SECTION 11. WATER LOADS
193. § 29.519 (Amendment 29-30) HULL TYPE ROTORCRAFT:
WATER-BASED, AMPHIBIAN.
a. Explanation.
(1) This regulation provides design criteria for amphibian rotorcraft with hull provisions.
(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 for a description of sea state 4 conditions).
(3) A rotor lift of two-thirds of the rotorcraft weight may be applied during landing impact.
(4) Vertical landing conditions are specified as:
(i) Zero forward speed.
(il) Likely pitch and roll attitudes.
(IlI) Vertical descent velocity ≥ 6.5 FPS.
(5) Forward speed landing conditions are specified as:
(i) Forward velocities of zero to 30 knots (a 30-knot limit may be reduced if it can be demonstrated that the maximum forward velocity selected would not be exceeded in a normal one-engine-out landing).
(її) Likely pitch, roll, and yaw attitudes.
(ili) Vertical descent velocity ≥ 6.5 FPS.
(6) Auxiliary float immersion conditions are specified to be applied unless it can be shown that full immersion is unlikely. If full immersion is unlikely, the highest float buoyancy load is specified that considers loading of the float immersed to create restoring moments which compensate for upsetting moments caused by side wind, asymmetrical rotorcraft loading, water wave action, and rotorcraft inertia.
b. Procedures.
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(1) Tests should be conducted to establish procedures for water entry. These tests should include determination of optimum pitch attitude and forward velocity for landing in a calm sea as well as entry procedures for the highest sea state to be demonstrated (e.g., the recommended part of the wave on which to land and direction of landing relative to crest/trough direction).
(2) 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:
Vertical Landing Conditions.
(A) Zero forward velocity.
(B) The optimum pitch attitude as determined in Paragraph 193b(1) with consideration for pitch attitude variations that would reasonably be expected to occur in service.
(C) Vertical descent velocity of 6.5 FPS or greater.
(D) Likely roll attitudes.
(i¡) Forward Speed Landing Conditions.
(A) Forward velocities of zero 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).
(B) The optimum pitch attitude as determined in Paragraph 193b(1) with consideration for pitch attitude variations that would reasonably be expected to occur in service.
(C) Vertical descent velocity of 6.5 FPS or greater.
(D) Likely roll and yaw attitudes.
(3) Landing load factors may be determined by--
(1) Landing gear drop tests for limited amphibian;
(il) Water drop tests for amphibian; or
(ili) Analysis based on tests.
Par 193 305
- AC 29-2B 7/30/97
(4) Water load distribution should be determined by tests or analysis based on tests.
(5) Auxiliary float loads should be determined by full immersion or restoring moments required to react upsetting moments caused by side wind, asymmetrical rotorcraft loading, water wave action, and rotorcraft inertia. Auxiliary float loads may be determined by analysis. Load distributions should be determined by tests or analysis based on tests.
194. § 29.521 (Amendment 29-3) FLOAT LANDING CONDITIONS.
a. Explanation. This is an optional requirement for float operations, and it applies only when float operations are requested. The regulation specifies vertical loads, aft loads, and side loads to be applied to the float installations. The two loading conditions to be applied are:
(1) Up-load Condition.
(i> A vertical load appropriate to a landing load factor determined under § 29.473(b).
(®) The resultant water reaction passes vertically through the aircraft CG.
(iii) An aft load equal to 25 percent of the vertical load.
(2) Side-load Condition.
(i> A vertical load equal to 75 percent of the vertical load for the up-load condition.
(ї) Vertical load equally divided among the floats.
(iii) A side load at each float equal to 25 percent of the vertical load at each float.
b. Procedures.
(1) The vertical load factor is determined by drop tests in accordance with §§ 29.473(b) and 29.725. The floats may be drop tested, or they may be assumed to have the same load factor as wheeled gear which have been drop tested.
(2) Structural substantiation may be accomplished by either static tests or analysis using the specified loads. The load distribution on the floats may be realistically based on hydrostatic pressure distributions or conservative pressure distributions.
306 Par 193
195.-203. RESERVED.
Intentionally Left Blank Par 195 307 (thru 315)
204. RESERVED.
Intentionally Left Blank 316 Par 204
SECTION 12. MAIN COMPONENT REQUIREMENTS
205. § 29.547 (Amendment 29-4) MAIN ROTOR STRUCTURE.
a. Explanation. This regulation requires the main rotor structure to be designed to the static load requirements of §§ 29.337 through 29.351 (vertical maneuvering loads, vertical and horizontal gust loads, and yawing maneuver loads). In addition, the main rotor blades, hubs, and flapping hinges are specified to be designed for impact forces of each blade against its stop during ground operation and for specified limit torque at any rotational speed including zero. The torque forces (from the drive system) are distributed to the rotor blades as specified.
b. Procedures.
(1) Substantiation in compliance with this standard is accomplished by application of the flight loads of §§ 29.337 through 29.351 and the torque loads of § 29.361 to the rotor structure by stress analyses and/or static tests. The use of wind tunnel data as well as flight loads survey data may be used to generate and/or check the external load magnitudes and distributions.
(2) Where new materials are used in the main rotor structure, such as composites containing plastics, the effects of temperature and humidity are to be considered in accordance with § 29.603, and the effects of uncertainties in manufacturing processes or inspection methods are to be considered in accordance with § 29.619.
(3) The design impact forces of each blade must be imposed against the blade stop or stops. Impact loads from 2 to 3 g's have been commonly used to provide rotor structure protection against blades impacting against lower (droop) stops. Different values may be used for flapping and lag stops as determined by a rational basis.
Appropriate monitoring of the blades, hubs, flapping hinges, and stops during laboratory tests, ground endurance tests, and flight tests should ensure that the stops are sufficient for ground operation loads (taxiing, backing, etc.), training, and offshore platform landings. Taxiing should consider typical obstacles such as pavement edges, ropes, air lines, and so forth. The design torque loads are derived as prescribed.
205A. § 29.547 (Amendment 29-40) MAIN ROTOR AND TAIL ROTOR STRUCTURE.
a. Explanation. Amendment 29-40 revised § 29.547 to add requirements to perform a design assessment. FAR 29.547 (a) and (b) set forth a definition of a rotor and its associated components and requires a design assessment to be performed.
The intent of these paragraphs is to identify the critical components and/or clarify their Par 205 317
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.
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(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.
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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.
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(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