CHAPTER 5. ROTORBURST REQUIREMENTS
5-1. What are the requirements for turbine or turbopropeller installations in small airplanes?
a. Compliance with 14 CFR, part 23, § 23.903(b)(1) requires the applicant to take design precautions to minimize the hazards to the airplane, including the airframe, due to an engine rotor failure. A Small Airplane Direct orate policy letter dated April 7, 1989, defines “minimize” in the context of § 23.903 as “to reduce to a minimum, decrease to the least practical amount. The 'least practical amount' obviously has to include consideration of what is attainable with current technology and materials.”
b. Industry practice for part 23 turbine-powered airplanes is to use discrete source damage analysis to evaluate design precaution options. Applicants have used the discrete source damage evaluations to select the most effective, appropriate design precautions and to substantiate that the selected design precautions minimize the rotorburst hazard to the least practical amount. The design pre cautions selected by these applicants to show compliance to § 23.903(b)(1) have provided sufficient residual strength capability to demonstrate compliance with the discrete source damage requirements in 14 CFR, part 25, § 25.571(e)(3).
5-2. I thought there was not a discrete source damage requirement in part 23?
a. Part 23 does not include a discrete source damage requirement in the fatigue regulations. However, small airplane applicants have used, and the FAA has accepted, discrete source damage analysis as a quantitative method of substantiating that selected design precautions minimize the rotorburst hazard to the least practical amount.
b. AC 20-128A, Design Considerations for Minimizing Hazards caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor Failure, provides a means, but not the only acceptable means, of showing compliance to § 23.903(b)(1). The AC uses discrete source damage criteria to include airframe hazards in a safety assessment and risk analysis. The correct interpretation of the AC is that the discrete source damage evaluation is required for all categories of airplanes. If you choose to show compliance using AC 20- 128A, include airframe hazards, via discrete source damage analysis, in the rotorburst safety assessment and risk analysis.
CHAPTER 6. FLIGHT WITH KNOWN CRACKS
6-1. What is the purpose of this policy?
This policy is provided as guidance for Aircraft Certification Office (ACO) engineers when the FAA’s Flight Standards Service requests engineering support. The Flight Standards Service may request support from an ACO when addressing a known cracking situation in an individual airplane.
6-2. What is the policy for operati onal flight with known cracks?
It is the Small Airplane Directorate’s policy to not allow continued operational flight with known cracks in primary structure. An airplane with a known crack in the primary structure no longer meets its type design and may no longer possess its type design strength.
6-3. Are there any exceptions to this policy?
a. The exceptions to this policy, to be taken on a case-by-case basis, are the following:
(1) Substantiation that the cracks are not in primary structure.
(2) Substantiation of the ability of single load path structure with known cracks to carry ultimate loads. Only in unusual circumstances, such as the difficulty of an operator in obtaining replacement parts, will this be allowed. The time allowed for replacement or repair must be as short as practical.
(3) Substantiation that the cracks are in fail-safe structure. The ability to sustain ultimate load with the maximum permissible crack must be substantiated. Temporary repairs, such as "stop drilling," should be specified. The conditions of a temporary repair, and the time allowed before a permanent repair is required, may be less restrictive than for single load path structure.
b. The Aircraft Certification Office (ACO) responsible for the airplane’s type certificate should make each specific approval for flight with known cracks. The ACO issuing the approval should coordinate the approval with the Small Airplane Directorate Standards Office.
6-4. Which airplanes are covered by this policy?
This policy applies to all normal, utility, acrobatic, or commuter category airplanes, regardless of certification basis. This policy provides accepta ble methods for any airplane with a known crack, even those whose certification basis does not include specific fatigue requirements, to substantiate a limited time for continued operational flight. The policy does not expand current fatigue requirements to existing designs; it provides methods to substantiate that, in certain unusual circumstances, an airplane with a known crack may be acceptable for limited operational flight.
6-5. What will the FAA consider before allowing operational flight with known cracks?
Continued operational flight with known cracks in a structure whose failure may be catastrophic, in general, is not permissible. However, there are certain unusual circumstances under which flight is permissible for a limited period of time with known cracks in the structure. Any approval for flight with known cracks should meet all the following criteria.
a. There must be an unusual need for allowing flight with known cracks. Unusual needs are those that arise from such things as the scarcity of parts or operator hardship. For example, unavailability of parts or tooling to support an immediate part replacement would qualify as an unusual need. Another would be the need for operational flight to a repair facility from an area where insufficient technical or maintenance resources are available to perform a repair. Deferral of corre ctive action should not be approached as either business as usual or as a matter of convenience to operators.
b. The level of safety provided by the certification basis must not be compromised by the allowance to operate with a known crack. At all times during the expected crack growth phase, the airplane must meet its limit and ultimate load requirements with the crack present. Part modification, replacement, or repair must be implemented if the structure's strength, with the crack present, is anticipated to fall below these load requirements before the next inspection of the structure. Specifically, this includes the following:
(1) Cracked structure must support limit load without detrimental permanent deformation (§ 23.305(a)).
(2) Cracked structure must support ultimate load for three seconds without failure (§ 23.305(b)).
(3) Airplanes certified as fail-safe should remain fail-safe with the damage present.
This means that the cracked structure must be able to bear ultimate load and meet the fail-safe criterion specified in the airplane’s certification basis.
(4) Any other certification requirement identified by the ACO as being potentially violated by repair deferral should be addressed.
(5) Single failure requirements remain in effect with the damage present (e.g., § 23.629 (g)(h)).
c. There should be a high degree of confidence in the evaluation used to justify the particular case. Regardless of the certification basis of the airplane, a fracture mechanics assessment must be performed for the cracked structure before further flight. The evaluation should determine that the material, location, and type of damage is such that a reliable fracture mechanics assessment can be performed. Th e following items are especially critical to a thorough evaluation:
(1) An investigation into the cause of the crack. This investigation should include a complete characterization of the crack location, size, and shape. The investigation should estimate the time of crack initiation and growth rate, using microscopic mapping techniques when possible. Identify the presence of any corrosion or manufacturing or material defect.
Review the results of any previous inspections of the cracked part to find any indications of abnormalities. The investigation should also review the assumptions used in the cracked structure’s design. This shoul d include a comparison of the load spectra used in the type certification fatigue evaluation and actual usage, validation of internal loads, and verification of material properties.
(2) Static strength and fracture mechanics analysis. The fracture mechanics analysis must provide a high degree of confidence that the crack will not grow faster than the predicted rate with the expected usage until the next inspection, or until it is repaired. All residual strength substantiations should be shown by a conservative analysis, by test, or by test combined with analysis. Analysis is allowed only where experience shows that analysis is reliable.
(3) Inspection intervals.
(a) This assessment must establish a viable inspection program for the cracked structure. The inspection program should provide a high level of confidence that the crack will not be allowed to grow to a size that would compromise the safety provided by the certification basis. The probability of detection and reliability of the inspection must be high. The inspection method should be suitable for the situation, and the cracked structure should be adequately accessible. The inspection progr am, along with any other limitations, must be placed in the mandatory section of the airplane’s maintenance manual for the period of time in which the cracked structure remains in operation.
(b) Inspection intervals should be conservatively determined based on approved crack growth data and a fracture mechanics based evaluation. The inspection interval is established by calculating the crack growth time from the detected crack size (or maximum of detected size range for non-visual) for the Non-Destructive Inspection (NDI) method used to detect critical crack size and dividing that number of flight hours or flights by a conservative value. Inspection intervals should provide for at least two opportunities to detect crack growth before the crack size reduces the capability of the structure below that provided by the certification basis.
(4) Temporary intervention. If possible, some intervention, such as stop drilling of the crack tip, should be implemented to slow crack growth (stop drilling is not considered a repair).
(5) Single-Load path structures. Flight with known cracks is strongly discouraged for structure where applied loads are eventually distributed through a single member, the failure of which would result in the loss of the structural capability to carry the applied loads. If cracking is found in this type of structure, the part must be replaced or repaired before further flight. Exceptions may be permitted only if all other criteria of this policy are met, stringent inspection requirements are specified, and a conservative maximum permissible crack size able to sustain ultimate load is specified. A combination of a conservative fracture mechanics analysis correlated to test results, or conservative fracture mechanics analysis with added scatter factors, must be performed to allow continued operation with cracks.
(6) Multiple-Load path structures. For redundant structures in which, with the failure of individual elements, the applied loads would be safely distributed to other load carrying members, adjacent structure must be inspected and be free of cracks or other damage.
Various combinations of analysis and test, including that provided at the time of original certification, may be considered as ample substantiation.
(7) Structures with multiple-site damage, corrosion and stress corrosion cracks.
Continued operational flight with known cracks in instances of multiple-site damage, cracks in the presence of corrosion, or stress corrosion cracks should be avoided.
d. The approval should provide a limited time for continued flight with known cracks. Under no circumstances can allowance for operation with known cracks be considered more than a temporary condition. The limited period must be based on the need for the deferral of the repair, replacement, or modification. Once the unusual need for the deferral has passed, the deferral must be eliminated, and the part repaired or replaced.
e. The ACO approving flight with known cracks should coordinate each approval with the Small Airplane Directorate Standards Office. The approval to allow flight with known cracks must be clearly made, documented, and approved by the FAA before any flight with known cracks. The ACO responsible for the type certificate should make each specific
approval for flight with known cracks in coordination with the Small Airplane Directorate Standards Office.
REFERENCES
1. FAA Report No. ACE-100-01, Fatigue Evaluation of Empennage, Forward Wing and Winglets/Tip Fins on Part 23 Airplanes, dated February 15, 1994.
2. DOT/FAA/CT-91/20 General Aviation Aircraft – Normal Acceleration Data Analysis and Collection Project, dated February 1993.
3. FAA Report No. AFS-120-73-2, Fatigue Evaluation of Wing and Associated Structure on Small Airplanes, dated May 1973.
4. FAA Memorandum, Method of Establishing Flight Load Spectra for Safe-Life Fatigue Analysis, dated March 25, 2005.
5. A Review of Fatigue Scatter Factors with Particular Reference to Civil Aircraft, M. B.
Benoy, Civil Aviation Authority of Australia, July 1978.
6. DOT/FAA/AR-MMPDS-01, Metallic Materials Properties Development and Standardization, dated January 2003.
7. FAA Memorandum, Example Cumulative Damage Calculation for Safe-Life Fatigue Analysis, dated February 7, 2005.
8. User’s Instructions for Computer Program to Calculate Fatigue Safe-Life (Unfactored) for Small Airplanes, dated March 1996.
9. Metal Fatigue, J. A. Pope (Editor), Chapman and Hall Ltd, dated 1959.
10. A Critical Review of Strategies Used to Deal with Metal Fatigue, by Bob Eastin, Federal Aviation Administration, Presented at the International Committee on Aeronautical Fatigue (ICAF) 2003.
11. Commercial Airplane Certification Process Study, An Evaluation of Selected Aircraft Certification, Operations, and Maintenance Process, Federal Aviation Administration, March 2002.
Note: You may obtain References 1, 3, 4, 7, and 8 from the Federal Aviation Administration, Small Airplane Directorate Standards Office, telephone number 816-329-4111.
Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA A1-1. What flight and ground load spectra are presented in this appendix?
The following list summarizes the flight and ground load spectra presented in this appendix.
Flight Spectra, General Usage, Single-Engine Unpressurized Description Graphic Data Tabulated Data Comments Gust spectra for single-engine, unpressurized operations, including basic instruction, personal, executive, and acrobatic usage.
Figure A1-1 Table A1-1 These spectra should also be used for pressurized single-engine airplanes that spend a significant amount of flight time at low altitudes.
Maneuver spectra for singleengine basic instruction usage.
Figure A1-2 Table A1-2 Maneuver spectra for singleengine personal usage.
Figure A1-3 Table A1-3 An airplane in the personal usage category has a single, reciprocating engine with 185 horsepower or less.
Maneuver spectra for singleengine executive usage.
Figure A1-4 Table A1-4 An airplane in the executive usage category has a single, reciprocating engine with more than 185 horsepower. The executive usage category also includes unpressurized, singleengine turboprop airplanes.
Maneuver spectra for singleengine acrobatic usage.
Figure A1-5 Table A1-5 Applicable to typical acrobatic category airplanes, nz = +6g / -3g Flight Spectra, General Usage, Twin-Engine Unpressurized Description Graphic Data Tabulated Data Comments Gust spectra for twin-engine, unpressurized operations, including instruction, and general usage.
Figure A1-6 Table A1-6 These spectra should also be used for pressurized twin-engine airplanes that may spend a significant amount of flight time at low altitudes, 7,000 feet and below. An example of this type of operation is short flight duration commuter and air taxi operations.
Maneuver spectra for twinengine instruction usage.
Figure A1-7 Table A1-7 Maneuver spectra for twinengine general usage.
Figure A1-8 Table A1-8 These spectra should also be used for pressurized and unpressurized twin-engine airplanes that may be operated in short flight duration commuter and air taxi operations.
A1-1
Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)
Flight Spectra, General Usage, Single-Engine and Twin-Engine Pressurized Description Graphic Data Tabulated Data Comments Gust spectra for pressurized usage.
Figure A1-9 Table A1-9 Maneuver spectra for pressurized usage.
Figure A1-10 Table A1-10 Flight Spectra, Special Usage – Agricultural (Aerial Application) Description Graphic Data Tabulated Data Comments Gust spectra for agricultural or aerial application usage.
Figure A1-11 Table A1-11 Maneuver spectra for agricultural or aerial application usage.
Figure A1-12 Table A1-12 Flight Spectra, Special Usage – Survey (Pipeline Patrol) Description Graphic Data Tabulated Data Comments Gust spectra for low-level survey or pipeline patrol usage.
Figure A1-13 Table A1-13 Maneuver spectra for lowlevel survey or pipeline patrol usage.
Figure A1-14 Table A1-14 Ground Spectra Description Graphic Data Tabulated Data Comments Landing Impact Figure A1-15 Taxi Figure A1-16 A1-2
Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED) A1-2. How do I compute the gu st limit load factor used in the gust load spectra?
a. The equation for computing the gust load limit load factor has evolved through the regulatory history of the Civil Air Regulation Part 3 (CAR 3) and 14 CFR Part 23. These changes to the gust load formula incorporate the improving knowledge of airplane response to atmospheric turbulence. The gust load spectra presented in this AC are normalized to the gust load formula in use before Amendment 23-7 to 14 CFR Part 23. Regardless of the certification basis of your airplane, you must calculate a gust load limit load factor (for use in developing fatigue loading spectra only) using the same equation used to derive the gust spectra.
b. Formula for computing the incremental gust load limit load factor:
U K V m anLLF = W498 S where U = 30.0, nominal gust velocity in feet per second 1 1 ⎛ 4 for W/S < 16 lbs / ft2K = ⎟ ⎠ ⎞⎜ ⎝ W S2
2.67 for W/S > 16 lbs / ft2= 1.33
⎛⎜ ⎝ − 3 W ⎞⎟ ⎠ 4 S
W / S = Wing loading at maximum weight, lbs / ft 2 V = Airplane structural design cruise speed, VC, KEAS m = Wing lift curve slope, C Lα, rad-1 A1-3. How are gust and maneuver load spectra ‘normalized’?
a. We normalized the gust spectra to the incremental gust limit load factor computed in Paragraph A1-2. We expressed the gust spectra in terms of the gust load factor ratio:
an = IncrementalGustLoad Factorat OperatingWeight anLLF IncrementalDesignLimitGustLoad FactoratMaximumGrossWeight
b. We normalized the maneuver spectra to the incremental maneuver limit load factor.
We expressed the maneuver spectra in terms of the maneuver load factor ratio:
an = IncrementalManeuverLoad Factorat OperatingWeight anLLF IncrementalDesignLimitManeuverLoad FactoratMaximumGrossWeight A1-3
Acceleration Fraction, an / anLLF Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)
Cumulative Frequency of Exceedance per Nautical Mile 100
10-1
10-2
10-3
10-4
10-5
10-6
Figure A1-1 Gust Spectra For Single-Engine Unpressurized Operations Including Basic Instruction, Personal, Executive, and Acrobatic Usage See Table A1-1 for Tabulated Spectra 8
1.5
1.0 -1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2 A1-4
Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)
Table A1-1 Gust Spectra For Single-Engine Unpressurized Operations Including Basic Instruction, Personal, Executive, and Acrobatic Usage Acceleration Fraction Cumulative Frequency of Exceedance per Nautical Mile Acceleration Fraction Cumulative Frequency of Exceedance per Nautical Mile
0.10 7.99040E-01 -0.10 8.86903E-01
0.15 2.39762E-01 -0.15 2.90400E-01
0.20 8.26537E-02 -0.20 7.93185E-02
0.25 2.59957E-02 -0.25 2.50615E-02
0.30 9.43467E-03 -0.30 8.64457E-03
0.35 3.74392E-03 -0.35 3.13716E-03
0.40 1.56672E-03 -0.40 1.17560E-03
0.45 6.76810E-04 -0.45 4.49619E-04
0.50 2.96910E-04 -0.50 1.74053E-04
0.55 1.31101E-04 -0.55 6.77344E-05
0.60 5.79043E-05 -0.60 2.63806E-05
0.65 2.55611E-05 -0.65 1.02760E-05
0.70 1.12764E-05 -0.70 4.00187E-06
0.75 4.97072E-06 -0.75 1.55782E-06
0.80 2.18708E-06 -0.80 6.05622E-07
0.85 9.58226E-07 -0.85 2.34685E-07
0.90 4.15769E-07 -0.90 9.01733E-08
0.95 1.76305E-07 -0.95 3.38733E-08
1.00 7.05919E-08 -1.00 1.19395E-08 A1-5
Acceleration Fraction, an / anLLF Appendix 1 APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)
Cumulative Frequency of Exceedance per Nautical Mile 100
10-1
10-2
10-3
10-4
10-5
10-6
Figure A1-2 Maneuver Spectra For Single-Engine Basic Instruction Usage See Table A1-2 for Tabulated Spectra 8
1.5
1.0 -1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2 A1-6