should be sealed to ensure that leaked flammable fluids or vapors do not contact the hot exhaust and cause a fire. The seal design should be reviewed to ensure that the sealing material is fireproof, is chemically compatible with the relevant fuels and vapors, is durable and is functionally adequate. If the design review is not conclusive for compliance purposes, then the seal system should be bench tested under pressure while undergoing critical service loads and motions to ensure no leakage occurs. An analysis should be conducted to determine the structural effects on the exhaust system of the worse case restricted backfire (typically a shock wave analysis can be used to determine the peak internal pressure and, the resultant load on the exhaust system.) If structural failure would occur, based on the analysis, either the backfire restriction should be reduced or the exhaust design should be structurally improved to eliminate the failure.
(10) For compliance with § 27.859(j), each heater's fuel system design must be reviewed to ensure that compliance with the powerplant fuel system requirements of Part 27 that are necessary for safe operation to be achieved. An equivalent safety finding should be made if an application is received that requests partial compliance or non-compliance with the powerplant fuel system requirements of Part 27. The finding should ensure that the safety intent of § 27.859() is achieved. Analysis, engineering evaluation, testing, or a combination should be used to substantiate the heater fuel system design. Heater fuel system components that, by leakage or other failures, can induce flammable fluids or vapors into the ventilating air stream should be shrouded by drainable, fireproof shrouds.
(11) For compliance with § 27.859(k), the drain system design should be reviewed to identify parts that may be subjected to high temperature and parts that may be subjected to hazardous ice accumulation in service. The high temperature parts should be evaluated using the methods of compliance for heater exhausts (reference Paragraph 9, above and Paragraph 549 of this AC). Drains that would be stopped up from ice accumulation should be protected by relocation, size, shields, heating, or a combination to ensure hazardous fluids and vapors are properly drained away.
361. § 27,861 FIRE PROTECTION OF STRUCTURE, CONTROLS, AND OTHER PARTS.
a. Explanation.
(1) As stated in the rule, parts essential to a controlled landing that would be affected by a powerplant fire are to be protected so they can perform their essential functions for at least 5 minutes under any foreseeable powerplant fire condition.
(2) To achieve the objective of the rule, essential parts of the rotorcraft as defined by the rule are to be isolated from a powerplant fire by a firewall (§ 27.1191) or must be protected so they can perform their essential functions for at least 5 minutes under any foreseeable powerplant fire condition.
460 Par 360
(3) Insufficient protection to provide enough time for a controlled landing would represent an unsafe feature or characteristic for the rotorcraft design.
(4) Section 27.1193(d) requires each cowling and engine compartment covering to be at least fire resistant. Also, § 27.1193(e) requires that each part of the cowling or engine compartment covering, subject to high temperature due to its nearness (proximity) to exhaust system parts or exhaust gas impingement, must be fireproof.
(5) In addition, § 27.1194 requires that all surfaces aft of and near powerplant compartments, other than tail surfaces not subject to heat, flames, or sparks emanating trom a powerplant compartment, be at least fire resistant.
b. Procedures.
(1) If each part described in the rule is isolated completely by firewalls, compliance is obtainable.
(2) If each part described by the rule is made of fireproof material, such as steel, compliance is obtained.
(3) If any part described by the rule does not comply with 361b(1) or (2), it shall be proven that it will perform its function under the prescribed conditions. Compliance may be demonstrated by the following criteria:
(i) The parts shall have a positive margin of safety for the appropriate flight and landing condition, including appropriate engine power conditions, under any foreseeable powerplant fire condition. The time interval under consideration here is the time necessary to complete an emergency descent (as described in the flight manual) and landing from the maximum operating altitude for which certification is requested. In no case is the total time interval to be less than 5 minutes.
(il) The factors affecting the time interval should include the maximum height above the terrain, the maximum operating altitude, the flight manual recommendations for rate of descent, and a reasonable time for recognizing a powerplant fire.
(iii) The factors affecting the change in physical characteristics (strength primarily, but stiffness may also be a factor) of the parts are the temperature of the part, time interval at the elevated temperature, size, and heat absorption or rejection.
(iv) The factors affecting the temperature of the part are location and distance from the fire and flames and temperature of the flames (2,000° F +50° F should be used unless proven to be inapplicable).
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(V) The rule requires substantiations for any foreseeable powerplant fire condition. Each rotorcraft design is unique and an evaluation of each design is necessary to establish the fire and flight conditions under consideration.
(vi) A very brief and simple example of compliance noted here may be helpful. This example pertains to a single-engine rotorcraft with the engine mounted on top at the fuselage centerline. The engine is supported by all steel tubular mounts.
The fuselage panel serves as a work deck as well as a firewall. A 15-minute duration is appropriate for this design. A representative panel of the firewall (deck) skin may be subjected to the autorotational flight loads and the landing load. A flame from an appropriate-sized burner, measuring 2,000° F t50° F at the skin surface, should impinge on the loaded panel for 15 minutes. The panel may deform but must remain intact and sustain the appropriate load. The flame should not penetrate the panel skin.
(vii) Other rotorcraft designs may have engines located on top of the fuselage under the main rotor. If cowls or firewalls do not isolate the rotors and essential controls, it must be determined by a rational analysis or by temperature measurement that the rotor and essential controls will perform their functions. Air flow through the rotor and factors noted in Paragraphs 361b(3)(ii, (3) iii), and (3)(iv) are important to an analysis.
361A. § 27.861 (Amendment 27-26) FIRE PROTECTION OF STRUCTURE, CONTROLS, AND OTHER PARTS.
a. Explanation.
(1) Amendment 27-26 revised the regulation to allow use of parts made from standard fireproof materials of known acceptable dimensions in areas affected by powerplant fires without further proof or qualification. Previously, the standard imposed a performance criteria regardless of the materials and part dimensions used.
(2) "Fireproof" and "fire resistant" are defined in FAR Part 1, § 1.1.
b. Procedures.
(1) A part of acceptable geometry made of steel, or another fireproof material, may be used to comply with the standard.
(2) A material system, panel, or assembly would be equivalent to steel provided it successfully completes the flammability tests described in Paragraph 361b(3) (vi) of this AC, for Category B rotorcraft adjusted for the time period appropriate to the rotorcraft application.
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(3) It is appropriate to further define "fire resistant." A material system, panel, or assembly would be equivalent to aluminum (fire resistant) if it successfully completes the following flammability test. Locate a specimen (approximately 8 inches square) of the material system panel, assembly, or part at approximately a 45° angle to a horizontal line. Apply the limit or normal operating load. Impinge a 2000 +150° F (1093 + 83°C) flame on the article for at least 5 minutes duration. Flame penetration of the test article is not allowed. In addition, the part or component should be able to perform its intended function, as installed, during and after the test.
362. § 27,863 (Amendment 27-16) FLAMMABLE FLUID FIRE PROTECTION
a. Background.
(1) The development of § 27.863 can be traced through CARs 6.485 and
6.486, and subsequent Amendment 27-16.
(2) Investigation of several accidents disclosed evidence of in-flight fires caused by leakage of flammable fluids to ignition sources. The revisions to § 27.863 adopted by Amendment 27-16 require significantly more attention to overall fire protection and prevention.
b. Explanation.
(1) Prior to Amendment 27-16, this rule only required either a means to prevent ignition of flammable fluids or vapors or a means to control any resulting fire. Isolation of flammable fluids and vapors from ignition sources by shrouding or sealing was the normal method of compliance. With Amendment 27-16, the rule further requires the assumption that these means fail or are ineffective and a fire does occur. Means to minimize the consequence of these fires must be provided. Specifically identified considerations must include the flammability of any combustible or absorbing materials, electrical faults, malfunction of protective devices, and so forth.
(2) The rule does not require the entire rotorcraft to be a "designated fire zone."
c. Methods of Compliance.
(1) To minimize the probability of ignition of fluids and vapors after single failure of a component or systems, the following methods may be used:
(i) Shroud and drain flammable fluid systems (including steel fluid lines, fittings, etc.) and provide the systems with fuel and vapor seals with respect to potential ignition sources (electrical wiring and equipment, hot bleed air lines, etc.).
(ii) Provide other effective separation, ventilation, or overheat shutdown devices, etc., to preclude ignition.
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AC 27-1A 7130/97
(iii) Ensure that electrical equipment in the areas subject to flammable fluids and vapors is either hermetically sealed or has been tested and shown to be free of ignition capability. Paragraph 621 of this advisory circular describes acceptable guidance for such laboratory testing.
(iv) Place a restricting orifice in fluid pressure lines routed to instruments and transducers.
(v) Ensure fluid lines are not located so as to be subject to abrasion during normal operations. Cargo compartments should be evaluated for potential line damage due to cargo movement.
(2) To minimize hazards if ignition occurs:
(i) Provide fireproof designs, firewall isolation, or equivalent means for critical structure, equipment, and personnel areas.
(її) Consider fire detection, extinguishment, shutoff valves, fire suppression systems, etc.
(3) In considering compliance, the actual protective measures may be related to the situation, considering the quantity and flammability characteristic of the fluid, the fire damage tolerance of the area, and the means available to the crew to minimize hazards from the fire. If action by the crew is necessary, a quick acting means (not necessarily fire detectors) must be provided to alert the crew in the event of a fire.
Details of any action required by the crew should be included in the Rotorcraft Flight Manual.
(4) Compliance with § 27.863(d) requires, a minimum, type design data defining each area where flammable fluids or vapors might escape.
363.-372. RESERVED.
464 (thru 474) Par 362
SECTION 22. EXTERNAL LOAD ATTACHING MEANS.
373. § 27.865 (Amendment 27-11) EXTERNAL LOAD ATTACHING MEANS.
a. Explanation.
(1) If certification for external load operations is requested, the rule requires that the external load attaching means be substantiated by test or analysis for a limit static load equal to or greater than 2.5 times the maximum external load for which certification is requested. The factor of 2.5 times the maximum external load was established as a minimum strength requirement by Part 133 operations to account for loading effects of sling-load angles up to 30° from the vertical. Allowance for reducing the 30° angle is provided if substantiated.
(2) The rule requires that a quick-release device be installed on one of the pilot's primary controls so the pilot can quickly release the external load during an emergency situation. In addition, a backup manual mechanical control for the quick-release device is required to be readily accessible to either the pilot or another crewmember.
(3) The rule requires appropriate placards or markings stating the maximum authorized external load.
b. Procedures.
(1) The maximum external load for which authorization is requested should not exceed the rated capacity of the quick-release device. The quick-release device should be strength tested (with FAA/AUTHORITY witness) if it is not produced to a recognized industry or military standard.
(2) Substantiation of external loading requirements must include any direction making an angle of 30° (with the exception of directions having a forward component).
(reference § 27.865(a).)
The sling-load angle (i.e., the angle between the vertical direction and the sling-load cable supporting the external load) should not exceed an angle of 30° to minimize the cable tension load.
(ii) The 30° angle may be reduced if an operating limitation is established limiting external load operations to such angles for which compliance has been shown or if the reduced angle cannot be exceeded in service. The lesser angle should be substantiated by flight testing.
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АС 27-1A 7/30/97
(3) The external load releasing system is specified to include a quick release device installed on one of the pilot's primary controls. It is usually installed on the cyclic stick to allow the pilot to release the load with minimum distraction after maneuvering the load into the release position.
(4) A manual mechanical control for the quick-release device is specified to be installed and be readily accessible to the pilot or to another crewmember. A sufficient amount of slack should be provided in the control cable to permit complete cargo movement without tripping the cargo release.
373A. § 27.865 (Amendment 27-26) EXTERNAL LOAD ATTACHING MEANS.
a. Explanation. Amendment 27-26 added two requirements to § 27.865:
(1) Section 27.865(a) is clarified to allow use of a design factor less than
2.5g's, for rotorcraft load combinations A, B, and C non-human external cargo applications provided the lower load factor is not likely to be exceeded by virtue of the rotorcraft characteristics and capability. That is, the rotorcraft design factors may be used for the cargo device system.
(2) Section 27.865(d) was added to clarify and specify the fatigue requirements for the external cargo attaching means.
b. Procedures. All of the policy material pertaining to this section remains in effect with the following additions:
(1) For § 27.865(a), if a design limit load factor less than 2.5g's is requested, the applicant should provide a rational analysis and/or a flight operations data base that clearly shows that the load factor requested is unlikely to be exceeded in service.
NOTE: § 27.337(b) requires use of 2.0 g's as a minimum.
(2) § 27.865(d), all failures of the cargo attaching means (and the associated critical components) that are likely to be hazardous to the rotorcraft should be identified by an acceptable means, such as a Failure Mode Effect Analysis (FMEA). The critical components associated with these failures modes should receive a fatigue analysis and/or test to ensure that the likelihood of a failure occurring is minimized. In the majority of cases, an analysis using the methods of AC 20-95, "Fatigue Evaluation of Rotorcraft Structure", will be sufficient. If any component has a service life and/or mandatory inspection these components and each mandatory life should be identified, approved, and placed in the airworthiness limitations section of the maintenance manual or Instructions for Continued Airworthiness. See Paragraph 729 (§27.1529) of the AC for information on these manuals.
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374.-382. RESERVED.
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383. RESERVED.
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7/30/97 АС 27-1A SECTION 23. MISCELLANEOUS (DESIGN AND CONSTRUCTION)
384. § 27.871 LEVELING MARKS.
a. Explanation. Reference marks are required for leveling the rotorcraft on the ground. These marks are necessary for accurate determination of weight and balance effects, particularly after modifications to the basic rotorcraft.
b. Procedures.
(1) Reference marks are sometimes provided in pairs, one high in the cabin and one low. The plumb weight is suspended from the high mark by an appropriate mechanical attachment, and the lower mark is used to level the rotorcraft by centering the plumb weight. The lower reference mark should be a raised or depressed target symbol and shall be applied to a permanent structural component or permanently attached plate in a readily accessible location. Seat tracks, floors, or door sills which are attached with permanent fasteners are typical locations.
(2) Horizontal reference marks, for support of bubble levels, may also be used, particularly for smaller rotorcraft.
385. § 27.873 BALLAST PROVISIONS.
a. Explanation.
(1) This rule requires that ballast provisions prevent inadvertent ballast shifting while in flight or as a result of a landing. Shifting of the ballast may cause a hazardous change in the center of gravity thereby affecting rotorcraft controllability.
(2) Other rules noted here allow removable and fixed ballast and require markings or placards to prevent overloading the ballast installation.
Section 27.29 specifies that the rotorcraft empty weight will include any fixed ballast. Section 27.31 allows the use of removable ballast to comply with the flight requirements. However, ballast may not be adjusted (moved, reduced, or increased) in flight.
(її) Section 27.1541 requires conspicuous and durable markings or placards. Section 27.1557 requires placards stating allowable maximum weight, distributed loading, if necessary, and other appropriate limitations for ballast installation.
(3) Section 27.1583(c) concerns Rotorcraft Flight Manual instructions and information about removable ballast or loading information. The instructions must be Par 384 487
included in the operating limitations section of the flight manual to allow ready observance of the limitations.
b. Procedures.
(1) The ballast installation may be substantiated by analysis or by static test.
The design ultimate load may be derived from flight, landing, or minor crash conditions load factors specified in the rules. Substantiation by analysis will require use of the fitting factor prescribed by § 27.625 where appropriate. If static tests are to be conducted, a test plan should be prepared, submitted for evaluation, and agreed upon prior to the test.
(2) Ballast installations in the aft part of the fuselage and tail boom may be subject to significant landing condition angular inertia load factors as well as the usual linear load factors.
(3) Substantiation methods and procedures acceptable for the airframe substantiation may be used for the ballast installation as well.
(4) Removable ballast will require attention to ensure the ballast is secured easily and properly and will remain secured under the appropriate ballast design load factor requirements. The flight manual instructions should be evaluated for compliance with § 27.1583(c) by flight test and airframe personnel.
(5) The installation must be designed and placarded or marked for the maximum allowable ballast load and for other appropriate loading limits. Normally compliance with § 27.1541 is accomplished with a drawing review by airframe personnel along with a MIDO or FSDO compliance and conformity inspection. An additional compliance inspection by airframe personnel can be conducted if desired.
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386.-396. RESERVED.
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