U.S. Department of Transportation Federal Aviation Administration Advisory Circular Subject: GUIDANCE MATERIAL
FOR AIRCRAFT ENGINE LIFE-
LIMITED PARTS REQUIREMENTS Date: 02/24/17 Initiated By: ANE-111 AC No: AC 33.70-1 Change: 1
1. Purpose. This advisory circular (AC) change clarifies the Interim Surface Damage Tolerance and Service Damage Monitoring requirements. The AC provides guidance for demonstrating compliance with the engine life-limited parts integrity requirements of Title 14, Code of Federal Regulations (14 CFR) 33.70.
2. Principal Changes.
a. Paragraphs 4.a – 4.c, 8.d.(7)(d), 8.d.(7)(e), 8.d.(7)(e)1, 8.d.(7)(e)2, 3, and 4, 8.e.(1),
8.e.(2), and 11.a. are changed.
b. The AC change num ber and the date of the change are shown at the top of each applicable page. The change bar in the right or left margin indicates where the change is located.
The changes described may shift the original text.
Page Control Chart Removed Pages Dated Inserted Pages Dated 2 7/31/09 2 02/24/17 15 7/31/09 15 02/24/17 16 7/31/09 16 02/24/17 22 7/31/09 22 02/24/17
3. Website Availability. T o access this AC electronically, go to the AC library at http://www.faa.gov/regulations_policies/advisory_circulars /.
Robert Ganley Acting Manager, Engine & Propeller Directorate, ANE-100 Aircraft Certification Service
Subject: GUIDANCE MATERIAL FOR
AIRCRAFT ENGINE LIFE-LIMITED PARTS
REQUIREMENTS Date: 7/31/09 Initiated By: ANE-111 AC No: 33.70-1
1. PURPOSE. This advis ory circular (AC) provi des definitions, guidance, and acceptable methods, but not the only methods, that may be used to demonstrate compliance with the engine life-limited parts integrity requirements of § 33.70 of Title 14 of the Code of Federal Regulations (14 CFR part 33). Section 33.70 contains requ irements applicable to the design and life management of propulsion system life-limited parts including high-energy rotating parts.
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.
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. The FAA will consider other methods of demonstrating compliance that an applicant may elect to present. 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 acceptable method of compliance 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. On the other hand, if the FAA becomes aware of circumstances that convince us that following this AC would not result in compliance with the applicable regulations, we will not be bound by the terms of 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 33.75, Safety analysis.
4. RELATED GUIDANCE.
a. AC 33.4-1, Instructions for Continued Airworthiness.
b. AC 33.4-2, Instructions for Continued Airworthiness: In-Service Inspection of Safety Critical Turbine Engine Parts at Piece-Part Opportunity.
c. AC 33.14-1, Damage Tolerance for High Energy Turbine Engine Rotors (Amdt. 33-10, Eff. 3/26/84).
5. DEFINITIONS. For the purpose of this AC, the following definitions apply:
a. Approved life. The mandatory replacement life of a part that is approved by the Administrator and listed in the Airworthiness Limitation Section (ALS) of the Instructions for Continued Airworthiness (ICA).
b. Attributes. Inherent characteris tics of a finished part that determines its capability to achieve the approved life without failure.
c. Damage tolerance. An element of the lif e management process that recognizes the potential existence of component imperfections, which 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, and nondestructive inspection.
d. Design Target Risk (DTR). The relative risk of a failure caused by material, manufacturing or service induced anomalies and the standard against which probabilistic assessment results (stated in terms of component event rates and/or engine level event rates) are 2
compared. 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 with the same level of inputs. 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 the predicted number of future events with safety and reliability ramifications.
e. Engine life-limited parts. Engine rotor and major static structural parts whose primary failure is likely to result in a hazardous engine effect. For the purposes of § 33.70, a hazardous engine effect is any of the conditions listed in § 33.75.
f. Engine flight cycle. The flight profil e or combination of profiles on which the approved life is based.
g. Engineering plan. A plan that includes the assumptions, technical data and actions required to establish and maintain the life capability of an engine life-limited part. The engineering plan is established and executed as part of the pre and post-certification activities.
h. Failure. Separation of the part into two or mo re pieces or into a condition so that it is no longer whole or complete. Examples of failures include disk or case bursts.
i. Fixed process. A manufacturing method used to produce a part that, once established, cannot be changed without engineering approval and alteration of the engineering plan. A fixed process ensures that a part is produced by a consistent, stable and repeatable process.
j. Feature. A unique location, stru ctural shape, or surface of a component or part. For a disk, a feature would be blade slots, bore or web locations. Other examples are grooves, slots, or holes. Each unique feature is usually produced using a different primary machining process and practice.
k. ICA. Instructions for Continued Airworthiness as required by § 33.4.
l. Life limit. An operational se rvice exposure limit characterized by the application of a finite number of flights or flight cycles. For rotating parts, it is equal to the minimum number of flight cycles required to initiate a crack equal to approximately 0.030 inches in length by 0.015 inches in depth. For life-limited pressure-loaded static parts, the life limit may be based on the crack initiation life plus a portion of the residual crack growth life.
m. Life management. A series of interrela ted engineering, manufacturing, and service support activities that ensure that life-limited engine parts are removed from service prior to the development of a hazardous condition.
n. Likely to result. Given that the part has failed, regardless of the probability of occurrence, what are the possible consequences on the engine and/or aircraft? The word “possible” used in the context of this definition, means it can occur but does not include consequences which are so remote that they are, for practical purposes, impossible.
o. 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 a detectable crack initiation typically within 10E+05 cycles. A detectable crack initiation is defined as 0.030 inches in length by 0.015 inches in depth.
p. Manufacturing plan. A plan that identifies the part specific manufacturing process constraints which must be included in the manufacturing definition (drawings, procedures, specifications, etc.) necessary to consistently produce each engine life-limited part with the attributes required by the engineering plan.
q. Primary failure. Failure of a part that is not the result of prior failure of another part or system.
r. Repairable limits. Damage that can be repa ired so the part retains its current approved life limit.
s. Residual crack growth life. The number of cycles required for a crack of a specified starting size to grow to failure.
t. Safe life. A cyclic fatigue based process in which life-limited components are designed, manufactured, substantiated, and maintained 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 (that is, crack initiation). The safe life approach only applies to parts which define crack initiation as the limit of the useful life, such as rotating parts.
u. Serviceable limits. Damage that can be to lerated without impacting the approved life of the part.
v. Service management plan. A plan that de fines the in-service maintenance processes and the limitations to repair associated with each engine life-limited part such that the part will maintain the attributes required by the engineering plan. These processes and limitations become part of the ICA.
6. INTRODUCTION.
a. Since the failure of an engine life-limited part is likely to result in a hazardous engine effect as defined in § 33.75, applicants should meet specific integrity requirements by executing a series of life management activities. The life management requirements, as defined in § 33.70, necessitate the development and execution of an engineering plan, a manufacturing plan, and a service management plan for each life-limited part. These three plans form a closed-loop system that links the assumptions made in the engineering plan, to how the part is manufactured, and to how the part is maintained in service. Engineering, manufacturing, and service management must function as an integrated system and recognize the effects of actions in one area on the entire system.
b. The engineering plan defines the assumptions, technical data, and actions required to establish and maintain the life capability of each part. The engineering plan is established prior to the introduction of the part into service.
c. To produce parts with the characteristics required by the engineering plan, the applicant should have a consistent and repeatable manufacturing method. The manufacturing plan highlights the parameters that are significant to attaining the life of the part and that should not be changed without verification and engineering approval. The parameters generally involve the manufacturing process steps, controls, and constraints, such as the drawings, procedures, specifications, and machining instructions required to produce and inspect a part using a fixed process to meet the design intent as defined by the engineering plan.
d. The service management plan ensures that the operational service assumptions and life determined by the method in the engineering plan remain valid. It also defines the part limitations associated with service maintenance, overhaul, and repair. These service limitations are conveyed to the maintenance facilities through the ICA. The service management plan applies a comparable level of control to the service aspects of part life management as the manufacturing plan does for the manufacturing processes.
7. GENERAL.
a. Life System Approval.
(1) For an applicant to use a life system, the FAA must first approve the system. The applicant should submit a formal written document to the FAA for approval that identifies the techniques and controls to be used to establish and maintain the life limits. This document will remain on file with the FAA. Proposed changes to life system content must be submitted to the FAA for approval. Modifications mu st be approved prior to use.
(2) The FAA will evaluate the applicant’s technical approach, process controls, experience, and conformance to accepted practices. Based on this evaluation, the FAA will assign a “safety factor” which may reduce the approved life-limits. Safety factors can range from 1.0 (no life reduction) to a substantial reduction.
b. Certification Process Overview. The applicant determines the safe life based on the approved techniques on file with the FAA. The FAA will approve, if appropriate, an approved life that may be less than the safe life. A list of life-limited parts and the approved life for each part must be included in the ALS of the ICA as required by § 33.4. The approved life is the part’s mandatory retirement life.
c. Identification of Engine Life-Limited Parts.
(1) Engine life-limited parts are those engine rotating and major static structural parts whose primary failure is likely to result in a hazardous engine effect. A hazardous engine effect is any of the conditions listed in § 33.75. Engi ne life-limited rotating parts usually include, but may not be limited to, disks, spools, spacers, hubs, and shafts. Static structural parts usually include, but may not be limited to, high-pressure cases and non-redundant engine mount components. To ensure a hazardous engine effect does not occur, an operating limitation or life 5
limit must be established for each engine life-limited part. The operating limitation or life limit should specify the maximum allowable number of flight cycles when the part must be removed from service. Identification of engine life-limited parts per § 33.70 is focused on parts whose primary damage mechanism is controlled by low cycle fatigue.
(2) If a part is made of various sub-parts that are finally integrated in an inseparable manner into a unique part, and any one of the sub-parts is identified as an engine life-limited part, then the entire part is treated as an engine life-limited part.
d. Attributes of a Part. Attributes include, but are not limited to: size, shape, material mechanical properties, material microstructure, material anomalies, residual stress, surface condition, and geometric tolerances. Processes such as alloy melting practice, ingot conversion to billet or bar, forging, casting, machining, welding, coating, shot peening, finishing, assembly, inspection, storage, repair, maintenance, overhaul and handling may influence the attributes of the finished part. Environmental conditions experienced in service may also affect the attributes.
e. Content of a Plan. The engineering, manufacturing, and service management plans should provide clear information for the management of the engine life-limited part. “Plan” in the context of § 33.70 does not necessarily mean that all the required technical information must be contained in a single document. When relevant information exists elsewhere, the plan may reference, as appropriate, documents such as drawings, material specifications, and process specifications. These references should be clear, uniquely identify the referenced document, and allow the history of the individual part number to be traced. The referenced documents should be available for examination if required by the appropriate authority.
8. GUIDANCE FOR DEFINING AN ENGINEERING PLAN.
a. Introduction. The engineering plan consists of a set of comprehensive life assessment processes and technologies that ensure each engine life-limited part can be removed from service before hazardous engine effects can occur. These processes and technologies address the design, test validation, and certification requirements. The plan defines those manufacturing and field management processes and attributes that must be controlled to ensure the established life is achieved and maintained during service deployment.
b. Elements of the Engineering Plan. The e ngineering plan should address the following subjects:
(1) Analytical and empirical engineering processes used to determine the safe life.
These processes include:
The projected or actual aircraft flight profiles used to establish the safe life.
The safe life should reflect service usage profiles associated with specific applications and account for environmental factors.
The mechanical and aerodynamic loads that will be applied to the part.
Thermal and structural analyses.
(2) Structured specimen, component and engine development, and certification testing to confirm the part operating conditions and to enhance confidence in the safe life.
(3) Identification of the attributes which must be provided during manufacture of the part and maintained during service operation. Any in-service inspections identified as critical elements to the overall part integrity should be incorporated into the service management plan.
c. Establishment of the Part Life. Determining the life capability of an engine life-limited part involves the consideration of many factors, each of which may have a significant influence on the final results. There can also be considerab le variation in the methods used to establish the life limits for engine parts, many of which provide acceptable results. Paragraphs 8.d. and 8.e of this AC outline the essential elements of a system which are required to determine satisfactory life limits.
d. Establishment of Safe Life for Rotating Parts. The following figure illustrates a representative process used to establish the safe life for rotating parts:
Figure 1. Representative Process to Establish Approved Life Limits.
Performance Programs
METHODS &
MATERIALS DATA
LIFE ESTIMATION
DEVELOPMENT AND
VALIDATION TESTS
SERVICE LIFE AND
PRODUCT ASSURANCE Analysis Programs Secondary Air Flow programs Analytical Heat Transfer Programs Experimental Stress & Vibration Residual Stress, Aero-elasticity Material & Laboratory Data, Static Rig Tests to Understand Behavior Damage Tolerance Assessments Flight Profiles Selection Engine performance Basic Design Data – Primary Air System, RPM, Temperature, Pressure, Mass Flow Internal, Secondary Air System Data: Flow, Static Pressure Loads, Temperature Component Transient Temperature Estimations Stress & Vibration Analysis Life Estimation:
Routine, Special Circumstances Approved Life Certification Analytical Stress & Vibration Programs Aircraft & Engine Requirements Performance Measurements Secondary Air System Flow, Pressure Ratio, Temperature Measurements Component Temperature & Internal Air System Measurements Vibratory Stress Measurements Cyclic Rig Tests Cyclic Engine Tests Changes in Requirements Flight Profile Monitoring Design Changes
PRODUCT
ASSURANCE
PROCEDURES
APPLY TO ALL
ACTIVITIES Service Damage Monitoring Manufacturing Changes Inspection, Natural Occurrences, Planned Component & Engine Time Expired Parts, Special Circumstances 7/31/09 AC 33.70-1
(1) Flight Profiles.
(a) To establish a safe life, the applicant needs to establish an appropriate flight profile (or combination of profiles) and consider the expected range of ambient conditions and operational variations to determine the service environment. The engine flight cycles should include the various flight segments that describe a complete flight (or flights). For example, for fixed wing aircraft applications this may include segments such as start, idle, taxi, takeoff, climb, cruise, approach, landing, thrust reverse and shutdown. The power requirements for each flight segment are based on the power required by the aircraft application(s). The hold times at the various flight segments should correspond to the limiting installation variables (aircraft weight, climb rates, etc.). Prior to service, the applicant should predict the flight profiles and the corresponding flight segments based on anticipated operator usage, and require coordination with the aircraft manufacturer and potential operators. During service, actual flight profiles may be recorded and may differ from those originally predicted which may alter or reduce the safe life.
Profiles should be conservative and based on realistic use of full rated power. When the engine manufacturer cannot obtain adequate knowledge of the operation of the aircraft, especially for older aircraft, the flight profile(s) should be based on conservative projections.
(b) The applicant should validate and maintain the accuracy of the engine flight cycle over the life of the design. The extent of the validation depends on the approach taken in the development of the engine flight cycle. For example, a conservative flight cycle where all the variables are placed at the most life damaging value would require minimum validation. A flight cycle that attempts to accurately represent the actual flight profile, but is inherently less conservative, would require more extensive validation. Applicants may apply further refinements to the engine flight cycle when significant field operational data is obtained.
(c) Flight profiles are defined as the power required versus time, and serve as the input or the boundary conditions for the next step.
(2) Engine Performance. For each flight condition or flight segment, the applicant should determine the corresponding internal performance parameters, i.e., rotor speeds, internal pressures and temperatures. The performance parameters should be adjusted for production tolerances, control tolerances, installation trim procedures, as well as for engine deterioration that can be expected between heavy maintenance intervals. The applicant should also consider the range of ambient temperature and takeoff altitude conditions encountered during the engine’s service life, as well as the impact of cold and hot engine starts.
(3) Secondary Air System Analysis.
(a) The applicant must establish the pressure, thermal and mechanical loads that must be sustained by the life-limited parts. This requires heat transfer and structural analyses that depend on boundary conditions which are derived from the internal performance parameters.
The boundary conditions are determined for each flight condition using analytical and empirical engineering processes.
(b) The applicant may use aerodynamic design software based on well established aerodynamic design principles and theory to determine the flow path interstage data—gas temperatures, pressures and mass flow. In the case of a compressor and turbine, this should be 9
the detailed interstage data associated with each individual blade and vane row. Based on this information, the applicant should determine the gas loads and air temperatures applied to the airfoils.
(c) The interstage data serves as boundary conditions for the secondary airflow system analysis. This software predicts the air temperature, pressure and air flow within the engine case and the rotor cavities, including the cavity recirculation, hot gas ingestion and leakage paths used as the boundary conditions for the heat transfer analysis. The cavity pressures are also used as boundary conditions for the structural analysis.
(d) The applicant must ensure the accuracy of the secondary airflow system software by calibrating the software using engine and rig testing.
(4) Heat Transfer Analysis. The applicant must determine the temperature levels associated with each life-limited part to ensure the cyclic stresses and strains produced by the thermal loads/temperature gradients are addressed adequately. To ensure the most damaging stresses are determined, the applicant must determine the component steady state and transient temperatures for each segment of the flight profile so that a complete temperature profile for each flight profile is determined. The heat transfer analyses usually require the use of sophisticated software tools which use the boundary conditions determined by the aerodynamic design and internal flow analyses. The analy tical model should incorporate the appropriate thermal material properties and requires knowledge of the various heat transfer coefficients derived from the boundary conditions and empirical correlations. The applicant should correlate and verify experimentally the steady state and transient component temperatures during engine development testing.
(5) Stress Analysis.
(a) Stress analysis is part of the process used to identify the life-limiting locations and the life limit for each life-limited part. The stress analysis techniques include analytical and empirical processes that are used to determine the stress-strain distributions for each life-limited component. These techniques must be capable of identifying the cyclic stresses and strains for each potential life-limiting location and of evaluating the effects of engine rotor speed, mechanical and pressure loads, thermal gradients, assembly preloads and loads applied by mating parts for the entire flight cycle. The applicant should focus particular attention on the concentrated stress locations such as bores, holes, changes in section, flanges, welds and attachment slots. The level of analysis should be sufficient to fully define the stress-strain state versus time for the entire flight profile and should be capable of predicting the three dimensional (3-D) stress-strain state. This is usually accomplished with the use of two dimensional (2-D) axisymmetric and three dimensional finite element techniques.
(b) The applicant should validate stress analysis techniques by experimental measurements or other acceptable means.
(c) The applicant should conduct studies that identify the proper use of stress analysis techniques, and identify the steps required to achieve accurate results. For finite element techniques, examples include:
Situations that require the use of axisymmetric (2-D) techniques;
Situations that require the use of non-axisymmetric (3-D) techniques;
Mesh density requirements;
Control of element aspect ratio, and
Method of computing potential stress errors.
The stress analysis techniques must be capable of determining the actual true minimum and maximum stresses-strains at the life-limiting locations. This determination should account for the non-linear inelastic material behavior based on the actual stress-stain curves. An acceptable alternative would be for the applicant to ensure the stress analysis techniques used to analyze each part are consistent with those used to analyze the fatigue tests used to establish the life of the part. If the stress analysis techniques are based on purely elastic behavior, then the applicant should ensure the fatigue test results are representative of engine operating conditions. The applicant may use various notch models, such as Neuber or Glinka, or inelastic finite element techniques.
(6) Crack Initiation Life.
(a) The applicant should employ a procedure that combines the stress-strain and temperature histories with the appropriate material and cyclic fatigue test data to establish the low cycle fatigue crack initiation life for the minimum property part. Also, the applicant should consider plasticity and creep related effects.
(b) The crack initiation life model for each material is based on test data obtained from cyclic fatigue testing of laboratory, sub-component/feature based tests, or component tests.
The model should account for the manufacturing processes that affect the low cyclic fatigue capability, including fabrication from the production grade materials and manufacturing methods. The applicant should perform sufficient testing to evaluate the effects of elevated temperatures and hold times, as well as interaction with other material failure mechanisms such as high-cycle fatigue and creep. Appropriate service experience gained through a successful program of parts retirement, precautionary sampling inspections, or cyclic component tests may be included to calibrate the life prediction system. The life prediction system should also account for environmental effects, such as vibration, corrosion and cumulative damage.
(c) Test data should be reduced using appropriate statistical techniques to account for inherent scatter and should be expressed in terms of an acceptable part risk level, which is 1 cracked part out of 1000 (B.1 life) or alternately 1 in 740 (-3 sigma). Part fatigue life should be quoted in terms of the minimum number of flight cycles required to initiate a crack approximately 1/32" or 0.030" (0.762 mm) in length.
(d) When the fatigue life is based on a cyclic test of a specific part or parts, the applicant must correct the test results for inherent fatigue scatter and for the differences between the test and engine operation. The factors used to account for scatter should be justified. To use this approach, applicants should design the test to reproduce, as nearly as possible, the engine critical operating conditions in terms of the cyclic stresses, strains, and temperature level at the life-limiting feature, such as the, bore, rim, or blade attachment slot. The applicant should use sufficient analytical and empirical tools so the differences between the engine conditions and the 11