required. It shall not be possible inadvertently to fill the fuel tank expansion space when the airplane is in the normal ground attitude.

§ 3.444 Fuel tank sump.

(a) Each tank shall be provided with a drainable sump having a capacity of not less than 0.25 percent of the tank capacity or 16 gallon, whichever is the greater. It shall be acceptable to dispense with the sump if the fuel system is provided with a sediment bowl permitting ground inspection. The sediment bowl shall also be accessible for drainage. The capacity of the sediment chamber shall not be less than 1 ounce per each 20 gallons of the fuel tank capacity.

(b) If a fuel tank sump is provided, the capacity specified in paragraph (a) of this section shall be effective with the airplane in the normal ground attitude and in all normal flight attitudes.

(c) If a separate sediment bowl is provided in lieu of a tank sump, the fuel tank outlet shall be so located that, when the airplane is in the normal ground attitude, water will drain from all portions of the tank to the sediment bowl. § 3.445 Fuel tank filler connection.

(a) Fuel tank filler connections shall be marked as specified in § 3.767.

(b) Provision shall be made to prevent the entrance of spilled fuel into the fuel tank compartment or any portions of the airplane other than the tank itself. The filler cap shall provide a fuel-tight seal for the main filler opening. However, small openings in the fuel tank cap for venting purposes or to permit passage of a fuel gauge through the cap shall be permissible.

§ 3.446 Fuel tank vents and carburetor vapor vents.

(a) Fuel tanks shall be vented from the top portion of the expansion space. Vent outlets shall be so located and constructed as to minimize the possibility of their being obstructed by ice or other foreign matter. The vent shall be so constructed as to preclude the possibility of siphoning fuel during normal operation. The vent shall be of sufficient size to permit the rapid relief of excessive differences of pressure between the interior and exterior of the tank. Air spaces of tanks the outlets of which are interconnected shall also be

interconnected. There shall be no undrainable points in the vent line where moisture is apt to accumulate with the airplane in either the ground or level flight attitude. Vents shall not terminate at points where the discharge of fuel from the vent outlet will constitute a fire hazard or from which fumes may enter personnel compartments.

(b) Carburetors which are provided with vapor elimination connections shall be provided with a vent line which will lead vapors back to one of the airplane fuel tanks. If more than one fuel tank is provided and it is necessary to use these tanks in a definite sequence for any reason, the vapor vent return line shall lead back to the fuel tank which must be used first unless the relative capacities of the tanks are such that return to another tank is preferable.

§3.447-A Fuel tank vents.

Provision shall be made to prevent excessive loss of fuel during acrobatic maneuvers including short periods of inverted flight. It shall not be possible for fuel to siphon from the vent when normal flight has been resumed after having executed any acrobatic maneuver for which the airplane is intended. § 3.448 Fuel tank outlet.

The fuel tank outlet shall be provided with a screen of from 8 to 16 meshes per inch. If a finger strainer is used, the length of the strainer shall not be less than 4 times the outlet diameter. The diameter of the strainer shall not be less than the diameter of the fuel tank outlet. Finger strainers shall be accessible for inspection and cleaning.

FUEL PUMPS

§ 3.449 Fuel pump and pump installation.

(a) If fuel pumps are provided to maintain a supply of fuel to the engine. at least one pump for each engine shall be directly driven by the engine. Fuel pumps shall be adequate to meet the flow requirements of the applicable portions of §§ 3.433-3.436.

(b) Emergency fuel pumps shall be provided to permit supplying all engines with fuel in case of the failure of any one engine-driven pump, except that if an engine fuel injection pump which has been certificated as an integral part of the engine is used, an emergency pump is not required. Emergency pumps shall

be available for immediate use in case of the failure of any other pump. If both the normal pump and emergency pump operate continuously, means shall be provided to indicate to the crew when either pump is malfunctioning.

§ 3.449-1 Fuel injection pump (FAA interpretations which apply to

§ 3.449 (b)).

The phrase "fuel injection pump" means a pump that supplies proper flow and pressure conditions for fuel injection 1 when such injection is not accomplished in a carburetor.

[Supp. 32, 23 F. R. 7481, Sept. 26, 1958]

LINES, FITTINGS, AND ACCESSORIES

§ 3.550 Fuel system lines and fittings.

(See § 3.638.) (a) Fuel lines shall be t 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 airplane between which relative motion could exist shall incorporate provisions for flexibility.

(c) Provisions for flexibility in fuel lines which may be under pressure and subjected to axial loading shall employ flexible hose assemblies rather than hose-clamp connections.

(d) Flexible hose shall be of an approved type or shall be shown to be suittable for the particular application.

(e) Flexible hoses which might be adversely affected by exposure to high temperatures shall not be employed in 3 locations where excessive temperatures # will exist during operation or after engine shutdown.

[Amdt. 3-2, 22 F. R. 5562, July 16, 1957] § 3.551 Fuel valves.

(a) Means shall be provided to permit the flight personnel to shut off rapidly the flow of fuel to any engine individually in flight. Valves provided for this

1 Fuel injection is a special form of carburetion: the charging of air or gas with volatile carbon compounds. It is either an intermittent charging of air by discrete, metered quantities of fuel such as occurs in a Diesel cylinder or it is a continuous charging of air by fuel, the fuel flow being proportioned to the airflow through the engine. Examples of continuous injection are injections into the supercharger section of a reciprocating engine or into the combustion chambers of a turbine engine.

purpose shall be located on the side of the fire wall most remote from the engine.

(b) Means shall be provided to guard against inadvertent operation of the shutoff valves and to make it possible for the flight personnel to reopen the valves rapidly after they have been closed.

(c) Valves shall be provided with either positive stops or "feel" in the on and off positions and shall be supported in such a manner that loads resulting from their operation or from accelerated flight conditions are not transmitted to the lines connected to the valve. Valves shall be so installed that the effect of gravity and vibration will tend to turn their handles to the open rather than the closed position.

(d) Fuel valve handles and their connections to the valve mechanism shall incorporate design features to minimize the possibility of incorrect installation. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-5, 24 F.R. 7067, Sept. 1, 1959] § 3.552 Fuel strainer.

A fuel strainer shall be provided between the fuel tank outlet and the carburetor inlet. If an engine-driven fuel pump is provided, the strainer shall be located between the tank outlet and the engine-driven pump inlet. The strainer shall be accessible for drainage and cleaning, and the strainer screen shall be removable.

DRAINS AND INSTRUMENTS

§ 3.553 Fuel system drains.

Drains shall be provided to permit safe drainage of the entire fuel system and shall incorporate means for locking in the closed position. The provisions for drainage shall be effective in the normal ground attitude.

critical operating conditions and the maximum oil consumption of the engine under the same conditions, plus a suitable margin to assure adequate system circulation and cooling.

§ 3.561-1 "Capacity" (FAA interpretations which apply to § 3.561).

The word "capacity" as used in § 3.561 is interpreted by the Administrator as follows:

(a) Only the usable fuel system capacity need be considered.

(b) In a conventional oil system (no transfer system provided) only the usable oil tank capacity shall be considered. The quantity of oil in the engine oil lines, the oil radiator, or in the feathering reserve shall not be included. When an oil transfer system is installed, and the transfer pump is so located that it can pump some of the oil in the transfer lines into the main engine oil tanks, the quantity of oil in these lines which can be pumped by the transfer pump may be added to the oil capacity.

[Supp. 1, 12 F. R. 3438, May 28, 1947, as amended by Amdt. 1, 14 F. R. 36, Jan. 5, 1949]

§ 3.562 Oil cooling.

See 3.581 and pertinent sections.

Oil tanks shall be capable of withstanding without failure all vibration, inertia, and fluid loads to which they might be subjected in operation. Flexible oil tank liners shall be of an acceptable type.

§ 3.564 Oil tank tests.

Oil tank tests shall be the same as fuel tank tests (see § 3.441), except as follows:

(a) The applied pressure shall be 5 p. s. i. for all tank constructions instead of those specified in § 3.441 (a).

(b) In the case of tanks with nonmetallic liners, the test fluid shall be oil rather than fuel as specified in § 3.441 (d) and the slosh test on a specimen liner shall be conducted with oil at a temperature of 250° F.

[21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-2, 22 F.R. 5562, July 16, 1957] § 3.565 Oil tank installation.

Oil tank installations shall comply with the requirements of § 3.442 (a) and (b).

§ 3.566 Oil tank expansion space.

Oil tanks shall be provided with an expansion space of not less than 10 percent of the tank capacity or 1/2 gallon, whichever is greater. It shall not be possible inadvertently to fill the oil tank expansion space when the airplane is in the normal ground attitude.

§ 3.567 Oil tank filler connection.

Oil tank filler connections shall be marked as specified in § 3.767.

§ 3.568 Oil tank vent.

(a) Oil tanks shall be vented to the engine crankcase from the top of the expansion space in such a manner that the vent connection is not covered by oil under any normal flight conditions. Oil tank vents shall be so arranged that condensed water vapor which might freeze and obstruct the line cannot accumulate at any point.

(b) Category A: Provision shall be made to prevent hazardous loss of oil during acrobatic maneuvers including short periods of inverted flight.

§ 3.569 Oil tank outlet.

The oil tank outlet shall not be enclosed or covered by any screen or other guard which might impede the flow of oil. The diameter of the oil tank outlet shall not be less than the diameter of the engine oil pump inlet. (See also § 3.577.)

LINES, FITTINGS, AND ACCESSORIES

§ 3.570 Oil system lines, fittings, and accessories.

Oil lines shall comply with the provisions of § 3.550, except that the inside diameter of the engine oil inlet and outlet lines shall not be less than the diameter of the corresponding engine oil pump inlet and outlet.

filter element will not jeopardize the continued operation of the engine oil supply system.

§ 3.574 Oil system drains.

Drains shall be provided to permit safe drainage of the entire oil system and shall incorporate means for positive locking in the closed position.

§ 3.575 Engine breather lines.

(a) Engine breather lines shall be so arranged that condensed water vapor which might freeze and obstruct the line cannot accumulate at any point. Breathers shall discharge in a location I which will not constitute a fire hazard in case foaming occurs and so that oil emitted from the line will not impinge the pilot's windshield. upon The breather shall not discharge into the engine air induction system.

(b) Category A: In the case of acrobatic type airplanes, provision shall be made to prevent excessive loss of oil from the breather during acrobatic maneuvers including short periods of inverted flight.

§ 3.576 Oil system instruments.

See § 3.655, 3.670, 3.671, and 3.674. 3.577

Propeller feathering system.

If the propeller feathering system is dependent upon the use of the engine oil upply, provision shall be made to trap a [uantity of oil in the tank in case the upply becomes depleted due to failure of ny portion of the lubricating system ther than the tank itself. The quantity f oil so trapped shall be sufficient to acomplish the feathering operation and hall be available only to the feathering ump. The ability of the system to acɔmplish feathering when the supply of il has fallen to the above level shall be emonstrated.

critical ground, water, and flight operating conditions. If the tests are conducted under conditions which deviate from the highest anticipated summer air temperature (see § 3.583), the recorded powerplant temperatures shall be corrected in accordance with the provisions of $3.584 and 3.585. The corrected temperatures determined in this manner shall not exceed the maximum established safe values. The fuel used during the cooling tests shall be of the minimum octane number approved for the engines involved, and the mixture settings shall be those appropriate to the operating conditions. The test procedures shall be as outlined in §§ 3.586 and 3.587.

§ 3.582-1 Water taxiing tests (FAA interpretations which apply to

§ 3.582).

No water taxiing tests need be conducted on aircraft certificated under this part, except in the case of flying boats which may reasonably be expected to be taxied for extended periods.

[Supp. 10. 16 F. R. 3291, Apr. 14, 1951] § 3.583

Maximum anticipated summer air temperatures.

The maximum anticipated summer air temperature shall be considered to be 100° F. at sea level and to decrease from this value at the rate of 3.6° F. per thousand feet of altitude above sea level. § 3.583-1 Powerplant winterization equipment (FAA interpretations which apply to § 3.583).

(a) Cooling test results for winterization installations may be corrected to any temperature desired by the manufacturer rather than the conventional 100° F. hot day. For example, if a manufacturer chooses to demonstrate cooling to comply with requirements for a 50° or 60° F. day with winterization equipment installed, he may do so. In such a case the sea level temperature for correction purposes should be considered to be the value elected by the manufacturer with a rate of temperature drop of 3.6° F. per thousand feet above sea level.

(b) Cooling tests and temperature correction methods should be the same as for conventional cooling tests.

(c) The airplane flight manual should clearly indicate that winterization equipment must be removed whenever the temperature reaches the limit for which adequate cooling has been demonstrated.

The cockpit should also be placarded accordingly. In addition, the airplane should be equipped with an ambient air temperature gauge or, alternatively, a cylinder head, barrel, or oil inlet temperature gauge (depending upon which is critical).

(d) If practical, winterization equipment such as baffles for oil radiators or for engine cooling air openings should be marked clearly to indicate the limiting temperature at which this equipment should be removed.

(e) Since winterization equipment is often supplied in kit form, accompanied by instructions for its installation, suitable information regarding temperature limitations should be included in the installation instructions for such kits. [Supp. 10, 16 F. R. 3291, Apr. 14, 1951] § 3.584 Correction factor for cylinder head, oil inlet, carburetor air, and engine coolant inlet temperatures. These temperatures shall be corrected by adding the difference between the maximum anticipated summer air temperature and the temperature of the ambient air at the time of the first occurrence of maximum head, air, oil, or coolant temperature recorded during the cooling test.

§ 3.586

Cooling test procedure for single-engine airplanes.

The engine cooling tests shall be conducted by stabilizing the engine temperature in flight with the engines operating at not less than 75 percent of the maximum continuous power rating. After engine temperatures have stabilized, the climb shall be started at the lowest practicable altitude and continued for one minute with the engines operating at the take-off rating. At the end of 1 minute, the climb shall be continued at maximum continuous power until at least 5 minutes after the occurrence of the highest temperature recorded. The climb shall not be con

ducted at a speed greater than the best rate-of-climb speed with maximum continuous power unless:

(a) The slope of the flight path at the speed chosen for the cooling test is equal to or greater than the minimum required angle of climb (see § 3.85 (a)), and

(b) A cylinder head temperature indicator is provided as specified in § 3.675. [21 F.R. 3339, May 22, 1956, as amended by Amdt. 3-2, 22 F.R. 5562, July 16, 1957] § 3.587 Cooling test procedure for multiengine airplanes.

(a) Airplanes which meet the minimum one-engine-inoperative climb performance specified in § 3.85(b). The engine cooling test for these airplanes shall be conducted with the airplane in the configuration specified in § 3.85 (b). except that the operating engine (s) shall be operated at maximum continuous power or at full throttle when above the critical altitude. Temperatures of the operating engines shall be stabilized in flight with the engines operating at no less than 75 percent of the maximum continuous power rating. After stabilizing temperatures in flight, the climb shall be started at the lower of the two following altitudes and shall be continued until at least 5 minutes after the highest temperature has been recorded:

(1) 1,000 feet below the engine critical altitude or at the lowest practicable altitude (when applicable).

(2) 1,000 feet below the altitude at which the single-engine-inoperative rate of climb is 0.02 Vso2.

The climb shall be conducted at a speed not in excess of the highest speed at which compliance with the climb requirement of § 3.85 (b) can be shown. However, if the speed used exceeds the speed for best rate of climb with one engine inoperative, a cylinder head temperature indicator shall be provided as specified in § 3.675.

(b) Airplanes which cannot meet the minimum one-engine-inoperative climb performance specified in § 3.85 (b). The engine cooling test for these airplanes shall be the same as in paragraph (a) of this section, except that after stabilizing temperatures in flight, the climb (or descent, in the case of airplanes with zero or negative one-engineinoperative rate of climb) shall be commenced at as near sea level as practicable and shall be conducted at the best rate