(vi) The maximum compressive load measured between the pelvis and the lumbar column of the (ATD) does not exceed 6.67 kN (1,500 lbs.).
(vii) Each upper torso restraint strap remains on the ATD shoulder during impact.
(viii) The pelvic restraint remains on the ATD pelvis during impact.
f. Procedures for Evaluating Impact Pulse Shapes.
(1) Acceptable Evaluation Method. Data for evaluating the impact pulse shape are obtained from an accelerometer which measures the acceleration on the test fixture or sled at the seat location or equivalent location in the direction parallel to the line of inertial response shown in Figure 29.562-1 of this guidance. The impact pulses intended for the tests discussed in this guidance have an isosceles triangle shape.
These ideal pulses are considered minimum test conditions. Since the actual acquired test pulses will normally differ from the ideal, it may be necessary to evaluate the acquired test pulses to insure the minimum requirements are satisfied.
(2) An acceptable method to evaluate the pulse shape should use the following steps:
(i) Extend the calibration baseline (zero G)
(ii) Locate the maximum peak deceleration (Gp) indicated on the plot.
(iii) Construct reference lines parallel to the baseline at levels of 0.1 Gp,
0.9 Gp, and 1.0 Gp.
(iv) Construct an onset line through the intersection points of the 0.1 Gp and 0.9 Gp reference lines with the increasing (onset) portion of the data plot. The data plot should not return to zero G between the two points selected.
(v) Locate the intersection points of the onset line with the baseline and with the 1.0 Gp reference line. The interval between these two points, measured along the time axis of the data plot, is considered the rise time (tr) of the test impact pulse.
(vi) The rise time of the test impact pulse should not exceed the value of (tr) given in Figure AC 29.562-1 for each test.
(vii) The area under the data plot curve within the rise time of the test impact pulse, Vra, should represent at least one half of the impact velocity given in Figure AC 29.562-1 for each test. If the value of peak acceleration measured in the test exceeds the level given in Figure AC 29.562-1 by no more than 10 percent, the pelvis to lumbar spinal column force and the upper torso restraint force measured in the test may Page C - 89
AC 29-2C, Chg 1 2/12/03 be adjusted by multiplying the measured values by the ratio of the peak acceleration given in Figure AC 29.562-1, divided by the measured peak acceleration, if necessary.
(viii) The magnitude of Gp should equal or exceed the minimum G given in Figure AC 29.562-1 for each test.
(ix) The area under the data plot curve from the intersection point of the onset line and the zero G baseline and a time not more than twice the appropriate rise time specified in Figure AC 29.562-1, plus 30 percent of the rise time later, should represent at least the impact during the test.
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Sled Impact Pulse
0.0 -1 Gp 9 Gp 50 - 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 29.562-7. Impulse Shape Sled Impact Pulse
0.0 Ideal Pulse Vra
- - - + Tred Gp 50 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 29.562-8. Change in Velocity (Vra) Page C - 91
AC 29-2C, Chg 4 5/1/2014 AC 29.563. § 29.563 (Amendment 29-12) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-12 included ce rtification requirements for ditching approvals. The rotorcraft must be able to sustain an emergency landing in water as prescribed by § 29.801(e).
b. Procedures. Refer to paragraph AC 29.801, § 29.801, for procedures.
AC 29.563A. § 29.563 (Amendment 29-30) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-30 added spec ific structural conditions to be considered to support the overall ditching requirements of § 29.801. These conditions are to be applied to rotorcraft for which over-water operations and associated ditching approvals are requested.
(1) The forward speed landing conditions are specified as:
(i) The rotorcraft should contact the most critical wave for reasonable, probable water conditions in the likely pitch, roll, and yaw attitudes.
(ii) The forward velocity relative to wave surface should be in a range of 0 to 30 knots with a vertical descent rate of not less than 5 FPS relative to the mean water surface.
NOTE: A forward velocity of less than 30 k nots may be used for multiengine rotorcraft if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown.
(iii) Rotor lift of not more than two-thirds of the design maximum weight may be used to act through the CG throughout the landing impact.
(2) For floats fixed or deployed before water contact, the auxiliary or emergency float conditions are specified in § 29.563(b)(1). Loads for a fully immersed float should be applied (unless it is shown that full immersion is unlikely). If full immersion is unlikely, loads resulting from restoring moments are specified for sidewind and unsymmetrical rotorcraft landing.
(3) Floats deployed after water contact are normally considered fully immersed during and after full inflation. An exception would be when the inflation interval is long enough that full immersion of the inflated floats does not occur; e.g., deceleration of the rotorcraft during water impact and natural buoyancy of the hull prevent full immersion loads on the fully inflated floats.
b. Procedures.
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(1) The rotorcraft support structure, structure-float attachments, and floats should be substantiated for rational limit and ultimate ditching loads.
(2) The most severe wave heights for which approval is desired are to be considered. A minimum of Sea State 4 condition wave heights should be considered (reference paragraph AC 29.801 (§ 29.801) for a description of Sea State 4 conditions).
(3) The landing structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions:
(i) Forward velocities of 0 to 30 knots (or a reduced maximum forward velocity if it can be demonstrated that a lower maximum velocity would not be exceeded in a normal one-engine-out landing).
(ii) The rotorcraft pitch attitude that would reasonably be expected to occur in service. Autorotation flight tests or one- engine-inoperative flight tests, as applicable, should be used to confirm the attitude selected. This information should be included in the Type Inspection Report.
(iii) Likely roll and yaw attitudes.
(iv) Vertical descent velocity of 5 FPS or greater.
(4) Landing load factors and water load distribution may be determined by water drop tests or analysis based on tests.
(5) Auxiliary or emergency float loads should be determined by full immersion or the use of restoring moments required to react upsetting moments caused by sidewind, asymmetrical rotorcraft landing, water wave action, rotorcraft inertia, and probable structure damage and punctures considered under § 29.801. Auxiliary or emergency float loads may be determined by tests or analysis based on tests.
(6) Floats deployed after initial water contact are required to be substantiated by tests or analysis for the specified immersion loads (same as for (5) above and for the specified combined vertical and drag loads).
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AC 29-2C, Chg 3 9/30/2008
SUBPART C - STRENGTH REQUIREMENTS
FATIGUE EVALUATION AC 29.571. § 29.571 FATIGUE EVALUATION OF FLIGHT STRUCTURE.
a. Explanation. An evaluation is required to assure structural reliability of the rotorcraft in flight. This evaluation may take the form of either tests or analysis. During the certification process, fatigue testing is more effective than analysis alone in identifying and preventing cracking that may occur during service. Analysis used for substantiation should be validated by tests. AC 27 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to the safe life methodology. A safe life may be assigned or the structure may be determined to be fail safe or a combination of these may be used. AC 29 MG 11 contains background information and acceptable means of compliance with the requirements pertaining to fatigue and flaw tolerance.
b. Procedures.
(1) The fatigue evaluation requires consideration of the following factors:
(i) Identification of the structure/components to be considered.
(ii) Th e stress during operating conditions.
(iii) The operating spectrum or frequency of occurrence.
(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual s trength of the cracked structure.
(2) Since the design limits, e.g., rotor RPM (maximum and minimum), airspeed, and blade angles (thrust, weight, etc.) affect the fatigue life of the rotor system, it is necessary that flight conditions be conducted at limits that are appropriate for the particular rotorcraft and at the correct combination of these limits. It will be the responsibility of flight test personnel to determine that the flight strain program includes conditions of flight at the various combinations of rotor RPM, airspeed, thrust, etc., that will be representative of the limits used in service. The flight test personnel should assure that the severity of the maneuvers to be investigated is such that actual service use will not be more severe. Flight test verification may be achieved through:
(i) Flying a representative set of maneuvers with the applicant’s pilot in the test aircraft at noncritical combinations of weight, CG, and speed. (An FAA/AUTHORITY letter for specific test authorization would ordinarily be required.)
(ii) Flying a representative set of maneuvers with the applicant’s pilot in a similar (certified) model to assess and agree upon the required maneuvers, control C – 94
deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.
(iii) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.
(iv) Observation of telemetered flight data to assure desired control deflections, rates, and aircraft attitudes.
(v) Some combinations of items b(2)(i) through b(2)(iv) above.
(3) Assessing the operation spectrum and the flight loads or strain measurement program will involve airframe, propulsion, and flight test personnel.
(4) Variation in the operating or loading spectrum among models, and variations in the spectrum for a particular model rotorcraft, should be evaluated.
Figure AC 27 MG 11-7 contains typical flight load measurement program conditions to be investigated. An example of a twin turbine spectrum is presented in Figure AC 27 MG 11-9. The tables should be used only as a guide and should be modified as necessary for each particular rotorcraft design.
(5) The difference in loading spectrum for different models that may be anticipated is illustrated by comparing the percentage of time assigned to level flight conditions, specifically 0.8 VH to 1.0 VH for three different rotorcraft designs where VH is the maximum airspeed at maximum continuous power in level flight. The first column applies to a single-piston-engine powered small rotorcraft used in utility operations. The second was obtained from data for a single-turbine-engine powered seven-place small business and utility rotorcraft. The third was obtained from data for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in the table below for the turbine utility business and turbine transport rotorcraft are examples of a particular design. The high percentage of time shown in this flight regime could be unconservative for some designs, especially if the stresses under these design conditions produce an infinite fatigue life for the particular component. The fatigue spectrum percentage of occurrences may be modified according to the intended operation usage of the rotorcraft. However, a conservative application should be considered. This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.
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AC 29-2C, Chg 1 2/12/2003 FIGURE AC 29.571-1 Comparison Percent of Time in Level Flight Turbine Piston Utility Twin Turbine Utility Business Transport
0.8 VNE 25% 0.8 VH 16% 0.8 VH 15%
1.0 VH 15% 0.9 VH 21% 0.9 VH 20%
1.0 VNE 3% 1.0 VH 24% 1.0 VH 38% Total 43% 61% 73% This variation illustrates the “tailoring” of the loading spectrum for the type of rotorcraft and the anticipated usage.
(6) External cargo operations are a unique and demanding operation. A “logging” operator may use 50 maximum power applications per flight hour to move logs from a cutting site to a hauling site. Power is used to accelerate, decelerate, or hover prior to load release. Lifting loads over an obstruction or natural barrier is another example of very frequent high power applications for takeoff and for hovering over the release area. Similar types of operations require flight loads data to assess the effects on fatigue critical components.
(7) Frequently the applicant may request approval of a gross weight for an external cargo configuration that exceeds the standard configuration gross weight. The external cargo VNE is typically significantly lower than the standard configuration VNE possibly due to adverse effects on flight loads at the increased weight.
(8) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives will be affected. The assessment may be done by calculating an “external cargo configuration” service life for each critical component. The lowest service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum is generally the approved service life. This procedure avoids prorating the operating time between the two types of operations. This procedure is necessary since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations.
(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if appropriate. This procedure will avoid delays associated with reinstallation and calibration of equipment.
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AC 29.571A. §29.571 (Amendment 29-28) FATIGUE EVALUATION OF STRUCTURE.
a. Explanation. Amendment 29- 28 adds a requirement to substantiate tolerance to flaws during the fatigue evaluation of structure. A flaw tolerant safe- life evaluation or a fail-safe (residual strength after flaw growth) evaluation is required by § 29.571(b) unless “the applicant establishes that these fatigue flaw tolerant methods for a particular structure cannot be achieved within the limitations of geometry, inspectability, and good design practices.”
b. Procedures.
(1) AC 29 MG 11 provides acceptable general procedures for complying with Amendment 29-28.
(2) Specific rotorcraft drive system gear fatigue evaluation procedures, which supplement AC 29 MG 11, follow:
(i) Fatigue test evidence is necessary for the fatigue evaluation of gears.
The test evidence should be provided by rotating tests of complete gearbox specimens operating under power. The tests provide the basis for analysis leading to the establishment of safe-life.
(ii) The tests are conducted specifically for the purpose of gear tooth evaluation, and components subjected to the tests do not have to be considered serviceable on completion of the test. Excessive wear on bearings and shafts and marking (including spalling) of bearings and gear teeth are acceptable provided no fatigue damage is evident on the gear teeth. However, fatigue damage other than tooth fatigue should be considered for test validity and the integrity of the affected part confirmed as necessary.
(iii) The test conditions (torque versus number of cycles) should permit the setting of mean strength curve(s) to be associated with each primary gear in the drive train. The minimum test condition should encompass those power levels for which repeated application in service is expected under normal conditions. The S -n curve(s), for the material and type of gear, should be reduced by a factor of safety to take into account material and manufacturing variability. The factored curve will then be used in conjunction with the flight power spectrum to determine a life (limited or unlimited) for the gears in the primary drive system.
(iv) Special procedures, which do not affect fatigue evaluation of the gear teeth, may be allowed to facilitate completion of the test provided they have been justified and they do not affect life determination. These include periodic interruption for inspection, replacement of non-critical parts and the use of special lubricants, special cooling systems, and methods to prevent unrepresentative deflections at the test torque levels.
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AC 29-2C, Chg 4 12/2/2011
(v) From evidence in relation to the strength of steel gears of conventional design, it is accepted that adequate fatigue strength can be demonstrated by the use of the above safety factor of 1.4 for a single test, 1.35 for two tests, 1.32 for three tests, and 1.3 for four or more tests. Where several tests are to be conducted, specimens should be selected from different manufacturing batches if practicable.
(vi) Demonstration of infinite life for gear teeth will normally require tests of a minimum of 107 cycles duration at factored power levels. Use of shorter duration tests should be justified.
AC 29.571B. § 29.571 (Amendment 29-55) FATIGUE TOLERANCE EVALUATION OF METALLIC STRUCTURE.
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.571 Amendment 29-55 of the FAA regulations dealing with the fatigue tolerance evaluation of transport category rotorcraft metallic structure. This guidance applies to conventional metallic materials. (Corresponding guidance for composite structure can be found in AC 29–2C MG 8, supplemented by AC 20-107B.
Note: once § 29.573 is effective, AC 29.573 will be the current guidance for composite structures.) The fatigue evaluation procedures outlined in this advisory material are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue tolerance evaluation is desirable, it is recognized that in such a complex area, new design features and methods of fabrication, new approaches to fatigue tolerance evaluation, and new configurations may require variations and deviations from the procedures described herein. It should be noted that § 29.571 requires that the methodology used by the applicant be approved by the FAA to assure compliance with the regulatory requirements.
b. Special Considerations. The unique performance capabilities of rotorcraft and their typical operational environment make fatigue tolerance evaluations both complex and critically important. Due to the many rotating elements inherent in their design, rotorcraft structures are potentially subject to damaging cyclic stresses in practically every regime of flight. The complexity of the fatigue loading is compounded by the fact that rotorcraft are highly maneuverable and are utilized for many widely varying roles.
Corrosion and other environmental damages are not uncommon in rotorcraft operations; neither are inadvertent damages from maintenance that is typically frequent and intensive. For these reasons, special attention should be focused on the fatigue tolerance evaluation of rotorcraft structure.
c. Background.
(1) Fatigue of rotorcraft dynamic components was first addressed in the 1950’s by means of a Safe-Life methodology. The application of this methodology, as Page C - 98
described in AC 27-1B MG 11, has proven to be successful in providing an adequate level of reliability for transport category rotorcraft. However, it was recognized in the 1980’s that higher levels of reliability might be realized by taking into account the fatigue strength-reducing effects of damage that experience has shown can occur in manufacture or in operational service. The introduction of composites led the manufacturers and regulatory authorities to develop a robust Safe-Life methodology by taking into account the specific static and fatigue strength-reducing effects of aging, temperature, moisture absorption, impact damage, and recognition of an accepted industry standard. Furthermore, where clearly visible damages resulted from impact or other sources, inspection programs were developed to maintain safety. In parallel, crack growth methodology has been successfully used for solving short-term airworthiness problems in metallic structures of rotorcraft, and as the certification basis for civil and military transport aircraft applications. These advances in design, analytical methods, and industry practices made it feasible to address certain types of damage, which could result in fatigue failure. Consistent with this, the regulatory requirements of § 29.571 were substantially revised by Amendment 28. While many years have passed since its introduction, Amendment 28 has had little exposure to use for certification of completely new rotorcraft designs. However, the general understanding of rotorcraft fatigue tolerance evaluation has developed considerably in the interim and an additional amendment was determined to be appropriate. The latest Amendment 29- 55 of part 29 and the associated revisions to advisory material were introduced to improve the currency and understanding of the rule and clarify the differing approaches and methods available for accomplishing fatigue tolerance evaluation of rotorcraft metallic structure.
(2) This guidance provides material with respect to the fatigue tolerance requirements for metallic structure and is supplemented by AC 27-1B MG 11 for evaluations using the Safe-Life methodology and other general fatigue considerations.
d. Introduction.
(1) Definitions. The following definitions are applicable when used within the context of this guidance material.
(i) As-manufactured structure is a structure that passes the applicable quality control process and has been found to conform to an approved design within the allowable tolerances.
(ii) Barely Detectable Flaw (BDF) is the worst-case flaw that is expected to remain on the structure for its operational life.
(iii) Catastrophic failure is an event that could prevent continued safe flight and landing.
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AC 29-2C, Chg 4 12/2/2011
(iv) Clearly Detectable Flaw (CDF) is the worst-case detectable flaw that would not be expected to remain in place for a significant period of time without corrective action.
(v) Damage is a detrimental change to the condition of the structure or assembly. In the context of this guidance material it is used as a generic term to describe all types of flaws including those caused by environmental effects and accidental damage arising in manufacture, maintenance or operation.
(vi) Damage Tolerance is the attribute of the structure that permits it to retain its required residual strength without detrimental structural deformation for a period of un-repaired use after the structure has sustained a given level of fatigue, corrosion, accidental, or discrete source damage.
(vii) Discrete flaw is a flaw that is not inherent in the design and is caused by an external action, such as corrosion, scratches, gouges, nicks, fretting, wear, impact, and potentially cracks initiated by fatigue.
(viii) Fatigue is a degradation process of a structure subject to repeated loads that may involve four phases (e.g., nucleation of many micro-cracks, coalescence of some micro-cracks to one major macro-crack, stable crack growth, unstable crack growth) and immediate failure. The boundaries between these phases are, in practice, not always easily defined. Crack initiation methods (e.g., using the S -N curve and the Miner’s Rule) are generally used to address the first two phases. Linear Fracture Mechanics methods (e.g., using da/dn - ΔK and fracture toughness data) are generally used for the latter two phases.
(ix) Fatigue Loads are repeated loads, which induce a repeated variation of stress versus time in a structure.
(x) Fatigue Tolerance is the ability of a structure, either in an as-manufactured or damaged condition, to tolerate specified operational loading for a given period of use without initiating cracks, and assuming they initiate, tolerate their growth, without failure, under specified residual strength loads.
(xi) Flaw is an imperfection, defect, or blemish and may be either discrete or intrinsic.
(xii) Inspection interval is the maximum period of usage allowed for a structure between inspections. At the end of this period, the structure is inspected and if there is no damage detected, the structure may be returned to service for another inspection interval.
(xiii) Intrinsic flaw is a flaw that is inherent in the design and manufacture of the part, situated within it or peculiar to it, such as inclusions, cracks, forging laps, or porosity.
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