**FLOAT SEAPLANES**

§ 3.273 Landing with inclined reactions.

\(a\) The vertical component of the limit load factor shall be 4.2
except that it need not exceed a

value given by the following formula:

n = 3.0 + 0.133 W/S

\(b\) The propeller axis (or equivalent reference line) shall be assumed
to be horizontal and the resultant water reaction to be acting in the
plane of symmetry and passing through the center of gravity of the
airplane, but inclined so

that its horizontal component is equal to one fourth of its vertical
component. Inertia forces shall be assumed to act in a direction
parallel to the water reaction.

\(c\) Factors of safety. For the design of float attachment members,
including the members necessary to complete a rigid brace truss through
the fuselage, the factor of safety shall be 1.85. For the remaining
structural members, the factor of safety shall be 1.5.

§ 3.275 Landing with vertical reactions.

\(a\) The limit load factor shall be 4.33 acting vertically, except that
it need not exceed a value given by the following formula:

n = 3.0 + 0.133 W/S

\(b\) The propeller axis (or equivalent reference line) shall be assumed
to be horizontal, and the resultant water reaction to be vertical and
passing through the center of gravity of the airplane.

\(c\) Factors of safety. The factors of safety shall be the same as
those specified in §3.273(c).

§ 3.277 Landing with side load. The vertical component of the limit load
factor shall be 4.0. The propeller axis (or equivalent reference line)
shall be assumed to be horizontal and the resultant water reaction shall
be assumed to be in the vertical plane which passes through the center
of gravity of the airplane and is perpendicular to the propeller axis.
The vertical load shall be applied through the keel or keels of the
float or floats and evenly divided between the floats when twin floats
are used. A side load equal to one-fourth of the vertical load shall be
applied along a line approximately halfway between the bottom of the
keel and the level of the water line at rest. When twin floats are used,
the entire side load specified shall be applied to the float on the side
from which the water reaction originates.

§ 3.278 Supplementary load conditions. Each main float of a float
seaplane shall be capable of carrying the following loads when supported
at the attachment fittings as installed on the airplane:

\(a\) A limit load, acting upward, applied at the bow end of float and
of magnitude equal to that portion of the airplane weight normally
supported by the particular float.

\(b\) A limit load, acting upward, applied at the stern of magnitude
equal to 0.8 times that portion of the airplane weight normally
supported by the particular float.

\(c\) A limit load, acting upward, applied at the step and of magnitude
equal to 1.5 times that portion of the airplane weight normally
supported by the particular float.

§ 3.279 Bottom loads.

\(a\) Main seaplane float bottoms shall be designed to withstand the
following local pressures:

\(1\) A limit pressure of at least 10 pounds per square inch over that
portion of the bottom lying between the first step and a section 25
percent of the distance from the step to the bow.

\(2\) A limit pressure of at least 5 pounds per square inch over that
portion of the bottom lying between the section 25 percent of the
distance from the strip to the bow and a section 75 percent of the
distance from the step to the bow.

\(3\) A limit pressure of at least 3 pounds per square inch over that
portion of the bottom aft of the step (aft of main step if more than one
step is used).

\(b\) The local pressures determined in paragraph (a) (1), (2) and (3)
of this section shall be applied over local areas in such a manner as to
cause the maximum loads in local structure such as bottom plating and
stringers.

\(c\) For the purpose of designing frames, keels, and chine structure,
distributed bottom pressures equal to one-half of the local values
specified above shall be applied over the entire specified bottom areas.

**WING-TIP FLOAT AND SEA WING LOADS**

§ 3.280 Wing-tip float loads. Wing-tip floats and their attachment,
including the wing structure, shall be analyzed for each of the
following conditions:

\(a\) A limit load acting vertically up at the completely submerged
center of buoyancy and equal to 3 times the completely submerged
displacement.

\(b\) A limit load inclined upward at 45 degrees to the rear and acting
through the completely submerged center of buoyancy and equal to 3 times
the completely submerged displacement.

\(c\) A limit load acting parallel to the water surface (laterally)
applied at the center of area of the side view and equal to 1.5 times
the completely submerged displacement.

§ 3.281 Wing structure. The primary wing structure shall incorporate
sufficient extra strength to insure that failure of wing-tip float
attachment members occurs before the wing structure is damaged.

§ 3.282 Sea wing loads. Sea wing design loads shall be based on suitable
test data.

**SUBPART D---DESIGN AND CONSTRUCTION**

**GENERAL**

§ 3.291 General. The suitability of all questionable design details or
parts having an important bearing on safety in operation shall be
established by tests.

§ 3.292 Materials and workmanship. The suitability and durability of all
materials used in the airplane structure shall be established on the
basis of experience or tests. All materials used in the airplane
structure shall conform to approved specifications which will insure
their having the strength and other properties assumed in the design
data. All workmanship shall be of a high standard.

§ 3.293 Fabrication methods. The methods of fabrication employed in
constructing the airplane structure shall be such as to produce
consistently sound structure. When a fabrication process such as gluing,
spot welding, or heat-treating requires close control to attain this
objective, the process shall be performed in accordance with an approved
process specification.

§ 3.294 Standard fastenings. All bolts, pins, screws, and rivets used in
the structure shall be of an approved type. The use of an approved
locking device or method is required for all such bolts, pins, and
screws. Self-locking nuts shall not be used on bolts subject to rotation
during the operation of the airplane.

§ 3.295 Protection. All members of the structure shall be suitably
protected against deterioration or loss of strength in service due to
weathering, corrosion, abrasion, or other causes. In seaplanes, special
precaution shall be taken against corrosion from salt water,
particularly where parts made from different metals are in close
proximity. Adequate provisions for ventilation and drainage of all parts
of the structure shall be made.

§ 3.296 Inspection provisions. Adequate means shall be provided to
permit the close examination of such parts of the airplane as require
periodic inspection, adjustments for proper alignment and functioning,
and lubrication of moving parts.

**STRUCTURAL PARTS**

§ 3.301 Material strength properties and design values. Material
strength properties shall be based on a sufficient number of tests of
material conforming to specifications to establish design values on a
statistical basis. The design values shall be so chosen that the
probability of any structure being understrength because of material
variations is extremely remote. Values contained in ANC-5 and ANC-18
shall be used unless shown to be inapplicable in a particular case.

Note: ANC-5, \"Strength of Aircraft Elements\" and ANC-18, \"Design of
Wood Aircraft Structures\" are published by the Army-Navy-Civil
Committee on Aircraft Design Criteria and may be obtained from the
Government Printing Office, Washington 25, D.C.

§ 3.302 Special factors. Where there may be uncertainty concerning the
actual strength of particular parts of the structure or where the
strength is likely to deteriorate in service prior to normal
replacement, increased factors of safety shall be provided to insure
that the reliability of such parts is not less than the rest of the
structure as specified in §§ 3.303-3.306.

§ 3.303 Variability factor. For parts whose strength is subject to
appreciable variability due to uncertainties in manufacturing processes
and inspection methods, the factor of safety shall be increased
sufficiently to make the probability of any part being under-strength
from this cause extremely remote. Minimum variability factors (only the
highest pertinent variability factor need be considered) are set forth
in §§ 3.304-3.306.

§ 3.304 Castings.

\(a\) Where visual inspection only is to be employed, the variability
factor shall be 2.0.

\(b\) The variability factor may be reduced to 1.25 for ultimate loads
and 1.15 for limit loads when at least three sample castings are tested
to show compliance with these factors, and all sample and production
castings are visually and radiographically inspected in accordance with
an approved inspection specification.

\(c\) Other inspection procedures and variability factors may be used if
found satisfactory by the Administrator.

§ 3.305 Bearing factors.

\(a\) The factor of safety in bearing at bolted or pinned joints shall
be suitably increased to provide for the following conditions:

\(1\) Relative motion in operation (control surface and system joints
are covered in §§ 3.327-3.347).

\(2\) Joints with clearance (free fit) subject to pounding or vibration.

\(b\) Bearing factors need not be applied when covered by other special
factors.

§ 3.306 Fitting factor. Fittings are denied as parts such as end
terminals used to join one structural member to another. A multiplying
factor of safety of at least 1.15 shall be used in the analysis of all
fittings the strength of which is not proved by limit and ultimate load
tests in which the actual stress conditions are simulated in the fitting
and the surrounding structure. This factor applies to all portions of
the fitting, the means of attachment, and bearing on the

members joined. In the case of integral fittings, the part shall be
treated as a fitting up to the point where the section properties become
typical of the member. The fitting factor need not be applied where a
type of joint design based on

comprehensive test data is used. The following are examples: continuous
joints in metal plating, welded joints, and scarf joints in wood, all
made in accordance with approved practices.

§ 3.307 Fatigue strength. The structure shall be designed, insofar as
practicable, to avoid points of stress concentration where variable
stresses above the fatigue limit are likely to occur in normal service.

**FLUTTER AND VIBRATION**

§ 3.311 Flutter and vibration prevention measures. Wings, tail, and
control surfaces shall be free from flutter, airfoil divergence, and
control reversal from lack of rigidity, for all conditions of operation
within the limit V-n envelope, and the following detail requirements
shall apply:

\(a\) Adequate wing torsional rigidity shall be demonstrated by tests or
other methods found suitable by the Administrator.

\(b\) The mass balance of surfaces shall be such as to preclude flutter.

\(c\) The natural frequencies of all main structural components shall be
determined by vibration tests or other methods found satisfactory by the
Administrator.

**WINGS**

§ 3.317 Proof of strength. The strength of stressed-skin wings shall be
substantiated by load tests or by combined structural analysis and
tests.

§ 3.318 Ribs.

\(a\) The strength of ribs in other than stressed-skin wings shall be
proved by test to at least 125 percent of the ultimate loads for the
most severe loading conditions, unless a rational load analysis and test
procedure is employed and the tests cover the variability of the
particular type of construction.

\(b\) The effects of ailerons and high lift devices shall be properly
accounted for. Rib tests shall simulate conditions in the airplane with
respect to torsional rigidity of spars, fixity conditions, lateral
support, and attachment to

spars.

§ 3.319 External bracing. When wires are used for external lift bracing
they shall be double unless the design provides for a lift-wire-cut
condition. Rigging loads shall be taken into account in a rational or
conservative manner. The end connections of brace wires shall be such as
to minimize restraint against bending or vibration. When brace struts of
large fineness ratio are used, the aerodynamic forces on such struts
shall be taken into account.

§ 3.320 Covering. Strength tests of fabric covering shall be required
unless approved grades of cloth, methods of support, attachment, and
finishing are employed. Special tests shall be required when it appears
necessary to account

for the effects of unusually high design air speeds, slipstream
velocities, or other unusual conditions.

**CONTROL SURFACES (FIXED AND MOVABLE)**

§ 3.327 Proof of strength. Limit load tests of control surfaces are
required. Such tests shall include the horn or fitting to which the
control system is attached. In structural analyses, rigging loads due to
wire bracing shall be taken into account in a rational or conservative
manner.

§ 3.328 Installation. Movable tail surfaces shall be so installed that
there is no interference between the surfaces or their bracing when each
is held in its extreme position and all others are operated through
their full angular movement.

When an adjustable stabilizer is used, stops shall be provided which, in
the event of failure of the adjusting mechanism, will limit its travel
to a range permitting safe flight and landing.

§ 3.329 Hinges. Control surface hinges, excepting ball and roller
bearings, shall incorporate a multiplying factor of safety of not less
than 6.67 with respect to the ultimate bearing strength of the softest
material used as a bearing. For hinges incorporating ball or roller
bearings, the approved rating of the bearing shall not be exceeded.
Hinges shall provide sufficient strength and rigidity for loads parallel
to the hinge line.

**CONTROL SYSTEMS**

§ 3.335 General. All controls shall operate with sufficient ease,
smoothness, and positiveness to permit the proper performance of their
function and shall be so arranged and identified as to provide
convenience in operation and prevent the possibility of confusion and
subsequent inadvertent operation. (See § 3.384 for cockpit controls.)

§ 3.336 Primary flight controls.

\(a\) Primary flight controls are defined as those used by the pilot for
the immediate control of the pitching, rolling, and yawing of the
airplane.

\(b\) For two-control airplanes the design shall be such as to minimize
the likelihood of complete loss of the lateral directional control in
the event of failure of any connecting or transmitting element in the
control system.

§ 3.337 Trimming controls. Proper precautions shall be taken against the
possibility of inadvertent, improper, or abrupt tab operations. Means
shall be provided to indicate to the pilot the direction of control
movement relative to airplane motion and the position of the trim device
with respect of the range of adjustment. The means used to indicate the
direction of the control movement shall be adjacent to the control, and
the means used to indicate the position of the trim device shall be
easily visible to the pilot and so located and operated as to preclude
the possibility of confusion. Trimming devices shall be capable of
continued normal operation notwithstanding the failure of any one
connecting or transmitting element in the primary flight control system.
Tab controls shall be irreversible unless the tab is properly balanced
and possesses no unsafe flutter characteristics. Irreversible tab
systems shall provide adequate rigidity and reliability in the portion
of the system from the tab to the attachment of the irreversible unit to
the airplane

structure.

§ 3.338 Wing flap controls. The controls shall be such that when the
flap has been placed in any position upon which compliance with the
performance requirements is based, the flap will not move from that
position except upon further

adjustment of the control or the automatic operation of a flap load
limiting device. Means shall be provided to indicate the flap position
to the pilot. If any flap position other than fully retracted or
extended is used to show compliance

with the performance requirements, such means shall indicate each such
position. The rate of movement of the flaps in response to the operation
of the pilot's control, or of an automatic device shall not be such as
to result in unsatisfactory flight or performance characteristics under
steady or changing conditions of air speed, engine power, and airplane
attitude (See § 3.109 (b) and (c).)

§ 3.339 Flap interconnection.

\(a\) The motion of flaps on opposite sides of the plane of symmetry
shall be synchronized by a mechanical

interconnection, unless the airplane is demonstrated to have safe flight
characteristics while the flaps are retracted on one side and extended
on the other.

\(b\) Where an interconnection is used, in the case of multiengine
airplanes, it shall be designed to account for the unsymmetrical loads
resulting from flight with the engines on one side of the plane of
symmetry inoperative and the

remaining engines at take-off power. For single engine airplanes, it may
be assumed that 100 percent of the critical air load acts on one side
and 70 percent on the other.

§ 3.340 Stops. All control systems shall be provided with stops which
positively limit the range of motion of the control surfaces. Stops
shall be so located in the system that wear, slackness, or take-up
adjustments will not appreciably affect the range of surface travel.
Stops shall be capable of withstanding the loads corresponding to the
design conditions for the control system.

§ 3.341 Control system locks. When a device is provided for locking a
control surface while the airplane is on the ground or water:

\(a\) The locking device shall be so installed as to provide
unmistakable warning to the pilot when it is engaged, and

\(b\) Means shall be provided to preclude the possibility of the lock
becoming engaged during flight.

§ 3.342 Proof of strength. Tests shall be conducted to prove compliance
with limit load requirements. The direction of test loads shall be such
as to produce the most severe loading of the control system structure.
The tests shall include all fittings, pulleys, and brackets used to
attach the control system to the primary structure. Analyses or
individual load tests shall be conducted to demonstrate compliance with
the multiplying factor of safety requirements specified for control
system joints subjected to angular motion.

§ 3.343 Operation test. An operation test shall be conducted by
operating the controls from the pilot compartment with the entire system
so loaded as to correspond to the limit air loads on the surface. In
this test there shall be no

jamming, excessive friction, or excessive deflection.

**CONTROL SYSTEM DETAILS**

§ 3.344 General. All control systems and operating devices shall be so
designed and installed as to prevent jamming, chafing, or interference
as a result of inadequate clearances or from cargo, passengers, or loose
objects. Special precautions shall be provided in the cockpit to prevent
the entry of foreign objects into places where they might jam the
controls. Provisions shall be made to prevent the slapping of cables or
tubes against parts of the airplane.

§ 3.345 Cable systems. Cables, cable fittings, turnbuckles, splices, and
pulleys shall be in accordance with approved specifications. Cables
smaller than 1/8-inch diameter shall not be used in primary control
systems. The design of

cable systems shall be such that there will not be hazardous change in
cable tension throughout the range of travel under operating conditions
and temperature variations. Pulley types and sizes shall correspond to
the cables with which they are used, as specified on the pulley
specification. All pulleys shall be provided with satisfactory guards
which shall be closely fitted to prevent the cables becoming misplaced
or fouling, even when slack. The pulleys shall lie in the plane passing
through the cable within such limits that the cable does not rub against
the pulley flange. Fairleads shall be so installed that they are not
required to cause a change in cable direction of more than 3 degrees.
Clevis pins (excluding those not subject to load or motion) retained
only by cotter pins shall not be employed in the control system.
Turnbuckles shall be attached to parts having angular motion in such a
manner as to prevent positively binding throughout the range of travel.
Provisions for visual inspection shall be made at all fairleads,
pulleys, terminals, and turnbuckles.

§ 3.346 Joints. Control system joints subject to angular motion in
push-pull systems, excepting ball and roller bearing systems, shall
incorporate a multiplying factor of safety of not less than 3.33 with
respect to the ultimate bearing strength of the softest material used as
a bearing. This factor may be reduced to 2.0 for such joints in cable
control systems. For ball or roller bearings the approved rating of the
bearing shall not be exceeded.

§ 3.347 Spring devices. The reliability of any spring devices used in
the control system shall be established by tests simulating service
conditions, unless it is demonstrated that failure of the spring will
not cause flutter or unsafe flight characteristics.

**LANDING GEAR**

**SHOCK ABSORBERS**

§ 3.351 Tests. Shock absorbing elements in main, nose, and tail wheel
units shall be substantiated by the tests specified in the following
section. In addition, the shock absorbing ability of the landing gear in
taxiing must be demonstrated in the operational tests of § 3.146.

§ 3.352 Shock absorption tests.

\(a\) It shall be demonstrated by energy absorption tests that the limit
load factors selected for design in

accordance with § 3.243 will not be exceeded in landings with the limit
descent velocity specified in that section.

\(b\) In addition, a reserve of energy absorption shall be demonstrated
by a test in which the descent velocity is at least 1.2 times the limit
descent velocity. In this test there shall be no failure of the shock
aborting unit, although

yielding of the unit will be permitted. Wing lift equal to the weight of
the airplane may be assumed for purposes of this test.

§ 3.353 Limit drop tests.

\(a\) If compliance with the specified limit landing conditions of §
3.352 (a) is demonstrated by free drop tests, these shall be conducted
on the complete airplane, or on units consisting of wheel, tire, and
shock absorber in their proper relation, from free drop heights not less
than the following:

h (inches) = 3.6 (W/S)0.5

except that the free drop height shall not be less than 9.2 inches and
need not be greater than 18.7 inches.

\(b\) In simulating the permissible wing lift in free drop tests, the
landing gear unit shall be dropped with an effective mass equal to:

![](media/image1.png){width="1.09375in" height="0.3333333333333333in"}

where

*We* = the effective weight to be used in the drop test.

h = specified height of drop in inches.

d = deflection under impact of the tire (at the approved inflation
pressure) plus the vertical component of the axle travel relative to the
drop mass. The value of d used in the computation of *We* shall not
exceed the value actually obtained in the drop tests.

W = WM or main gear units, and shall be equal to the static weight on
the particular unit with the airplane in the level attitude (with the
nose wheel clear, in the case of nose wheel clear, in the case of nose
wheel type airplanes).

W = WT for tail gear units, and shall be equal to the static weight on
the tail unit with the airplane in the tail down attitude.

W = WN for nose wheel units, and shall be equal to the static reaction
which will exist at the nose wheel when the mass of the airplane is
concentrated at the center of gravity and exerts a force of 1.0g
downward and 0.33g forward.

L = ratio of assumed wing lift to airplane weight, not greater than
0.667.

The attitude in which the landing gear unit is drop tested shall be such
as to simulate the airplane landing condition which is critical from the
standpoint of energy to be absorbed by the particular unit.

§ 3.354 Limit load factor determination. In determining the limit
airplane inertia load factor n from the free drop test described above,
the following formula shall be used:

![](media/image2.png){width="0.9166666666666666in"
height="0.3333333333333333in"}

where

nj = the load factor developed in the drop test, i.e., the acceleration
(dv/dt) in g's recorded in the drop test, plus 1.0.

The value of n so determined shall not be greater than the limit inertia
load factor used in the landing conditions, § 3.243.

§ 3.355 Reserve energy absorption drop tests. If compliance with the
reserve energy absorption condition specified in § 3.352 (b) is
demonstrated by free drop tests, the drop height shall be not less than
1.44 times the drop height

specified in § 3.353. In simulating wing lift equal to the airplane
weight, the units shall be dropped with an effective mass equal to

![](media/image3.png){width="0.9479166666666666in"
height="0.3541666666666667in"}

where the symbols and other details are the same as in § 3.353.

**RETRACTING MECHANISM**

§ 3.356 General. The landing gear retracting mechanism and supporting
structure shall be designed for the maximum load factors in the flight
conditions when the gear is in the retracted position. It shall also be
designed for the combination of friction, inertia, brake torque, and air
loads occurring during retraction at any air speed up to 1.6Vs1, flaps
retracted and any load factors up to those specified for the flaps
extended condition, § 3.190. The landing gear

and retracting mechanism, including the wheel well doors, shall
withstand flight loads with the landing gear extended at any speed up to
at least 1.6 Vs1 flaps retracted. Positive means shall be provided for
the purpose of maintaining the

wheels in the extended position.

§ 3.357 Emergency operation. When other than manual power for the
operation of the landing gear is employed, an auxiliary means of
extending the landing gear shall be provided.

§ 3.358 Operation test. Proper functioning of the landing gear
retracting mechanism shall be demonstrated by operation tests.

§ 3.359 Position indicator and warning device. When retractable landing
wheels are used, means shall be provided for indicating to the pilot
when the wheels are secured in the extreme positions. In addition,
landplanes shall be provided with an aural or equally effective warning
device which shall function continuously after the throttle is closed
until the gear is down and locked.

§ 3.360 Control. See § 3.384.

**WHEELS AND TIRES**

§ 3.361 Wheels.

\(a\) Main landing gear wheels (i.e., those nearest the airplane center
of gravity) shall be of an approved type.

\(b\) The rated static load of each main wheel shall not be less than
the design weight for ground loads (§ 3.242) divided by the number of
main wheels. Nose wheels shall have been tested for an ultimate radial
load not less than the maximum nose wheel ultimate load obtained in the
ground loads requirements, and for corresponding side and burst loads.

§ 3.362 Tires. A landing gear wheel may be equipped with any make or
type of tire, provided that the approved tire rating is not exceeded
under the following conditions:

\(a\) Load on main wheel tires equal to the airplane weight divided by
the number of wheels,

\(b\) Load on nose wheel tires (to be compared with the dynamic rating
established for such tires) equal to the reaction obtained at the nose
wheel, assuming the mass of the airplane concentrated at the center of
gravity and exerting a force of 1.0g downward and 0.31g forward, the
reactions being distributed to the nose and main wheels by the principle
of statics with the drag reaction at the ground applied only at those
wheels having brakes. When specially constructed tires are used to
support an airplane, the wheels shall be plainly and conspicuously
marked to that effect. Such marking shall include the make, size, number
of plies, and identification marking of the proper tire.

Note: Approved ratings are those assigned by the Tire and Rim
Association or by the Administrator.

**BRAKES**

§ 3.363 Brakes. Brakes shall be installed which are adequate to prevent
the airplane from rolling on a paved runway while applying take-off
power to the critical engine, and of sufficient capacity to provide
adequate speed control

during taxiing without the use of excessive pedal or hand forces.

**SKIS**

§ 3.364 Skis. Skis shall be of an approved typed. The approved rating of
the skis shall not be less than the maximum weight of the airplane on
which they are installed.

§ 3.365 Installation.

\(a\) When type certificated skis are installed, the installation shall
be made in accordance with the ski or airplane manufacturer's
recommendations which shall have been approved by the Administrator.
When other than type certificated skis are installed, data shall be
submitted to the Administrator showing a dimensional drawing of the
proposed method of attaching the skis, the sizes and material of the
restraining members and attachment fittings.

\(b\) In addition to such shock cord(s) as may be provided, front and
rear check cables shall be used on skis not equipped with special
stabilizing devices.

§ 3.366 Tests.

\(a\) If the airplane is of a model not previously approved with the
specific ski installation, it shall satisfactorily pass a ground
inspection of the installation, demonstrate satisfactory landing and
taxiing characteristics, and comply with such flight tests as are found
necessary to indicate that the airplane's flight characteristics are
satisfactory with the skis installed.

\(b\) If the airplane is of a model previously approved with the
specific ski installation, it need pass satisfactorily only a ground
inspection of the installation.

**HULLS AND FLOATS**

§ 3.371 Buoyancy (main seaplane floats).

\(a\) Main seaplane floats shall have a buoyancy in excess of that
required to support the maximum weight of the airplane in fresh water as
follows:

\(1\) 80 percent in the case of single floats.

\(2\) 90 percent in the case of double floats.

\(b\) Main seaplane floats for use on airplanes of 2,500 pounds or more
maximum weight shall contain at least 5 watertight compartments of
approximately equal volume. Main seaplane floats for use on airplanes of
less than 2,500 pounds maximum weight shall contain at least four such
compartments.

§ 3.372 Buoyancy (boat seaplanes). The hulls of boat seaplanes and
amphibians shall be divided into watertight compartments in accordance
with the following requirements:

\(a\) In airplanes of 5,000 pounds or more maximum weight, the
compartments shall be so arranged that, with any two adjacent
compartments flooded, the hull and auxiliary floats (and tires, if used)
will retain sufficient buoyancy to support the maximum weight of the
airplane in fresh water.

\(b\) In airplanes of 1,500 to 5,000 pounds maximum weight, the
compartments shall be so arranged that, with any one compartment
flooded, the hull and auxiliary floats (and tires, if used) will retain
sufficient buoyancy to support the maximum weight of the airplane in
fresh water.

\(c\) In airplanes of less than 1,500 pounds maximum weight, watertight
subdivision of the hull is not required.

\(d\) Bulkheads may have watertight doors for the purpose of
communication between compartments.

§ 3.373 Water stability. Auxiliary floats shall be so arranged that when
completely submerged in fresh water, they will provide a righting moment
which is at least 1.5 times the upsetting moment caused by the airplane
being tilted. A greater degree of stability may be required by the
Administrator in the case of large flying boats, depending on the height
of the center of gravity above the water level, the area and location of
wings and tail surfaces, and other considerations.

**FUSELAGE**

**PILOT COMPARTMENT**

§ 3.381 General.

\(a\) The arrangement of the pilot compartment and its appurtenances
shall provide a satisfactory degree of safety and assurance that the
pilot will be able to perform all his duties and operate the controls in
the correct manner without unreasonable concentration and fatigue.

\(b\) The primary flight control units listed on Figure 3-14, excluding
cables and control rods, shall be so located with respect to the
propellers that no portion of the pilot or controls lies in the region
between the plane of rotation of any inboard propeller and the surface
generated by a line passing through the center of the propeller hub and
making an angle of 5° forward or aft of the plane of rotation of the
propeller.

§ 3.382 Vision. The pilot compartment shall be arranged to afford the
pilot a sufficiently extensive, clear, and undistorted view for the safe
operation of the airplane. During flight in a moderate rain condition,
the pilot shall have an adequate view of the flight path in normal
flight and landing, and have sufficient protection from the elements so
that his vision is not unduly impaired. This may be accomplished by
providing an openable window or by a means for maintaining a portion of
the windshield in a clear condition without continuous attention by the
pilot. The pilot compartment shall be free of glare and reflections
which would interfere with the pilot's vision. For airplanes intended
for night operation, the demonstration of these qualities shall include
night flight tests.

§ 3.383 Pilot windshield and windows. All glass panes shall be of a
nonsplintering safety type.

§ 3.384 Cockpit controls.

\(a\) All cockpit controls shall be so located and, except for those the
function of which is obvious, identified as to provide convenience in
operation including provisions to prevent the possibility of confusion
and consequent inadvertent operations. (See Fig. 3-14 for required sense
of motion of cockpit controls.) The controls shall be so located and
arranged that when seated it will be readily possible for the pilot to
obtain full and unrestricted movement of each control without
interference form either his clothing or the cockpit structure.

\(b\) Identical power-plant controls for the several engines in the case
of multiengine airplanes shall be so located as to prevent any
misleading impression as to the engines of which they relate.

+-----------------------------------+----------------------------------+
| > Control                         | > Movement and actuation         |
+-----------------------------------+----------------------------------+
| > **Primary:**                    |                                  |
+-----------------------------------+----------------------------------+
| > Aileron                         | > Right (clockwise) for right    |
|                                   | > wing down.                     |
+-----------------------------------+----------------------------------+
| > Elevator                        | > Rearward for nose up.          |
+-----------------------------------+----------------------------------+
| > Rudder                          | > Right pedal forward for nose   |
|                                   | > right.                         |
+-----------------------------------+----------------------------------+
| > **Power plant:**                |                                  |
+-----------------------------------+----------------------------------+
| > Throttle                        | > Forward to open.               |
+-----------------------------------+----------------------------------+

Figure 3-14 Cockpit Controls

§ 3.385 Instruments and markings. See § 3.661 relative to instrument
arrangement. The operational markings, instructions, and placards
required for the instruments and controls are specified in §§ 3.756 to
3.765.

**EMERGENCY PROVISIONS**

§ 3.386 Protection. The fuselage shall be designed to give reasonable
assurance that each occupant, if he makes proper use of belts or harness
for which provisions are made in the design, will not suffer serious
injury during minor crash conditions as a result of contact of any
vulnerable part of his body with any penetrating or relatively solid
object, although it is accepted that parts of the airplane may be
damaged.

\(a\) The ultimate accelerations to which occupants are assumed to be
subjected shall be as follows:

+-----------------------+----------------------+----------------------+
|                       | > N, U               | > A                  |
+-----------------------+----------------------+----------------------+
| > Upward              | > 3.0g               | > 4.5g               |
+-----------------------+----------------------+----------------------+
| > Forward             | > 9.0g               | > 9.0g               |
+-----------------------+----------------------+----------------------+
| > Sideward            | > 1.5g               | > 1.5g               |
+-----------------------+----------------------+----------------------+

\(b\) For airplanes having retractable landing gear, the fuselage in
combination with other portions of the structure shall be designed to
afford protection of the occupants in a wheels-up landing with moderate
descent velocity.

\(c\) If the characteristics of an airplane are such as to make a
turn-over reasonably probable, the fuselage of such an airplane in
combination with other portions of the structure shall be designed to
afford protection of the occupants in a complete turn-over.

Note: In § 3.386 (b) and (c), a vertical ultimate acceleration of 3g and
a friction coefficient of 0.5 at the ground may be assumed.

§ 3.387 Exits.

\(a\) Closed cabins on airplanes carrying more than 5 persons shall be
provided with emergency exits consisting of movable windows or panels or
of additional external doors which provide a clear and unobstructed
opening, the minimum dimensions of which shall be such that a
19-by-26-inch ellipse may be completely inscribed therein. The exits
shall be readily accessible, shall not require exceptional agility of a
person using them, and shall be distributed so as to facilitate egress
without crowding in all probable attitudes resulting from a crash. The
method of opening shall be simple and obvious, and the exits shall be so
arranged and marked as to be readily located and operated even in
darkness. Reasonable provisions shall be made against the jamming of
exits as a result of fuselage deformation. The proper functioning of
exits shall be demonstrated by tests.

\(b\) The number of emergency exits required is as follows:

\(1\) Airplanes with a total seating capacity of more than 5 persons,
but not in excess of 15, shall be provided with at least one emergency
exit or one suitable door in addition to the main door specified in §
3.389. This emergency exit, or second door, shall be on the opposite
side of the cabin from the main door.

\(2\) Airplanes with a seating capacity of more than 15 persons shall be
provided with emergency exits or doors in addition to those required in
paragraph (b) (1) of this section. There shall be one such additional
exit or door located either in the top or side of the cabin for every
additional 7 persons or fraction thereof above 15, except that not more
than four exits, including doors, will be required if the arrangement
and dimensions are suitable for quick evacuation of all occupants.

\(c\) If the pilot compartment is separated from the cabin by a door
which is likely to block the escape in the event of a minor crash, it
shall have its own exit, but such exit shall not be considered as an
emergency exit for the

passengers.

\(d\) In categories U and A exits shall be provided which will permit
all occupants to bail out quickly with parachutes.

§ 3.388 Fire precautions---

\(a\) Cabin interiors. Only materials which are flash resistant shall be
used. In compartments where smoking is to be permitted, the materials of
the cabin lining, floors, upholstery, and furnishings shall be
flame-resistant. Such compartments shall be equipped with an adequate
number of self contained ash trays. All other compartments shall be
placarded against smoking.

\(b\) Combustion heaters. Gasoline operated combustion heater
installations shall comply with applicable parts of the power-plant
installation requirements covering fire hazards and precautions. All
applicable requirements concerning fuel tanks, lines, and exhaust
systems shall be considered.

**PERSONNEL AND CARGO ACCOMMODATIONS**

§ 3.389 Doors. Closed cabins on all airplanes carrying passengers shall
be provided with at least one adequate and easily accessible external
door. No passenger door shall be so located with respect to the
propeller discs as to endanger persons using the door.

§ 3.390 Seats and berths---

\(a\) Passenger seats and berths. All seats and berths and supporting
structure shall be designed for a passenger weight of 170 pounds (190
pounds with parachute for the acrobatic and utility categories) and the
maximum load factors corresponding to all specified flight and ground
load conditions including the emergency conditions of § 3.386.

\(b\) Pilot seats. Pilot seats shall be designed for the reactions
resulting from the application of the pilot forces to the primary flight
controls as specified in § 3.231.

\(c\) Categories U and A. All seats designed to be occupied in the U and
A categories under §3.74 (c) (4) shall be designed to accommodate
passengers wearing parachutes.

§ 3.391 Safety belt or harness provisions. Provisions shall be made at
all seats and berths for the installation of belts or harness of
sufficient strength to comply with the emergency conditions of § 3.386.

§ 3.392 Cargo compartments. Each cargo compartment shall be designed for
the placarded maximum weight of contents and critical load distributions
at the appropriate maximum load factors corresponding to all specified
flight and ground load conditions. Suitable provisions shall be made to
prevent the contents of cargo compartments form becoming a hazard by
shifting. Such provisions shall be adequate to protect the passengers
from injury by the contents of any cargo compartment when the ultimate
forward acting accelerating force is 4.5g.

§ 3.393 Ventilation. All passenger and crew compartments shall be
suitably ventilated. Carbon monoxide concentration shall not exceed 1
part in 20,000 parts of air.

**MISCELLANEOUS**

§ 3.401 Leveling marks. Leveling marks shall be provided for leveling
the airplane on the ground.

**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 airplanes equipped with
metal propellers, the magnitude of the propeller blade 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 takeoff

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 blade 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 taxiing 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 such a
manner that the failure of any one component will not result in the loss
of the power of more than one engine. Unless other provisions are made
in order 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.

§ 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 feed systems. The fuel flow rate for
gravity feed systems (main and reserve supply) shall be 1.2 pounds per
hour for each take-off horsepower or 150 percent of the actual take-off
fuel consumption of the engine, whichever is greater.

§ 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
instruction 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 Vso.

\(4\) 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. The fuel system shall be so
arranged as to minimize the possibility of the formation of vapor lock
in the system under all normal conditions of operation.

§ 3.439 Flow between interconnected tanks. In the case of gravity feed
systems with tanks who's 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 not be less than 1
gallon for each seven maximum continuous rated horsepower for which the
airplane is certificated. 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 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. Subsequent tanks shall be production tested to at least
0.5 psi. 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 anxle 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 support for tanks shall not be such as to
concentrate the loads resulting from the weight of the fluid 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 fluids. If flexible tank
liners are employed, 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 protection of
the liner at such points or unless the construction of the liner itself
provides such protection.

\(b\) Tank compartments shall be ventilated and drained to prevent the
accumulation of inflammable fluids or vapors. Compartments adjacent of
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. The sump may be dispensed with if the fuel
system is provided with a sediment bowl which will permit visual ground
inspection for accumulation of water or other foreign material. The
sediment bowl shall also be readily 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 above
shall be effective with the airplane in the normal ground attitude.

\(c\) If a separate sediment bowl is provided, the fuel tank outlet
shall be so located that water will drain from all portions of the tank
to the outlet when the airplane is in the ground attitude.

§ 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,
unless the engine-driven pumps have been approved with the engines, in
which case emergency pumps need not be provided. Similarly, if an engine
fuel injection pump which has been certificated as an integral part of
the engine is used, an emergency pump will not be required. Emergency
pumps shall be capable of

complying with the same flow requirements as are prescribed for the main
pumps. Hand emergency pumps shall not require excessive effort for their
continued operation at the rate of 60 complete cycles (120 single
strokes) per minute. Emergency pumps shall be available for immediate
use in case of the failure of any other pump.

**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.

§ 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.

§ 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 oil capacity of the system shall not be less
than 1 gallon for every 25 gallons of fuel capacity. However, in no case
shall the oil capacity be less than 1 gallon for each 75 maximum
continuous horsepower of the engine(s) involved unless lower quantities
can be substantiated.

§ 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 pound
psi.

\(b\) In the case of tanks with nonmetabllic 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 ½
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 an 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 setting 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,
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.

§ 3.585 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 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 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-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. Primary and alternate air intakes may
open within the cowling only if that portion of the cowling is isolated
from the engine accessory section by means of a fireproof diaphragm.
Alternate air intakes shall be located in a sheltered position.
Supplying air to the engine through the alternate air intake system or
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 antiicing 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.

§ 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 in applicable.

§ 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 exchanges 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 air, fluids, or flame can pass from the engine
compartment to other portions of the airplane. All openings in the fire
wall or shroud shall be sealed with closefitting fireproof grommets,
bushings, or fire-wall fittings.

\(b\) 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 3.759. 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.

§ 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 control 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.

**SUBPART F - EQUIPMENT**

§ 3.651 General. The equipment specified in § 3.655 shall be the minimum
installed when the airplane is submitted to determine its compliance
with the airworthiness requirements. Such additional equipment as is
necessary for a specific type of operation is specified in other
pertinent parts of the Civil Air Regulations, but, where necessary, its
installation and that of the items mentioned in § 3.655 is covered
herein.

§ 3.652 Functional and installational requirements. Each item of
equipment which is essential to the safe operation of the airplane shall
be found by the Administrator to perform adequately the functions for
which it is to be used, shall function properly when installed, and
shall be adequately labeled as to its identification, function,
operational limitations, or any combination of these, whichever is
applicable. Items of equipment for which type certification is required
shall have been certificated in accordance with the provisions of Part
15 of this chapter (or previous regulations) and such other parts as may
be applicable.

**BASIC EQUIPMENT**

§ 3.655 Required basic equipment. The following table shows the basic
equipment items required for type and airworthiness certification of an
airplane:

\(a\) Flight and navigational instruments.

\(1\) Air-speed indicator (see § 3.663).

\(2\) Altimeter.

\(3\) Magnetic direction indicator (see § 3.666)

\(b\) Power-plant instruments---(1) For each engine or tank. (i) Fuel
quantity indicator (see § 3.672).

\(ii\) Oil pressure indicator.

\(iii\) Oil temperature indicator.

\(iv\) Tachometer.

\(2\) For each engine or tank (if required in reference section). (i)
Carburetor air temperature indicator (see § 3.676).

\(ii\) Coolant temperature indicator (if liquid cooled engines used).

\(iii\) Cylinder head temperature indicator (see § 3.675).

\(iv\) Fuel pressure indicator (if pump-fed engines used).

\(v\) Manifold pressure indicator (if altitude engines used).

\(vi\) Oil quantity indicator (see § 3.674).

\(c\) Electrical equipment (if required by reference section).

\(1\) Master switch arrangement (see § 3.688).

\(2\) Adequate source(s) of electrical energy (see §§ 3.682 and 3.685).

\(3\) Electrical protective devices (see § 3.690).

\(d\) Miscellaneous equipment.

\(1\) Certificated safety belts for all occupants (see Part 15 of this
chapter).

\(2\) Airplane Flight Manual (see § 3.777).

**INSTRUMENTS; INSTALLATION**

**GENERAL**

§ 3.661 Arrangement and visibility of instrument installations.

\(a\) Flight, navigation, and power-plant instruments for use by each
pilot shall be easily visible to him.

\(b\) On multiengine airplanes, identical power-plant instruments for
the several engines shall be so located as to, prevent any confusion as
to the engines to which they relate.

§ 3.662 Instrument panel vibration characteristics. Vibration
characteristics of the instrument panel shall not be such as to impair
the accuracy of the instruments or to cause damage to them.

**FLIGHT AND NAVIGATIONAL INSTRUMENTS**

§ 3.663 Air-speed indicating system. This system shall be so installed
that the air-speed indicator shall indicate true air speed at sea level
under standard conditions to within an allowable installational error of
not more than plus or minus 3 percent of the calibrated air speed or 5
miles per hour, whichever is greater, throughout the operating range of
the airplane with flaps up from Vc to 1.3 Vs1 and with flaps at 1.3 Vs1.
The calibration shall be made in flight.

§ 3.664 Air-speed indicator-marking. The air-speed indicator shall be
marked as specified in § 3.757.

§ 3.665 Static air vent system. All instruments provided with static air
case connections shall be so vented that the influence of airplane
speed, the opening and closing of windows, air-flow variation, moisture,
or other foreign matter will not seriously affect their accuracy.

§ 3.666 Magnetic direction indicator. The magnetic direction indicator
shall be so installed that its accuracy shall not be excessively
affected by the airplane's vibration or magnetic fields. After the
direction indicator has been compensated, the installation shall be such
that the deviation in level flight does not exceed 10 degrees on any
heading. A suitable calibration placard shall be provided as specified
in § 3.758.

§ 3.667 Automatic pilot system. If an automatic pilot system is
installed:

\(a\) The actuating (servo) devices shall be of such design that they
can, when necessary, be positively disengaged from operating the control
system or be overpowered by the human pilot to enable him to maintain
satisfactory control of the airplane.

\(b\) A satisfactory means shall be provided to indicate readily to the
pilot the alignment of the actuating device in relation to the control
system which it operates, except when automatic synchronization is
provided.

\(c\) The manually operated control(s) for the system's operation shall
be readily accessible to the pilot.

\(d\) The automatic pilot system shall be of such design and so adjusted
that it cannot produce loads in the control system and surfaces greater
than those for which they were designed.

§ 3.668 Gyroscopic indicators (air-driven type). All air-driven
gyroscopic instruments installed in airplanes which are certificated for
instrument flight operations shall derive their energy from a reliable
suction source of sufficient capacity to maintain their required
accuracy at all speeds above the best rate-of-climb speed. In addition
the system shall be so installed as to preclude malfunctioning due to
rain, oil, or other detrimental elements. On multiengine airplanes,

the following detail requirements shall be applicable:

\(a\) Two sources actuated by separate means shall be provided, either
one of which shall be of sufficient capacity to operate all of the
air-driven gyroscopic instruments with which the airplane is equipped,
with the airplane in normal cruising attitude at 65 percent maximum
continuous power.

\(b\) A suitable means shall be provided in the attendant installation
where the source lines connect into a common line to select either
suction air source for the proper functioning of the instruments should
failure of one source or a breakage of one source line occur. When an
automatic means to permit simultaneous air flow is provided in the
system, a suitable method for maintaining suction shall be provided. In
order to indicate which source of energy has failed, a visual means
shall be provided to indicate this condition to the flight crew.

§ 3.669 Suction gauge. A suction gauge shall be provided and so
installed as to indicate readily to the flight crew while in flight the
suction in inches of mercury which is being applied to the air-driven
types of gyroscopic instruments. This gauge shall be connected to the
instruments by a suitable system.

**POWER-PLANT INSTRUMENTS**

§ 3.670 Operational markings. Instruments shall be marked as specified
in § 3.759.

§ 3.671 Instrument lines. Power-plant instrument lines shall comply with
the provisions of § 3.550. In addition, instrument lines carrying
inflammable fluids or gases under pressure shall be provided with
restricted orifices or other

safety devices at the source of the pressure to prevent escape of
excessive fluid or gas in case of line failure.

§ 3.672 Fuel quantity indicator. Means shall be provided to indicate to
the flight personnel the quantity of fuel in each tank during flight.
Tanks, the outlets and air spaces of which are interconnected, may be
considered as one tank and need not be provided with separate
indicators. Exposed sight gauges shall be so installed and guarded as to
preclude the possibility of breakage or damage. Fuel quantity indicators
shall be calibrated to read zero during level flight when the quantity
of fuel remaining in the tank is equal to the unusable fuel supply as
defined by § 3.437. \[12 F.R. 3438\]

§ 3.673 Fuel flowmeter system. When a fuel flowmeter system is installed
in the fuel line(s), the metering component shall be of such design as
to include a suitable means for bypassing the fuel supply in the event
that malfunctioning of the metering component offers a severe
restriction to fuel flow.

§ 3.674 Oil quantity indicator. Ground means, such as a slick gauge,
shall be provided to indicate the quantity of oil in each tank. If an
oil transfer system or a reserve oil supply system is installed, means
shall be provided to indicate to the flight personnel during flight the
quantity of oil in each tank.

§ 3.675 Cylinder head temperature indicating system for air-cooled
engines. A cylinder head temperature indicator shall be provided for
each engine on airplanes equipped with cowl flaps. In the case of
airplanes which do not have cowl flaps, an indicator shall be provided
if compliance with the provisions of § 3.581 is demonstrated at a speed
in excess of the speed of best rate of climb.

§ 3.676 Carburetor air temperature indicating system. A carburetor air
temperature indicating system shall be provided for each altitude engine
equipped with a preheater which is capable of providing a heat rise in
excess of 60°F.

**ELECTRICAL SYSTEMS AND EQUIPMENT**

§ 3.681 Installation.

\(a\) Electrical systems in airplanes shall be free from hazards in
themselves, in their method of operation, and in

their effects on other parts of the airplane. Electrical equipment shall
be of a type and design adequate for the use intended. Electrical
systems shall be installed in such a manner that they are suitably
protected from fuel, oil, water, other detrimental substances, and
mechanical damage.

\(b\) Items of electrical equipment required for a specific type of
operation are listed in other pertinent parts of the Civil Air
Regulations.

**BATTERIES**

§ 3.682 Batteries. When an item of electrical equipment which is
essential to the safe operation of the airplane is installed, the
battery required shall have sufficient capacity to supply the electrical
power necessary for dependable operation of the connected electrical
equipment.

§ 3.683 Protection against acid. If batteries are of such a type that
corrosive substance may escape during servicing or flight, means such as
a completely enclosed compartment shall be provided to prevent such
substances from coming in contact with other parts of the airplane which
are essential to safe operation. Batteries shall be accessible for
servicing and inspection on the ground.

§ 3.684 Battery vents. The battery container or compartment shall be
vented in such manner that gases released by the battery are carried
outside the airplane.

**GENERATORS**

§ 3.685 Generator. Generators shall be capable of delivering their
continuous rated power.

§ 3.686 Generator controls. Generator voltage control equipment shall be
capable of dependably regulating the generator output within rated
limits.

§ 3.687 Reverse current cut-out. A generator reverse current cut-out
shall disconnect the generator from the battery and other generators
when the generator is developing a voltage of such value that current
sufficient to cause malfunctioning can flow into the generator.

**MASTER SWITCH**

§ 3.688 Arrangement. If electrical equipment is installed, a master
switch arrangement shall be provided which will

disconnect all sources of electrical power from the main distribution
system at a point adjacent to the power sources.

§ 3.689 Master switch installation. The master switch or its controls
shall be so installed that it is easily discernible and accessible to a
member of the crew in flight.

**PROTECTIVE DEVICES**

§ 3.690 Fuses or circuit breakers. If electrical equipment is installed,
protective devices (fuses or circuit breakers) shall be installed in the
circuits to all electrical equipment, except that such items need not be
installed in the main circuits of starter motors or in other circuits
where no hazard is presented by their omission.

§ 3.691 Protective devices installation. Protective devices in circuits
essential to safety in flight shall be so located and identified that
fuses may be replaced or circuit breakers reset readily in flight.

§ 3.692 Square fuses. If fuses are used, one spare of each rating or 50
percent spare fuses of each rating, whichever is greater, shall be
provided.

**ELECTRIC CABLES**

§ 3.693 Electric cables. If electrical equipment is installed, the
connecting cables used shall be in accordance with recognized standards
for electric cable of a slow burning type and of suitable capacity.

**SWITCHES**

§ 3.694 Switches. Switches shall be capable of carrying their rated
current and shall be of such construction that there is sufficient
distance or insulating material between current carrying parts and the
housing so that vibration in flight will not cause shorting.

§ 3.695 Switch installation. Switches shall be so installed as to be
readily accessible to the appropriate crew member and shall be suitably
labeled as to operation and the circuit controlled.

**INSTRUMENT LIGHTS**

§ 3.696 Instrument lights. If instrument lights are required, they shall
be of such construction that there is sufficient distance or insulating
material between current carrying parts and the housing so that
vibration in flight will not cause shorting. They shall provide
sufficient illumination to make all instruments and controls easily
readable and discernible, respectively.

§ 3.697 Instrument light installation. Instrument lights shall be
installed in such a manner that their direct rays are shielded from the
pilot's eyes. Direct rays shall not be reflected from the windshield or
other surfaces into the pilot's eyes.

**LANDING LIGHTS**

§ 3.698 Landing lights. If landing lights are installed, they shall be
of an acceptable type.

§ 3.699 Landing light installation. Landing lights shall be so installed
that there is no dangerous glare visible to the pilot and also so that
the pilot is not seriously affected by halation. They shall be installed
at such a location that they provide adequate illumination for night
landing.

**POSITION LIGHTS**

§ 3.700 Type. If position lights are installed, they shall be of a type
certificated in accordance with Part 15 of this chapter, or shall comply
with the pertinent provisions of that part.

§ 3.701 Forward position light installation. Forward position lights
shall be so installed that, with the airplane in normal flying position,
the red light is displayed on the left side and the green light on the
right side, each showing unbroken light between two vertical planes the
dihedral angle of which is 100 degrees when measured to the left and
right, respectively, of the airplane from dead ahead. The lights shall
be spaced laterally as far apart as practicable.

§ 3.702 Rear position light installation. The rear position light shall
be mounted as far aft as practicable and so installed that unbroken
light is directed symmetrically aft in such a manner that the axis of
the maximum cone of

illumination is parallel to the flight path. In addition, the
intersection of the two planes forming dihedral angle A given in Part 15
of this chapter shall be vertical.

§ 3.703 Flashing rear position lights. If red and white flashing lights
are used, in addition to meeting the installation requirements in §
3.702, they shall be located close together.

**ANCHOR LIGHTS**

§ 3.704 Anchor light. When an anchor light is required for seaplanes and
amphibians, at least one light shall be provided and it shall be capable
of showing a white light for at least 2 miles at night under clear
atmosphere conditions.

§ 3.705 Anchor light installation. Anchor lights shall be so installed
that they will show the maximum unbroken light practicable when the
airplane is moored or drifting on the water. Externally hung lights are
permitted.

**SAFETY EQUIPMENT; INSTALLATION**

§ 3.711 Marking. Required safety equipment which the crew is expected to
operate at a time of emergency, such as flares and automatic life raft
releases, shall be readily accessible and plainly marked as to its
method of operation. When such equipment is carried in lockers,
compartments, or other storage places, such storage places shall be
marked for the benefit of passengers and crew.

§ 3.712 De-icers. When pneumatic deicers are installed, the installation
shall be in accordance with approved data. Positive means shall be
provided for the deflation of the pneumatic boots.

§ 3.713 Flare requirements. When parachute flares are required, they
shall be of a type certificated in accordance with Part 15 of this
chapter.

§ 3.714 Flare installation. Parachute flares shall be releasable from
the pilot compartment and so installed that danger of accidental
discharge is reduced to a minimum. The installation shall be
demonstrated in flight to eject flares satisfactorily, except in those
cases where inspection indicates a ground test will be adequate. If the
flares are ejected so that recoil loads are involved, structural
provisions for such loads shall be made.

§ 3.715 Safety belts. Safety belts shall be of a type certificated in
accordance with Part 15 of this chapter. They shall be so attached that
no part of the anchorage will fail at a lower load than specified in §
3.386.

**EMERGENCY FLOTATION AND SIGNALING EQUIPMENT**

§ 3.716 Rafts and life preservers. An approved life raft or approved
life preserver, when required by other parts of the Civil Air
Regulations, is one approved by either the Administrator, the Bureau of
Marine Inspection and Navigation, the United States Army Air Forces, or
the Bureau of Aeronautics, Navy Department.

§ 3.717 Installation. When such emergency equipment is required, it
shall be so installed as to be readily available to the crew and
passengers. Rafts released automatically or by the pilot shall be
attached to the airplane by

means of a line to keep them adjacent to the airplane. The strength of
the line shall be such that it will break before submerging the empty
raft.

§ 3.718 Signaling device. Signaling devices, when required by other
parts of the Civil Air Regulations, shall be accessible, function
satisfactorily, and be free from any hazard in their operation.

**RADIO EQUIPMENT; INSTALLATION**

§ 3.721 General. Radio equipment and installations in the airplane shall
be free from hazards in themselves, in their method of operation, and in
their effects on their components of the airplane.

**MISCELLANEOUS EQUIPMENT; INSTALLATION**

§ 3.725 Accessories for multiengine airplanes. Engine driven accessories
essential to the safe operation of the airplane shall be so distributed
among two or more engines that the failure of any one engine will not
impair the safe operation of the airplane by the malfunctioning of these
accessories.

**HYDRAULIC SYSTEMS**

§ 3.726 General. Hydraulic systems and elements shall be so designed as
to withstand, without exceeding the yield point, any structural loads
which might be imposed in addition to the hydraulic loads.

§ 3.727 Tests. Hydraulic systems shall be substantiated by proof
pressure tests. When proof tested, no part of the hydraulic system shall
fail, malfunction, or experience a permanent set. The proof load of any
system shall be 15

times the maximum operating pressure of that system.

§ 3.728 Accumulators. Hydraulic accumulators or pressurized reservoirs
shall not be installed on the engine side of the fire wall, except when
they form an integral part of the engine or propeller.

**SUBPART G---OPERATING LIMITATIONS AND**

**INFORMATION**

§ 3.735 General. Means shall be provided to inform adequately the pilot
and other appropriate crew members of all operating limitations upon
which the type design is based. Any other information concerning the
airplane found by the Administrator to be necessary for safety during
its operation shall also be made available to the crew. (See §§ 3.755
and 3.777.)

**LIMITATIONS**

§ 3.737 Limitations. The operating limitations specified in §§
3.738-3.750 and any similar limitations shall be established for any
airplane and made available to the operator as further described in §§
3.755-3.780, unless its

design is such that they are unnecessary for safe operation.

**AIR SPEED**

§ 3.738 Air speed. Air-speed limitations shall be established as set
forth in §§ 3.739-3.743.

§ 3.739 Never-exceed speed (Vne). This speed shall not exceed the lesser
of the following:

\(a\) 0.9 Vd chosen in accordance with § 3.184.

\(b\) 0.9 times the maximum speed demonstrated in accordance with §
3.159, but shall not be less than 0.9 times the minimum value of Vd
permitted by § 3.184.

§ 3.740 Maximum structural cruising speed (Vno). This operating
limitation shall be:

\(a\) Not greater than Vc chosen in accordance with § 3.184.

\(b\) Not greater than 0.89 times Vne established under § 3.739.

\(c\) Not less than the minimum Vc permitted in § 3.184.

§ 3.741 Maneuvering speed (Vp). (See § 3.184.)

§ 3.742 Flaps-extended speed (Vfe).

\(a\) This speed shall not exceed the lesser of the following:

\(1\) The design flap speed, Vf chosen in accordance with § 3.190.

\(2\) The design flap speed chosen in accordance with § 3.223, but shall
not be less than the minimum value of design flap speed permitted in §§
3.190 and 3.223.

\(b\) Additional combinations of flap setting, air speed, and engine
power may be established, provided the structure has been proven for the
corresponding design conditions.

§ 3.743 Minimum control speed (Vmc).(See § 3.111.)

**POWER PLANT**

§ 3.744 Power plant. The power plant limitations in §§ 3.745 through
3.747 shall be established and shall not exceed the corresponding limits
established as a part of the type certification of the engine and
propeller installed in the airplane.

§ 3.745 Take-off operation.

\(a\) Maximum rotational speed (revolutions per minute).

\(b\) Maximum permissible manifold pressure (if applicable).

\(c\) The time limit upon the use of the corresponding power.

\(d\) Where the time limit of paragraph (c) of this section exceeds 2
minutes, the maximum allowable temperatures for cylinder head, oil, and
coolant outlet if applicable.

§ 3.746 Maximum continuous operation,

\(a\) Maximum rotational speed (revolutions per minute).

\(b\) Maximum permissible manifold pressure (if applicable).

\(c\) Maximum allowable temperatures for cylinder head, oil, and coolant
outlet if applicable.

§ 3.747 Fuel octane rating. The minimum octane rating of fuel required
for satisfactory operation of the power plant at the limits of §§ 3.745
and 3.746.

**AIRPLANE WEIGHT**

§ 3.748 Airplane weight. The airplane weight and center of gravity
limitations are those required to be determined by § 3.71.

**MINIMUM FLIGHT CREW**

§ 3.749 Minimum flight crew. The minimum flight crew shall be
established as that number of persons required for the safe operation of
the airplane during any contact flight as determined by the availability
and satisfactory operation of all necessary controls by each operator
concerned.

**TYPES OF OPERATION**

§ 3.750 Types of operation. The type of operation to which the airplane
is limited shall be established by the category in which it has been
found eligible for certification and by the equipment installed. (See
Parts 42 and 43 of this chapter.)

**MARKINGS AND PLACARDS**

§ 3.755 Markings and placards.

\(a\) The markings and placards specified are required for all
airplanes. Placards shall be displayed in a conspicuous place and both
shall be such that they cannot be easily erased, disfigured, or
obscured. Additional informational placards and instrument markings
having a direct and important bearing on safe operation may be required
by the Administrator when unusual design, operating, or handling
characteristics so warrant.

\(b\) When an airplane is certificated in more than one category, the
applicant shall select one category on which all placards and markings
on the airplane shall be based. The placard and marking information for
the other categories in which the airplane is certificated shall be
entered in the Airplane Flight Manual. A reference to this information
shall be included on a placard which shall also indicate the category on
which the airplane placards and markings are based.

**INSTRUMENT MARKINGS**

§ 3.756 Instrument markings. The instruments listed in §§ 3.757-3.761
shall have the following limitations marked thereon. When these markings
are placed on the cover glass of the instrument, adequate provision
shall be made to maintain the correct alignment of the glass cover with
the face of the dial. All arcs and lines shall be of sufficient width
and so located as to be clearly and easily visible to the pilot.

§ 3.757 Air-speed indicator.

\(a\) True indicated air speed shall be used:

\(1\) The never-exceed speed, Vne---a radial red line (see § 3.739).

\(2\) The caution range---a yellow arc extending from the red line in
(1) above to the upper limit of the green arc specified in (3) below.

\(3\) The normal operating range---a green arc with the lower limit at
Vs1, as determined in § 3.82 with maximum weight, landing gear and wing
flaps retracted, and the upper limit at the maximum structural cruising
speed established in § 3.740.

\(4\) The flap operating range---a white arc with the lower limit at Vso
as determined in § 3.82 at the maximum weight, and the upper limit at
the flaps-extended speed in § 3.742.

\(b\) When the never-exceed and maximum structural cruising speeds vary
with altitude, means shall be provided which will indicate the
appropriate limitations to the pilot throughout the operating altitude
range.

§ 3.758 Magnetic direction indicator. A placard shall be installed on or
in close proximity to the magnetic direction indicator which contains
the calibration of the instrument in a level flight attitude with
engine(s) operating and radio receiver(s) on or off (which shall be
stated). The calibration readings shall be those to known magnetic
headings in not greater than 30-degree increments.

§ 3.759 Power-plant instruments. All required power-plant instruments
shall be marked with a red radial line at the maximum and minimum (if
applicable) indications for safe operation. The normal operating ranges
shall be marked with a green arc which shall not extend beyond the
maximum and minimum limits for continuous operation. Take-off and
precautionary ranges shall be marred with a yellow arc.

§ 3.760 Oil quantity indicators. Indicators shall be suitably marked in
sufficient increments so that they will readily and accurately indicate
the quantity of oil.

§ 3.761 Fuel quantity indicator. When the unusable fuel supply for any
tank exceeds 1 gallon or 5 percent of the tank capacity, whichever is
greater, a red band shall be placed on the indicator extending from the
calibrated zero reading (see § 3.437) to the lowest reading obtainable
in the level flight attitude, and a suitable notation in the Airplane
Flight Manual shall be provided to indicate the flight personnel that
the fuel remaining in the tank when the quantity indicator reaches zero
cannot be used safely in flight. (See § 3.672.)

**CONTROL MARKINGS**

§ 3.762 General. All cockpit controls, with the exception of the primary
flight controls, shall be plainly marked as to their function and method
of operation.

§ 3.763 Aerodynamic controls. The secondary controls shall be suitably
marked to comply with §§ 3.337 and 3.338.

§ 3.764 Power-plant fuel controls.

\(a\) Controls for fuel tank selector valves shall be marked to indicate
the position corresponding to each tank and to all existing cross feed
positions.

\(b\) When more than one fuel tank is provided, and if safe operation
depends upon the use of tanks in a specific sequence, the fuel tank
selector controls shall be marked adjacent to or on the control to
indicate to the flight personnel the order in which the tanks must be
used.

\(c\) On multiengine airplanes, controls for engine valves shall be
marked to indicate the position corresponding to each engine.

\(d\) The capacity of each tank shall be indicated adjacent to or on the
fuel tank selector control.

§ 3.765 Accessory and auxiliary controls.

\(a\) When a retractable landing gear is used, the indicator required in
§ 3.359 shall be marked in such a manner that the pilot can ascertain at
all times when the wheels are secured in the extreme positions.

\(b\) Emergency controls shall be colored red and clearly marked as to
their method of operation.

**MISCELLANEOUS**

§ 3.766 Baggage compartments, ballast location, and special seat loading
limitations.

\(a\) Each baggage or cargo compartment and ballast location shall bear
a placard which states the maximum allowable weight of contents and, if
applicable, any special limitation of contents due to loading
requirements, etc.

\(b\) When the maximum permissible weight to be carried in a seat is
less than 170 pounds (see § 3.74), a placard shall be permanently
attached to the seat structure which states the maximum allowable weight
of occupants to be carried.

§ 3.767 Fuel, oil, and coolant filler openings. The following
information shall be marked on or adjacent to the filler cover in each
case:

\(a\) The word \"fuel,\" the minimum permissible fuel octane number for
the engines installed, and the usable fuel tank capacity. (See § 3.437.)

\(b\) The word \"oil\" and the oil tank capacity.

\(c\) The name of the proper coolant fluid and the capacity of the
coolant system.

§ 3.768 Emergency exit placards. Emergency exit placards and operating
controls shall be colored red. A placard shall be located adjacent to
the control(s) which clearly indicates it to be an emergency exit and
describes the

method of operation. (See § 3.387.)

§ 3.769 Approved flight maneuvers---

\(a\) Category N. A placard shall be provided in front of and in clear
view of the pilot stating: \"No acrobatic maneuvers including spins
approved.\"

\(b\) Category U. A placard shall be provided in front of and in clear
view of the pilot stating: \"No acrobatic maneuvers approved, except
those listed in the Airplane Flight Manual.\"

\(c\) Category A. A placard shall be provided in clear view of the pilot
which lists all approved acrobatic maneuvers and the recommended entry
air speed for each. If inverted flight maneuvers are not approved, the
placard shall bear a notation to this effect.

§ 3.770 Airplane category placard. A placard shall be provided in front
of and in clear view of the pilot stating: \"This airplane must be
operated as a \-\-\-\-\-\-\-\-\-\-\-\-\-\-\-\-\-- or
\-\-\-\-\-\-\-\-\-\-\-\-\-\-\-- category airplane in compliance with the
Airplane Flight Manual.\"

**AIRPLANE FLIGHT MANUAL**

§ 3.777 Airplane Flight Manual. An Airplane Flight Manual shall be
furnished with each airplane. The portions of this document listed below
shall be verified and approved by the Administrator, and shall be
segregated, identified, and clearly distinguished from portions not so
approved. Additional items of information having a direct and important
bearing on safe operation may be required by the Administrator when
unusual design, operating, or handling characteristics so warrant.

§ 3.778 Operating limitations---

\(a\) Airspeed limitations. Sufficient information shall be included to
permit proper marking of the airspeed

limitations on the indicator as required in § 3.757. It shall also
include the design, maneuvering speed, and the maximum safe air speed at
which the landing gear can be safely lowered. In addition to the above
information, the significance of the air speed limitations and of the
color coding used shall be explained.

\(b\) Power-plant limitations. Sufficient information shall be included
to outline and explain all power-plant limitations (see § 3.744) and to
permit marking the instruments as required in § 3.759.

\(c\) Weight. The following information shall be included:

\(1\) Maximum weight for which the airplane has been certificated,

\(2\) Airplane empty weight and center of gravity location,

\(3\) Useful load,

\(4\) The composition of the useful load, including the total weight of
fuel and oil with tanks full.

\(d\) Load distribution.

\(1\) All authorized center of gravity limits shall be stated. If the
available space for loading the airplane is adequately placarded or so
arranged that any reasonable distribution of the useful load listed in
weight above will not result in a center of gravity location outside of
the stated limits, this section need not include any other information
than the statement of center of gravity limits.

\(2\) In all other cases this section shall also include adequate
information to indicate satisfactory loading combinations which will
assure maintaining the center of gravity position within approved
limits.

\(e\) Maneuvers. All authorized maneuvers and the appropriate air-speed
limitations as well as all unauthorized maneuvers shall be included in
accordance with the following:

\(1\) Normal category. All acrobatic maneuvers, including spins, are
unauthorized. If the airplane has been demonstrated to be
characteristically incapable of spinning in accordance with § 3.124 (d),
a statement to this

effect shall be entered here.

\(2\) Utility category. All authorized maneuvers demonstrated in the
type flight tests shall be listed, together with recommended entry
speeds. All other maneuvers are not approved. If the airplane has been
demonstrated to be characteristically incapable of spinning in
accordance with § 3.124 (d), a statement to this effect shall be entered
here.

\(3\) Acrobatic category. All approved flight maneuvers demonstrated in
the type flight tests shall be included, together with recommended entry
speeds.

\(f\) Flight load factor. The positive limit load factors made good by
the airplane's structure shall be described here in terms of
accelerations.

\(g\) Flight crew. When a flight crew of more than one is required to
operate the airplane safely, the number and functions of this minimum
flight crew shall be included.

§ 3.779 Operating procedures. This section shall contain information
concerning normal and emergency procedures and other pertinent
information peculiar to the airplane's operating characteristics which
are necessary to

safe operation.

§ 3.780 Performance information.

\(a\) Information relative to the following items of performance shall
be included:

\(1\) The stalling speed, Vso, at maximum weight,

\(2\) The stalling speed, Vs1, at maximum weight and with landing gear
and wing flaps retracted,

\(3\) The take-off distance determined in accordance with § 3.84,
including the air speed at the 50-foot height, and the airplane
configuration, if pertinent,

\(4\) The landing distance determined in accordance with § 3.86,
including the airplane configuration, if pertinent,

\(5\) The steady rate of climb determined in accordance with § 3.85 (a),
(c), and, as appropriate, (b), including the air speed, power, and
airplane configuration, if pertinent.

\(b\) The effect of variation in (a) (2) with angle of bank up to 60
degrees shall be included.

\(c\) The calculated approximate effect of variations in subparagraphs
(3), (4) and (5) of this paragraph with altitude and temperature shall
be included.

**SUBPART H---IDENTIFICATION DATA**

§ 3.791 Name plate. A name plate shall be securely attached to and
located in the pilot compartment which shall contain:

\(a\) The manufacturer's name and address.

\(b\) Model and serial numbers.

\(c\) Date of manufacture.

\(d\) Type certificate number.

\(e\) Production certificate number, (if pertinent).

§ 3.792 Airworthiness certificate number. The identifying symbols and
registration numbers shall be permanently affixed to the airplane
structure in compliance with § 43.10 (c) of this chapter.