SECTION 13. EMERGENCY LANDING CONDITIONS
218. § 27.561 GENERAL.
a. Explanation.
(1) The occupants should be protected as prescribed from serious injury during an emergency/minor crash landing on water or land for the conditions prescribed in the standard. The standard states that each occupant should be given every reasonable chance of escaping serious injury in a minor crash landing.
(2) The standard in § 27.561(b)(3) specifies certain ultimate inertia load factors but allows a lesser downward vertical load factor by virtue of a 5 FPS ultimate rate of descent.
(3) In addition, the occupants shall be protected from items of mass inside the cabin as well as outside the cabin. For example, a cabin fire extinguisher must be restrained for the load factors prescribed in this section (reference § 27.1411(b)(2).) A transmission or engine must be restrained to the load factors in § 27.561(b)(3) if located above or behind the occupants.
(4) For aircraft equipped with retractable landing gear, the landing gear is retracted for compliance with the standards.
b. Procedures.
(1) The design criteria report or another similar report of the rotorcraft structural limits should contain the (ultimate) minor crash condition load factors.
(2) Section 27.785 (Paragraph 336 of this AC) concerns application of this design standard to seats, berths, belts, and harnesses.
(3) The ultimate design landing and maneuvering load factors may exceed the minor crash condition load factors. The highest load factor derived shall be used.
For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 27.337.
(il) The ultimate vertical landing load tactors derived trom 99 27.471 through 27.521, whichever is appropriate for the design, may exceed the 4.0g down load factor in this section. The rotorcraft landing case design limit contact velocity shall range from 6.5 to 8.3 FPS (see §§ 27.473 and 27.725).
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(4) As specified in (b)(3) (iv) of the standard, the downward load factor is 4.0 or a lower design load factor may be used.
The lower load factor relates to a rotorcraft impacting a flat, hard landing surface at 5 FPS (ultimate) vertical rate of descent. The load factor derived for each unique design is a tunction of the aircraft impact/crushing characteristics.
(i) The 4.0g down load factor case is related to either a fixed or retractable gear rotorcraft. This condition is not dependent on impact characteristics of the rotorcraft.
(ill) As noted in Paragraph (3) above, the design landing load factors may exceed each of the two previous cases and would then become the prominent design (vertical load) parameter for seats, transmissions, fire extinguishers, etc.
(5) Items of mass such as fire extinguishers, nav-com equipment, liferafts, engines, and/or transmissions shall be restrained for the appropriate load factors.
(6) Cargo/baggage compartments separated from the passenger compartment shall be designed for load factors specified in § 27.787. The conditions in § 27.561 are excepted from that standard.
218A. § 27.561 (Amendment 27-25) EMERGENCY LANDING CONDITIONS - GENERAL.
a. Explanation. Amendment 27-25 adds or increases the design static load factor of § 27.561 in two areas:
(1) The design static load factors for the cabin in § 27.561(b)(3) are increased in concert with the dynamic test requirements of new § 27.562.
(2) Design static load factors are added in § 27.561(c) for external items of mass located above and/or behind the crew and passenger compartment.
b. Procedures. The procedures of Paragraph 218, § 27.561, continue to apply except the new load factors of § 27.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented.
218B. § 27.561 (Amendment 27-30) EMERGENCY LANDING
CONDITIONS - GENERAL
a. Explanation.
(1) Amendment 27-30 adds Paragraph (d) which lists specific load factors for the fuselage structure in the area of internal fuel tanks located below the passenger 272 Par 218
floor level. For other locations, the fuselage structure is to be designed to resist crash impact loads prescribed in § 27.561(b)(3) for fuel tanks located within the cabin area; or § 27.561(c) for fuel tanks located behind or above the occupant area. These load factors are provided to prevent crash induced fuel tank ballistic hazards to occupants and to also protect the fuel tank from rupture as prescribed. The landing gear must be retracted if the rotorcraft is equipped with retractable gear.
(2) Section 27.952(b) provides specific load factors for the fuel tanks which are identical to the load factors stated in § 27.561. Paragraph 447 of this document provides information and guidance for § 27.952 and may be used in conjunction with this paragraph.
b. Procedures. The procedures of Paragraphs 218 and 218A continue to apply except new load factors are established for fuel tanks located below the passenger floor level. Each fuel tank and its installation are subject to the loads stated in the standard.
The load factors are determined by the fuel tank location.
218C. § 27,561 (Amendment 27-32) EMERGENCY LANDING CONDITIONS - GENERAL.
a. Explanation. Amendment 27-32 adds a new rearward emergency load factor of 1.5g to both §§ 27.561(b)(3)(v) and 27.561(c)(5). The addition of the 1.5g rearward load factor in § 27.561(b)(3)(v) is to provide an aft ultimate load condition for substantiation of the restraints required for retention of both occupants and significant items of mass inside the cabin that could otherwise come loose and cause injuries in an emergency landing. The addition of the 1.5g rearward load factor to § 27.561(c)(5) is to provide an aft ultimate load condition for substantiation of the support structure for retention of significant items of mass above and forward of the occupied volume(s) of the rotorcraft that could otherwise come loose and injure an occupant in an emergency landing. Amendment 27-32 also increases the forward, sideward, and downward emergency load factors of § 27.561(c)(2), (c)(3), and (c)(4), respectively; for retention of items of mass above and behind the occupied volumes) that could otherwise come loose and injure an occupant in an emergency landing.
b. Procedures. The procedures of Paragraphs 218, 218A, and 218B continue to apply except the newly specified load factors must be used. A list of the significant items of mass to be considered should be compiled by the applicant and approved by the certifying authority.
219. § 27.562 (Amendment 27-25) EMERGENCY LANDING DYNAMIC CONDITIONS.
a. Explanation. Amendment 27-25 adds new requirements for the dynamic testing of all seats in rotorcraft.
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b. Procedures. AC 20-137, "Dynamic Evaluation of Seat Restraint Systems and Occupant Restraint for Rotorcraft (Normal and Transport)i provides procedures for complying with § 27. 562 using the 170-pound anthropomorphic test dummy specified in § 27.562(b). Those seats not occupied for takeoff and landing, and so placarded and identified in the rotorcraft flight manual (RFM), may be excluded from compliance.
220. § 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 338 of this AC for procedures.
220A. § 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.
(її) 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 274 Par 219
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.
(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 338 (§ 27.801) of this AC 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).
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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221.-229. RESERVED.
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SECTION 14. FATIGUE EVALUATION
230. § 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.
(1) Advisory Circular 20-95 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 729 of this document 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:
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.
(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 Par 230 285
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.
AC 20-95, Paragraph 7, entitled "Loading Spectrum," contains the statement that Table 1 (of the circular) contains typical percent of occurrences for various flight conditions for a single-piston-engine powered small rotorcraft used in utility operations.
In addition, the table 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 Vu to 1.0 Vu for three different rotorcraft designs as shown in Table 230-1. (V. is the maximum airspeed at maximum continuous power in level flight.) The first column was obtained from Table 1, AC 20-95 which applies to a 286 Par 230
single-piston-engine powered small rotorcraft used in utility operations. The second column is appropriate tor 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 Table 230-1 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 in AC 20-95 shall be modified according to the intended operational 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.
Table 230-1 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 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).
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АС 27-1A 7/30/97
(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.
230A. $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 AC 29-2A, Section 230A are required for fatigue tolerance evaluation. AC 29-2A, Section 230A 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 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.
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(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.
231.-240. RESERVED.
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SECTION 15, DESIGN AND CONSTRUCTION - GENERAL
241. $ 27.601 DESIGN.
a. Explanation.
(1) This rule requires that no design features or details be used that experience has shown to be hazardous or unreliable.
(2) Further, the rule requires that the suitability of each questionable design detail and part must be established by tests.
b. Procedures.
(1) This rule is met partially by a review of service history of earlier model rotorcraft, or for a new model, review of service experience of models with similar design features. Specifically, this rule covers "features or details" such as the following:
Seat track-to-seat interface fittings. These fittings should have adequate locking devices to prevent both premature structural failure and premature unlatching.
(¡¡) Seat belt and harness should be of a type and construction that service experience has shown to be easy to don and unlatch and remove. They should also be of a type that is reliable, does not interfere with egress, and does not sustain unnecessary wear and tear under normal operations.
(iii) Metallic parts less than a certain thickness gauge and composite materials less than a certain number of plies should not be used. The minimum thickness and number of plies should be based to a large degree on service (normal wear and tear) experience with similar designs.
(2) The effects of service wear on the loading of critical components should be considered. Flight testing, ground testing, and analyses may be used in these considerations.
(3) Tests are required for details and parts which the applicant chooses to use after questions have arisen concerning their suitability.
242. § 27.603 (Amendment 27-16) MATERIALS.
a. Explanation. The rule requires that the suitability and durability of materials, the failure of which could adversely affect safety, must be determined by three-fold considerations:
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