to read as follows:

Sec.  23.3  Airplane categories.

* * * * *
    (d) The commuter category is limited to multiengine airplanes that
have a seating configuration, excluding pilot seats, of 19 or less, and
a maximum certificated takeoff weight of 19,000 pounds or less. * * *
* * * * *
    5. Amend Sec.  23.45 by revising the introductory text of paragraph
(h) to read as follows:

Sec.  23.45  General.

* * * * *
    (h) For multiengine turbojet powered airplanes over 6,000 pounds in
the normal, utility, and acrobatic category and commuter category
airplanes the following also apply:
* * * * *
    6. Amend Sec.  23.49 by revising the section heading and the
introductory text of paragraphs (a) and (c) to read as follows:

Sec.  23.49  Stalling speed.

    (a) VSO (maximum landing flap configuration) and
VS1 are the stalling speeds or the minimum steady flight
speeds, in knots (CAS), at which the airplane is controllable with--
     * * *
    (c) Except as provided in paragraph (d) of this section,
VSO at maximum weight may not exceed 61 knots for--
* * * * *
    7. Amend Sec.  23.51 by revising paragraph (b)(1) introductory text
and paragraph (c) introductory text to read as follows:

Sec.  23.51  Takeoff speeds.

* * * * *
    (b) * * *
    (1) For multiengine airplanes, the highest of--
     * * *
    (c) For normal, utility, and acrobatic category multiengine
turbojet airplanes of more than 6,000 pounds maximum weight and
commuter category airplanes, the following apply:
* * * * *
    8. Amend Sec.  23.53 by revising paragraph (c) to read as follows:

Sec.  23.53  Takeoff performance.

* * * * *
    (c) For normal, utility, and acrobatic category multiengine
turbojet airplanes of more than 6,000 pounds maximum weight and
commuter category airplanes, takeoff performance, as required by
Sec. Sec.  23.55 through 23.59, must be determined with the operating
engine(s) within approved operating limitations.
    9. Amend Sec.  23.55 by revising the introductory text to read as
follows:

Sec.  23.55  Accelerate-stop distance.

    For normal, utility, and acrobatic category multiengine turbojet
airplanes of more than 6,000 pounds maximum weight and commuter
category airplanes, the accelerate-stop distance must be determined as
follows:
* * * * *
    10. Amend Sec.  23.57 by revising the introductory text to read as
follows:

Sec.  23.57  Takeoff path.

    For normal, utility, and acrobatic category multiengine turbojet
airplanes of more than 6,000 pounds maximum weight and commuter
category airplanes, the takeoff path is as follows:
* * * * *
    11. Amend Sec.  23.59 by revising the introductory text to read as
follows:

Sec.  23.59  Takeoff distance and takeoff run.

    For normal, utility, and acrobatic category multiengine turbojet
airplanes of more than 6,000 pounds maximum weight and commuter
category airplanes, the takeoff distance and, at the option of the
applicant, the takeoff run, must be determined.
* * * * *
    12. Amend Sec.  23.61 by revising the introductory text to read as
follows:

Sec.  23.61  Takeoff flight path.

    For normal, utility, and acrobatic category multiengine turbojet
airplanes of more than 6,000 pounds maximum weight and commuter
category airplanes, the takeoff flight path must be determined as
follows:
* * * * *
    13. Amend Sec.  23.63 by revising the introductory text of
paragraphs (c) and (d) to read as follows:

Sec.  23.63  Climb: General.

* * * * *
    (c) For reciprocating engine-powered airplanes of more than 6,000
pounds maximum weight, single-engine turbines, and multiengine turbine
airplanes of 6,000 pounds or less maximum weight in the normal,
utility, and acrobatic category, compliance must be shown at weights as
a function of airport altitude and ambient temperature, within the
operational limits established for takeoff and landing, respectively,
with--
* * * * *
    (d) For multiengine turbine airplanes over 6,000 pounds maximum
weight in the normal, utility, and acrobatic category and commuter
category airplanes, compliance must be shown at weights as a function
of airport altitude

and ambient temperature within the operational limits established for
takeoff and landing, respectively, with--
* * * * *
    14. Amend Sec.  23.67 by:
    a. Revising paragraph (b) introductory text and (b)(1) introductory
text;
    b. Redesignating paragraph (c) as paragraph (d)
    c. Revising newly redesignated paragraph (d) introductory text,
paragraph (d)(2) introductory text, paragraph (d)(3) introductory text,
and paragraph (d)(4) introductory text; and
    d. Adding new paragraph (c).
    The revisions and addition read as follows:

Sec.  23.67  Climb: One-engine inoperative.

* * * * *
    (b) For normal, utility, and acrobatic category reciprocating
engine-powered airplanes of more than 6,000 pounds maximum weight, and
turbopropeller-powered airplanes in the normal, utility, and acrobatic
category--
    (1) The steady gradient of climb at an altitude of 400 feet above
the takeoff may be no less than 1 percent with the--
* * * * *
    (c) For normal, utility, and acrobatic category turbojet engine-
powered airplanes of 6,000 pounds or less maximum weight--
    (1) The steady gradient of climb at an altitude of 400 feet above
the takeoff may be no less than 1.2 percent with the--
    (i) Critical engine inoperative;
    (ii) Remaining engine(s) at takeoff power;
    (iii) Landing gear retracted;
    (iv) Wing flaps in the takeoff position(s); and
    (v) Climb speed equal to that achieved at 50 feet in the
demonstration of Sec.  23.53.
    (2) The steady gradient of climb may not be less than 0.75 percent
at an altitude of 1,500 feet above the takeoff surface, or landing
surface, as appropriate, with the--
    (i) Critical engine inoperative:
    (ii) Remaining engine(s) at not more than maximum continuous power;
    (iii) Landing gear retracted;
    (iv) Wing flaps retracted; and
    (v) Climb speed not less than 1.2 VS1.
    (d) For turbojet powered airplanes over 6,000 pounds maximum weight
in the normal, utility and acrobatic category and commuter category
airplanes, the following apply:
* * * * *
    (2) Takeoff; landing gear retracted. The steady gradient of climb
at an altitude of 400 feet above the takeoff surface must be at least
2.0 percent of two-engine airplanes, 2.3 percent for three-engine
airplanes, and 2.6 percent for four-engine airplanes with--
* * * * *
    (3) Enroute. The steady gradient of climb at an altitude of 1,500
feet above the takeoff or landing surface, as appropriate, must be at
least 1.2 percent for two-engine airplanes, 1.5 percent for three-
engine airplanes, and 1.7 percent for four-engine airplanes with--
* * * * *
    (4) Discontinued approach. The steady gradient of climb at an
altitude of 400 feet above the landing surface must be at least 2.1
percent for two-engine airplanes, 2.4 percent for three-engine
airplanes, and 2.7 percent for four-engine airplanes, with--
* * * * *
    15. Revise Sec.  23.73 to read as follows:

Sec.  23.73  Reference landing approach speed.

    (a) For normal, utility, and acrobatic category reciprocating
engine-powered airplanes of 6,000 pounds or less maximum weight, the
reference landing approach speed, VREF, may not be less than
the greater of VMC, determined in Sec.  23.149(b) with the
wing flaps in the most extended takeoff position, and 1.3
VS1.
    (b) For normal, utility, and acrobatic category turbine powered
airplanes of 6,000 pounds or less maximum weight, turboprops of more
than 6,000 pounds maximum weight, and reciprocating engine-powered
airplanes of more than 6,000 pounds maximum weight, the reference
landing approach speed, VREF, may not be less than the
greater of VMC, determined in Sec.  23.149(c), and 1.3
VS1.
    (c) For normal, utility, and acrobatic category turbojet engine-
powered airplanes of more than 6,000 pounds maximum weight and commuter
category airplanes, the reference landing approach speed,
VREF, may not be less than the greater of 1.05
VMC, determined in Sec.  23.149(c), and 1.3 VS1.
    16. Amend Sec.  23.77 by revising the introductory text of
paragraphs (b) and (c) to read as follows:

Sec.  23.77  Balked landing.

* * * * *
    (b) Each normal, utility, and acrobatic category reciprocating
engine-powered and single engine turbine powered airplane of more than
6,000 pounds maximum weight, and multiengine turbine engine-powered
airplane of 6,000 pounds or less maximum weight in the normal, utility,
and acrobatic category must be able to maintain a steady gradient of
climb of at least 2.5 percent with--
* * * * *
    (c) Each normal, utility, and acrobatic multiengine turbine powered
airplane over 6,000 pounds maximum weight and each commuter category
airplane must be able to maintain a steady gradient of climb of at
least 3.2 percent with--
* * * * *
    17. Amend Sec.  23.175 by adding a new paragraph (b)(3) to read as
follows:

Sec.  23.175  Demonstration of static longitudinal stability.

* * * * *
    (b) * * *
    (3) Maximum speed for stability characteristics, VFC/MFC.
VFC/MFC may not be less than a speed midway
between VMO/MMO and VDF/MDF
except that, for altitudes where Mach number is the limiting factor,
MFC need not exceed the Mach number at which effective speed
warning occurs.
* * * * *
    18. Amend Sec.  23.177 by revising paragraphs (a), (b), and (d) to
read as follows:

Sec.  23.177  Static directional and lateral stability.

    (a)(1) The static directional stability, as shown by the tendency
to recover from a wings level sideslip with the rudder free, must be
positive for any landing gear and flap position appropriate to the
takeoff, climb, cruise, approach, and landing configurations. This must
be shown with symmetrical power up to maximum continuous power, and at
speeds from 1.2 VS1 up to the landing gear or wing flap
operating limit speeds, or VNO or VFC/
MFC, whichever is appropriate.
    (2) The angle of sideslip for these tests must be appropriate to
the type of airplane. The rudder pedal force may not reverse at larger
angles of sideslip, up to that at which full rudder is used or a
control force limit in Sec.  23.143 is reached, whichever occurs first,
and at speeds from 1.2 VS1 to VO.
    (b)(1) The static lateral stability, as shown by the tendency to
raise the low wing in a sideslip with the aileron controls free, may
not be negative for any landing gear and flap position appropriate to
the takeoff, climb, cruise, approach, and landing configurations. This
must be shown with symmetrical power from idle up to 75 percent of
maximum continuous power at speeds from 1.2 VS1 in the
takeoff configuration(s) and at speeds from 1.3 VS1 in other
configurations, up to the maximum allowable airspeed for the
configuration being investigated, (VFE,

VLE, VNO, VFC/MFC,
whichever is appropriate) in the takeoff, climb, cruise, descent, and
approach configurations. For the landing configuration, the power must
be that necessary to maintain a 3-degree angle of descent in
coordinated flight.
    (2) The static lateral stability may not be negative at 1.2
VS1 in the takeoff configuration, or at 1.3 VS1
in other configurations.
    (3) The angle of sideslip for these tests must be appropriate to
the type of airplane, but in no case may the constant heading sideslip
angle be less than that obtainable with a 10 degree bank or, if less,
the maximum bank angle obtainable with full rudder deflection or 150
pound rudder force.
* * * * *
    (d)(1) In straight, steady slips at 1.2 VS1 for any
landing gear and flap position appropriate to the takeoff, climb,
cruise, approach, and landing configurations, and for any symmetrical
power conditions up to 50 percent of maximum continuous power, the
aileron and rudder control movements and forces must increase steadily,
but not necessarily in constant proportion, as the angle of sideslip is
increased up to the maximum appropriate to the type of airplane.
    (2) At larger slip angles, up to the angle at which the full rudder
or aileron control is used or a control force limit contained in Sec.
23.143 is reached, the aileron and rudder control movements and forces
may not reverse as the angle of sideslip is increased.
    (3) Rapid entry into, and recovery from, a maximum sideslip
considered appropriate for the airplane may not result in
uncontrollable flight characteristics.
    19. Amend Sec.  23.181 by revising paragraph (b) to read as
follows:

Sec.  23.181  Dynamic stability.

* * * * *
    (b) Any combined lateral-directional oscillations (``Dutch roll'')
occurring between the stalling speed and the maximum allowable speed
appropriate to the configuration of the airplane with the primary
controls in both free and fixed position, must be damped to 1/10
amplitude in:
    (1) Seven (7) cycles below 18,000 feet, and
    (2) Thirteen (13) cycles from 18,000 feet to the certified maximum
altitude.
* * * * *
    20. Amend Sec.  23.201 by revising paragraphs (d) and (e) and by
adding a new paragraph (f) to read as follows:

Sec.  23.201  Wings level stall.

* * * * *
    (d) During the entry into and the recovery from the maneuver, it
must be possible to prevent more than 15 degrees of roll or yaw by the
normal use of controls except as provided for in paragraph (e) of this
section.
    (e) For airplanes approved with a maximum operating altitude above
25,000 feet, during the entry into and the recovery from stalls
performed above 25,000 feet, it must be possible to prevent more than
25 degrees of roll or yaw by the normal use of controls.
    (f) Compliance with the requirements of this section must be shown
under the following conditions:
    (1) Wing flaps: Retracted, fully extended, and each intermediate
normal operating position, as appropriate for the phase of flight.
    (2) Landing gear: Retracted and extended as appropriate for the
altitude.
    (3) Cowl flaps: Appropriate to configuration.
    (4) Spoilers/speedbrakes: Retracted and extended unless they have
little to no effect at low speeds.
    (5) Power:
    (i) Power/Thrust off; and
    (ii) For reciprocating engine powered airplanes: 75 percent maximum
continuous power. However, if the power-to-weight ratio at 75 percent
of maximum continuous power results in nose-high attitudes exceeding 30
degrees, the test must be carried out with the power required for level
flight in the landing configuration at maximum landing weight and a
speed of 1.4 VSO, except that the power may not be less than
50 percent of maximum continuous power; or
    (iii) For turbine engine powered airplanes: The maximum engine
thrust, except that it need not exceed the thrust necessary to maintain
level flight at 1.6 VS1 (where VS1 corresponds to
the stalling speed with flaps in the approach position, the landing
gear retracted, and maximum landing weight).
    (6) Trim at 1.5 VS1 or the minimum trim speed, whichever
is higher.
    (7) Propeller: Full increase r.p.m. position for the power off
condition.
    21. Amend Sec.  23.203 by revising paragraph (c) to read as
follows:

Sec.  23.203  Turning flight and accelerated turning stalls.

* * * * *
    (c) Compliance with the requirements of this section must be shown
under the following conditions:
    (1) Wings flaps: Retracted, fully extended, and each intermediate
normal operating position as appropriate for the phase of flight.
    (2) Landing gear: Retracted and extended as appropriate for the
altitude.
    (3) Cowl flaps: Appropriate to configuration.
    (4) Spoilers/speedbrakes: Retracted and extended unless they have
little to no effect at low speeds.
    (5) Power:
    (i) Power/Thrust off; and
    (ii) For reciprocating engine powered airplanes: 75 percent maximum
continuous power. However, if the power-to-weight ratio at 75 percent
of maximum continuous power results in nose-high attitudes exceeding 30
degrees, the test may be carried out with the power required for level
flight in the landing configuration at maximum landing weight and a
speed of 1.4 VSO, except that the power may not be less than
50 percent of maximum continuous power; or
    (iii) For turbine engine powered airplanes: The maximum engine
thrust, except that it need not exceed the thrust necessary to maintain
level flight at 1.6 VS1 (where VS1 corresponds to
the stalling speed with flaps in the approach position, the landing
gear retracted, and maximum landing weight).
    (6) Trim: The airplane trimmed at 1.5 VS1.
    (7) Propeller: Full increase rpm position for the power off
condition.
    22. Revise Sec.  23.251 to read as follows:

Sec.  23.251  Vibration and buffeting.

    (a) There may be no vibration or buffeting severe enough to result
in structural damage, and each part of the airplane must be free from
excessive vibration, under any appropriate speed and power conditions
up to VD/MD, or VDF/MDF for
turbojets. In addition, there may be no buffeting in any normal flight
condition, including configuration changes during cruise, severe enough
to interfere with the satisfactory control of the airplane or cause
excessive fatigue to the flight crew. Stall warning buffeting within
these limits is allowable.
    (b) There may be no perceptible buffeting condition in the cruise
configuration in straight flight at any speed up to VMO/
MMO, except stall buffeting, which is allowable.
    (c) For airplanes with MD greater than M 0.6 and a
maximum operating altitude greater than 25,000 feet, the positive
maneuvering load factors at which the onset of perceptible buffeting
occurs must be determined with the airplane in the cruise configuration
for the ranges of airspeed or Mach number, weight, and altitude for
which the airplane is to be certificated. The envelopes of load factor,
speed, altitude, and weight must provide a sufficient range of speeds
and load factors for

normal operations. Probable inadvertent excursions beyond the
boundaries of the buffet onset envelopes may not result in unsafe
conditions.
    23. Amend Sec.  23.253 by revising paragraphs (b)(1) and (b)(2),
and by adding a new paragraph (b)(3) to read as follows:

Sec.  23.253  High speed characteristics.

* * * * *
    (b) * * *
    (1) Exceptional piloting strength or skill;
    (2) Exceeding VD/MD, or VDF/
MDF for turbojet, the maximum speed shown under Sec.
23.251, or the structural limitations; and
    (3) Buffeting that would impair the pilot's ability to read the
instruments or to control the airplane for recovery.
* * * * *
    24. Section 23.255 is added to subpart B to read as follows:

Sec.  23.255  Out of trim characteristics.

    For airplanes with an MD greater than M 0.6 and that
incorporate a trimmable horizontal stabilizer, the following
requirements for out-of-trim characteristics apply:
    (a) From an initial condition with the airplane trimmed at cruise
speeds up to VMO/MMO, the airplane must have
satisfactory maneuvering stability and controllability with the degree
of out-of-trim in both the airplane nose-up and nose-down directions,
which results from the greater of the following:
    (1) A three-second movement of the longitudinal trim system at its
normal rate for the particular flight condition with no aerodynamic
load (or an equivalent degree of trim for airplanes that do not have a
power-operated trim system), except as limited by stops in the trim
system, including those required by Sec.  23.655(b) for adjustable
stabilizers; or
    (2) The maximum mis-trim that can be sustained by the autopilot
while maintaining level flight in the high speed cruising condition.
    (b) In the out-of-trim condition specified in paragraph (a) of this
section, when the normal acceleration is varied from +l g to the
positive and negative values specified in paragraph (c) of this
section, the following apply:
    (1) The stick force versus g curve must have a positive slope at
any speed up to and including VFC/MFC; and
    (2) At speeds between VFC/MFC and
VDF/MDF, the direction of the primary
longitudinal control force may not reverse.
    (c) Except as provided in paragraphs (d) and (e) of this section,
compliance with the provisions of paragraph (a) of this section must be
demonstrated in flight over the acceleration range as follows:
    (1) -1 g to +2.5g; or
    (2) 0 g to 2.0g, and extrapolating by an acceptable method to -1g
and +2.5g.
    (d) If the procedure set forth in paragraph (c)(2) of this section
is used to demonstrate compliance and marginal conditions exist during
flight test with regard to reversal of primary longitudinal control
force, flight tests must be accomplished from the normal acceleration
at which a marginal condition is found to exist to the applicable limit
specified in paragraph (b)(1) of this section.
    (e) During flight tests required by paragraph (a) of this section,
the limit maneuvering load factors, prescribed in Sec. Sec.  23.333(b)
and 23.337, need not be exceeded. In addition, the entry speeds for
flight test demonstrations at normal acceleration values less than 1g
must be limited to the extent necessary to accomplish a recovery
without exceeding VDF/MDF.
    (f) In the out-of-trim condition specified in paragraph (a) of this
section, it must be possible from an overspeed condition at
VDF/MDF to produce at least 1.5g for recovery by
applying not more than 125 pounds of longitudinal control force using
either the primary longitudinal control alone or the primary
longitudinal control and the longitudinal trim system. If the
longitudinal trim is used to assist in producing the required load
factor, it must be shown at VDF/MDF that the
longitudinal trim can be actuated in the airplane nose-up direction
with the primary surface loaded to correspond to the least of the
following airplane nose-up control forces:
    (1) The maximum control forces expected in service, as specified in
Sec. Sec.  23.301 and 23.397.
    (2) The control force required to produce 1.5g.
    (3) The control force corresponding to buffeting or other phenomena
of such intensity that it is a strong deterrent to further application
of primary longitudinal control force.
    25. Amend Sec.  23.561 by adding new paragraphs (e)(1) and (e)(2)
to read as follows:

Sec.  23.561  General.

* * * * *
    (e) * * *
    (1) For turbojet engines mounted inside the fuselage, aft of the
cabin, it must be shown by test or analysis that the engine and
attached accessories, and the engine mounting structure--
    (i) Can withstand a forward acting static ultimate inertia load
factor of 18.0g plus the maximum takeoff engine thrust; or
    (ii) The airplane structure is designed to deflect the engine and
its attached accessories away from the cabin should the engine mounts
fail.
    (2) [Reserved]
    26. Amend Sec.  23.562 by revising paragraphs (a) introductory
text, (b) introductory text, and (c)(5)(ii) to read as follows:

Sec.  23.562  Emergency landing dynamic conditions.

    (a) Each seat/restraint system for use in a normal, utility, or
acrobatic category airplane, or in a commuter category turbojet powered
airplane, must be designed to protect each occupant during an emergency
landing when--
* * * * *
    (b) Except for those seat/restraint systems that are required to
meet paragraph (d) of this section, each seat/restraint system for crew
or passenger occupancy in a normal, utility, or acrobatic category
airplane, or in a commuter category turbojet powered airplane, must
successfully complete dynamic tests or be demonstrated by rational
analysis supported by dynamic tests, in accordance with each of the
following conditions. These tests must be conducted with an occupant
simulated by an anthropomorphic test dummy (ATD) defined by 49 CFR part
572, subpart B, or an FAA-approved equivalent, with a nominal weight of
170 pounds and seated in the normal upright position.
* * * * *
    (c) * * *
    (5) * * *
    (ii) The value of HIC is defined as--
    [GRAPHIC] [TIFF OMITTED] TP17AU09.000

Where:

t1 is the initial integration time, expressed in seconds,
t2 is the final integration time, expressed in seconds,
and a(t) is the total acceleration vs. time curve for the head
expressed as a multiple of g (units of gravity).
* * * * *
    27. Amend Sec.  23.571 by adding a new paragraph (d) to read as
follows:

Sec.  23.571  Metallic pressurized cabin structures.

* * * * *
    (d) If certification for operation above 41,000 feet is requested,
a damage tolerance evaluation of the fuselage pressure boundary per
Sec.  23.573(b) must be conducted and the evaluation must factor in the
environmental requirements of Sec.  23.841.
    28. Amend Sec.  23.573 by adding a new paragraph (c) to read as
follows:

Sec.  23.573  Damage tolerance and fatigue evaluation of structure.

* * * * *
    (c) If certification for operation above 41,000 feet is requested,
the damage tolerance evaluation of this paragraph for the fuselage
pressure boundary must factor in the requirements of Sec.  23.841.
    29. Amend Sec.  23.574 by adding a new paragraph (c) to read as
follows:

Sec.  23.574  Metallic damage tolerance and fatigue evaluation of
commuter category airplanes.

* * * * *
    (c) If certification for operation above 41,000 feet is requested,
the damage tolerance evaluation of this paragraph for the fuselage
pressure boundary must factor in the requirements of Sec.  23.841.
    30. Amend Sec.  23.629 by revising paragraphs (b)(1), (b)(3),
(b)(4), and (c) to read as follows:

Sec.  23.629  Flutter.

* * * * *
    (b) * * *
    (1) Proper and adequate attempts to induce flutter have been made
within the speed range up to VD/MD;
* * * * *
    (3) A proper margin of damping exists at VD/
MD, or VDF/MDF for turbojet airplanes;
and
    (4) As VD/MD (or VDF/
MDF for turbojet airplanes) is approached, there may not be
a large or rapid reduction in damping.
    (c) Any rational analysis used to predict freedom from flutter,
control reversal and divergence must cover all speeds up to 1.2
VD/MD, or 1.2 VDF/MDF for
turbojet airplanes.
* * * * *
    31. Amend Sec.  23.703 by revising the introductory text and
paragraph (b) to read as follows:

Sec.  23.703  Takeoff warning system.

    For all airplanes with a maximum weight more than 6,000 pounds and
all turbojet airplanes, unless it can be shown that a lift or
longitudinal trim device that affects the takeoff performance of the
airplane would not give an unsafe takeoff configuration when selected
out of an approved takeoff position, a takeoff warning system must be
installed and meet the following requirements:
* * * * *
    (b) For the purpose of this section, an unsafe takeoff
configuration is the inability to rotate or the inability to prevent an
immediate stall after rotation.
    32. Amend Sec.  23.735 by revising paragraph (e) to read as
follows:

Sec.  23.735  Brakes.

* * * * *
    (e) For airplanes required to meet Sec.  23.55, the rejected
takeoff 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 conservative rational analysis of the sequence of events expected
during a rejected takeoff at the design takeoff weight.
    (2) Instead of a rational analysis, the kinetic energy absorption
requirements for each main wheel brake assembly may be derived from the
following formula--

KE = 0.0443 WV\2\/N

Where:

KE = Kinetic energy per wheel (ft.-lbs.);
W = Design takeoff weight (lbs.);
V = Ground speed, in knots, associated with the maximum value of
V1 selected in accordance with Sec.  23.51(c)(1);
N = Number of main wheels with brakes.

    33. Amend Sec.  23.777 by revising paragraph (d) to read as
follows:

Sec.  23.777  Cockpit controls.

* * * * *
    (d) When separate and distinct control levers are co-located (such
as located together on the pedestal), the control location order from
left to right must be power (thrust) lever, propeller (rpm control),
and mixture control (condition lever and fuel cut-off for turbine-
powered airplanes). Power (thrust) levers must be at least one inch
higher or longer than propeller (rpm control) or mixture controls to
make them more prominent. Carburetor heat or alternate air control must
be to the left of the throttle or at least eight inches from the
mixture control when located other than on a pedestal. Carburetor heat
or alternate air control, when located on a pedestal, must be aft or
below the power (thrust) lever. Supercharger controls must be located
below or aft of the propeller controls. Airplanes with tandem seating
or single-place airplanes may utilize control locations on the left
side of the cabin compartment; however, location order from left to
right must be power (thrust) lever, propeller (rpm control), and
mixture control.
* * * * *
    34. Amend Sec.  23.807 by adding a new paragraph (e)(3) to read as
follows:

Sec.  23.807  Emergency exits.

* * * * *
    (e) * * *
    (3) In lieu of paragraph (e)(2) of this section, if any side exit
or exits cannot be above the waterline, a device may be placed at each
of such exit(s) prior to ditching. This device must slow the inflow of
water when such exit(s) is opened with the airplane in a ditching
emergency. For commuter category airplanes, the clear opening of such
exit or exits must meet the requirements defined in paragraph (d) of
this section.
    35. Amend Sec.  23.831 by adding paragraphs (c) and (d) to read as
follows:

Sec.  23.831  Ventilation.

* * * * *
    (c) For turbojet powered pressurized airplanes, under normal
operating conditions and in the event of any probable failure
conditions of any system which would adversely affect the ventilating
air, the ventilation system must provide reasonable passenger comfort.
The ventilation system must also provide a sufficient amount of
uncontaminated air to enable the crew members to perform their duties
without undue discomfort or fatigue and to provide reasonable passenger
comfort. For normal operating conditions, the ventilation system must
be designed to provide each occupant with at least 0.55 pounds of fresh
air per minute. In the event of the loss of one source of fresh air,
the supply of fresh airflow must not be less than 0.4 pounds per minute
for any period exceeding five minutes.
    (d) Other probable and improbable Environmental Control System
failure conditions that adversely affect the passenger and crew
compartment environmental conditions may not affect crew performance so
as to result in a hazardous condition, and no occupant shall sustain
permanent physiological harm.

    36. Amend Sec.  23.841 by revising paragraphs (a) and (b)(6), and
by adding paragraphs (c), (d), and (e) to read as follows:

Sec.  23.841  Pressurized cabins.

    (a) If certification for operation above 25,000 feet is requested,
the airplane must be able to maintain a cabin pressure altitude of not
more than 15,000 feet, in the event of any probable failure condition
in the pressurization system. During the decompression, the cabin
altitude shall not exceed 15,000 feet for more than 10 seconds and
25,000 feet for any duration.
    (b) * * *
    (6) Warning indication at the pilot station to indicate when the
safe or preset pressure differential is exceeded and when a cabin
pressure altitude of 10,000 feet is exceeded. The 10,000 foot cabin
altitude warning may be increased up to 15,000 feet for operations from
high altitude airfields (10,000 to 15,000 feet) provided:
    (i) The landing or the take off modes (normal or high altitude) are
clearly indicated to the flight crew.
    (ii) Selection of normal or high altitude airfield mode requires no
crew action beyond normal pressurization system operation.
    (iii) The pressurization system is designed to ensure cabin
altitude does not exceed 10,000 feet when in flight above flight level
(FL) 250.
    (iv) The pressurization system and cabin altitude warning system is
designed to ensure cabin altitude warning at 10,000 feet when in flight
above FL250.
* * * * *
    (c) If certification for operation above 41,000 feet and not more
than 45,000 feet is requested,
    (1) The airplane must prevent cabin pressure altitude from
exceeding the following after decompression from any probable
pressurization system failure in conjunction with any undetected,
latent pressurization system failure condition:
    (i) If depressurization analysis shows that the cabin altitude does
not exceed 25,000 feet, the pressurization system must prevent the
cabin altitude from exceeding the cabin altitude-time history shown in
Figure 1 of this section.
    (ii) Maximum cabin altitude is limited to 30,000 feet. If cabin
altitude exceeds 25,000 feet, the maximum time the cabin altitude may
exceed 25,000 feet is 2 minutes; time starting when the cabin altitude
exceeds 25,000 feet and ending when it returns to 25,000 feet.
    (2) The airplane must prevent cabin pressure altitude from
exceeding the following after decompression from any single
pressurization system failure in conjunction with any probable fuselage
damage:
    (i) If depressurization analysis shows that the cabin altitude does
not exceed 37,000 feet, the pressurization system must prevent the
cabin altitude from exceeding the cabin altitude-time history shown in
Figure 2 of this section.
    (ii) Maximum cabin altitude is limited to 40,000 feet. If cabin
altitude exceeds 37,000 feet, the maximum time the cabin altitude may
exceed 25,000 feet is 2 minutes; time starting when the cabin altitude
exceeds 25,000 feet and ending when it returns to 25,000 feet.
    (3) In showing compliance with paragraphs (c)(1) and (c)(2) of this
section, it may be assumed that an emergency descent is made by an
approved emergency procedure. A 17-second crew recognition and reaction
time must be applied between cabin altitude warning and the initiation
of an emergency descent. Fuselage structure, engine and system failures
are to be considered in evaluating the cabin decompression.
[GRAPHIC] [TIFF OMITTED] TP17AU09.001

    Note: For Figure 1, time starts at the moment cabin altitude
exceeds 10,000 feet during decompression.

[GRAPHIC] [TIFF OMITTED] TP17AU09.002

    Note: For Figure 2, time starts at the moment cabin altitude
exceeds 10,000 feet during decompression.

    (d) If certification for operation above 45,000 feet and not more
than 51,000 feet is requested--
    (1) Pressurized cabins must be equipped to provide a cabin pressure
altitude of not more than 8,000 feet at the maximum operating altitude
of the airplane under normal operating conditions.
    (2) The airplane must prevent cabin pressure altitude from
exceeding the following after decompression from any failure condition
not shown to be extremely improbable:
    (i) Twenty-five thousand (25,000) feet for more than 2 minutes, or
    (ii) Forty thousand (40,000) feet for any duration.
    (3) Fuselage structure, engine and system failures are to be
considered in evaluating the cabin decompression.
    (4) In addition to the cabin altitude indicating means in (b)(6) of
this section, an aural or visual signal must be provided to warn the
flight crew when the cabin pressure altitude exceeds 10,000 feet.
    (5) The sensing system and pressure sensors necessary to meet the
requirements of (b)(5), (b)(6), and (d)(4) of this section and Sec.
23.1447(e), must, in the event of low cabin pressure, actuate the
required warning and automatic presentation devices without any delay
that would significantly increase the hazards resulting from
decompression.
    (e) If certification for operation above 41,000 feet is requested,
additional damage-tolerance requirements are necessary to prevent
fatigue damage that could result in a loss of pressure that exceeds the
requirements of paragraphs (c) and (d) of this section. Sufficient full
scale fatigue test evidence must be provided to demonstrate that this
type of pressure loss due to fatigue cracking will not occur within the
Limit of Validity of the Maintenance program for the airplane. In
addition, a damage tolerance evaluation of the fuselage pressure
boundary must be performed assuming visually detectable cracks and the
maximum damage size for which the requirements of paragraphs (c) and
(d) of this section can be met. Based on this evaluation, inspections
or other procedures must be established and included in the Limitations
Section of the Instructions for Continued Airworthiness required by
Sec.  23.1529.
    37. Amend Sec.  23.853 by revising paragraph (d)(2) to read as
follows:

Sec.  23.853  Passenger and crew compartment interiors.

* * * * *
    (d) * * *
    (2) Lavatories must have ``No Smoking'' or ``No Smoking in
Lavatory'' placards located conspicuously on each side of the entry
door.
* * * * *
    38. Add a new Sec.  23.856 to read as follows:

Sec.  23.856  Thermal/Acoustic insulation materials.

    Thermal/acoustic insulation material installed in the fuselage must
meet the flame propagation test requirements of part II of Appendix F
to this part, or other approved equivalent test requirements. This
requirement does not apply to ``small parts,'' as defined in part I of
Appendix F of this part.
    39. Amend Sec.  23.903 by revising paragraph (b)(2) to read as
follows:

Sec.  23.903  Engines.

* * * * *
    (b) * * *
    (2) For engines embedded in the fuselage behind the cabin, the
effects of a fan exiting forward of the inlet case (fan disconnect)
must be addressed, the passengers must be protected, and the airplane
must have the ability to maintain controlled flight and landing.
* * * * *
    40. Amend Sec.  23.1141 by adding a new paragraph (h) to read as
follows:

Sec.  23.1141  Powerplant controls: General.

* * * * *
    (h) Electronic engine control system installations must meet the
requirements of Sec.  23.1309.
    41. Amend Sec.  23.1165 by revising paragraph (f) to read as
follows:

Sec.  23.1165  Engine ignition systems.

* * * * *
    (f) In addition, for commuter category airplanes, each turbine
engine ignition system must be an essential electrical load.
    42. Amend Sec.  23.1193 by revising paragraph (g) to read as
follows:

Sec.  23.1193  Cowling and nacelle.

* * * * *
    (g) In addition, for all turbojet airplanes and commuter category
airplanes, the airplane must be designed so that no fire originating in
any engine compartment can enter, either through openings or by burn
through, any other region where it would create additional hazards.
    43. Amend Sec.  23.1195 by revising the introductory text of
paragraph (a) and by revising paragraph (a)(2) to read as follows:

Sec.  23.1195  Fire extinguishing systems.

    (a) For all turbojet airplanes and commuter category airplanes,
fire extinguishing systems must be installed and compliance shown with
the following:
* * * * *
    (2) The fire extinguishing system, the quantity of the
extinguishing agent, the rate of discharge, and the discharge
distribution must be adequate to extinguish fires. An individual ``one
shot'' system may be used, except for engine(s) embedded in the
fuselage, where a ``two-shot'' system is required.
* * * * *
    44. Amend Sec.  23.1197 by revising the introductory text to read
as follows:

Sec.  23.1197  Fire extinguishing agents.

    For all turbojet airplanes and commuter category airplanes, the
following applies:
* * * * *
    45. Amend Sec.  23.1199 by revising the introductory text to read
as follows:

Sec.  23.1199  Extinguishing agent containers.

    For all turbojet airplanes and commuter category airplanes, the
following applies:
* * * * *
    46. Amend Sec.  23.1201 by revising the introductory text to read
as follows:

Sec.  23.1201  Fire extinguishing systems materials.

    For all turbojet airplanes and commuter category airplanes, the
following apply:
* * * * *
    47. Revise Sec.  23.1301 by revising paragraphs (b) and (c) and by
removing paragraph (d) to read as follows:

Sec.  23.1301  Function and installation.

* * * * *
    (b) Be labeled as to its identification, function, or operating
limitations, or any applicable combination of these factors; and
    (c) Be installed according to limitations specified for that
equipment.
    48. Amend Sec.  23.1303 by revising paragraph (c) to read as
follows:

Sec.  23.1303  Flight and navigation instruments.

* * * * *
    (c) A magnetic direction indicator.
* * * * *
    49. Amend Sec.  23.1305 by adding a new paragraph (f) to read as
follows:

Sec.  23.1305  Powerplant instruments.

* * * * *
    (f) Powerplant indicators must either provide trend or rate-of-
change information, or have the ability to:
    (1) Allow the pilot to assess necessary trend information quickly,
including if and when this information is needed during engine restart;
    (2) Allow the pilot to assess how close the indicated parameter is
relative to a limit;
    (3) Forewarn the pilot before the parameter reaches an operating
limit; and
    (4) For multiengine airplanes, allow the pilot to quickly and
accurately compare engine-to-engine data.
    50. Revise Sec.  23.1307 to read as follows:

Sec.  23.1307  Miscellaneous equipment.

    The equipment necessary for an airplane to operate at the maximum
operating altitude and in the kinds of operations (e.g., part 91, part
135) and meteorological conditions for which certification is requested
and is approved in accordance with Sec.  23.1559 must be included in
the type design.
    51. Revise Sec.  23.1309 to read as follows:

Sec.  23.1309  Equipment, systems, and installations.

    The requirements of this section, except as identified below, are
applicable, in addition to specific design requirements of part 23, to
any equipment or system as installed in the airplane. This section is a
regulation of general requirements. It does not supersede any specific
requirements contained in another section of part 23. This section
should be used to determine software and hardware development assurance
levels. This section does not apply to the performance, flight
characteristics requirements of subpart B of this part, and structural
loads and strength requirements of subparts C and D of this part, but
it does apply to any system on which compliance with the requirements
of subparts B, C, D, and E of this part are based. The flight structure
such as wing, empennage, control surfaces and their simple, or simple
and conventional systems, the fuselage, engine mounting, and landing
gear and their related primary attachments are excluded. For example,
it does not apply to an airplane's inherent stall characteristics or
their evaluation of Sec.  23.201, but it does apply to a stick pusher
(stall barrier) system installed to attain compliance with Sec.
23.201.
    (a) The airplane equipment and systems must be designed and
installed so that:
    (1) Those required for type certification or by operating rules, or
whose improper functioning would reduce safety, perform as intended
under the airplane operating and environmental conditions, including
radio frequency energy and the effects (both direct and indirect) of
lightning strikes.
    (2) Those required for type certification or by operating rules and
other equipment and systems do not adversely affect the safety of the
airplane or its occupants, or the proper functioning of those covered
by paragraph (a)(1) of this section.
    (3) For minor, major, hazardous, or catastrophic failure
condition(s), the results of certification testing must not be
inconsistent with the results of the safety analysis process.
    (b) The airplane systems and associated components for the
appropriate classes of airplane, considered separately and in relation
to other systems, must be designed and installed so that:
    (1) Each catastrophic failure condition is extremely improbable and
does not result from a single failure;
    (2) Each hazardous failure condition is extremely remote;
    (3) Each major failure condition is remote; and
    (4) Each failure condition meets the relationship among airplane
classes, probabilities, severity of failure condition(s), and software
and complex hardware development assurance levels shown in Appendix K
of this part.
    (5) Compliance with the requirements of paragraph (b)(2) of this
section may be shown by analysis and, where necessary, by appropriate
ground, flight, or simulator tests. The analysis must consider--
    (i) Possible modes of failure, including malfunctions and damage
from external sources;
    (ii) The probability of multiple failures and the probability of
undetected faults;
    (iii) The resulting effects of the airplane and occupants,
considering the stage of flight and operating conditions; and
    (iv) The crew warning cues, corrective action required, and the
crew's capability of determining faults.
    (c) Functional failure condition(s) that are classified as minor do
not require a quantitative analysis, but verification by a design and
installation appraisal is required.
    (d) Systems with major failure condition(s)--
    (1) May be verified by a qualitative analysis, if the systems are
simple, simple and conventional, or conventional and redundant.
    (2) Must be verified by a qualitative and quantitative analysis, if
the systems

do not meet the condition(s) prescribed in paragraph (d)(1) of this
section.
    (e) Systems with hazardous or catastrophic failure condition(s)--
    (1) May be verified by a qualitative and quantitative analysis, if
the systems are simple and conventional.
    (2) Must be verified by a qualitative and quantitative analysis if
the systems are not simple and conventional.
    (f) Information concerning an unsafe system operating condition(s)
must be provided to the crew to enable them to take appropriate
corrective action. A warning indication must be provided if immediate
corrective action is required. Systems and controls, including
indications and annunciations must be designed to minimize crew errors,
which could create additional hazards.
    52. Add a new Sec.  23.1310 to read as follows:

Sec.  23.1310  Power source capacity and distribution.

    (a) Each item of equipment, each system, and each installation
whose functioning is required by this chapter and that requires a power
supply is an ``essential load'' on the power supply. The power sources
and the system must be able to supply the following power loads in
probable operating combinations and for probable durations:
    (1) Loads connected to the power distribution system with the
system functioning normally.
    (2) Essential loads after failure of--
    (i) Any one engine on two-engine airplanes, or
    (ii) Any two engines on an airplane with three or more engines, or
    (iii) Any power converter or energy storage device.
    (3) Essential loads for which an alternate source of power is
required, as applicable, by the operating rules of this chapter, after
any failure or malfunction in any one power supply system, distribution
system, or other utilization system.
    (b) In determining compliance with paragraph (a)(2) of this
section, the power loads may be assumed to be reduced under a
monitoring procedure consistent with safety in the kinds of operations
authorized. Loads not required in controlled flight need not be
considered for the two-engine-inoperative condition on airplanes with
three or more engines.
    53. Amend Sec.  23.1311 by revising paragraphs (a)(5), (a)(6),
(a)(7), and paragraph (b) to read as follows:

Sec.  23.1311  Electronic display instrument systems.

    (a) * * *
    (5) Have an independent magnetic direction indicator and an
independent secondary mechanical altimeter, airspeed indicator, and
attitude instrument or electronic display parameters for the altitude,
airspeed, and attitude that are independent from the airplane's primary
electrical power system. These secondary instruments may be installed
in panel positions that are displaced from the primary positions
specified by Sec.  23.1321(d), but must be located where they meet the
pilot's visibility requirements of Sec.  23.1321(a).
    (6) Incorporate sensory cues that provide a quick glance sense of
rate and, when appropriate, trend information to the pilot.
    (7) Incorporate equivalent visual displays of the instrument
markings required by Sec. Sec.  23.1541 through 23.1553, or visual
displays that alert the pilot to abnormal operational values or
approaches to established limitation values, for each parameter
required to be displayed by this part.
    (b) The electronic display indicators, including their systems and
installations, and considering other airplane systems, must be designed
so that one display of information essential for continued safe flight
and landing will be available within one second to the crew with a
single pilot action or by automatic means for continued safe operation,
after any single failure or probable combination of failures.
* * * * *
    54. Amend Sec.  23.1323 by revising paragraph (e) to read as
follows:

Sec.  23.1323  Airspeed indicating system.

* * * * *
    (e) In addition, for normal, utility, and acrobatic category
multiengine turbojet airplanes of more than 6,000 pounds maximum weight
and commuter category airplanes, each system must be calibrated to
determine the system error during the accelerate-takeoff ground run.
The ground run calibration must be determined--
    (1) From 0.8 of the minimum value of V1 to the maximum
value of V2, considering the approved ranges of altitude and
weight, and
    (2) The ground run calibration must be determined assuming an
engine failure at the minimum value of V1.
* * * * *
    55. Amend Sec.  23.1331 by revising paragraph (c) to read as
follows:

Sec.  23.1331  Instruments using a power source.

* * * * *
    (c) For certification for Instrument Flight Rules (IFR) operations
and for the heading, altitude, airspeed, and attitude, there must be at
least:
    (1) Two independent sources of power (not driven by the same engine
on multiengine airplanes), and a manual or an automatic means to select
each power source; or
    (2) An additional display of parameters for heading, altitude,
airspeed, and attitude that is independent from the airplane's primary
electrical power system.
    56. Amend Sec.  23.1353 by revising paragraph (h) to read as
follows:

Sec.  23.1353  Storage battery design and installation.

* * * * *
    (h) In the event of a complete loss of the primary electrical power
generating system, the battery must be capable of providing electrical
power to those loads that are essential to continued safe flight and
landing for:
    (1) At least 30 minutes for airplanes that are certificated with a
maximum altitude of 25,000 feet or less, and
    (2) At least 60 minutes for airplanes that are certificated with a
maximum altitude over 25,000 feet.
    57. Revise Sec.  23.1443 to read as follows:

Sec.  23.1443  Minimum mass flow of supplemental oxygen.

    (a) If the airplane is to be certified above 40,000 feet, a
continuous flow oxygen system must be provided for each passenger and
crewmember.
    (b) If continuous flow oxygen equipment is installed, an applicant
must show compliance with the requirements of either paragraphs (b)(1)
and (b)(2) or paragraph (b)(3) of this section:
    (1) For each passenger, the minimum mass flow of supplemental
oxygen required at various cabin pressure altitudes may not be less
than the flow required to maintain, during inspiration and while using
the oxygen equipment (including masks) provided, the following mean
tracheal oxygen partial pressures:
    (i) At cabin pressure altitudes above 10,000 feet up to and
including 18,500 feet, a mean tracheal oxygen partial pressure of 100mm
Hg when breathing 15 liters per minute, Body Temperature, Pressure,
Saturated (BTPS) and with a tidal volume of 700cc with a constant time
interval between respirations.
    (ii) At cabin pressure altitudes above 18,500 feet up to and
including 40,000 feet, a mean tracheal oxygen partial pressure of
83.8mm Hg when breathing 30 liters per minute, BTPS, and with a tidal
volume of 1,100cc with a constant time interval between respirations.

[GRAPHIC] [TIFF OMITTED] TP17AU09.003

    (2) For each flight crewmember, the minimum mass flow may not be
less than the flow required to maintain, during inspiration, a mean
tracheal oxygen partial pressure of 149mm Hg when breathing 15 liters
per minute, BTPS, and with a maximum tidal volume of 700cc with a
constant time interval between respirations.
    (3) The minimum mass flow of supplemental oxygen supplied for each
user must be at a rate not less than that shown in the following figure
for each altitude up to and including the maximum operating altitude of
the airplane.
    (c) If demand equipment is installed for use by flight crewmembers,
the minimum mass flow of supplemental oxygen required for each flight
crewmember may not be less than the flow required to maintain, during
inspiration, a mean tracheal oxygen partial pressure of 122mm Hg up to
and including a cabin pressure altitude of 35,000 feet, and 95 percent
oxygen between cabin pressure altitudes of 35,000 and 40,000 feet, when
breathing 20 liters per minutes BTPS. In addition, there must be means
to allow the crew to use undiluted oxygen at their discretion.
    (d) If first-aid oxygen equipment is installed, the minimum mass
flow of oxygen to each user may not be less than 4 liters per minute,
STPD. However, there may be a means to decrease this flow to not less
than 2 liters per minute, STPD, at any cabin altitude. The quantity of
oxygen required is based upon an average flow rate of 3 liters per
minute per person for whom first-aid oxygen is required.
    (e) As used in this section:
    (1) BTPS means Body Temperature, and Pressure, Saturated (which is
37 [deg]C, and the ambient pressure to which the body is exposed, minus
47mm Hg, which is the tracheal pressure displaced by water vapor
pressure when the breathed air becomes saturated with water vapor at 37
[deg]C).
    (2) STPD means Standard Temperature and Pressure, Dry (which is 0
[deg]C at 760mm Hg with no water vapor).
    58. Amend Sec.  23.1445 by adding a new paragraph (c) to read as
follows:

Sec.  23.1445  Oxygen distribution system.

* * * * *
    (c) If the flight crew and passengers share a common source of
oxygen, a means to separately reserve the minimum supply required by
the flight crew must be provided.
    59. Amend Sec.  23.1447 by adding a new paragraph (g) to read as
follows:

Sec.  23.1447  Equipment standards for oxygen dispensing units.

* * * * *
    (g) If the airplane is to be certified for operation above 40,000
feet, a quick-donning oxygen mask system, with a pressure demand, mask
mounted regulator must be provided for the flight crew. This dispensing
unit must be immediately available to the flight crew when seated at
their station and installed so that it:
    (1) Can be placed on the face from its ready position, properly
secured, sealed, and supplying oxygen upon demand, with one hand,
within five seconds and without disturbing eyeglasses or causing delay
in proceeding with emergency duties, and
    (2) Allows, while in place, the performance of normal communication
functions.
    60. Amend Sec.  23.1505 by revising paragraph (c) to read as
follows:

Sec.  23.1505  Airspeed limitations.

* * * * *
    (c) Paragraphs (a) and (b) of this section do not apply to turbine
airplanes or the airplanes for which a design diving speed
VD/MD is established under Sec.  23.335(b)(4).
For those airplanes, a maximum operating limit speed (VMO/
MMO airspeed or Mach number, whichever is critical at a
particular altitude) must be established as a speed that may not be
deliberately exceeded in any regime of flight (climb, cruise, or
descent) unless a higher speed is authorized for flight test or pilot
training operations. VMO/MMO must be established
so that it is not greater than the design cruising speed VC/
MC and so that it is sufficiently below VD/
MD, or VDF/MDF for turbojets, and the
maximum speed shown under Sec.  23.251 to make it highly improbable
that the latter speeds will be inadvertently exceeded in operations.
The speed margin between VMO/MMO and
VD/MD,

or VDF/MDF for turbojets, may not be less than
that determined under Sec.  23.335(b), or the speed margin found
necessary in the flight tests conducted under Sec.  23.253.
    61. Revise Sec.  23.1525 to read as follows:

Sec.  23.1525  Kinds of operation.

    The kinds of operation authorized (e.g., VFR, IFR, day, night, part
91, part 135) and the meteorological conditions (e.g., icing) to which
the operation of the airplane is limited or from which it is
prohibited, must be established appropriate to the installed equipment.
    62. Amend Sec.  23.1545 by revising paragraph (d) to read as
follows:

Sec.  23.1545  Airspeed indicator.

* * * * *
    (d) Paragraphs (b)(1) through (b)(4) and paragraph (c) of this
section do not apply to airplanes for which a maximum operating speed
VMO/MMO is established under Sec.  23.1505(c).
For those airplanes, there must either be a maximum allowable airspeed
indication showing the variation of VMO/MMO with
altitude or compressibility limitations (as appropriate), or a radial
red line marking for VMO/MMO must be made at the
lowest value of VMO/MMO established for any
altitude up to the maximum operating altitude for the airplane.
    63. Amend Sec.  23.1555 by adding a new paragraph (d)(3) to read as
follows:

Sec.  23.1555  Control markings.

* * * * *
    (d) * * *
    (3) For fuel systems having a calibrated fuel quantity indication
system complying with Sec.  23.1337(b)(1) and accurately displaying the
actual quantity of usable fuel in each selectable tank, no fuel
capacity placards outside of the fuel quantity indicator are required.
* * * * *
    64. Amend Sec.  23.1559 by adding a new paragraph (d) to read as
follows:

Sec.  23.1559  Operating limitations placard.

* * * * *
    (d) The placard(s) required by this section need not be lighted.
    65. Amend Sec.  23.1563 by adding a new paragraph (d) to read as
follows:

Sec.  23.1563  Airspeed placards.

* * * * *
    (d) The airspeed placard required by this section need not be
lighted if the landing gear operating speed is indicated on the
airspeed indicator or other lighted area such as the landing gear
control and the airspeed indicator has features such as low speed
awareness that provide ample warning prior to VMC.
    66. Amend Sec.  23.1567 by adding a new paragraph (e) to read as
follows:

Sec.  23.1567  Flight maneuver placard.

* * * * *
    (e) The placards required by this section need not be lighted.
    67. Amend Sec.  23.1583 as follows:
    a. Revise the introductory text of paragraphs (c)(3) and (c)(4);
    b. Redesignate paragraphs (c)(4)(iii) and (c)(4)(iv) as paragraphs
(c)(4)(ii)(A) and (c)(4)(ii)(B); and
    c. Revise paragraph (c)(5) introductory text to read as follows:

Sec.  23.1583  Operating limitations.

* * * * *
    (c) * * *
    (3) For reciprocating engine-powered airplanes of more than 6,000
pounds maximum weight, single-engine turbines, and multiengine turbine
airplanes 6,000 pounds or less maximum weight in the normal, utility,
and acrobatic category, performance operating limitations as follows--
* * * * *
    (4) For normal, utility, and acrobatic category multiengine
turbojet powered airplanes over 6,000 pounds and commuter category
airplanes, the maximum takeoff weight for each airport altitude and
ambient temperature within the range selected by the applicant at
which--
* * * * *
    (5) For normal, utility, and acrobatic category multiengine
turbojet powered airplanes over 6,000 pounds and commuter category
airplanes, the maximum landing weight for each airport altitude within
the range selected by the applicant at which--
* * * * *
    68. Amend Sec.  23.1585 by revising paragraph (f) introductory text
to read as follows:

Sec.  23.1585  Operating procedures.

* * * * *
    (f) In addition to paragraphs (a) and (c) of this section, for
normal, utility, and acrobatic category multiengine turbojet powered
airplanes over 6,000 pounds, and commuter category airplanes, the
information must include the following:
* * * * *
    69. Amend Sec.  23.1587 by revising paragraph (d) introductory text
to read as follows:

Sec.  23.1587  Performance information.

* * * * *
    (d) In addition to paragraph (a) of this section, for normal,
utility, and acrobatic category multiengine turbojet powered airplanes
over 6,000 pounds, and commuter category airplanes, the following
information must be furnished--
* * * * *
    70. Amend Appendix F to Part 23 by:
    a. Redesignating the existing text as Part I and adding a new Part
I heading;
    b. Removing the introductory paragraph; and
    c. Adding a new Part II.
    The additions read as follows:

APPENDIX F TO PART 23--TEST PROCEDURE

Part I--Acceptable Test Procedure for Self-Extinguishing Materials for
Showing Compliance With Sec. Sec.  23.853, 23.855 and 23.1359

* * * * *

Part II--Test Method To Determine the Flammability and Flame
Propagation Characteristics of Thermal/Acoustic Insulation Materials

    Use this test method to evaluate the flammability and flame
propagation characteristics of thermal/acoustic insulation when
exposed to both a radiant heat source and a flame.
    (a) Definitions.
    ``Flame propagation'' means the furthest distance of the
propagation of visible flame towards the far end of the test
specimen, measured from the midpoint of the ignition source flame.
Measure this distance after initially applying the ignition source
and before all flame on the test specimen is extinguished. The
measurement is not a determination of burn length made after the
test.
    ``Radiant heat source'' means an electric or air propane panel.
    ``Thermal/acoustic insulation'' means a material or system of
materials used to provide thermal and/or acoustic protection.
Examples include fiberglass or other batting material encapsulated
by a film covering and foams.
    ``Zero point'' means the point of application of the pilot
burner to the test specimen.
    (b) Test apparatus.

[GRAPHIC] [TIFF OMITTED] TP17AU09.004

    (1) Radiant panel test chamber. Conduct tests in a radiant panel
test chamber (see figure F1 above). Place the test chamber under an
exhaust hood to facilitate clearing the chamber of smoke after each
test. The radiant panel test chamber must be an enclosure 55 inches
(1397 mm) long by 19.5 (495 mm) deep by 28 (710 mm) to 30 inches
(maximum) (762 mm) above the test specimen. Insulate the sides,
ends, and top with a fibrous ceramic insulation, such as Kaowool MTM
board. On the front side, provide a 52 by 12-inch (1321 by 305 mm)
draft-free, high-temperature, glass window for viewing the sample
during testing. Place a door below the window to provide access to
the movable specimen platform holder. The bottom of the test chamber
must be a sliding steel platform that has provision for securing the
test specimen holder in a fixed and level position. The chamber must
have an internal chimney with exterior dimensions of 5.1 inches (129
mm) wide, by 16.2 inches (411 mm) deep by 13 inches (330 mm) high at
the opposite end of the chamber from the radiant energy source. The
interior dimensions must be 4.5 inches (114 mm) wide by 15.6 inches
(395 mm) deep. The chimney must extend to the top of the chamber
(see figure F2).
[GRAPHIC] [TIFF OMITTED] TP17AU09.005

    (2) Radiant heat source. Mount the radiant heat energy source in
a cast iron frame or equivalent. An electric panel must have six, 3-
inch wide emitter strips. The emitter strips must be perpendicular
to the length of the panel. The panel must have a radiation surface
of 12 \7/8\ by 18 \1/2\ inches (327 by 470 mm). The panel must be
capable of operating at temperatures up to 1300 [deg]F (704 [deg]C).
An air propane panel must be made of a porous refractory material
and have a radiation surface of 12 by 18 inches (305 by 457 mm). The
panel must be capable of operating at temperatures up to 1,500
[deg]F (816 [deg]C). See figures 3a and 3b.

[GRAPHIC] [TIFF OMITTED] TP17AU09.006

    (i) Electric radiant panel. The radiant panel must be 3-phase
and operate at 208 volts. A single-phase, 240 volt panel is also
acceptable. Use a solid-state power controller and microprocessor-
based controller to set the electric panel operating parameters.
    (ii) Gas radiant panel. Use propane (liquid petroleum gas--2.1
UN 1075) for the radiant panel fuel. The panel fuel system must
consist of a venturi-type aspirator for mixing gas and air at
approximately atmospheric pressure. Provide suitable instrumentation
for monitoring and controlling the flow of fuel and air to the
panel. Include an air flow gauge, an air flow regulator, and a gas
pressure gauge.
    (iii) Radiant panel placement. Mount the panel in the chamber at
30 degrees to the horizontal specimen plane, and 7\1/2\ inches above
the zero point of the specimen.
    (3) Specimen holding system.
    (i) The sliding platform serves as the housing for test specimen
placement. Brackets may be attached (via wing nuts) to the top lip
of the platform in order to accommodate various thicknesses of test
specimens. Place the test specimens on a sheet of Kaowool MTM board
or 1260 Standard Board (manufactured by Thermal Ceramics and
available in Europe), or equivalent, either resting on the bottom
lip of the sliding platform or on the base of the brackets. It may
be necessary to use multiple sheets of material based on the
thickness of the test specimen (to meet the sample height
requirement). Typically, these non-combustible sheets of material
are available in \1/4\ inch (6 mm) thicknesses. See figure F4. A
sliding platform that is deeper than the 2-

inch (50.8 mm) platform shown in figure F4 is also acceptable as
long as the sample height requirement is met.
[GRAPHIC] [TIFF OMITTED] TP17AU09.007

    (ii) Attach a \1/2\ inch (13 mm) piece of Kaowool MTM board or
other high temperature material measuring 41\1/2\ by 8\1/4\ inches
(1054 by 210 mm) to the back of the platform. This board serves as a
heat retainer and protects the test specimen from excessive
preheating. The height of this board must not impede the sliding
platform movement (in and out of the test chamber). If the platform
has been fabricated such that the back side of the platform is high
enough to prevent excess preheating of the specimen when the sliding
platform is out, a retainer board is not necessary.
    (iii) Place the test specimen horizontally on the non-
combustible board(s). Place a steel retaining/securing frame
fabricated of mild steel, having a thickness of \1/8\ inch (3.2 mm)
and overall dimensions of 23 by 13\1/8\ inches (584 by 333 mm) with
a specimen opening of 19 by 10\3/4\ inches (483 by 273 mm) over the
test specimen. The front, back, and right portions of the top flange
of the frame must rest on the top of the sliding platform, and the
bottom flanges must pinch all 4 sides of the test specimen. The
right bottom flange must be flush with the sliding platform. See
figure F5.
[GRAPHIC] [TIFF OMITTED] TP17AU09.008

    (4) Pilot Burner. The pilot burner used to ignite the specimen
must be a BernzomaticTM commercial propane venturi torch with an
axially symmetric burner tip and a propane supply tube with an
orifice diameter of 0.006 inches (0.15 mm). The length of the burner
tube must be 2\7/8\ inches (71 mm). The propane flow must be
adjusted via gas pressure through an in-line regulator to produce a
blue inner cone length of \3/4\ inch (19 mm). A \3/4\ inch (19 mm)
guide (such as a thin strip of metal) may be soldered to the top of
the burner to aid in setting the flame height. The overall flame
length must be approximately 5 inches long (127 mm). Provide a way
to move the burner out of the ignition position so that the flame is
horizontal and at least 2 inches (50 mm) above the specimen plane.
See figure F6.
[GRAPHIC] [TIFF OMITTED] TP17AU09.009

    (5) Thermocouples. Install a 24 American Wire Gauge (AWG) Type K
(Chromel- Alumel) thermocouple in the test chamber for temperature
monitoring. Insert it into the chamber through a small hole drilled
through the back of the chamber. Place the thermocouple so that it
extends 11 inches (279 mm) out from the back of the chamber wall,
11\1/2\ inches (292 mm) from the right side of the chamber wall, and
is 2 inches (51 mm) below the radiant panel. The use of other
thermocouples is optional.
    (6) Calorimeter. The calorimeter must be a one-inch cylindrical
water-cooled, total heat flux density, foil type Gardon Gage that
has a range of 0 to 5 BTU/ft\2\ -second (0 to 5.7 Watts/cm\2\).
    (7) Calorimeter calibration specification and procedure.
    (i) Calorimeter specification.
    (A) Foil diameter must be 0.25  0.005 inches (6.35
 0.13 mm).
    (B) Foil thickness must be 0.0005  0.0001 inches
(0.013  0.0025 mm).
    (C) Foil material must be thermocouple grade Constantan.
    (D) Temperature measurement must be a Copper Constantan
thermocouple.
    (E) The copper center wire diameter must be 0.0005 inches (0.013
mm).
    (F) The entire face of the calorimeter must be lightly coated
with ``Black Velvet'' paint having an emissivity of 96 or greater.
    (ii) Calorimeter calibration.
    (A) The calibration method must be by comparison to a like
standardized transducer.
    (B) The standardized transducer must meet the specifications
given in paragraph VI(b)(6) of this appendix.
    (C) Calibrate the standard transducer against a primary standard
traceable to the National Institute of Standards and Technology
(NIST).
    (D) The method of transfer must be a heated graphite plate.
    (E) The graphite plate must be electrically heated, have a clear
surface area on each side of the plate of at least 2 by 2 inches (51
by 51 mm), and be \1/8\ inch  \1/16\ inch thick (3.2
 1.6 mm).
    (F) Center the 2 transducers on opposite sides of the plates at
equal distances from the plate.
    (G) The distance of the calorimeter to the plate must be no less
than 0.0625 inches (1.6 mm), nor greater than 0.375 inches (9.5 mm).
    (H) The range used in calibration must be at least 0-3.5 BTUs/
ft\2\ second (0-3.9 Watts/cm\2\) and no greater than 0-5.7 BTUs/
ft\2\ second (0-6.4 Watts/cm\2\).
    (I) The recording device used must record the 2 transducers
simultaneously or at least within \1/10\ of each other.
    (8) Calorimeter fixture. With the sliding platform pulled out of
the chamber, install the calorimeter holding frame and place a sheet
of non-combustible material in the bottom of the sliding platform
adjacent to the holding frame. This will prevent heat losses during
calibration. The frame must be 13\1/8\ inches (333 mm) deep (front
to back) by 8 inches (203 mm) wide and must rest on the top of the
sliding platform. It must be fabricated of \1/8\ inch (3.2 mm) flat
stock steel and have an opening that accommodates a \1/2\ inch (12.7
mm) thick piece of refractory board, which is level with the top of
the sliding platform. The board must have three 1-inch (25.4 mm)
diameter holes drilled through the board for calorimeter insertion.
The distance to the radiant panel surface from the centerline of the
first hole (``zero'' position) must be 7\1/2\  \1/8\
inches (191  3 mm). The distance between the centerline
of the first hole to the centerline of the second hole must be 2
inches (51 mm). It must also be the same distance from the
centerline of the second hole to the centerline of the third hole.
See figure F7. A calorimeter holding frame that differs in
construction is acceptable as long as the height from the centerline
of the first hole to the radiant panel and the distance between
holes is the same as described in this paragraph.

[GRAPHIC] [TIFF OMITTED] TP17AU09.010

    (9) Instrumentation. Provide a calibrated recording device with
an appropriate range or a computerized data acquisition system to
measure and record the outputs of the calorimeter and the
thermocouple. The data acquisition system must be capable of
recording the calorimeter output every second during calibration.
    (10) Timing device. Provide a stopwatch or other device,
accurate to  1 second/hour, to measure the time of
application of the pilot burner flame.
    (c) Test specimens.
    (1) Specimen preparation. Prepare and test a minimum of three
test specimens. If an oriented film cover material is used, prepare
and test both the warp and fill directions.
    (2) Construction. Test specimens must include all materials used
in construction of the insulation (including batting, film, scrim,
tape, etc.). Cut a piece of core material such as foam or
fiberglass, and cut a piece of film cover material (if used) large
enough to cover the core material. Heat sealing is the preferred
method of preparing fiberglass samples, since they can be made
without compressing the fiberglass (``box sample''). Cover materials
that are not heat sealable may be stapled, sewn, or taped as long as
the cover material is over-cut enough to be drawn down the sides
without compressing the core material. The fastening means should be
as continuous as possible along the length of the seams. The
specimen thickness must be of the same thickness as installed in the
airplane.
    (3) Specimen Dimensions. To facilitate proper placement of
specimens in the sliding platform housing, cut non-rigid core
materials, such as fiberglass, 12\1/2\ inches (318 mm) wide by 23
inches (584 mm) long. Cut rigid materials, such as foam, 11\1/2\
 \1/4\ inches (292 mm  6 mm) wide by 23
inches (584 mm) long in order to fit properly in the sliding
platform housing and provide a flat, exposed surface equal to the
opening in the housing.
    (d) Specimen conditioning. Condition the test specimens at 70
 5 [deg]F (21  2 [deg]C) and 55 percent
 10 percent relative humidity, for a minimum of 24 hours
prior to testing.
    (e) Apparatus Calibration.
    (1) With the sliding platform out of the chamber, install the
calorimeter holding frame. Push the platform back into the chamber
and insert the calorimeter into the first hole (``zero'' position).
See figure F7. Close the bottom door located below the sliding
platform. The distance from the centerline of the calorimeter to the
radiant panel surface at this point must be 7\1/2\ inches  \1/8\ (191 mm  3). Before igniting the radiant
panel, ensure that the calorimeter face is clean and that there is
water running through the calorimeter.
    (2) Ignite the panel. Adjust the fuel/air mixture to achieve 1.5
BTUs/feet\2\ -second  5 percent (1.7 Watts/cm\2\  5 percent) at the ``zero'' position. If using an electric
panel, set the power controller to achieve the proper heat flux.
Allow the unit to reach steady state (this may take up to 1 hour).
The pilot burner must be off and in the down position during this
time.
    (3) After steady-state conditions have been reached, move the
calorimeter 2 inches (51 mm) from the ``zero'' position (first hole)
to position 1 and record the heat flux. Move the calorimeter to
position 2 and record the heat flux. Allow enough time at each
position for the calorimeter to stabilize. Table 1 depicts typical
calibration values at the three positions.

                       Table 1--Calibration Table
------------------------------------------------------------------------
            Position               BTU's/feet\2\ sec      Watts/cm\2\
------------------------------------------------------------------------
``Zero'' Position...............                 1.5                 1.7
Position 1......................      1.51-1.50-1.49      1.71-1.70-1.69
Position 2......................           1.43-1.44           1.62-1.63
------------------------------------------------------------------------

     (4) Open the bottom door, remove the calorimeter and holder
fixture. Use caution as the fixture is very hot.
    (f) Test Procedure.
    (1) Ignite the pilot burner. Ensure that it is at least 2 inches
(51 mm) above the top of the platform. The burner must not contact
the specimen until the test begins.
    (2) Place the test specimen in the sliding platform holder.
Ensure that the test sample surface is level with the top of the
platform. At ``zero'' point, the specimen surface must be 7\1/2\
inches  \1/8\ inch (191 mm  3) below the
radiant panel.
    (3) Place the retaining/securing frame over the test specimen.
It may be necessary (due to compression) to adjust the sample (up or
down) in order to maintain the distance from the sample to the
radiant panel (7\1/2\ inches  \1/8\ inch (191 mm  3) at ``zero'' position). With film/fiberglass assemblies, it
is critical to make a slit in the film cover to purge any air
inside. This allows the operator to maintain the proper test
specimen position (level with the top of the platform) and to allow
ventilation of gases during testing. A longitudinal slit,
approximately 2 inches (51 mm) in length, must be centered 3 inches
 \1/2\ inch (76 mm  13 mm) from the left
flange of the securing frame. A utility knife is acceptable for
slitting the film cover.
    (4) Immediately push the sliding platform into the chamber and
close the bottom door.

    (5) Bring the pilot burner flame into contact with the center of
the specimen at the ``zero'' point and simultaneously start the
timer. The pilot burner must be at a 27 degree angle with the sample
and be approximately \1/2\ inch (12 mm) above the sample. See figure
F7. A stop, as shown in figure F8, allows the operator to position
the burner correctly each time.
[GRAPHIC] [TIFF OMITTED] TP17AU09.011

    (6) Leave the burner in position for 15 seconds and then remove
to a position at least 2 inches (51 mm) above the specimen.
    (g) Report.
    (1) Identify and describe the test specimen.
    (2) Report any shrinkage or melting of the test specimen.
    (3) Report the flame propagation distance. If this distance is
less than 2 inches, report this as a pass (no measurement required).
    (4) Report the after-flame time.
    (h) Requirements.
    (1) There must be no flame propagation beyond 2 inches (51 mm)
to the left of the centerline of the pilot flame application.
    (2) The flame time after removal of the pilot burner may not
exceed 3 seconds on any specimen.
    71. Add a new Appendix K to part 23 to read as follows:

Appendix K to Part 23--Relationship Among Airplane Classes,
Probabilities, Severity of Failure Conditions, and Software and Complex
Hardware Development Assurance Levels

--------------------------------------------------------------------------------------------------------------------------------------------------------

--------------------------------------------------------------------------------------------------------------------------------------------------------
     Classification of failure          No safety effect              Minor                  Major                Hazardous             Catastrophic
             conditions
--------------------------------------------------------------------------------------------------------------------------------------------------------
 Allowable qualitative probability       No probability             Probable                 Remote            Extremely remote     Extremely improbable
                                           requirement
--------------------------------------------------------------------------------------------------------------------------------------------------------
Effect on Airplane.................  No effect on            Slight reduction in     Significant reduction  Large reduction in     Normally with hull
                                      operational             functional              in functional          functional             loss.
                                      capabilities or         capabilities or         capabilities or        capabilities or
                                      safety.                 safety margins.         safety margins.        safety margins.
Effect on Occupants................  Inconvenience for       Physical discomfort     Physical distress to   Serious or fatal       Multiple fatalities
                                      passengers.             for passengers.         passengers, possibly   injury to an
                                                                                including injuries.    occupant.
Effect on Flight Crew..............  No effect on flight     Slight increase in      Physical discomfort    Physical distress or   Fatal Injury or
                                      crew.                   workload or use of      or a significant       excessive workload     incapacitation.
                                                              emergency procedures.   increase in workload.  impairs ability to
                                                                                perform tasks.
--------------------------------------------------------------------------------------------------------------------------------------------------------
        Classes of Airplanes         Allowable Quantitative Probabilities and Software (SW) and Complex Hardware (HW) Development Assurance Levels (Note
                                                                                2)
--------------------------------------------------------------------------------------------------------------------------------------------------------
              Class I
(Typically SRE under 6,000).                                 HW Development                                  P=C, S=D.              S=D.                   S=C.
                                      Assurance Levels
                                      Requirement.
              Class II
(Typically MRE, STE, or MTE under    No Probability or SW &  <10-3, Note 1, P=D....  <10-5, Notes 1 & 4,    <10-6, Notes 4, P=C,   <10-7, Note 3, P=C,
 6,000).                     HW Development                                  P=C, S=D.              S=C.                   S=C.
                                      Assurance Levels
                                      Requirement.

             Class III
(Typically SRE, STE, MRE, & MTE      No Probability or SW &  <10-3, Note 1, P=D....  <10-5, Notes 1 & 4,    <10-7, Notes 4, P=C,   <10-8, Note 3, P=B,
 equal or over 6,000).       HW Development                                  P=C, S=D.              S=C.                   S=C.
                                      Assurance Levels
                                      Requirement.
              Class IV
(Typically Commuter Category)......  No Probability or SW &  <10-3, Note 1, P=D....  <10-5, Notes 1 & 4,    <10-7, Notes 4, P=B,   <10-9, Note 3, P=A,
                                      HW Development                                  P=C, S=D.              S=C.                   S=B.
                                      Assurance Levels
                                      Requirement.
--------------------------------------------------------------------------------------------------------------------------------------------------------
Note 1: Numerical values indicate an order of probability range and are provided here as a reference.
Note 2: The alphabets denote the typical SW and HW Development Assurance Levels for Primary System (P) and Secondary System (S). For example, HW or SW
  Development Assurance Level A on Primary System is noted by P=A.
Note 3: At airplane function level, no single failure will result in a Catastrophic Failure Condition.
Note 4: Secondary System (S) may not be required to meet probability goals. If installed, S must meet stated criteria.
Acronyms: SRE--single, reciprocating engine, MRE--multiple, reciprocating engines, STE--single, turbine engine, MTE--multiple, turbine engines, SW--
  software, HW--hardware.

    Issued in Washington, DC, on August 6, 2009.
Dorenda D. Baker,
Director, Aircraft Certification Service, Office of Aviation Safety.
[FR Doc. E9-19350 Filed 8-14-09; 8:45 am]

BILLING CODE 4910-13-P


[Federal Register Volume 78, Number 65 (Thursday, April 4, 2013)]
[Notices]
[Pages 20327-20328]
[FR Doc No: 2013-07843]

-----------------------------------------------------------------------

DEPARTMENT OF HEALTH AND HUMAN SERVICES

Food and Drug Administration

[Docket No. FDA-2013-N-0001]

Joint Meeting of the Advisory Committee for Reproductive Health
Drugs and the Drug Safety and Risk Management Advisory Committee;
Amendment of Notice

AGENCY:  Food and Drug Administration, HHS.

ACTION: Notice.

-----------------------------------------------------------------------

    The Food and Drug Administration (FDA) is announcing an amendment
to the notice of the joint meeting of the Advisory Committee for
Reproductive Health Drugs and the Drug Safety and Risk Management
Advisory Committee. This meeting was announced in the Federal Register
of March 14, 2013 (78 FR 16271-16272). The amendment is being made to
reflect a change in the Agenda portion of the document. There are no
other changes.

FOR FURTHER INFORMATION CONTACT:  Kalyani Bhatt, Center for Drug
Evaluation and Research, Food and Drug Administration, 10903 New
Hampshire Ave., Bldg. 31 rm. 2417, Silver Spring, MD 20993-0002, 301-
796-9001, FAX:<PHONE_NUMBER><EMAIL_ADDRESS>or use the FDA Advisory
Committee Information Line, 1-800-741-8138<PHONE_NUMBER> in the
Washington DC area), and follow the prompts to the desired center or
product area. Please call the Information Line for up-to-date
information on this meeting.

SUPPLEMENTARY INFORMATION:  In the Federal Register of March 14, 2013,
FDA announced that a joint meeting of the Advisory Committee for
Reproductive Health Drugs and the Drug Safety and Risk Management
Advisory Committee would be held on April 18, 2013. On page 16272, in
the first column, the Agenda portion of the document is changed to read
as follows:
    Agenda: The committee will discuss the efficacy and safety of new
drug application (NDA) 22219, AVEED (testosterone undecanoate)
intramuscular injection, submitted by Endo Pharmaceutical Solutions,
Inc., for the proposed indication of replacement

therapy in adult males for conditions associated with a deficiency or
absence of testosterone. The safety discussion will focus on
postmarketing reports of oil microembolism in the lungs and potential
anaphylactic reactions. In addition to AVEED, other approved
testosterone injectable products will be referenced, especially in
regard to oil microembolism and potential anaphylactic reactions
reported for those products.
    This notice is issued under the Federal Advisory Committee Act (5
U.S.C. app. 2) and 21 CFR part 14, relating to the advisory committees.

    Dated: March 27, 2013.
Jill Hartzler Warner,
Acting Associate Commissioner for Special Medical Programs.
[FR Doc. 2013-07843 Filed 4-3-13; 8:45 am]
BILLING CODE 4160-01-P