Subject: DAMAGE TOLERANCE OF HOLE Date: 8/28/09 AC No: 33.70-2 FEATURES IN HIGH-ENERGY TURBINE Initiated by: ANE-111
ENGINE ROTORS
1. Purpose.
a. This advisory circular (AC) provides definitions, guidance, and acceptable methods, but not the only methods, that may be used to demonstrate compliance with requirements in § 33.70 of Title 14 of the Code of Federal Regulations related to the integrity of engine life-limited parts.
Section 33.70 contains requirements applicable to the design and life management of propulsion system life-limited parts, including high-energy rotating parts.
b. This AC presents a damage tolerance approach which can be used to address manufacturing and operationally-induced anomalies in turbine engine rotating part hole features.
This approach can be readily integrated with the existing “safe-life” process for high-energy rotors to produce an enhanced life management process. This approach does not replace existing safe-life methodology but supplements it. In the context of damage tolerance, this AC is not intended to allow operation beyond the component manual life limit set using the existing safelife approach which limits the useful rotor life to the minimum number of flight cycles required to initiate a crack. Rotor failure modes for which full containment of high-energy debris can be demonstrated are excluded from the procedures outlined in this AC and need not be accounted for in the overall risk assessment.
2. Applicability.
a. The guidance provided in this document is directed to engine manufacturers, modifiers, foreign regulatory authorities, part manufacturers who hold Parts Manufacturer Approval (PMA) authority, and Federal Aviation Administration (FAA) designated engineering representatives.
Within this AC the term “engine manufacturer” refers to any person who holds a design approval of an engine life-limited part.
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b. This material is neither mandatory nor regulatory in nature and does not constitute a regulation. It describes acceptable means, but not the only means, for demonstrating compliance with the applicable regulations. We (“the FAA”) will consider other methods an applicant may present to demonstrate compliance. Terms such as “should,” “shall,” “may,” and “must” are used only in the sense of ensuring applicability of this particular method of compliance when the method in this document is used. While these guidelines are not mandatory, they are derived from extensive FAA and industry experience in determining compliance with the relevant regulations. If we find that following this AC would not result in compliance with the applicable regulations, we will not be bound by this AC, and we may require additional substantiation as the basis for finding compliance.
c. This material does not change, create any additional, authorize changes in, or permit deviations from existing regulatory requirements.
3. Related Regulations.
a. Section 33.4, Instructions for Continued Airworthiness.
b. Section 33.15, Materials.
c. Section 33.19, Durability.
d. Section 33.27, Turbine, compressor, fan, and turbo-supercharger rotors.
e. Section 33.63, Vibration.
f. Section 33.70, Engine life-limited parts.
g. Section 33.75, Safety analysis.
4. References and Related Reading.
a. FAA Documents.
(1) AC 33-3, “Turbine and Compressor Rotor Type Certification Substantiation Procedures,” September 9, 1968.
(2) AC 33.14-1, “Damage Tolerance for High Energy Turbine Engine Rotors,” January 8, 2001. Note: AC 33.14-1 will be replaced by AC 33.70-3 since the final rule on Aircraft Engine Standards for Engine Life-limited Parts (72 FR 50856), which published on September 4, 2007, removed § 33.14 from part 33.
(3) AC 33.70-1, “Guidance Material for Aircraft Engine Life-Limited Parts Requirements,” July 31, 2009.
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(4) FAA Report number DOT/FAA/AR-06/3 “Guidelines to Minimize Manufacturing Induced Anomalies in Critical Rotating Parts” (available at www.tc.faa.gov/its/worldpac/techrpt/ar06-3.pdf).
b. Industry Documents.
(1) “Development of Anomaly Distributions for Aircraft Engine Titanium Disk Alloys.” Technical paper presented by the AIA Rotor Integrity Sub-Committee at the American Institute of Aeronautics and Astronautics Conference in April 1997.
(2) “Development of Anomaly Distributions for Machined Holes in Aircraft Engine Rotors.” Technical paper presented at ASME Turbo Expo in Barcelona, Spain in May 2006.
(3) SAE/FAA Committee On Uncontained Turbine Engine Rotor Events, Report No.
AIR 1537, Data Period 1962-1975.
(4) SAE/FAA Committee On Uncontained Turbine Engine Rotor Events, Report No.
AIR 4003, Data Period 1976-1983.
(5) SAE/FAA Committee On Uncontained Turbine Engine Rotor Events, Report No.
SP-1270, Data Period 1984-1989.
5. Definitions. The following definitions apply within this AC:
a. Component Feature. A unique location, structural shap e, or surface of a component or part. For a disk, component features are blade slots, bore or web locations. Other examples are grooves, slots, or holes. Usually different machining processes and practices are required to produce each unique feature.
b. Component Event Rate. The number of events for a given rotor component stage for each engine cycle, averaged over the projected life of the component.
c. Damage Tolerance. An element of the life manage ment process that recognizes the potential existence of component imperfections that are the result of inherent material structure, material processing, component design, manufacturing, or usage. Damage tolerance addresses this situation through the incorporation of fracture resistant design, fracture-mechanics, process control, or nondestructive inspection.
d. Default Probability of Detection (POD) Values. Values representing mean probabilities of detecting anomalies of various types and sizes. These values are accepted as valid provided the inspection is conducted consistent with good industry practice.
e. Design Target Risk (DTR). The relative risk of failure caused by material, manufacturing, and service induced anomalies and the standard against which probabilistic assessment results (stated in terms of feature, component, and engine level event rates) are compared.
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f. Engine Event Rate. The cumulative number of events predicted for a given anomaly type for each engine cycle, for all rotor life-limited parts in a given engine, calculated over the projected life of those components.
g. Event. A rotor structural part separation, fa ilure, or burst, with no assessment of consequence.
h. Focused Inspection. A component inspection in which the inspector has been instructed to pay special attention to specific component feature(s) or location(s) and the necessary specialized processing instructions have been provided.
i. Full Field Inspection. A component inspection without special attention to any specific component feature(s) or location(s).
j. Hard Time Inspection Interval. The number of engine cy cles since new or since the most recent inspection after which a rotor part must be made available and must receive the inspection specified in the Airworthiness Limitations Section (ALS) of the Instructions for Continued Airworthiness (ICA).
k. Hazardous Engine Effects. Any of the effects listed in § 33.75(g)(2). Failure of rotating parts usually results in the non-containment of high-energy debris, which is the primary hazardous engine effect addressed by this AC. Disks, hubs, impellers, large rotating seals, and other similar large rotating components represent potential sources of high-energy debris.
l. Hole Feature. A set of one or more circular hole s at the same axial and radial location where any given hole falls at least partially within the circumferential (diametral) shadow of the largest hole and its size is greater than 75 percent of the diameter of the largest hole. (See Appendix 5).
m. Inspection Opportunity. An occasion when an engine is disassembled to at least the modular level and the hardware in question is accessible for inspection, whether or not the hardware has been reduced to the piece part level.
n. Low Cycle Fatigue (LCF). The process of progressive and permanent local structural deterioration occurring in a material subject to cyclic variations, in stress and strain, of sufficient magnitude and number of repetitions. The process will culminate in detectable crack initiation typically within 105 cycles. A detectable crack initiation is defined as 0.030 inches in length by
0.015 inches in depth.
o. Manufacturing Anomaly. A surface-related imperf ection introduced during the manufacturing process (subsequent to melting) that is considered potentially detrimental to the structural integrity of a life-limited rotating part during its service life.
p. Maintenance Exposure Interval. A statistical distribution which defines the number of flight cycles between engine shop visits. This interval provides an opportunity to inspect the iv
33.70-2 8/28/09 engine internal parts because the module or parts have been disassembled as a result of normal maintenance activity.
q. Mean POD. The 50-percent confidence level POD versus anomaly size curve.
r. Module Available. An individual module removed from the engine.
s. Module. A combination of assemblies, subassemblies, and parts contained in one package, or arranged to be installed in one maintenance action.
t. Operationally-Induced Anomaly. A surface imperfection introduced during service operations (includes flight operations and maintenance) that is potentially detrimental to the structural integrity of the life-limited rotating part.
u. Part Available. A part that can be inspected, as required by the ALS of the ICA, without any further disassembly. Depending on the inspection requirements, some parts may require a fully disassembled “piece part” condition, while other parts may be available for inspection while still in the assembled module.
v. Probabilistic Risk Assessment. A fracture-mechanics base d simulation procedure that uses statistical techniques to mathematically model and combine the influence of two or more variables to estimate the likelihood of various outcomes for a product. Since not all variables may be considered or may not be capable of being accurately quantified, the numerical predictions are used on a comparative basis to evaluate various options. Results from these analyses are typically used for design optimization to meet a predefined target or to conduct parametric studies. This type of procedure differs from an absolute risk analysis, which attempts to consider all significant variables and is used to quantify, on an absolute basis, the predicted number of future events with safety and reliability consequences.
w. Probability of Detection. A quantitative statistical measure for detecting a particular type of anomaly (flaw) over a range of sizes using a specific nondestructive inspection technique under specific conditions. Typically, the mean POD curve is used.
x. Rotor Life-Limited Parts. Engine rotor parts whose primary failure is likely to result in a hazardous engine effect, which is any of the conditions listed in § 33.75.
y. Safe Life. A low cycle fatigue-based process in which life-limited components are designed, manufactured, and substantiated to have a specified service life or life limit, which is stated in operating flight cycles, operating hours, or both. The “safe life approach” requires that parts be removed from service prior to the development of an unsafe condition (i.e., crack initiation). When a component reaches its published life limit, it is retired from service. The safe life approach only applies to parts which define crack initiation as the limit of the useful life such as rotating parts.
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z. Soft Time Inspection Interval. The number of engine cycles since new, or the most recent inspection, after which a rotor part in an available module must receive the inspection specified in the ALS of the ICA.
aa. Stage. The rotor structure that supports and is attached to a singl e aerodynamic blade row.
6. Background.
a. As required by § 33.70(a), applicants must perform appropriate damage tolerance assessments to address the potential for failure from material, manufacturing, and service induced anomalies within the approved life of the part. This is necessary because service experience with gas turbine engines has demonstrated that material, manufacturing and serviceinduced anomalies do occur, and they can potentially degrade the structural integrity of rotor life-limited parts. Several approaches are currently available to address this requirement and have been introduced by AC 33.70-1 and AC 33.14-1. The methodology described in Chapter 2 of this AC was first introduced in AC 33.14-1 to address the potential presence of hard alpha anomalies in titanium rotors.
b. The complex and time-consuming nature of the effort required to develop the technical elements associated with this method—for example, data collection, analysis, methodology development, acceptable design targets, and validation—has led to a decision to introduce the surface damage tolerance assessment methodology for manufacturing and in-service anomalies in phases. Phase 1 of this methodology, introduced in this AC, presents a damage tolerance approach which can be used to address manufacturing and operationally-induced anomalies in circular hole features (see Appendix 5) in rotor parts. We will introduce other phases in future ACs that will address the remaining surface features.
Francis A. Favara Manager, Engine & Propeller Directorate Aircraft Certification Service vi
TABLE OF CONTENTS
CHAPTER 1. ENHANCED LIFE MANAGEMENT PROCESS 1
CHAPTER 2. DAMAGE TOLER ANCE ASSESSMENTS 3
CHAPTER 3. “SOFT TIME INSPECTION” ROTOR LIFE MANAGEMENT 10
APPENDIX 1. MANUFACTURING-I NDUCED ANOMALY CALIBRATION TEST CASE 13
APPENDIX 2. DEFAULT ANOM ALY DISTRIBUTION CURVES 21
APPENDIX 3. DEFAULT POD APPLICABILITY 25
APPENDIX 4. DEFAULT POD CURVES 28
APPENDIX 5. DEFINITION OF A HOLE FOR MANUFACTURING ANOMALY ASSESSMENTS 30
APPENDIX 6. MANUFACTURING CREDITS 31
APPENDIX 7. SURFACE (MANUFACT URING INDUCTED ANOMALY) DAMAGE TOLERANCE ASSESSMENT 34
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CHAPTER 1. ENHANCED LIFE MANAGEMENT PROCESS
1. Based on historical experience, the safe-life philosophy has worked well and provides a solid base for further enhancement. Modifications to the industry standard life management procedures should augment, not replace, current approaches, which are based on the safe-life philosophy.
a. A new element, damage tolerance, has been added to the existing conventional life management process to define the enhanced life management process. We expect enhanced life management to further minimize the occurrence of uncontained rotor failures and thus improve flight safety (see figure 1-1).
b. Use of the enhanced life management process, shown in figure 1-1, will result in damage tolerance assessments being conducted on critical rotating part designs. These will be in the form of fracture-mechanics-based probabilistic risk assessments, the results of which will be compared to specified design target risk values. Designs that satisfy these DTR values will be considered in compliance with the damage tolerance requirements of § 33.70. Engine manufacturers will have a variety of options available to achieve the DTR. These options include, but are not limited to:
• component redesign,
• material changes,
• material process improvements,
• manufacturing process improvements,
• manufacturing inspection improvements,
• enhanced in-service inspections, and
• life limit reductions.
33.70-2 8/28/09 Figure 1-1: Enhanced Life Management Process Safe Life Prediction Life Management Process Development & Validation Tes t s Methods & Material Data Service Life & Product Assurance Damage Tol er an c e Assessment New Element
Added to
Existing Safe
Life Process
33.70-2 8/28/09 CHAPTER 2. DAMAGE TOLERANCE ASSESSMENTS
1. Approach .
a. As described in AC 33.70-1, Probabilistic Damage Tolerance Risk Assessments (PDTRA) is an acceptable method to assess the ability of a part to tolerate anomalies. The results of these assessments will provide the basis for evaluating the relative damage tolerance capabilities of candidate part designs and will also allow the engine manufacturer to balance its designs for both enhanced reliability and customer impact. The results will be compared against design target risk values (see paragraph 6, “Design Target Risk,” in this chapter).
b. The need for further risk reducing actions on the part of the engine manufacturer is based on whether the design under consideration satisfies the desired DTR values. If the targets are met, then the design is in compliance with § 33.70. The manufacturer has a variety of options available to achieve the required DTR and may conduct quantitative parametric studies to determine the influence of key variables such as inspection methods, inspection frequency, hardware geometry, hardware processing, material selection and life limit reduction. The manufacturer may make changes to the design, or field management of the part, or both, to achieve the DTR values.
c. PDTRA will usually be performed during the engine component detail design phase.
Paragraph 2 of this chapter defines an assessment methodology applicable to manufacturing and service-induced anomalies and provides a process to refine a design to achieve the desired DTR values.
2. Methodology.
a. Probabilistic risk assessments may be conducted using a variety of methods, such as Monte Carlo simulation or numerical integration techniques. When performing a PDTRA, use of the standardized inputs and default data presented in this AC will achieve consistent assessment results.
b. A list of standardized inputs is provided under paragraph “c” below. Default input data is described in paragraphs 3 and 4 of this chapter. Use of this default data in probabilistic assessments requires no further demonstration of applicability or accuracy. However, when this data is used, the applicant should follow the guidelines accompanying this data to ensure applicability. Use of inputs other than these default values may require additional validation to verify applicability, adequacy, and accuracy.
c. Probabilistic risk assessments should incorporate the following inputs as part of a basic analysis: component stress, surface area, material properties, crack growth (propagation) life, anomaly distribution, design service life, inspection PODs, and maintenance exposure rate.
(1) Input. The following data sets are defined as basic input elements.
(a) Stress. This variable influences crack propagation life. Data input should encompass the most limiting operational principal stresses. Engine cycle(s) for design LCF certification, or actual usage (if known), should be used to establish the stress distribution.
33.70-2 8/28/09 Surface assessments should incorporate the appropriate surface and near surface stress distributions, including the effects of stress concentrations and vibratory stresses.
Note: The method described in this AC has been calibrated against industry experience without considering the effect of shot peening and the resulting beneficial residual stresses on the crack propagation lives. As a result, with this method it is inappropriate to include the beneficial effects of shot peening.
(b) Surface Area. This variable influences the probability of having a manufacturing-induced anomaly on the surface of a specific component/component feature. For hole features, this corresponds to the total surface area of one hole in a given feature.
(c) Material Cyclic Crack Growth Rate (CCGR) Data. Use average crack growth rate properties of the base material as the default condition to calculate crack propagation life.
(d) Propagation Life. Nu mber of cycles for a given size anomaly to grow to a critical size. This is based on knowledge of part stresses, including the vibratory stresses, temperature, geometry, stress gradient, anomaly orientation, and material properties. Linear elastic fracture mechanics should be used for the calculation of propagation life. Default conditions should assume anomalies to be in the worst orientation to the stress field.
(e) Anomaly Distribution. For manufacturing-induced assessments, use the default anomaly distributions outlined in paragraph 3 of this chapter or FAA-approved companyspecific data. Company-specific data should be developed using the same process employed to derive the default anomaly distributions. For manufacturing-induced anomaly assessments, this process is described in a technical paper titled “Development of Anomaly Distributions for Machined Holes in Aircraft Engine Rotors” - see paragraph b.(2) in “References and Related Reading.” The paper also describes the process required to establish the potential benefits (e.g., reduced anomaly size/frequency distributions) derived from the imposition of enhanced surface inspections during the manufacturing process. Note: In all assessments, anomalies will be treated as sharp propagating cracks from the first stress cycle.
(f) POD. Default PODs and instructio ns on use of individual company values are contained in paragraph 5 of this chapter. For surface assessments, consider the effects of fluorescent penetrant inspection (FPI) and eddy current inspection (ECI) as appropriate.
(g) Maintenance Exposure Interval. For the purpose of assessing inspection benefits, model exposure interval curves for the engine, module, or component in question in the analysis.
(2) Calibration.
(a) Each engine manufacturer should calibrate its analytical prediction tools by conducting the industry test case detailed in Appendix 1.
(b) The manufacturing-induced anomaly test case in Appendix 1 consists of a probabilistic analysis of a simulated titanium ring disk with a row of ½" diameter holes, using specified inputs and scenarios. Test case resu lts in the ranges from 2.58E-04 to 3.22E-04 events 4
33.70-2 8/28/09 within the service life of the part (for the “no inspection” case) and from 7.29E-05 to 2.25E-04 (for the “with in-service inspection” case) are considered acceptable. Test case results outside of these ranges may indicate deficiencies with either the probabilistic assessment technique or the assumptions made.
(3) Output: Manufacturing-Induced Anomaly Assessments. Holes.
(a) Applicability. This pr ocedure applies to circular holes in rotor life-limited parts with the exception of the bore hole. The bore hole is specifically exempted because the machining anomalies observed in field events originating at the bore hole do not resemble the anomaly types observed in drilled hole features, which suggests different damage mechanisms.
(b) Hole Feature Assessments. The probabilistic assessments and prediction of event potential for a given hole feature should be calculated over the entire anticipated service life of the part. This result should be expressed as the number of predicted events for each cycle and should be designated as the predicted “hole feature event rate.” The predicted hole feature event rate should then be compared to the hole feature level DTR value to assess design acceptability.
(4) General Comments.
(a) The FAA encourages engine manufacturers to incorporate fracture resistant design concepts where possible. Credit may be given for fracture resistant or burst resistant engine design features, which clearly demonstrate, through both analysis and test, a reduction in the potential for rotor failure due to the presence of unanticipated anomalies. Credit may be determined using the probabilistic risk assessment methods contained with this AC and by data which supports the fracture resistant features of the design.
(b) Because the design of an aircraft turbine engine rotor is a lengthy process involving numerous iterations, each of which can substantially alter the initial design risk value, it is important that the design feature satisfy the appropriate DTR value at the time of engine certification. Some risk assessments will be conducted several years after the engine enters service due to in-service problems, changes in engine service usage, improvements in analytical predictive tools, or changes in engine component materials. In these instances, it is important to note that the DTR values should be satisfied throughout the life of the hardware.
(c) Etch inspection is widely used in the industry to evaluate material quality on rotor part surfaces. While etch inspections are not yet included in the damage tolerance assessment, they have been identified as the foremost means of detecting non-geometric anomalies (worked layers) in holes by the FAA Report number DOT/FAA/AR-06/3 “Guidelines to Minimize Manufacturing Induced Anomalies in rotor life limited parts” – see paragraph a.(3) in “References and Related Reading.” As such, etch inspections currently in-place perform a vital role and should not be discontinued because of any damage tolerance analysis set forth in this AC.
(d) Crack propagation assessment guidelines for consistent treatment of out of phase stress vs. temperature history are included below. These guidelines apply only to the probabilistic damage tolerance analysis.