c. Procedures.
(1) Failure Analysis. The first stage of the design assessment should be the failure analysis, by which all the hazardous and catastrophic failure modes are identified. The failure analysis may consist of a structured, inductive bottom -up analysis, which is used to evaluate the effects of failures on the system and on the aircraft for each possible item or component failure. When properly f ormatted, it will aid in identifying latent failures and the possible causes of each failure mode. The failure analysis should take into consideration all reasonably conceivable failure modes in accordance with the following:
(i) Each item/component function(s).
(ii) Item/component failure modes and their causes.
(iii) The most critical operational phase/mode associated with the failure mode.
(iv) The effects of the failure mode on the item/component under analysis, the secondary effects on the rotors and on the rotor drive system, on other systems, and on the rotorcraft. Combined effects of failures should be analyzed where a primary failure is likely to result in a secondary failure.
(v) The safety device or health monitoring means by which occurring or incipient failure modes are detected, or their effects mitigated. The analysis should consider the safety system failure.
(vi) The compensating provision(s) made available to circumvent or mitigate the effects of the failure mode (see also paragraph c(2) below)
(vii) The failure condition severity classification according to the definitions given in paragraph b above.
When deemed necessary for particular system failures of interest, the above analysis may be supplemented by a structured, deductive top-down analysis, which is used to determine which failure modes contribute to the system failure of interest.
Dormant failure modes should be analyzed in conjunction with at least one other failure mode for the specific component or an interfacing component. This latter failure mode should be selected to represent a failure combination with potential worst case consequences.
When significant doubt exists as to the effects of a failure, these effects may be required to be verified by tests.
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(2) Evaluation of Hazardous and Catastrophic Failures: The second stage of the design assessment is to summarize the hazardous and catastrophic failures and appropriately substantiate the compensating provisions which are made available to minimize the likelihood of their occurrence. Those failure conditions that are more severe should have a lower likelihood of occurrence associated with them than those that are less severe. The applicant should obtain early concurrence of the cognizant certificating authority with the compensating provisions for each hazardous or catastrophic failure.
Compensating provisions may be selected from one or more of those listed below, but not necessarily limited to this list.
(i) Design features; i.e., safety factors, part derating criteria, redundancies, etc.
(ii) A high level of integrity: All parts with catastrophic failure modes and critical characteristics are to be identified as Critical Parts and be subject to a Critical Parts Plan (see AC 29.602). Where a high level of integrity is used as a compensating provision, parts with a hazardous failure mode which would prevent continued safe flight may be included in a Critical Parts Plan or subjected to other enhancements to the normal control procedures for parts.
(iii) Fatigue tolerance evaluation.
(iv) Flight limitations.
(v) Emergency procedures.
(vi) An inspection or check that would detect the failure mode or evidence of conditions that could cause the failure mode.
(vii) A preventive maintenance action to minimize the likelihood of occurrence of the failure mode including replacement actions and verification of serviceability of items which may be subject to a dormant failure mode.
(viii) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.
(ix) Safety devices or health monitoring means beyond those identified in
(vi) and (vii) above.
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AC 29.549. § 29.549 (Amendment 29-26) FUSELAGE AND ROTOR PYLON.
a. Explanation. This regulation requires that the fuselage and rotor pylon (including the tail fin, if any) be designed to withstand the flight loads of §§ 29.337 through 29.351, the ground loads of §§ 29.235, 29.471 through 29.497, skid loads of § 29.501, ski loads of § 29.505, water loads of § 29.521, and rotor loads of § 29.547(d) and (e). The ski and water loads pertain to optional features.
(1) Consideration is also required of --
(i) Aux iliary rotor thrust;
(ii) The torque reaction of each rotor drive system; and
(iii) Balancing air and inertia loads.
(2) Each eng ine mount and adjacent fuselage must be substantiated as prescribed. In addition, if 2 ½-minute power is used, “each engine mount and adjacent structure must be designed to withstand the loads resulting from a limit torque equal to
1.25 times the mean torque for 2 ½-minute power combined with 1g flight loads.”
Amendment 29-26 extended paragraph (e) of the standard to 2 ½-minute “OEI power.”
b. Procedures. Compliance with this standard is accomplished by application of the specified aircraft loads including engine torque to the fuselage and rotor pylon structure by stress analyses and/or static tests. Drive system torque factors to be used are noted in § 29.547 for the main rotor structure as well as in § 29.549(e).
AC 29.551. § 29.551 AUXILIARY LIFTING SURFACES.
a. Explanation. This regulation specifies that auxiliary lifting surfaces be designed to withstand critical flight and ground loads derived for conditions specified and any “other critical condition expected in normal operation.” Stub wings would comply with this standard.
b. Procedures. The surface design loads are derived from the conditions specified. Conservative aerodynamic data, including load distributions, may be used in place of data derived from wind tunnel or instrumented flight testing of the exact aerodynamic shapes involved. Special attention should be placed on concentrated load effects from fuel tanks or other large mass items that may be located in lifting surfaces.
These types of load concentrations are to be considered in conjunction with inertia and aerodynamic loads.
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SUBPART C – STRENGTH REQUIREMENTS
EMERGENCY LANDING CONDITIONS AC 29.561. § 29.561 EMERGENCY LANDI NG CONDITIONS - GENERAL.
a. Explanation.
(1) Occupant protection. 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.
In addition, the occupants must 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. A transmission or engine mu st be restrained to the load factors in § 29.561(b)(3) if located adjacent to, above, or behind the occupants.
(2) Load factor determination. Secti on 29.561(b)(3) specifies certain ultimate inertial load factors but allows a lesser downward vertical load factor by virtue of a 5 FPS ultimate rate of descent at maximum design weight.
(3) Retractable landing gear. For roto rcraft equipped with retractable landing gear only the retracted configuration must be considered.
(4) Fuel tank protection.
(i) Underfloor fuel tanks are specifically addressed in § 29.561(d). The fuselage structure must be designed to resist crash impact loads prescribed in § 29.561(b)(3) and to also protect the fuel tank from rupture as prescribed. The landing gear must be retracted if the rotorcraft is equipped with retractable gears.
(ii) Section 29.963(b), a general rule tank design standard, also refers to § 29.561. This standard specif ies that each tank and its installation must be designed or protected to retain fuel without leakage under the emergency landing conditions in § 29.561. Section 29.963 of this AC relates to this standard.
(5) External load considerations. The load factors of § 29.561 and the criteria of § 29.562 are not directly applicable to external load systems. This is because in emergency crash scenarios that involve external loads, the external load is neither typically subjected to the same minor crash loads (§ 29.561) as is the rotorcraft hull and its internal occupants nor are all of the occupant protection criteria (§ 29.562) needed or practicable to apply. Appropriate safety for external load carriage systems is provided by the criteria of § 29.865. Safety standar ds for external load attaching means are provided in § 29.865.
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AC 29-2C, Chg 4 5/1/2014
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 29.785 (section 29.875 of this AC) concerns application of this design standard to seats (berths, litters), belts, and harnesses.
(3) The ultimate design landing and maneuvering load factors may exceed the minor crash condition load factors. The hi ghest load factor derived must be used.
(i) For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 29.337.
(ii) The ultimate vertical landing load factors derived from §§ 29.471 through
29.521, whichever are appropriate for the design, may exceed the 4.0g down load factor in this section. The rotorcraft landing case design limit contact velocity must be at least
6.5 FPS (see §§ 29.473 and 29.725).
(4) As specified in § 29.561(b)(3)(iv), the downward load factor is 4.0, or a lower design load factor may be used at maximum design weight.
(i) 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 function of the rotorcraft impact and crushing characteristics.
(ii) The 4.0g down load factor case is related to either a fixed or retractable gear rotorcraft. This condi tion is not dependent on impact characteristics of the rotorcraft.
(iii) As noted in paragraph b.(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, and so forth.
(5) Items of mass such as fire extinguishers, radio equipment, life rafts, engines, and transmissions must be restrained for the appropriate load factors.
(6) Cargo or baggage compartments separated from the passenger compartment must be designed for load factors specified in § 29.787. The conditions in § 29.561 are excepted from that standard.
(7) Each fuel tank and its installation are subject to the loads stated in this standard whether “under floor” or located elsewhere. (See § 29.963(b) also.) Underfloor fuel tanks are specifically addressed in § 29.561(d); however, an acceptable means of compliance with CAR 7.261 which is identical to and preceded § 29.561(d) is quoted here for information.
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Notes: Fuselage keels whose design and structural strength are such as to resist crash impacts associated with the emergency landing conditions of § 7.260 (§ 29.561) without extreme distortion which might tend to rupture the fuel tank may be considered to comply with the requirements of this section (7.261).
Puncture resistant “bladder” fuel cells that are adequately designed and also protected from the stated impact loads imposed on the fuselage may also satisfy the standards.
(8) For rotorcraft with retractable landing gear, alternative landing gear positions and the resulting effects on potential fuel release should be evaluated.
AC 29.561A. § 29.561 (Amendment 29-29) EMERGENCY LANDING CONDITIONS - GENERAL.
a. Explanation. Amendment 29-29 adds or increases the design static load factors of § 29.561 in three different areas. The addition of these load factors eliminates the 5 FPS descent velocity criteria of unamended § 29.561(b)(3).
(1) The design static load factors for the cabin in § 29.561(b)(3) are increased in concert with the dynamic test requirements of new § 29.562.
(2) Design static load factors are added in § 29.561(c) for external items of mass located above or behind the crew and passenger compartment.
(3) The static load factors, which were formerly only referenced in § 29.561(d), are now included explicitly in § 29.561(d) for substantiation of internal fuel tanks which are below the passenger floor.
b. Procedures. The procedures in se ction 29.561 of this AC continue to apply except the new load factors of § 29.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented. In addition, each fuel tank and its installation are subject to specific load factors that are based on the fuel tank location.
(1) The crash impact load factors for the airframe structure surrounding the underfloor fuel tanks are specified in § 29.561(d). The fuselage structure must be designed to resist the specified crash impact loads and to help protect the fuel tank from rupture. If equipped with retr actable landing gear, the effects of the landing gear on fuel system rupture should be considered in both the retracted and unretracted configurations.
(2) Section 29.952(b) (see section 29.952 of this AC) specifies the design load factors for crash resistant fuel systems in an otherwise survivable impact. This section relates to § 29.561(d) as follows. The § 29.952 load factors are for the fuel tanks, other Page C - 53
AC 29-2C, Chg 4 5/1/2014 significant mass items in the fuel system, and their attachment to the rotorcraft airframe for both occupant survivability and retention of fuel in a survivable impact; whereas, the § 29.561(d) load factors only apply to the rotorcraft airframe surrounding the underfloor fuel tanks and their installation for the same reasons. These two sets of load factors are not additive. They are applied separately (as design ultimate load factors) to the portions of the rotorcraft to which they are specified to apply. The application of the § 29.561(d) load factors is described as follows. The loads generated by § 29.561(d) are intended to be applied to the airframe structure surrounding the fuel cell to ensure that the entire airframe structure provides the appropriate level of crash resistance (i.e., stiffness, crushability, crushing rate, energy absorption capability, etc.) and to ensure that the airframe structure’s failure modes (e.g., buckling, creation of sharp edges, structural spears, etc.) are such that fuel cell rupture (and the resultant post crash fire potential) is mitigated to the maximum practicable extent in a otherwise survivable emergency landing. Each fuel cell (and major fuel cell component) creates an applied load on the airframe in an emergency landing condition. These loads are determined by multiplying the worst case mass of the fuel cell (i.e., a full fuel cell) by the load factors of § 29.561(d). These loads are then applied (ut ilizing the appropriate design load paths) to the airframe structure surrounding the fuel cell to help design the structure for optimal crash resistance. Added stiffness effects for both a full and less than full fuel cell should be considered in the design process. A significantly less than full fuel cell will typically not have any significant stiffness effects, since in a less than full condition the fuel cell cannot typically transfer load hydraulically.
(3) The minor crash ultimate load factors for doors and others emergency exits are specified in §§ 29.783(d) and 29.809(e). The related inertial forces are not applicable to cargo or to service doors (not suitable for use as an exit in an emergency).
If any item of mass installed in the cabin can possibly interact with the fuselage and cause higher deformation, then § 29.561 (b)(3) loads factors should be applied to the design of doors and emergency exits.
AC 29.561B. § 29.561 (Amendment 29-38) EMERGENCY LANDING CONDITIONS - GENERAL.
a. Explanation. Amendment 29-38 adds a new rearward emergency load factor of
1.5g to both §§ 29.561(b)(3)(v) and 29.561(c)(5). The addition of the 1.5g rearward load factor in § 29.561(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 rear ward load factor to § 29.561(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 29-38 also increases the forward, sideward, and downward emergency load factors of § 29.561(c)(2), (c)(3), and (c)(4), respectively, for retention of items of mass above and behind the occupied volume(s) that could otherwise come loose and injure an occupant in an emergency landing.
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b. Procedures. The procedures in secti ons 29.561 and 29.561A of this AC 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.
Note: For doors and emergency exit design, when applicable, the rearward load factor to consider is in § 29.561(b)(3)(v).
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AC 29.562 § 29.562 (Amendment 29-29) EMERGENCY LANDING DYNAMIC CONDITIONS.
a. Explanation. Amendment 29 -29 adds new requirements for the dynamic testing of all seats in rotorcraft. This paragr aph is rewritten to incorporate the guidance previously documented in AC 20-137 dated 3/30/92.
b. Background. Improved occupant restraint in civil rotorcraft is addressed in Amendments 29-29 to the airworthiness standards, which add two dynamic crash impact design conditions for seat and occupant restraint systems. This amendment also prescribes a shoulder harness for each occupant and adopts human impact injury criteria as a measure for occupant protection for the dynamic crash impact conditions.
In addition, these amendments significantly improve occupant protection for normal category rotorcraft in a survivable emergency landing. This advisory material addresses the dynamic test conditions and the related pass-fail injury criteria for the dynamic test conditions. This material pertains to single as well as multiple seats and tandem arrangements of the seats in rotorcraft.
(1) Dynamic test methods. This guidance focuses on the use of dynamic test methods for evaluating the performance of rotorcraft seats, restraint systems, and certain related interior systems for demonstrating structural strength and the ability of those systems to protect an occupant from possible injuries in an emergency landing environment represented by the standard. These test methods differ from static test methods, which are limited to demonstrating only the structural strength of the seat or restraint system under ultimate load for at least 3 seconds. This guidance contains sources for appropriate test procedures and provides some insight into the logic of these procedures. It also defines, in part, test facility and equipment characteristics necessary for conducting these tests.
(2) Standardized test methods. Dynamic tests are often conducted at a specially equipped facility, one other than that owned by the designer or manufacturer of the test article. To obtain consistent test results, it is necessary to specify the critical test procedures in detail in the test plan, and then carefully follow these procedures when conducting the tests. This guidance defi nes certain critical procedures for accomplishing the tests of the seat and restraint systems and assessing the data obtained in the tests. Many of these procedures are accepted as standards by government and commercial test facilities and have been modified in this guidance only as necessary for the specific testing of rotorcraft systems.
(3) Relationship of dynamic tests to design standards. This guidance describes test procedures useful in assessing the performance of a rotorcraft seat, Page C - 55
AC 29-2C, Chg 1 2/12/03 restraint system, and interior system. Howeve r, it is impractical to conduct sufficient tests for assessing the performance of the system throughout its entire range of possible uses in unique interior arrangements. The seat, restraint system, and related interior system should be designed for the range of occupants and environments for which it is expected to perform, not just for the dynamic test conditions described in this guidance.
(i) Occupant size. The dynamic tests are conducted with a specific, acceptable, standard anthropomorphic test dummy (ATD) representing a 50th percentile male occupant. Energy absorbing syst ems, restraint system loads and anchorage locations, seat adjustments, seat pitch (for multiple seat rows), head strike envelopes, etc., are typical factors directly influenced by occupant size.
(ii) Seat position and location. T he tests should be sufficient to represent the range of performance expected of a seat and restraint system. A seat, especially an adjustable flight crew seat, should be qualified for those positions approved for take-off and landing. As with static test procedures the seat is also tested to the most critical condition for the dynamic tests. For an adjustable flight crew seat, as an example, the full-up position and longitudinal impact case are expected to be the critical condition.
But these dynamic tests and occupant injury assessment provide a systems approach to qualification. It is therefore necessary to test adjustable seats at the design position for the ATD. Two tests would be required to demonstrate compliance with the strength standards and with the occupant injury criteria. Alternatively adjusting the flight crew seat to its highest position with the interior features, such as an instrument panel shield, raised to maintain the proper perspective or relation to the ATD, is considered an acceptable test procedure for demonstrating compliance with the structural and occupant injury requirements for the seat and its location in a particular cockpit arrangement.
(iii) Test conditions. Only two minimum impact tests are described in the dynamic test procedures discussed in this guidance. These procedures address the tests needed to demonstrate compliance for a typical seat and restraint system installation. Additional tests may be nece ssary to demonstrate compliance for other types or variations of seat and restraint system installations. For example, while only one lateral load direction is specified in the tests, the system should perform properly when similarly loaded from either side.
(iv) Floor deformation. The test procedures require evaluating the effect of certain sidewall or floor deformation. The seat and its attachments and the restraint system should also perform properly if no floor deformation is present.
(v) Head impact. Should such contact occur, head impact with a seat back or the interior of the rotorcraft is evaluated by using a Head Injury Criterion (HIC), which can be measured directly in the tests discussed in this guidance or in supplementary tests of the interior. The design of the inte rior should protect the head Page C - 56
(vi) Emergency egress. Standards for emergency evacuation of the rotorcraft are contained in FAR Part 29. The objective is to allow each occupant to leave the seat and rapidly evacuate the rotorcraft using an exit on either side of the rotorcraft.
c. Dynamic Test Methods and Facilities.
(1) General. A minimum of two dynamic tests are used to assess the performance of the rotorcraft seat, restraint system, and related interior system. The seat, the restraint, and the nearby interior all function together as a system to protect the occupant during emergency landing. The specific test conditions are shown in Figure AC 29.562-1. Ex planations of the test conditions are as follows:
(i) Test 1. The test determines the protection provided when the impact environment is such that the resulting predominant impact load component (vertical) is directed along the spinal column of the occupant in combination with a horizontal (longitudinal) component. Protection against spinal injury is important and it may be necessary to provide energy absorbing (load limiting) or attenuation capability in the seat system in order to comply with the human injury criteria specified in § 29.562 (c)(7).
(ii) Test 2. The test determines the protection provided in an impact where the predominant impact load component is in the longitudinal direction in combination with a lateral component. Evaluation of head injury protection is important in this test if the head could strike some interior portion of the rotorcraft or a forward seat. Chest or spinal column injury, which might result from the upper torso restraint, is also evaluated in this test.
(iii) Tests 1 and 2. These test conditions are also significant for the structural strength of the system. Both tests should be used to assess submarining (where the seat belt slips above the ATD pelvis) and rollout of the upper torso restraint system particularly with single, diagonal torso restraint belts. Since external crash forces frequently cause significant structural deformation, simulated floor deformation is specified for the tests to prove the seat design can accommodate the relative deformation between the seat and the floor or sidewall and still function without imposing excessive loads on the seat, the attachment fittings, or floor tracks.
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