AC 35.37-1B

Subject: Propeller Fatigue Limits and Date: 3/24/11 AC No: 35.37-1B Evaluation Initiated By: ANE-111

1. Purpose . This advisory circular (AC) provides guidance and describes methods, but not the only methods, for demonstrating compliance with § 35.37, Propeller fatigue limits and evaluation, of Title 14 of the Code of Federal Regulations (14 CFR).

2. Applicability.

a. The guidance provided in this document is directed to propeller manufacturers, modifiers, and foreign regulatory authorities.

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. Cancellation . This AC cancels AC 35.37-1A, Guidance Material for Fatigue Limit Tests and Composite Blade Substantiation; September 17, 2001.

4. Related Reference Material.

a. Regulations. 14 CFR §§ 35.4, 35.37, 23.907, and 25.907.

b. Advisory Circulars. Please check the FAA’s website at http://www.faa.gov/regulations_policies/ for the latest revision of the following documents:

(1) AC 35-1, Certification of Propellers; December 19, 2008.

(1) AC 20-66B, Propeller Vibration and Fatigue; March 4, 2011.

(2) AC 20-107B, Composite Aircraft Structure; September 8, 2009.

(3) AC 21-26, Quality Control for the Manufacture of Composite Structures; June 26, 1989.

(4) AC 23-13A, Fatigue, Fail-Safe, and Damage Tolerance Evaluation of Metallic Structure for Normal, Utility, Acrobatic, and Commuter Category Airplanes’; September 29, 2005

(5) AC 23-20, Acceptance Guidance on Material Procurement and Process Specifications for Polymer Matrix Composite Systems; September 19, 2003.

(6) AC 25.571-1C, Damage Tolerance and Fatigue Evaluation of Structure; April 29, 1998.

c. Military Specifications. MIL-HDBK-17-1F, June 17, 2002.

5. Definitions. For the purposes of this AC, the following definitions apply.

a. Damage tolerance. The attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of damage from fatigue, corrosion, accident, or a discrete source.

b. End of life condition. The physical condition of the component when it has sustained the maximum extent of damage but maintains sufficient residual strength to meet all airworthiness loading requirements. This condition is defined during certification.

c. Fail-safe. The attribute of the structure that permits it to retain its required residual strength for a period of use without repair after the failure or partial failure of a principal structural element.

d. Flaw. A pre-existing anomaly in the structure that is either created during manufacture or has resulted from damage to the structure introduced after manufacture. Generally, flaws can be geometrically quantified by non-destructive inspection methods. The term flaw is widely used in fracture mechanics disciplines for assessing damage tolerance.

e. Life factor. The factor applied to increase or decrease the number of cycles on an S-N (see definition of S-N in paragraph 5.l. below) diagram as needed to properly account for test sample size and variations in material properties, environmental effects, service deterioration, 2

repairs, and manufacturing anomalies. This is often used in combination with a load factor, as shown in figure 1 in the Appendix.

f. Limit load. The maximum load expected in service.

g. Load factor. A factor applied to increase or decrease load or stress level on a fatigue life diagram (S-N diagram) as needed to properly account for test sample size and variations in material properties, environmental effects, service deterioration, repairs, and manufacturing anomalies. This is often used in combination with a life factor, as shown in figure 1 in the Appendix.

h. Principal structural element. The element that contributes significantly to the carrying of propeller loads and whose integrity is essential in maintaining the overall structural integrity of the propeller.

i. Safe-life. The number of events, such as stress cycles, flights, landings, or flight hours, during which there is a low probability that the strength will degrade below its design value due to fatigue damage.

j. Scatter factor. A life reduction factor used in the interpretation of fatigue analysis and test results.

k. S-N Diagram. A plot of stress, strain or load (S) against the number of cycles (N) required to cause failure of specimens in a fatigue test. Also known as a stress life curve or a Wöhler curve.

l. Stress ratio (R). For repeated stress cycles, this is the ratio of the minimum stress to the maximum stress, R = min/max.

6. Discussion . Propellers are continuously subjected to steady and vibratory stresses under many different operating conditions on an airplane, both while in flight and on the ground.

Therefore, § 35.37 requires a fatigue evaluation and the determination of fatigue limits. The data developed to show compliance with § 35.37 is also used to support installation of the propeller on an airplane as required by §§ 23.907 and 25.907 for an initial airplane type certificate (TC) or for modified installations through an amended TC or supplemental type certificate (STC). Since the rate of accumulation of stress cycles for propeller blades, hubs, and other propeller components is very high, the design goal, whenever possible, should be to show that stresses are below the component or material endurance limit. However, not all materials have a welldefined endurance limit. In addition, the stresses developed during maneuvers, ground operation, ground air ground (GAG) cycles, or in other areas of the airplane operating envelope, may cause damage. The applicant should evaluate the accumulation of this damage to determine if propeller components are life-limited or require mandatory inspections or to determine if the propeller is suitable for use on an airplane.

a. Fatigue Limits. Establishing propeller fatigue limits is one step in the structural evaluation for certification and approval for use on an airplane. Fatigue limits refer to the 3

material and structural databases used to support the fatigue evaluation of the propeller on an airplane. Fatigue limits may take many forms, such as Goodman diagrams to assess whether or not stresses are below the endurance limit; S-N curves for safe-life evaluations; and crack growth curves (da/dn curves) and delamination growth curves for damage tolerance evaluations. The fatigue limits are developed with an understanding of the propeller component, material, failure mechanism, environment, and loading, and may also include the effect of service damage. When determining the fatigue limits under § 35.37, the applicant should also consider the possible airplane operating environment. The development of fatigue limits can be supported by previous testing, past experience, and acceptable published data, when available.

b. Damage Accumulation. The fatigue limits are developed to support a damage accumulation algorithm, such as Miner’s rule for safe-life calculations or a crack growth or damage growth algorithm for damage tolerance calculations. The damage accumulation algorithm is used for the fatigue evaluation required by § 35.37 and may also be used as part of showing compliance with the airplane vibration requirements in §§ 23.907 and 25.907. Verify the damage accumulation algorithm by previous testing, past experience, and acceptable published data.

c. Applicable Components.

(1) Components. For propeller certification, fatigue limits are established for the propeller hub, blades, blade retention components, and components whose failure due to fatigue may cause a hazardous propeller effect. Examples of components whose failure may cause a hazardous propeller effect are the pitch change piston pressure cylinder, counterweights, and pitch control components. For items such as bearings, which are typically part of the hub assembly, the fatigue limits are generally established for the assembly, not for the individual components.

(2) Regions. Assess each applicable component to determine if it has multiple critical regions or if it requires different limits for different regions. Blades are an example of a component that has different limits for different regions, typically blade shank, mid-blade, and tip. The blade steady stresses vary substantially from blade root to blade tip; the blades should have fatigue limits established for each of the stress regions. Also, metallic blades may have different material limits to account for local working effects due to shotpeening or cold rolling.

d. Propeller Loads. The loads applied during the fatigue limit tests are derived from the consideration of the steady and vibratory propeller loading conditions that occur on the intended airplane and engine installation throughout its life or on a typical airplane. The loads applied during fatigue limit tests reflect the fatigue data to be generated. For example, the loads establish the stress ratios (R) for coupon specimen tests and set the amplitude and direction for full-scale testing. Also, the load magnitude and direction should be established in a manner that represents the loading the propeller will experience in service.

(1) Operating Spectrum Conditions. The applied loads are derived from an airplane operating spectrum. The operating spectrum depends on the category and operation of the airplane and includes both flight and ground operation.

(a) Flight conditions include conditions that occur with each flight such as:

1 Take-off;

2 Climb;

3 Cruise;

4 Descent;

5 Approach;

6 Landing; and 7 Reverse thrust.

(b) Transient airplane flight conditions are associated with maneuvers such as:

1 Banked turns;

2 Side-slip;

3 Pull-ups;

4 Push-overs;

5 Rudder kicks; and 6 Gusts.

(c) Special flight conditions are specific to a mission such as:

1 Fire-fighting;

2 Aerobatic maneuvers;

3 Emergency conditions; and 4 Training maneuvers.

(d) Ground conditions include, but are not limited to:

1 Taxi;

2 Operation in cross-winds; and 5

3 Maintenance checks including, but not limited to, run-up governor function checks and propeller feather checks.

(e) Engine load conditions include the loads generated by the engine and transmitted to the propeller, such as firing impulses from reciprocating engines.

(2) Operating Spectrum Source. The airplane operating spectrum, when available, may be obtained from the airplane company for the intended application. When the airplane operating spectrum is not available, the spectrum may be developed from design assumptions and design experience on a typical or the intended airplane. AC 23-13A provides airplanerelated examples of operating spectrums and fatigue assessments. For use with propellers, airplane loads would need to be converted to propeller once per revolution loads using analytical methods such as those addressed in AC 20-66B.

e. Component Degradation. The fatigue limits should account for likely service deterioration, variations in material properties, manufacturing anomalies, and environmental effects. Many methods are available to account for component degradation, such as using a life or load factor to reduce the fatigue limits or apply to fatigue test loads. Reductions in fatigue limits may be applied as life or load factors based on the manufacturer’s service experience and test database with the components, as shown in figure 1 in the Appendix and defined in paragraphs 5.e. and g. The reductions may also be developed by appropriate specimen tests or by testing intentionally degraded full-scale components.

(1) Manufacturing Anomalies. Propellers manufactured to a process specification have established parameters to control manufacturing variables. These parameters should be considered for their impact on structural integrity.

(a) For metal propellers, these include, but are not limited to, surface finish and machining marks.

(b) For composite propellers, these include, but are not limited to, fiber misalignment, fiber content, porosity, and delaminations.

(c) For adhesive bond joints, these include, but are not limited to, adhesive thickness, voids, and disbonds.

(2) Service Damage. Propellers are exposed to and encounter a certain amount of service damage from corrosion, erosion, stone strikes, and handling damage. Composites, in addition, may be damaged by small bird strikes and hail impact. The damage to a composite should consider both visible surface damage and hidden internal damage, such as delaminations.

Consider the extent of tolerance to such damage and the method of demonstration.

(3) Environmental Degradation. Airplanes can spend their entire operational life in a severe climatic zone. Therefore, the strength degradation of the material system should be established. When assessing operational environments, consider the following:

(a) High temperature;

(b) Humidity;

(c) Low temperature;

(d) Thermal cycling;

(e) Ultraviolet light; and

(f) Aviation chemicals.

f. Repairs. Repairs change the condition of the propeller as manufactured. Therefore, the manufacturer should consider the impact of planned repairs on the structural integrity of the propeller.

g. Endurance limit. When conducting coupon tests, test the coupons to failure under a combination of steady and vibratory loads. The steady loading should represent that anticipated in service for the propeller. Select the vibratory loading component to facilitate the generation of a fatigue or S-N curve, over a broad range of stress cycles, to define an endurance limit.

(1) The endurance limit for composites is normally projected to be at least 500 million cycles.

(2) The endurance limit for aluminum is normally projected to be at least 100 million cycles.

h. Accelerated Testing. Attempts to compress testing time by using high-frequency vibratory loading should be approached with caution to prevent the introduction of nonrepresentative temperatures, failure modes, and fatigue lives. The vibratory loading may be at any frequency if internal heating does not produce harmful effects and representative failure modes are realized. Accelerated testing should include some form of temperature monitoring to prevent overheating the specimen. Cooling may be required to prevent overheating the specimen.

i. Coupon Tests. We recommend establishing an S-N curve from coupon tests for determining fatigue strength and statistical strength distribution of propeller materials. Establish the S-N curve for both the low cycle fatigue (LCF) and high cycle fatigue (HCF) regimes. The S-N curve should be representative of the propeller’s material system. Develop S-N curves by using acceptable published data, when available, or by testing a sufficient number of specimens that represent the propeller’s material system manufacturing processes.

(1) Metals. For metals, this should include the following effects:

(a) Surface finish;

(b) Cold rolling;

(c) Shot peening;

(d) Corrosion inhibitors; and

(e) Surface treatments, such as plating.

(2) Composites. For composites, this should include the following effects:

(a) Fiber, resin, and sizing;

(b) Ply stacking sequence;

(c) Ply orientation; and

(d) Manufacturing processes.

(3) Bond Joints. For bond joints, this should include the following effects:

(a) Adhesive materials;

(b) Surface preparation;

(c) Adhesive thickness; and

(d) Manufacturing processes.

j. Full-scale Testing.

(1) Test Specimen. Full-scale components should be manufactured to represent the type design and should be fatigue tested at combinations of steady and vibratory loads as needed to support subsequent evaluations. Also, to address the issues of safe-life, damage tolerance, and continued airworthiness, manufacture the specimens to include specific manufacturing anomalies and likely service damage. The extent of such defects/damage should be consistent with inspection techniques employed in service.

(2) Failure Criteria. Applicants should establish rational failure criteria. This will differ for various types of construction and materials. Composites tend to fail between layers first, while metals tend to fail through the thickness of the part and have much smaller permissible defect sizes.

(a) For composites, as shown in figure 2 in the Appendix, fatigue failure is markedly progressive, starting at initiation (region I); through a damage growth or delamination growth phase (region II); to an advanced damage state in which large delamination and 8

secondary failure modes and locations form throughout the component (region III). Failure criteria such as a specific loss of stiffness in the structure, visible damage, and delamination area should be applied.

(b) For metallic components, the failure criteria is typically a detectable crack.

The selected failure criteria will be a factor in establishing the component life since the component should maintain sufficient residual strength throughout its life.

(c) When using stiffness loss as failure criteria, address the effect on component resonant frequency to determine that the critical speed placement is acceptable as this may affect the continued airworthiness of the component. Vibratory loads may be magnified as component natural frequencies approach the frequency of the load.

(3) Component Monitoring. Components should be strain gauged and have load cells, as appropriate, for testing to monitor stresses and loads. When a test rig applies load with an amplitude actuator, the input amplitude may require adjustment to account for test rig and/or component wear-in or degradation to ensure that the test load is maintained. Examine the specimens regularly for cracks, delamination, or other degradation. When using stiffness as the failure criteria, it should be measured for each specimen at the outset of each test and periodically monitored throughout the test. The monitoring should be more frequent during the failure process. When appropriate, continue testing after the component has failed, according to the failure criteria, to demonstrate damage growth characteristics and residual strength capacity.

This may include testing to the limit load conditions after LCF or HCF tests are completed to demonstrate residual strength at the end of life condition.

k. Data Reliability.

(1) Mean S-N Diagram. Define the mean trend line for the data for each critical location from the mean steady and vibratory stress for all coupons/components tested.

Extrapolate this mean trend line out to an asymptote representing the HCF endurance limit and extrapolated back to a low number of cycles. The coupon-derived S-N curve shape may be used to supplement extrapolation through the failure points when the coupon tests duplicate the failure mechanism.

(2) Reliability. As dynamically loaded components, the structural integrity of propellers is generally governed by their fatigue strength, rather than their static strength.

Accordingly, a reliability at least as good as the “A” basis of a normal distribution should be demonstrated. That is a reliability of 99 percent, with a 95 percent confidence level.

(3) Number of Test Specimens. There is no set number of full-scale test specimens that should be fatigue tested for each critical propeller section if the required reliability is satisfied.

(a) For “A” basis reliability, the mean endurance limit (E50) should be reduced by a factor (k), governed by the selected sample size (n) and the standard deviation () data.

Specifically, the 99 percentile endurance limit (E99) may be expressed: E99 = E50 -k ; in which 9

(k) is a function of the sample size (number of specimens tested). The “A” basis values may also be defined by other distributions such as Weibull or Log Normal, depending upon the fit. Refer to MIL-HDBK-17-1C for details.

(b) For the normal strength distribution, 0.95 confidence, (k) varies with sample size as follows:

Table 1. Normal Strength Distribution Factors Sample size Normal Distribution Reduction Factor n k 2 37.094 3 10.553 4 7.042 5 5.741 6 5.062 7 4.642 8 4.354 9 4.143 10 3.981 REF. Table 8.5.11, MIL-HDBK-17-1F

(c) The manufacturer’s in-house methodology, development data, and analytic expertise, along with limited full-scale testing, may be used as an alternative to substantiate the required reliability.

(4) Working Curve. The working curve is the design curve. It is developed from the “A” basis curve by dividing by the appropriately justified load or life factors to account for the manufacturing variations, environment, and service conditions. Develop factors for composite materials for each new composite material. These materials include, but are not limited to, fiber, resin, coating, and adhesives.

7. Composite Component Fatigue Limits.

a. Discussion. An increasing number of propeller blades are constructed of composite materials. Due to the anisotropic characteristics of composite materials, component design and verification of design differ significantly from that of metal structures. Propeller blade fatigue loading also differs significantly from most other structures. Past practices for metal structures or other composite structures may not be adequate when dealing with composite propeller structure and fatigue loading. Although this discussion focuses on propeller blades, the principles may be applied to any composite structure. With the added flexibility of composite component design, include damage tolerance and fail-safe principles whenever possible. When 10

damage tolerance methods are applied, the component should be designed using damage tolerance and fail-safe principles.

(1) Failure Progression. Figure 2 in the Appendix illustrates the typical progression of fatigue degradation and reduction in residual strength of a typical composite component. Region I is the damage initiation phase. Region II is the damage or delamination growth phase. This is the region where delamination growth propagates steadily and predictably. Region III is in an advanced damage state. Large delamination and secondary failure modes and locations form throughout the laminate. In region III, component stiffness and residual strength degrade rapidly. This AC contains guidance for fatigue limits to be used in a fatigue evaluation to ensure that region III is not reached in service.

(2) Fatigue Limits. Develop fatigue limits to support a fatigue evaluation based on a safe-life approach or a damage tolerance approach. This AC provides guidance for both of these approaches. When appropriate, the fatigue limits should address permissible repairs that alter the as-manufactured condition of the propeller.

b. Safe-Life Fatigue Limits. Use a two-phase approach to develop the fatigue strength characteristics for composite components to be used in a safe-life-fatigue evaluation. The first phase is coupon specimen testing for both static properties and fatigue properties to supplement the fatigue (S-N) curve shape, standard deviation, and statistical distribution. The second phase is full-scale specimen testing to establish the strength level, statistical distribution, failure location, and failure mode and mechanism for the material system and geometry of the propeller.

For blades, this may involve testing of metallic to composite bonds, shank, mid-blade, and tip specimens. Figure 3 in the Appendix illustrates some types of test rigs that have been used to assess various areas of a blade. The development of the fatigue properties may be supported by previous testing, past experience, and acceptable published data, when available.

(1) Loading.

(a) Two types of loading may be needed for testing full-scale composite structures.

Constant loading defines the S-N curve. Spectrum loading addresses cumulative loading effects, since damage depends on both LCF and HCF loads. These loads include GAG cycles, low– occurrence, high-amplitude maneuver limit loads, and high-occurrence take-off, climb, and cruise loads. The fatigue characteristics may also depend on the order in which the loads are applied. This should be considered when establishing the test load spectrum. This effect should be addressed by performing many sets of load blocks.

(b) LCF and HCF are closely related and both of these should be addressed. Fullscale testing of composite components usually emphasizes LCF initially since composite S-N diagrams have a shallow slope. However, unless the coupons duplicate the mode of damage, coupon tests have limited value in areas with complex geometry or metallic to composite bonds.

(c) Constant amplitude HCF failures are generated in the 106 to 108 cycle range.

To induce failures in this cycle range, the alternating component of the load(s) is generally increased by a load factor from design or operating values. The steady component of load is 11

generally not increased to induce failures. If no failure occurs, a no failure point, known as a run-out point, is generated. Specimens that have completed a run-out should not be re-tested to generate failure points at higher load levels. The run-out specimen may have accumulated damage during previous testing that could bias the test results. Use a new specimen for testing at a higher load level to generate a failure point.

(d) Constant amplitude LCF failures are generated in less than 106 cycle range.

Failures may be induced by increasing the number of cycles by a life factor on the applied load or by increasing the vibratory component of the applied load by a load factor, or a combination of both. The steady component of load is generally not increased to induce failures. Subsequent data evaluation should account for the cycle or load increase required to generate the failures.

Use load factors with caution, as the failure mode may change while testing at increased load.

(e) Spectrum load testing is often used to verify the damage accumulation algorithm for subsequent fatigue evaluations. Miner’s rule, with appropriate consideration for scatter factor, has proven to be suitable for calculating component life. A load spectrum includes operating conditions representative of the intended operating conditions for the propeller.

Within this spectrum, periodically apply a high amplitude low cycle maneuver load. To account for scatter, apply life factors instead of load factors which may alter the failure mode. For blades, the load spectrum may include elements of a typical flight consisting of the following:

1 Start-up;

2 Taxi out;

3 Run-up to take-off thrust;

4 Maximum once-per-revolution (1P) vibratory load;

5 Climb;

6 Maneuver and other transient loads, as applicable;

7 Cruise;

8 Descent;

9 Landing;

1 0 Reverse thrust;

1 1 Taxi back; and 1 2 Shut-down.