airflow becomes inadequate for safe operation.

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-2, 22 F.R. 5569, July 16, 1957]

§ 6.384 Fire protection of structure, controls, and other parts.

All structure, controls, rotor mechanism, and other parts essential to a controlled landing of the rotorcraft which would be affected by powerplant fires shall either be of fireproof construction or shall be otherwise protected, so that they can perform their essential functions for at least 5 minutes under all foreseeable powerplant fire conditions. (See also §§ 6.480 and 6.483 (a).)

§ 6.400

Scope and general design.

(a) The powerplant installation shall be considered to include all components of the rotorcraft which are necessary for its propulsion with the exception of the structure of the main and auxiliary rotors. It shall also be considered to include all components which affect the control of the major propulsive units or which affect their safety of operation between normal inspections or overhaul periods. (See §§ 6.604 and 6.613 for instrument installation and marking.) The general provisions of paragraphs (b) through (d) of this section shall be applicable.

(b) All components of the powerplant installation shall be constructed, arranged, and installed in a manner which will assure their continued safe operation between normal inspections or overhaul periods.

(c) Accessibility shall be provided to permit such inspection and maintenance as is necessary to assure continued airworthiness.

(d) Electrical interconnections shall be provided to prevent the existence of differences of potential between major components of the powerplant installation and other portions of the rotorcraft. § 6.400-1 Powerplant installation components (FAA interpretations which apply to § 6.400).

The term "all components" includes engines and their parts, appurtenances, and accessories which are furnished by the engine manufacturer and all other components of the powerplant installation which are furnished by the rotorcraft manufacturer. For example: Fuel pumps, lines, valves, and other components of the fuel system which are integral parts of the type certificated engine are also components of the rotorcraft powerplant installation.

[Supp. 16, 23 F. R. 9018, Nov. 20, 1958]

(a) All engines shall be type certificated in accordance with the provisions of Part 13 of this subchapter.

(b) Engine cooling fan blade protection. If an engine cooling fan is installed, means shall be provided to protect the rotorcraft and to permit a safe landing in the event of a fan blade failure. Compliance shall be shown with any one of the provisions of subparagraphs (1) through (3) of this paragraph.

(1) It shall be demonstrated that the fan blades will be contained in the event of failure;

(2) The fan is so located that a fan blade failure will not jeopardize the safety of the rotorcraft or its occupants;

or

(3) It shall be demonstrated that the fan blade can withstand an ultimate load of 1.5 times the centrifugal force resulting from engine rpm limited by either:

(i) The engine terminal rpm which can occur under uncontrolled conditions,

or

(ii) An overspeed limiting device. [21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-3, 23 F.R. 2592, Apr. 19, 1958] § 6.402 Engine vibration.

The engine shall be installed to preclude harmful vibration of any of the engine parts or of any of the components of the rotorcraft. It shall be demonstrated by means of a vibration investigation that the addition of the rotor and

the rotor drive system to the engine does not result in modification of engine vibration characteristics to the extent that the principal rotating portions of the engine are subjected to excessive vibratory stresses. It shall also be demonstrated that no portion of the rotor drive system is subjected to excessive vibratory stresses.

ROTOR DRIVE SYSTEM

§ 6.410 Rotor drive mechanism.

The rotor drive mechanism shall incorporate a unit which will automatically disengage the engine from the main and auxiliary rotors in the event of power failure. The rotor drive mechanism shall be so arranged that all rotors necessary for control of the rotorcraft in autorotative flight will continue to be driven by the main rotor(s) after disengagement of the engine from the main and auxiliary rotors. If a torque limiting device is employed in the rotor drive system (see § 6.250 (f)), such device shall be located to permit continued control of the rotorcraft after it becomes operative. § 6.411 Rotor brakes.

If a means is provided to control the rotation of the rotor drive system independent of the engine, the limitations on the use of such means shall be specified, and the control for this means shall be guarded to prevent inadvertent operation.

§ 6.412 Rotor drive and control mechanism endurance tests.

(a) The rotor drive and control mechanism shall be tested for not less than 100 hours. The test shall be conducted on the rotorcraft, and the power shall be absorbed by the actual rotors to be installed, except that the use of other ground or flight test facilities with any other appropriate method of power absorption shall be acceptable provided that all conditions of support and vibration closely simulate the conditions which would exist during a test on the actual rotorcraft. The endurance tests shall include the tests prescribed in paragraphs (b) through (g) of this section. At the conclusion of the endurance testing, all parts shall be in a serviceable condition.

(b) A 60-hour portion of the endurance test shall be run at not less than the maximum continuous engine speed in conjunction with maximum continu

ous engine power. In this test the main rotor shall be set in the position which will give maximum longitudinal cyclic pitch change to simulate forward flight. The auxiliary rotor controls shall be in the position for normal operation under the conditions of the test.

(c) A 30-hour portion of the endurance test shall be run at not less than 90 percent of maximum continuous engine speed and 75 percent of maximum continuous engine power. The main and auxiliary rotor controls during this test shall be in the position for normal operation under the conditions of the test.

(d) A 10-hour portion of the endurance test shall be run at not less than take-off engine power and speed. The main and auxiliary rotor controls shall be in the normal position for vertical ascent during this test.

(e) The portions of the endurance test prescribed in paragraphs (b) and (c) of this section shall be conducted in intervals of not less than 30 minutes and may be accomplished either on the ground or in flight. The portion of the endurance test prescribed in paragraph (d) of this section may be conducted in intervals of 5 minutes or more.

(f) At intervals of not more than every 5 hours during the endurance tests prescribed in paragraphs (b), (c), and (d) of this section the engine shall be stopped rapidly enough to allow the engine and rotor drive to be automatically disengaged from the rotors.

(g) There shall be accomplished under the operating conditions specified in paragraph (b) of this section 500 complete cycles of lateral control and 500 complete cycles of longitudinal control of the main rotors, and 500 complete cycles of control of all auxiliary rotors. A complete control cycle shall be considered to involve movement of the controls from the neutral position, through both extreme positions, and back to the neutral position, except that control movement need not produce loads or flapping motions exceeding the maximum loads or motions encountered in flight. The control cycling may be accomplished during the testing prescribed in paragraph (b) of this section or may be accomplished separately. § 6.413

Additional tests.

Such additional dynamic, endurance, and operational tests or vibratory investigations shall be conducted as are

found necessary by the Administrator to substantiate the airworthiness of the rotor drive mechanism.

§ 6.414 Shafting critical speed.

The critical speeds of all shafting shall be determined by actual demonstration, except that analytical methods shall be acceptable for determining these speeds if the Administrator finds that reliable methods of analysis are available for the particular design. If the critical speeds lie within or close to the operating ranges for idling, power-on, and autorotative conditions, it shall be demonstrated by tests that the resultant stresses are within safe limits. If analytical methods are used and indicate that no critical speeds lie within the permissible operating ranges, the margins between the calculated critical speeds and the limits of the permissible operating ranges shall be adequate to allow for possible variations of the computed values from actual values.

(a) The fuel supply system shall be arranged so that, in so far as practicable, the entire fuel supply can be utilized in the maximum inclination of the fuselage for any sustained conditions of flight, and so that the feed ports will not be uncovered during normal maneuvers involving moderate rolling or sideslipping. On rotorcraft with more than one fuel tank (see § 6.422(e)) the system shall feed fuel promptly after one tank is turned off and another tank is turned on, and there shall be installed in addition to the fuel quantity indicator (see § 6.604(d)) a warning device to indicate when the fuel in any tank becomes low.

NOTE: The fuel in any tank is considered to be low when there remains approximately a five-minute supply with the rotorcraft in the most critical sustained flight attitude.

(b) Fuel system independence. The design of the fuel system for multiengine rotorcraft shall be such as to permit fuel to be supplied to each engine

through a system independent of all portions of the systems supplying fuel to other engines, except that separate fuel tanks need not be provided for each engine. The following features shall be provided if a single fuel tank is employed on a multiengine rotorcraft:

(1) Independent tank outlets for each engine. Each outlet shall incorporate a shutoff valve at the tank. This valve may also serve as the firewall shutoff valve required by § 6.426 provided the line between the valve and the engine compartment does not contain a hazardous amount of fuel which can drain into the engine compartment.

(2) At least 2 vents arranged to minimize the possibility of both vents becoming obstructed simultaneously.

(3) Filler caps designed to minimize the possibility of incorrect installation or loss in flight.

(4) The fuel system from the tank outlet to the engine shall be entirely independent of any portion of the system supplying fuel to the other engine(s).

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959]

§ 6.421

Unusable fuel supply.

The unusable fuel supply in each tank shall be that quantity at which the first evidence of malfunctioning occurs in any sustained flight condition at the most critical weight and center of gravity position within the approved limitations. The unusable fuel supply shall be determined for each tank used in normal operation. (See also §§ 6.104, 6.736, and 6.741(f).)

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959]

§ 6.422 Fuel tank construction and installation.

Fuel tanks shall be designed and installed in accordance with the provisions of paragraphs (a) through (e) of this section.

(a) Fuel tanks shall be capable of withstanding without failure all vibration, inertia, fluid, and structural loads to which they may be subjected in operation.

(b) Fuel tanks shall be capable of withstanding, without failure or leakage, an internal pressure equal to the pressure developed during the maximum limit acceleration with full tanks, except that in no case shall the minimum

internal pressure be less than 3.5 lb./sq. in. for conventional type tanks or less than 2.0 lb./sq. in. for bladder type tanks.

(c) Fuel tanks of 10 gallons or greater capacity shall incorporate internal baffles unless external support is provided to resist surging.

(d) Fuel tanks shall be separated from the engine compartment by a fire wall. At least one-half inch clear air space shall be provided between the tank and fire wall.

(e) Spaces adjacent to the surfaces of fuel tanks shall be ventilated so that fumes cannot accumulate in the tank compartment in case of leakage. If two or more tanks have their outlets interconnected, they shall be considered as one tank. The air spaces in such tanks shall be interconnected to prevent the flow of fuel from one tank to another as a result of a difference in pressure in the respective tank air spaces.

(a) Expansion space. Fuel tanks shall be provided with an expansion space of not less than 2 percent of the tank capacity. It shall not be possible to fill the fuel tank expansion space inadvertently when the rotorcraft is in the normal ground attitude.

(b) Sump. Each fuel tank shall incorporate a sump and drain located at the point in the tank which is the lowest when the rotorcraft is in the normal ground attitude. The main fuel supply shall not be drawn from the bottom of the sump.

(c) Filler connection. The design of fuel tank filler connections shall be such as to prevent the entrance of fuel into the fuel tank compartment or to any other portion of the rotorcraft other than the tank itself. (See also § 6.738 (b) (1).)

(d) Vents. Fuel tanks shall be vented from the top portion of the expansion space in such a manner that venting of the tank is effective under all normal flight conditions. The air vents shall be arranged to minimize the possibility of stoppage by dirt or ice formation.

(e) Outlet. Fuel tank outlets shall be provided with large-mesh finger strainers.

§ 6.424 Fuel pumps.

If a mechanical pump is employed, an emergency pump shall also be installed to be available for immediate use in case

of failure of the mechanical pump. The emergency pump shall be actuated automatically or operated continuously such that sufficient fuel pressure will be maintained to prevent engine stoppage after failure of the mechanical pump. Means shall be provided for indication to the pilot when the emergency system is in operation. Pumps of appropriate capacity may also be used for pumping fuel from an auxiliary tank to a main fuel tank. Mechanical pump systems shall be so arranged that they cannot feed from more than one tank at a time. [21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-4, 24 F.R. 7073, Sept. 1, 1959]

§ 6.425 Fuel system lines and fittings.

(a) Fuel lines shall be installed and supported to prevent excessive vibration and to withstand loads due to fuel pressure and due to accelerated flight conditions.

(b) Fuel lines which are connected to components of the rotorcraft between which relative motion could exist shall incorporate provisions for flexibility.

(c) Flexible hose shall be of an approved type.

(d) All fuel lines and fittings shall be of sufficient size so that the fuel flow, with the fuel being supplied to the carburetor at the minimum pressure for proper carburetor operation, is not less than the following:

(1) For gravity feed systems: 1.5 times the normal flow required to operate the engine at take-off power;

(2) For pump systems: 1.25 times the normal flow required to operate the engine at take-off power.

(e) Rotorcraft with suction lift fuel systems or other fuel system features conducive to vapor formation shall be demonstrated to be free from vapor lock when using fuel at a temperature of 110° F. under critical operating conditions.

(f) A test for proof of compliance with the applicable flow requirements shall be conducted.

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-2, 22 F.R. 5569, July 16, 1957] § 6.426

Valves.

A positive and quick-acting valve which will shut off all fuel to each engine individually shall be provided. The control for this valve shall be within easy reach of appropriate flight per

One or more accessible drains shall be provided at the lowest point in the fuel system to drain completely all parts of the system when the rotorcraft is in its normal position on level ground. Such drains shall discharge clear of all parts of the rotorcraft and shall be equipped with safety locks to prevent accidental opening.

§ 6.429 Fuel quantity indicator.

The fuel quantity indicator (see § 6.613 (b)) shall be installed to indicate clearly to the flight crew the quantity of fuel in each tank while in flight. When two or more tanks are closely interconnected by a gravity feed system and vented, and when it is impossible to feed from each tank separately, only one fuel quantity indicator need be installed. If exposed sight gauges are employed they shall be installed and guarded to preclude the possibility of breakage or damage.

(a) Each engine shall be provided with an independent oil system capable of supplying the engine with an appropriate quantity of oil at a temperature not exceeding the maximum which has been established as safe for continuous operation. (For oil system instruments see §§ 6.604 and 6.735.)

(b) The usable oil capacity shall not be less than the product of the endurance of the rotorcraft under critical operating conditions and the maximum oil consumption of the engine under the same conditions, to which product a suitable margin shall be added to assure adequate circulation and cooling of

the oil system. In lieu of a rational analysis of rotorcraft endurance and oil consumption, the usable oil capacity of 1 gallon for each 40 gallons of usable fuel quantity shall be considered acceptable. (See also § 6.101 (d) (3).)

(c) The ability of the oil cooling provisions to maintain the oil inlet temperature to the engine at or below the maximum established value shall be demonstrated by flight tests.

§ 6.441

Oil tank construction and installation.

Oil tanks shall be designed and installed in accordance with the provisions of paragraphs (a) through (e) of this section.

(a) Oil tanks shall be capable of withstanding without failure all vibration, inertia, fluid, and structural loads to which they may be subjected in operation.

(b) Oil tanks shall be capable of withstanding without failure or leakage an internal pressure of 5 lb./sq. in.

(c) Oil tanks shall be provided with an expansion space of not less than 10 percent of the tank capacity, nor less than one-half gallon. It shall not be possible inadvertently to fill the oil tank expansion space when the rotorcraft is in the normal ground attitude.

(d) Oil tanks shall be vented.

(e) Provision shall be made in the filler opening to prevent oil overflow from entering the compartment in which the oil tank is located. (See also § 6.738 (b) (2).)

[21 F.R. 10291, Dec. 22, 1956, as amended by Amdt. 6-2, 22 F.R. 5569, July 16, 1957] § 6.442 Oil lines and fittings.

(a) Oil lines shall be supported to prevent excessive vibration.

(b) Oil lines which are connected to components of the rotorcraft between which relative motion could exist shall incorporate provisions for flexibility.

(c) Flexible hose shall be of an approved type.

(d) Oil lines shall have an inside diameter not less than the inside diameter of the engine inlet or outlet, and shall have no splices between connections.

§ 6.443 Oil drains.

One or more accessible drains shall be provided at the lowest point in the oil