be installed. The endurance tests shall be conducted in 10-hour test cycles composed of the tests prescribed in subparagraphs (2) through (10) of this paragraph. Compliance with the endurance tests prescribed in this paragraph will be accepted for helicopter engine certification in lieu of the endurance testing specified in Part 13 of this subchapter. (The other phases of helicopter engine certification such as vibration, calibration, detonation, operation, and engine inspection will of course require compliance in accordance with Part 13 of this subchapter.)

(2) Take-off power run. The take-off power run shall consist of one hour of alternate runs of 5 minutes at take-off power and speed, and 5 minutes at as low an engine idle speed as practicable. The engine shall be declutched from the rotor drive system and the rotor brake, if furnished and so intended, shall be applied during the first minute of the idle run. During the remaining 4 minutes of the idle run, the clutch shall be engaged so that the engine drives the rotors at the minimum practical rpm. Acceleration of the engine and the rotor drive system shall be accomplished at the maximum rate. When declutching the engine, it shall be decelerated at a rate sufficiently rapid to permit the operation of the overrunning clutch. In the absence of a take-off rating, maximum continuous power and speed shall be substituted for take-off power and speed.

(3) Maximum continuous run. Three hours of continuous operation at maximum continuous power and speed as follows:

(i) During the run, the main rotor controls shall be operated at a minimum of 15 times each hour through the main rotor pitch positions of full vertical thrust, maximum forward thrust component, maximum aft thrust component, maximum left thrust component, and maximum right thrust component, except that the control movements need not produce loads or blade flapping motion exceeding the maximum loads or motions encountered in flight.

(ii) The directional controls shall be operated at a minimum of 15 times each hour through the control extremes of maximum right turning torque, neutral torque as required by the power applied to the main rotor, and maximum left turning torque.

(iii) Each control position shall be held at maximum for at least 10 seconds and the rate of change of control position shall be at least as rapid as for normal operation.

(4) 90 percent maximum continuous run. One hour of continuous operation at 90 percent maximum continuous power at maximum continuous speed.

(5) 80 percent maximum continuous run. One hour of continuous operation at 80 percent maximum continuous power and speed.

(6) 60 percent maximum continuous run. Two hours of continuous operation at 60 percent maximum continuous power at minimum desired cruising speed or at 90 percent maximum continuous speed, whichever speed is lower.

It

(7) Engine malfunctioning run. shall be determined whether malfunctioning of such components as the engine fuel or ignition systems or unequal power output from the various engines can result in dynamic conditions which might be detrimental to the drive system. If so, a suitable number of hours of operation shall be accomplished under such conditions, one hour of which shall be included in each cycle, and the remaining hours accomplished at the conclusion of the 20 cycles. If no detrimental condition results, an additional hour of operation as prescribed in subparagraph (2) of this paragraph shall be substituted.

(8) Overspeed run. One hour of continuous operation at 110 percent maximum continuous speed at maximum continuous power. In the event that the engine(s) installed are limited by the manufacturer to an overspeed of less than 110 percent of maximum continuous speed for the periods required, the speed employed shall be the highest speed permissible with the engine(s) involved.

(9) Rotor control positions. Whenever the rotor controls are not being cycled during the tie-down tests, the rotor shall be operated to produce each of the maximum thrust positions for the percentages of test time as follows, except that the control positions need not produce loads or blade flapping motion exceeding the maximum loads or motions encountered in flight, using the procedures of subparagraph (3) of this paragraph:

(i) Full vertical thrust, 20 percent.

(ii) Forward thrust component, 50 percent.

(iii) Right thrust component, 10 percent.

(iv) Left thrust component, 10 percent.

en

(v) Aft thrust component, 10 percent. (10) Clutch and brake engagements. A total of at least 400 clutch and brake engagements including the engagements of paragraph (a) (2) of this section shall be made during the take-off power runs and, as necessary, at each change of power and speed throughout the test. In each clutch engagement, the shaft on the driven side of the clutch shall be accelerated from rest. The clutch gagements shall be accomplished at the speed and by the method prescribed in the operations manual. During deceleration after each clutch engagement, the engine (s) shall be stopped rapidly enough to allow the engine(s) to be automatically disengaged from the rotor(s) and/or rotor-drive(s). If a rotor brake is installed for the purpose of stopping the rotor, the clutch, during brake engagements, shall be disengaged above 40 percent maximum continuous rotor speed and the rotor(s) allowed to decelerate to 40 percent maximum continuous rotor speed at which time the rotor brake shall be applied. If the clutch design does not permit stopping the rotor(s) with the engine running, or if no clutch is provided, the engine shall be stopped before each application of the rotor brake, and then immediately restarted after the rotors have stopped.

(b) Overspeed test. After completion of the 200-hour tie-down test and without intervening major disassembly, the rotor drive system shall be subjected to 50 overspeed runs, each 30±3 seconds in duration at 120 percent maximum continuous speed. Overspeed runs shall be alternated with stabilizing runs of 1 to 5 minutes duration each at from 60 to 80 percent maximum continuous speed. Acceleration and deceleration shall be accomplished in a period not longer than 10 seconds, and the time for changing speeds shall not be deducted from the specified time for the overspeed runs. Overspeed runs should be made with the rotor(s) in the flattest pitch at which smooth operation can be obtained. In the event that the engine (s) installed is limited by the engine manufacturer to an overspeed of less than 120 percent of maximum continuous speed for the

periods required, the speed employed shall be the highest speed permissible with the engine(s) involved.

(c) Critical component reliability tests. Components within the rotor drive system, the failure of which will result in an uncontrolled landing, components essential to the phasing of the rotors on multirotor rotorcraft, or as a driving link for essential control of rotors in autorotation, and components ! common to more than one engine on multiengine rotorcraft, shall be designed to have a level of safety equivalent to the main rotors. Components which are affected by flight maneuvering and gust loads shall be additionally investigated for the same flight conditions as the main rotor(s). The service life of such parts shall be determined by fatigue tests or by other methods found acceptable by the Administrator.

(d) Special tests. Rotor drive systems designed to operate at two or more gear ratios shall be subjected to special testing and durations found necessary by the Administrator to substantiate the airworthiness of the rotor drive system.

(e) Category A; gear box bench tests. Each gear box employed in the rotor drive system shall be tested for 150 hours at 110 percent of its maximum continuous power and speed. The components employed in this test need not be the same as those employed in the tests of paragraph (a) of this section.

§ 7.406 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.

§ 7.407 Critical shafting speeds.

An investigation shall be made to determine that the critical speeds of all shafting lie outside the range of permissible engine speeds under idling, poweron, and autorotative conditions. If critical vibration conditions (persistent or momentary) are found in the entire | range of operations from and including clutch engagement to maximum overspeed, either during acceleration or deceleration, it shall be demonstrated in the rotorcraft that such vibration is within safe limits. Such demonstration may be made during the endurance testing (see 7.405(a)), in which case the i

1

(a) The fuel system shall be constructed and arranged in such a manner as to assure a flow of fuel to each engine at a rate and pressure which have been established for proper engine functioning under all normal conditions, including all maneuvers for which the rotorcraft is intended. (For fuel system instruments see § 7.604.)

(b) The fuel system shall be so ar=ranged that no one engine or fuel pump can draw fuel from more than one tank at a time unless means are provided to prevent introducing air into the system. § 7.411 Fuel system independence.

(a) Category A. The design of the fuel system shall comply with the requirements of § 7.401 (b). Unless other provisions are made in compliance with this requirement, the fuel system shall be arranged to permit the supply of fuel to each engine through a system independent of any portion of a system supplying fuel to any other engine.

(b) Category B. The design of the fuel system for multiengine rotorcraft shall be arranged to permit supplying fuel to each engine through a system independent of all portions of systems supplying fuel to the other engines, except that separate fuel tanks need not be provided for each engine.

§ 7.412 Fuel lines in personnel and

cargo compartments.

(a) Fuel lines shall not pass through portions of the rotorcraft intended to carry personnel or cargo, unless they are so located or protected by drained and ventilated shrouds or other means which will assure that in case of leakage, fuel and fumes will be carried safely overboard. Means shall be provided to permit the flight personnel to shut off the supply of fuel to such lines without af

(a) The fuel supply system shall be arranged so that, insofar as practicable, the entire fuel supply can be utilized in the maximum inclinations of the rotorcraft 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, the system shall feed promptly when the fuel supply becomes low in one tank and another tank is turned on (see § 7.438).

(b) The ability of the fuel system to provide the required fuel flow rate shall be demonstrated when the rotorcraft is in the attitude which represents the most adverse sustained condition, from the standpoint of fuel feed, which the rotorcraft is designed to attain. The demonstration may be accomplished by a ground test of the rotorcraft or on a representative operating mock-up of the fuel system. The following conditions are applicable to such demonstration:

(1) Category A: The critical attitude (or attitudes) selected shall be verified by a flight demonstration which takes into consideration all operating speeds, power settings, accelerated maneuvers, and engine inoperative conditions.

(2) The quantity of fuel in the tank being considered shall not exceed the amount established as the unusable fuel supply for that tank, as determined by demonstrating compliance with the provisions of § 7.416 (see also §§ 7.418 and 7.613 (b)), together with whatever minimum quantity of fuel it may be necessary to add for the purpose of conducting the flow test.

(3) The fuel shall be delivered to the engine at a pressure not less than the minimum inlet pressure established for proper engine operation in accordance with Part 13 of this subchapter.

(4) If a fuel flowmeter is provided, the meter shall be blocked during the flow test and the fuel shall flow through the meter by-pass.

(c) The fuel flow rate required for the demonstration specified in paragraph (b) of this section shall be as follows:

(1) For gravity feed systems: The fuel flow rate shall be 150 percent of the actual take-off power fuel consumption of the engine.

(2) For pump systems: The fuel flow rate shall be 0.9 pound per hour for each rated take-off horsepower or 125 percent of the actual take-off fuel consumption of the engine, whichever is greater.

§ 7.414 Pump systems.

(a) The fuel flow rate specified in § 7.413 (c) shall be applicable to both the primary engine-driven pump and to emergency pumps. The fuel flow rate shall be available when the pump is running at the speed at which it normally would be operating during take-off. In the case of hand-operated pumps, the speed required shall be not more than 60 complete cycles (120 single strokes) per minute.

(b) Emergency pumps shall be provided to permit supplying all engines with fuel in case of failure of any one main fuel pump, except in the case of installations in which the only fuel pump necessary is an engine fuel injection or fuel metering pump which is approved as an integral part of the engine.

(c) Category A: If the arrangement of the fuel system necessitates a fuel boost pump to maintain operating fuel flow and pressure for the range of altitudes and temperatures in which flight is expected, a duplicate boost pump shall be provided to serve as an emergency pump in case of failure of the main boost pump. Each pump shall be capable of supplying fuel flow at the rate specified in § 7.413 (c).

(d) Category A: Where the provisions of paragraph (c) of this section are applicable, and both boost pumps are dependent upon a common source of power, it shall be possible with these components inoperative to maintain cruising fuel flow and pressure for all engines. The limiting weights, speeds, and altitudes shall be demonstrated and the results recorded in the operating procedures portion of the Rotorcraft Flight Manual.

§ 7.415 Transfer systems.

The provisions of § 7.413 shall apply to transfer systems, except that the re

quired fuel flow rate for the engine or engines involved shall be established upon the basis of maximum continuous power and its corresponding speed.

§ 7.416 Determination of 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 §§ 7.104 and

and 7.613 (b).)

§ 7.417 Fuel system hot weather operation.

(a) The fuel system shall be so arranged as to minimize the possibility of the formation of vapor in the system under all normal conditions of operation. Rotorcraft with suction lift fuel systems or systems which have features likely to produce vapor shall be demonstrated to be free from vapor lock when using fuel at a temperature of 110° F. under critical operating conditions.

(b) Category A: To prove satisfactory hot weather operation the rotorcraft shall be climbed from the altitude of the airport chosen by the applicant to an altitude of 5,000 feet above the terrain, or to the altitude at which the rotorcraft is expected to operate, whichever is greater. There shall be no evidence of vapor lock or other malfunctioning. The climb test shall be conducted under the following conditions:

(1) All engines shall operate at maximum continuous power, except that take-off power shall be used at the beginning of the demonstration for the maximum time interval for which takeoff power is approved for use on the rotorcraft.

(2) The weight shall be with full fuel tanks, minimum crew, and only such ballast as is required to maintain the center of gravity within allowable limits.

(3) The speed of climb shall be the speed for best rate of climb under the conditions of the test.

(4) The fuel temperature shall be not less than 110° F. at the beginning of the demonstration.

(c) Category A: The test prescribed in paragraph (b) of this section shall be performed either in flight or on the

ground closely simulating flight conditions. If a flight test is performed in weather sufficiently cold to interfere with the proper conduct of the test, the fuel tank surfaces, fuel lines, and other fuel system parts subjected to cooling action from cold air shall be insulated to simulate, insofar as practicable, flight in hot weather.

§ 7.418 Flow between interconnected tanks.

(a) Where tank outlets are interconnected and permit flow through the interconnection due to gravity or flight accelerations, it shall not be possible for fuel to flow between tanks in quantities sufficient to cause an overflow of fuel from the tank vent when the rotorcraft is operated in any sustained flight condition.

(b) If it is possible to pump fuel from one tank to another in flight, the design of the fuel tank vents and the fuel transfer system shall be such that structural damage to tanks will not occur in the event of overfilling. In addition, means shall be provided to warn the crew before overflow through the vents occurs.

FUEL TANK CONSTRUCTION AND
INSTALLATION

(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 and their installation shall be designed or protected so as to retain the fuel supply without leakage when the rotorcraft is subjected to the emergency landing conditions specified under § 7.260.

(c) Flexible fuel tank liners shall be of an approved type or shall be shown to be suitable for the particular application.

(d) Integral type fuel tanks shall be provided with facilities for inspection and repair of the tank interior.

§ 7.421 Fuel tank tests.

(a) Fuel tanks shall be capable of withstanding the following pressure tests without failure or leakage. It shall be acceptable to apply the pressures in a manner simulating the actual pressure distribution in service (where this is practicable).

(1) Conventional metal tanks, nonmetallic tanks the walls of which are not

supported by the rotorcraft structure, and integral tanks shall be submitted to a pressure of 3.5 psi unless the pressure developed during the maximum limit acceleration or emergency deceleration (see § 7.260) of the rotorcraft with a full tank exceeds this value, in which case a hydrostatic head, or equivalent test, shall be applied to duplicate the acceleration loads insofar as possible, except that the pressure need not exceed 3.5 psi on surfaces not exposed to the acceleration loading.

(2) Nonmetallic tanks the walls of which are supported by the rotorcraft structure shall be submitted to the following tests:

(i) A pressure test of at least 2.0 psi. The test may be conducted on the tank alone in conjunction with the test specified in subdivision (ii) of this subparagraph.

(ii) A pressure test, with the tank mounted in the rotorcraft structure, equivalent to the load developed by the reaction of the contents, when the tanks are full, during the maximum limit acceleration or emergency deceleration (see § 7.260) of the rotorcraft, except that the pressure need not exceed 2.0 psi on the surfaces not exposed to the acceleration loading.

(b) Tanks with large unsupported or unstiffened flat areas or other design or construction features the failure or deformation of which could cause fuel leakage shall be capable of withstanding the following test, or other equivalent test, without leakage or failure:

(1) The complete tank assembly together with its supports shall be subjected to a vibration test when mounted in a manner simulating the actual

installation.

(2) The tank assembly shall be vibrated for 25 hours while filled twothirds full of water or any suitable fluid. The amplitude of vibration shall not be less than one thirty-second of an inch, unless otherwise substantiated.

(3) The frequency of vibration shall be 90 percent of the maximum continuous rated speed of the engine unless some other frequency within the normal operating range of speeds of the engine or rotor system is more critical, in which case the latter speed shall be employed and the time of test shall be adjusted to accomplish the same number of vibration cycles.