(i) Impact pulse shape (Paragraph d.(2)).
(ii) HIC results for all ATD exposed to secondary head impact with interior components of the rotorcraft (Paragraph d.(3)), or head strike paths and velocities if secondary head impact is likely for future use in unique interiors (Paragraph c.(3)(iii)).
(iii) Impact velocity (Paragraph d.(4)).
(iv) Upper torso restraint system load if applicable (Paragraph d.(5)).
(v) Compressive load between the pelvis and the lumbar column (Paragraph d.(6)).
(vi) Retention of upper torso restraint straps if applicable (Paragraph d.(7)).
(vii) Retention of lap safety belt (Paragraph d.(8)).
(viii) Femur thigh loads, optional measurement.
(ix) Seat attachment (Paragraph d.(10)).
(x) Seat deformation (Paragraph d.(11)).
(xi) Seat attachment reaction time histories (Paragraph f.).
(4) Dynamic Impact Test – Pass/Fail Criteria: The dynamic impact tests should demonstrate that:
(i) The seat structure remains intact that is attached to the tracks or fittings, etc.
(ii) The occupant retention system is capable of carrying the dynamic loads.
(iii) The seat permanent deformations are within defined limits and will not significantly impede an occupant from releasing the torso restraints, standing and exiting the seat.
(iv) If the ATD’s head is exposed to impact during the test, a HIC of 1,000 is not exceeded. Data may be obtained for use with other unique installations.
(v) Where upper torso restraint straps are used, tension loads in individual straps do not exceed 7.78 kN (1,750 lbs.). If dual straps are used for Page C - 77
AC 27-1B, Chg 1 2/12/03 restraining the upper torso, the total strap tension load does not exceed 8.90kN (2,000 lbs).
(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 27.562-1 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 27.562-1 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 27.562-1 for each test. If the value of peak acceleration measured in the test Page C - 78
exceeds the level given in Figure AC 27.562-1 by no more than 10 percent, the pelvis to lumbar spinal column force and the upper torso restraint force measured in the test may be adjusted by multiplying the measured values by the ratio of the peak acceleration given in Figure AC 27.562-1, divided by the measured peak acceleration, if necessary.
(viii) The magnitude of Gp should equal or exceed the minimum G given in Figure AC 27.562-1 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 27.562-1, plus 30 percent of the rise time later, should represent at least the impact during the test.
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AC 27-1B, Chg 1 2/12/03 Sled Impact Pulse
0.0 T1 GP, .9 Gp Gp 50 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 27.562-7. Impulse Shape Sled Impact Pulse
0.0 Ideal Pulse Vra
- - - + Tred Gp 100 150 200 250 300 350 Time (Milliseconds) FIGURE AC 27.562-8. Change in Velocity (Vra) Page C - 80
AC 27.563 § 27.563 (Amendment 27-11) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 27-11 included certification requirements for ditching approvals. The rotorcraft must be able to sustain an emergency landing in water as prescribed by § 27.801(e).
b. Procedures. Refer to paragraph AC 27.801 for procedures.
AC 27.563A § 27.563(Amendment 27-26) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 27-26 added specific structural conditions to be considered to support the overall ditching requirements of § 27.801. 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 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 knots 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 § 27.563(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 27.801 (§ 27.801) 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 § 27.801. 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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SUBPART C - STRENGTH REQUIREMENTS
FATIGUE EVALUATION AC 27.571. § 27.571 (Amendment 27-26) 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.
(1) Chapter 3 AC 27 MG 11 contains background information and acceptable means of compliance with the requirements. A safe life may be assigned or the structure may be fail safe as prescribed or a combination of these may be used.
(2) Mandatory inspections, service life (replacement times) etc., determined in complying with the standard shall be placed in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness (also called Maintenance Manual). See Appendix A of FAR Part 27, paragraphs A27.4 and paragraph AC 27.1529 for information.
(3) Amendment 27-26 amended the standard to require evaluation of the landing gear and their related primary attachments.
(4) Amendment 27-26 also amended the standard to require evaluation of ground-air-ground cycles on the rotorcraft, and if applicable, of external cargo operations. Previously external cargo operations were evaluated whenever the rotorcraft cargo combination exceeded the “standard” maximum certificated gross weight, and the CG range specified in § 27.25(c). If these limits were not exceeded, an evaluation was not required by the standard prior to Amendment 27-26.
b. Procedures.
(1) The fatigue evaluation requires consideration of the following factors:
(i) Identification of the structure/components to be considered.
(ii) The stress during operating conditions.
(iii) The operating spectrum or frequency of occurrence including frequency of ground-air-ground cycles, as well as external cargo operations.
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AC 27-1B, Chg 3 9/30/2008
(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual strength 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 engineering and flight test personnel to determine that the flight strain program proposal 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. Verification that proposed maneuvers are suitable may be achieved by:
(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.) If the procedure is used, the applicant should provide adequate preliminary flight strain data from development or other tests to confirm a cleared (non-critical) flight envelope for conduct of these representative maneuvers.
(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 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.
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(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 as shown in figure AC 27.571-1. (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 column is appropriate for a single- turbine-engine powered seven-place small business and utility rotorcraft. The third column is appropriate for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in figure AC 27.571-1 for the turbine utility business and turbine transport rotorcraft are examples of particular designs. The high percentage of time shown in this level 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 should 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.
FIGURE AC 27.571-1 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%
(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) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives, inspections, etc., will be affected. The assessment may be done by calculating an “external cargo configuration” service life for each critical component. The lowest Page C - 85
AC 27-1B, Chg 3 9/30/2008 service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum or from frequent ground-air-ground cycles is generally the approved service life or replacement time. Since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations, this procedure could be used if an “operational cycles” equation for equivalent flight hours is not approved (see (8) below).
(8) The Airworthiness Limitations Section of the maintenance manual shall contain the required information derived from complying with the standard. If an “operational cycles” equation for “equivalent flight hours” is approved under the standard, the equation is included in this approved section of the manual.
(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if applicable. The ground-air-ground cycle is inherent in these conditions. This procedure will avoid delays associated with reinstallation and calibration of equipment.
AC 27.571A. §27.571 (Amendment 27-33) FATIGUE EVALUATION OF FLIGHT STRUCTURE FOR CATEGORY A CERTIFICATION.
a. Explanation. Amendment 27-33 added Appendix C to specify the requirements for Category A certification of normal category rotorcraft. The requirement for fatigue tolerance evaluation will require test evidence to support the analysis.
b. Procedures. For Category A certification, the tests specified in paragraph AC 29.571A are required for fatigue tolerance evaluation.
Paragraph AC 29.571A is repeated in this section.
(1) 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.
(2) 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.
(3) 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 test conditions, should at a minimum, encompass those power levels for which repeated application inservice is expected under normal circumstances. The S -N curve(s), for the material and type of gear, should be reduced by a factor of safety to Page C - 86
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.
(4) 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, etc., 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.
(5) 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.
(6) The 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.
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AC 27-1B, Chg 4 12/1/2011 AC 27.573. § 27.573 (Amendment 27-47) DAMAGE TOLERANCE AND
FATIGUE EVALUATION OF COMPOSITE ROTORCRAFT
STRUCTURES
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 27.573, Amendment 27-47, Title 14 of the Code of Federal Regulations (CFR) dealing with the damage tolerance and fatigue evaluation of normal category composite rotorcraft structures. Paragraph f.(6) specifically addresses the advisory guidance applying to damage tolerance and fatigue evaluation as required by § 27.573, Amendment 27-47. Some information contained in AC 27-1B, MG 8 (Amendment 27-30) is repeated and updated, as appropriate, to preserve the “building block” approach for analyses of composite rotorcraft structure for compliance to § 27.573, Amendment 27-47 . (Supplemental guidance can be found in AC 20-107, “Composite Aircraft Structure.”) These procedures address the substantiation requirements for composite material system constituents, composite material systems, and composite structures common to rotorcraft. A uniform approach to composite structural substantiation is desirable, but it is recognized that in a continually developing technical area, which has diverse industrial roots both in aerospace and in other industries, variations and deviations from the procedures described here may be necessary. Deviations from this advisory material should be coordinated in advance with the Rotorcraft Directorate.
b. Special Considerations. Since rotorcraft structure is configured uniquely and is inherently subjected to severe cyclic stresses, special consideration is required for the substantiation of all rotorcraft structure, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations are attained during the type certification process for all structure with special emphasis on composite structure because of its unique structural characteristics, manufacturing quality and operational considerations, and failure mechanisms.
c. Background.
(1) Historically, rotorcraft have required unique, conservative structural substantiation because of unique configuration effects, unique loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testing required by these effects. Rotorcraft structural static strength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft structural fatigue substantiation is significantly different from fixed wing fatigue substantiation. Since AC 20-107, as developed, applies to both fixed wing aircraft and rotorcraft, it, of necessity, was finalized in a broad generic form. Accordingly, a need to supplement AC 20-107 for rotorcraft was recognized during type certification programs. One significant difference in traditional rotorcraft fatigue substantiation programs and fixed wing fatigue programs is the use of multiple component fatigue tests for rotorcraft programs rather than just one full-scale test. Also, constant amplitude, accelerated load tests are typically used rather than spectrum tests because of the high frequency loads Page C - 88