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CAR PART 3-1949 - SUBPART E--POWER-PLANT INSTALLATIONS; RECIPROCATING ENGINES
GENERAL
3.411 Components.
ENGINES AND PROPELLERS
3.415 Engines.
3.416 Propellers.
3.417 Propeller vibration.
3.418 Propeller pitch and speed limitations.
3.419 Speed limitations for fixed pitch propellers, ground adjustable pitch
propellers, and automatically varying pitch propellers which cannot be
controlled in flight.
3.420 Speed and pitch limitations for controllable pitch propellers without
constant speed controls.
3.421 Variable pitch propellers with constant speed controls.
3.422 Propeller clearance.
FUEL SYSTEM
3.429 General.
ARRANGEMENT
3.430 Fuel system arrangement.
3.431 Multiengine fuel system arrangement.
3.432 Pressure cross feed arrangements.
OPERATION
3.433 Fuel flow rate.
3.434 Fuel flow rate for gravity systems.
3.435 Fuel flow rate for pump systems.
3.436 Fuel flow rate for auxiliary fuel systems and fuel transfer systems.
3.437 Determination of unusable fuel supply and fuel system operation on low
fuel.
3.438 Fuel system hot weather operation.
3.439 Flow between interconnected tanks.
FUEL TANKS
3.440 General.
3.441 Fuel tank tests.
3.442 Fuel tank installation.
3.443 Fuel tank expansion space.
3.444 Fuel tank sump.
3.445 Fuel tank filler connection.
3.446 Fuel tank vents and carburetor vapor vents.
3.447-A Fuel tank vents.
3.448 Fuel tank outlet.
FUEL PUMPS
3.449 Fuel pump and pump installation
LINES, FITTINGS, AND ACCESSORIES
3.550 Fuel system lines, fittings, and accessories.
3.551 Fuel valves.
3.552 Fuel Strainer.
DRAINS AND INSTRUMENTS
3.553 Fuel system drains.
3.554 Fuel system instruments.
OIL SYSTEM
3.561 Oil system.
3.562 Oil cooling.
OIL TANKS
3.563 Oil tanks.
3.564 Oil tank tests.
3.565 Oil tank installation.
3.566 Oil tank expansion space.
3.567 Oil tank filler connection.
3.568 Oil tank vent.
3.569 Oil tank outlet.
LINES, FITTINGS, AND ACCESSORIES
3.570 Oil system lines, fittings, and accessories.
3.571 Oil valves.
3.572 Oil radiators.
3.573 Oil filters.
3.574 Oil system drains.
3.575 Engine breather lines.
3.576 Oil system instruments.
3.577 Propeller feathering system.
COOLING
3.581 General.
TESTS
3.582 Cooling tests.
3.583 Maximum anticipated summer air temperatures.
3.584 Correction factor for cylinder head, oil inlet, carburetor air, and
engine coolant inlet temperatures.
3.585 Correction factor for cylinder barrel temperatures.
3.586 Cooling test procedure for single engine airplanes.
3.587 Cooling test procedure for multi engine airplanes.
LIQUID COOLING SYSTEMS
3.588 Independent systems.
3.589 Coolant tank.
3.590 Coolant tank tests.
3.591 Coolant tank installation.
3.592 Coolant tank connection.
3.593 Coolant lines, fittings, and accessories.
3.594 Coolant radiators.
3.595 Cooling system drains.
3.596 Cooling system instruments.
INDUCTION SYSTEM
3.605 General.
3.606 Induction system de-icing and anti-icing provisions.
3.607 Carburetor de-icing fluid flow rate.
3.608 Carburetor fluid de-icing system capacity.
3.609 Carburetor fluid de-icing system detail design.
3.610 Carburetor air preheater design.
3.611 Induction system ducts.
3.612 Induction system screens.
EXHAUST SYSTEM
3.615 General.
3.616 Exhaust manifold.
3.617 Exhaust heat exchangers.
3.618 Exhaust heat exchangers used in ventilating air heating systems.
FIRE WALL AND COWLING
3.623 Fire walls.
3.624 Fire wall construction.
3.625 Cowling.
POWER-PLANT CONTROLS AND ACCESSORIES CONTROLS
3.627 Power-plant controls.
3.628 Throttle controls.
3.629 Ignition switches.
3.630 Mixture controls.
3.631 Propeller speed and pitch controls.
3.632 Propeller feathering controls.
3.633 Fuel system controls.
3.634 Carburetor air preheat controls.
ACCESSORIES
3.635 Power-plant accessories.
3.636 Engine battery ignition systems.
POWER-PLANT FIRE PROTECTION
3.637 Power-plant fire protection.

CAR PART 3-1949 - SUBPART E
SUBPART E--POWER-PLANT INSTALLATIONS; RECIPROCATING ENGINES
GENERAL

§ 3.411 Components. (a) The power-plant installation shall be considered
to include all components of the airplane which are necessary for its
propulsion. It shall also be considered to include all components which
affect the control of the major propulsive units or which affect their
continued safety of operation.
(b) All components of the power-plant installation shall be
constructed, arranged, and installed in a manner which will assure the
continued safe operation of the airplane and power plant. Accessibility
shall be provided to permit such inspection and maintenance as is necessary
to assure continued airworthiness.

ENGINES AND PROPELLERS

§ 3.415 Engines. Engines installed in certificated airplanes shall be of
a type which has been certificated in accordance with the provisions of Part
13 of this chapter.
§ 3.416 Propellers. (a) Propellers installed in certificated airplanes
shall be of a type which has been certificated in accordance with the
provisions of Part 14 of this chapter.
(b) The maximum engine power and propeller shaft rotational speed
permissible for use in the particular airplane involved shall not exceed the
corresponding limits for which the propeller has been certificated.
§ 3.417 Propeller vibration. In the case of propellers with metal blades
or other highly stressed metal components, the magnitude of the critical
vibration stresses under all normal conditions of operation shall be
determined by actual measurements or by comparison with similar
installations for which such measurements have been made. The vibration
stresses thus determined shall not exceed values which have been
demonstrated to be safe for continuous operation. Vibration tests may be
waived and the propeller installation accepted on the basis of service
experience, engine or ground tests which show adequate margins of safety, or
other considerations which satisfactorily substantiate its safety in this
respect. In addition to metal propellers, the Administrator may require that
similar substantiation of the vibration characteristics be accomplished for
other types of propellers, with the exception of conventional fixed-pitch
wood propellers.
§ 3.418 Propeller pitch and speed limitations. The propeller pitch and
speed shall be limited to values which will assure safe operation under all
normal conditions of operation and will assure compliance with the
performance requirements specified in §§ 3.81-3.86.
§ 3.419 Speed limitations for fixed-pitch propellers, ground adjustable
pitch propellers, and automatically varying pitch propellers which cannot be
controlled in flight. (a) During take-off and initial climb at best
rate-of-climb speed, the propeller, in the case of fixed pitch or ground
adjustable types, shall restrain the engine to a speed not exceeding its
maximum permissible take-off speed and, in the case of automatic
variable-pitch types, shall limit the maximum governed engine revolutions
per minute to a speed not exceeding the maximum permissible take-off speed.
In demonstrating compliance with this provision the engine shall be operated
at full throttle or the throttle setting corresponding to the maximum
permissible take-off manifold pressure.
(b) During a closed throttle glide at the placard, "never-exceed
speed" (see § 3.739), the propeller shall not cause the engine to rotate at
a speed in excess of 110 percent of its maximum allowable continuous speed.
§ 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:
(1) Propeller blades set in the lowest possible pitch and the
governor inoperative.
(2) Engine operating at take-off manifold pressure with the airplane
stationary and with no wind.
§ 3.422 Propeller clearance. 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.

FUEL SYSTEM

§ 3.429 General. The fuel system shall be constructed and arranged in a
manner to assure the provision of fuel to each engine at a flow rate and
pressure adequate for proper engine functioning under all normal conditions
of operation, including all maneuvers and acrobatics for which the airplane
is intended.

ARRANGEMENT

§ 3.430 Fuel system arrangement. Fuel systems shall be so arranged as to
permit any one fuel pump to draw fuel from only one tank at a time. Gravity
feed systems shall not supply fuel to any one engine from more than one tank
at a time unless the tank air spaces are interconnected in such a manner as
to assure that all interconnected tanks will feed equally. (See also §
3.439.)
§ 3.431 Multiengine fuel system arrangement. The fuel systems of
multiengine airplanes shall be arranged to permit operation in at least one
configuration in such a manner that the failure of any one component 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. (NOTE: It is not necessarily intended that fuel
tanks proper be separate for each engine if a common tank is provided with
separate outlets and the remainder of the fuel system is independent.)
§ 3.432 Pressure cross feed arrangements. Pressure cross feed lines
shall not pass through portions of the airplane devoted to carrying
personnel or cargo, unless means are provided to permit the flight personnel
to shut off the supply of fuel to these lines, or unless any joints,
fittings, or other possible sources of leakage installed in such lines are
enclosed in a fuel- and fume-proof enclosure which is ventilated and drained
to the exterior of the airplane. Bare tubing need not be enclosed but shall
be protected where necessary against possible inadvertent damage.

OPERATION

§ 3.433 Fuel flow rate. The ability of the fuel system to provide the
required fuel flow rate and pressure shall be demonstrated when the airplane
is in the attitude which represents the most adverse condition from the
standpoint of fuel feed and quantity 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 systems. The fuel flow rate for
gravity systems (main and reserve supply) shall be 150 percent of the actual
take-off consumption of the engine.
§ 3.435 Fuel flow rate for pump systems. 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 .
(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.

FUEL TANKS

§ 3.440 General. Fuel tanks shall be capable of withstanding without
failure any vibration, 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
per cent of the tank capacity or 1 gallon, whichever is greater, a placard
and 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 psi 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 psi 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 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 1/32 of an inch while filled 2/3 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 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
projections which are apt to cause wear of the liner, unless provisions are
made for the protection of the liner at such points or 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 single-engine
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.
§ 3.443 Fuel tank expansion space. 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 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
1/16 gallon, whichever is 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 portion 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.

LINES, FITTINGS, AND ACCESSORIES

§ 3.550 Fuel system lines, fittings, and accessories. Fuel lines shall
be installed and supported in a manner which will prevent excessive
vibration and will be adequate to withstand loads due to fuel pressure and
accelerated flight conditions. Lines which are connected to components of
the airplane between which relative motion might exist shall incorporate
provisions for flexibility. Flexible hose shall be of an acceptable type.
§ 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 purpose shall be located on the side of the
fire wall most remote from the engine.
(b) Shut-off valves shall be so constructed as to make it possible
for the flight personnel to reopen the valves rapidly after they have once
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.
§ 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.
§ 3.554 Fuel system instruments. (See § 3.655 and §§ 3.670 through
3.673.)

OIL SYSTEM

§ 3.561 Oil system. Each engine shall be provided with an independent
oil system capable of supplying the engine with an ample quantity of oil at
a temperature not exceeding the maximum which has been established as safe
for continuous operation. The usable oil tank capacity shall not be less
than the product of the endurance of the airplane under critical operating
conditions and the maximum oil consumption of the engine under the same
conditions, plus a suitable margin to assure adequate circulation and
cooling.

§ 3.562 Oil cooling. (See § 3.581 and pertinent sections.)

OIL TANKS

§ 3.563 Oil tanks. 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. 3.5 psi pressure specified in § 3.441 (a) shall be 5 pounds
psi.
(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.
§ 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.
§ 3.571 Oil valves. (See § 3.637.)
§ 3.572 Oil radiators. Oil radiators and their support shall be capable
of withstanding without failure any vibration, inertia, and oil pressure
loads to which they might normally be subjected.
§ 3.573 Oil filters. If the engine is equipped with an oil filter, the
filter shall be constructed and installed in such a manner that complete
blocking of the flow through the 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
which will not constitute a fire hazard in case foaming occurs and so that
oil emitted from the line will not impinge upon the pilot's windshield. 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 supply, provision shall be made
to trap a quantity of oil in the tank in case the supply becomes depleted
due to failure of any portion of the lubricating system other than the tank
itself. The quantity of oil so trapped shall be sufficient to accomplish the
feathering operation and shall be available only to the feathering pump. The
ability of the system to accomplish feathering when the supply of oil has
fallen to the above level shall be demonstrated.

COOLING

§ 3.581 General. The power-plant cooling provisions shall be capable of
maintaining the temperatures of all power-plant components, engine parts,
and engine fluids (oil and coolant), at or below the maximum established
safe values under critical conditions of ground and flight operation.

TESTS

§ 3.582 Cooling tests. Compliance with the provisions of § 3.581 shall
be demonstrated under 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
power-plant 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.
§ 3.583 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.
§ 3.584 Correction factor for cylinder head, oil inlet, carburetor air,
an 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.
§ 3.585 Correction factor for cylinder barrel temperatures. Cylinder
barrel temperatures shall be corrected by adding 0.7 of the difference
between the maximum anticipated summer air temperature and the temperature
of the ambient air at the time of the first occurrence of the maximum
cylinder barrel temperature recorded during the cooling test.
§ 3.586 Cooling test procedure for single-engine airplanes. This test
shall be conducted by stabilizing engine temperatures in flight and then
starting at the lowest practicable altitude and climbing for 1 minute at
take-off power. 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 conducted 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.
§ 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. 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 .
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-engine-inoperative rate of climb)
shall be commenced at as near sea level as practicable and shall be
conducted at the best rate-of-climb speed (or the speed of minimum rate of
descent, in the case of airplanes with zero or negative
one-engine-inoperative rate of climb).

LIQUID COOLING SYSTEMS

§ 3.588 Independent systems. Each liquid cooled engine shall be provided
with an independent cooling system. The cooling system shall be so arranged
that no air or vapor can be trapped in any portion of the system, except the
expansion tank, either during filling or during operation.
§ 3.589 Coolant tank. A coolant tank shall be provided. The tank
capacity shall not be less than 1 gallon plus 10 percent of the cooling
system capacity. Coolant tanks shall be capable of withstanding without
failure all vibration, inertia, and fluid loads to which they may be
subjected in operation. Coolant tanks shall be provided with an expansion
space of not less than 10 percent of the total cooling system capacity. It
shall not be possible inadvertently to fill the expansion space with the
airplane in the normal ground attitude.
§ 3.590 Coolant tank tests. Coolant tank tests shall be the same as fuel
tank tests (see § 3.441), except as follows:
(a) The 3.5 pounds per square inch pressure test of § 3.441 (a)
shall be replaced by the sum of the pressure developed during the maximum
ultimate acceleration with a full tank or a pressure of 3.5 pounds per
square inch, whichever is greater, plus the maximum working pressure of the
system.
(b) In the case of tanks with nonmetallic liners, the test fluid
shall be coolant rather than fuel as specified in § 3.441 (d), and the slosh
test on a specimen liner shall be conducted with coolant at operating
temperature.
§ 3.591 Coolant tank installation. Coolant tanks shall be supported in a
manner so as to distribute the tank loads over a large portion of the tank
surface. Pads shall be provided to prevent chafing between the tank and the
support. Material used for padding shall be nonabsorbent or shall be treated
to prevent the absorption of inflammable fluids.
§ 3.592 Coolant tank filler connection. Coolant tank filler connections
shall be marked as specified in § 3.767. Provisions shall be made to prevent
the entrance of spilled coolant into the coolant tank compartment or any
portions of the airplane other than the tank itself. Recessed coolant filler
connections shall be drained and the drain shall discharge clear of all
portions of the airplane.
§ 3.593 Coolant lines, fittings, and accessories. Coolant lines shall
comply with the provisions of § 3.550, except that the inside diameter of
the engine coolant inlet and outlet lines shall not be less than the
diameter of the corresponding engine inlet and outlet connections.
§ 3.594 Coolant radiators. Coolant radiators shall be capable of
withstanding without failure any vibration, inertia, and coolant pressure
loads to which they may normally be subjected. Radiators shall be supported
in a manner which will permit expansion due to operating temperatures and
prevent the transmittal of harmful vibration to the radiator. If the coolant
employed is inflammable, the air intake duct to the coolant radiator shall
be so located that flames issuing from the nacelle in case of fire cannot
impinge upon the radiator.
§ 3.595 Cooling system drains. One or more drains shall be provided to
permit drainage of the entire cooling system, including the coolant tank,
radiator, and the engine, when the airplane is in the normal ground
attitude. Drains shall discharge clear of all portions of the airplane and
shall be provided with means for positively locking the drain in the closed
position. Cooling system drains shall be accessible.
§ 3.596 Cooling system instruments. See §§ 3.655, 3.670, and 3.671.

INDUCTION SYSTEM

§ 3.605 General. (a) The engine air induction system shall permit
supplying an adequate quantity of air to the engine under all conditions of
operation.
(b) Each engine shall be provided with at least two separate air
intake sources, except that in the case of an engine equipped with a fuel
injector only one air intake source need be provided, if the air intake,
opening, or passage is unobstructed by a screen, filter, or other part on
which ice might form and so restrict the air flow as to affect adversely
engine operation. It shall be permissible for primary air intakes to open
within the cowling only if that portion of the cowling is isolated from the
engine accessory section by means of a fire-resistant diaphragm or if
provision is made to prevent the emergence of backfire flames. Alternate air
intakes shall be located in a sheltered position and shall not open within
the cowling unless they are so located that the emergence of backfire flames
will not result in a hazard. Supplying air to the engine through the
alternate air intake system of the carburetor air preheater shall not result
in the loss of excessive power in addition to the power lost due to the rise
in the temperature of the air.
§ 3.606 Induction system de-icing and anti-icing provisions. The engine
air induction system shall incorporate means for the prevention and
elimination of ice accumulations in accordance with the provisions in this
section. It shall be demonstrated that compliance with the provisions
outlined in the following paragraphs can be accomplished when the airplane
is operating in air at a temperature of 30° F. when the air is free of
visible moisture.
(a) Airplanes equipped with sea level engines employing conventional
venturi carburetors shall be provided with a preheater capable of providing
a heat rise of 90° F. when the engine is operating at 75 percent of its
maximum continuous power.
(b) Airplanes equipped with altitude engines employing conventional
venturi carburetors shall be provided with a preheater capable of providing
a heat rise of 120° F. when the engine is operating at 75 percent of its
maximum continuous power.
(c) Airplanes equipped with altitude engines employing carburetors
which embody features tending to reduce the possibility of ice formation
shall be provided with a preheater capable of providing a heat rise of 100°
F. when the engine is operating at 60 percent of its maximum continuous
power. However, the preheater need not provide a heat rise in excess of 40°
F. if a fluid de-icing system complying with the provisions of §§
3.607-3.609 is also installed.
(d) Airplanes equipped with sea level engines employing carburetors
which embody features tending to reduce the possibility of ice formation
shall be provided with a sheltered alternate source of air. The preheat
supplied to this alternate air intake shall be not less than that provided
by the engine cooling air downstream of the cylinders.
§ 3.607 Carburetor de-icing fluid flow rate. The system shall be capable
of providing each engine with a rate of fluid flow, expressed in pounds per
hour of not less than 2.5 multiplied by the square root of the maximum
continuous power of the engine. This flow shall be available to all engines
simultaneously. The fluid shall be introduced into the air induction system
at a point close to, and upstream from, the carburetor. The fluid shall be
introduced in a manner to assure its equal distribution over the entire
cross section of the induction system air passages.
§ 3.608 Carburetor fluid de-icing system capacity. The fluid de-icing
system capacity shall not be less than that required to provide fluid at the
rate specified in § 3.607 for a time equal to 3 percent of the maximum
endurance of the airplane. However, the capacity need not in any case exceed
that required for 2 hours of operation nor shall it be less than that
required for 20 minutes of operation at the above flow rate. If the
available preheat exceeds 50° F. but is less than 100° F., it shall be
permissible to decrease the capacity of the system in proportion to the heat
rise available in excess of 50° F.
§ 3.609 Carburetor fluid de-icing system detail design. Carburetor fluid
de-icing systems shall comply with provisions for the design of fuel
systems, except as specified in §§ 3.607 and 3.608, unless such provisions
are manifestly inapplicable.
§ 3.610 Carburetor air preheater design. Means shall be provided to
assure adequate ventilation of the carburetor air preheater when the engine
is being operated in cold air. The preheater shall be constructed in such a
manner as to permit inspection of exhaust manifold parts which it surrounds
and also to permit inspection of critical portions of the preheater itself.
§ 3.611 Induction system ducts. Induction system ducts shall be provided
with drains which will prevent the accumulation of fuel or moisture in all
normal ground and flight attitudes. No open drains shall be located on the
pressure side of turbo-supercharger installations. Drains shall not
discharge in a location which will constitute a fire hazard. Ducts which are
connected to components of the airplane between which relative motion may
exist shall incorporate provisions for flexibility.
§ 3.612 Induction system screens. If induction system screens are
employed, they shall be located upstream from the carburetor. It shall not
be possible for fuel to impinge upon the screen. Screens shall not be
located in portions of the induction system which constitute the only
passage through which air can reach the engine, unless the available preheat
is 100° F. or over and the screen is so located that it can be de-iced by
the application of heated air. De-icing of screens by means of alcohol in
lieu of heated air shall not be acceptable.

EXHAUST SYSTEM

§ 3.615 General. (a) The exhaust system shall be constructed and
arranged in such a manner as to assure the safe disposal of exhaust gases
without the existence of a hazard of fire or carbon monoxide contamination
of air in personnel compartments.
(b) Unless suitable precautions are taken, exhaust system parts
shall not be located in close proximity to portions of any systems carrying
inflammable fluids or vapors nor shall they be located under portions of
such systems which may be subject to leakage. All exhaust system components
shall be separated from adjacent inflammable portions of the airplane which
are outside the engine compartment by means of fireproof shields. Exhaust
gases shall not be discharged at a location which will cause a glare
seriously affecting pilot visibility at night, nor shall they discharge
within dangerous proximity of any fuel or oil system drains. All exhaust
system components shall be ventilated to prevent the existence of points of
excessively high temperature.
§ 3.616 Exhaust manifold. Exhaust manifolds shall be made of fireproof,
corrosion-resistant materials, and shall incorporate provisions to prevent
failure due to their expansion when heated to operating temperatures.
Exhaust manifolds shall be supported in a manner adequate to withstand all
vibration and inertia loads to which they might be subjected in operation.
Portions of the manifold which are connected to components between which
relative motion might exist shall incorporate provisions for flexibility.
§ 3.617 Exhaust heat exchangers. (a) Exhaust heat exchangers shall be
constructed and installed in such a manner as to assure their ability to
withstand without failure all vibration, inertia, and other loads to which
they might normally be subjected. Heat exchangers shall be constructed of
materials which are suitable for continued operation at high temperatures
and which are adequately resistant to corrosion due to products contained in
exhaust gases.
(b) Provisions shall be made for the inspection of all critical
portions of exhaust heat exchangers, particularly if a welded construction
is employed. Heat exchangers shall be ventilated under all conditions in
which they are subject to contact with exhaust gases.
§ 3.618 Exhaust heat exchangers used in ventilating air heating systems.
Heat exchangers of this type shall be so constructed as to preclude the
possibility of exhaust gases entering the ventilating air.

FIRE WALL AND COWLING

§ 3.623 Fire walls. All engines, auxiliary power units, fuel burning
heaters, and other combustion equipment which are intended for operation in
flight shall be isolated from the remainder of the airplane by means of fire
walls, or shrouds, or other equivalent means.
§ 3.624 Fire wall construction. (a) Fire walls and shrouds shall be
constructed in such a manner that no hazardous quantity of liquids, gases,
or flame could pass from the engine compartment to other portions of the
airplane. All openings in the fire wall or shroud shall be sealed tight with
fireproof grommets, bushings, or fire-wall fittings, except that, such seals
of fire-resistant materials shall be acceptable for use on single-engine
airplanes and multiengine airplanes not required to comply with § 3.85 (b)
or § 3.85a (b), if such airplanes are equipped with engine(s) having a
volumetric displacement of 1,000 cubic inches or less; and if the openings
in the fire walls or shrouds are such that, without seals, the passage of a
hazardous quantity of flame could not result.
(b) Fire walls and shrouds shall be constructed of fireproof
material and shall be protected against corrosion. The following materials
have been found to comply with this requirement:
(1) Heat- and corrosion-resistant steel 0.015 inch thick,
(2) Low carbon steel, suitably protected against corrosion, 0.018
inch thick.
§ 3.625 Cowling. (a) Cowling shall be constructed and supported in such
a manner as to be capable of resisting all vibration, inertia, and air loads
to which it may normally be subjected. Provision shall be made to permit
rapid and complete drainage of all portions of the cowling in all normal
ground and flight attitudes. Drains shall not discharge in locations
constituting a fire hazard.
(b) Cowling shall be constructed of fire-resistant material. All
portions of the airplane lying behind openings in the engine compartment
cowling shall also be constructed of fire-resistant materials for a distance
of at least 24 inches aft of such openings. Portions of cowling which are
subjected to high temperatures due to proximity to exhaust system ports or
exhaust gas impingement shall be constructed of fireproof material.

POWER-PLANT CONTROLS AND ACCESSORIES
CONTROLS

§ 3.627 Power-plant controls. Power-plant controls shall comply with the
provisions of §§ 3.384 and 3.762. Controls shall maintain any necessary
position without constant attention by the flight personnel and shall not
tend to creep due to control loads or vibration. Flexible controls shall be
of an acceptable type. Controls shall have adequate strength and rigidity to
withstand operating loads without failure or excessive deflection.
§ 3.628 Throttle controls. A throttle control shall be provided to give
independent control for each engine. Throttle controls shall afford a
positive and immediately responsive means of controlling the engine(s).
Throttle controls shall be grouped and arranged in such a manner as to
permit separate control of each engine and also simultaneous control of all
engines.
§ 3.629 Ignition switches. Ignition switches shall provide control for
each ignition circuit on each engine. It shall be possible to shut off
quickly all ignition on multiengine airplanes, either by grouping of the
individual switches or by providing a master ignition control. If a master
control is provided, suitable means shall be incorporated to prevent its
inadvertent operation.
§ 3.630 Mixture controls. If mixture controls are provided, a separate
control shall be provided for each engine. The controls shall be grouped and
arranged in such a manner as to permit both separate and simultaneous
control of all engines.
§ 3.631 Propeller speed and pitch controls. (See also § 3.421 (a).) If
propeller speed or pitch controls are provided, the controls shall be
grouped and arranged in such a manner as to permit control of all
propellers, both separately and together. The controls shall permit ready
synchronization of all propellers on multiengine airplanes.
§ 3.632 Propeller feathering controls. If propeller feathering controls
are provided, a separate control shall be provided for each propeller.
Propeller feathering controls shall be provided with means to prevent
inadvertent operation.
§ 3.633 Fuel system controls. Fuel system controls shall comply with
requirements of § 3.551 (c).
§ 3.634 Carburetor air preheat controls. Separate controls shall be
provided to regulate the temperature of the carburetor air for each engine.

ACCESSORIES

§ 3.635 Power-plant accessories. Engine-driven accessories shall be of a
type satisfactory for installation on the engine involved and shall utilize
the provisions made on the engine for the mounting of such units. Items of
electrical equipment subject to arcing or sparking shall be installed so as
to minimize the possibility of their contact with any inflammable fluids or
vapors which might be present in a free state.

§ 3.636 Engine battery ignition systems. (a) Battery ignition systems
shall be supplemented with a generator which is automatically made available
as an alternate source of electrical energy to permit continued engine
operation in the event of the depletion of any battery.
(b) The capacity of batteries and generators shall be sufficient to
meet the simultaneous demands of the engine ignition system and the greatest
demands of any of the airplane's electrical system components which may draw
electrical energy from the same source. Consideration shall be given to the
condition of an inoperative generator, and to the condition of a completely
depleted battery when the generator is running at its normal operating
speed. If only one battery is provided, consideration shall also be given to
the condition in which the battery is completely depleted and the generator
is operating at idling speed.
(c) Means shall be provided to warn the appropriate flight personnel
if malfunctioning of any part of the electrical system is causing the
continuous discharging of a battery used for engine ignition. (See § 3.629
for ignition switches.)

POWER-PLANT FIRE PROTECTION

§ 3.637 Power-plant fire protection. Suitable means shall be provided to
shut off the flow in all lines carrying inflammable fluids into the engine
compartment on multiengine airplanes required to comply with the provisions
of § 3.85 (b).

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