The base scatter factors applicable to test results are: BSF1= 3.0, and BSF2 = (see paragraph 7b(5) of this AC). If the applicant can meet the requirements of paragraph 7b(3) of this AC, he may use BSF1 or, at his option, BSF2.
(2) The base scatter factor, BSF1, is associated with test results of one representative test specimen.
() Justification for use of BSF1. BSF1 may only be used if the following criteria are met:
(1) Understanding of load paths and failure modes. Service and test experience of similar in-service components that were designed using similar design criteria and methods should demonstrate that the load paths and potential failure modes of the components are well understood.
(ії) Control of design, material, and manufacturing process quality.
The applicant should demonstrate that his quality system (e.g., design, process control, and material standards) ensures the scatter in fatigue properties is controlled, and that the design of the fatigue critical areas of the part account for the material scatter.
(111) Representativeness of the test specimen.
(A) The test article should be full scale (component or subcomponent) and represent that portion of the production aircraft requiring test. All differences between the test article and production article should be accounted for either by analysis supported by test evidence or by testing itself.
B Construction details, such as bracket attachments, clips, etc., should be accounted for, even though the items themselves may be non-load bearing.
Points of load application and reaction should accurately reflect those of the aircraft, ensure correct behavior of the test article, and guard against uncharacteristic failures.
(D) Systems used to protect the structure against environmental degradation can have a negative effect on fatigue life and therefore should be included as part of the test article.
(4) Adjustments to base scatter factor BSF1. Having satisfied the criteria of paragraph 7b(3), justifying the use of BSF1, the base value of 3.0 should be adjusted to Par 7 13
account for the following considerations, as necessary, where not wholly taken into account by design analysis. As a result of the adjustments, the final scatter factor may be less than, equal to, or greater than 3.0.
Material fatigue scatter. Material properties should be investigated up to a 99% probability of survival and a 95% level of confidence.
(11) Spectrum severity. Test load spectrum should be derived based on a spectrum sensitive analysis accounting for variations in both utilization (i.e. aircraft weight, cg etc.) and occurrences/size of loads. The test loads spectrum applied to the structure should be demonstrated to be conservative when compared to the usage expected in service.
(iii) Number of representative test specimens. Well established statistical methods should be used that associate the number of items tested with the distribution chosen, to obtain an adjustment to the base scatter factor.
(5) If the applicant cannot satisfy the intent of all of paragraph 7b(3) of this AC, BSF2 should be used.
(1) The applicant should propose scatter factor BSF2 based on careful consideration of the following issues: the required level of safety, the number of representative test specimens, how representative the test is, expected fatigue scatter, type of repeated load test, the accuracy of the test loads spectrum, spectrum severity, and the expected service environmental conditions.
In no case should the value of BSF2 be less than 3.0.
(б) Resolution of test loadings to actual loadings. The applicant may use a number of different approaches to reduce both the number of load cycles and number of test set-ups required. These include, but are not limited to, spectrum blocking (e.g., a change in the spectrum load sequence to reduce the total number of test setups); high load clipping (e.g., the reduction of the highest spectrum loads to a level such that the beneficial effects of compression yield are reduced or eliminated); and low load truncation (e.g., the removal of non-damaging load cycles to simplify the spectrum). Due to the modifications to the flightby-flight loading sequence caused by these changes, the applicant should propose either analytical or empirical approaches to quantify an adjustment to the number of test cycles which represents the difference between the test spectrum and assumed flight-by-flight spectrum. In addition, an adjustment to the number of test cycles may be justified by raising or lowering the test load levels, as long as appropriate data supports the applicant's position.
Other effects to be considered are different failure locations, different response to fretting 14 Par 7
conditions, temperature effects, etc. The analytical approach should use well established methods or be supported by test evidence.
c. Replacement times. Replacement times should be established for parts with established safe-lives and should, under § 25.571(a)(3), be included in the information prepared under § 25.1529. These replacement times can be extended if additional data indicates an extension is warranted. Important factors that should be considered for such extensions include, but are not limited to, the following:
() Comparison of original evaluation with service experience.
(2) Recorded load and stress data. Recorded load and stress data entails instrumenting airplanes in service to obtain a representative sampling of actual loads and stresses experienced. The data to be measured include airspeed, altitude, and load factor versus time data; or airspeed, altitude, and strain ranges versus time data; or similar data. The data, obtained by instrumenting airplanes in service, provide a basis for correlating the estimated loading spectrum with the actual service experience.
Additional analyses and tests. If test data and analyses based on repeated load tests of additional specimens are obtained, a re-evaluation of the established safe-life can be made.
Tests of parts removed from service. Repeated load tests of replaced parts can be utilized to reevaluate the established safe-life. The tests should closely simulate service loading conditions. Repeated load testing of parts removed from service is especially useful where recorded load data obtained in service are available, since the actual loading experienced by the part prior to replacement is known.
(5) Repair or rework of the structure. In some cases, repair or rework of the structure can gain further life.
d. Type design developments and changes. For design developments or design changes involving structural configurations similar to those of a design already shown to comply with the applicable provisions of § 25.571(c), it might be possible to evaluate the variations in critical portions of the structure on a comparative basis. Typical examples would be redesign of the wing structure for increased loads, and the introduction in pressurized cabins of cutouts having different locations or different shapes, or both. This evaluation should involve analysis of the predicted stresses of the redesigned primary structure and correlation of the analysis with the analytical and test results used in showing compliance of the original design with § 25.571(c).
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e. Environmental effects such as temperature and humidity should be considered in the damage tolerance and fatigue analysis and should be demonstrated through suitable testing.
8. DISCRETE SOURCE DAMAGE.
General. The purpose of this section is to establish FAA guidelines for consistent selection of load conditions for residual strength substantiation in showing compliance with § 25.571(e), Damage tolerance (discrete source) evaluation. The intent of these guidelines is to define load conditions that will not be exceeded with a satisfactory level of confidence on the flight during which the specified incident of § 25.571(e) occurs. In defining these load conditions, consideration has been given to the expected damage to the airplane, the anticipated response of the pilot at the time of the incident, and the actions of the pilot to avoid severe load environments for the remainder of the flight consistent with his knowledge that the airplane may be in a damaged state. With these considerations in mind, the following ultimate loading conditions should be used to establish residual strength of the damaged structure.
b. The maximum extent of immediately obvious damage from discrete sources (§ 25.571(e)) should be determined and the remaining structure shown, with an acceptable level of confidence, to have static strength for the maximum load (considered as ultimate load) expected during completion of the flight.
C. The ultimate loading conditions should not be less than those developed from the following conditions:
(1) At the time of the incident:
(i) The maximum normal operating differential pressure, multiplied by a 1.1 factor, plus the expected external aerodynamic pressures during 1g level flight, combined with 1g flight loads.
(11) The airplane, assumed to be in 1g level flight, should be shown to be able to survive any maneuver or any other flight path deviation caused by the specified incident of § 25.571(e), taking into account any likely damage to the flight controls and pilot normal corrective action.
(2) Following the incident:
(1) Seventy percent (70%) limit flight maneuver loads and, separately, 40 percent of the limit gust velocity (vertical and lateral) at the specified speeds, 16 Par 7
each combined with the maximum appropriate cabin differential pressure (including the expected external aerodynamic pressure).
(i1) The airplane must be shown by analysis to be free from flutter up to VD/Mp with any change in structural stiffness resulting from the incident.
a→ JAMES V. DEVANY Acting Manager, Transport Airplane Directorate Aircraft Certification Service, ANM-100 17 *U.S. Government Printing Office: 1997 - 514-000/50019