthermore, these data, particularly for Lead Cracks, are characterized by exponential Crack growth behaviour. Because of this common characteristic, the DSTO framework can use Lead Crack growth data to provide reasonable (i.e. not overly conservative) lower-bound estimates of typical Crack growth lives of components, starting from small natural discontinuities and continuing up to Crack sizes (thus encompassing short-to-long Crack growth) that just meet the residual strength requirements. Scatter factors based on engineering judgement are then applied to these estimates to determine the maximum allowable service life (safe life limit). The aim of the paper is to present the framework of assumptions and observations used in conjunction with a unique measure of the initiating discontinuity and a simple Crack growth law to predict a lower bound fatigue life estimate.

  • Fatigue-initiating discontinuities in aircraft structures
    2011
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and fatigue Cracks are examples from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

Rjh Wanhill - One of the best experts on this subject based on the ideXlab platform.

  • Typical fatigue-initiating discontinuities in metallic aircraft structures
    International Journal of Fatigue, 2012
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept, based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and the fatigue Cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

  • the Lead Crack fatigue lifing framework
    International Journal of Fatigue, 2011
    Co-Authors: Lorrie Molent, Simon Barter, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. Furthermore, these data, particularly for Lead Cracks, are characterized by exponential Crack growth behaviour. Because of this common characteristic, the DSTO framework can use Lead Crack growth data to provide reasonable (i.e. not overly conservative) lower-bound estimates of typical Crack growth lives of components, starting from small natural discontinuities and continuing up to Crack sizes (thus encompassing short-to-long Crack growth) that just meet the residual strength requirements. Scatter factors based on engineering judgement are then applied to these estimates to determine the maximum allowable service life (safe life limit). The aim of the paper is to present the framework of assumptions and observations used in conjunction with a unique measure of the initiating discontinuity and a simple Crack growth law to predict a lower bound fatigue life estimate.

  • Fatigue-initiating discontinuities in aircraft structures
    2011
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and fatigue Cracks are examples from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

Simon Barter - One of the best experts on this subject based on the ideXlab platform.

  • fatigue Crack growth lessons from thirty five years of the royal australian air force f a 18 a b hornet aircraft structural integrity program
    International Journal of Fatigue, 2020
    Co-Authors: Ben Main, R. Singh, L. Molent, Simon Barter
    Abstract:

    Abstract As the Royal Australian Air Force (RAAF) retires its fleet of F/A-18 A and B model Hornets, it is opportune to reflect upon some of the in-service fatigue Crack growth lessons accrued from 35 years of the aircrafts Aircraft Structural Integrity Program (ASIP). In this paper, the authors have selected a small number of the key lessons and observations arising from real-world fatigue problems in the RAAF Hornet fleet that have contributed to advances in the state of the art of fatigue Crack growth measurement and prediction. The importance of the Lead Crack concept, quantitative fractography, fatigue testing, nucleating discontinuities, the physically short Crack regime, as well as individual aircraft tracking are given prominence. It is hoped that this review of the metal fatigue challenges for a highly stressed lightweight structure that have driven research into fatigue Crack growth will be valuable to those aspiring to make impactful contributions to the field and the sustainment of aircraft fleets worldwide.

  • Fatigue Cracking from a corrosion pit in an aircraft bulkhead
    Engineering Failure Analysis, 2014
    Co-Authors: Simon Barter, Lorrie Molent
    Abstract:

    Abstract For aircraft, the repair and management of corrosion can be a major through life cost driver as well as an availability degrader. A common form of corrosion in aircraft structural alloys is pitting corrosion. This paper describes fatigue Cracking in a laboratory environment that had initiated from a significant corrosion pit in a locally shot-peened aluminium alloy 7050-T7451 bulkhead of an F/A-18 Hornet aircraft. Extensive study of this material under Hornet type loading has shown that if the initial equivalent pre-Crack size is determined, then the fatigue Crack growth follows the Lead Crack form, as was the case for this corrosion-pit initiated Crack in the absence of retardation affects. This is a valuable observation which should allow the growth of Cracks from corrosion pits to be determined through fatigue Crack modelling without the need to account for the potential of corrosion-assisted fatigue effects, which are postulated here to be insignificant for combat-type aircraft. The main outcome of the investigation was that whilst the physical size of the corrosion pit was large (442 μm deep) its effect as a fatigue Crack initiator i.e. its effective pre-Crack size (EPS), was significantly smaller (i.e. approximately 10 μm deep).

  • Typical fatigue-initiating discontinuities in metallic aircraft structures
    International Journal of Fatigue, 2012
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept, based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, has been developed by the DSTO for metallic primary airframe components. This framework is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force fleet. Like the original Damage Tolerance concept, developed by the United States Air Force, this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and the fatigue Cracks that have grown from them are taken from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

  • the Lead Crack fatigue lifing framework
    International Journal of Fatigue, 2011
    Co-Authors: Lorrie Molent, Simon Barter, Rjh Wanhill
    Abstract:

    Abstract A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. Furthermore, these data, particularly for Lead Cracks, are characterized by exponential Crack growth behaviour. Because of this common characteristic, the DSTO framework can use Lead Crack growth data to provide reasonable (i.e. not overly conservative) lower-bound estimates of typical Crack growth lives of components, starting from small natural discontinuities and continuing up to Crack sizes (thus encompassing short-to-long Crack growth) that just meet the residual strength requirements. Scatter factors based on engineering judgement are then applied to these estimates to determine the maximum allowable service life (safe life limit). The aim of the paper is to present the framework of assumptions and observations used in conjunction with a unique measure of the initiating discontinuity and a simple Crack growth law to predict a lower bound fatigue life estimate.

  • Fatigue-initiating discontinuities in aircraft structures
    2011
    Co-Authors: Simon Barter, Lorrie Molent, Rjh Wanhill
    Abstract:

    A fatigue lifing framework using a Lead Crack concept has been developed by the DSTO for metallic primary airframe components. The framework is based on years of detailed inspection and analysis of fatigue Cracks in many specimens and airframe components, and is an important additional tool for determining aircraft component fatigue lives in the Royal Australian Air Force (RAAF) fleet. Like the original Damage Tolerance (DT) concept developed by the United States Air Force (USAF), this framework assumes that fatigue Cracking begins as soon as an aircraft enters service. However, there are major and fundamental differences. Instead of assuming initial Crack sizes and deriving early Crack growth behaviour from back-extrapolation of growth data for long Cracks, the DSTO framework uses data for real Cracks growing from small discontinuities inherent to the material and the production of the component. To this end, this paper examines the types of discontinuities that initiate fatigue Cracks in typical metallic airframe structures. These discontinuities and fatigue Cracks are examples from coupon, component and full-scale tests, and also from service aircraft, including commercial transport aircraft and high performance military aircraft.

L. Molent - One of the best experts on this subject based on the ideXlab platform.

  • fatigue Crack growth lessons from thirty five years of the royal australian air force f a 18 a b hornet aircraft structural integrity program
    International Journal of Fatigue, 2020
    Co-Authors: Ben Main, R. Singh, L. Molent, Simon Barter
    Abstract:

    Abstract As the Royal Australian Air Force (RAAF) retires its fleet of F/A-18 A and B model Hornets, it is opportune to reflect upon some of the in-service fatigue Crack growth lessons accrued from 35 years of the aircrafts Aircraft Structural Integrity Program (ASIP). In this paper, the authors have selected a small number of the key lessons and observations arising from real-world fatigue problems in the RAAF Hornet fleet that have contributed to advances in the state of the art of fatigue Crack growth measurement and prediction. The importance of the Lead Crack concept, quantitative fractography, fatigue testing, nucleating discontinuities, the physically short Crack regime, as well as individual aircraft tracking are given prominence. It is hoped that this review of the metal fatigue challenges for a highly stressed lightweight structure that have driven research into fatigue Crack growth will be valuable to those aspiring to make impactful contributions to the field and the sustainment of aircraft fleets worldwide.

  • Using the Lead Crack framework to reduce durability test duration
    The Aeronautical Journal, 2019
    Co-Authors: L. Molent, R. Singh
    Abstract:

    ABSTRACTAircraft full-scale fatigue tests are expensive and time-consuming to conduct but are a critical item on the certification path of any aircraft design or modification. This paper outlines a proposal that trades cycling hours for increased detail in the teardown of a metallic test article. A method for determining the equivalent demonstrated Crack size (and Crack growth curve) at the mandated test life utilising the Lead Crack framework is demonstrated. It is considered that the test duration can be significantly reduced, whilst still achieving all the desired outcomes of a certification program.

  • Fatigue 2010 The Lead fatigue Crack concept for aircraft structural integrity
    2010
    Co-Authors: L. Molent
    Abstract:

    Over many years of quantitative fractographic examination of fatigue Cracking from in-service and full-scale fatigue tests of metallic airframe components, it has been consistently observed that the largest Cracks formed have grown in an approximately exponential manner. These Crack growth observations range from the initiation of Cracks and their early growth from a few micrometers through to many millimetres in length. It appears that these Lead Cracks commence growing shortly after the airframe is introduced to the loading environment. Furthermore these Cracks usually initiate from production-induced or, less frequently, inherent material discontinuities. Based on these two observations, an aircraft lifing methodology that is based on the results of fatigue testing programs utilising the Lead Crack concept has been developed and implemented as an additional tool in the determination of aircraft component fatigue lives in several Royal Australian Air Force (RAAF) fleet types. In this paper the Lead Crack concept is developed and its strengths and weaknesses are discussed. Examples of Crack growth behaviour that are considered typical and representative of Lead Cracks are presented.

  • The Lead fatigue Crack concept for aircraft structural integrity
    Procedia Engineering, 2010
    Co-Authors: L. Molent, S.a. Barter
    Abstract:

    Abstract Over many years of quantitative fractographic examination of fatigue Cracking from in-service and full-scale fatigue tests of metallic airframe components, it has been consistently observed that the largest Cracks formed have grown in an approximately exponential manner. These Crack growth observations range from the initiation of Cracks and their early growth from a few micrometers through to many millimetres in length. It appears that these Lead Cracks commence growing shortly after the airframe is introduced to the loading environment. Furthermore these Cracks usually initiate from production-induced or, less frequently, inherent material discontinuities. Based on these two observations, an aircraft lifing methodology that is based on the results of fatigue testing programs utilising the Lead Crack concept has been developed and implemented as an additional tool in the determination of aircraft component fatigue lives in several Royal Australian Air Force (RAAF) fleet types. In this paper the Lead Crack concept is developed and its strengths and weaknesses are discussed. Examples of Crack growth behaviour that are considered typical and representative of Lead Cracks are presented.