Appendix 1 specimens with stress concentration factor, as applicable. A reduced S-N curve and the loading spectrum of Paragraph 6b should be used in determining replacement times.
(iii) Figure 1 represents the method of constructing a typical S-N curve from the fatigue test data. Four to six full-scale specimens have commonly been used in past sate-life programs to determine mean S-N curves. The applicant may propose a specific number of specimens that may be evaluated with respect to the proposed methodology in determining the acceptability of the structural substantiation program.
Page A1- 14
Appendix 1 * simple safe life ** enhanced safe life *** X is commonly 3 sigma for pristine specimens but may be
- MEAN CURVE IHROUCH TEST DATA less for flawed specimen testing STRESS (S) OR LOAD (P) WEAN CURVE THROUGH TEST DATA (1 lowed specimens) ** **** XX% varies from 50% to 80% based on expected material and usage variablity
• Denotes bottom of scatter REDUCED CURVE (MEAN MINUS "X" SIGMA) *** - (pristine specimens)*_ - REDUCED CURVE (XXx OF MEAN) ****
NUMBER OF CYCLES FIGURE APX1-1 S-N OR P-N CURVE USAGE Page A1-15
Appendix 1
(iv) Additional coupon tests have been used, when necessary, to identify more completely test scatter due to material properties variability. When the agreed-upon number of specimens is tested at various acceptable load levels, a reduction by three mean standard deviations, sigma (o), necessary to account for material properties, should be made to establish a reduced S-N curve for determination of retirement times. This reduction may take into account the fact that flawed specimens were tested to preclude a dual penalty situation. Reduction by two mean standard deviations rather than three, may be used if justified by appropriate design features such as multiple elements or unmistakable flaw indications or by material properties that provide benign types of failure modes.
(v) "Run-outs" (specimens that do not fail) may be treated in one of three ways: deleted from data (conservative), considered as a failure at maximum test cycles (conservative), or considered in a rational manner to develop a more realistic "o." To further account for variability in fatigue strength not totally accounted for by this reduction in S-N data, the lesser of this reduction, 80 percent of mean, or bottom of scatter, may be used as a limit, where justified, depending on material properties.
(3) The procedures in Paragraph 7a of AC 20-107A provide criteria for substantiating safe flaw growth for composite structure.
(4) Safe crack growth (fracture mechanics) substantiation should show that the damage growth rate under the repeated loads expected in service (between the time at which the damage becomes initially detectable and the time at which the extent of damage reaches the value for residual strength evaluation) provides a practical basis for development of the inspection programs and procedures described in Paragraph 7h of this Appendix. For multiengine load paths, a minimum of three inspection intervals is recommended between the initially detectable damage time and the time when residual strength is reduced to design limit load by crack growth. For single element structures, a minimum of four inspection intervals is recommended. The repeated loads should be defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where it is significant.
(5) For flaw tolerance to achieve an improvement in safety over safe-life for composite structures, the procedures of Paragraph 7a of AC 20-107A are recommended. For flaw tolerance to achieve and improvement in safety over simple safe-life for metallic structures, the following testing criteria are recommended to supplement flaw tolerant safe-life analysis using flawed specimen S-N curves or crack growth analyses using appropriate stress intensity factors and da/dn data from coupon tests:
(i) Flaw Tolerant (Enhanced) Safe-Life Testing Criteria. Test full-scale specimens with flaws or a mix of full-scale and coupon specimens with flaws to obtain Page A1- 16
Appendix 1 S-N data. Plot the data as shown in Figure APX1-1. Utilize the S-N data and loading spectrum of Paragraph 6b in substantiating a crack-free life or in arriving at a replacement time by cumulative damage analysis means. Where practical (as allowed by the number of damaging cycles in the loading spectrum), spectrum testing may be used for fatigue substantiation in lieu of S-N testing followed by analysis. The replacement time established should be included in the airworthiness limitations section of the document established under § 29.1529.
(ii) Fail-Safe Crack Growth (Fracture Mechanics) Substantiation. Test two or more specimens to obtain crack propagation data using either a realistic load spectrum or an accelerated load (spectrum or single) associated with the use of propagation theory and data after cracks have been initiated. Unless a more rational method with an equivalent level-of-satety is applied for, the tollowing methods of setting inspection intervals should be applied. In all cases, the inspection methods and intervals should adequately consider variables such as inspectability, type of inspection, crack growth behavior, and other scheduled maintenance considerations.
(A) For single element (load path) structure, plot the data and set the inspection as shown in Figure APX1-2.
(1) Set the initial inspection at L1/3.
(2) Set the repetitive inspection intervals at L2/4.
Page A1-17
Appendix 1
FLAW SIZE CRITICAL CRACK LENGTH FOR LIMIT LOAD
FIELD DETECTABLE FLAW SIZE
FLAW SIZE CONTROLLED I
BY MANUFACTURING QUALITY LIFE FIGURE APX1-2 CRACK GROWTH FOR SINGLE ELEMENT STRUCTURE Page A1- 18
Appendix 1 (B) For multi-element load path structure:
(1) Test all paths simultaneously with a 0.05-inch thick crack (or size detectable by manufacturing quality control procedures) in the critical element at the start of tests.
(2) Note when field detectable cracking occurs.
(3) Note when one element fails.
(4) Note when the residual strength of the remaining elements decreases to limit load due to crack growth.
(5) From Figure APX1-3, set initial inspection at L, /3.
(6) Set repetitive inspection intervals at (L2 + Lp)/3.
NOTE: If partial failure of the critical element is not detectable, then L2 becomes zero.
Page A1-19
Appendix 1
FLAW SIZE
FAILURE OF
CRITICAL CRACK LENGTH FOR LIMIT LOAD CRITICAL MEMBER
FOR REMAINING MEMBERS I- OR ELEMENT
AFTER FAILURE OF CRITICAL MEMBER MEMBER ELEMENTS
DETECTABLEFLAWSIZE FLAW GROWTH IN CRITICAL REMAINING
FLAW GROWTH LIFE FIGURE APX1-3 CRACK GROWTH FOR REMAINING ELEMENTS
OF MULTIPLE ELEMENT STRUCTURE Page A1- 20
Appendix 1
(iii) Other test and inspection programs may be used rather than those of Paragraphs 7g(5)(i) and 7g(5)(ii), if shown to have comparable or better probability of assuring that a catastrophic fatigue failure will not occur.
(6) The flaw tolerance characteristics can also be shown analytically by rellable or conservative methods such as the following:
(i) By demonstrating quantitative relationships with structure already verified as damage tolerant;
(Il) By demonstrating that the damage would be detected before it reaches the value for residual strength evaluation; or
(iii) By demonstrating that the repeated loads and limit load stresses do not exceed those of previously verified designs of similar configuration, materials, and inspectability.
h. Inspection. Detection of flaws before they become dangerous is the ultimate control in ensuring the flaw tolerance characteristics of the structure. Therefore, the applicant should provide sufficient guidance information to assist operators in establishing the frequency, extent, and methods of inspection of the critical structure, and this kind of information should, under § 29.571(a)(2), be included in the maintenance manual required by § 29.1529. Due to the inherent, complex interactions of the many parameters affecting flaw tolerance, such as operating practices, environmental effects, load sequence on flaw growth, and variations in inspection methods, related operational experience should be taken into account in establishing inspection procedures. Comparative analysis can be used to guide the changes from successful past practice, when necessary. Therefore, maintenance and inspection requirements should recognize the dependence on experience and should be specified in a document that provides for revision as a result of operational experience, such as the one containing the operator's FAA/AUTHORITY-approved structural inspection program developed through the Maintenance Review Board (MRB) procedures for FAR Part 121 operators.
COMBINATION OF REPLACEMENT TIME AND FLAW GROWTH EVALUATION.
It may be possible to extend the replacement time of safe-life components that exhibit limited flaw tolerance capability by using a combination of the safe-life and flaw growth characteristics as described elsewhere in this Appendix and by assigning both a replacement time and inspection period to these components. The replacement time may then be based on the combined probability of not initiating a fatigue crack at or before the replacement time and the probability that the crack, if initiated, will be detected prior to catastrophic failure or loss of limit load (or maximum attainable load, whichever is less) carrying capability. The probability of detection should be based on consideration of the inspection effectiveness, the inspection intervals, and the fatigue life remaining after an obvious partial failure. A lower strength reduction factor Page A1-21
Appendix 1 commensurate with this probability of detection may then be used in the determination of the replacement time.
Page A1-22
Appendix 1 Attachment 1 - FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED
1. GROUND CONDITIONS.
a. Normal start.
b. Rapid increases of RPM on ground to maximum power-on RPM of main rotor.
c. Taxiing with full cyclic control.
d. Landing run (if applicable).
e. Braking (if applicable).
f. Normal shutdown.
g. Special ground checks (if applicable).
2. IN-GROUND-EFFECT (IGE) MANEUVERS.
a. Hovering.
(1) Steady with rotor at maximum side of RPM tolerance.
(2) Steady with rotor at minimum side of RPM tolerance.
(3) 90° right turn.
(4) 90° left turn.
(5) Control reversal.
(i) Longitudinal.
(ii) Lateral.
(iii) Rudder.
(6) Sideward flight.
(1) Right.
(її) › Left.
(ili) Rearward flight.
b. Maneuvering.
(1) Jump takeoff.
(2) Normal takeoff and accelerate to climb airspeed.
(3) Normal approach and landing.
(1) Multiengine.
(ii) One-engine-inoperative.
(4) Full autorotational landing.
Page A1-23
Appendix 1
3. FORWARD FLIGHT-POWER ON.
a. Level flight.
(1) 40 percent VH.
(i) Minimum side of main rotor RPM tolerance (RPM +)
(il) Maximum side of main rotor RPM tolerance (RPM -)
(2) 60 percent VH.
(i) (RPM +)
(il) (RPM -)
(3) 80 percent VH.
(i) (RPM +) (I!) (RPM -)
(4) V
(i) (RPM +)
(ii) (RPM -)
(5) VNE:
(1) (RPM +) (i1) (RPM -)
b. Maneuvers.
(1) Full power climbs.
(i) All engines operative.
(ii) One-engine-inoperative.
(2) Cyclic pull-ups.
(i) 60 percent VH.
(ii) 90 percent VH.
(3) Normal acceleration from climb airspeed to 90 percent VH.
(4) Turns.
(1) Right at 60 percent Vu and 90 percent VH.
(ii) Left at 60 percent Vu and 90 percent VH.
(5) Control reversals at 90 percent VH.
(1) Longitudinal.
(il) Lateral.
(ili) Rudder.
(6) Deceleration from 90 percent Vu to descent airspeed.
(7) Part power descent.
(i) All engines.
(il) One engine out.
4. POWER TRANSITIONS.
a. All engines operating to one engine out.
(1) In full power climb.
(2) At 90 percent VH.
Page A1- 24
Appendix 1
b. One engine out to all engines operating in powered descent.
c. All engines operating to autorotation.
(1) At 60 percent VH.
(2) At maximum forward transition speed.
d. Stabilized autorotation to all engines operating at normal autorotation airspeed.
5. AUTOROTATION.
a. Stabilized.
(1) At 70 percent VNE:
(2) At NE•
b. Turns at 70 percent and 100 percent VNE:
(1) Right.
(2) Left.
c. Cyclic pull-up.
d. Control reversals.
(1) Longitudinal.
(2) Lateral.
(3) Rudder.
The flight and ground conditions of this table may be used as practical in developing operating load spectra for the range of projected aircraft usage such as:
(1) Short missions with high-power cycles and ground-air-ground (GAG) cycles (such as used in logging operations).
(2) Medium range missions of 1 to 2 hours that include three or more GAG or LO-HI power cycles per mission.
(3) Long range missions that consider the longest practical range of the aircraft and consider two or more GAG (or high power) cycles per mission (such as used for remote offshore drilling platforms).
If aircraft usages that produce a large number of high-power cycles (approximately 50 per hour) or high GAG cycles (approximately 25 per hour) such as logging operations are not included in the fatigue tolerance substantiation, this should be noted in the RFM and Maintenance Manual airworthiness limitations section (reference Paragraph 6a).
Page A1-25