Correlation of 3D fatigue crack growth in residual stress bearing materials

Transcription

1 Hill Engineering, LLC Solutions Engineering for structural Aircraft Structures integrity Correlation of 3D fatigue crack growth in residual stress bearing materials 2013 AFGROW Workshop Sep 10-11, Layton, UT Michael R. Hill, Daniel Stuart University of California, Davis Adrian DeWald, John VanDalen Hill Engineering, LLC Collaborator: Jeff Bunch The Boeing Company Sponsors: Stephanie Flanagan, Kristina Langer Air Force Research Laboratory

2 Acknowledgement of earlier presentation 2012 ASIP Conference, San Antonio Hill Engineering, LLC Solutions Engineering for structural Aircraft Structures integrity Design and analysis of engineered residual stress surface treatments for enhancement of aircraft structure Michael R. Hill, Adrian DeWald, John VanDalen Hill Engineering, LLC Jeff Bunch The Boeing Company Stephanie Flanagan, Kristina Langer Air Force Research Laboratory DISTRIBUTION STATEMENT A. Approved for public release; distribution is unlimited. ASIP 2012 San Antonio, TX November 27-29, Hill Engineering, LLC 2

3 Outline New residual stress engineering methods Residual stress measurement Residual stress modeling Use of these methods in F-22 laser shock peening (LSP) program Residual stress measurement Residual stress modeling Crack growth prediction for residual stress bearing structure Correlation of predicted crack growth with test F-22 DO-30 test articles Lug element crack growth tests Frame crack growth tests Application to aluminum cold expanded holes Initial validation work Summary 2013 Hill Engineering, LLC 3

4 Innovations and conditions leading to progress Processes that provide deep residual stress Cold expanded holes Laser shock peening Others (LPB, shot peening, water jet peening, deep rolling, ) New residual stress engineering methods Contour method measurements reliable 2D stress mapping Process modeling rapid design calculations for 3D residual stress New interface to crack growth analysis software Accurate residual stress intensity factors Arbitrary growth of cracks in three dimensions Right conditions for technological progress Sustained DoD investment in and support for technology development (USAF, NAVAIR, FAA) SPO and OEM engagement in solving sustainment issues 2013 Hill Engineering, LLC 4

5 Background (Polin, et al, 2011 ASIP) Polin, Bunch, Caruso, McClure, Full Scale Component Tests to Validate the Effects of Laser Shock Peening, 2011 ASIP Conference F-22 LSP Validation Testing Laser Shock Peening (LSP) is a surface treatment that imparts surface compressive residual stresses to improve the fatigue life of metallic components widely used on aircraft engine turbine blades F-22 has implemented LSP as a structural retrofit on the wingattach lugs for crack initiation benefit Fracture critical wing carry through structure Multi-phase test program defined benefit on lug elements & full scale frames using factory processes Frame 5 Lug Frame 5 Frame 2 WELDED FORWARD BOOM LSP Area 5

6 New residual stress engineering method: Contour Contour method generates a 2D map of residual stress normal to a plane M.B. Prime (2001) Contour method steps (illustrated for 2D body) Part contains unknown RS (a) Cut part in two: stress release deformation (b) Measure deformation of cut surfaces Apply reverse of average deformation to finite element model of body (c) Map of RS normal to surface determined Same procedure holds for 3D Cut measure FEM 2D residual stress map 2013 Hill Engineering, LLC 6

7 Contour method example Cut the part Wire EDM typical Clamp part rigidly Measure surface deformation CMM or laser scanner typical Measure a grid of points on both cut surfaces Analyze experimental data Filter out noise Average data from both surfaces Compute residual stress FEA model Displacement boundary condition EDM cut applied to FEA Model y Surface measure Measured surface deformation 2013 Hill Engineering, LLC 7 x

8 Contour application for F-22 LSP program Contour Plane Rectangular plates and beams ~1 LSP area ~20 ~3 Geometry coupon Frames ~2 Lug elements 2013 Hill Engineering, LLC 8

9 Use of contour on F-22 LSP program Contour Plane Rectangular plates and beams with chamfered corners ~1 LSP area Large numbers of test articles (~ 20) Quantitative outcomes: Stress at corners Effects of process parameters Residual Stress 0 LSP surface Residual stress Process 1 Process 2 Process 4 Depth from corner 2013 Hill Engineering, LLC 9

10 Use of contour on F-22 LSP program Geometry coupon Small number of test articles (~ 5) Quantitative outcomes: Stress at corners with intersecting radii Effects of LSP intensity LSP applied in radiused channel Residual Stress 0 Higher intensity LSP surface Lower intensity LSP surface 2013 Hill Engineering, LLC 10

11 Use of contour on F-22 LSP program Lug elements Small number of test articles (~ 1) Quantitative outcomes: Residual stress in fatigue test articles Geometric similarity to airframe LSP surface Residual Stress 0 ~ Hill Engineering, LLC 11

12 Use of contour on F-22 LSP program Frames Single test article (Frame 5) Quantitative outcome: Stress in airframe ~20 LSP surface Residual Stress Hill Engineering, LLC 12

13 Contour results: RS changes with geometry Aircraft Frames Lug elements Residual stress Beams Lug elements Frames Geometry coupon Rectangular plates and beams Depth from corner Residual stress engineering requires transferability with geometric scale 2013 Hill Engineering, LLC 13

14 New RS engineering tool: ERS-Toolbox Unique semi-empirical approach Based on concept of eigenstrain (ε*) ε* is function of process parameters and material ε* is determined experimentally (from contour data) Build library of ε* distributions (ERS-Toolbox ) Motivation: rapid calculations for design Integrates with a commercial finite element package(s) Process spec: Surface treatment type Process parameters Processed area Model application steps Create FE mesh Specify surface treatment Determine corresponding ε* (software tool) Solve for equilibrium stress state Model outputs full field RS and distortion The ERS-Toolbox provides geometric transferability Output: Residual stress and distortion ERS-Toolbox FE solver Hill Engineering, LLC 14

15 Use of ERS-Toolbox on F-22 LSP program Predicted residual stress in range of test articles Solid model Model fidelity improved with measurement results Initial predictions based on flat-plate measurements Blind predictions of Geometry coupon, Lug element Comparison of measurement and prediction data enables model updates Model outputs support sound engineering decisions Manage tradeoffs between more compressive stress (generally good) versus more tensile stress and distortion (generally bad) High vs low intensity Large vs small processed area Process spec: LSP treatment LSP parameters Processed area Manage locations and magnitude of tensile stress FE mesh Output: Residual stress and distortion ERS-Toolbox 2013 Hill Engineering, LLC 15

16 Example residual stress prediction Geometry coupon Comparison of results on single plane Measured Model result is full field Model output, σ min Model output Residual Stress Hill Engineering, LLC 16

17 Residual stress predictions, all LSP frames s Frame 5 Frame 4 Frame 3 Frame 2 Residual Stress Frame 5 0 Frame 2 Frame 3 Frame Hill Engineering, LLC 17

18 New interface to fatigue crack growth software Commercial fracture mechanics code Arbitrary crack shape 3D crack shape evolution Analysis inputs Applied stress spectrum (FE model) Provided by Boeing F-22 Residual stress field (FE model, ERS-Toolbox ) Linear superposition of applied and RS Material model Initial flaw OEM Design info Applied stress OEM component design Solid model FE mesh ERS-Toolbox Residual stress Predict performance Material model Initial flaw 2013 Hill Engineering, LLC 18

19 Predicted crack growth behavior in Frames Growth occurs on same plane in Baseline and LSP models Baseline prediction: planar, roughly quarter-elliptical shape LSP prediction: planar, bulging shape Similar behavior for all Frames Predicted crack shape evolution Baseline Observed crack shape for LSP (Frame 2 test article) LSP Dashed line shows extent of fatigue crack growth From: Polin et. al, 2011 ASIP conference 2013 Hill Engineering, LLC 19

20 Lug element correlation (test data, predictions) Computed mean crack size from a and c (i.e., (c a) 0.5 ) Life normalized so all have same starting size equivalent to a o = c o =0.100 RED = Baseline BLUE = LSP over crack Mean crack size (c a) 0.5 (inch) Lug Element Normalized Life LE-1 LE-2 LE-10 LE-13 LE-14 LE-20 Calculated LE-66 LE-68 Calculated a o = c o = 0.100" 2013 Hill Engineering, LLC 20

21 Lug element: Comparison with basic analysis Typical handbook type of analysis Quarter elliptic shape, stress intensity factor assumptions Baseline: very similar to current analysis LSP: significant over-prediction (RS has large effect at a and c ) Mean crack size (c a) 0.5 (inch) Baseline Lug Element Current analysis (Natural crack shape) LSP over crack Basic analysis (Quarter-ellipse) Normalized Life LE-1 LE-2 LE-10 LE-13 LE-14 LE-20 QE WF Calculated LE-66 LE-68 QE WF Calculated a o = c o = 0.100" LSP Baseline 2013 Hill Engineering, LLC 21

22 Frame 2 correlation Limited test data RED = Baseline BLUE = LSP over crack Mean crack size (c a) 0.5 (inch) Frame 2 Normalized Life Coupon 7A Calculated Coupon 1BR Calculated a o = c o = 0.100" 2013 Hill Engineering, LLC 22

23 Frame 4 correlation RED = Baseline BLUE = LSP over crack Mean crack size (c a) 0.5 (inch) Frame 4 Normalized Life Coupon 7B1 Coupon 7B2 Coupon 5D1 Coupon 5D2 Coupon 4C Calculated Coupon 4A Coupon 4B Calculated a o = c o = 0.100" 2013 Hill Engineering, LLC 23

24 Predicted LSP life improvement factors RED = Baseline a o = c o = 0.050" BLUE = LSP over crack Mean crack size (c a) 0.5 (inch) Frame 2 LIF = 24X Normalized Life Mean crack size (c a) 0.5 (inch) LIF = 6X Frame 3 Normalized Life Mean crack size (c a) 0.5 (inch) Frame 4 LIF = 6X Normalized Life Mean crack size (c a) 0.5 (inch) Frame 5 For LSP no growth from 0.050" flaw LIF very large Normalized Life 2013 Hill Engineering, LLC 24

25 Adapt approach to crack growth from CX holes To do Extend ERS-Toolbox to cold expanded holes Validate against fatigue crack growth data Exploratory work conducted with support from A-10 ASIP under Modernization III Measured residual stress in range of CX conditions Initial validation against fatigue crack growth data (data developed earlier, under FAA funding) Follow-on work with A-10 and T-38 being sponsored by AFRL RS measurements in retired flight hardware Statistically significant populations of data Measurements in straight holes and countersink 2013 Hill Engineering, LLC 25

26 Residual stress measurements in CX holes Range of material Range of hole size Range of interference Range of edge margin 0.25 hole 0.50 hole 0.75 hole Repeated measurements (statistical bounds) Incorporated data into ERS-Toolbox ERS-toolbox Residual Stress Hill Engineering, LLC 26

27 Observations on initial CX crack growth predictions Good correlation in as-machined coupons No growth in CX coupons software update coming Continue analysis if some points on crack front arrest Update coming later this year Used knock-down on residual stress to get growth ~ 90% of measured residual stress Reasonable correlation, but short life (not enough residual stress) 2013 Hill Engineering, LLC 27

28 As-Machined - Crack shape evolution Free vs QE Constrained crack shape 2013 Hill Engineering, LLC 28

29 As-Machined Effect of initial shape QE vs. Triangular starting flaw Photo from: Clark and Johnson, IJF 25(2), T7451, 0.25 plate, AM 2013 Hill Engineering, LLC 29

30 AM R=0.1 validation: Crack growth history NASGRO BEASY, TRI β R BEASY, TRI, β R BEASY, QE 2013 Hill Engineering, LLC 30

31 Crack shape evolution AM vs. CX CX prediction uses 90% of measured RS Full RS leaves crack pinned at bore; awaiting software update Photo from: Clark and Johnson, IJF 25(2), T7451, 0.25 plate, 4% CX 2013 Hill Engineering, LLC 31

32 Summary and conclusions Described application of residual stress engineering tools Residual stress measurement Residual stress modeling Crack growth prediction Described prediction outputs for F-22 LSP program Predicted fatigue crack growth consistent with DO-30 test data Crack shape evolution Crack growth history All F-22 frames predicted to have significant LSP life improvement 6X to >25X for starting flaw Described extension to CX holes, work continuing Demonstrated progress on validation of residual stress engineering Is validation complete? No! Should we keep pressing forward with validation? Yes! Goal: capability available for use in solving sustainment issues available to support OEMs, ASIP managers, and analysis staff 2013 Hill Engineering, LLC 32

33 Acknowledgements Funding for this work from AFRL Decade of UC Davis work on laser shock peening (1999 to 2009) Collaboration with LLNL and MIC Collaborations with F-22 SPO and OEMs SPO: John McClure Boeing: Co-author, Jim Pillers, Bob Franz Lockheed-Martin: Tom Brussat, Dale Ball US DoD SBIR Programs Design Tools for Fatigue Life Prediction in Surface Treated Aerospace Components, 2004 to Present Design/Life Prediction Tools for Aircraft Structural Components with Engineered Residual Stresses, 2008 NAVAIR: Ravi Ravindranath AFRL: Mike Shepard, Pam Kobryn, Rollie Dutton, Stephanie Flanagan, Kristina Langer Pratt & Whitney: Bob Morris, Dave Murphy Collaboration with A-10 and T-38 Programs Mark Thomsen, Bob Pilarczyk, Tim Allred, Mike Blinn FAA Rotorcraft Damage Tolerance Program FAA: John Bakuckas; MS State: Jim Newman, Jr.; Sikorsky: Mike Urban 2013 Hill Engineering, LLC 33

34 Historical Residual Stress Engineering applied as a band-aid to repair design deficiencies Emerging Residual Stress Engineering is a conventional technology that assures performance Thank You 3035 Prospect Park Drive, Suite 180 Rancho Cordova, CA (916)

35 Abstract (to be removed) The presentation focuses on correlation of 3D fatigue crack growth in residual stress bearing aircraft structural details, comparing test data against predictions based on linear elastic fracture mechanics. The main focus of the presentation is a review of recent work regarding laser shock peened titanium wing attach lugs. Test information includes data on residual stress present in test articles in addition to more typical information (test article geometry, load spectrum, etc.). Test data are reported as fatigue crack growth as a function of applied cycles. An initial calculation of full field residual stress in test articles due to laser shock peening was developed using finite element analysis (ABAQUS and ERSToolbox ). Crack growth from an initial flaw, due to applied cyclic and residual stresses, was predicted using boundary element analysis (BEASY ), by repeatedly calculating stress intensity factors and incrementally extending the crack. The methods allow for arbitrary crack front evolution in 3D. Comparisons of observed and predicted crack growth behavior provide useful insight. The final element of the presentation describes current work to extend the modeling capability to highstrength aluminum lower wing areas, with fatigue crack growth from cold expanded holes Hill Engineering, LLC 35