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FAA FAR Part 25 D
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Subpart D--Design and
Construction
General
25.601 General.
25.603 Materials.
25.605 Fabrication methods.
25.607 Fasteners.
25.609 Protection of structure.
25.611 Accessibility provisions.
25.613 Material strength properties and design values.
25.619 Special factors.
25.621 Casting factors.
25.623 Bearing factors.
25.625 Fitting factors.
25.629 Aeroelastic stability requirements.
25.631 Bird strike damage.
Control
Surfaces
25.651 Proof of strength.
25.655 Installation.
25.657 Hinges.
Control Systems
25.671 General.
25.672 Stability augmentation and automatic and
power-operated systems.
25.675 Stops.
25.677 Trim systems.
25.679 Control system gust locks.
25.681 Limit load static tests.
25.683 Operation tests.
25.685 Control system details.
25.689 Cable systems.
25.693 Joints.
25.697 Lift and drag devices, controls.
25.699 Lift and drag device indicator.
25.701 Flap interconnection.
25.703 Takeoff warning system.
Landing Gear
25.721 General.
25.723 Shock absorption tests.
25.725 Limit drop tests.
25.727 Reserve energy absorption drop tests.
25.729 Retracting mechanism.
25.731 Wheels.
25.733 Tires
25.735 Brakes.
25.737 Skis.
Floats and
Hulls
25.751 Main float buoyancy.
25.753 Main float design.
25.755 Hulls.
Personnel
and Cargo Accommodations
25.771 Pilot compartment.
25.772 Pilot compartment doors.
25.773 Pilot compartment view.
25.775 Windshields and windows.
25.777 Cockpit controls.
25.779 Motion and effect of cockpit controls.
25.781 Cockpit control knob shape.
25.783 Doors.
25.785 Seats, berths, safety belts, and harnesses.
25.787 Stowage compartments.
25.789 Retention of items of mass in passenger and crew
compartments and galleys.
25.791 Passenger information signs.
25.793 Floor surfaces.
Emergency
Provisions
25.801 Ditching.
25.803 Emergency evacuation.
25.807 Passenger emergency exits.
25.809 Emergency exit arrangement.
25.810 Emergency egress assist means and escape routes.
25.811 Emergency exit marking.
25.812 Emergency lighting.
25.813 Emergency exit access.
25.815 Width of aisle.
25.817 Maximum number of seats abreast.
25.819 Lower deck surface compartments (including
galleys).
Ventilation and
Heating
25.831 Ventilation.
25.832 Cabin ozone concentration.
25.833 Heating systems.
Pressurization
25.841 Pressurized cabins.
25.843 Tests for pressurized cabins.
Fire Protection
25.851 Fire extinguishers.
25.853 Compartment interiors.
25.854 Lavatory fire protection.
25.855 Cargo and baggage compartments.
25.857 Cargo compartment classification.
25.858 Cargo compartment fire detection systems.
25.859 Combustion heater fire protection.
25.863 Flammable fluid fire protection.
25.865 Fire protection of flight controls, engine mounts,
and other flight structure.
25.867 Fire protection: other components.
25.869 Fire protection: systems.
Miscellaneous
25.871 Leveling means.
25.875 Reinforcement near propellers.
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General:
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Sec. 25.601 General.
The airplane may not have design features or details that
experience has shown to be hazardous or unreliable. The
suitability of each questionable design detail and part must
be established by tests.
Sec. 25.603 Materials.
The suitability and durability of materials used for
parts, the failure of which could adversely affect safety,
must--
(a) Be established on the basis of experience or
tests;
(b) Conform to approved specifications (such as industry
or military specifications, or Technical Standard Orders)
that ensure their having the strength and other properties
assumed in the design data; and
(c) Take into account the effects of environmental
conditions, such as temperature and humidity, expected in
service.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-38, 41 FR 55466, Dec. 20 1976; Amdt.
25-46, 43 FR 50595, Oct. 30, 1978]
Sec. 25.605 Fabrication methods.
(a) The methods of fabrication used must produce a
consistently sound structure. If a fabrication process (such
as gluing, spot welding, or heat treating) requires close
control to reach this objective, the process must be
performed under an approved process specification.
(b) Each new aircraft fabrication method must be
substantiated by a test program.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-46, 43 FR 50595, Oct. 30, 1978]
Sec. 25.607 Fasteners.
(a) Each removable bolt, screw, nut, pin, or other
removable fastener must incorporate two separate locking
devices if--
(1) Its loss could preclude continued flight and landing
within the design limitations of the airplane using normal
pilot skill and strength; or
(2) Its loss could result in reduction in pitch, yaw, or
roll control capability or response below that required by
Subpart B of this chapter.
(b) The fasteners specified in paragraph (a) of this
section and their locking devices may not be adversely
affected by the environmental conditions associated with the
particular installation.
(c) No self-locking nut may be used on any bolt subject
to rotation in operation unless a nonfriction locking device
is used in addition to the self-locking device.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 25.609 Protection of structure.
Each part of the structure must--
(a) Be suitably protected against deterioration or loss
of strength in service due to any cause, including--
(1) Weathering;
(2) Corrosion; and
(3) Abrasion; and
(b) Have provisions for ventilation and drainage where
necessary for protection.
Sec. 25.611 Accessibility provisions.
Means must be provided to allow inspection (including
inspection of principal structural elements and control
systems), replacement of parts normally requiring
replacement, adjustment, and lubrication as necessary for
continued airworthiness. The inspection means for each item
must be practicable for the inspection interval for the
item. Nondestructive inspection aids may be used to inspect
structural elements where it is impracticable to provide
means for direct visual inspection if it is shown that the
inspection is effective and the inspection procedures are
specified in the maintenance manual required by Sec.
25.1529.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 25.613 Material strength properties and design
values.
(a) Material strength properties must be based on enough
tests of material meeting approved specifications to
establish design values on a statistical basis.
(b) Design values must be chosen to minimize the
probability of structural failures due to material
variability. Except as provided in paragraph (e) of this
section, compliance with this paragraph must be shown by
selecting design values which assure material strength with
the following probability:
(1) Where applied loads are eventually distributed
through a single member within an assembly, the failure of
which would result in loss of structural integrity of the
component, 99 percent probability with 95 percent
confidence.
(2) For redundant structure, in which the failure of
individual elements would result in applied loads being
safely distributed to other load carrying members, 90
percent probability with 95 percent confidence.
(c) The effects of temperature on allowable stresses used
for design in an essential component or structure must be
considered where thermal effects are significant under
normal operating conditions.
(d) The strength, detail design, and fabrication of the
structure must minimize the probability of disastrous
fatigue failure, particularly at points of stress
concentration.
(e) Greater design values may be used if a "premium
selection" of the material is made in which a specimen of
each individual item is tested before use to determine that
the actual strength properties of that particular item will
equal or exceed those used in design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt.
25-72, 55 FR 29776, July 20, 1990]
Sec. 25.615 [Removed. Amdt. 25-72, 55 FR
29776, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the
removed text is set out below.
Sec. 25.615 Design properties.
(a) Design properties outlined in MIL-HDBK-5 may be used
subject to the following conditions:
(1) Where applied loads are eventually distributed
through a single member within an assembly, the failure of
which would result in the loss of the structural integrity
of the component involved, the guaranteed minimum design
mechanical properties ("A" values) when listed in MIL-HDBK-5
must be met.
(2) Redundant structures, in which the failure of
individual elements would result in applied loads being
safely distributed to other load-carrying members, may be
designed on the basis of the "90 percent probability ("B"
values)" when listed in MIL-HDBK-5.
(b) Design values greater than the guaranteed minimums
required by paragraph (a) of this section may be used where
only guaranteed minimum values are normally allowed if a
"premium selection" of the material is made in which a
specimen of each individual item is tested before use to
determine that the actual strength properties of that
particular item will equal or exceed those used in
design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5674. Apr. 8, 1970]
Sec. 25.619 Special factors.
The factor of safety prescribed in Sec. 25.303 must be
multiplied by the highest pertinent special factor of safety
prescribed in Secs. 25.621 through 25.625 for each part of
the structure whose strength is--
(a) Uncertain;
(b) Likely to deteriorate in service before normal
replacement; or (c) Subject to appreciable variability
because of uncertainties in manufacturing processes or
inspection methods.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 25.621 Casting factors.
(a) General. The factors, tests, and inspections
specified in paragraphs (b) through (d) of this section must
be applied in addition to those necessary to establish
foundry quality control. The inspections must meet approved
specifications. Paragraphs (c) and (d) of this section apply
to any structural castings except castings that are pressure
tested as parts of hydraulic or other fluid systems and do
not support structural loads.
(b) Bearing stresses and surfaces. The casting factors
specified in paragraphs (c) and (d) of this section--
(1) Need not exceed 1.25 with respect to bearing stresses
regardless of the method of inspection used; and
(2) Need not be used with respect to the bearing surfaces
of a part whose bearing factor is larger than the applicable
casting factor.
(c) Critical castings. For each casting whose failure
would preclude continued safe flight and landing of the
airplane or result in serious injury to occupants, the
following apply:
(1) Each critical casting must--
(i) Have a casting factor of not less than 1.25; and
(ii) Receive 100 percent inspection by visual,
radiographic, and magnetic particle or penetrant inspection
methods or approved equivalent nondestructive inspection
methods.
(2) For each critical casting with a casting factor less
than 1.50, three sample castings must be static tested and
shown to meet--
(i) The strength requirements of Sec. 25.305 at an
ultimate load corresponding to a casting factor of 1.25;
and
(ii) The deformation requirements of Sec. 25.305 at a
load of 1.15 times the limit load.
(3) Examples of these castings are structural attachment
fittings, parts of flight control systems, control surface
hinges and balance weight attachments, seat, berth, safety
belt, and fuel and oil tank supports and attachments, and
cabin pressure valves.
(d) Noncritical castings. For each casting other than
those specified in paragraph (c) of this section, the
following apply:
(1) Except as provided in paragraphs (d) (2) and (3) of
this section, the casting factors and corresponding
inspections must meet the following table:
Casting factor -> Inspection
2.0 or more -> 100 percent visual.
Less than 2.0 but more than 1.5 -> 100 percent visual,
and magnetic particle or penetrant or equivalent
nondestructive inspection methods.
1.25 through 1.50 -> 100 percent visual, magnetic
particle or penetrant, and radiographic, or approved
equivalent nondestructive inspection methods.
(2) The percentage of castings inspected by nonvisual
methods may be reduced below that specified in paragraph
(d)(1) of this section when an approved quality control
procedure is established.
(3) For castings procured to a specification that
guarantees the mechanical properties of the material in the
casting and provides for demonstration of these properties
by test of coupons cut from the castings on a sampling
basis--
(i) A casting factor of 1.0 may be used; and
(ii) The castings must be inspected as provided in
paragraph (d)(1) of this section for casting factors of
"1.25 through 1.50" and tested under paragraph (c)(2) of
this section.
Sec. 25.623 Bearing factors.
(a) Except as provided in paragraph (b) of this section,
each part that has clearance (free fit), and that is subject
to pounding or vibration, must have abearing factor large
enough to provide for the effects of normal relative
motion.
(b) No bearing factor need be used for a part for which
any larger special factor is prescribed.
Sec. 25.625 Fitting factors.
For each fitting (a part or terminal used to join one
structural member to another), the following apply:
(a) For each fitting whose strength is not proven by
limit and ultimate load tests in which actual stress
conditions are simulated in the fitting and surrounding
structures, a fitting factor of at least 1.15 must be
applied to each part of--
(1) The fitting;
(2) The means of attachment; and
(3) The bearing on the joined members.
(b) No fitting factor need be used--
(1) For joints made under approved practices and based on
comprehensive test data (such as continuous joints in metal
plating, welded joints, and scarf joints in wood); or
(2) With respect to any bearing surface for which a
larger special factor is used.
(c) For each integral fitting, the part must be treated
as a fitting up to the point at which the section properties
become typical of the member. (d) For each seat, berth,
safety belt, and harness, the fitting factor specified in
Sec. 25.785(f)(3) applies.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5674, Apr. 8, 1970; Amdt.
25-72, 55 FR 29776, July 20, 1990]
Sec. 25.629 Aeroelastic stability
requirements.
(a) General. The aeroelastic stability evaluations
required under this section include flutter, divergence,
control reversal and any undue loss of stability and control
as a result of structural deformation. The aeroelastic
evaluation must include whirl modes associated with any
propeller or rotating device that contributes significant
dynamic forces. Compliance with this section must be shown
by analyses, wind tunnel tests, ground vibration tests,
flight tests, or other means found necessary by the
Administrator.
(b) Aeroelastic stability envelopes. The airplane must be
designed to be free from aeroelastic instability for all
configurations and design conditions within the aeroelastic
stability envelopes as follows:
(1) For normal conditions without failures, malfunctions,
or adverse conditions, all combinations of altitudes and
speeds encompassed by the VD/MD versus altitude envelope
enlarged at all points by an increase of 15 percent in
equivalent airspeed at both constant Mach number and
constant altitude. In addition, a proper margin of stability
must exist at all speeds up to VD/MD and, there must be no
large and rapid reduction in stability as VD/MD is
approached. The enlarged envelope may be limited to Mach 1.0
when MD is less than 1.0 at all design altitudes, and
(2) For the conditions described in Sec. 25.629(d) below,
for all approved altitudes, any airspeed up to the greater
airspeed defined by;
(i) The VD/MD envelope determined by Sec. 25.335(b);
or,
(ii) An altitude-airspeed envelope defined by a 15
percent increase in equivalent airspeed above VC at constant
altitude, from sea level to the altitude of the intersection
of 1.15 VC with the extension of the constant cruise Mach
number line, MC, then a linear variation in equivalent
airspeed to MC+.05 at the altitude of the lowest VC/MC
intersection; then, at higher altitudes, up to the maximum
flight altitude, the boundary defined by a .05 Mach increase
in MC at constant altitude.
(c) Balance weights. If concentrated balance weights are
used, their effectiveness and strength, including supporting
structure, must be substantiated.
(d) Failures, malfunctions, and adverse conditions. The
failures, malfunctions, and adverse conditions which must be
considered in showing compliance with this section are:
(1) Any critical fuel loading conditions, not shown to be
extremely improbable, which may result from mismanagement of
fuel.
(2) Any single failure in any flutter damper system.
(3) For airplanes not approved for operation in icing
conditions, the maximum likely ice accumulation expected as
a result of an inadvertent encounter.
(4) Failure of any single element of the structure
supporting any engine, independently mounted propeller
shaft, large auxiliary power unit, or large externally
mounted aerodynamic body (such as an external fuel
tank).
(5) For airplanes with engines that have propellers or
large rotating devices capable of significant dynamic
forces, any single failure of the engine structure that
would reduce the rigidity of the rotational axis.
(6) The absence of aerodynamic or gyroscopic forces
resulting from the most adverse combination of feathered
propellers or other rotating devices capable of significant
dynamic forces. In addition, the effect of a single
feathered propeller or rotating device must be coupled with
the failures of paragraphs (d)(4) and (d)(5) of this
section.
(7) Any single propeller or rotating device capable of
significant dynamic forces rotating at the highest likely
overspeed.
(8) Any damage or failure condition, required or selected
for investigation by Sec. 25.571. The single structural
failures described in paragraphs (d)(4) and (d)(5) of this
section need not be considered in showing compliance with
this section if;
(i) The structural element could not fail due to discrete
source damage resulting from the conditions described in
Sec. 25.571(e), and
(ii) A damage tolerance investigation in accordance with
Sec. 25.571(b) shows that the maximum extent of damage
assumed for the purpose of residual strength evaluation does
not involve complete failure of the structural element.
(9) Any damage, failure, or malfunction considered under
Secs. 25.631, 25.671, 25.672, and 25.1309.
(10) Any other combination of failures, malfunctions, or
adverse conditions not shown to be extremely improbable.
(e) Flight flutter testing. Full scale flight flutter
tests at speeds up to VDF/MDF must be conducted for new type
designs and for modifications to a type design unless the
modifications have been shown to have an insignificant
effect on the aeroelastic stability. These tests must
demonstrate that the airplane has a proper margin of damping
at all speeds up to VDF/MDF, and that there is no large and
rapid reduction in damping as VDF/MDF, is approached. If a
failure, malfunction, or adverse condition is simulated
during flight test in showing compliance with paragraph (d)
of this section, the maximum speed investigated need not
exceed VFC/MFC if it is shown, by correlation of the flight
test data with other test data or analyses, that the
airplane is free from any aeroelastic instability at all
speeds within the altitude airspeed envelope described in
paragraph (b)(2) of this section.
[57 FR 28949, June 29, 1992]
Sec. 25.631 Bird strike damage.
The empennage structure must be designed to assure
capability of continued safe flight and landing of the
airplane after impact with an 8-pound bird when the velocity
of the airplane (relative to the bird along the airplane's
flight path) is equal to VC at sea level, selected under
Sec. 25.335(a). Compliance with this section by provision of
redundant structure and protected location of control system
elements or protective devices such as splitter plates or
energy absorbing material is acceptable. Where compliance is
shown by analysis, tests, or both, use of data on airplanes
having similar structural design is acceptable.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
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Control
Surfaces:
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Sec. 25.651 Proof of strength.
(a) Limit load tests of control surfaces are required.
These tests must include the horn or fitting to which the
control system is attached.
(b) Compliance with the special factors requirements of
Secs. 25.619 through 25.625 and 25.657 for control surface
hinges must be shown by analysis or individual load
tests.
Sec. 25.655 Installation.
(a) Movable tail surfaces must be installed so that there
is no interference between any surfaces when one is held in
its extreme position and the others are operated through
their full angular movement.
(b) If an adjustable stabilizer is used, it must have
stops that will limit its range of travel to the maximum for
which the airplane is shown to meet the trim requirements of
Sec. 25.161.
Sec. 25.657 Hinges.
(a) For control surface hinges, including ball, roller,
and self-lubricated bearing hinges, the approved rating of
the bearing may not be exceeded. For nonstandard bearing
hinge configurations, the rating must be established on the
basis of experience or tests and, in the absence of a
rational investigation, a factor of safety of not less than
6.67 must be used with respect to the ultimate bearing
strength of the softest material used as a bearing.
(b) Hinges must have enough strength and rigidity for
loads parallel to the hinge line.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
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Control
Systems:
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Sec. 25.671 General.
(a) Each control and control system must operate with the
ease, smoothness, and positiveness appropriate to its
function.
(b) Each element of each flight control system must be
designed, or distinctively and permanently marked, to
minimize the probability of incorrect assembly that could
result in the malfunctioning of the system.
(c) The airplane must be shown by analysis, tests, or
both, to be capable of continued safe flight and landing
after any of the following failures or jamming in the flight
control system and surfaces (including trim, lift, drag, and
feel systems), within the normal flight envelope, without
requiring exceptional piloting skill or strength. Probable
malfunctions must have only minor effects on control system
operation and must be capable of being readily counteracted
by the pilot.
(1) Any single failure, excluding jamming (for example,
disconnection or failure of mechanical elements, or
structural failure of hydraulic components, such as
actuators, control spool housing, and valves).
(2) Any combination of failures not shown to be extremely
improbable, excluding jamming (for example, dual electrical
or hydraulic system failures, or any single failure in
combination with any probable hydraulic or electrical
failure).
(3) Any jam in a control position normally encountered
during takeoff, climb, cruise, normal turns, descent, and
landing unless the jam is shown to be extremely improbable,
or can be alleviated. A runaway of a flight control to an
adverse position and jam must be accounted for if such
runaway and subsequent jamming is not extremely
improbable.
(d) The airplane must be designed so that it is
controllable if all engines fail. Compliance with this
requirement may be shown by analysis where that method has
been shown to be reliable.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 25.672 Stability augmentation and automatic and
power-operated systems.
If the functioning of stability augmentation or other
automatic or power-operated systems is necessary to show
compliance with the flight characteristics requirements of
this part, such systems must comply with Sec. 25.671 and the
following:
(a) A warning which is clearly distinguishable to the
pilot under expected flight conditions without requiring his
attention must be provided for any failure in the stability
augmentation system or in any other automatic or
power-operated system which could result in an unsafe
condition if the pilot were not aware of the failure.
Warning systems must not activate the control systems.
(b) The design of the stability augmentation system or of
any other automatic or power-operated system must permit
initial counteraction of failures of the type specified in
Sec. 25.671(c) without requiring exceptional pilot skill or
strength, by either the deactivation of the system, or a
failed portion thereof, or by overriding the failure by
movement of the flight controls in the normal sense.
(c) It must be shown that after any single failure of the
stability augmentation system or any other automatic or
power-operated system--
(1) The airplane is safely controllable when the failure
or malfunction occurs at any speed or altitude within the
approved operating limitations that is critical for the type
of failure being considered;
(2) The controllability and maneuverability requirements
of this part are met within a practical operational flight
envelope (for example, speed, altitude, normal acceleration,
and airplane configurations) which is described in the
Airplane Flight Manual; and
(3) The trim, stability, and stall characteristics are
not impaired below a level needed to permit continued safe
flight and landing.
[Amdt. 25-23, 35 FR 5675 Apr. 8, 1970]
Sec. 25.673 [Removed. Amdt. 25-72, 55 FR
29777, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the
removed text is set
out below.
Sec. 25.673 Two-control airplanes.
Two-control airplanes must be able to continue safely in
flight and landing if any one connecting element in the
directional-lateral flight control system fails.
Sec. 25.675 Stops.
(a) Each control system must have stops that positively
limit the range of motion of each movable aerodynamic
surface controlled by the system.
(b) Each stop must be located so that wear, slackness, or
take-up adjustments will not adversely affect the control
characteristics of the airplane because of a change in the
range of surface travel.
(c) Each stop must be able to withstand any loads
corresponding to the design conditions for the control
system.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
Sec. 25.677 Trim systems.
(a) Trim controls must be designed to prevent inadvertent
or abrupt operation and to operate in the plane, and with
the sense of motion, of the airplane.
(b) There must be means adjacent to the trim control to
indicate the direction of the control movement relative to
the airplane motion. In addition, there must be clearly
visible means to indicate the position of the trim device
with respect to the range of adjustment.
(c) Trim control systems must be designed to prevent
creeping in flight. Trim tab controls must be irreversible
unless the tab is appropriately balanced and shown to be
free from flutter.
(d) If an irreversible tab control system is used, the
part from the tab to the attachment of the irreversible unit
to the airplane structure must consist of a rigid
connection.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.679 Control system gust locks.
(a) There must be a device to prevent damage to the
control surfaces (including tabs), and to the control
system, from gusts striking the airplane while it is on the
ground or water. If the device, when engaged, prevents
normal operation of the control surfaces by the pilot, it
must--
(1) Automatically disengage when the pilot operates the
primary flight controls in a normal manner; or
(2) Limit the operation of the airplane so that the pilot
receives unmistakable warning at the start of takeoff.
(b) The device must have means to preclude the
possibility of it becoming inadvertently engaged in
flight.
Sec. 25.681 Limit load static tests.
(a) Compliance with the limit load requirements of this
Part must be shown by tests in which--
(1) The direction of the test loads produces the most
severe loading in the control system; and
(2) Each fitting, pulley, and bracket used in attaching
the system to the main structure is included.
(b) Compliance must be shown (by analyses or individual
load tests) with the special factor requirements for control
system joints subject to angular motion.
Sec. 25.683 Operation tests.
It must be shown by operation tests that when portions of
the control system subject to pilot effort loads are loaded
to 80 percent of the limit load specified for the system and
the powered portions of the control system are loaded to the
maximum load expected in normal operation, the system is
free from--
(a) Jamming;
(b) Excessive friction; and
(c) Excessive deflection.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.685 Control system details.
(a) Each detail of each control system must be designed
and installed to prevent jamming, chafing, and interference
from cargo, passengers, loose objects, or the freezing of
moisture.
(b) There must be means in the cockpit to prevent the
entry of foreign objects into places where they would jam
the system.
(c) There must be means to prevent the slapping of cables
or tubes against other parts.
(d) Sections 25.689 and 25.693 apply to cable systems and
joints.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
Sec. 25.689 Cable systems.
(a) Each cable, cable fitting, turnbuckle, splice, and
pulley must be approved. In addition--
(1) No cable smaller than 1/8 inch in diameter may be
used in the aileron, elevator, or rudder systems; and
(2) Each cable system must be designed so that there will
be no hazardous change in cable tension throughout the range
of travel under operating conditions and temperature
variations.
(b) Each kind and size of pulley must correspond to the
cable with which it is used. Pulleys and sprockets must have
closely fitted guards to prevent the cables and chains from
being displaced or fouled. Each pulley must lie in the plane
passing through the cable so that the cable does not rub
against the pulley flange.
(c) Fairleads must be installed so that they do not cause
a change in cable direction of more than three degrees.
(d) Clevis pins subject to load or motion and retained
only by cotter pins may not be used in the control
system.
(e) Turnbuckles must be attached to parts having angular
motion in a manner that will positively prevent binding
throughout the range of travel.
(f) There must be provisions for visual inspection of
fairleads, pulleys, terminals, and turnbuckles.
Sec. 25.693 Joints.
Control system joints (in push-pull systems) that are
subject to angular motion, except those in ball and roller
bearing systems, must have a special 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 joints in cable control systems. For
ball or roller bearings, the approved ratings may not be
exceeded.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29777, July 20,
1990]
Sec. 25.697 Lift and drag devices, controls.
(a) Each lift device control must be designed so that the
pilots can place the device in any takeoff, en route,
approach, or landing position established under Sec.
25.101(d). Lift and drag devices must maintain the selected
positions, except for movement produced by an automatic
positioning or load limiting device, without further
attention by the pilots.
(b) Each lift and drag device control must be designed
and located to make inadvertent operation improbable. Lift
and drag devices intended for ground operation only must
have means to prevent the inadvertant operation of their
controls in flight if that operation could be hazardous.
(c) The rate of motion of the surfaces in response to the
operation of the control and the characteristics of the
automatic positioning or load limiting device must give
satisfactory flight and performance characteristics under
steady or changing conditions of airspeed, engine power, and
airplane attitude.
(d) The lift device control must be designed to retract
the surfaces from the fully extended position, during steady
flight at maximum continuous engine power at any speed below
VF +9.0 (knots).
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970, as amended by
Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-57, 49 FR
6848, Feb. 23, 1984]
Sec. 25.699 Lift and drag device indicator.
(a) There must be means to indicate to the pilots the
position of each lift or drag device having a separate
control in the cockpit to adjust its position. In addition,
an indication of unsymmetrical operation or other
malfunction in the lift or drag device systems must be
provided when such indication is necessary to enable the
pilots to prevent or counteract an unsafe flight or ground
condition, considering the effects on flight characteristics
and performance.
(b) There must be means to indicate to the pilots the
takeoff, en route, approach, and landing lift device
positions.
(c) If any extension of the lift and drag devices beyond
the landing position is possible, the controls must be
clearly marked to identify this range of extension.
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.701 Flap and slat interconnection.
(a) Unless the airplane has safe flight characteristics
with the flaps or slats retracted on one side and extended
on the other, the motion of flaps or slats on opposite sides
of the plane of symmetry must be synchronized by a
mechanical interconnection or approved equivalent means.
(b) If a wing flap or slat interconnection or equivalent
means is used, it must be designed to account for the
applicable unsymmetrical loads, including those resulting
from flight with the engines on one side of the plane of
symmetry inoperative and the remaining engines at takeoff
power.
(c) For airplanes with flaps or slats that are not
subjected to slipstream conditions, the structure must be
designed for the loads imposed when the wing flaps or slats
on one side are carrying the most severe load occurring in
the prescribed symmetrical conditions and those on the other
side are carrying not more than 80 percent of that load.
(d) The interconnection must be designed for the loads
resulting when interconnected flap or slat surfaces on one
side of the plane of symmetry are jammed and immovable while
the surfaces on the other side are free to move and the full
power of the surface actuating system is applied.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29777, July 20,
1990]
Sec. 25.703 Takeoff warning system.
A takeoff warning system must be installed and must meet
the following requirements:
(a) The system must provide to the pilots an aural
warning that is automatically activated during the initial
portion of the takeoff roll if the airplane is in a
configuration, including any of the following, that would
not allow a safe takeoff:
(1) The wing flaps or leading edge devices are not within
the approved range of takeoff positions.
(2) Wing spoilers (except lateral control spoilers
meeting the requirements of Sec. 25.671), speed brakes, or
longitudinal trim devices are in a position that would not
allow a safe takeoff.
(b) The warning required by paragraph (a) of this section
must continue until--
(1) The configuration is changed to allow a safe
takeoff;
(2) Action is taken by the pilot to terminate the takeoff
roll;
(3) The airplane is rotated for takeoff; or
(4) The warning is manually deactivated by the pilot.
(c) The means used to activate the system must function
properly throughout the ranges of takeoff weights,
altitudes, and temperatures for which certification is
requested.
[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]
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Landing
Gear:
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Sec. 25.721 General.
(a) The main landing gear system must be designed so that
if it fails due to overloads during takeoff and landing
(assuming the overloads to act in the upward and aft
directions), the failure mode is not likely to cause--
(1) For airplanes that have passenger seating
configuration, excluding pilots seats, of nine seats or
less, the spillage of enough fuel from any fuel system in
the fuselage to constitute a fire hazard; and
(2) For airplanes that have a passenger seating
configuration, excluding pilots seats, of 10 seats or more,
the spillage of enough fuel from any part of the fuel system
to constitute a fire hazard.
(b) Each airplane that has a passenger seating
configuration excluding pilots seats, of 10 seats or more
must be designed so that with the airplane under control it
can be landed on a paved runway with any one or more landing
gear legs not extended without sustaining a structural
component failure that is likely to cause the spillage of
enough fuel to constitute a fire hazard.
(c) Compliance with the provisions of this section may be
shown by analysis or tests, or both.
[Amdt. 25-32, 37 FR 3969, Feb. 24, 1972]
Sec. 25.723 Shock absorption tests.
(a) It must be shown that the limit load factors selected
for design in accordance with Sec. 25.473 for takeoff and
landing weights, respectively, will not be exceeded. This
must be shown by energy absorption tests except that
analyses based on earlier tests conducted on the same basic
landing gear system which has similar energy absorption
characteristics may be used for increases in previously
approved takeoff and landing weights.
(b) The landing gear may not fail in a test,
demonstrating its reserve energy absorption capacity,
simulating a descent velocity of 12 f.p.s. at design landing
weight, assuming airplane lift not greater than the airplane
weight acting during the landing impact.
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970, as amended by
Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55 FR
29777, July 20, 1990]
Sec. 25.725 Limit drop tests.
(a) If compliance with Sec. 25.723(a) is shown by free
drop tests, these tests must be made on the complete
airplane, or on units consisting of a wheel, tire, and shock
absorber, in their proper positions, from free drop heights
not less than--
(1) 18.7 inches for the design landing weight conditions;
and
(2) 6.7 inches for the design takeoff weight
conditions.
(b) If airplane lift is simulated by air cylinders or by
other mechanical means, the weight used for the drop must be
equal to W. If the effect of airplane lift is represented in
free drop tests by an equivalent reduced mass, the landing
gear must be dropped with an effective mass equal to
[...??equation goes here]
(c) The drop test attitude of the landing gear unit and
the application of appropriate drag loads during the test
must simulate the airplane landing conditions in a manner
consistent with the development of a rational or
conservative limit load factor value.
(d) The value of d used in the computation of We in
paragraph (b) of this section may not exceed the value
actually obtained in the drop test.
(e) The limit inertia load factor n must be determined
from the free drop test in paragraph (b) of this section
according to the following formula:
[...??equation goes here]
(f) The value of n determined in paragraph (e) of this
section may not be more than the limit inertia load factor
used in the landing conditions in Sec. 25.473.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.727 Reserve energy absorption drop
tests.
(a) If compliance with the reserve energy absorption
condition specified in Sec. 25.723(b) is shown by free drop
tests, the drop height may not be less than 27 inches.
(b) If airplane lift is simulated by air cylinders or by
other mechanical means, the weight used for the drop must be
equal to W. If the effect of airplane lift is represented in
free drop tests by an equivalent reduced mass, the landing
gear must be dropped with an effective mass,
[...??equation goes here]
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.729 Retracting mechanism.
(a) General. For airplanes with retractable landing gear,
the following apply:
(1) The landing gear retracting mechanism, wheel well
doors, and supporting structure, must be designed for--
(i) The loads occurring in the flight conditions when the
gear is in the retracted position,
(ii) The combination of friction loads, inertia loads,
brake torque loads, air loads, and gyroscopic loads
resulting from the wheels rotating at a peripheral speed
equal to 1.3 Vs (with the flaps in takeoff position at
design takeoff weight), occurring during retraction and
extension at any airspeed up to 1.6 Vs1 (with the flaps in
the approach position at design landing weight), and
(iii) Any load factor up to those specified in Sec.
25.345(a) for the flaps extended condition.
(2) Unless there are other means to decelerate the
airplane in flight at this speed, the landing gear, the
retracting mechanism, and the airplane structure (including
wheel well doors) must be designed to withstand the flight
loads occurring with the landing gear in the extended
position at any speed up to 0.67 VC.
(3) Landing gear doors, their operating mechanism, and
their supporting structures must be designed for the yawing
maneuvers prescribed for the airplane in addition to the
conditions of airspeed and load factor prescribed in
paragraphs (a) (1) and (2) of this section.
(b) Landing gear lock. There must be positive means to
keep the landing gear extended, in flight and on the
ground.
(c) Emergency operation. There must be an emergency means
for extending the landing gear in the event of--
(1) Any reasonably probable failure in the normal
retraction system; or
(2) The failure of any single source of hydraulic,
electric, or equivalent energy supply.
(d) Operation test. The proper functioning of the
retracting mechanism must be shown by operation tests.
(e) Position indicator and warning device. If a
retractable landing gear is used, there must be a landing
gear position indicator (as well as necessary switches to
actuate the indicator) or other means to inform the pilot
that the gear is secured in the extended (or retracted)
position. This means must be designed as follows:
(1) If switches are used, they must be located and
coupled to the landing gear mechanical systems in a manner
that prevents an erroneous indication of "down and locked"
if the landing gear is not in a fully extended position, or
of "up and locked" if the landing gear is not in the fully
retracted position. The switches may be located where they
are operated by the actual landing gear locking latch or
device.
(2) The flightcrew must be given an aural warning that
functions continuously, or is periodically repeated, if a
landing is attempted when the landing gear is not locked
down.
(3) The warning must be given in sufficient time to allow
the landing gear to be locked down or a go-around to be
made.
(4) There must not be a manual shut-off means readily
available to the flightcrew for the warning required by
paragraph (e)(2) of this section such that it could be
operated instinctively, inadvertently, or by habitual
reflexive action.
(5) The system used to generate the aural warning must be
designed to eliminate false or inappropriate alerts.
(6) Failures of systems used to inhibit the landing gear
aural warning, that would prevent the warning system from
operating, must be improbable.
(f) Protection of equipment in wheel wells. Equipment
that is essential to safe operation of the airplane and that
is located in wheel wells must be protected from the
damaging effects of--
(1) A bursting tire, unless it is shown that a tire
cannot burst from overheat; and
(2) A loose tire tread, unless it is shown that a loose
tire tread cannot cause damage.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt.
25-42, 43 FR 2323, Jan. 16, 1978; Amdt. 25-72, 55 FR 29777,
July 20, 1990; Amdt. 25-75, 56 FR 63762, Dec. 5,
1991]
Sec. 25.731 Wheels.
(a) Each main and nose wheel must be approved.
(b) The maximum static load rating of each wheel may not
be less than the corresponding static ground reaction
with--
(1) Design maximum weight; and
(2) Critical center of gravity.
(c) The maximum limit load rating of each wheel must
equal or exceed the maximum radial limit load determined
under the applicable ground load requirements of this
part.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-72, 55 FR 29777, July 20, 1990]
Sec. 25.733 Tires.
(a) When a landing gear axle is fitted with a single
wheel and tire assembly, the wheel must be fitted with a
suitable tire of proper fit with a speed rating approved by
the Administrator that is not exceeded under critical
conditions and with a load rating approved by the
Administrator that is not exceeded under--
(1) The loads on the main wheel tire, corresponding to
the most critical combination of airplane weight (up to
maximum weight) and center of gravity position, and
(2) The loads corresponding to the ground reactions in
paragraph (b) of this section, on the nose wheel tire,
except as provided in paragraphs (b)(2) and (b)(3) of this
section.
(b) The applicable ground reactions for nose wheel tires
are as follows:
(1) The static ground reaction for the tire corresponding
to the most critical combination of airplane weight (up to
maximum ramp weight) and center of gravity position with a
force of 1.0g acting downward at the center of gravity. This
load may not exceed the load rating of the tire.
(2) The ground reaction of the tire corresponding to the
most critical combination of airplane weight (up to maximum
landing weight) and center of gravity position combined with
forces of 1.0g downward and 0.31g forward acting at the
center of gravity. The reactions in this case must be
distributed to the nose and main wheels by the principles of
statics with a drag reaction equal to 0.31 times the
vertical load at each wheel with brakes capable of producing
this ground reaction. This nose tire load may not exceed 1.5
times the load rating of the tire.
(3) The ground reaction of the tire corresponding to the
most critical combination of airplane weight (up to maximum
ramp weight) and center of gravity position combined with
forces of 1.0g downward and 0.20g forward acting at the
center of gravity. The reactions in this case must be
distributed to the nose and main wheels by the principles of
statics with a drag reaction equal to 0.20 times the
vertical load at each wheel with brakes capable of producing
this ground reaction. This nose tire load may not exceed 1.5
times the load rating of the tire.
(c) When a landing gear axle is fitted with more than one
wheel and tire assembly, such as dual or dual-tandem, each
wheel must be fitted with a suitable tire of proper fit with
a speed rating approved by the Administrator that is not
exceeded under critical conditions, and with a load rating
approved by the Administrator that is not exceeded by--
(1) The loads on each main wheel tire, corresponding to
the most critical combination of airplane weight (up to
maximum weight) and center of gravity position, when
multiplied by a factor of 1.07; and
(2) Loads specified in paragraphs (a)(2), (b)(1), (b)(2),
and (b)(3) of this section on each nose wheel tire.
(d) Each tire installed on a retractable landing gear
system must, at the maximum size of the tire type expected
in service, have a clearance to surrounding structure and
systems that is adequate to prevent unintended contact
between the tire and any part of the structure or
systems.
(e) For an airplane with a maximum certificated takeoff
weight of more than 75,000 pounds, tires mounted on braked
wheels must be inflated with dry nitrogen or other gases
shown to be inert so that the gas mixture in the tire does
not contain oxygen in excess of 5 percent by volume, unless
it can be shown that the tire liner material will not
produce a volatile gas when heated or that means are
provided to prevent tire temperatures from reaching unsafe
levels.
[Amdt. 25-48, 44 FR 68752, Nov. 29, 1979, as amended
by Amdt. 25-72, 55 FR 29777, July 20, 1990; Amdt. 25-78, 58
FR 11781, Feb. 26, 1993]
Sec. 25.735 Brakes.
(a) Each brake must be approved.
(b) The brake system and associated systems must be
designed and constructed so that if any electrical,
pneumatic, hydraulic, or mechanical connecting or
transmitting element (excluding the operating pedal or
handle) fails, or if any single source of hydraulic or other
brake operating energy supply is lost, it is possible to
bring the airplane to rest under conditions specified in
Sec. 25.125, with a mean deceleration during the landing
roll of at least 50 percent of that obtained in determining
the landing distance as prescribed in that section.
Subcomponents within the brake assembly, such as brake drum,
shoes, and actuators (or their equivalents), shall be
considered as connecting or transmitting elements, unless it
is shown that leakage of hydraulic fluid resulting from
failure of the sealing elements in these subcomponents
within the brake assembly would not reduce the braking
effectiveness below that specified in this paragraph.
(c) Brake controls may not require excessive control
force in their operation.
(d) The airplane must have a parking control that, when
set by the pilot, will without further attention, prevent
the airplane from rolling on a paved, level runway with
takeoff power on the critical engine.
(e) If antiskid devices are installed, the devices and
associated systems must be designed so that no single
probable malfunction will result in a hazardous loss of
braking ability or directional control of the airplane.
(f) The brake kinetic energy capacity rating of each main
wheel-brake assembly may not be less than the kinetic energy
absorption requirements determined under either of the
following methods:
(1) The brake kinetic energy absorption requirements must
be based on a rational analysis of the sequence of events
expected during operational landings at maximum landing
weight. This analysis must include conservative values of
airplane speed at which the brakes are applied, braking
coefficient of friction between tires and runway,
aerodynamic drag, propeller drag or power-plant forward
thrust, and (if more critical) the most adverse single
engine or propeller malfunction.
(2) Instead of a rational analysis, the kinetic energy
absorption requirements for each main wheel brake assembly
may be derived from the following formula, which assumes an
equal distribution of braking between main wheels:
[...??equation goes here]
(g) The minimum stalling speed rating of each main
wheel-brake assembly (that is, the initial speed used in the
dynamometer tests) may not be more than the V used in the
determination of kinetic energy in accordance with paragraph
(f) of this section, assuming that the test procedures for
wheel brake assemblies involve a specified rate of
deceleration, and, therefore, for the same amount of kinetic
energy, the rate of energy absorption (the power absorbing
ability of the brake) varies inversely with the initial
speed.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt.
25-48, 44 FR 68742, Nov. 29, 1979; Amdt. 25-72, 55 FR 29777,
July 20, 1990]
Sec. 25.737 Skis.
Each ski must be approved. The maximum limit load rating
of each ski must equal or exceed the maximum limit load
determined under the applicable ground load requirements of
this part. Floats and Hulls
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Floats
and Hulls:
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Sec. 25.751 Main float buoyancy.
Each main float must have--
(a) A buoyancy of 80 percent in excess of that required
to support the maximum weight of the seaplane or amphibian
in fresh water; and
(b) Not less than five watertight compartments
approximately equal in volume.
Sec. 25.753 Main float design.
Each main float must be approved and must meet the
requirements of Sec. 25.521.
Sec. 25.755 Hulls.
(a) Each hull must have enough watertight compartments so
that, with any two adjacent compartments flooded, the
buoyancy of the hull and auxiliary floats (and wheel tires,
if used) provides a margin of positive stability great
enough to minimize the probability of capsizing in rough,
fresh water.
(b) Bulkheads with watertight doors may be used for
communication between compartments.
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Personnel
and Cargo Accommodations:
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Sec. 25.771 Pilot compartment.
(a) Each pilot compartment and its equipment must allow
the minimum flight crew (established under Sec. 25.1523) to
perform their duties without unreasonable concentration or
fatigue.
(b) The primary controls listed in Sec. 25.779(a),
excluding cables and control rods, must be located with
respect to the propellers so that no member of the minimum
flight crew (established under Sec. 25.1523), or part of the
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 making an
angle of five degrees forward or aft of the plane of
rotation of the propeller.
(c) If provision is made for a second pilot, the airplane
must be controllable with equal safety from either pilot
seat.
(d) The pilot compartment must be constructed so that,
when flying in rain or snow, it will not leak in a manner
that will distract the crew or harm the structure.
(e) Vibration and noise characteristics of cockpit
equipment may not interfere with safe operation of the
airplane.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-4, 30 FR 6113, Apr. 30, 1965]
Sec. 25.772 Pilot compartment doors.
For an airplane that has a maximum passenger seating
configuration of more than 20 seats and that has a lockable
door installed between the pilot compartment and the
passenger compartment:
(a) The emergency exit configuration must be designed so
that neither crewmembers nor passengers need use that door
in order to reach the emergency exits provided for them;
and
(b) Means must be provided to enable flight crewmembers
to directly enter the passenger compartment from the pilot
compartment if the cockpit door becomes jammed.
[Doc. No. 24344, Admt. 25-72, 55 FR 29777, July 20,
1990]
Sec. 25.773 Pilot compartment view.
(a) Nonprecipitation conditions. For nonprecipitation
conditions, the following apply:
(1) Each pilot compartment must be arranged to give the
pilots a sufficiently extensive, clear, and undistorted
view, to enable them to safely perform any maneuvers within
the operating limitations of the airplane, including taxiing
takeoff, approach, and landing.
(2) Each pilot compartment must be free of glare and
reflection that could interfere with the normal duties of
the minimum flight crew (established under Sec. 25.1523).
This must be shown in day and night flight tests under
nonprecipitation conditions.
(b) Precipitation conditions. For precipitation
conditions, the following apply:
(1) The airplane must have a means to maintain a clear
portion of the windshield, during precipitation conditions,
sufficient for both pilots to have a sufficiently extensive
view along the flight path in normal flight attitudes of the
airplane. This means must be designed to function, without
continuous attention on the part of the crew, in--
(i) Heavy rain at speeds up to 1.6 Vs1 with lift and drag
devices retracted; and
(ii) The icing conditions specified in Sec. 25.1419 if
certification with ice protection provisions is
requested.
(2) The first pilot must have--
(i) A window that is openable under the conditions
prescribed in paragraph (b)(1) of this section when the
cabin is not pressurized, provides the view specified in
that paragraph, and gives sufficient protection from the
elements against impairment of the pilot's vision; or
(ii) An alternate means to maintain a clear view under
the conditions specified in paragraph (b)(1) of this
section, considering the probable damage due to a severe
hail encounter.
(c) Internal windshield and window fogging. The airplane
must have a means to prevent fogging of the internal
portions of the windshield and window panels over an area
which would provide the visibility specified in paragraph
(a) of this section under all internal and external ambient
conditions, including precipitation conditions, in which the
airplane is intended to be operated.
(d) Fixed markers or other guides must be installed at
each pilot station to enable the pilots to position
themselves in their seats for an optimum combination of
outside visibility and instrument scan. If lighted markers
or guides are used they must comply with the requirements
specified in Sec. 25.1381.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt.
25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 29778,
July 20, 1990]
Sec. 25.775 Windshields and windows.
(a) Internal panes must be made of nonsplintering
material.
(b) Windshield panes directly in front of the pilots in
the normal conduct of their duties, and the supporting
structures for these panes, must withstand, without
penetration, the impact of a four-pound bird when the
velocity of the airplane (relative to the bird along the
airplane's flight path) is equal to the value of VC, at sea
level, selected under Sec. 25.335(a).
(c) Unless it can be shown by analysis or tests that the
probability of occurrence of a critical windshield
fragmentation condition is of a low order, the airplane must
have a means to minimize the danger to the pilots from
flying windshield fragments due to bird impact. This must be
shown for each transparent pane in the cockpit that--
(1) Appears in the front view of the airplane;
(2) Is inclined 15 degrees or more to the longitudinal
axis of the airplane; and
(3) Has any part of the pane located where its
fragmentation will constitute a hazard to the pilots.
(d) The design of windshields and windows in pressurized
airplanes must be based on factors peculiar to high altitude
operation, including the effects of continuous and cyclic
pressurization loadings, the inherent characteristics of the
material used, and the effects of temperatures and
temperature differentials. The windshield and window panels
must be capable of withstanding the maximum cabin pressure
differential loads combined with critical aerodynamic
pressure and temperature effects after any single failure in
the installation or associated systems. It may be assumed
that, after a single failure that is obvious to the flight
crew (established under Sec. 25.1523), the cabin pressure
differential is reduced from the maximum, in accordance with
appropriate operating limitations, to allow continued safe
flight of the airplane with a cabin pressure altitude of not
more than 15,000 feet.
(e) The windshield panels in front of the pilots must be
arranged so that, assuming the loss of vision through any
one panel, one or more panels remain available for use by a
pilot seated at a pilot station to permit continued safe
flight and landing.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt.
25-38, 41 FR 55466, Dec. 20, 1976]
Sec. 25.777 Cockpit controls.
(a) Each cockpit control must be located to provide
convenient operation and to prevent confusion and
inadvertent operation.
(b) The direction of movement of cockpit controls must
meet the requirements of Sec. 25.779. Wherever practicable,
the sense of motion involved in the operation of other
controls must correspond to the sense of the effect of the
operation upon the airplane or upon the part operated.
Controls of a variable nature using a rotary motion must
move clockwise from the off position, through an increasing
range, to the full on position.
(c) The controls must be located and arranged, with
respect to the pilots' seats, so that there is full and
unrestricted movement of each control without interference
from the cockpit structure or the clothing of the minimum
flight crew (established under Sec. 25.1523) when any member
of this flight crew, from 5'2'' to 6'3'' in height, is
seated with the seat belt and shoulder harness (if provided)
fastened.
(d) Identical powerplant controls for each engine must be
located to prevent confusion as to the engines they
control.
(e) Wing flap controls and other auxiliary lift device
controls must be located on top of the pedestal, aft of the
throttles, centrally or to the right of the pedestal
centerline, and not less than 10 inches aft of the landing
gear control.
(f) The landing gear control must be located forward of
the throttles and must be operable by each pilot when seated
with seat belt and shoulder harness (if provided)
fastened.
(g) Control knobs must be shaped in accordance with Sec.
25.781. In addition, the knobs must be of the same color,
and this color must contrast with the color of control knobs
for other purposes and the surrounding cockpit.
(h) If a flight engineer is required as part of the
minimum flight crew (established under Sec. 25.1523), the
airplane must have a flight engineer station located and
arranged so that the flight crewmembers can perform their
functions efficiently and without interfering with each
other.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-46, 43 FR 50596, Oct. 30, 1978]
Sec. 25.779 Motion and effect of cockpit
controls.
Cockpit controls must be designed so that they operate in
accordance with the following movement and actuation:
(a) Aerodynamic controls:
(1) Primary.Controls Motion and effect
- Aileron Right (clockwise) for right wing down
- Elevator Rearward for nose up
- Rudder Right pedal forward for nose right
(2) Secondary.Controls Motion and effect
- Flaps (or auxiliary lift devices) Forward for flaps up;
rearward for flaps down
- Trim tabs (or equivalent) Rotate to produce similar
rotation of the airplane about an axis parallel to the axis
of the control
(b) Powerplant and auxiliary controls:
(1) Powerplant Motion and effect
- Power or thrust Forward to increase forward thrust and
rearward to increase rearward thrust
- PropellersForward to increase rpm
- Mixture Forward or upward for rich.
- Carburetor air heat Forward or upward for cold.
- Supercharger Forward or upward for low blower. For
turbosuperchargers, forward, upward, or clockwise, to
increase pressure.
(2) Auxiliary Motion and effect
- Landing gear Controls Down to extend
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-72, 55 FR 29778, July 20, 1990]
Sec. 25.781 Cockpit control knob shape.
Cockpit control knobs must conform to the general shapes
(but not necessarily the exact sizes or specific
proportions) in the following figure:
- Flap Control Knob [ ...Illustration appears here...
]
- Landing Gear Control Knob [ ...Illustration appears
here... ]
- Mixture Control Knob [ ...Illustration appears here...
]
- Supercharger Control Knob [ ...Illustration appears
here... ]
- Power or Thrust Knob [ ...Illustration appears here...
]
- Propeller Control Knob[ ...Illustration appears
here... ]
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-72, 55 FR 29778, July 20, 1990]
Sec. 25.783 Doors.
(a) Each cabin must have at least one easily accessible
external door.
(b) There must be a means to lock and safeguard each
external door against opening in flight (either
inadvertently by persons or as a result of mechanical
failure or failure of a single structural element either
during or after closure). Each external door must be
openable from both the inside and the outside, even though
persons may be crowded against the door on the inside of the
airplane. Inward opening doors may be used if there are
means to prevent occupants from crowding against the door to
an extent that would interfere with the opening of the door.
The means of opening must be simple and obvious and must be
arranged and marked so that it can be readily located and
operated, even in darkness. Auxiliary locking devices may be
used.
(c) Each external door must be reasonably free from
jamming as a result of fuselage deformation in a minor
crash.
(d) Each external door must be located where persons
using them will not be endangered by the propellers when
appropriate operating procedures are used.
(e) There must be a provision for direct visual
inspection of the locking mechanism to determine if external
doors, for which the initial opening movement is not inward
(including passenger, crew, service, and cargo doors), are
fully closed and locked. The provision must be discernible
under operational lighting conditions by appropriate
crewmembers using a flashlight or equivalent lighting
source. In addition, there must be a visual warning means to
signal the appropriate flight crewmembers if any external
door is not fully closed and locked. The means must be
designed such that any failure or combination of failures
that would result in an erroneous closed and locked
indication is improbable for doors for which the initial
opening movement is not inward.
(f) External doors must have provisions to prevent the
initiation of pressurization of the airplane to an unsafe
level if the door is not fully closed and locked. In
addition, it must be shown by safety analysis that
inadvertent opening is extemely improbable.
(g) Cargo and service doors not suitable for use as
emergency exits need only meet paragraphs (e) and (f) of
this section and be safeguarded against opening in flight as
a result of mechanical failure or failure of a single
structural element.
(h) Each passenger entry door in the side of the fuselage
must qualify as a Type A, Type I, or Type II passenger
emergency exit and must meet the requirements of Secs.
25.807 through 25.813 that apply to that type of passenger
emergency exit.
(i) If an integral stair is installed in a passenger
entry door that is qualified as a passenger emergency exit,
the stair must be designed so that under the following
conditions the effectiveness of passenger emergency egress
will not be impaired:
(1) The door, integral stair, and operating mechanism
have been subjected to the inertia forces specified in Sec.
25.561(b)(3), acting separately relative to the surrounding
structure.
(2) The airplane is in the normal ground attitude and in
each of the attitudes corresponding to collapse of one or
more legs of the landing gear.
(j) All lavatory doors must be designed to preclude
anyone from becoming trapped inside the lavatory, and if a
locking mechanism is installed, it be capable of being
unlocked from the outside without the aid of special
tools.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-15, 32 FR 13262, Sept. 20, 1967; Amdt.
25-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-54, 45 FR 60173,
Sept. 11, 1980; Amdt. 25-72, 55 FR 29780, July 20,
1990]
Sec. 25.785 Seats, berths, safety belts, and
harnesses.
(a) A seat (or berth for a nonambulant person) must be
provided for each occupant who has reached his or her second
birthday.
(b) Each seat, berth, safety belt, harness, and adjacent
part of the airplane at each station designated as
occupiable during takeoff and landing must be designed so
that a person making proper use of these facilities will not
suffer serious injury in an emergency landing as a result of
the inertia forces specified in Secs. 25.561 and 25.562.
(c) Each seat or berth must be approved.
(d) Each occupant of a seat that makes more than an
18-degree angle with the vertical plane containing the
airplane centerline must be protected from head injury by a
safety belt and an energy absorbing rest that will support
the arms, shoulders, head, and spine, or by a safety belt
and shoulder harness that will prevent the head from
contacting any injurious object. Each occupant of any other
seat must be protected from head injury by a safety belt
and, as appropriate to the type, location, and angle of
facing of each seat, by one or more of the following:
(1) A shoulder harness that will prevent the head from
contacting any injurious object.
(2) The elimination of any injurious object within
striking radius of the head.
(3) An energy absorbing rest that will support the arms,
shoulders, head, and spine.
(e) Each berth must be designed so that the forward part
has a padded end board, canvas diaphragm, or equivalent
means, that can withstand the static load reaction of the
occupant when subjected to the forward inertia force
specified in Sec. 25.561. Berths must be free from corners
and protuberances likely to cause injury to a person
occupying the berth during emergency conditions.
(f) Each seat or berth, and its supporting structure, and
each safety belt or harness and its anchorage must be
designed for an occupant weight of 170 pounds, considering
the maximum load factors, inertia forces, and reactions
among the occupant, seat, safety belt, and harness for each
relevant flight and ground load condition (including the
emergency landing conditions prescribed in Sec. 25.561). In
addition--
(1) The structural analysis and testing of the seats,
berths, and their supporting structures may be determined by
assuming that the critical load in the forward, sideward,
downward, upward, and rearward directions (as determined
from the prescribed flight, ground, and emergency landing
conditions) acts separately or using selected combinations
of loads if the required strength in each specified
direction is substantiated. The forward load factor need not
be applied to safety belts for berths.
(2) Each pilot seat must be designed for the reactions
resulting from the application of the pilot forces
prescribed in Sec. 25.395.
(3) The inertia forces specified in Sec. 25.561 must be
multiplied by a factor of 1.33 (instead of the fitting
factor prescribed in Sec. 25.625) in determining the
strength of the attachment of each seat to the structure and
each belt or harness to the seat or structure.
(g) Each seat at a flight deck station must have a
restraint system consisting of a combined safety belt and
shoulder harness with a single-point release that permits
the flight deck occupant, when seated with the restraint
system fastened, to perform all of the occupant's necessary
flight deck functions. There must be a means to secure each
combined restraint system when not in use to prevent
interference with the operation of the airplane and with
rapid egress in an emergency.
(h) Each seat located in the passenger compartment and
designated for use during takeoff and landing by a flight
attendant required by the operating rules of this chapter
must be:
(1) Near a required floor level emergency exit, except
that another location is acceptable if the emergency egress
of passengers would be enhanced with that location. A flight
attendant seat must be located adjacent to each Type A
emergency exit. Other flight attendant seats must be evenly
distributed among the required floor level emergency exits
to the extent feasible.
(2) To the extent possible, without compromising
proximity to a required floor level emergency exit, located
to provide a direct view of the cabin area for which the
flight attendant is responsible.
(3) Positioned so that the seat will not interfere with
the use of a passageway or exit when the seat is not in
use.
(4) Located to minimize the probability that occupants
would suffer injury by being struck by items dislodged from
service areas, stowage compartments, or service
equipment.
(5) Either forward or rearward facing with an energy
absorbing rest that is designed to support the arms,
shoulders, head, and spine.
(6) Equipped with a restraint system consisting of a
combined safety belt and shoulder harness unit with a single
point release. There must be means to secure each restraint
system when not in use to prevent interference with rapid
egress in an emergency.
(i) Each safety belt must be equipped with a metal to
metal latching device.
(j) If the seat backs do not provide a firm handhold,
there must be a handgrip or rail along each aisle to enable
persons to steady themselves while using the aisles in
moderately rough air.
(k) Each projecting object that would injure persons
seated or moving about the airplane in normal flight must be
padded.
(l) Each forward observer's seat required by the
operating rules must be shown to be suitable for use in
conducting the necessary enroute inspection.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29780, July 20,
1990]
Sec. 25.787 Stowage compartments.
(a) Each compartment for the stowage of cargo, baggage,
carry-on articles, and equipment (such as life rafts), and
any other stowage compartment must be designed for its
placarded maximum weight of contents and for the critical
load distribution at the appropriate maximum load factors
corresponding to the specified flight and ground load
conditions, and to the emergency landing conditions of Sec.
25.561(b), except that the forces specified in the emergency
landing conditions need not be applied to compartments
located below, or forward, of all occupants in the airplane.
If the airplane has a passenger seating configuration,
excluding pilots seats, of 10 seats or more, each stowage
compartment in the passenger cabin, except for underseat and
overhead compartments for passenger convenience, must be
completely enclosed.
(b) There must be a means to prevent the contents in the
compartments from becoming a hazard by shifting, under the
loads specified in paragraph (a) of this section. For
stowage compartments in the passenger and crew cabin, if the
means used is a latched door, the design must take into
consideration the wear and deterioration expected in
service.
(c) If cargo compartment lamps are installed, each lamp
must be installed so as to prevent contact between lamp bulb
and cargo.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-32, 37 FR 3969, Feb. 24, 1972; Amdt.
25-38, 41 FR 55466, Dec. 20, 1976; Amdt. 25-51, 45 FR 7755,
Feb. 4, 1980]
Sec. 25.789 Retention of items of mass in passenger
and crew compartments and galleys.
(a) Means must be provided to prevent each item of mass
(that is part of the airplane type design) in a passenger or
crew compartment or galley from becoming a hazard by
shifting under the appropriate maximum load factors
corresponding to the specified flight and ground load
conditions, and to the emergency landing conditions of Sec.
25.561(b).
(b) Each interphone restraint system must be designed so
that when subjected to the load factors specified in Sec.
25.561(b)(3), the interphone will remain in its stowed
position.
[Amdt. 25-32, 37 FR 3969, Feb. 24, 1972, as amended
by Amdt. 25-46, 43 FR 50596, Oct. 30, 1978]
Sec. 25.791 Passenger information signs and
placards.
(a) If smoking is to be prohibited, there must be at
least one placard so stating that is legible to each person
seated in the cabin. If smoking is to be allowed, and if the
crew compartment is separated from the passenger
compartment, there must be at least one sign notifying when
smoking is prohibited. Signs which notify when smoking is
prohibited must be operable by amember of the flightcrew
and, when illuminated, must be legible under all probable
conditions of cabin illumination to each person seated in
the cabin.
(b) Signs that notify when seat belts should be fastened
and that are installed to comply with the operating rules of
this chapter must be operable by a member of the flightcrew
and, when illuminated, must be legible under all probable
conditions of cabin illumination to each person seated in
the cabin.
(c) A placard must be located on or adjacent to the door
of each receptacle used for the disposal of flammable waste
materials to indicate that use of the receptacle for
disposal of cigarettes, etc., is prohibited.
(d) Lavatories must have "No Smoking" or "No Smoking in
Lavatory" placards conspicuously located on or adjacent to
each side of the entry door.
(e) Symbols that clearly express the intent of the sign
or placard may be used in lieu of letters.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29780, July 20,
1990]
Sec. 25.793 Floor surfaces.
The floor surface of all areas which are likely to become
wet in service must have slip resistant properties.
[Amdt. 25-51, 45 FR 7755, Feb. 4, 1980]
|
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Emergency
Provisions:
|
|
Sec. 25.801 Ditching.
(a) If certification with ditching provisions is
requested, the airplane must meet the requirements of this
section and Secs. 25.807(e), 25.1411, and 25.1415(a).
(b) Each practicable design measure, compatible with the
general characteristics of the airplane, must be taken to
minimize the probability that in an emergency landing on
water, the behavior of the airplane would cause immediate
injury to the occupants or would make it impossible for them
to escape.
(c) The probable behavior of the airplane in a water
landing must be investigated by model tests or by comparison
with airplanes of similar configuration for which the
ditching characteristics are known. Scoops, flaps,
projections, and any other factor likely to affect the
hydrodynamic characteristics of the airplane, must be
considered.
(d) It must be shown that, under reasonably probable
water conditions, the flotation time and trim of the
airplane will allow the occupants to leave the airplane and
enter the liferafts required by Sec. 25.1415. If compliance
with this provision is shown by buoyancy and trim
computations, appropriate allowances must be made for
probable structural damage and leakage. If the airplane has
fuel tanks (with fuel jettisoning provisions) that can
reasonably be expected to withstand a ditching without
leakage, the jettisonable volume of fuel may be considered
as buoyancy volume.
(e) Unless the effects of the collapse of external doors
and windows are accounted for in the investigation of the
probable behavior of the airplane in a water landing (as
prescribed in paragraphs (c) and (d) of this section), the
external doors and windows must be designed to withstand the
probable maximum local pressures.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-72, 55 FR 29781, July 20, 1990]
Sec. 25.803 Emergency evacuation.
(a) Each crew and passenger area must have emergency
means to allow rapid evacuation in crash landings, with the
landing gear extended as well as with the landing gear
retracted, considering the possibility of the airplane being
on fire.
(b) [Reserved]
(c) For airplanes having a seating capacity of more than
44 passengers, it must be shown that the maximum seating
capacity, including the number of crewmembers required by
the operating rules for which certification is requested,
can be evacuated from the airplane to the ground under
simulated emergency conditions within 90 seconds. Compliance
with this requirement must be shown by actual demonstration
using the test criteria outlined in appendix Jof this part
unless the Administrator finds that a combination of
analysis and testing will provide data equivalent to that
which would be obtained by actual demonstration.
(d) [Reserved]
(e) [Reserved]
[Doc. No. 5066, 29 FR 18291 Dec. 24, 1964, as amended
by Amdt. 25-15, 32 FR 13262, Sept. 20, 1967; Amdt. 25-20, 34
FR 5544, Mar. 22, 1969; Amdt. 25-32, 37 FR 3969, Feb. 24,
1972; Amdt. 25-46, 43 FR 50596, Oct. 30, 1978; Amdt. 25-72,
55 FR 29781, July 20, 1990]
Sec. 25.805 [Removed. 55 FR 29781, July 20,
1990]
EDITORIAL NOTE: For the convenience of the user, the
removed text is set out below.
Sec. 25.805 Flight crew emergency exits.
Except for airplanes with a passenger capacity of 20 or
less in which the proximity of passenger emergency exits to
the flight crew area offers a convenient and readily
accessible means of evacuation for the flight crew, the
following apply:
(a) There must be either one exit on each side of the
airplane or a top hatch, in the flight crew area.
(b) Each exit must be of sufficient size and must be
located so as to allow rapid evacuation of the crew. An exit
size and shape of other than at least 19 by 20 inches
unobstructed rectangular opening may be used only if exit
utility is satisfactorily shown, by a typical flight
crewmember, to the Administrator.
Sec. 25.807 Emergency exits.
(a) Type. For the purpose of this part, the types of
exits are defined as follows:
(1) Type I. This type is a floor level exit with a
rectangular opening of not less than 24 inches wide by 48
inches high, with corner radii not greater than one-third
the width of the exit.
(2) Type II. This type is a rectangular opening of not
less than 20 inches wide by 44 inches high, with corner
radii not greater than one-third the width of the exit. Type
II exits must be floor level exits unless located over the
wing, in which case they may not have a step-up inside the
airplane of more than 10 inches nor a step-down outside the
airplane of more than 17 inches.
(3) Type III. This type is a rectangular opening of not
less than 20 inches wide by 36 inches high, with corner
radii not greater than one-third the width of the exit, and
with a step-up inside the airplane of not more than 20
inches. If the exit is located over the wing, the step-down
outside the airplane may not exceed 27 inches.
(4) Type IV. This type is a rectangular opening of not
less than 19 inches wide by 26 inches high, with corner
radii not greater than one-third the width of the exit,
located over the wing, with a step-up inside the airplane of
not more than 29 inches and a step-down outside the airplane
of not more than 36 inches.
(5) Ventral. This type is an exit from the passenger
compartment through the pressure shell and the bottom
fuselage skin. The dimensions and physical configuration of
this type of exit must allow at least the same rate of
egress as a Type I exit with the airplane in the normal
ground attitude, with landing gear extended.
(6) Tail cone. This type is an aft exit from the
passenger compartment through the pressure shell and through
an openable cone of the fuselage aft of the pressure shell.
The means of opening the tailcone must be simple and obvious
and must employ a single operation.
(7) Type A. This type is a floor level exit with a
rectangular opening of not less than 42 inches wide by 72
inches high with corner radii not greater than one-sixth of
the width of the exit.
(b) Step down distance. Step down distance, as used in
this section, means the actual distance between the bottom
of the required opening and a usable foot hold, extending
out from the fuselage, that is large enough to be effective
without searching by sight or feel.
(c) Over-sized exits. Openings larger than those
specified in this section, whether or not of rectangular
shape, may be used if the specified rectangular opening can
be inscribed within the opening and the base of the
inscribed rectangular opening meets the specified step-up
and step-down heights.
(d) Passenger emergency exits. Except as provided in
paragraphs (d) (3) through (7) of this section, the minimum
number and type of passenger emergency exits is as
follows:
(1) For passenger seating configurations of 1 through 299
seats:
Emergency exits for each side of the fuselage
Passenger seating configuration
(crewmember seats not Type Type Type Type included)III
III IV
1 through 9 1
10 through 19 1
20 through 39 1 1
40 through 791 1
80 through 1091 2
110 through 1392 1
140 through 1792 2
Additional exits are required for passenger seating
configurations greater than 179 seats in accordance with the
following table:
Additional emergency exits Increase in passenger
(each side seating of configuration fuselage) allowed
Type A 110
Type I 45
Type II 40
Type III 35
(2) For passenger seating configurations greater than 299
seats, each emergency exit in the side of the fuselage must
be either a Type A or Type I. Apassenger seating
configuration of 110 seats is allowed for each pair of Type
A exits and a passenger seating configuration of 45 seats is
allowed for each pair of Type I exits.
(3) If a passenger ventral or tail cone exit is installed
and that exit provides at least the same rate of egress as a
Type III exit with the airplane in the most adverse exit
opening condition that would result from the collapse of one
or more legs of the landing gear, an increase in the
passenger seating configuration beyond the limits specified
in paragraph (d) (1) or (2) of this section may be allowed
as follows:
(i) For a ventral exit, 12 additional passenger
seats.
(ii) For a tail cone exit incorporating a floor level
opening of not less than 20 inches wide by 60 inches high,
with corner radii not greater than one-third the width of
the exit, in the pressure shell and incorporating an
approved assist means in accordance with Sec. 25.809(h), 25
additional passenger seats.
(iii) For a tail cone exit incorporating an opening in
the pressure shell which is at least equivalent to a Type
III emergency exit with respect to dimensions, step-up and
step-down distance, and with the top of the opening not less
than 56 inches from the passenger compartment floor, 15
additional passenger seats.
(4) For airplanes on which the vertical location of the
wing does not allow the installation of overwing exits, an
exit of at least the dimensions of a Type III exit must be
installed instead of each Type IV exit required by
subparagraph (1) of this paragraph.
(5) An alternate emergency exit configuration may be
approved in lieu of that specified in paragraph (d) (1) or
(2) of this section provided the overall evacuation
capability is shown to be equal to or greater than that of
the specified emergency exit configuration.
(6) The following must also meet the applicable emergency
exit requirements of Secs. 25.809 through 25.813:
(i) Each emergency exit in the passenger compartment in
excess of the minimum number of required emergency
exits.
(ii) Any other floor level door or exit that is
accessible from the passenger compartment and is as large or
larger than a Type II exit, but less than 46 inches
wide.
(iii) Any other passenger ventral or tail cone exit.
(7) For an airplane that is required to have more than
one passenger emergency exit for each side of the fuselage,
no passenger emergency exit shall be more than 60 feet from
any adjacent passenger emergency exit on the same side of
the same deck of the fuselage, as measured parallel to the
airplane's longitudinal axis between the nearest exit
edges.
(e) Ditching emergency exits for passengers. Ditching
emergency exits must be provided in accordance with the
following requirements whether or not certification with
ditching provisions is requested:
(1) For airplanes that have a passenger seating
configuration of nine seats or less, excluding pilots seats,
one exit above the waterline in each side of the airplane,
meeting at least the dimensions of a Type IV exit.
(2) For airplanes that have a passenger seating
configuration of 10 seats or more, excluding pilots seats,
one exit above the waterline in a side of the airplane,
meeting at least the dimensions of a Type III exit for each
unit (or part of a unit) of 35 passenger seats, but no less
than two such exits in the passenger cabin, with one on each
side of the airplane. The passenger seat/exit ratio may be
increased through the use of larger exits, or other means,
provided it is shown that the evacuation capability during
ditching has been improved accordingly.
(3) If it is impractical to locate side exits above the
waterline, the side exits must be replaced by an equal
number of readily accessible overhead hatches of not less
than the dimensions of a Type III exit, except that for
airplanes with a passenger configuration of 35 seats or
less, excluding pilots seats, the two required Type III side
exits need be replaced by only one overhead hatch.
(f) Flightcrew emergency exits. For airplanes in which
the proximity of passenger emergency exits to the flightcrew
area does not offer a convenient and readily accessible
means of evacuation of the flightcrew, and for all airplanes
having a passenger seating capacity greater than 20,
flightcrew exits shall be located in the flightcrew area.
Such exits shall be of sufficient size and so located as to
permit rapid evacuation by the crew. One exit shall be
provided on each side of the airplane; or, alternatively, a
top hatch shall be provided. Each exit must encompass an
unobstructed rectangular opening of at least 19 by 20 inches
unless satisfactory exit utility can be demonstrated by a
typical crewmember.
[Doc. No. 24344, Admt. 25-72, 55 FR 29781, July 20,
1990]
Sec. 25.809 Emergency exit arrangement.
(a) Each emergency exit, including a flight crew
emergency exit, must be a movable door or hatch in the
external walls of the fuselage, allowing unobstructed
opening to the outside.
(b) Each emergency exit must be openable from the inside
and the outside except that sliding window emergency exits
in the flight crew area need not be openable from the
outside if other approved exits are convenient and readily
accessible to the flight crew area. Each emergency exit must
be capable of being opened, when there is no fuselage
deformation--
(1) With the airplane in the normal ground attitude and
in each of the attitudes corresponding to collapse of one or
more legs of the landing gear; and
(2) Within 10 seconds measured from the time when the
opening means is actuated to the time when the exit is fully
opened.
(c) The means of opening emergency exits must be simple
and obvious and may not require exceptional effort. Internal
exit-opening means involving sequence operations (such as
operation of two handles or latches or the release of safety
catches) may be used for flight crew emergency exits if it
can be reasonably established that these means are simple
and obvious to crewmembers trained in their use.
(d) If a single power-boost or single power-operated
system is the primary system for operating more than one
exit in an emergency, each exit must be capable of meeting
the requirements of paragraph (b) of this section in the
event of failure of the primary system. Manual operation of
the exit (after failure of the primary system) is
acceptable.
(e) Each emergency exit must be shown by tests, or by a
combination of analysis and tests, to meet the requirements
of paragraphs (b) and (c) of this section.
(f) There must be a means to lock each emergency exit and
to safeguard against its opening in flight, either
inadvertently by persons or as a result of mechanical
failure. In addition, there must be a means for direct
visual inspection of the locking mechanism by crewmembers to
determine that each emergency exit, for which the initial
opening movement is outward, is fully locked.
(g) There must be provisions to minimize the probability
of jamming of the emergency exits resulting from fuselage
deformation in a minor crash landing.
(h) When required by the operating rules for any large
passenger-carrying turbojet-powered airplane, each ventral
exit and tailcone exit must be- (1) Designed and constructed
so that it cannot be opened during flight; and (2) Marked
with a placard readable from a distance of 30 inches and
installed at a conspicuous location near the means of
opening the exit, stating that the exit has been designed
and constructed so that it cannot be opened during
flight.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-15, 32 FR 13264, Sept. 20, 1967; Amdt.
25-32, 37 FR 3970, Feb. 24, 1972; Amdt. 25-34, 37 FR 25355,
Nov. 30, 1972; Amdt. 25-46, 43 FR 50597, Oct. 30, 1978;
Amdt. 25-47, 44 FR 61325, Oct. 25, 1979; Amdt. 25-72, 55 FR
29782, July 20, 1990]
Sec. 25.810 Emergency egress assist means and escape
routes.
(a) Each nonoverwing landplane emergency exit more than 6
feet from the ground with the airplane on the ground and the
landing gear extended and each nonoverwing Type A exit must
have an approved means to assist the occupants in descending
to the ground.
(1) The assisting means for each passenger emergency exit
must be a self supporting slide or equivalent; and, in the
case of a Type A exit, it must be capable of carrying
simultaneously two parallel lines of evacuees. In addition,
the assisting means must be designed to meet the following
requirements:
(i) It must be automatically deployed and deployment must
begin during the interval between the time the exit opening
means is actuated from inside the airplane and the time the
exit is fully opened. However, each passenger emergency exit
which is also a passenger entrance door or a service door
must be provided with means to prevent deployment of the
assisting means when it is opened from either the inside or
the outside under nonemergency conditions for normal
use.
(ii) It must be automatically erected within 10 seconds
after deployment is begun.
(iii) It must be of such length after full deployment
that the lower end is self-supporting on the ground and
provides safe evacuation of occupants to the ground after
collapse of one or more legs of the landing gear.
(iv) It must have the capability, in 25-knot winds
directed from the most critical angle, to deploy and, with
the assistance of only one person, to remain usable after
full deployment to evacuate occupants safely to the
ground.
(v) For each system installation (mockup or airplane
installed), five consecutive deployment and inflation tests
must be conducted (per exit) without failure, and at least
three tests of each such |