Use this URL to cite or link to this record in EThOS: https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.794810
Title: Extending fatigue life of aircraft fuselage structures using laser-peening
Author: Busse, David Osman
ISNI:       0000 0004 8501 0979
Awarding Body: Cranfield University
Current Institution: Cranfield University
Date of Award: 2017
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Abstract:
Fatigue of airframe structures is a constant challenge to aircraft manufacturers when designing, maintaining and repairing new and aging metallic components. Laser-Peening (LP) is a highly flexible and controllable surface treatment and relatively new to manufacturers of large civil aircraft which demonstrated that it can extend the fatigue and crack growth life in aluminium alloys by introducing deep compressive Residual Stresses (RS). Currently there is no application of LP to any components of large civil aircraft. The aim of this research was to demonstrate and explore different LP strategies that can produce significant extension of the fatigue and crack growth performance of aircraft fuselage structures using Laser-Peening. Two representative samples made from 2000 series aluminium alloy were designed to represent features of the fuselage: A Centre Cracked Tension (CCT) panel made of 1.6 mm thick 2524-T3 represented the fuselage skin. Single overlap Lap-Joints (LJ) of 2.5 mm thick 2024-T3 aluminium with titanium Hi-Lok bolts arrayed in 5 columns and 3 rows embodied longitudinal LJ of aircraft fuselages. Both test samples were laser-peened without protective coating (LPwC) using a range of LP strategies in which LP process parameters and spatial arrangements of laser-peened areas were systematically varied. RS fields were measured before fatigue testing under constant amplitude loading. RS measurements used Incremental Centre Hole Drilling (ICHD) and X-ray and Neutron diffraction techniques. Laser-peening produced peak compressive RS of 200 - 350 MPa and compression stress penetration depths between 700-1000 μm. These values are superior to RS profiles induced by Shot-Peening. The value of peak compression stress and penetration depth depends on LP process parameters and on the LP layout. The latter defines the location and size of the laser-peened areas. A study of the effect of different LP strategies to establish the most effective LP treatment to enhance crack growth life of fuselage skins was performed using a Finite Element based crack growth model. The model was first used to introduce balanced RS fields into a cracked CCT sample. The effective stress intensity factor range (ΔKeff) and effective R-ratios (Reff) were then calculated as the crack tip progressed through the sample. Subsequently, fatigue crack growth rates and lives were computed using Walker's empirical crack growth law. The accuracy of the model was demonstrated by comparison with crack growth test results from laser-peened CCT-samples. Results of the parameter study showed that an increase in the level of compression within the LPS increased life most significantly. Increased width of peen stripe increased the life while increasing the distance of the stripe from the starting position of the crack tip decreased the life. Four different LP strategies were applied to LJ samples. Subsequent fatigue testing demonstrated fatigue life improvements of between 1.14 to 3.54, depending on the LP strategy. The LP layout was identified as a key parameter determining the fatigue life. It was found that when small LP areas were used, to leave as much elastic material as possible between the peened areas, larger compressive stresses and minimised balancing tensile stresses were produced. Observations of fatigue fractures on joint samples showed that crack initiation occurred remote from the fastener holes, either in regions of fretting fatigue in peened areas or in regions of balancing tensile stress adjacent to peen boundaries. Optimum fatigue lives occurred when both fracture types occurred in the same sample. Striation spacing measurement and analysis showed that compressive residual stresses had little or no effect on fatigue growth rates at crack lengths < 600 µm. The majority of fatigue life extension was achieved during initiation and crack growth < 600 μm. The obtained results established evidence of how aircraft fuselage structures made of conventional 2000 series aluminium-copper alloys can be effectively laser-peened to produced fatigue life improvements and also of how to avoid any detrimental reductions in fatigue life which can also occur when LP is applied randomly. The generated research conclusions are applicable to other metals, geometries and components.
Supervisor: Irving, Phil E. ; Ganguly, Supriyo Sponsor: Not available
Qualification Name: Thesis (Ph.D.) Qualification Level: Doctoral
EThOS ID: uk.bl.ethos.794810  DOI: Not available
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