design integrity to show that the basic airworthiness requirements which are applicable to the rotors will be met.
A design assessment of the rotors should be carried out in order to substantiate that they are of a safe design and that compensating provisions are made available to prevent failures classified as hazardous and catastrophic in the sense specified in Paragraph b below. In carrying out the design assessment, the results of the certification ground and flight testing (including any failures or degradation) should be taken into consideration. Previous service experience with similar designs should also be taken into account (see also FAR 29.601(a)).
b. Definitions.
For the purposes of this assessment, failure conditions may be classified according to the severity of their effects as follows:
(1) Minor. Failure conditions which would not significantly reduce rotorcraft safety, and which involve crew actions that are well within the crew capabilities. Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as routine flight plan changes, or some inconvenience to occupants.
(2) Major. Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew work load or in conditions impairing crew efficiency, or discomfort to occupants, possibly including injuries.
(3) Hazardous. Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be -
(i) A large reduction in safety margins or functional capabilities.
(ii) Physical distress or higher workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely.
(ili) Serious or fatal injury to a relatively small number of the occupants.
(iv) Loss of ability to continue safe flight to a suitable landing site.
(4) Catastrophic. Failure conditions which would prevent a safe landing.
318 Par 205A
(5) Minimize. Reduce to a minimum, decrease to the least possible amount that can be shown to be both technically feasible and economically justifiable to the certification authority.
(6) Health Monitoring. Equipment, techniques and/or procedures by which selected incipient failure or degradation can be determined.
c. Procedures.
(1) Eailure 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 formatted, 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:
(• Each item/component function(s).
(її) Item/component failure modes and their causes.
(¡i) 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 2 below)
(vii) The failure condition severity classification according to the definitions given in (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.
Par 205A 319
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.
(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.
Design features; i.e., safety factors, part derating criteria, redundancies, etc.
(ii) A high level of integrity.
(ili) 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.
(vili) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.
(x) Safety devices or health monitoring means beyond those identified in
(vi) and (vii) above.
320 Par 205A
206. § 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) Auxiliary rotor thrust;
(ii) The torque reaction of each rotor drive system; and
(iii) Balancing air and inertia loads.
(2) Each engine 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 "OEl 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 Paragraph (e) of this standard.
207. § 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.
208.-217. RESERVED Par 206 321 (thru 331)
SECTION 13. EMERGENCY LANDING CONDITIONS
218. § 29.561 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) Section 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) 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 must be restrained to the load factors in § 29.561(b) (3) if located adjacent to, above, or behind the occupants.
(4) For aircraft equipped with retractable landing gear, the landing gear must be retracted for compliance.
(5) Fuel tank protection.
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 specifies that each tank and its installation must be designed or protected to retain fuel without leakage under the emergency landing conditions in § 29.561. Paragraph 454 of this AC relates to this standard.
(6) The minor crash conditions contained in § 29.561(b)(3) must also be considered in designing doors and exits (§ 29.783(d) and (g), and § 29.809(e)).
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.
332 Par 218
(2) Section 29.785 (Paragraph 336 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 highest load factor derived must be used.
For example, for light weight conditions, the ultimate maneuvering load factor may be 5.25g as specified in § 29.337.
(її) 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.
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/crushing characteristics.
(II) 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.
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, literafts, engines, and/or transmissions must be restrained for the appropriate load factors.
(6) Cargo/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.)
Under-floor 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.
NOTE: Fuselage keels whose design and structural strength are such as to resist crash impacts associated with the emergency landing conditions of Par 218 333
§ 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.
218A. § 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:
(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 and/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 of Paragraph 218, § 29.561, 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.
218B. § 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 rearward 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 volumes) 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 334 Par 218
items of mass above and behind the occupied volume(s) that could otherwise come loose and injure an occupant in an emergency landing.
b. Procedures. The procedures of Paragraphs 218 and 218A 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. § 29.562 EMERGENCY LANDING DYNAMIC CONDITIONS.
a. Explanation. Amendment 29-29 adds new requirements for the dynamic testing of all seats in rotorcraft.
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 § 29. 562 using the 170-pound anthropomorphic test dummy specified in § 29.562(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. § 29.563 (Amendment 29-12) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-12 included certification 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 337, § 29.801, for procedures.
220A. § 29.563 (Amendment 29-30) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-30 added specific 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:
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 O to 30 knots with a vertical descent rate of not less than 5 FPS relative to the mean water surface.
Par 218B 335
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 § 29.563(b)(i). 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.
(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 337 (§ 29.801) 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:
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).
(i¡) 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.
336 Par 220A
(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).
Par 220A 337
221.-229. RESERVED.
Intentionally Left Blank Par 221 338 (thru 346)
7/30/97 АС 29-2B SECTION 14. FATIGUE EVALUATION
230. § 29.571 FATIGUE EVALUATION OF FLIGHT STRUCTURE.
a. Explanation. An evaluation is required to assure structural reliability of the rotorcraft in flight. 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 sate as prescribed.
b. Procedures.
(1) The fatigue evaluation requires consideration of the following factors:
Identification of the structure/components to be considered.
(її) The stress during operating conditions.
(Ill) The operating spectrum or frequency of occurrence.
(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 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 deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.
(ill) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.
Par 230 347