Report

FATIGUE AND DAMAGE TOLERANCE ASSESSMENT OF AIRCRAFT STRUCTURE UNDER UNCERTAINTY Lorens S. Goksel 5/1/2013 Committee Members: Dr. Seung-Kyum Choi, Chair Dr. Roger Jiao Dr. David Scott Outline 2 Introduction Research Questions Probability of Failure (PF) Predictable Range of Risk Risk Mitigation Damage Tolerance Risk Assessment (DTRA) Comparison Proposed Framework Validation Example Conclusion Introduction 3 Research Question DTRA Validation Conclusion Purpose Create a tool for an engineer that assesses the life of a component using ‘cradle-to-grave’ approach which includes Manufacturing defects Design loading conditions Component failure mitigation approaches Methodology needs to provide economic solutions Determine how cracks in a component grow with variation of parameters Obtain an optimal range for inspection and refurbishment Introduction 4 Research Question DTRA Validation Conclusion What is Fatigue? Fatigue is the degradation of materials due to repeated loads Degradation occurs due to Mechanically Load induced loads rate Caustic environmental effects Introduction 5 Research Question DTRA Validation Conclusion Examples of Fatigue Degradation Caustic Environment 3X Less Life Sump Tank Lab Air Introduction 6 Research Question DTRA Validation Conclusion Examples of Fatigue Degradation Load Rate Wind only conditions provide more than twice the life compared to Ground-Air-Ground Introduction 7 Research Question DTRA Validation Conclusion Sources of Fatigue? Material in-homogeneity (voids, inclusions, etc.) Damage (scratches, stress concentration) Stress Risers Manufacturing quality is essential for good fatigue life! Introduction 8 Research Question DTRA Validation Conclusion Damage Tolerance/Crack Growth Crack Growth assumes the material has some initial defect Cracks are two dimensional Failure occurs at critical crack length (fracture) Introduction 9 Research Question DTRA Validation Conclusion How is Risk Related to Fatigue? Each time there is an accumulation of damage, the chance of failure becomes a little higher. Failure is considered the last stage of crack growth i.e. Fracture Usually occurs when critical crack length is reached (potentially catastrophic to system) Introduction 10 Research Question DTRA Validation Conclusion Research Question #1 How does one determine the Probability of Failure for aerospace structures? Hypothesis: When the Fatigue loads exceeds the material strength, failure occurs. Probability of this occurrence depends on the occurrence and size of both the load and material strength. Introduction 11 Research Question DTRA Validation Conclusion Research Question #1 Environmental Input Flight Design Case Probability Will see this load level every flight Probability Residual Strength Extreme Rare Occurrence Residual Strength Distributions Strength Strength Interference = Probability of Failure Interference = Probability of Failure Introduction 12 Research Question DTRA Validation Conclusion Research Question #2 How can one predict risk failure based on a crack growing for aerospace systems? Hypothesis: By knowing the material properties, geometry of crack, and all load conditions, and started from the smallest computational crack size. Introduction 13 Research Question DTRA Validation Conclusion Research Question #2 Slow Crack Growth Predictable (Paris) Area Fast Crack Growth Introduction 14 Research Question DTRA Validation Conclusion Research Question #3 How can one mitigate crack growth risk, economically? Hypothesis: By detecting a crack before it reaches a critical length, but during its predictable growth period. Introduction 15 Research Question DTRA Validation Conclusion Summary Need to understand how crack grows in a part with certain parameters Need for a method that can provide an optimal range for inspections All need to account for: Probability of failure Predict risk associated with failure Minimize Failures Damage Tolerance Risk Assessment incorporates all Introduction 16 Research Question DTRA Validation Conclusion Comparison to Other Methods White [64] proposed risk analysis Includes loading history, material properties and flaw size Indicates gradual increase is fast crack growth area Wang [65] performed risk analysis at bolted connection. Approach: at what crack length can one start inspections based on an acceptable risk level Neglects to provide a range of inspection periods Introduction 17 Research Question DTRA Validation Conclusion Comparison to Other Methods Grooteman [64] Equivalent initial flaw size to and probability of detection curves to determine optimum inspection intervals Computationally arduous Cavallini and Lazzeri [65] Probabilistic Investigation for Safe Aircraft (PISA) Accounts for Initial Flaw Size Material Variability Probability of Detection Computation limitation cannot provide risk associated with small cracks Research Question Introduction 18 DTRA Validation Conclusion Proposed Framework Step 1: Specify Geometry, Loading Conditions and Material Statistic Properties (Crack Growth) Obtain Step 2: Discontinuity Check Obtain Residual Strength based on distribution more clear data Step 3: Obtain Probability of Failures, Probability of Detection Setup the DTRA Introduction 19 Research Question DTRA Validation Conclusion Proposed Framework – Step 1 Assume Initial Flaw Size Grow Flaw Until Critical Crack Length Final Crack Initial Flaw Obtain residual strength PDF based on variability of fracture toughness Lays foundation for residual strength distribution needed for PF Introduction 20 Research Question DTRA Validation Conclusion Proposed Framework – Step 2 Phantom Distribution Obtain residual strength PDF based on variability of fracture toughness Yes Discontinuities? No Create ‘Phantom’ Distribution Self-check to increase resolution Residual strength due only to local loading Introduction 21 Research Question DTRA Validation Conclusion Proposed Framework – Step 3 Discontinuities? No Intersect Flight Design with Residual Strength Case RQ #1 Answer Plot Probability of Failures for each crack interval RQ #3 Answer Insert Probability of Detection for crack size Flight Design Introduction DTRA Validation Conclusion Proposed Framework – Setup DTRA Probability of Detection 99% 10-7 Probability of Failure 22 Research Question Probability of Failure at Each Crack Length Conservative Optimal High Risk 10-50 Time Introduction 23 Research Question DTRA Validation Conclusion Summary Assume Initial Flaw Size Grow Flaw Until Critical Crack Length Obtain residual strength PDF based on variability of fracture toughness Discontinuities? No Intersect Flight Design with Residual Strength Case RQ #1 Answer Plot Probability of Failures for each crack interval Insert Probability of Detection for crack size in plot RQ #2 & #3 Answer Optimal Area determined based on DTRA, PD and FAA minimum allowable Yes Create ‘Phantom’ Distribution Introduction 24 Research Question DTRA Validation Conclusion Validation: Engine Nacelle Inlet External Loads (Aerodynamic Loads) When is the optimum time to inspect the nacelle inlet for fatigue cracks? Internal Loading (Engine Noise) Step 1 Introduction 25 Research Question DTRA Validation Conclusion Internal Loads Need to determine most pertinent loading mode: longitudinal vs. circumferential Loading is assumed only to act in hoop direction, thus circumferential natural frequency examined Step 1 Introduction 26 Research Question DTRA Validation Conclusion Internal Loads FEM & Hand Method Engine Specification Internal stresses derived using standard static techniques for hoop load conditions Internal Pressure Step 1 Introduction 27 Research Question DTRA Validation Conclusion Crack Growth Assume manufacturing flaw Flaw is two dimensional Use previous internal loading Determine Residual Strength at some crack length Assume normal material distribution Step 1 Introduction 28 Research Question DTRA Validation Conclusion Crack Growth Critical crack Length This progression only accounts for internal loads Initial crack Length Step 1 Introduction 29 Research Question DTRA Validation Conclusion Determine POF Critical Crack Flight Design Case Failure accounts for internal and external loads Phantom Distribution Each failure accounts for crack growth iteration Failure Region Steps 2 &3 Introduction 30 Research Question DTRA Validation Conclusion Risk Mitigation There is a 90% chance… Each crack length is associated with a flight time …this crack length can be found Step 3 Introduction 31 Research Question DTRA Validation Conclusion Damage Tolerance Risk Assessment FAA Minimum 90% Certainty of Flaw Detection Probabilities of Failure The optimal inspection range Step 3 Introduction 32 Research Question DTRA Validation Conclusion Contributions Single Visual Aid that accounts for Manufacturing defects as initial flaw size from processes (machining, castings) Material strength variability (fracture toughness assumed to conform under statistical distribution) Aircraft maneuver variability (Passenger vs. fighter jet, extreme value distribution) Flaw detection resolution (Type of material, minimum desired crack detection size, non-destructive techniques) Introduction 33 Research Question DTRA Validation Conclusion Further Research Further Research Account for bulging effects (crack growth more arduous under cylindrical shape) Hammershock Condition (backpressure pulse results in shock during supersonic flight) Statistical range of initial flaws Acknowledgements 34 Advisor: Dr. Seung-Kyum Choi Reading Committee: Dr. Roger Jiao Dr. David Scott Funding: Gulfstream Aerospace References 35 http://www.tamarackhti.com/tools/FADT_capabilities.asp http://avstop.com/maint/corrosion/ch5.html http://www.vgblogger.com/tom-clancys-hawx-briefing-extrememaneuvers-and-enhanced-reality-system-explained/4329/ http://matdl.org/failurecases/images/thumb/9/91/SchenectadyShi p.png/500px-SchenectadyShip.png http://pressurevesseltech.asmedigitalcollection.asme.org/data/Jour nals/JPVTAS/926532/pvt_134_6_061213_f002.png http://wwwold.me.gatech.edu/jonathan.colton/me4210/castdefect.pdf Fatigue and Damage Tolerance Assessment of Aircraft Structure Under Uncertainty, Goksel, L