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a. Overview. The process to show propeller loads by similarity involves a review of the vibration stress survey from the baseline airplane. The review should identify trends in load variation and the conditions that caused the maximum loads during ground and flight operation.
The review should also evaluate the target airplane, identifying the probable sources of loads.
The vibratory loads may be due to engine or aerodynamic excitation forces. The evaluation should show that the operating environments vary considerably in complexity from single reciprocating engine installations to multi-engine turbine installations.
b. Aerodynamic Environment. The propeller aerodynamic environment should be evaluated by a substantiated analysis to compute the flow into the propeller from the baseline airplane to the target airplane. AQ analysis has proven to be suitable for the evaluation of some installations. The parameter AQ is the product of angle of flow (A) into the propeller multiplied by the airplane dynamic pressure (Q):
AQ = ψ x (1/2ρV2) In which: ψ - total inflow angle into the propeller, degrees ρ - air density, lb-sec2/ft4 V - air speed, ft/sec The AQ is proportional to the propeller vibratory loads in flight due to angular inflow. For more information on aerodynamic excitations, refer to Appendix 1.
c. General Considerations.
(1) An approved propeller vibration stress survey and evaluation showing compliance with §§23.907 or 25.907 should have been conducted on the baseline airplane to form the basis for the target airplane.
(2) The vibratory loads associated with aerodynamic excitation forces measured in flight on the baseline airplane should be below the endurance limits of the propeller for normal flight conditions. Similarity should not be used when the loads measured in flight on the baseline airplane are close to the endurance limits of the propeller. Analyses and measurement to show similarity are generally inadequate when load margins are close to the endurance limits of the propeller.
(3) The baseline airplane and target airplane should have equivalent engines and propellers that are equivalent in vibration.
(a) Reciprocating engine model differences may cause significant variations in propeller loads. Therefore, for reciprocating engines the baseline engine model and ratings and the target engine model and ratings should be equivalent.
(b) Turbine engine model number variations associated with component differences not affecting the vibration characteristics of the propeller may be acceptable.
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(4) The engine and propeller control systems should be rigged similarly so that the propeller loading on the target airplane is shown to be less than that of the baseline airplane in reverse thrust, feather, taxi, ground and flight operation.
(5) The power and RPM ratings for the target airplane should not exceed that of the baseline airplane. Limitations and placards should be the same or more restrictive.
(6) Similarity should not be used for acrobatic airplanes due to potentially significant differences between installations in gyroscopic and aerodynamic propeller loading.
(7) Similarity should not be used when there are service life limits associated with the propeller on the baseline airplane.
(8) The target airplane should be of the same category (normal, utility, agricultural use) as the baseline airplane or one with a less severe operating environment.
(9) All propeller airworthiness limitations from the baseline airplane should be applied to the target airplane.
(10) The engine mounts and the flexibility of the support structure should be equivalent for the baseline airplane and the target airplane.
10. NOTE 9 VIBRATION APPROVALS. Some propeller TC data sheets contain a section called Note 9. This note simplifies the vibration compliance for propellers under §23.907 for normal category, single reciprocating engine airplanes in tractor configuration. Note 9 cannot be used as substantiation for vibration approval of propellers for any other installation.
a. Applicability. Note 9 lists the propeller-engine combinations that the Propeller CMACO and the propeller manufacturer have approved as suitable for operation on normal category, single reciprocating engine airplanes in tractor configuration. These propeller-engine combinations have been shown by test to have vibration loads dominated by the reciprocating engine excitations. Therefore, the aerodynamic vibratory environment created by a single engine tractor configuration airplane does not significantly affect the propeller vibration.
b. Use. To use Note 9 for approval the following should be shown:
(1) All placards and other restrictions for the propeller-engine combination are followed.
(2) The installation is as shown in Note 9.
(3) The hub, blade, and engine models are identical to those listed in Note 9.
(4) The airplane is a normal category single reciprocating engine tractor configuration.
(5) The applicability has been confirmed by the Propeller CMACO.
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c. Not Applicable. Note 9 does not apply to the following airplanes:
(1) Restricted category (including agricultural airplanes);
(2) Utility category;
(3) Acrobatic category;
(4) Transport category;
(5) Turbine engine;
(6) Multi-engine; or
(7) Pusher airplanes.
11. PROPELLER FATIGUE EVALUATION. For the propeller, the propeller fatigue evaluation establishes the fatigue life, mandatory replacement times (life limits), and, in some cases, mandatory inspections for components due to fatigue. For the airplane, the propeller fatigue evaluation establishes operating and airworthiness limitations that may be required for safe operation of the propeller. Although a uniform approach to fatigue evaluation is desirable, the complexity of the problem makes a uniform approach unfeasible. New design features, methods of fabrication, approaches to fatigue evaluation and configurations may require procedures other than those described in this AC. In addition, many different phenomena influence the fatigue life of the propeller. Therefore, assessing and assuring the fatigue life of the propeller should begin at the earliest stages of design and should end with the fatigue evaluation on the airplane.
a. Design Goals. Since the rate of accumulation of load cycles for propeller blades, hubs, and other propeller components is very high, the design goal should be to show that loads are below the component or material endurance limit whenever possible. However, all materials do not have a well defined endurance limit, and the loads developed during maneuvers, ground operation, ground air ground (GAG) cycles and in other areas of the airplane operating envelope may cause damage. The accumulation of this damage should be considered when determining if propeller components are life limited or require mandatory inspections, or if the propeller is suitable for use on the airplane.
b. General. There are a number of different approaches to fatigue evaluation. This AC provides safe-life and damage tolerance approaches. The approaches presented in this AC are suitable to both metallic structures and composite structures. The method used for the fatigue evaluation is also affected by the material and failure mode. Regardless of the approach, the fatigue evaluation should include the following elements:
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(1) Applicable Components. A fatigue evaluation should be performed on the hub, blades, blade retention components and any other propeller component whose failure due to fatigue could be hazardous or catastrophic to the airplane. The following are examples of components whose failure may be hazardous or catastrophic to the airplane:
(a) Pitch change piston pressure cylinder (dome);
(b) Counterweights; and
(c) Pitch control components.
(2) Locations. When identifying which locations should be evaluated, consider areas prone to probable damage, such as corrosion, denting, gouging, wear, erosion, bird impact, and other foreign object damage. Also consider, as necessary, the following:
(a) Results of stress analyses;
(b) Static tests;
(c) Fatigue tests;
(d) Strain gage surveys;
(e) Tests of similar structural configurations; and
(f) Service experience. Service experience has shown that special attention should be focused on the design details of important discontinuities such as the adhesive bond between a composite blade and a metallic blade retention, hub mounting faces, bolt holes, dowel pin holes, and blade bearing retention.
(3) Effects. Consider the effects of material variability and environmental conditions on the strength and durability properties of the propeller materials.
(4) Fracture Modes. Identify fracture modes for the structural components. The components should be assessed to establish appropriate damage criteria in relation to the ability to be inspected and damage characteristics from initial detectability to fracture.
(5) Damage Accumulation. Appropriate and substantiated damage accumulation algorithms should be selected. Examples of these are Miner’s rule for safe-life calculations or a crack or damage growth algorithm for damage tolerance calculations. The damage accumulation algorithm should be verified by previous testing, past experience, and acceptable published data, when available.
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(6) Instructions for Continued Airworthiness (14 CFR part 35 Appendix A). Mandatory replacement periods and inspection intervals must be included in the Airworthiness Limitation Section of the Instruction of Continued Airworthiness. If the vibration survey and fatigue evaluation indicate that certain operating conditions or ranges should be limited, those limitations should be included in the appropriate propeller and airplane manuals (refer to section 12 of this AC).
(7) Component Degradation and Repair. The fatigue evaluation and resonant frequency placement should account for likely service deterioration, variations in material properties, manufacturing anomalies, environmental effects, and permissible repairs. For methods to account for component degradation and repairs, refer to AC 35.37-1A. For example, the frequency of an aluminum blade may increase as the blade width and thickness decrease with erosion, and the frequency of composite blades may decrease as material is added when the blade is repaired.
c. Airplane Operating Spectrum. The airplane operation spectrum depends on the category and operation of the airplane. The overall airplane operating spectrum involves the combination of all ground and flight conditions in the operation of the airplane throughout its life. The operating spectrum should be determined from ground and flight test data on the intended airplane and engine combination with the installed propeller. The propeller loads are then determined for each airplane operating condition in the spectrum.
(1) The elements of the operating spectrum include normal flight conditions that occur with each flight:
(a) Take-off;
(b) Climb;
(c) Cruise;
(d) Descent;
(e) Approach; and
(f) Landing and reverse.
(2) The elements of the operating spectrum also include transient airplane flight conditions:
(a) Maneuvers (banked turns, side-slip, pull-ups, push-overs, rudder kicks, etc.);
(b) Gusts;
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(c) Special flight conditions specific to a mission (fire-fighting, acrobatic, agricultural, etc.);
(d) Emergency conditions;
(e) Airplane limit load conditions; and
(f) Training maneuvers.
(3) The operating spectrum also includes ground operating conditions:
(a) Taxi;
(b) Operation in cross winds; and
(c) Maintenance checks.
(4) The airplane applicant should provide the airplane operating spectrum for the intended application. The propeller manufacturer may develop additional operating spectrum information to supplement the airplane data. Loads associated with portions of the operating spectrum may not be directly measurable, such as some severe gust conditions, limit load conditions, and some emergency conditions that may threaten the safety of the airplane. These conditions should then be extrapolated or derived based on the available test data. When the airplane operating spectrum is not available, the spectrum information should be based on the design assumptions and design and service experience for the intended airplane and engine application.
(5) The number of occurrences of each operating condition and the duration of each operating condition are also associated with the airplane operating spectrum. Table 1 provides elements of an operating spectrum using a transport category airplane as an example. As shown, most of the flight conditions occur within daily normal flight operation of the airplane. Some operating conditions occur once in the life of the airplane, such as an extreme yaw or extreme high “g” maneuver. The propeller fatigue evaluation should consider all of these operating conditions.
(6) The airplane operating spectrum should include loads associated with LCF ground-airground (GAG) cycles that occur with each flight. Within each flight there is a maximum and minimum load. Each GAG cycle is capable of causing fatigue damage and should be considered in the test planning, data acquisition, and fatigue evaluation. Figure 3 illustrates in schematic form the load variation and the GAG cycle of a normal flight.
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(7) The airplane applicant should define the airplane operating spectrum so it can be used as necessary to thoroughly evaluate the propeller loading throughout the intended flight envelope and to conduct a fatigue evaluation. Some installations may have substantial load margin; in these cases, only the maximum and minimum load levels associated with the operating spectrum would be needed for the fatigue evaluation. Other installations may have operating conditions that produce load amplitudes that cause fatigue damage to the propeller blades or other applicable propeller components. Those operating conditions that cause fatigue damage should be included in the airplane operating spectrum and fatigue evaluation as necessary.
(8) After identifying the airplane operating spectrum, the stress or load levels at each of the operating conditions should be determined, using the measured vibratory stress data and associated analytical extrapolation and interpolation methods.
d. Safe-Life Evaluation. The safe-life approach to fatigue evaluation is based on the principle that the repeated loads can be sustained throughout the intended life of the propeller, during which there is a low probability that the strength will degrade below its design value due to fatigue.
(1) Goodman and S-N diagrams are usually used for the safe-life evaluation. These diagrams should be developed from an appropriate combination of coupon tests and full-scale component tests, as required by §35.37 and described in AC 35.37-1A.
(2) To determine the safe-life, reduce the component fatigue life by an appropriate scatter factor that accounts for the variability of the fatigue evaluation process. To determine the fatigue life, combine the loads generated for the airplane operating spectrum with the fatigue data, using a damage summation algorithm (safe-life evaluation). A scatter factor of three or greater should be used for metallic structure, unless there is substantial justification for a lower scatter factor.
Mandatory replacement times should be established for parts with safe-lives. Mandatory replacement times must be included in the propeller Airworthiness Limitations Section of the propeller Instructions for Continued Airworthiness. The mandatory replacement times may also be included in the airplane Airworthiness Limitations Section of the airplane Instructions for Continued Airworthiness.
(3) When the propeller loads are below the endurance limits defined in the Goodman diagram, fatigue damage is not accumulated and damage summation is not necessary. This generally does not occur. The loads generated during extreme and emergency maneuvers and GAG cycles may result in loads that are above the fatigue limit, accumulating fatigue damage. In these cases, summation procedures such as Miner’s rule may be employed, if properly substantiated. Small deviations in the loads at some conditions in the operating spectrum may have appreciable influence on the calculated fatigue life. Therefore, the safe-life should be established carefully, and a life sensitivity to variations in the loads should be performed. The scatter factor should be increased when the fatigue life is highly sensitive to load variations.
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(4) When all loads are below the endurance limits established for the component, the component is said to have unlimited life. In addition, when the safe-life of a component is shown to be greater then 70,000 hours, and it is shown that the component will be retired from service for reasons other than fatigue prior to 70,000 hours, the component may be said to have unlimited life.
(5) Miner’s rule may be applied to composite materials when sufficiently substantiated.
Since the application of Miner’s rule to composites may be highly unconservative, the safe-life should be established using a substantiated scatter factor. A scatter factor of ten or greater should be used for composite structures, unless there is substantial justification for a lower scatter factor.
The applicability of Miner’s rule summation for composites should be verified by testing for the full-scale structure, using a load spectrum established from the flight test. For blades, the load spectrum may include loads from typical airplane operation consisting of start-up, taxi out, runup to take-off thrust and maximum 1P vibratory load, climb, cruise, descent, landing, reverse thrust, taxi back, and shut-down. Within this spectrum a high amplitude low cycle maneuver load should be applied periodically.
e. Damage Tolerance Evaluation. This AC assumes that when damage tolerance methods are applied the component has been designed using damage tolerance and fail-safe principles. The damage tolerance methods are based on the principle that damage is inherent in the structure or inflicted in service and may grow with the repeated application of loads, and the propeller or propeller components will be inspected at intervals to assess the extent of damage. When damage reaches the maximum permissible flaw size, the propeller or propeller component will be retired.
(1) Inspection Interval. For damage tolerance methods the inspection interval is determined by the relationship between the time the damage reaches maximum permissible flaw size as defined during certification (detectable damage) and the end of life condition. The maximum permissible flaw size is established during certification by considering the inspection method, the inspection interval, and the end of life condition. The inspection interval permits multiple opportunities, usually three, to find the damage before the component reaches the end of life condition. The inspection method should also be evaluated to determine the probability of detection (POD). Inspection methods should have a POD of 90% probability with 90% confidence. When the POD is less than 90% probability with 90% confidence, the inspection frequency should be increased. The component should be removed from service when damage is detected at the maximum permissible flaw size. The Airworthiness Limitations Section of the Instructions for Continued Airworthiness establishes these inspections as mandatory.
(2) Damage Growth Data. The damage tolerance evaluation can be applied to both metallic and composite materials.
(a) With metallic materials the damage is generally a crack. The damage growth is characterized by da/dn curves in which the crack growth rate is plotted against a stress intensity factor.
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(b) With composites the modes of damage accumulation are typically more dispersed and should be characterized for each unique full-scale composite structure. One mode of damage accumulation noted in laminated composites is delamination, which can be characterized in growth rate curves as shown in Figure 4. AC 35.37-1A provides an approach to the development of delamination growth rate curves.
(c) The damage tolerance data and evaluation should include:
1. Structural details, elements, and sub-components of critical structural areas tested in accordance with §35.37 to define the sensitivity of the structure to damage growth.
2. Environmental effects on the flaw growth characteristics. The environment assumed should be appropriate to the expected service usage.
3. Loading representative of anticipated service usage.
4. Testing including damage levels (including impact damage) typical of those that may occur during fabrication, assembly, and in-service.
5. Test articles fabricated and assembled in accordance with production specifications and processes associated with the type design, so that the test articles are representative of production structure.
(3) Verification. The applicability of the damage tolerance assessment should be verified by spectrum loading for the full-scale structure, using loads established from the flight test. The detectable damage size and location should be established and be consistent with the inspection techniques employed during manufacture and in service. Flaw/damage growth data should be obtained by repeated load cycling of intrinsic flaws or mechanically introduced damage. The damage growth model should be validated by tests of full-scale components.
(4) Residual Strength. The residual strength of the component should be demonstrated on full-scale damaged components at the end of life condition. It should also be shown that the stiffness properties have not changed beyond acceptable levels with the maximum extent of damage. The end of life condition is established in conjunction with the service life. Therefore, the component in its end of life condition is still safe.
12. AIRPLANE AND PROPELLER OPERATING AND AIRWORTHINESS LIMITATIONS. Each propeller or airplane operating and airworthiness limitation necessary for safe operation of the airplane and propeller should be appropriately documented.
Documentation includes, but is not limited to, the following:
a. Airplane Airworthiness Limitations Section, Airplane Instructions for Continued Airworthiness, for life limits and mandatory inspections;
b. Airplane Instructions for Continued Airworthiness;
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c. Airplane Flight Manual;
d. Airplane Placards;
e. Propeller Airworthiness Limitations Section, Propeller Instructions for Continued Airworthiness, for life limits and mandatory inspections;
f. Propeller Instructions for Continued Airworthiness; and
g. Instructions for Installing and Operating the Propeller.
Original signed by JJP on 9/17/01 Jay J. Pardee Manager, Engine and Propeller Directorate Aircraft Certification Service
TABLE 1
EXAMPLE TRANSPORT AIRPLANE OPERATING SPECTRUM Time Events Stress Stress No. Condition GW Yaw Bank Flaps Load RPM Torque V Event per cycles per cycles per deg. deg. deg. g's % % KCAS sec. 70k hrs 70k hrs Flight Typical Flight 1 Taxi no crosswind 1 0 0 0 1 70 GI 0 100 35000 4.90E+07 466.66667 2 Taxi 15kt crwind 1 0 0 0 1 70 GI 0 100 35000 4.90E+07 466.66667 3 Taxi 25kt crwind 1 0 0 0 1 70 GI 0 100 35000 4.90E+07 466.66667 4 TO roll 1 0 0 15 1 100 100 0-100 40 105000 8.40E+07 800 5 TO rotation 1 0 0 15 1 100 100 110 2 105000 4.20E+06 40 6 Climb A 1 0 0 15 1 100 100 110 30 105000 6.30E+07 600 7 Climb B 1 0 0 15 1 100 100 130 40 105000 8.40E+07 800 8 Climb Spectrum Contained in spectrum data below 10 Cruise Spectrum Contained in spectrum data below 13 Descent Spec. Contained in spectrum data below 15 Approach Spec. Contained in spectrum data below 17 Reverse max 1 0 0 20 1 100 70 120 10 35000 7.00E+06 66.666667 18 Reverse 1/2 1 0 0 20 1 100 35 120 10 35000 7.00E+06 66.666667 Vertical Maneuver Spectrum 19 Vertical Maneuver 2 0 0 0 2.6 100 70 180 15 5.6 1.68E+03 0.016 20 Vertical Maneuver 2 0 0 0 2.2 100 70 180 18 47 1.69E+04 0.1611429 21 Vertical Maneuver 2 0 0 0 1.8 100 70 180 22 576 2.53E+05 2.4137143 22 Vertical Maneuver 2 0 0 0 1.4 100 70 180 25 23838 1.19E+07 113.51429 23 Vertical Maneuver 2 0 0 0 1 100 70 180 30 122354 7.34E+07 699.16571 24 Vertical Maneuver 2 0 0 0 -1.4 100 70 180 25 23838 1.19E+07 113.51429 25 Vertical Maneuver 2 0 0 0 -1.8 100 70 180 22 576 2.53E+05 2.4137143 26 Vertical Maneuver 2 0 0 0 -2.2 100 70 180 18 47 1.69E+04 0.1611429 27 Vertical Maneuver 2 0 0 0 -2.8 100 70 180 15 5.6 1.68E+03 0.016 28 Vertical Maneuver 2 0 0 0 2.6 80 70 200 15 5.6 1.34E+03 0.0128 29 Vertical Maneuver 2 0 0 0 2.2 80 70 200 18 47 1.35E+04 0.1289143 30 Vertical Maneuver 2 0 0 0 1.8 80 70 200 22 576 2.03E+05 1.9309714 31 Vertical Maneuver 2 0 0 0 1.4 80 70 200 25 23838 9.54E+06 90.811429 etc.
Lateral Gust - Yaw Spectrum 72 Lateral Gust - Yaw 2 9.12 0 0 1 100 70 180 0.5 34 3.40E+02 0.0032381 73 Lateral Gust - Yaw 2 7.45 0 0 1 100 70 180 0.5 364 3.64E+03 0.0346667 74 Lateral Gust - Yaw 2 5.23 0 0 1 100 70 180 0.5 7324 7.32E+04 0.6975238 75 Lateral Gust - Yaw 2 3.12 0 0 1 100 70 180 0.5 56398 5.64E+05 5.3712381 76 Lateral Gust - Yaw 2 0 0 0 1 100 70 180 0.5 345626 3.46E+06 32.916762 77 Lateral Gust - Yaw 2 -3.12 0 0 1 100 70 180 0.5 56398 5.64E+05 5.3712381 78 Lateral Gust - Yaw 2 -5.23 0 0 1 100 70 180 0.5 7324 7.32E+04 0.6975238 79 Lateral Gust - Yaw 2 -7.45 0 0 1 100 70 180 0.5 364 3.64E+03 0.0346667 80 Lateral Gust - Yaw 2 -9.12 0 0 1 100 70 180 0.5 34 3.40E+02 0.0032381 etc.
Vertical Gusts etc.
Extreme Maneuvers 213 Limit Yaw 1 32 0 0 1 100 100 170 3.3 1 6.60E+01 0.0006286 214 Limit Pull out 1 0 0 0 3 100 100 150 3.2 1 6.40E+01 0.0006095 215 Rudder kick 1 21 0 0 1 100 100 150 2.3 4 1.84E+02 0.0017524 etc.
FIGURE 1 Establishment of an RPM Restriction for a Cross Wind Condition 0
0.2
0.4
0.6
0.8 1
1.2
1.4 0 200 400 600 800 1000 1200 1400 1600 Rotational Speed - RPM Relative Stress Level Measured Stress Stress adjusted for downward frequency shift Stress adjusted for upward frequency shift RPM avoidance band without accounting for variabiltyRPM avoidance bandwith accounting for variabilty Endurance LImit FIGURE 2 First Blade Mode Propeller Critical Speed Variability 0 20 40 60 80 100 120 0 200 400 600 800 1000 1200 1400 1600 Rotational Speed - RPM Frequency - Hz 950 RPM Accounting for:
Maximum paint and thickness 1050 RPM Accounting for:
Wear to minimum dimensionsNominal Blade 1000 RPM Predicted Blade Frequencies 1P 4P 3P 2P