duct the balked landing climb with whatever r. p. m, is possible without exceeding the engine take-off limitations with the low pitch setting determined in accordance with § 3.419 (a). (b) In cases where the airplane is to be operated using either the water injection or dry take-off power ratings of the engines, the low pitch stop setting shall be determined on the basis of whichever rating will result in the lower pitch. This will generally be the “dry” rating. In instances where the airplane is intended to be operated only at the water injection take-off power ratings of the engines, the low pitch stop for the propellers should be determined on that basis. These settings are to be deterImined in the usual manner with the airplane static unless there are unconventional features in the propeller installation requiring this determination by some other means. (c) In cases where dual engines drive a single propeller through free wheeling clutches, the setting of the low pitch stop should be such that the propeller will not overspeed when take-off power is applied to one engine at an airplane speed of V2. [Supp. 10, 16 F. R. 3291, Apr. 14, 1951] § 3.420 Speed and pitch limitations for controllable pitch propellers without constant speed controls. The stops or other means incorporated in the propeller mechanism to restrict the pitch range shall limit (a) the lowest possible blade pitch to a value which will assure compliance with the provisions of § 3.419 (a), and (b) the highest possible blade pitch to a value not lower than the flattest blade pitch with which compliance with the provisions of § 3.419 (b) can be demonstrated. § 3.421 Variable pitch propellers with constant speed controls. (a) Suitable means shall be provided at the governor to limit the speed of the propeller. Such means shall limit the maximum governed engine speed to a value not exceeding its maximum permissible take-off revolutions per minute. (b) The low pitch blade stop, or other means incorporated in the propeller mechanism to restrict the pitch range, shall limit the speed of the engine to a value not exceeding 103 percent of the maximum permissible take-off revolutions per minute under the following conditions: With the airplane loaded to the maximum weight and most adverse center of gravity position and the propeller in the most adverse pitch position, propeller clearances shall not be less than the following, unless smaller clearances are properly substantiated for the particular design involved: (a) Ground clearance. (1) Seven inches (for airplanes equipped with nose wheel type landing gears) or 9 inches (for airplanes equipped with tail wheel type landing gears) with the landing gear statically deflected and the airplane in the level, normal take-off, or taxying attitude, whichever is most critical. (2) In addition to subparagraph (1) of this paragraph, there shall be positive clearance between the propeller and the ground when, with the airplane in the level take-off attitude, the critical tire is completely deflated and the corresponding landing gear strut is completely bottomed. (b) Water clearance. A minimum clearance of 18 inches shall be provided unless compliance with §3.147 can be demonstrated with lesser clearance. (c) Structural clearance. (1) One inch radial clearance between the blade tips and the airplane structure, or whatever additional radial clearance is necessary to preclude harmful vibration of the propeller or airplane. (2) One-half inch longitudinal clearance between the propeller blades or cuffs and stationary portions of the airplane. Adequate positive clearance shall be provided between other rotating portions of the propeller or spinner and stationary portions of the airplane. § 3.422-1 Propeller clearance on tricycle gear airplanes (FAA interpretations which apply to § 3.422 (a) (1)). In determining minimum propeller clearance for aircraft equipped with tricycle gear, dynamic effects need not be considered. [Supp. 10, 16 F. R. 3291, Apr. 14, 1951] (a) The fuel systems of multiengine airplanes shall be arranged to permit operation in at least one configuration in such manner that the failure of any one component except the fuel tanks will not result in the loss of power of more than one engine and will not require immediate action by the pilot to prevent the loss of power of more than one engine. Unless other provisions are made to comply with this requirement, the fuel system shall be arranged to permit supplying fuel to each engine through a system entirely independent of any portion of the system supplying fuel to the other engines. (b) If multiengine aircraft employ a single fuel tank or series of fuel tanks interconnected to function as a single fuel tank, the following provisions shall apply: (1) Independent tank outlets to each engine. Each outlet shall incorporate a shutoff valve at the tank. This valve may also serve as the fire wall shutoff valve required by $3.551 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 two vents arranged to minimize the possibility of both vents becoming obstructed simultaneously. (3) Filler cap(s) designed to minimize the possibility of incorrect installation or loss in flight. (4) The remainder of 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. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7066, Sept. 1, 1959] § 3.431-1 Multiengine single tank fuel system (FAA policies which apply to § 3.431). If the shutoff valve also serves as a firewall shutoff valve, the line between the valve and the engine compartment should not contain more than 1 quart of fuel which can drain into the engine compartment. (Secs. 313(a), 601, 603, 72 Stat. 752, 775, 776: 49 U.S.C. 1354 (a), 1421, 1423) [Supp. 35, 24 F.R. 7067, Sept. 1, 1959] § 3.432 Pressure cross feed arrange ments. of unusable fuel in the tank. During this test fuel shall be delivered to the engine at the applicable flow rate (see $$ 3.434-3.436) and at a pressure not less than the minimum required for proper carburetor operation. A suitable mock-up of the system, in which the most adverse conditions are simulated, may be used for this purpose. The quantity of fuel in the tank being tested shall not exceed the amount established as the unusable fuel supply for that tank as determined by demonstration of compliance with the provisions of § 3.437 (see also §§ 3.440 and 3.672), plus whatever minimum quantity of fuel it may be necessary to add for the purpose of conducting the flow test. If a fuel flowmeter is provided, the meter shall be blocked during the flow test and the fuel shall flow through the meter bypass. § 3.434 Fuel flow rate for gravity sys tems. The fuel flow rate for gravity systems (main and reserve supply) shall be 150 percent of the actual take-off fuel con-sumption of the engine. §3.435 Fuel flow rate for pump sys tems. The fuel flow rate for pump systems (main and reserve supply) shall be 0.9 pound per hour for each take-off horsepower or 125 percent of the actual take-off fuel consumption of the engine. whichever is greater. This flow rate shall be applicable to both the primary engine-driven pump and the emergency pumps and shall be available when the pump is running at the speed at which it would normally be operating during take-off. In the case of hand-operated pumps, this speed shall be considered to be not more than 60 complete cycles (120 single strokes) per minute. §3.436 Fuel flow rate for auxiliary fuel systems and fuel transfer systems. The provisions of § 3.434 or § 3.435, whichever is applicable, shall also apply to auxiliary and transfer systems with the exception that the required fuel flow rate shall be established upon the basis of maximum continuous power and speed instead of take-off power and speed. A lesser flow rate shall be acceptable. however, in the case of a small auxiliary tank feeding into a large main tank. provided a suitable placard is installed to require that the auxiliary tank must only be opened to the main tank when a predetermined satisfactory amount of fuel still remains in the main tank. § 3.437 Determination of unusable fuel supply and fuel system operation on low fuel. (a) The unusable fuel supply for each tank shall be established as not less than the quantity at which the first evidence of malfunctioning occurs under the conditions specified in this section. (See also § 3.440.) In the case of airplanes equipped with more than one fuel tank, any tank which is not required to feed the engine in all of the conditions specified in this section need be investigated only for those flight conditions in which it shall be used and the unusable fuel supply for the particular tank in question shall then be based on the most critical of those conditions which are found to be applicable. In all such cases, information regarding the conditions under which the full amount of usable fuel in the tank can safely be used shall be made available to the operating personnel by means of a suitable placard or instructions in the Airplane Flight Manual. (b) Upon presentation of the airplane for test, the applicant shall stipulate the quantity of fuel with which he chooses to demonstrate compliance with this provision and shall also indicate which of the following conditions is most critical from the standpoint of establishing the unusable fuel supply. He shall also indicate the order in which the other conditions are critical from this standpoint: (1) Level flight at maximum continuous power or the power required for level flight at Vc, whichever is less. (2) Climb at maximum continuous power at the calculated best angle of climb at minimum weight. (3) Rapid application of power and subsequent transition to best rate of climb following a power-off glide at 1.3 Vo (4) Sideslips and skids in level flight, climb, and glide under the conditions specified in subparagraphs (1), (2), and (3) of this paragraph, of the greatest severity likely to be encountered in normal service or in turbulent air. (c) In the case of utility category airplanes, there shall be no evidence of malfunctioning during the execution of all approved maneuvers included in the Airplane Flight Manual. During this test the quantity of fuel in each tank shall not exceed the quantity established as the unusable fuel supply, in accordance with paragraph (b) of this section, plus 0.03 gallon for each maximum continuous horsepower for which the airplane is certificated. (d) In the case of acrobatic category airplanes, there shall be no evidence of malfunctioning during the execution of all approved maneuvers included in the Airplane Flight Manual. During this test the quantity of fuel in each tank shall not exceed that specified in paragraph (c) of this section. (e) If an engine can be supplied with fuel from more than one tank, it shall be possible to regain the full power and fuel pressure of that engine in not more than 10 seconds (for single-engine airplanes) or 20 seconds (for multiengine airplanes) after switching to any full tank after engine malfunctioning becomes apparent due to the depletion of the fuel supply in any tank from which the engine can be fed. Compliance with this provision shall be demonstrated in level flight. (f) There shall be no evidence of malfunctioning during take-off and climb for 1 minute at the calculated attitude of best angle of climb at take-off power and minimum weight. At the beginning of this test the quantity of fuel in each tank shall not exceed that specified in paragraph (c) of this section. § 3.438 Fuel system hot weather operation. Airplanes 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. § 3.439 Flow between interconnected tanks. In the case of gravity feed systems with tanks whose outlets are interconnected, 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 airplane is operated as specified in § 3.437 (a) and the tanks are full. inertia, and fluid and structural loads to which they may be subjected in operation. Flexible fuel tank liners shall be of an acceptable type. Integral type fuel tanks shall be provided with adequate facilities for the inspection and repair of the tank interior. The total usable capacity of the fuel tanks shall be sufficient for not less than one-half hour operation at rated maximum continuous power (see § 3.74(d)). The unusable capacity shall be considered to be the minimum quantity of fuel which will permit compliance with the provisions of § 3.437. The fuel quantity indicator shall be adjusted to account for the unusable fuel supply as specified in § 3.672. If the unusable fuel supply in any tank exceeds 5 percent of the tank capacity or 1 gallon, whichever is greater, a placard and a suitable notation in the Airplane Flight Manual shall be provided to indicate to the flight personnel that the fuel remaining in the tank when the quantity indicator reads zero cannot be used safely in flight. The weight of the unusable fuel supply shall be included in the empty weight of the airplane. § 3.441 Fuel tank tests. (a) Fuel tanks shall be capable of withstanding the following pressure tests without failure or leakage. These pressures may be applied in a manner simulating the actual pressure distribution in service: (1) Conventional metal tanks and nonmetallic tanks whose walls are not supported by the airplane structure: A pressure of 3.5 p. s. i. or the pressure developed during the maximum ultimate acceleration of the airplane with a full tank, whichever is greater. (2) Integral tanks: The pressure developed during the maximum limit acceleration of the airplane with a full tank, simultaneously with the application of the critical limit structural loads. (3) Nonmetallic tanks the walls of which are supported by the airplane structure: Tanks constructed of an acceptable basic tank material and type of construction and with actual or simulated support conditions shall be subjected to a pressure of 2 p. s. i. for the first tank of a specific design. The supporting structure shall be designed for the critical loads occurring in the flight or landing strength conditions combined with the fuel pressure loads resulting from the corresponding accelerations. (b) (1) Tanks with large unsupported or unstiffened flat areas shall be capable s of withstanding the following tests without leakage or failure. The complete tank assembly, together with its supports, shall be subjected to a vibration test when mounted in a manner simulating the actual installation. The tank assembly shall be vibrated for 25 hours at a total amplitude of not less than 2 of an inch while filled 3 full of water. 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 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. (2) In conjunction with the vibration test, the tank assembly shall be rocked through an angle of 15° on either side of E the horizontal (30° total) about an axis - parallel to the axis of the fuselage. The assembly shall be rocked at the rate of 16 to 20 complete cycles per minute. (c) Integral tanks which incorporate methods of construction and sealing not previously substantiated by satisfactory test data or service experience shall be capable of withstanding the vibration test specified in paragraph (b) of this section. (d) (1) Tanks with nonmetallic liners shall be subjected to the sloshing portion of the test outlined under paragraph (b) of this section with fuel at room temperature. (2) In addition, a specimen liner of the same basic construction as that to be used in the airplane shall, when installed in a suitable test tank, satisfactorily withstand the slosh test with fuel at a temperature of 110° F. § 3.442 Fuel tank installation. (a) The method of supporting tanks shall not be such as to concentrate the loads resulting from the weight of the fuel in the tanks. Pads shall be provided to prevent chafing between the tank and its supports. Materials employed for padding shall be nonabsorbent or shall be treated to prevent the absorption of fuels. If flexible tank liners are employed, they = shall be of an approved type, and they shall be so supported that the liner is not required to withstand fluid loads. Interior surfaces of compartments for such liners shall be smooth and free of pro jections which are apt to cause wear of the liner, unless provisions are made for the protection of the liner at such points cr unless the construction of the liner itself provides such protection. A positive pressure shall be maintained within the vapor space of all bladder cells under all conditions of operation including the critical condition of low air speed and rate of descent likely to be encountered in normal operation. (b) Tank compartments shall be ventilated and drained to prevent the accumulation of inflammable fluids or vapors. Compartments adjacent to tanks which are an integral part of the airplane structure shall also be ventilated and drained. (c) Fuel tanks shall not be located on the engine side of the fire wall. Not less than one-half inch of clear air space shall be provided between the fuel tank and the fire wall. No portion of engine nacelle skin which lies immediately behind a major air egress opening from the engine compartment shall act as the wall of an integral tank. Fuel tanks shall not be located in personnel compartments, except in the case of singleengine airplanes. In such cases fuel tanks the capacity of which does not exceed 25 gallons may be located in personnel compartments, if adequate ventilation and drainage are provided. In all other cases, fuel tanks shall be isolated from personnel compartments by means of fume and fuel proof enclosures. § 3.442-1 Bladder type fuel cells located in a personnel compartment (FAA interpretations which apply to § 3.442). In the case where a bladder type fuel cell having a fuel capacity in excess of 25 gallons is located in a personnel compartment, a separate fume and fuel proof enclosure for the fuel cell and its retaining shell is not deemed necessary provided the retaining shell is at least equivalent to a conventional metal fuel tank in structural integrity and fume and fuel tightness. The shell surrounding the tank should be adequately drained to the exterior of the airplane. [Supp. 10, 16 F. R. 3291, Apr. 14, 1951] Fuel tanks shall be provided with an expansion space of not less than 2 percent of the tank capacity, unless the tank vent discharges clear of the aircraft in which case no expansion space will be |