b. Stress Survey Tests. When testing indicates a stress peak is within the propeller operating range, further testing may be necessary to determine if rpm restrictions, such as placards, will be needed to avoid the resonant conditions. The test program should be modified to obtain further detail regarding the extent of the stress peak. When testing indicates a stress peak is just beyond the propeller operating range, conduct testing sufficiently above the operating range to determine the maximum stresses that could occur from potential overspeed conditions. Overspeed conditions may result from conditions such as overspeed governor checks, transients, or tachometers that are not properly calibrated.
c. Identification of the Forces. When a significant resonant condition is found during the vibratory stress survey, conduct sufficient testing to permit the identification of the vibratory force acting on the blades. Typical forces are the firing impulses from a reciprocating engine;
aerodynamic forces from ground cross-winds; and aerodynamic forces from the wake of the airplane on pusher installations. The aerodynamic forces act at integer multiples of the propeller rotational speed known as “P” (per-propeller revolution) orders. The term “2P” means two force cycles per-propeller revolution. The reciprocating engine forces act at multiples of the engine speed and may be integer or fractional multiples of the rotational speed. The engine forces are known as “E” (per-engine revolution) orders. The significant vibratory forces are identified to assess the potential variability in the magnitude of the force to determine the extent of any operating restriction. For example, the stress peak caused by ground cross-winds will increase significantly as the cross-winds increase. Therefore, operating in a 15 knot rear-quartering crosswind may be acceptable, while operating in winds above 15 knots may not.
d. Identification of the Blade Modes. When a significant resonant condition is found during the stress survey, conduct sufficient testing to permit the identification of the blade natural frequency and vibratory mode shape. Do this by comparing the measured resonance frequency with analytical predictions, test results, and previous test experience. In addition, the measured strain distribution on the blade may also be used, either directly or filtered at the resonant frequency, to provide further information to identify the blade vibratory mode shape.
Use the significant blade natural frequencies to assess the potential variability in the natural frequency to determine the extent of any operating restriction. For example, the natural frequencies of an aluminum blade at the maximum thickness permitted within the manufacturing tolerances will be different from the natural frequencies of this blade after it has worn to the minimum thickness permitted in service. The shift in natural frequency will shift the location of the resonant condition and resulting stress peak. Also, composite blade natural frequencies may be lower for blades that have been repaired, have multiple coats of paint, or both.
e. Operating Restrictions.
(1) When an operating restriction, such as a placard, is required to avoid a stress peak due to a resonant condition, the maximum and minimum permissible rpm values should consider such items as manufacturing variability, service wear, repairs, and the excitation force variability. These factors may be assessed by test, analysis, service experience or any other suitable method. Placards are typically set 100 rpm higher and lower than the peak stress rpm or as shown by testing.
(2) For example, figure 1-1 in Appendix 1 shows the measured vibratory stress for a blade during an rpm traverse on the ground in a cross-wind. A stress peak is found at 1000 rpm.
A review of calculated blade frequencies in the critical speed diagram in figure 1-2 in Appendix 1 shows that this stress peak is due to a 2P aerodynamic force and a first bending mode natural frequency critical speed. If the rpm restriction was set based on only the measured data, figure 1-1's upper and lower bounds would be set to 1100 and 900 rpm, respectively. This restricted operating band is not sufficient when blade manufacturing and wear limits are taken into account. Figure 1-2 shows how calculated blade frequencies shift the stress peak up and down to account for frequency variability. The critical speed shift is then used to shift the stress peak in figure 1-1 assuming that the shape of the stress peak remains the same. A blade with the maximum manufacturing tolerances and paint will lower the critical speed to 950 rpm, and blade wear will increase the critical speed to 1050 rpm. This increases the upper and lower bounds of the rpm restriction to 1150 and 850 rpm, respectively.
(3) In addition, assess the magnitude of the stress amplitude to determine if further restrictions are needed. The operating restriction may prohibit operation above a maximum rotational speed or may prohibit continuous operation in a restricted rotational speed range to enable the pilot to pass through that range.
f. Continued Airworthiness Monitoring. We r ecommend the following procedures to maintain rpm restrictions and vibratory stress levels.
(1) Periodically check and calibrate the airplane tachometer.
(2) Check and replace the engine vibration dampers at engine overhaul, if needed.
(3) Assess the natural frequency of propeller blades at intervals over the life of the fleet. This assessment should determine if any service use dependent changes in the blade's natural frequency exist that were not considered during propeller certification. This may be done by static frequency tests or similarity to existing designs when service use dependent changes are known.
(4) Evaluate the natural frequency of propeller blades to assess the effect of repairs on frequency and rpm restrictions.
9. P ropeller Flutter. The propeller should be shown to be free from the harmful effects of flutter. This is done by evaluating the measured vibratory response and showing that flutter does not exist within the operational envelope of the propeller. The harmful effects of flutter are high blade loads that result in unacceptable fatigue life. Propeller flutter is typically associated with torsional vibration of a blade. It typically occurs during high power operation on the ground or during reverse operation on landing. If flutter is found within the operational envelope of the propeller, further testing and fatigue evaluation may be necessary to show that the flutter conditions will not cause harmful effects within the operational life of the propeller. Limitations may be needed to avoid flutter or to limit exposure to flutter.
AC 20-66B CHG 1
10. Similarity. The propeller vibration stresses may be determined by similarity. To determine similarity, the operating environment is compared from one airplane to another. When shown to be sufficiently similar, the vibratory load measurement and the fatigue evaluation of the baseline airplane apply to the target airplane. Similarity is generally not used to determine propeller vibration loads on large multi-engine airplanes or acrobatic airplanes, due to the inability to accurately evaluate the aerodynamic environment. Approvals based on similarity are only valid when based on a direct comparison to measured data. If airplane “B” is approved based on similarity to airplane “A,” airplane “C” cannot be approved based on similarity to airplane “B;”
only comparisons to airplane “A,” the baseline airplane, are valid.
a. Overview. Showing 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 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. Evaluate the propeller aerodynamic environment 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.
AQ is the product of angle of flow into the propeller (A) 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 AQ is proportional to the propeller vibratory loads in flight due to angular inflow. For more information on aerodynamic excitations, refer to Appendix 2.
c. General Considerations.
(1) An approved propeller vibration stress survey and evaluation showing compliance with §§ 23.907 (amendment 63 and earlier), 23.2400 (amendment 64), or 25.907 conducted on the baseline airplane forms 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 proposals become less reliable 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.
AC 20-66B CHG 1
(3) The baseline airplane and target airplane should have engines and propellers that have equivalent vibration characteristics.
(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 considered to have equivalent vibration characteristics.
(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.
11. Note 9 Vibration Approvals. Some propeller TC data sheets contain a section called “Note 9.” This note simplifies the vibration compliance for aluminum-bladed propellers under §§ 23.907 (amendment 63 and earlier) and 23.2400 (amendment 64) for normal category, single reciprocating engine airplanes in a 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 CMCB 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:
AC 20-66B CHG 1
(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 CMCB.
(6) The propeller contains conventional aluminum blades.
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.
12. 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. Consider the accumulation of this damage 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. This AC provides safe- life and damage tolerance approaches to fatigue evaluation. These approaches are suitable for 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:
(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 gauge 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. Assess the components to establish appropriate damage criteria in relation to the ability to be inspected and damage characteristics from initial detectability to fracture.
(5) Damage Accumulation. Select appropriate and substantiated damage accumulation algorithms. Examples are Miner’s rule for safe-life calculations or a crack or damage growth 16
AC 20-66B CHG 1 algorithm for damage tolerance calculations. Verify the damage accumulation algorithm by previous testing, past experience, and acceptable published data, when available.
(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 (ALS) of the Instructions for Continued Airworthiness (ICA). If the vibration survey and fatigue evaluation indicate that certain operating conditions or ranges should be limited, include those limitations in the appropriate propeller and airplane manuals (refer to paragraph 13 of this AC).
(7) Component Degradation and Repair. The fatigue evaluation and critical speed 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 35.37-1B. 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 operating 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 the following normal flight conditions that occur with each flight:
(a) Take-off;
(b) Climb;
(c) Cruise;
(d) Descent. When investigating the actual vibration behavior of each propeller, the applicant should include the operating conditions that correspond to descent with the power levers at flight idle position and with airspeeds near the maximum operating limit speed (VMO) for airplanes equipped with turboprop engines. Experience has shown that such conditions may cause unacceptable cyclic loads and vibrations on some parts of the propeller. Since aerodynamic loads differ depending on the position of the engine-propeller assembly on the airplane, the applicant should investigate the propellers’ vibration behavior at all enginepropeller assembly positions.
(e) Approach; and
(f) Landing and reverse.
AC 20-66B CHG 1
(2) The elements of the operating spectrum also include the following transient airplane flight conditions:
(a) Maneuvers (for example, banked turns, side-slip, pull-ups, push-overs, or rudder kicks);
(b) Gusts;
(c) Special flight conditions specific to a mission (for example, fire-fighting, acrobatic, or agricultural);
(d) Emergency conditions;
(e) Airplane limit load conditions; and
(f) Training maneuvers.
(3) The operating spectrum also includes the following ground operating conditions:
(a) Taxi;
(b) Operation in cross winds; and
(c) Maintenance checks, such as governor or feather checks.
(4) When applicable, special operations such as water bomber or fire bomber operations should also be included.
(5) 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, base the spectrum information on the design assumptions and design and service experience for the intended airplane and engine application.
(6) The number of occurrences and duration of each operating condition are also associated with the airplane operating spectrum. Table 1-1 in Appendix 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, such as an extreme yaw or extremely high “g” maneuver, occur once in the life of the airplane. The propeller fatigue evaluation should consider all of these operating conditions.
(7) The airplane operating spectrum should include loads associated with low cycle fatigue (LCF) 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 1-3 in 18
AC 20-66B CHG 1 Appendix 1 illustrates in schematic form the load variation and the GAG cycle of a normal flight.
(8) The airplane applicant should define the airplane operating spectrum so it can be used to 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 are 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.
(9) After identifying the airplane operating spectrum, determine the stress or load levels at each of the operating conditions, 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 calculated 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 35.37-1B.
(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). Use a scatter factor of three or greater for metallic structure, unless there is substantial justification for a lower scatter factor.
Mandatory replacement times should be established for parts with safe-lives as appropriate.
Mandatory replacement times must be included in the propeller airworthiness limitation section of the ICA’s for both the propeller and the airplane.
(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 used, 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.
(4) When all stresses are below the endurance limits established for the component, the component is said to have unlimited life. In addition, when the applicant shows that the component will be safely retired from service for reasons other than fatigue before a safe-life of 19
AC 20-66B CHG 1 70,000 hours for commuter and transport airplanes and of at least 3 times the service life for general aviation airplanes, 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 inaccurate, the safe-life should be established using a substantiated scatter factor. A scatter factor of 10 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 loads during 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. W hen damage tolerance methods are applied, the component should be designed using damage tolerance and fail-safe principles. The damage tolerance methods are based on the principles 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 damage before the component reaches the end of life. The inspection method should also be evaluated to determine the probability of detection (POD). Inspection methods should have a POD of 90% with 90% confidence. When the POD is less than 90% 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 ALS of the ICA 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.
(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 1-4. AC 35.37-1B 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. Verify the applicability of the damage tolerance assessment 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. Obtain flaw/damage growth data by repeated load cycling of intrinsic flaws or mechanically introduced damage. Validate the damage growth model by tests of full-scale components.
(4) Residual Strength. Demonstrate the residual strength of the component on fullscale damaged components at the end of life condition. Also, show 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.
13. 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;
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 21
g. Instructions for Installing and Operating the Propeller.
14. A dvisory Circular Feedback Form. For your convenience, the AC Feedback Form is the last page of this AC. Note any deficiencies found, clarifications needed, or suggested improvements regarding the contents of this AC on the Feedback Form.