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29.1 Applicability.
(a)
This part prescribes airworthiness standards for the
issue of type certificates, and changes to those
certificates, for transport category rotorcraft.
(b)
Transport category rotorcraft must be certificated
in accordance with either the Category A or Category
B requirements of this part. A multiengine
rotorcraft may be type certificated as both Category
A and Category B with appropriate and different
operating limitations for each category.
(c)
Rotorcraft with a maximum weight greater than 20,000
pounds and 10 or more passenger seats must be type
certificated as Category A rotorcraft.
(d)
Rotorcraft with a maximum weight greater than 20,000
pounds and nine or less passenger seats may be type
certificated as Category B rotorcraft provided the
Category A requirements of Subparts C, D, E, and F
of this part are met.
(e)
Rotorcraft with a maximum weight of 20,000 pounds or
less but with 10 or more passenger seats may be type
certificated as Category B rotorcraft provided the
Category A requirements of 29.67(a)(2), 29.87,
29.1517, and subparts C, D, E, and F of this part
are met.
(f)
Rotorcraft with a maximum weight of 20,000 pounds or
less and nine or less passenger seats may be type
certificated as Category B rotorcraft.
(g)
Each person who applies under Part 21 for a
certificate or change described in paragraphs (a)
through (f) of this section must show compliance
with the applicable requirements of this part.
For
each rotorcraft, each applicant must show that each
occupant's seat is equipped with a safety belt and
shoulder harness that meets the requirements of
paragraphs (a), (b), and (c) of this section.
(a)
Each occupant's seat must have a combined safety
belt and shoulder harness with a single-point
release. Each pilot's combined safety belt and
shoulder harness must allow each pilot, when seated
with safety belt and shoulder harness fastened, to
perform all functions necessary for flight
operations. There must be a means to secure belts
and harnesses, when not in use, to prevent
interference with the operation of the rotorcraft
and with rapid egress in an emergency.
(b)
Each occupant must be protected from serious head
injury by a safety belt plus a shoulder harness that
will prevent the head from contacting any injurious
object.
(c)
The safety belt and shoulder harness must meet the
static and dynamic strength requirements, if
applicable, specified by the rotorcraft type
certification basis.
(d)
For purposes of this section, the date of
manufacture is either—
(1)
The date the inspection acceptance records, or
equivalent, reflect that the rotorcraft is complete
and meets the AFRO-CAA-Approved Type Design Data; or
(2)
The date that the foreign civil airworthiness
authority certifies the rotorcraft is complete and
issues an original standard airworthiness
certificate, or equivalent, in that country.
General
29.21 Proof of
compliance.
Each
requirement of this subpart must be met at each
appropriate combination of weight and center of
gravity within the range of loading conditions for
which certification is requested. This must be
shown—
(a)
By tests upon a rotorcraft of the type for which
certification is requested, or by calculations based
on, and equal in accuracy to, the results of
testing; and
(b)
By systematic investigation of each required
combination of weight and center of gravity, if
compliance cannot be reasonably inferred from
combinations investigated.
(a)
Maximum weight. The maximum weight (the
highest weight at which compliance with each
applicable requirement of this part is shown) or, at
the option of the applicant, the highest weight for
each altitude and for each practicably separable
operating condition, such as takeoff, enroute
operation, and landing, must be established so that
it is not more than—
(1)
The highest weight selected by the applicant;
(2)
The design maximum weight (the highest weight at
which compliance with each applicable structural
loading condition of this part is shown); or
(3)
The highest weight at which compliance with each
applicable flight requirement of this part is shown.
(b)
Minimum weight. The minimum weight (the
lowest weight at which compliance with each
applicable requirement of this part is shown) must
be established so that it is not less than—
(1)
The lowest weight selected by the applicant;
(2)
The design minimum weight (the lowest weight at
which compliance with each structural loading
condition of this part is shown); or
(3)
The lowest weight at which compliance with each
applicable flight requirement of this part is shown.
(c)
Total weight with jettisonable external load.
A total weight for the rotorcraft with a
jettisonable external load attached that is greater
than the maximum weight established under paragraph
(a) of this section may be established for any
rotorcraft-load combination if—
(1)
The rotorcraft-load combination does not include
human external cargo,
(2)
Structural component approval for external load
operations under either 29.865 or under equivalent
operational standards is obtained,
(3)
The portion of the total weight that is greater than
the maximum weight established under paragraph (a)
of this section is made up only of the weight of all
or part of the jettisonable external load,
(4)
Structural components of the rotorcraft are shown to
comply with the applicable structural requirements
of this part under the increased loads and stresses
caused by the weight increase over that established
under paragraph (a) of this section, and
(5)
Operation of the rotorcraft at a total weight
greater than the maximum certificated weight
established under paragraph (a) of this section is
limited by appropriate operating limitations under
29.865 (a) and (d) of this part.
The
extreme forward and aft centers of gravity and,
where critical, the extreme lateral centers of
gravity must be established for each weight
established under 29.25. Such an extreme may not lie
beyond—
(a)
The extremes selected by the applicant;
(b)
The extremes within which the structure is proven;
or
(c)
The extremes within which compliance with the
applicable flight requirements is shown.
(a)
The empty weight and corresponding center of gravity
must be determined by weighing the rotorcraft
without the crew and payload, but with—
(1)
Fixed ballast;
(2)
Unusable fuel; and
(3)
Full operating fluids, including—
(i)
Oil;
(ii)
Hydraulic fluid; and
(iii) Other fluids required for normal operation of
rotorcraft systems, except water intended for
injection in the engines.
(b)
The condition of the rotorcraft at the time of
determining empty weight must be one that is well
defined and can be easily repeated, particularly
with respect to the weights of fuel, oil, coolant,
and installed equipment.
Removable ballast may be used in showing compliance
with the flight requirements of this subpart.
(a)
Main rotor speed limits. A range of main
rotor speeds must be established that—
(1)
With power on, provides adequate margin to
accommodate the variations in rotor speed occurring
in any appropriate maneuver, and is consistent with
the kind of governor or synchronizer used; and
(2)
With power off, allows each appropriate autorotative
maneuver to be performed throughout the ranges of
airspeed and weight for which certification is
requested.
(b)
Normal main rotor high pitch limit (power on).
For rotorcraft, except helicopters required to
have a main rotor low speed warning under paragraph
(e) of this section, it must be shown, with power on
and without exceeding approved engine maximum
limitations, that main rotor speeds substantially
less than the minimum approved main rotor speed will
not occur under any sustained flight condition. This
must be met by—
(1)
Appropriate setting of the main rotor high pitch
stop;
(2)
Inherent rotorcraft characteristics that make unsafe
low main rotor speeds unlikely; or
(3)
Adequate means to warn the pilot of unsafe main
rotor speeds.
(c)
Normal main rotor low pitch limit (power off).
It must be shown, with power off, that—
(1)
The normal main rotor low pitch limit provides
sufficient rotor speed, in any autorotative
condition, under the most critical combinations of
weight and airspeed; and
(2)
It is possible to prevent overspeeding of the rotor
without exceptional piloting skill.
(d)
Emergency high pitch. If the main rotor high
pitch stop is set to meet paragraph (b)(1) of this
section, and if that stop cannot be exceeded
inadvertently, additional pitch may be made
available for emergency use.
(e)
Main rotor low speed warning for helicopters.
For each single engine helicopter, and each
multiengine helicopter that does not have an
approved device that automatically increases power
on the operating engines when one engine fails,
there must be a main rotor low speed warning which
meets the following requirements:
(1)
The warning must be furnished to the pilot in all
flight conditions, including power-on and power-off
flight, when the speed of a main rotor approaches a
value that can jeopardize safe flight.
(2)
The warning may be furnished either through the
inherent aerodynamic qualities of the helicopter or
by a device.
(3)
The warning must be clear and distinct under all
conditions, and must be clearly distinguishable from
all other warnings. A visual device that requires
the attention of the crew within the cockpit is not
acceptable by itself.
(4)
If a warning device is used, the device must
automatically deactivate and reset when the
low-speed condition is corrected. If the device has
an audible warning, it must also be equipped with a
means for the pilot to manually silence the audible
warning before the low-speed condition is corrected.
29.45 General.
(a)
The performance prescribed in this subpart must be
determined—
(1)
With normal piloting skill and;
(2)
Without exceptionally favorable conditions.
(b)
Compliance with the performance requirements of this
subpart must be shown—
(1)
For still air at sea level with a standard
atmosphere and;
(2)
For the approved range of atmospheric variables.
(c)
The available power must correspond to engine power,
not exceeding the approved power, less—
(1)
Installation losses; and
(2)
The power absorbed by the accessories and services
at the values for which certification is requested
and approved.
(d)
For reciprocating engine-powered rotorcraft, the
performance, as affected by engine power, must be
based on a relative humidity of 80 percent in a
standard atmosphere.
(e)
For turbine engine-powered rotorcraft, the
performance, as affected by engine power, must be
based on a relative humidity of—
(1)
80 percent, at and below standard temperature; and
(2)
34 percent, at and above standard temperature plus
50 °F.
Between these two temperatures, the relative
humidity must vary linearly.
(f)
For turbine-engine-power rotorcraft, a means must be
provided to permit the pilot to determine prior to
takeoff that each engine is capable of developing
the power necessary to achieve the applicable
rotorcraft performance prescribed in this subpart.
(a)
For each Category A helicopter, the hovering
performance must be determined over the ranges of
weight, altitude, and temperature for which takeoff
data are scheduled—
(1)
With not more than takeoff power;
(2)
With the landing gear extended; and
(3)
At a height consistent with the procedure used in
establishing the takeoff, climbout, and rejected
takeoff paths.
(b)
For each Category B helicopter, the hovering
performance must be determined over the ranges of
weight, altitude, and temperature for which
certification is requested, with—
(1)
Takeoff power;
(2)
The landing gear extended; and
(3)
The helicopter in ground effect at a height
consistent with normal takeoff procedures.
(c)
For each helicopter, the out-of-ground effect
hovering performance must be determined over the
ranges of weight, altitude, and temperature for
which certification is requested with takeoff power.
(d)
For rotorcraft other than helicopters, the steady
rate of climb at the minimum operating speed must be
determined over the ranges of weight, altitude, and
temperature for which certification is requested
with—
(1)
Takeoff power; and
(2)
The landing gear extended.
(a)
The takeoff data required by 29.53, 29.55, 29.59,
29.60, 29.61, 29.62, 29.63, and 29.67 must be
determined—
(1)
At each weight, altitude, and temperature selected
by the applicant; and
(2)
With the operating engines within approved operating
limitations.
(b)
Takeoff data must—
(1)
Be determined on a smooth, dry, hard surface; and
(2)
Be corrected to assume a level takeoff surface.
(c)
No takeoff made to determine the data required by
this section may require exceptional piloting skill
or alertness, or exceptionally favorable conditions.
The
takeoff performance must be determined and scheduled
so that, if one engine fails at any time after the
start of takeoff, the rotorcraft can—
(a)
Return to, and stop safely on, the takeoff area; or
(b)
Continue the takeoff and climbout, and attain a
configuration and airspeed allowing compliance with
29.67(a)(2).
(a)
The TDP is the first point from which a continued
takeoff capability is assured under 29.59 and is the
last point in the takeoff path from which a rejected
takeoff is assured within the distance determined
under 29.62.
(b)
The TDP must be established in relation to the
takeoff path using no more than two parameters;
e.g., airspeed and height, to designate the TDP.
(c)
Determination of the TDP must include the pilot
recognition time interval following failure of the
critical engine.
(a)
The takeoff path extends from the point of
commencement of the takeoff procedure to a point at
which the rotorcraft is 1,000 feet above the takeoff
surface and compliance with 29.67(a)(2) is shown. In
addition—
(1)
The takeoff path must remain clear of the
height-velocity envelope established in accordance
with 29.87;
(2)
The rotorcraft must be flown to the engine failure
point; at which point, the critical engine must be
made inoperative and remain inoperative for the rest
of the takeoff;
(3)
After the critical engine is made inoperative, the
rotorcraft must continue to the takeoff decision
point, and then attain VTOSS;
(4)
Only primary controls may be used while attaining VTOSSand
while establishing a positive rate of climb.
Secondary controls that are located on the primary
controls may be used after a positive rate of climb
and VTOSSare established but in no case
less than 3 seconds after the critical engine is
made inoperative; and
(5)
After attaining VTOSSand a positive rate
of a climb, the landing gear may be retracted.
(b)
During the takeoff path determination made in
accordance with paragraph (a) of this section and
after attaining VTOSSand a positive rate
of climb, the climb must be continued at a speed as
close as practicable to, but not less than, VTOSSuntil
the rotorcraft is 200 feet above the takeoff
surface. During this interval, the climb performance
must meet or exceed that required by 29.67(a)(1).
(c)
During the continued takeoff, the rotorcraft shall
not descend below 15 feet above the takeoff surface
when the takeoff decision point is above 15 feet.
(d)
From 200 feet above the takeoff surface, the
rotorcraft takeoff path must be level or positive
until a height 1,000 feet above the takeoff surface
is attained with not less than the rate of climb
required by 29.67(a)(2). Any secondary or auxiliary
control may be used after attaining 200 feet above
the takeoff surface.
(e)
Takeoff distance will be determined in accordance
with 29.61.
(a)
The elevated heliport takeoff path extends from the
point of commencement of the takeoff procedure to a
point in the takeoff path at which the rotorcraft is
1,000 feet above the takeoff surface and compliance
with 29.67(a)(2) is shown. In addition—
(1)
The requirements of 29.59(a) must be met;
(2)
While attaining VTOSSand a positive rate
of climb, the rotorcraft may descend below the level
of the takeoff surface if, in so doing and when
clearing the elevated heliport edge, every part of
the rotorcraft clears all obstacles by at least 15
feet;
(3)
The vertical magnitude of any descent below the
takeoff surface must be determined; and
(4)
After attaining VTOSSand a positive rate
of climb, the landing gear may be retracted.
(b)
The scheduled takeoff weight must be such that the
climb requirements of 29.67 (a)(1) and (a)(2) will
be met.
(c)
Takeoff distance will be determined in accordance
with 29.61.
29.61 Takeoff distance: Category A
(a)
The normal takeoff distance is the horizontal
distance along the takeoff path from the start of
the takeoff to the point at which the rotorcraft
attains and remains at least 35 feet above the
takeoff surface, attains and maintains a speed of at
least VTOSS, and establishes a positive
rate of climb, assuming the critical engine failure
occurs at the engine failure point prior to the
takeoff decision point.
(b)
For elevated heliports, the takeoff distance is the
horizontal distance along the takeoff path from the
start of the takeoff to the point at which the
rotorcraft attains and maintains a speed of at least
VTOSSand establishes a positive rate of
climb, assuming the critical engine failure occurs
at the engine failure point prior to the takeoff
decision point.
The
rejected takeoff distance and procedures for each
condition where takeoff is approved will be
established with—
(a)
The takeoff path requirements of 29.59 and 29.60
being used up to the TDP where the critical engine
failure is recognized and the rotorcraft is landed
and brought to a complete stop on the takeoff
surface;
(b)
The remaining engines operating within approved
limits;
(c)
The landing gear remaining extended throughout the
entire rejected takeoff; and
(d)
The use of only the primary controls until the
rotorcraft is on the ground. Secondary controls
located on the primary control may not be used until
the rotorcraft is on the ground. Means other than
wheel brakes may be used to stop the rotorcraft if
the means are safe and reliable and consistent
results can be expected under normal operating
conditions.
The
horizontal distance required to take off and climb
over a 50-foot obstacle must be established with the
most unfavorable center of gravity. The takeoff may
be begun in any manner if—
(a)
The takeoff surface is defined;
(b)
Adequate safeguards are maintained to ensure proper
center of gravity and control positions; and
(c)
A landing can be made safely at any point along the
flight path if an engine fails.
Compliance with the requirements of 29.65 and 29.67
must be shown at each weight, altitude, and
temperature within the operational limits
established for the rotorcraft and with the most
unfavorable center of gravity for each
configuration. Cowl flaps, or other means of
controlling the engine-cooling air supply, will be
in the position that provides adequate cooling at
the temperatures and altitudes for which
certification is requested.
(a)
The steady rate of climb must be determined—
(1)
With maximum continuous power;
(2)
With the landing gear retracted; and
(3)
At Vyfor standard sea level conditions
and at speeds selected by the applicant for other
conditions.
(b)
For each Category B rotorcraft except helicopters,
the rate of climb determined under paragraph (a) of
this section must provide a steady climb gradient of
at least 1:6 under standard sea level conditions.
(a)
For Category A rotorcraft, in the critical takeoff
configuration existing along the takeoff path, the
following apply:
(1)
The steady rate of climb without ground effect, 200
feet above the takeoff surface, must be at least 100
feet per minute for each weight, altitude, and
temperature for which takeoff data are to be
scheduled with—
(i)
The critical engine inoperative and the remaining
engines within approved operating limitations,
except that for rotorcraft for which the use of
30-second/2-minute OEI power is requested, only the
2-minute OEI power may be used in showing compliance
with this paragraph;
(ii)
The landing gear extended; and
(iii) The takeoff safety speed selected by the
applicant.
(2)
The steady rate of climb without ground effect, 1000
feet above the takeoff surface, must be at least 150
feet per minute, for each weight, altitude, and
temperature for which takeoff data are to be
scheduled with—
(i)
The critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, or at 30-minute
OEI power for rotorcraft for which certification for
use of 30-minute OEI power is requested;
(ii)
The landing gear retracted; and
(iii) The speed selected by the applicant.
(3)
The steady rate of climb (or descent) in feet per
minute, at each altitude and temperature at which
the rotorcraft is expected to operate and at any
weight within the range of weights for which
certification is requested, must be determined with—
(i)
The critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, and at 30-minute
OEI power for rotorcraft for which certification for
the use of 30-minute OEI power is requested;
(ii)
The landing gear retracted; and
(iii) The speed selected by the applicant.
(b)
For multiengine Category B rotorcraft meeting the
Category A engine isolation requirements, the steady
rate of climb (or descent) must be determined at the
speed for best rate of climb (or minimum rate of
descent) at each altitude, temperature, and weight
at which the rotorcraft is expected to operate, with
the critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, and at 30-minute
OEI power for rotorcraft for which certification for
the use of 30-minute OEI power is requested.
For
each category B helicopter, except multiengine
helicopters meeting the requirements of 29.67(b) and
the powerplant installation requirements of category
A, the steady angle of glide must be determined in
autorotation—
(a)
At the forward speed for minimum rate of descent as
selected by the applicant;
(b)
At the forward speed for best glide angle;
(c)
At maximum weight; and
(d)
At the rotor speed or speeds selected by the
applicant.
(a)
For each rotorcraft—
(1)
The corrected landing data must be determined for a
smooth, dry, hard, and level surface;
(2)
The approach and landing must not require
exceptional piloting skill or exceptionally
favorable conditions; and
(3)
The landing must be made without excessive vertical
acceleration or tendency to bounce, nose over,
ground loop, porpoise, or water loop.
(b)
The landing data required by 29.77, 29.79, 29.81,
29.83, and 29.85 must be determined—
(1)
At each weight, altitude, and temperature for which
landing data are approved;
(2)
With each operating engine within approved operating
limitations; and
(3)
With the most unfavorable center of gravity.
(a)
The LDP is the last point in the approach and
landing path from which a balked landing can be
accomplished in accordance with 29.85.
(b)
Determination of the LDP must include the pilot
recognition time interval following failure of the
critical engine.
(a)
For Category A rotorcraft—
(1)
The landing performance must be determined and
scheduled so that if the critical engine fails at
any point in the approach path, the rotorcraft can
either land and stop safely or climb out and attain
a rotorcraft configuration and speed allowing
compliance with the climb requirement of
29.67(a)(2);
(2)
The approach and landing paths must be established
with the critical engine inoperative so that the
transition between each stage can be made smoothly
and safely;
(3)
The approach and landing speeds must be selected by
the applicant and must be appropriate to the type of
rotorcraft; and
(4)
The approach and landing path must be established to
avoid the critical areas of the height-velocity
envelope determined in accordance with 29.87.
(b)
It must be possible to make a safe landing on a
prepared landing surface after complete power
failure occurring during normal cruise.
The
horizontal distance required to land and come to a
complete stop (or to a speed of approximately 3
knots for water landings) from a point 50 ft above
the landing surface must be determined from the
approach and landing paths established in accordance
with 29.79.
(a)
For each Category B rotorcraft, the horizontal
distance required to land and come to a complete
stop (or to a speed of approximately 3 knots for
water landings) from a point 50 feet above the
landing surface must be determined with—
(1)
Speeds appropriate to the type of rotorcraft and
chosen by the applicant to avoid the critical areas
of the height-velocity envelope established under
29.87; and
(2)
The approach and landing made with power on and
within approved limits.
(b)
Each multi-engined Category B rotorcraft that meets
the powerplant installation requirements for
Category A must meet the requirements of—
(1)
Sections 29.79 and 29.81; or
(2)
Paragraph (a) of this section.
(c)
It must be possible to make a safe landing on a
prepared landing surface if complete power failure
occurs during normal cruise.
For
Category A rotorcraft, the balked landing path with
the critical engine inoperative must be established
so that—
(a)
The transition from each stage of the maneuver to
the next stage can be made smoothly and safely;
(b)
From the LDP on the approach path selected by the
applicant, a safe climbout can be made at speeds
allowing compliance with the climb requirements of
29.67(a)(1) and (2); and
(c)
The rotorcraft does not descend below 15 feet above
the landing surface. For elevated heliport
operations, descent may be below the level of the
landing surface provided the deck edge clearance of
29.60 is maintained and the descent (loss of height)
below the landing surface is determined.
(a)
If there is any combination of height and forward
velocity (including hover) under which a safe
landing cannot be made after failure of the critical
engine and with the remaining engines (where
applicable) operating within approved limits, a
height-velocity envelope must be established for—
(1)
All combinations of pressure altitude and ambient
temperature for which takeoff and landing are
approved; and
(2)
Weight from the maximum weight (at sea level) to the
highest weight approved for takeoff and landing at
each altitude. For helicopters, this weight need not
exceed the highest weight allowing hovering
out-of-ground effect at each altitude.
(b)
For single-engine or multiengine rotorcraft that do
not meet the Category A engine isolation
requirements, the height-velocity envelope for
complete power failure must be established.
Flight
Characteristics
The
rotorcraft must—
(a)
Except as specifically required in the applicable
section, meet the flight characteristics
requirements of this subpart—
(1)
At the approved operating altitudes and
temperatures;
(2)
Under any critical loading condition within the
range of weights and centers of gravity for which
certification is requested; and
(3)
For power-on operations, under any condition of
speed, power, and rotor r.p.m. for which
certification is requested; and
(4)
For power-off operations, under any condition of
speed, and rotor r.p.m. for which certification is
requested that is attainable with the controls
rigged in accordance with the approved rigging
instructions and tolerances;
(b)
Be able to maintain any required flight condition
and make a smooth transition from any flight
condition to any other flight condition without
exceptional piloting skill, alertness, or strength,
and without danger of exceeding the limit load
factor under any operating condition probable for
the type, including—
(1)
Sudden failure of one engine, for multiengine
rotorcraft meeting Transport Category A engine
isolation requirements;
(2)
Sudden, complete power failure, for other
rotorcraft; and
(3)
Sudden, complete control system failures specified
in 29.695 of this part; and
(c)
Have any additional characteristics required for
night or instrument operation, if certification for
those kinds of operation is requested. Requirements
for helicopter instrument flight are contained in
appendix B of this part.
(a)
The rotorcraft must be safely controllable and
maneuverable—
(1)
During steady flight; and
(2)
During any maneuver appropriate to the type,
including—
(i)
Takeoff;
(ii)
Climb;
(iii) Level flight;
(iv)
Turning flight;
(v)
Glide; and
(vi)
Landing (power on and power off).
(b)
The margin of cyclic control must allow satisfactory
roll and pitch control at VNEwith—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Critical rotor r.p.m.; and
(4)
Power off (except for helicopters demonstrating
compliance with paragraph (e) of this section) and
power on.
(c)
A wind velocity of not less than 17 knots must be
established in which the rotorcraft can be operated
without loss of control on or near the ground in any
maneuver appropriate to the type (such as crosswind
takeoffs, sideward flight, and rearward flight),
with—
(1)
Critical weight;
(2)
Critical center of gravity; and
(3)
Critical rotor r.p.m.
(d)
The rotorcraft, after (1) failure of one engine, in
the case of multiengine rotorcraft that meet
Transport Category A engine isolation requirements,
or (2) complete power failure in the case of other
rotorcraft, must be controllable over the range of
speeds and altitudes for which certification is
requested when such power failure occurs with
maximum continuous power and critical weight. No
corrective action time delay for any condition
following power failure may be less than—
(i)
For the cruise condition, one second, or normal
pilot reaction time (whichever is greater); and
(ii)
For any other condition, normal pilot reaction time.
(e)
For helicopters for which a VNE(power-off) is
established under 29.1505(c), compliance must be
demonstrated with the following requirements with
critical weight, critical center of gravity, and
critical rotor r.p.m.:
(1)
The helicopter must be safely slowed to VNE(power-off),
without exceptional pilot skill after the last
operating engine is made inoperative at power-on VNE.
(2)
At a speed of 1.1 VNE(power-off), the margin of
cyclic control must allow satisfactory roll and
pitch control with power off.
(a)
Longitudinal, lateral, directional, and collective
controls may not exhibit excessive breakout force,
friction, or preload.
(b)
Control system forces and free play may not inhibit
a smooth, direct rotorcraft response to control
system input.
29.161 Trim
control.
The
trim control—
(a)
Must trim any steady longitudinal, lateral, and
collective control forces to zero in level flight at
any appropriate speed; and
(b)
May not introduce any undesirable discontinuities in
control force gradients.
The
rotorcraft must be able to be flown, without undue
pilot fatigue or strain, in any normal maneuver for
a period of time as long as that expected in normal
operation. At least three landings and takeoffs must
be made during this demonstration.
(a)
The longitudinal control must be designed so that a
rearward movement of the control is necessary to
obtain a speed less than the trim speed, and a
forward movement of the control is necessary to
obtain a speed more than the trim speed.
(b)
With the throttle and collective pitch held constant
during the maneuvers specified in 29.175 (a) through
(c), the slope of the control position versus speed
curve must be positive throughout the full range of
altitude for which certification is requested.
(c)
During the maneuver specified in 29.175(d), the
longitudinal control position versus speed curve may
have a negative slope within the specified speed
range if the negative motion is not greater than 10
percent of total control travel.
(a)
Climb. Static longitudinal stability must be
shown in the climb condition at speeds from 0.85 VY,
or 15 knots below VY, whichever is less,
to 1.2 VYor 15 knots above VY,
whichever is greater, with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Maximum continuous power;
(4)
The landing gear retracted; and
(5)
The rotorcraft trimmed at V Y.
(b)
Cruise. Static longitudinal stability must be
shown in the cruise condition at speeds from 0.7
V Hor 0.7 V NE,whichever is less, to 1.1
V Hor 1.1 V
NE,whichever is
less, with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power for level flight at 0.9 V Hor 0.9 V
NE,whichever is less;
(4)
The landing gear retracted, and
(5)
The rotorcraft trimmed at 0.9 V Hor 0.9 V
NE,whichever is less.
(c)
Autorotation. Static longitudinal stability
must be shown in autorotation at airspeeds from 0.5
times the speed for minimum rate of descent, or 0.5
times the maximum range glide speed for Category A
rotorcraft, to VNEor to 1.1 VNE(power-off)
if VNE(power-off) is established under
29.1505(c), and with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power off;
(4)
The landing gear—
(i)
Retracted; and
(ii)
Extended; and
(5)
The rotorcraft trimmed at appropriate speeds found
necessary by the Administrator to demonstrate
stability throughout the prescribed speed range.
(d)
Hovering. For helicopters, the longitudinal
cyclic control must operate with the sense,
direction of motion, and position as prescribed in
29.173 between the maximum approved rearward speed
and a forward speed of 17 knots with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power required to maintain an approximate constant
height in ground effect;
(4)
The landing gear extended; and
(5)
The helicopter trimmed for hovering.
Static directional stability must be positive with
throttle and collective controls held constant at
the trim conditions specified in 29.175 (a), (b),
and (c). Sideslip angle must increase steadily with
directional control deflection for sideslip angles
up to ±10° from trim. Sufficient cues must accompany
sideslip to alert the pilot when approaching
sideslip limits.
Any
short-period oscillation occurring at any speed from
VYto VNEmust be positively
damped with the primary flight controls free and in
a fixed position.
The
rotorcraft must have satisfactory ground and water
handling characteristics, including freedom from
uncontrollable tendencies in any condition expected
in operation.
The
rotorcraft must be designed to withstand the loads
that would occur when the rotorcraft is taxied over
the roughest ground that may reasonably be expected
in normal operation.
If
certification for water operation is requested, no
spray characteristics during taxiing, takeoff, or
landing may obscure the vision of the pilot or
damage the rotors, propellers, or other parts of the
rotorcraft.
The
rotorcraft may have no dangerous tendency to
oscillate on the ground with the rotor turning.
Each
part of the rotorcraft must be free from excessive
vibration under each appropriate speed and power
condition.
(a)
Strength requirements are specified in terms of
limit loads (the maximum loads to be expected in
service) and ultimate loads (limit loads multiplied
by prescribed factors of safety). Unless otherwise
provided, prescribed loads are limit loads.
(b)
Unless otherwise provided, the specified air,
ground, and water loads must be placed in
equilibrium with inertia forces, considering each
item of mass in the rotorcraft. These loads must be
distributed to closely approximate or conservatively
represent actual conditions.
(c)
If deflections under load would significantly change
the distribution of external or internal loads, this
redistribution must be taken into account.
Unless otherwise provided, a factor of safety of 1.5
must be used. This factor applies to external and
inertia loads unless its application to the
resulting internal stresses is more conservative.
(a)
The structure must be able to support limit loads
without detrimental or permanent deformation. At any
load up to limit loads, the deformation may not
interfere with safe operation.
(b)
The structure must be able to support ultimate loads
without failure. This must be shown by—
(1)
Applying ultimate loads to the structure in a static
test for at least three seconds; or
(2)
Dynamic tests simulating actual load application.
(a)
Compliance with the strength and deformation
requirements of this subpart must be shown for each
critical loading condition accounting for the
environment to which the structure will be exposed
in operation. Structural analysis (static or
fatigue) may be used only if the structure conforms
to those structures for which experience has shown
this method to be reliable. In other cases,
substantiating load tests must be made.
(b)
Proof of compliance with the strength requirements
of this subpart must include—
(1)
Dynamic and endurance tests of rotors, rotor drives,
and rotor controls;
(2)
Limit load tests of the control system, including
control surfaces;
(3)
Operation tests of the control system;
(4)
Flight stress measurement tests;
(5)
Landing gear drop tests; and
(6)
Any additional tests required for new or unusual
design features.
The
following values and limitations must be established
to show compliance with the structural requirements
of this subpart:
(a)
The design maximum and design minimum weights.
(b)
The main rotor r.p.m. ranges, power on and power
off.
(c)
The maximum forward speeds for each main rotor r.p.m.
within the ranges determined under paragraph (b) of
this section.
(d)
The maximum rearward and sideward flight speeds.
(e)
The center of gravity limits corresponding to the
limitations determined under paragraphs (b), (c),
and (d) of this section.
(f)
The rotational speed ratios between each powerplant
and each connected rotating component.
(g)
The positive and negative limit maneuvering load
factors.
(a)
The flight load factor must be assumed to act normal
to the longitudinal axis of the rotorcraft, and to
be equal in magnitude and opposite in direction to
the rotorcraft inertia load factor at the center of
gravity.
(b)
Compliance with the flight load requirements of this
subpart must be shown—
(1)
At each weight from the design minimum weight to the
design maximum weight; and
(2)
With any practical distribution of disposable load
within the operating limitations in the Rotorcraft
Flight Manual.
The
rotorcraft must be designed for—
(a)
A limit maneuvering load factor ranging from a
positive limit of 3.5 to a negative limit of −1.0;
or
(b)
Any positive limit maneuvering load factor not less
than 2.0 and any negative limit maneuvering load
factor of not less than −0.5 for which—
(1)
The probability of being exceeded is shown by
analysis and flight tests to be extremely remote;
and
(2)
The selected values are appropriate to each weight
condition between the design maximum and design
minimum weights.
The
loads resulting from the application of limit
maneuvering load factors are assumed to act at the
center of each rotor hub and at each auxiliary
lifting surface, and to act in directions and with
distributions of load among the rotors and auxiliary
lifting surfaces, so as to represent each critical
maneuvering condition, including power-on and
power-off flight with the maximum design rotor tip
speed ratio. The rotor tip speed ratio is the ratio
of the rotorcraft flight velocity component in the
plane of the rotor disc to the rotational tip speed
of the rotor blades, and is expressed as follows:
where—
V =The
airspeed along the flight path (f.p.s.);
a =The angle
between the projection, in the plane of symmetry, of
the axis of no feathering and a line perpendicular
to the flight path (radians, positive when axis is
pointing aft);
Ω=The angular velocity of rotor (radians per
second); and
R =The rotor
radius (ft.).
Each
rotorcraft must be designed to withstand, at each
critical airspeed including hovering, the loads
resulting from vertical and horizontal gusts of 30
feet per second.
(a)
Each rotorcraft must be designed for the loads
resulting from the maneuvers specified in paragraphs
(b) and (c) of this section, with—
(1)
Unbalanced aerodynamic moments about the center of
gravity which the aircraft reacts to in a rational
or conservative manner considering the principal
masses furnishing the reacting inertia forces; and
(2)
Maximum main rotor speed.
(b)
To produce the load required in paragraph (a) of
this section, in unaccelerated flight with zero yaw,
at forward speeds from zero up to 0.6 VNE—
(1)
Displace the cockpit directional control suddenly to
the maximum deflection limited by the control stops
or by the maximum pilot force specified in
29.397(a);
(2)
Attain a resulting sideslip angle or 90°, whichever
is less; and
(3)
Return the directional control suddenly to neutral.
(c)
To produce the load required in paragraph (a) of the
section, in unaccelerated flight with zero yaw, at
forward speeds from 0.6 VNEup to VNEor
VH, whichever is less—
(1)
Displace the cockpit directional control suddenly to
the maximum deflection limited by the control stops
or by the maximum pilot force specified in
29.397(a);
(2)
Attain a resulting sideslip angle or 15°, whichever
is less, at the lesser speed of VNEor VH;
(3)
Vary the sideslip angles of paragraphs (b)(2) and
(c)(2) of this section directly with speed; and
(4)
Return the directional control suddenly to neutral.
The
limit engine torque may not be less than the
following:
(a)
For turbine engines, the highest of—
(1)
The mean torque for maximum continuous power
multiplied by 1.25;
(2)
The torque required by 29.923;
(3)
The torque required by 29.927; or
(4)
The torque imposed by sudden engine stoppage due to
malfunction or structural failure (such as
compressor jamming).
(b)
For reciprocating engines, the mean torque for
maximum continuous power multiplied by—
(1)
1.33, for engines with five or more cylinders; and
(2)
Two, three, and four, for engines with four, three,
and two cylinders, respectively.
Each
auxiliary rotor, each fixed or movable stabilizing
or control surface, and each system operating any
flight control must meet the requirements of 29.395
through 29.399, 29.411, and 29.427.
29.395 Control
system.
(a)
The reaction to the loads prescribed in 29.397 must
be provided by—
(1)
The control stops only;
(2)
The control locks only;
(3)
The irreversible mechanism only (with the mechanism
locked and with the control surface in the critical
positions for the effective parts of the system
within its limit of motion);
(4)
The attachment of the control system to the rotor
blade pitch control horn only (with the control in
the critical positions for the affected parts of the
system within the limits of its motion); and
(5)
The attachment of the control system to the control
surface horn (with the control in the critical
positions for the affected parts of the system
within the limits of its motion).
(b)
Each primary control system, including its
supporting structure, must be designed as follows:
(1)
The system must withstand loads resulting from the
limit pilot forces prescribed in 29.397;
(2)
Notwithstanding paragraph (b)(3) of this section,
when power-operated actuator controls or power boost
controls are used, the system must also withstand
the loads resulting from the limit pilot forces
prescribed in 29.397 in conjunction with the forces
output of each normally energized power device,
including any single power boost or actuator system
failure;
(3)
If the system design or the normal operating loads
are such that a part of the system cannot react to
the limit pilot forces prescribed in 29.397, that
part of the system must be designed to withstand the
maximum loads that can be obtained in normal
operation. The minimum design loads must, in any
case, provide a rugged system for service use,
including consideration of fatigue, jamming, ground
gusts, control inertia, and friction loads. In the
absence of a rational analysis, the design loads
resulting from 0.60 of the specified limit pilot
forces are acceptable minimum design loads; and
(4)
If operational loads may be exceeded through
jamming, ground gusts, control inertia, or friction,
the system must withstand the limit pilot forces
specified in 29.397, without yielding.
(a)
Except as provided in paragraph (b) of this section,
the limit pilot forces are as follows:
(1)
For foot controls, 130 pounds.
(2)
For stick controls, 100 pounds fore and aft, and 67
pounds laterally.
(b)
For flap, tab, stabilizer, rotor brake, and landing
gear operating controls, the following apply
(R=radius in inches):
(1)
Crank wheel, and lever controls, [1 + R]/3 × 50
pounds, but not less than 50 pounds nor more than
100 pounds for hand operated controls or 130 pounds
for foot operated controls, applied at any angle
within 20 degrees of the plane of motion of the
control.
(2)
Twist controls, 80R inch-pounds.
Each
dual primary flight control system must be able to
withstand the loads that result when pilot forces
not less than 0.75 times those obtained under 29.395
are applied—
(a)
In opposition; and
(b)
In the same direction.
(a)
It must be impossible for the tail rotor to contact
the landing surface during a normal landing.
(b)
If a tail rotor guard is required to show compliance
with paragraph (a) of this section—
(1)
Suitable design loads must be established for the
guard: and
(2)
The guard and its supporting structure must be
designed to withstand those loads.
(a)
Horizontal tail surfaces and their supporting
structure must be designed for unsymmetrical loads
arising from yawing and rotor wake effects in
combination with the prescribed flight conditions.
(b)
To meet the design criteria of paragraph (a) of this
section, in the absence of more rational data, both
of the following must be met:
(1)
One hundred percent of the maximum loading from the
symmetrical flight conditions acts on the surface on
one side of the plane of symmetry, and no loading
acts on the other side.
(2)
Fifty percent of the maximum loading from the
symmetrical flight conditions acts on the surface on
each side of the plane of symmetry, in opposite
directions.
(c)
For empennage arrangements where the horizontal tail
surfaces are supported by the vertical tail
surfaces, the vertical tail surfaces and supporting
structure must be designed for the combined vertical
and horizontal surface loads resulting from each
prescribed flight condition, considered separately.
The flight conditions must be selected so that the
maximum design loads are obtained on each surface.
In the absence of more rational data, the
unsymmetrical horizontal tail surface loading
distributions described in this section must be
assumed.
(a)
Loads and equilibrium. For limit ground
loads—
(1)
The limit ground loads obtained in the landing
conditions in this part must be considered to be
external loads that would occur in the rotorcraft
structure if it were acting as a rigid body; and
(2)
In each specified landing condition, the external
loads must be placed in equilibrium with linear and
angular inertia loads in a rational or conservative
manner.
(b)
Critical centers of gravity. The critical
centers of gravity within the range for which
certification is requested must be selected so that
the maximum design loads are obtained in each
landing gear element.
(a)
For specified landing conditions, a design maximum
weight must be used that is not less than the
maximum weight. A rotor lift may be assumed to act
through the center of gravity throughout the landing
impact. This lift may not exceed two-thirds of the
design maximum weight.
(b)
Unless otherwise prescribed, for each specified
landing condition, the rotorcraft must be designed
for a limit load factor of not less than the limit
inertia load factor substantiated under 29.725.
(c)
Triggering or actuating devices for additional or
supplementary energy absorption may not fail under
loads established in the tests prescribed in 29.725
and 29.727, but the factor of safety prescribed in
29.303 need not be used.
Unless otherwise prescribed, for each specified
landing condition, the tires must be assumed to be
in their static position and the shock absorbers to
be in their most critical position.
Sections 29.235, 29.479 through 29.485, and 29.493
apply to landing gear with two wheels aft, and one
or more wheels forward, of the center of gravity.
(a)
Attitudes. Under each of the loading
conditions prescribed in paragraph (b) of this
section, the rotorcraft is assumed to be in each of
the following level landing attitudes:
(1)
An attitude in which each wheel contacts the ground
simultaneously.
(2)
An attitude in which the aft wheels contact the
ground with the forward wheels just clear of the
ground.
(b)
Loading conditions. The rotorcraft must be
designed for the following landing loading
conditions:
(1)
Vertical loads applied under 29.471.
(2)
The loads resulting from a combination of the loads
applied under paragraph (b)(1) of this section with
drag loads at each wheel of not less than 25 percent
of the vertical load at that wheel.
(3)
The vertical load at the instant of peak drag load
combined with a drag component simulating the forces
required to accelerate the wheel rolling assembly up
to the specified ground speed, with—
(i)
The ground speed for determination of the spin-up
loads being at least 75 percent of the optimum
forward flight speed for minimum rate of descent in
autorotation; and
(ii)
The loading conditions of paragraph (b) applied to
the landing gear and its attaching structure only.
(4)
If there are two wheels forward, a distribution of
the loads applied to those wheels under paragraphs
(b)(1) and (2) of this section in a ratio of 40:60.
(c)
Pitching moments. Pitching moments are
assumed to be resisted by—
(1)
In the case of the attitude in paragraph (a)(1) of
this section, the forward landing gear; and
(2)
In the case of the attitude in paragraph (a)(2) of
this section, the angular inertia forces.
(a)
The rotorcraft is assumed to be in the maximum
nose-up attitude allowing ground clearance by each
part of the rotorcraft.
(b)
In this attitude, ground loads are assumed to act
perpendicular to the ground.
For
the one-wheel landing condition, the rotorcraft is
assumed to be in the level attitude and to contact
the ground on one aft wheel. In this attitude—
(a)
The vertical load must be the same as that obtained
on that side under 29.479(b)(1); and
(b)
The unbalanced external loads must be reacted by
rotorcraft inertia.
(a)
The rotorcraft is assumed to be in the level landing
attitude, with—
(1)
Side loads combined with one-half of the maximum
ground reactions obtained in the level landing
conditions of 29.479(b)(1); and
(2)
The loads obtained under paragraph (a)(1) of this
section applied—
(i)
At the ground contact point; or
(ii)
For full-swiveling gear, at the center of the axle.
(b)
The rotorcraft must be designed to withstand, at
ground contact—
(1)
When only the aft wheels contact the ground, side
loads of 0.8 times the vertical reaction acting
inward on one side and 0.6 times the vertical
reaction acting outward on the other side, all
combined with the vertical loads specified in
paragraph (a) of this section; and
(2)
When the wheels contact the ground simultaneously—
(i)
For the aft wheels, the side loads specified in
paragraph (b)(1) of this section; and
(ii)
For the forward wheels, a side load of 0.8 times the
vertical reaction combined with the vertical load
specified in paragraph (a) of this section.
Under braked roll conditions with the shock
absorbers in their static positions—
(a)
The limit vertical load must be based on a load
factor of at least—
(1)
1.33, for the attitude specified in 29.479(a)(1);
and
(2)
1.0, for the attitude specified in 29.479(a)(2); and
(b)
The structure must be designed to withstand, at the
ground contact point of each wheel with brakes, a
drag load of at least the lesser of—
(1)
The vertical load multiplied by a coefficient of
friction of 0.8; and
(2)
The maximum value based on limiting brake torque.
(a)
General. Rotorcraft with landing gear with
two wheels forward and one wheel aft of the center
of gravity must be designed for loading conditions
as prescribed in this section.
(b)
Level landing attitude with only the forward
wheels contacting the ground. In this attitude—
(1)
The vertical loads must be applied under 29.471
through 29.475;
(2)
The vertical load at each axle must be combined with
a drag load at that axle of not less than 25 percent
of that vertical load; and
(3)
Unbalanced pitching moments are assumed to be
resisted by angular inertia forces.
(c)
Level landing attitude with all wheels contacting
the ground simultaneously. In this attitude, the
rotorcraft must be designed for landing loading
conditions as prescribed in paragraph (b) of this
section.
(d)
Maximum nose-up attitude with only the rear wheel
contacting the ground. The attitude for this
condition must be the maximum nose-up attitude
expected in normal operation, including autorotative
landings. In this attitude—
(1)
The appropriate ground loads specified in paragraph
(b)(1) and (2) of this section must be determined
and applied, using a rational method to account for
the moment arm between the rear wheel ground
reaction and the rotorcraft center of gravity; or
(2)
The probability of landing with initial contact on
the rear wheel must be shown to be extremely remote.
(e)
Level landing attitude with only one forward
wheel contacting the ground. In this attitude,
the rotorcraft must be designed for ground loads as
specified in paragraph (b)(1) and (3) of this
section.
(f)
Side loads in the level landing attitude. In
the attitudes specified in paragraphs (b) and (c) of
this section, the following apply:
(1)
The side loads must be combined at each wheel with
one-half of the maximum vertical ground reactions
obtained for that wheel under paragraphs (b) and (c)
of this section. In this condition, the side loads
must be—
(i)
For the forward wheels, 0.8 times the vertical
reaction (on one side) acting inward, and 0.6 times
the vertical reaction (on the other side) acting
outward; and
(ii)
For the rear wheel, 0.8 times the vertical reaction.
(2)
The loads specified in paragraph (f)(1) of this
section must be applied—
(i)
At the ground contact point with the wheel in the
trailing position (for non-full swiveling landing
gear or for full swiveling landing gear with a lock,
steering device, or shimmy damper to keep the wheel
in the trailing position); or
(ii)
At the center of the axle (for full swiveling
landing gear without a lock, steering device, or
shimmy damper).
(g)
Braked roll conditions in the level landing
attitude. In the attitudes specified in
paragraphs (b) and (c) of this section, and with the
shock absorbers in their static positions, the
rotorcraft must be designed for braked roll loads as
follows:
(1)
The limit vertical load must be based on a limit
vertical load factor of not less than—
(i)
1.0, for the attitude specified in paragraph (b) of
this section; and
(ii)
1.33, for the attitude specified in paragraph (c) of
this section.
(2)
For each wheel with brakes, a drag load must be
applied, at the ground contact point, of not less
than the lesser of—
(i)
0.8 times the vertical load; and
(ii)
The maximum based on limiting brake torque.
(h)
Rear wheel turning loads in the static ground
attitude. In the static ground attitude, and
with the shock absorbers and tires in their static
positions, the rotorcraft must be designed for rear
wheel turning loads as follows:
(1)
A vertical ground reaction equal to the static load
on the rear wheel must be combined with an equal
side load.
(2)
The load specified in paragraph (h)(1) of this
section must be applied to the rear landing gear—
(i)
Through the axle, if there is a swivel (the rear
wheel being assumed to be swiveled 90 degrees to the
longitudinal axis of the rotorcraft); or
(ii)
At the ground contact point if there is a lock,
steering device or shimmy damper (the rear wheel
being assumed to be in the trailing position).
(i)
Taxiing condition. The rotorcraft and its
landing gear must be designed for the loads that
would occur when the rotorcraft is taxied over the
roughest ground that may reasonably be expected in
normal operation.
(a)
General. Rotorcraft with landing gear with
skids must be designed for the loading conditions
specified in this section. In showing compliance
with this section, the following apply:
(1)
The design maximum weight, center of gravity, and
load factor must be determined under 29.471 through
29.475.
(2)
Structural yielding of elastic spring members under
limit loads is acceptable.
(3)
Design ultimate loads for elastic spring members
need not exceed those obtained in a drop test of the
gear with—
(i)
A drop height of 1.5 times that specified in 29.725;
and
(ii)
An assumed rotor lift of not more than 1.5 times
that used in the limit drop tests prescribed in
29.725.
(4)
Compliance with paragraph (b) through (e) of this
section must be shown with—
(i)
The gear in its most critically deflected position
for the landing condition being considered; and
(ii)
The ground reactions rationally distributed along
the bottom of the skid tube.
(b)
Vertical reactions in the level landing attitude.
In the level attitude, and with the rotorcraft
contacting the ground along the bottom of both
skids, the vertical reactions must be applied as
prescribed in paragraph (a) of this section.
(c)
Drag reactions in the level landing attitude.
In the level attitude, and with the rotorcraft
contacting the ground along the bottom of both
skids, the following apply:
(1)
The vertical reactions must be combined with
horizontal drag reactions of 50 percent of the
vertical reaction applied at the ground.
(2)
The resultant ground loads must equal the vertical
load specified in paragraph (b) of this section.
(d)
Sideloads in the level landing attitude. In
the level attitude, and with the rotorcraft
contacting the ground along the bottom of both
skids, the following apply:
(1)
The vertical ground reaction must be—
(i)
Equal to the vertical loads obtained in the
condition specified in paragraph (b) of this
section; and
(ii)
Divided equally among the skids.
(2)
The vertical ground reactions must be combined with
a horizontal sideload of 25 percent of their value.
(3)
The total sideload must be applied equally between
skids and along the length of the skids.
(4)
The unbalanced moments are assumed to be resisted by
angular inertia.
(5)
The skid gear must be investigated for—
(i)
Inward acting sideloads; and
(ii)
Outward acting sideloads.
(e)
One-skid landing loads in the level attitude.
In the level attitude, and with the rotorcraft
contacting the ground along the bottom of one skid
only, the following apply:
(1)
The vertical load on the ground contact side must be
the same as that obtained on that side in the
condition specified in paragraph (b) of this
section.
(2)
The unbalanced moments are assumed to be resisted by
angular inertia.
(f)
Special conditions. In addition to the
conditions specified in paragraphs (b) and (c) of
this section, the rotorcraft must be designed for
the following ground reactions:
(1)
A ground reaction load acting up and aft at an angle
of 45 degrees to the longitudinal axis of the
rotorcraft. This load must be—
(i)
Equal to 1.33 times the maximum weight;
(ii)
Distributed symmetrically among the skids;
(iii) Concentrated at the forward end of the
straight part of the skid tube; and
(iv)
Applied only to the forward end of the skid tube and
its attachment to the rotorcraft.
(2)
With the rotorcraft in the level landing attitude, a
vertical ground reaction load equal to one-half of
the vertical load determined under paragraph (b) of
this section. This load must be—
(i)
Applied only to the skid tube and its attachment to
the rotorcraft; and
(ii)
Distributed equally over 33.3 percent of the length
between the skid tube attachments and centrally
located midway between the skid tube attachments.
If
certification for ski operation is requested, the
rotorcraft, with skis, must be designed to withstand
the following loading conditions (where P is
the maximum static weight on each ski with the
rotorcraft at design maximum weight, and n is
the limit load factor determined under 29.473(b)):
(a)
Up-load conditions in which—
(1)
A vertical load of Pn and a horizontal load
of Pn/4 are simultaneously applied at the
pedestal bearings; and
(2)
A vertical load of 1.33 P is applied at the
pedestal bearings.
(b)
A side load condition in which a side load of 0.35
Pn is applied at the pedestal bearings in a
horizontal plane perpendicular to the centerline of
the rotorcraft.
(c)
A torque-load condition in which a torque load of
1.33 P (in foot-pounds) is applied to the ski
about the vertical axis through the centerline of
the pedestal bearings.
(a)
In dual-wheel gear units, 60 percent of the total
ground reaction for the gear unit must be applied to
one wheel and 40 percent to the other.
(b)
To provide for the case of one deflated tire, 60
percent of the specified load for the gear unit must
be applied to either wheel except that the vertical
ground reaction may not be less than the full static
value.
(c)
In determining the total load on a gear unit, the
transverse shift in the load centroid, due to
unsymmetrical load distribution on the wheels, may
be neglected.
29.519 Hull type rotorcraft: Water-based and amphibian.
(a)
General. For hull type rotorcraft, the
structure must be designed to withstand the water
loading set forth in paragraphs (b), (c), and (d) of
this section considering the most severe wave
heights and profiles for which approval is desired.
The loads for the landing conditions of paragraphs
(b) and (c) of this section must be developed and
distributed along and among the hull and auxiliary
floats, if used, in a rational and conservative
manner, assuming a rotor lift not exceeding
two-thirds of the rotorcraft weight to act
throughout the landing impact.
(b)
Vertical landing conditions. The rotorcraft
must initially contact the most critical wave
surface at zero forward speed in likely pitch and
roll attitudes which result in critical design
loadings. The vertical descent velocity may not be
less than 6.5 feet per second relative to the mean
water surface.
(c)
Forward speed landing conditions. The
rotorcraft must contact the most critical wave at
forward velocities from zero up to 30 knots in
likely pitch, roll, and yaw attitudes and with a
vertical descent velocity of not less than 6.5 feet
per second relative to the mean water surface. A
maximum forward velocity of less than 30 knots may
be used in design if it can be demonstrated that the
forward velocity selected would not be exceeded in a
normal one-engine-out landing.
(d)
Auxiliary float immersion condition. In
addition to the loads from the landing conditions,
the auxiliary float, and its support and attaching
structure in the hull, must be designed for the load
developed by a fully immersed float unless it can be
shown that full immersion of the float is unlikely,
in which case the highest likely float buoyancy load
must be applied that considers loading of the float
immersed to create restoring moments compensating
for upsetting moments caused by side wind,
asymmetrical rotorcraft loading, water wave action,
and rotorcraft inertia.
If
certification for float operation (including float
amphibian operation) is requested, the rotorcraft,
with floats, must be designed to withstand the
following loading conditions (where the limit load
factor is determined under 29.473(b) or assumed to
be equal to that determined for wheel landing gear):
(a)
Up-load conditions in which—
(1)
A load is applied so that, with the rotorcraft in
the static level attitude, the resultant water
reaction passes vertically through the center of
gravity; and
(2)
The vertical load prescribed in paragraph (a)(1) of
this section is applied simultaneously with an aft
component of 0.25 times the vertical component
(b)
A side load condition in which—
(1)
A vertical load of 0.75 times the total vertical
load specified in paragraph (a)(1) of this section
is divided equally among the floats; and
(2)
For each float, the load share determined under
paragraph (b)(1) of this section, combined with a
total side load of 0.25 times the total vertical
load specified in paragraph (b)(1) of this section,
is applied to that float only.
29.547 Main and tail rotor structure.
(a)
A rotor is an assembly of rotating components, which
includes the rotor hub, blades, blade dampers, the
pitch control mechanisms, and all other parts that
rotate with the assembly.
(b)
Each rotor assembly must be designed as prescribed
in this section and must function safely for the
critical flight load and operating conditions. A
design assessment must be performed, including a
detailed failure analysis to identify all failures
that will prevent continued safe flight or safe
landing, and must identify the means to minimize the
likelihood of their occurrence.
(c)
The rotor structure must be designed to withstand
the following loads prescribed in 29.337 through
29.341 and 29.351:
(1)
Critical flight loads.
(2)
Limit loads occurring under normal conditions of
autorotation.
(d)
The rotor structure must be designed to withstand
loads simulating—
(1)
For the rotor blades, hubs, and flapping hinges, the
impact force of each blade against its stop during
ground operation; and
(2)
Any other critical condition expected in normal
operation.
(e)
The rotor structure must be designed to withstand
the limit torque at any rotational speed, including
zero.
In
addition:
(1)
The limit torque need not be greater than the torque
defined by a torque limiting device (where
provided), and may not be less than the greater of—
(i)
The maximum torque likely to be transmitted to the
rotor structure, in either direction, by the rotor
drive or by sudden application of the rotor brake;
and
(ii)
For the main rotor, the limit engine torque
specified in 29.361.
(2)
The limit torque must be equally and rationally
distributed to the rotor blades.
(a)
Each fuselage and rotor pylon structure must be
designed to withstand—
(1)
The critical loads prescribed in 29.337 through
29.341, and 29.351;
(2)
The applicable ground loads prescribed in 29.235,
29.471 through 29.485, 29.493, 29.497, 29.505, and
29.521; and
(3)
The loads prescribed in 29.547 (d)(1) and (e)(1)(i).
(b)
Auxiliary rotor thrust, the torque reaction of each
rotor drive system, and the balancing air and
inertia loads occurring under accelerated flight
conditions, must be considered.
(c)
Each engine mount and adjacent fuselage structure
must be designed to withstand the loads occurring
under accelerated flight and landing conditions,
including engine torque.
(d)
[Reserved]
(e)
If approval for the use of 21/2-minute OEI power is
requested, each engine mount and adjacent structure
must be designed to withstand the loads resulting
from a limit torque equal to 1.25 times the mean
torque for 21/2-minute OEI power combined with 1g
flight loads.
Each
auxiliary lifting surface must be designed to
withstand—
(a)
The critical flight loads in 29.337 through 29.341,
and 29.351;
(b)
the applicable ground loads in 29.235, 29.471
through 29.485, 29.493, 29.505, and 29.521; and
(c)
Any other critical condition expected in normal
operation.
(a)
The rotorcraft, although it may be damaged in
emergency landing conditions on land or water, must
be designed as prescribed in this section to protect
the occupants under those conditions.
(b)
The structure must be designed to give each occupant
every reasonable chance of escaping serious injury
in a crash landing when—
(1)
Proper use is made of seats, belts, and other safety
design provisions;
(2)
The wheels are retracted (where applicable); and
(3)
Each occupant and each item of mass inside the cabin
that could injure an occupant is restrained when
subjected to the following ultimate inertial load
factors relative to the surrounding structure:
(i)
Upward—4g.
(ii)
Forward—16g.
(iii) Sideward—8g.
(iv)
Downward—20g, after the intended displacement of the
seat device.
(v)
Rearward—1.5g.
(c)
The supporting structure must be designed to
restrain under any ultimate inertial load factor up
to those specified in this paragraph, any item of
mass above and/or behind the crew and passenger
compartment that could injure an occupant if it came
loose in an emergency landing. Items of mass to be
considered include, but are not limited to, rotors,
transmission, and engines. The items of mass must be
restrained for the following ultimate inertial load
factors:
(1)
Upward—1.5g.
(2)
Forward—12g.
(3)
Sideward—6g.
(4)
Downward—12g.
(5)
Rearward—1.5g.
(d)
Any fuselage structure in the area of internal fuel
tanks below the passenger floor level must be
designed to resist the following ultimate inertial
factors and loads, and to protect the fuel tanks
from rupture, if rupture is likely when those loads
are applied to that area:
(1)
Upward—1.5g.
(2)
Forward—4.0g.
(3)
Sideward—2.0g.
(4)
Downward—4.0g.
(a)
The rotorcraft, although it may be damaged in a
crash landing, must be designed to reasonably
protect each occupant when—
(1)
The occupant properly uses the seats, safety belts,
and shoulder harnesses provided in the design; and
(2)
The occupant is exposed to loads equivalent to those
resulting from the conditions prescribed in this
section.
(b)
Each seat type design or other seating device
approved for crew or passenger occupancy during
takeoff and landing must successfully complete
dynamic tests or be demonstrated by rational
analysis based on dynamic tests of a similar type
seat in accordance with the following criteria. The
tests must be conducted with an occupant simulated
by a 170-pound anthropomorphic test dummy (ATD), as
defined by 49 CFR 572, Subpart B, or its equivalent,
sitting in the normal upright position.
(1)
A change in downward velocity of not less than 30
feet per second when the seat or other seating
device is oriented in its nominal position with
respect to the rotorcraft's reference system, the
rotorcraft's longitudinal axis is canted upward 60°
with respect to the impact velocity vector, and the
rotorcraft's lateral axis is perpendicular to a
vertical plane containing the impact velocity vector
and the rotorcraft's longitudinal axis. Peak floor
deceleration must occur in not more than 0.031
seconds after impact and must reach a minimum of
30g's.
(2)
A change in forward velocity of not less than 42
feet per second when the seat or other seating
device is oriented in its nominal position with
respect to the rotorcraft's reference system, the
rotorcraft's longitudinal axis is yawed 10° either
right or left of the impact velocity vector
(whichever would cause the greatest load on the
shoulder harness), the rotorcraft's lateral axis is
contained in a horizontal plane containing the
impact velocity vector, and the rotorcraft's
vertical axis is perpendicular to a horizontal plane
containing the impact velocity vector. Peak floor
deceleration must occur in not more than 0.071
seconds after impact and must reach a minimum of
18.4g's.
(3)
Where floor rails or floor or sidewall attachment
devices are used to attach the seating devices to
the airframe structure for the conditions of this
section, the rails or devices must be misaligned
with respect to each other by at least 10°
vertically (i.e., pitch out of parallel) and by at
least a 10° lateral roll, with the directions
optional, to account for possible floor warp.
(c)
Compliance with the following must be shown:
(1)
The seating device system must remain intact
although it may experience separation intended as
part of its design.
(2)
The attachment between the seating device and the
airframe structure must remain intact although the
structure may have exceeded its limit load.
(3)
The ATD's shoulder harness strap or straps must
remain on or in the immediate vicinity of the ATD's
shoulder during the impact.
(4)
The safety belt must remain on the ATD's pelvis
during the impact.
(5)
The ATD's head either does not contact any portion
of the crew or passenger compartment or, if contact
is made, the head impact does not exceed a head
injury criteria (HIC) of 1,000 as determined by this
equation.
Where: a(t) is the resultant acceleration at the
center of gravity of the head form expressed as a
multiple of g (the acceleration of gravity) and t2−
t1is the time duration, in seconds, of
major head impact, not to exceed 0.05 seconds.
(6)
Loads in individual shoulder harness straps must not
exceed 1,750 pounds. If dual straps are used for
retaining the upper torso, the total harness strap
loads must not exceed 2,000 pounds.
(7)
The maximum compressive load measured between the
pelvis and the lumbar column of the ATD must not
exceed 1,500 pounds.
(d)
An alternate approach that achieves an equivalent or
greater level of occupant protection, as required by
this section, must be substantiated on a rational
basis.
If
certification with ditching provisions is requested,
structural strength for ditching must meet the
requirements of this section and 29.801(e).
(a)
Forward speed landing conditions. The
rotorcraft must initially contact the most critical
wave for reasonably probable water conditions at
forward velocities from zero up to 30 knots in
likely pitch, roll, and yaw attitudes. The
rotorcraft limit vertical descent velocity may not
be less than 5 feet per second relative to the mean
water surface. Rotor lift may be used to act through
the center of gravity throughout the landing impact.
This lift may not exceed two-thirds of the design
maximum weight. A maximum forward velocity of less
than 30 knots may be used in design if it can be
demonstrated that the forward velocity selected
would not be exceeded in a normal one-engine-out
touchdown.
(b)
Auxiliary or emergency float conditions —(1)
Floats fixed or deployed before initial water
contact. In addition to the landing loads in
paragraph (a) of this section, each auxiliary or
emergency float, or its support and attaching
structure in the airframe or fuselage, must be
designed for the load developed by a fully immersed
float unless it can be shown that full immersion is
unlikely. If full immersion is unlikely, the highest
likely float buoyancy load must be applied. The
highest likely buoyancy load must include
consideration of a partially immersed float creating
restoring moments to compensate the upsetting
moments caused by side wind, unsymmetrical
rotorcraft loading, water wave action, rotorcraft
inertia, and probable structural damage and leakage
considered under 29.801(d). Maximum roll and pitch
angles determined from compliance with 29.801(d) may
be used, if significant, to determine the extent of
immersion of each float. If the floats are deployed
in flight, appropriate air loads derived from the
flight limitations with the floats deployed shall be
used in substantiation of the floats and their
attachment to the rotorcraft. For this purpose, the
design airspeed for limit load is the float deployed
airspeed operating limit multiplied by 1.11.
(2)
Floats deployed after initial water contact.
Each float must be designed for full or partial
immersion prescribed in paragraph (b)(1) of this
section. In addition, each float must be designed
for combined vertical and drag loads using a
relative limit speed of 20 knots between the
rotorcraft and the water. The vertical load may not
be less than the highest likely buoyancy load
determined under paragraph (b)(1) of this section.
Fatigue Evaluation
(a)
General. An evaluation of the strength of
principal elements, detail design points, and
fabrication techniques must show that catastrophic
failure due to fatigue, considering the effects of
environment, intrinsic/discrete flaws, or accidental
damage will be avoided. Parts to be evaluated
include, but are not limited to, rotors, rotor drive
systems between the engines and rotor hubs,
controls, fuselage, fixed and movable control
surfaces, engine and transmission mountings, landing
gear, and their related primary attachments. In
addition, the following apply:
(1)
Each evaluation required by this section must
include—
(i)
The identification of principal structural elements,
the failure of which could result in catastrophic
failure of the rotorcraft;
(ii)
In-flight measurement in determining the loads or
stresses for items in paragraph (a)(1)(i) of this
section in all critical conditions throughout the
range of limitations in 29.309 (including altitude
effects), except that maneuvering load factors need
not exceed the maximum values expected in
operations; and
(iii) Loading spectra as severe as those expected in
operation based on loads or stresses determined
under paragraph (a)(1)(ii) of this section,
including external load operations, if applicable,
and other high frequency power cycle operations.
(2)
Based on the evaluations required by this section,
inspections, replacement times, combinations
thereof, or other procedures must be established as
necessary to avoid catastrophic failure. These
inspections, replacement times, combinations
thereof, or other procedures must be included in the
airworthiness limitations section of the
Instructions for Continued Airworthiness required by
29.1529 and section A29.4 of appendix A of this
part.
(b)
Fatigue tolerance evaluation (including tolerance
to flaws). The structure must be shown by
analysis supported by test evidence and, if
available, service experience to be of fatigue
tolerant design. The fatigue tolerance evaluation
must include the requirements of either paragraph
(b)(1), (2), or (3) of this section, or a
combination thereof, and also must include a
determination of the probable locations and modes of
damage caused by fatigue, considering environmental
effects, intrinsic/discrete flaws, or accidental
damage. Compliance with the flaw tolerance
requirements of paragraph (b)(1) or (2) of this
section is required unless the applicant establishes
that these fatigue flaw tolerant methods for a
particular structure cannot be achieved within the
limitations of geometry, inspectability, or good
design practice. Under these circumstances, the
safe-life evaluation of paragraph (b)(3) of this
section is required.
(1)
Flaw tolerant safe-life evaluation. It must
be shown that the structure, with flaws present, is
able to withstand repeated loads of variable
magnitude without detectable flaw growth for the
following time intervals—
(i)
Life of the rotorcraft; or
(ii)
Within a replacement time furnished under section
A29.4 of appendix A to this part.
(2)
Fail-safe (residual strength after flaw growth)
evaluation. It must be shown that the structure
remaining after a partial failure is able to
withstand design limit loads without failure within
an inspection period furnished under section A29.4
of appendix A to this part. Limit loads are defined
in 29.301(a).
(i)
The residual strength evaluation must show that the
remaining structure after flaw growth is able to
withstand design limit loads without failure within
its operational life.
(ii)
Inspection intervals and methods must be established
as necessary to ensure that failures are detected
prior to residual strength conditions being reached.
(iii) If significant changes in structural stiffness
or geometry, or both, follow from a structural
failure or partial failure, the effect on flaw
tolerance must be further investigated.
(3)
Safe-life evaluation. It must be shown that
the structure is able to withstand repeated loads of
variable magnitude without detectable cracks for the
following time intervals—
(i)
Life of the rotorcraft; or
(ii)
Within a replacement time furnished under section
A29.4 of appendix A to this part.
(a)
The rotorcraft may have no design features or
details that experience has shown to be hazardous or
unreliable.
(b)
The suitability of each questionable design detail
and part must be established by tests.
(a)
Critical part. A critical part is a part, the
failure of which could have a catastrophic effect
upon the rotocraft, and for which critical
characterists have been identified which must be
controlled to ensure the required level of
integrity.
(b)
If the type design includes critical parts, a
critical parts list shall be established. Procedures
shall be established to define the critical design
characteristics, identify processes that affect
those characteristics, and identify the design
change and process change controls necessary for
showing compliance with the quality assurance
requirements of part 21 of this chapter.
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)
Meet approved specifications 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.
(a)
The methods of fabrication used must produce
consistently sound structures. If a fabrication
process (such as gluing, spot welding, or
heat-treating) requires close control to reach this
objective, the process must be performed according
to an approved process specification.
(b)
Each new aircraft fabrication method must be
substantiated by a test program.
29.607 Fasteners.
(a)
Each removable bolt, screw, nut, pin, or other
fastener whose loss could jeopardize the safe
operation of the rotorcraft must incorporate two
separate locking devices. The fastener and its
locking devices may not be adversely affected by the
environmental conditions associated with the
particular installation.
(b)
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.
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 to prevent the accumulation of corrosive,
flammable, or noxious fluids.
(a)
The rotorcraft structure must be protected against
catastrophic effects from lightning.
(b)
For metallic components, compliance with paragraph
(a) of this section may be shown by—
(1)
Electrically bonding the components properly to the
airframe; or
(2)
Designing the components so that a strike will not
endanger the rotorcraft.
(c)
For nonmetallic components, compliance with
paragraph (a) of this section may be shown by—
(1)
Designing the components to minimize the effect of a
strike; or
(2)
Incorporating acceptable means of diverting the
resulting electrical current to not endanger the
rotorcraft.
(d)
The electric bonding and protection against
lightning and static electricity must—
(1)
Minimize the accumulation of electrostatic charge;
(2)
Minimize the risk of electric shock to crew,
passengers, and service and maintenance personnel
using normal precautions;
(3)
Provide and electrical return path, under both
normal and fault conditions, on rotorcraft having
grounded electrical systems; and
(4)
Reduce to an acceptable level the effects of
lightning and static electricity on the functioning
of essential electrical and electronic equipment.
There must be means to allow close examination of
each part that requires—
(a)
Recurring inspection;
(b)
Adjustment for proper alignment and functioning; or
(c)
Lubrication.
(a)
Material strength properties must be based on enough
tests of material meeting specifications to
establish design values on a statistical basis.
(b)
Design values must be chosen to minimize the
probability of structural failure due to material
variability. Except as provided in paragraphs (d)
and (e) of this section, compliance with this
paragraph must be shown by selecting design values
that 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; and
(2)
For redundant structures, those 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 strength, detail design, and fabrication of the
structure must minimize the probability of
disastrous fatigue failure, particularly at points
of stress concentration.
(d)
Design values may be those contained in the
following publications (available from the Naval
Publications and Forms Center, 5801 Tabor Avenue,
Philadelphia, PA 19120) or other values approved by
the Administrator:
(1)
MIL—HDBK–5, “Metallic Materials and Elements for
Flight Vehicle Structure”.
(2)
MIL—HDBK–17, “Plastics for Flight Vehicles”.
(3)
ANC–18, “Design of Wood Aircraft Structures”.
(4)
MIL—HDBK–23, “Composite Construction for Flight
Vehicles”.
(e)
Other design values may be used if a selection of
the material is made in which a specimen of each
individual item is tested before use and it is
determined that the actual strength properties of
that particular item will equal or exceed those used
in design.
(a)
The special factors prescribed in 29.621 through
29.625 apply to each part of the structure whose
strength is—
(1)
Uncertain;
(2)
Likely to deteriorate in service before normal
replacement; or
(3)
Subject to appreciable variability due to—
(i)
Uncertainties in manufacturing processes; or
(ii)
Uncertainties in inspection methods.
(b)
For each part of the rotorcraft to which 29.621
through 29.625 apply, the factor of safety
prescribed in 29.303 must be multiplied by a special
factor equal to—
(1)
The applicable special factors prescribed in 29.621
through 29.625; or
(2)
Any other factor great enough to ensure that the
probability of the part being under strength because
of the uncertainties specified in paragraph (a) of
this section is extremely remote.
(a)
General. The factors, tests, and inspections
specified in paragraphs (b) and (c) 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 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 rotorcraft or result in serious
injury to any occupant, 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 (for
ferromagnetic materials) or penetrant (for
non-ferromagnetic materials) inspection methods or
approved equivalent 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 29.305 at an ultimate
load corresponding to a casting factor of 1.25; and
(ii)
The deformation requirements of 29.305 at a load of
1.15 times the limit load.
(d)
Non-critical 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 greater |
100 percent visual. |
|
Less than 2.0, greater than 1.5 |
100 percent visual, and magnetic particle
(ferromagnetic materials), penetrant
(non-ferromagnetic materials), or approved
equivalent inspection methods. |
|
1.25 through 1.50 |
100 percent visual, and magnetic particle
(ferromagnetic materials), penetrant
(non-ferromagnetic materials), and radiographic
or approved equivalent inspection methods. |
(2)
The percentage of castings inspected by non-visual
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.
(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 a
bearing factor large enough to provide for the
effects of normal relative motion.
(b)
No bearing factor need be used on a part for which
any larger special factor is prescribed.
For
each fitting (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); and
(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)
Each seat, berth, litter, safety belt, and harness
attachment to the structure must be shown by
analysis, tests, or both, to be able to withstand
the inertia forces prescribed in 29.561(b)(3)
multiplied by a fitting factor of 1.33.
Each
aerodynamic surface of the rotorcraft must be free
from flutter and divergence under each appropriate
speed and power condition.
The
rotorcraft must be designed to ensure capability of
continued safe flight and landing (for Category A)
or safe landing (for Category B) after impact with a
2.2-lb (1.0 kg) bird when the velocity of the
rotorcraft (relative to the bird along the flight
path of the rotorcraft) is equal to VNEor
VH(whichever is the lesser) at altitudes
up to 8,000 feet. Compliance must be shown by tests
or by analysis based on tests carried out on
sufficiently representative structures of similar
design.
(a)
For each rotor blade—
(1)
There must be means for venting the internal
pressure of the blade;
(2)
Drainage holes must be provided for the blade; and
(3)
The blade must be designed to prevent water from
becoming trapped in it.
(b)
Paragraphs (a)(1) and (2) of this section does not
apply to sealed rotor blades capable of withstanding
the maximum pressure differentials expected in
service.
(a)
The rotor and blades must be mass balanced as
necessary to—
(1)
Prevent excessive vibration; and
(2)
Prevent flutter at any speed up to the maximum
forward speed.
(b)
The structural integrity of the mass balance
installation must be substantiated.
There must be enough clearance between the rotor
blades and other parts of the structure to prevent
the blades from striking any part of the structure
during any operating condition.
29.663 Ground
resonance prevention means.
(a)
The reliability of the means for preventing ground
resonance must be shown either by analysis and
tests, or reliable service experience, or by showing
through analysis or tests that malfunction or
failure of a single means will not cause ground
resonance.
(b)
The probable range of variations, during service, of
the damping action of the ground resonance
prevention means must be established and must be
investigated during the test required by 29.241.
(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 any incorrect
assembly that could result in the malfunction of the
system.
(c)
A means must be provided to allow full control
movement of all primary flight controls prior to
flight, or a means must be provided that will allow
the pilot to determine that full control authority
is available prior to flight.
If
the functioning of stability augmentation or other
automatic or power-operated system is necessary to
show compliance with the flight characteristics
requirements of this part, the system must comply
with 29.671 of this part and the following:
(a)
A warning which is clearly distinguishable to the
pilot under expected flight conditions without
requiring the pilot's 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 is unaware 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
allow initial counteraction of failures without
requiring exceptional pilot skill or strength, by
overriding the failure by moving the flight controls
in the normal sense, and by deactivating the failed
system.
(c)
It must be show that after any single failure of the
stability augmentation system or any other automatic
or power-operated system—
(1)
The rotorcraft is safely controllable when the
failure or malfunction occurs at any speed or
altitude within the approved operating limitations;
(2)
The controllability and maneuverability requirements
of this part are met within a practical operational
flight envelope (for example, speed, altitude,
normal acceleration, and rotorcraft configurations)
which is described in the Rotorcraft Flight Manual;
and
(3)
The trim and stability characteristics are not
impaired below a level needed to allow continued
safe flight and landing.
Primary flight controls are those used by the pilot
for immediate control of pitch, roll, yaw, and
vertical motion of the rotorcraft.
Each
primary flight control system must provide for safe
flight and landing and operate independently after a
malfunction, failure, or jam of any auxiliary
interconnected control.
(a)
Each control system must have stops that positively
limit the range of motion of the pilot's controls.
(b)
Each stop must be located in the system so that the
range of travel of its control is not appreciably
affected by—
(1)
Wear;
(2)
Slackness; or
(3)
Take-up adjustments.
(c)
Each stop must be able to withstand the loads
corresponding to the design conditions for the
system.
(d)
For each main rotor blade—
(1)
Stops that are appropriate to the blade design must
be provided to limit travel of the blade about its
hinge points; and
(2)
There must be means to keep the blade from hitting
the droop stops during any operation other than
starting and stopping the rotor.
If
there is a device to lock the control system with
the rotorcraft on the ground or water, there must be
means to—
(a)
Automatically disengage the lock when the pilot
operates the controls in a normal manner, or limit
the operation of the rotorcraft so as to give
unmistakable warning to the pilot before takeoff;
and
(b)
Prevent the lock from engaging in flight.
(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.
It
must be shown by operation tests that, when the
controls are operated from the pilot compartment
with the control system loaded to correspond with
loads specified for the system, the system is free
from—
(a)
Jamming;
(b)
Excessive friction; and
(c)
Excessive deflection.
(a)
Each detail of each control system must be designed
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)
Cable systems must be designed as follows:
(1)
Cables, cable fittings, turnbuckles, splices, and
pulleys must be of an acceptable kind.
(2)
The design of cable systems must prevent any
hazardous change in cable tension throughout the
range of travel under any operating conditions and
temperature variations.
(3)
No cable smaller than1/8inch diameter may be used in
any primary control system.
(4)
Pulley kinds and sizes must correspond to the cables
with which they are used. The pulley-cable
combinations and strength values specified in
MIL-HDBK-5 must be used unless they are
inapplicable.
(5)
Pulleys must have close fitting guards to prevent
the cables from being displaced or fouled.
(6)
Pulleys must lie close enough to the plane passing
through the cable to prevent the cable from rubbing
against the pulley flange.
(7)
No fairlead may cause a change in cable direction of
more than three degrees.
(8)
No clevis pin subject to load or motion and retained
only by cotter pins may be used in the control
system.
(9)
Turnbuckles attached to parts having angular motion
must be installed to prevent binding throughout the
range of travel.
(10)
There must be means for visual inspection at each
fairlead, pulley, terminal, and turnbuckle.
(e)
Control system joints subject to angular motion must
incorporate the following special factors with
respect to the ultimate bearing strength of the
softest material used as a bearing:
(1)
3.33 for push-pull systems other than ball and
roller bearing systems.
(2)
2.0 for cable systems.
(f)
For control system joints, the manufacturer's
static, non-Brinell rating of ball and roller
bearings may not be exceeded.
(a)
Each control system spring device whose failure
could cause flutter or other unsafe characteristics
must be reliable.
(b)
Compliance with paragraph (a) of this section must
be shown by tests simulating service conditions.
Each
main rotor blade pitch control mechanism must allow
rapid entry into autorotation after power failure.
(a)
If a power boost or power-operated control system is
used, an alternate system must be immediately
available that allows continued safe flight and
landing in the event of—
(1)
Any single failure in the power portion of the
system; or
(2)
The failure of all engines.
(b)
Each alternate system may be a duplicate power
portion or a manually operated mechanical system.
The power portion includes the power source (such as
hydrualic pumps), and such items as valves, lines,
and actuators.
(c)
The failure of mechanical parts (such as piston rods
and links), and the jamming of power cylinders, must
be considered unless they are extremely improbable.
The
landing inertia load factor and the reserve energy
absorption capacity of the landing gear must be
substantiated by the tests prescribed in 29.725 and
29.727, respectively. These tests must be conducted
on the complete rotorcraft or on units consisting of
wheel, tire, and shock absorber in their proper
relation.
The
limit drop test must be conducted as follows:
(a)
The drop height must be at least 8 inches.
(b)
If considered, the rotor lift specified in 29.473(a)
must be introduced into the drop test by appropriate
energy absorbing devices or by the use of an
effective mass.
(c)
Each landing gear unit must be tested in the
attitude simulating the landing condition that is
most critical from the standpoint of the energy to
be absorbed by it.
(d)
When an effective mass is used in showing compliance
with paragraph (b) of this section, the following
formulae may be used instead of more rational
computations.
where:
W e=the
effective weight to be used in the drop test (lbs.).
W=W
Mfor main gear units (lbs.), equal to the static
reaction on the particular unit with the rotorcraft
in the most critical attitude. A rational method may
be used in computing a main gear static reaction,
taking into consideration the moment arm between the
main wheel reaction and the rotorcraft center of
gravity.
W=W Nfor nose
gear units (lbs.), equal to the vertical component
of the static reaction that would exist at the nose
wheel, assuming that the mass of the rotorcraft acts
at the center of gravity and exerts a force of 1.0
g downward and 0.25 g forward.
W=W tfor
tailwheel units (lbs.) equal to whichever of the
following is critical—
(1)
The static weight on the tailwheel with the
rotorcraft resting on all wheels; or
(2)
The vertical component of the ground reaction that
would occur at the tailwheel assuming that the mass
of the rotorcraft acts at the center of gravity and
exerts a force of 1 g downward with the
rotorcraft in the maximum nose-up attitude
considered in the nose-up landing conditions.
h =specified
free drop height (inches).
L =ratio of
assumed rotor lift to the rotorcraft weight.
d =deflection
under impact of the tire (at the proper inflation
pressure) plus the vertical component of the axle
travel (inches) relative to the drop mass.
n =limit
inertia load factor.
n j=the load
factor developed, during impact, on the mass used in
the drop test (i.e., the acceleration dv/dt
in g 's recorded in the drop test plus 1.0).
The
reserve energy absorption drop test must be
conducted as follows:
(a)
The drop height must be 1.5 times that specified in
29.725(a).
(b)
Rotor lift, where considered in a manner similar to
that prescribed in 29.725(b), may not exceed 1.5
times the lift allowed under that paragraph.
(c)
The landing gear must withstand this test without
collapsing. Collapse of the landing gear occurs when
a member of the nose, tail, or main gear will not
support the rotorcraft in the proper attitude or
allows the rotorcraft structure, other than landing
gear and external accessories, to impact the landing
surface.
For
rotorcraft with retractable landing gear, the
following apply:
(a)
Loads. The landing gear, retracting
mechanism, wheel well doors, and supporting
structure must be designed for—
(1)
The loads occurring in any maneuvering condition
with the gear retracted;
(2)
The combined friction, inertia, and air loads
occurring during retraction and extension at any
airspeed up to the design maximum landing gear
operating speed; and
(3)
The flight loads, including those in yawed flight,
occurring with the gear extended at any airspeed up
to the design maximum landing gear extended speed.
(b)
Landing gear lock. A positive means must be
provided to keep the gear extended.
(c)
Emergency operation. When other than manual
power is used to operate the gear, emergency means
must be provided for extending the 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.
(d)
Operation tests. The proper functioning of
the retracting mechanism must be shown by operation
tests.
(e)
Position indicator. There must be means to
indicate to the pilot when the gear is secured in
the extreme positions.
(f)
Control. The location and operation of the
retraction control must meet the requirements of
29.777 and 29.779.
(g)
Landing gear warning. An aural or equally
effective landing gear warning device must be
provided that functions continuously when the
rotorcraft is in a normal landing mode and the
landing gear is not fully extended and locked. A
manual shutoff capability must be provided for the
warning device and the warning system must
automatically reset when the rotorcraft is no longer
in the landing mode.
(a)
Each landing gear 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)
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.
Each
landing gear wheel must have a tire—
(a)
That is a proper fit on the rim of the wheel; and
(b)
Of a rating that is not exceeded under—
(1)
The design maximum weight;
(2)
A load on each main wheel tire equal to the static
ground reaction corresponding to the critical center
of gravity; and
(3)
A load on nose wheel tires (to be compared with the
dynamic rating established for those tires) equal to
the reaction obtained at the nose wheel, assuming
that the mass of the rotorcraft acts as the most
critical center of gravity and exerts a force of 1.0
g downward and 0.25 g forward, the
reactions being distributed to the nose and main
wheels according to the principles of statics with
the drag reaction at the ground applied only at
wheels with brakes.
(c)
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
contact between the tire and any part of the
structure or systems.
For
rotorcraft with wheel-type landing gear, a braking
device must be installed that is—
(a)
Controllable by the pilot;
(b)
Usable during power-off landings; and
(c)
Adequate to—
(1)
Counteract any normal unbalanced torque when
starting or stopping the rotor; and
(2)
Hold the rotorcraft parked on a 10-degree slope on a
dry, smooth pavement.
(a)
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.
(b)
There must be a stabilizing means to maintain the
ski in an appropriate position during flight. This
means must have enough strength to withstand the
maximum aerodynamic and inertia loads on the ski.
29.751 Main float buoyancy.
(a)
For main floats, the buoyancy necessary to support
the maximum weight of the rotorcraft in fresh water
must be exceeded by—
(1)
50 percent, for single floats; and
(2)
60 percent, for multiple floats.
(b)
Each main float must have enough water-tight
compartments so that, with any single main float
compartment flooded, the mainfloats will provide a
margin of positive stability great enough to
minimize the probability of capsizing.
(a)
Bag floats. Each bag float must be designed
to withstand—
(1)
The maximum pressure differential that might be
developed at the maximum altitude for which
certification with that float is requested; and
(2)
The vertical loads prescribed in 29.521(a),
distributed along the length of the bag over
three-quarters of its projected area.
(b)
Rigid floats. Each rigid float must be able
to withstand the vertical, horizontal, and side
loads prescribed in 29.521. An appropriate load
distribution under critical conditions must be used.
Water-based and amphibian rotorcraft.
The hull and auxiliary floats, if used, must have enough watertight
compartments so that, with any single compartment of
the hull or auxiliary floats flooded, the buoyancy
of the hull and auxiliary floats, and wheel tires if
used, provides a margin of positive water stability
great enough to minimize the probability of
capsizing the rotorcraft for the worst combination
of wave heights and surface winds for which approval
is desired.
29.757 Hull and auxiliary float strength.
The
hull, and auxiliary floats if used, must withstand
the water loads prescribed by 29.519 with a rational
and conservative distribution of local and
distributed water pressures over the hull and float
bottom.
For
each pilot compartment—
(a)
The compartment and its equipment must allow each
pilot to perform his duties without unreasonable
concentration or fatigue;
(b)
If there is provision for a second pilot, the
rotorcraft must be controllable with equal safety
from either pilot position. Flight and powerplant
controls must be designed to prevent confusion or
inadvertent operation when the rotorcraft is piloted
from either position;
(c)
The vibration and noise characteristics of cockpit
appurtenances may not interfere with safe operation;
(d)
Inflight leakage of rain or snow that could distract
the crew or harm the structure must be prevented.
(a)
Non-precipitation conditions. For
non-precipitation conditions, the following apply:
(1)
Each pilot compartment must be arranged to give the
pilots a sufficiently extensive, clear, and
undistorted view for safe operation.
(2)
Each pilot compartment must be free of glare and
reflection that could interfere with the pilot's
view. If certification for night operation is
requested, this must be shown by night flight tests.
(b)
Precipitation conditions. For precipitation
conditions, the following apply:
(1)
Each pilot must have a sufficiently extensive view
for safe operation—
(i)
In heavy rain at forward speeds up to V H;
and
(ii)
In the most severe icing condition for which
certification is requested.
(2)
The first pilot must have a window that—
(i)
Is openable under the conditions prescribed in
paragraph (b)(1) of this section; and
(ii)
Provides the view prescribed in that paragraph.
Windshields and windows must be made of material
that will not break into dangerous fragments.
Cockpit controls must be—
(a)
Located to provide convenient operation and to
prevent confusion and inadvertent operation; and
(b)
Located and arranged with respect to the pilot's
seats so that there is full and unrestricted
movement of each control without interference from
the cockpit structure or the pilot's clothing when
pilots from 5′2&inch; to 6′0&inch; in height are
seated.
Cockpit controls must be designed so that they
operate in accordance with the following movements
and actuation:
(a)
Flight controls, including the collective pitch
control, must operate with a sense of motion which
corresponds to the effect on the rotorcraft.
(b)
Twist-grip engine power controls must be designed so
that, for lefthand operation, the motion of the
pilot's hand is clockwise to increase power when the
hand is viewed from the edge containing the index
finger. Other engine power controls, excluding the
collective control, must operate with a forward
motion to increase power.
(c)
Normal landing gear controls must operate downward
to extend the landing gear.
(a)
Each closed cabin must have at least one adequate
and easily accessible external door.
(b)
Each external door must be located, and appropriate
operating procedures must be established, to ensure
that persons using the door will not be endangered
by the rotors, propellers, engine intakes, and
exhausts when the operating procedures are used.
(c)
There must be means for locking crew and external
passenger doors and for preventing their opening in
flight inadvertently or as a result of mechanical
failure. It must be possible to open external doors
from inside and outside the cabin with the
rotorcraft on the ground even though persons may be
crowded against the door on the inside of the
rotorcraft. The means of opening must be simple and
obvious and so arranged and marked that it can be
readily located and operated.
(d)
There must be reasonable provisions to prevent the
jamming of any external doors in a minor crash as a
result of fuselage deformation under the following
ultimate inertial forces except for cargo or service
doors not suitable for use as an exit in an
emergency:
(1)
Upward—1.5g.
(2)
Forward—4.0g.
(3)
Sideward—2.0g.
(4)
Downward—4.0g.
(e)
There must be means for direct visual inspection of
the locking mechanism by crewmembers to determine
whether the external doors (including passenger,
crew, service, and cargo doors) are fully locked.
There must be visual means to signal to appropriate
crewmembers when normally used external doors are
closed and fully locked.
(f)
For outward opening external doors usable for
entrance or egress, there must be an auxiliary
safety latching device to prevent the door from
opening when the primary latching mechanism fails.
If the door does not meet the requirements of
paragraph (c) of this section with this device in
place, suitable operating procedures must be
established to prevent the use of the device during
takeoff and landing.
(g)
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 inertial forces specified
in paragraph (d) of this section, acting separately
relative to the surrounding structure.
(2)
The rotorcraft is in the normal ground attitude and
in each of the attitudes corresponding to collapse
of one or more legs, or primary members, as
applicable, of the landing gear.
(h)
Non-jettisonable doors used as ditching emergency
exits must have means to enable them to be secured
in the open position and remain secure for emergency
egress in sea state conditions prescribed for
ditching.
(a)
Each seat, safety belt, harness, and adjacent part
of the rotorcraft at each station designated for
occupancy during takeoff and landing must be free of
potentially injurious objects, sharp edges,
protuberances, and hard surfaces and 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 inertial
factors specified in 29.561(b) and dynamic
conditions specified in 29.562.
(b)
Each occupant must be protected from serious head
injury by a safety belt plus a shoulder harness that
will prevent the head from contacting any injurious
object, except as provided for in 29.562(c)(5). A
shoulder harness (upper torso restraint), in
combination with the safety belt, constitutes a
torso restraint system as described in TSO-C114.
(c)
Each occupant's seat must have a combined safety
belt and shoulder harness with a single-point
release. Each pilot's combined safety belt and
shoulder harness must allow each pilot when seated
with safety belt and shoulder harness fastened to
perform all functions necessary for flight
operations. There must be a means to secure belt and
harness when not in use to prevent interference with
the operation of the rotorcraft and with rapid
egress in an emergency.
(d)
If seat backs do not have a firm handhold, there
must be hand grips or rails along each aisle to let
the occupants steady themselves while using the
aisle in moderately rough air.
(e)
Each projecting object that would injure persons
seated or moving about in the rotorcraft in normal
flight must be padded.
(f)
Each seat and its supporting structure must be
designed for an occupant weight of at least 170
pounds, considering the maximum load factors,
inertial forces, and reactions between the occupant,
seat, and safety belt or harness corresponding with
the applicable flight and ground-load conditions,
including the emergency landing conditions of
29.561(b). In addition—
(1)
Each pilot seat must be designed for the reactions
resulting from the application of the pilot forces
prescribed in 29.397; and
(2)
The inertial forces prescribed in 29.561(b) must be
multiplied by a factor of 1.33 in determining the
strength of the attachment of—
(i)
Each seat to the structure; and
(ii)
Each safety belt or harness to the seat or
structure.
(g)
When the safety belt and shoulder harness are
combined, the rated strength of the safety belt and
shoulder harness may not be less than that
corresponding to the inertial forces specified in
29.561(b), considering the occupant weight of at
least 170 pounds, considering the dimensional
characteristics of the restraint system
installation, and using a distribution of at least a
60-percent load to the safety belt and at least a
40-percent load to the shoulder harness. If the
safety belt is capable of being used without the
shoulder harness, the inertial forces specified must
be met by the safety belt alone.
(h)
When a headrest is used, the headrest and its
supporting structure must be designed to resist the
inertia forces specified in 29.561, with a 1.33
fitting factor and a head weight of at least 13
pounds.
(i)
Each seating device system includes the device such
as the seat, the cushions, the occupant restraint
system and attachment devices.
(j)
Each seating device system may use design features
such as crushing or separation of certain parts of
the seat in the design to reduce occupant loads for
the emergency landing dynamic conditions of 29.562;
otherwise, the system must remain intact and must
not interfere with rapid evacuation of the
rotorcraft.
(k)
For purposes of this section, a litter is defined as
a device designed to carry a non-ambulatory person,
primarily in a recumbent position, into and on the
rotorcraft. Each berth or litter must be designed to
withstand the load reaction of an occupant weight of
at least 170 pounds when the occupant is subjected
to the forward inertial factors specified in
29.561(b). A berth or litter installed within 15° or
less of the longitudinal axis of the rotorcraft must
be provided with a padded end-board, cloth
diaphragm, or equivalent means that can withstand
the forward load reaction. A berth or litter
oriented greater than 15° with the longitudinal axis
of the rotorcraft must be equipped with appropriate
restraints, such as straps or safety belts, to
withstand the forward reaction. In addition—
(1)
The berth or litter must have a restraint system and
must not have corners or other protuberances likely
to cause serious injury to a person occupying it
during emergency landing conditions; and
(2)
The berth or litter attachment and the occupant
restraint system attachments to the structure must
be designed to withstand the critical loads
resulting from flight and ground load conditions and
from the conditions prescribed in 29.561(b). The
fitting factor required by 29.625(d) shall be
applied.
(a)
Each cargo and baggage compartment must be designed
for its placarded maximum weight of contents and for
the critical load distributions at the appropriate
maximum load factors corresponding to the specified
flight and ground load conditions, except the
emergency landing conditions of 29.561.
(b)
There must be means to prevent the contents of any
compartment from becoming a hazard by shifting under
the loads specified in paragraph (a) of this
section.
(c)
Under the emergency landing conditions of 29.561,
cargo and baggage compartments must—
(1)
Be positioned so that if the contents break loose
they are unlikely to cause injury to the occupants
or restrict any of the escape facilities provided
for use after an emergency landing; or
(2)
Have sufficient strength to withstand the conditions
specified in 29.561, including the means of
restraint and their attachments required by
paragraph (b) of this section. Sufficient strength
must be provided for the maximum authorized weight
of cargo and baggage at the critical loading
distribution.
(d)
If cargo compartment lamps are installed, each lamp
must be installed so as to prevent contact between
lamp bulb and cargo.
(a)
If certification with ditching provisions is
requested, the rotorcraft must meet the requirements
of this section and 29.807(d), 29.1411 and 29.1415.
(b)
Each practicable design measure, compatible with the
general characteristics of the rotorcraft, must be
taken to minimize the probability that in an
emergency landing on water, the behavior of the
rotorcraft would cause immediate injury to the
occupants or would make it impossible for them to
escape.
(c)
The probable behavior of the rotorcraft in a water
landing must be investigated by model tests or by
comparison with rotorcraft of similar configuration
for which the ditching characteristics are known.
Scoops, flaps, projections, and any other factors
likely to affect the hydrodynamic characteristics of
the rotorcraft must be considered.
(d)
It must be shown that, under reasonably probable
water conditions, the flotation time and trim of the
rotorcraft will allow the occupants to leave the
rotorcraft and enter the liferafts required by
29.1415. If compliance with this provision is shown
by bouyancy and trim computations, appropriate
allowances must be made for probable structural
damage and leakage. If the rotorcraft 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 bouyancy 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 rotorcraft 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.
(a)
Each crew and passenger area must have means for
rapid evacuation in a crash landing, with the
landing gear (1) extended and (2) retracted,
considering the possibility of fire.
(b)
Passenger entrance, crew, and service doors may be
considered as emergency exits if they meet the
requirements of this section and of 29.805 through
29.815.
(c)
[Reserved]
(d)
Except as provided in paragraph (e) of this section,
the following categories of rotorcraft must be
tested in accordance with the requirements of
appendix D of this part to demonstrate that the
maximum seating capacity, including the crewmembers
required by the operating rules, can be evacuated
from the rotorcraft to the ground within 90 seconds:
(1)
Rotorcraft with a seating capacity of more than 44
passengers.
(2)
Rotorcraft with all of the following:
(i)
Ten or more passengers per passenger exit as
determined under 29.807(b).
(ii)
No main aisle, as described in 29.815, for each row
of passenger seats.
(iii) Access to each passenger exit for each
passenger by virtue of design features of seats,
such as folding or break-over seat backs or folding
seats.
(e)
A combination of analysis and tests may be used to
show that the rotorcraft is capable of being
evacuated within 90 seconds under the conditions
specified in 29.803(d) if the Administrator finds
that the combination of analysis and tests will
provide data, with respect to the emergency
evacuation capability of the rotorcraft, equivalent
to that which would be obtained by actual
demonstration.
(a)
For rotorcraft with passenger emergency exits that
are not convenient to the flight crew, there must be
flight crew emergency exits, on both sides of the
rotorcraft or as a top hatch, in the flight crew
area.
(b)
Each flight crew emergency exit must be of
sufficient size and must be located so as to allow
rapid evacuation of the flight crew. This must be
shown by test.
(c)
Each exit must not be obstructed by water or
flotation devices after a ditching. This must be
shown by test, demonstration, or analysis.
(a)
Type. For the purpose of this part, the types
of passenger emergency exit are as follows:
(1)
Type I. This type must have 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, in the passenger area in the
side of the fuselage at floor level and as ACAR away
as practicable from areas that might become
potential fire hazards in a crash.
(2)
Type II. This type is the same as Type I,
except that the opening must be at least 20 inches
wide by 44 inches high.
(3)
Type III. This type is the same as Type I,
except that—
(i)
The opening must be at least 20 inches wide by 36
inches high; and
(ii)
The exits need not be at floor level.
(4)
Type IV. This type must have 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, in the side of the fuselage
with a step-up inside the rotorcraft of not more
than 29 inches.
Openings with dimensions larger than those specified
in this section may be used, regardless of shape, if
the base of the opening has a flat surface of not
less than the specified width.
(b)
Passenger emergency exits; side-of-fuselage.
Emergency exits must be accessible to the passengers
and, except as provided in paragraph (d) of this
section, must be provided in accordance with the
following table:
|
Passenger seating capacity |
Emergency exits for each
side of the fuselage |
|
Type I |
Type II |
Type III |
Type IV |
|
1 through 10 |
|
|
|
1 |
|
11 through 19 |
|
|
1 or |
2 |
|
20 through 39 |
|
1 |
|
1 |
|
40 through 59 |
1 |
|
|
1 |
|
60 through 79 |
1 |
|
1 or |
2 |
(c)
Passenger emergency exits; other than
side-of-fuselage. In addition to the
requirements of paragraph (b) of this section—
(1)
There must be enough openings in the top, bottom, or
ends of the fuselage to allow evacuation with the
rotorcraft on its side; or
(2)
The probability of the rotorcraft coming to rest on
its side in a crash landing must be extremely
remote.
(d)
Ditching emergency exits for passengers. If
certification with ditching provisions is requested,
ditching emergency exits must be provided in
accordance with the following requirements and must
be proven by test, demonstration, or analysis unless
the emergency exits required by paragraph (b) of
this section already meet these requirements.
(1)
For rotorcraft that have a passenger seating
configuration, excluding pilots seats, of nine seats
or less, one exit above the waterline in each side
of the rotorcraft, meeting at least the dimensions
of a Type IV exit.
(2)
For rotorcraft that have a passenger seating
configuration, excluding pilots seats, of 10 seats
or more, one exit above the waterline in a side of
the rotorcraft 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
rotorcraft. However, where it has been shown through
analysis, ditching demonstrations, or any other
tests found necessary by the Administrator, that the
evacuation capability of the rotorcraft during
ditching is improved by the use of larger exits, or
by other means, the passenger seat to exit ratio may
be increased.
(3)
Flotation devices, whether stowed or deployed, may
not interfere with or obstruct the exits.
(e)
Ramp exits. One Type I exit only, or one Type
II exit only, that is required in the side of the
fuselage under paragraph (b) of this section, may be
installed instead in the ramp of floor ramp
rotorcraft if—
(1)
Its installation in the side of the fuselage is
impractical; and
(2)
Its installation in the ramp meets 29.813.
(f)
Tests. The proper functioning of each
emergency exit must be shown by test.
(a)
Each emergency exit must consist of a movable door
or hatch in the external walls of the fuselage and
must provide an unobstructed opening to the outside.
(b)
Each emergency exit must be openable from the inside
and from the outside.
(c)
The means of opening each emergency exit must be
simple and obvious and may not require exceptional
effort.
(d)
There must be means for locking each emergency exit
and for preventing opening in flight inadvertently
or as a result of mechanical failure.
(e)
There must be means to minimize the probability of
the jamming of any emergency exit in a minor crash
landing as a result of fuselage deformation under
the ultimate inertial forces in 29.783(d).
(f)
Except as provided in paragraph (h) of this section,
each land-based rotorcraft emergency exit must have
an approved slide as stated in paragraph (g) of this
section, or its equivalent, to assist occupants in
descending to the ground from each floor level exit
and an approved rope, or its equivalent, for all
other exits, if the exit threshold is more that 6
feet above the ground—
(1)
With the rotorcraft on the ground and with the
landing gear extended;
(2)
With one or more legs or part of the landing gear
collapsed, broken, or not extended; and
(3)
With the rotorcraft resting on its side, if required
by 29.803(d).
(g)
The slide for each passenger emergency exit must be
a self-supporting slide or equivalent, and must be
designed to meet the following requirements:
(1)
It must be automatically deployed, and deployment
must begin during the interval between the time the
exit opening means is actuated from inside the
rotorcraft 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
slide when the exit is opened from either the inside
or the outside under non-emergency conditions for
normal use.
(2)
It must be automatically erected within 10 seconds
after deployment is begun.
(3)
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 or part of the
landing gear.
(4)
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.
(5)
Each slide installation must be qualified by five
consecutive deployment and inflation tests conducted
(per exit) without failure, and at least three tests
of each such five-test series must be conducted
using a single representative sample of the device.
The sample devices must be deployed and inflated by
the system's primary means after being subjected to
the inertia forces specified in 29.561(b). If any
part of the system fails or does not function
properly during the required tests, the cause of the
failure or malfunction must be corrected by positive
means and after that, the full series of five
consecutive deployment and inflation tests must be
conducted without failure.
(h)
For rotorcraft having 30 or fewer passenger seats
and having an exit threshold more than 6 feet above
the ground, a rope or other assist means may be used
in place of the slide specified in paragraph (f) of
this section, provided an evacuation demonstration
is accomplished as prescribed in 29.803(d) or (e).
(i)
If a rope, with its attachment, is used for
compliance with paragraph (f), (g), or (h) of this
section, it must—
(1)
Withstand a 400-pound static load; and
(2)
Attach to the fuselage structure at or above the top
of the emergency exit opening, or at another
approved location if the stowed rope would reduce
the pilot's view in flight.
(a)
Each passenger emergency exit, its means of access,
and its means of opening must be conspicuously
marked for the guidance of occupants using the exits
in daylight or in the dark. Such markings must be
designed to remain visible for rotorcraft equipped
for over water flights if the rotorcraft is capsized
and the cabin is submerged.
(b)
The identity and location of each passenger
emergency exit must be recognizable from a distance
equal to the width of the cabin.
(c)
The location of each passenger emergency exit must
be indicated by a sign visible to occupants
approaching along the main passenger aisle. There
must be a locating sign—
(1)
Next to or above the aisle near each floor emergency
exit, except that one sign may serve two exits if
both exists can be seen readily from that sign; and
(2)
On each bulkhead or divider that prevents fore and
aft vision along the passenger cabin, to indicate
emergency exits beyond and obscured by it, except
that if this is not possible the sign may be placed
at another appropriate location.
(d)
Each passenger emergency exit marking and each
locating sign must have white letters 1 inch high on
a red background 2 inches high, be self or
electrically illuminated, and have a minimum
luminescence (brightness) of at least 160
microlamberts. The colors may be reversed if this
will increase the emergency illumination of the
passenger compartment.
(e)
The location of each passenger emergency exit
operating handle and instructions for opening must
be shown—
(1)
For each emergency exit, by a marking on or near the
exit that is readable from a distance of 30 inches;
and
(2)
For each Type I or Type II emergency exit with a
locking mechanism released by rotary motion of the
handle, by—
(i)
A red arrow, with a shaft at least three-fourths
inch wide and a head twice the width of the shaft,
extending along at least 70 degrees of arc at a
radius approximately equal to three-fourths of the
handle length; and
(ii)
The word “open” in red letters 1 inch high, placed
horizontally near the head of the arrow.
(f)
Each emergency exit, and its means of opening, must
be marked on the outside of the rotorcraft. In
addition, the following apply:
(1)
There must be a 2-inch colored band outlining each
passenger emergency exit, except small rotorcraft
with a maximum weight of 12,500 pounds or less may
have a 2-inch colored band outlining each exit
release lever or device of passenger emergency exits
which are normally used doors.
(2)
Each outside marking, including the band, must have
color contrast to be readily distinguishable from
the surrounding fuselage surface. The contrast must
be such that, if the reflectance of the darker color
is 15 percent or less, the reflectance of the
lighter color must be at least 45 percent.
“Reflectance” is the ratio of the luminous flux
reflected by a body to the luminous flux it
receives. When the reflectance of the darker color
is greater than 15 percent, at least a 30 percent
difference between its reflectance and the
reflectance of the lighter color must be provided.
(g)
Exits marked as such, though in excess of the
required number of exits, must meet the requirements
for emergency exits of the particular type.
Emergency exits need only be marked with the word
“Exit.”
For
transport Category A rotorcraft, the following
apply:
(a)
A source of light with its power supply independent
of the main lighting system must be installed to—
(1)
Illuminate each passenger emergency exit marking and
locating sign; and
(2)
Provide enough general lighting in the passenger
cabin so that the average illumination, when
measured at 40-inch intervals at seat armrest height
on the center line of the main passenger aisle, is
at least 0.05 foot-candle.
(b)
Exterior emergency lighting must be provided at each
emergency exit. The illumination may not be less
than 0.05 foot-candle (measured normal to the
direction of incident light) for minimum width on
the ground surface, with landing gear extended,
equal to the width of the emergency exit where an
evacuee is likely to make first contact with the
ground outside the cabin. The exterior emergency
lighting may be provided by either interior or
exterior sources with light intensity measurements
made with the emergency exits open.
(c)
Each light required by paragraph (a) or (b) of this
section must be operable manually from the cockpit
station and from a point in the passenger
compartment that is readily accessible. The cockpit
control device must have an “on,” “off,” and “armed”
position so that when turned on at the cockpit or
passenger compartment station or when armed at the
cockpit station, the emergency lights will either
illuminate or remain illuminated upon interruption
of the rotorcraft's normal electric power.
(d)
Any means required to assist the occupants in
descending to the ground must be illuminated so that
the erected assist means is visible from the
rotorcraft.
(1)
The assist means must be provided with an
illumination of not less than 0.03 foot-candle
(measured normal to the direction of the incident
light) at the ground end of the erected assist means
where an evacuee using the established escape route
would normally make first contact with the ground,
with the rotorcraft in each of the attitudes
corresponding to the collapse of one or more legs of
the landing gear.
(2)
If the emergency lighting subsystem illuminating the
assist means is independent of the rotorcraft's main
emergency lighting system, it—
(i)
Must automatically be activated when the assist
means is erected;
(ii)
Must provide the illumination required by paragraph
(d)(1); and
(iii) May not be adversely affected by stowage.
(e)
The energy supply to each emergency lighting unit
must provide the required level of illumination for
at least 10 minutes at the critical ambient
conditions after an emergency landing.
(f)
If storage batteries are used as the energy supply
for the emergency lighting system, they may be
recharged from the rotorcraft's main electrical
power system provided the charging circuit is
designed to preclude inadvertent battery discharge
into charging circuit faults.
(a)
Each passageway between passenger compartments, and
each passageway leading to Type I and Type II
emergency exits, must be—
(1)
Unobstructed; and
(2)
At least 20 inches wide.
(b)
For each emergency exit covered by 29.809(f), there
must be enough space adjacent to that exit to allow
a crewmember to assist in the evacuation of
passengers without reducing the unobstructed width
of the passageway below that required for that exit.
(c)
There must be access from each aisle to each Type
III and Type IV exit, and
(1)
For rotorcraft that have a passenger seating
configuration, excluding pilot seats, of 20 or more,
the projected opening of the exit provided must not
be obstructed by seats, berths, or other protrusions
(including seatbacks in any position) for a distance
from that exit of not less than the width of the
narrowest passenger seat installed on the
rotorcraft;
(2)
For rotorcraft that have a passenger seating
configuration, excluding pilot seats, of 19 or less,
there may be minor obstructions in the region
described in paragraph (c)(1) of this section, if
there are compensating factors to maintain the
effectiveness of the exit.
The
main passenger aisle width between seats must equal
or exceed the values in the following table:
|
Passenger seating capacity |
Minimum main passenger aisle width |
|
Less than 25 inches from floor (inches) |
25 Inches and more from floor (inches) |
|
10 or less |
12 |
15 |
|
11 through 19 |
12 |
20 |
|
20 or more |
15 |
20 |
1A
narrower width not less than 9 inches may be
approved when substantiated by tests found necessary
by the Administrator.
(a)
Each passenger and crew compartment must be
ventilated, and each crew compartment must have
enough fresh air (but not less than 10 cu. ft. per
minute per crewmember) to let crewmembers perform
their duties without undue discomfort or fatigue.
(b)
Crew and passenger compartment air must be free from
harmful or hazardous concentrations of gases or
vapors.
(c)
The concentration of carbon monoxide may not exceed
one part in 20,000 parts of air during forward
flight. If the concentration exceeds this value
under other conditions, there must be suitable
operating restrictions.
(d)
There must be means to ensure compliance with
paragraphs (b) and (c) of this section under any
reasonably probable failure of any ventilating,
heating, or other system or equipment.
Each
combustion heater must be approved.
(a)
Hand fire extinguishers. For hand fire
extinguishers the following apply:
(1)
Each hand fire extinguisher must be approved.
(2)
The kinds and quantities of each extinguishing agent
used must be appropriate to the kinds of fires
likely to occur where that agent is used.
(3)
Each extinguisher for use in a personnel compartment
must be designed to minimize the hazard of toxic gas
concentrations.
(b)
Built-in fire extinguishers. If a built-in
fire extinguishing system is required—
(1)
The capacity of each system, in relation to the
volume of the compartment where used and the
ventilation rate, must be adequate for any fire
likely to occur in that compartment.
(2)
Each system must be installed so that—
(i)
No extinguishing agent likely to enter personnel
compartments will be present in a quantity that is
hazardous to the occupants; and
(ii)
No discharge of the extinguisher can cause
structural damage.
For
each compartment to be used by the crew or
passengers—
(a)
The materials (including finishes or decorative
surfaces applied to the materials) must meet the
following test criteria as applicable:
(1)
Interior ceiling panels, interior wall panels,
partitions, galley structure, large cabinet walls,
structural flooring, and materials used in the
construction of stowage compartments (other than
underseat stowage compartments and compartments for
stowing small items such as magazines and maps) must
be self-extinguishing when tested vertically in
accordance with the applicable portions of appendix
F of Part 25 of this chapter, or other approved
equivalent methods. The average burn length may not
exceed 6 inches and the average flame time after
removal of the flame source may not exceed 15
seconds. Drippings from the test specimen may not
continue to flame for more than an average of 3
seconds after falling.
(2)
Floor covering, textiles (including draperies and
upholstery), seat cushions, padding, decorative and
non-decorative coated fabrics, leather, trays and
galley furnishings, electrical conduit, thermal and
acoustical insulation and insulation covering, air
ducting, joint and edge covering, cargo compartment
liners, insulation blankets, cargo covers, and
transparencies, molded and thermoformed parts, air
ducting joints, and trim strips (decorative and
chafing) that are constructed of materials not
covered in paragraph (a)(3) of this section, must be
self extinguishing when tested vertically in
accordance with the applicable portion of appendix F
of Part 25 of this chapter, or other approved
equivalent methods. The average burn length may not
exceed 8 inches and the average flame time after
removal of the flame source may not exceed 15
seconds. Drippings from the test specimen may not
continue to flame for more than an average of 5
seconds after falling.
(3)
Acrylic windows and signs, parts constructed in
whole or in part of elastometric materials, edge
lighted instrument assemblies consisting of two or
more instruments in a common housing, seat belts,
shoulder harnesses, and cargo and baggage tie down
equipment, including containers, bins, pallets,
etc., used in passenger or crew compartments, may
not have an average burn rate greater than 2.5
inches per minute when tested horizontally in
accordance with the applicable portions of appendix
F of Part 25 of this chapter, or other approved
equivalent methods.
(4)
Except for electrical wire and cable insulation, and
for small parts (such as knobs, handles, rollers,
fasteners, clips, grommets, rub strips, pulleys, and
small electrical parts) that the Administrator finds
would not contribute significantly to the
propagation of a fire, materials in items not
specified in paragraphs (a)(1), (a)(2), or (a)(3) of
this section may not have a burn rate greater than 4
inches per minute when tested horizontally in
accordance with the applicable portions of appendix
F of Part 25 of this chapter, or other approved
equivalent methods.
(b)
In addition to meeting the requirements of paragraph
(a)(2), seat cushions, except those on flight
crewmember seats, must meet the test requirements of
Part II of appendix F of Part 25 of this chapter, or
equivalent.
(c)
If smoking is to be prohibited, there must be a
placard so stating, and if smoking is to be allowed—
(1)
There must be an adequate number of self-contained,
removable ashtrays; and
(2)
Where the crew compartment is separated from the
passenger compartment, there must be at least one
illuminated sign (using either letters or symbols)
notifying all passengers when smoking is prohibited.
Signs which notify when smoking is prohibited must—
(i)
When illuminated, be legible to each passenger
seated in the passenger cabin under all probable
lighting conditions; and
(ii)
Be so constructed that the crew can turn the
illumination on and off.
(d)
Each receptacle for towels, paper, or waste must be
at least fire-resistant and must have means for
containing possible fires;
(e)
There must be a hand fire extinguisher for the
flight crewmembers; and
(f)
At least the following number of hand fire
extinguishers must be conveniently located in
passenger compartments:
|
Passenger capacity |
Fire extinguishers |
|
7 through 30 |
1 |
|
31 through 60 |
2 |
|
61 or more |
3 |
(a)
Each cargo and baggage compartment must be
construced of or lined with materials in accordance
with the following:
(1)
For accessible and inaccessible compartments not
occupied by passengers or crew, the material must be
at least fire resistant.
(2)
Materials must meet the requirements in
29.853(a)(1), (a)(2), and (a)(3) for cargo or
baggage compartments in which—
(i)
The presence of a compartment fire would be easily
discovered by a crewmember while at the crewmember's
station;
(ii)
Each part of the compartment is easily accessible in
flight;
(iii) The compartment has a volume of 200 cubic feet
or less; and
(iv)
Not withstanding 29.1439(a), protective breathing
equipment is not required.
(b)
No compartment may contain any controls, wiring,
lines, equipment, or accessories whose damage or
failure would affect safe operation, unless those
items are protected so that—
(1)
They cannot be damaged by the movement of cargo in
the compartment; and
(2)
Their breakage or failure will not create a fire
hazard.
(c)
The design and sealing of inaccessible compartments
must be adequate to contain compartment fires until
a landing and safe evacuation can be made.
(d)
Each cargo and baggage compartment that is not
sealed so as to contain cargo compartment fires
completely without endangering the safety of a
rotorcraft or its occupants must be designed, or
must have a device, to ensure detection of fires or
smoke by a crewmember while at his station and to
prevent the accumulation of harmful quantities of
smoke, flame, extinguishing agents, and other
noxious gases in any crew or passenger compartment.
This must be shown in flight.
(e)
For rotorcraft used for the carriage of cargo only,
the cabin area may be considered a cargo compartment
and, in addition to paragraphs (a) through (d) of
this section, the following apply:
(1)
There must be means to shut off the ventilating
airflow to or within the compartment. Controls for
this purpose must be accessible to the flight crew
in the crew compartment.
(2)
Required crew emergency exits must be accessible
under all cargo loading conditions.
(3)
Sources of heat within each compartment must be
shielded and insulated to prevent igniting the
cargo.
(a)
Combustion heater fire zones. The following
combustion heater fire zones must be protected
against fire under the applicable provisions of
29.1181 through 29.1191, and 29.1195 through
29.1203:
(1)
The region surrounding any heater, if that region
contains any flammable fluid system components
(including the heater fuel system), that could—
(i)
Be damaged by heater malfunctioning; or
(ii)
Allow flammable fluids or vapors to reach the heater
in case of leakage.
(2)
Each part of any ventilating air passage that—
(i)
Surrounds the combustion chamber; and
(ii)
Would not contain (without damage to other
rotorcraft components) any fire that may occur
within the passage.
(b)
Ventilating air ducts. Each ventilating air
duct passing through any fire zone must be
fireproof. In addition—
(1)
Unless isolation is provided by fireproof valves or
by equally effective means, the ventilating air duct
downstream of each heater must be fireproof for a
distance great enough to ensure that any fire
originating in the heater can be contained in the
duct; and
(2)
Each part of any ventilating duct passing through
any region having a flammable fluid system must be
so constructed or isolated from that system that the
malfunctioning of any component of that system
cannot introduce flammable fluids or vapors into the
ventilating air stream.
(c)
Combustion air ducts. Each combustion air
duct must be fireproof for a distance great enough
to prevent damage from backfiring or reverse flame
propagation. In addition—
(1)
No combustion air duct may communicate with the
ventilating air stream unless flames from backfires
or reverse burning cannot enter the ventilating air
stream under any operating condition, including
reverse flow or malfunction of the heater or its
associated components; and
(2)
No combustion air duct may restrict the prompt
relief of any backfire that, if so restricted, could
cause heater failure.
(d)
Heater controls; general. There must be means
to prevent the hazardous accumulation of water or
ice on or in any heater control component, control
system tubing, or safety control.
(e)
Heater safety controls. For each combustion
heater, safety control means must be provided as
follows:
(1)
Means independent of the components provided for the
normal continuous control of air temperature,
airflow, and fuel flow must be provided, for each
heater, to automatically shut off the ignition and
fuel supply of that heater at a point remote from
that heater when any of the following occurs:
(i)
The heat exchanger temperature exceeds safe limits.
(ii)
The ventilating air temperature exceeds safe limits.
(iii) The combustion airflow becomes inadequate for
safe operation.
(iv)
The ventilating airflow becomes inadequate for safe
operation.
(2)
The means of complying with paragraph (e)(1) of this
section for any individual heater must—
(i)
Be independent of components serving any other
heater whose heat output is essential for safe
operation; and
(ii)
Keep the heater off until restarted by the crew.
(3)
There must be means to warn the crew when any heater
whose heat output is essential for safe operation
has been shut off by the automatic means prescribed
in paragraph (e)(1) of this section.
(f)
Air intakes. Each combustion and ventilating
air intake must be where no flammable fluids or
vapors can enter the heater system under any
operating condition—
(1)
During normal operation; or
(2)
As a result of the malfunction of any other
component.
(g)
Heater exhaust. Each heater exhaust system
must meet the requirements of 29.1121 and 29.1123.
In addition—
(1)
Each exhaust shroud must be sealed so that no
flammable fluids or hazardous quantities of vapors
can reach the exhaust systems through joints; and
(2)
No exhaust system may restrict the prompt relief of
any backfire that, if so restricted, could cause
heater failure.
(h)
Heater fuel systems. Each heater fuel system
must meet the powerplant fuel system requirements
affecting safe heater operation. Each heater fuel
system component in the ventilating air stream must
be protected by shrouds so that no leakage from
those components can enter the ventilating air
stream.
(i)
Drains. There must be means for safe drainage
of any fuel that might accumulate in the combustion
chamber or the heat exchanger. In addition—
(1)
Each part of any drain that operates at high
temperatures must be protected in the same manner as
heater exhausts; and
(2)
Each drain must be protected against hazardous ice
accumulation under any operating condition.
Each
part of the structure, controls, and the rotor
mechanism, and other parts essential to controlled
landing and (for category A) flight that would be
affected by powerplant fires must be isolated under
29.1191, or must be—
(a)
For category A rotorcraft, fireproof; and
(b)
For Category B rotorcraft, fireproof or protected so
that they can perform their essential functions for
at least 5 minutes under any foreseeable powerplant
fire conditions.
(a)
In each area where flammable fluids or vapors might
escape by leakage of a fluid system, there must be
means to minimize the probability of ignition of the
fluids and vapors, and the resultant hazards if
ignition does occur.
(b)
Compliance with paragraph (a) of this section must
be shown by analysis or tests, and the following
factors must be considered:
(1)
Possible sources and paths of fluid leakage, and
means of detecting leakage.
(2)
Flammability characteristics of fluids, including
effects of any combustible or absorbing materials.
(3)
Possible ignition sources, including electrical
faults, overheating of equipment, and malfunctioning
of protective devices.
(4)
Means available for controlling or extinguishing a
fire, such as stopping flow of fluids, shutting down
equipment, fireproof containment, or use of
extinguishing agents.
(5)
Ability of rotorcraft components that are critical
to safety of flight to withstand fire and heat.
(c)
If action by the flight crew is required to prevent
or counteract a fluid fire (e.g. equipment shutdown
or actuation of a fire extinguisher), quick acting
means must be provided to alert the crew.
(d)
Each area where flammable fluids or vapors might
escape by leakage of a fluid system must be
identified and defined.
(a)
It must be shown by analysis, test, or both, that
the rotorcraft external load attaching means for
rotorcraft-load combinations to be used for nonhuman
external cargo applications can withstand a limit
static load equal to 2.5, or some lower load factor
approved under 29.337 through 29.341, multiplied by
the maximum external load for which authorization is
requested. It must be shown by analysis, test, or
both that the rotorcraft external load attaching
means and corresponding personnel carrying device
system for rotorcraft-load combinations to be used
for human external cargo applications can withstand
a limit static load equal to 3.5 or some lower load
factor, not less than 2.5, approved under 29.337
through 29.341, multiplied by the maximum external
load for which authorization is requested. The load
for any rotorcraft-load combination class, for any
external cargo type, must be applied in the vertical
direction. For jettisonable external loads of any
applicable external cargo type, the load must also
be applied in any direction making the maximum angle
with the vertical that can be achieved in service
but not less than 30°. However, the 30° angle may be
reduced to a lesser angle if—
(1)
An operating limitation is established limiting
external load operations to such angles for which
compliance with this paragraph has been shown; or
(2)
It is shown that the lesser angle can not be
exceeded in service.
(b)
The external load attaching means, for jettisonable
rotorcraft-load combinations, must include a
quick-release system to enable the pilot to release
the external load quickly during flight. The
quick-release system must consist of a primary quick
release subsystem and a backup quick release
subsystem that are isolated from one another. The
quick release system, and the means by which it is
controlled, must comply with the following:
(1)
A control for the primary quick release subsystem
must be installed either on one of the pilot's
primary controls or in an equivalently accessible
location and must be designed and located so that it
may be operated by either the pilot or a crewmember
without hazardously limiting the ability to control
the rotorcraft during an emergency situation.
(2)
A control for the backup quick release subsystem,
readily accessible to either the pilot or another
crewmember, must be provided.
(3)
Both the primary and backup quick release subsystems
must—
(i)
Be reliable, durable, and function properly with all
external loads up to and including the maximum
external limit load for which authorization is
requested.
(ii)
Be protected against electromagnetic interference
(EMI) from external and internal sources and against
lightning to prevent inadvertent load release.
(A)
The minimum level of protection required for
jettisonable rotorcraft-load combinations used for
nonhuman external cargo is a radio frequency field
strength of 20 volts per meter.
(B)
The minimum level of protection required for
jettisonable rotorcraft-load combinations used for
human external cargo is a radio frequency field
strength of 200 volts per meter.
(iii) Be protected against any failure that could be
induced by a failure mode of any other electrical or
mechanical rotorcraft system.
(c)
For rotorcraft-load combinations to be used for
human external cargo applications, the rotorcraft
must—
(1)
For jettisonable external loads, have a
quick-release system that meets the requirements of
paragraph (b) of this section and that—
(i)
Provides a dual actuation device for the primary
quick release subsystem, and
(ii)
Provides a separate dual actuation device for the
backup quick release subsystem;
(2)
Have a reliable, approved personnel carrying device
system that has the structural capability and
personnel safety features essential for external
occupant safety;
(3)
Have placards and markings at all appropriate
locations that clearly state the essential system
operating instructions and, for the personnel
carrying device system, ingress and egress
instructions;
(4)
Have equipment to allow direct intercommunication
among required crewmembers and external occupants;
(5)
Have the appropriate limitations and procedures
incorporated in the flight manual for conducting
human external cargo operations; and
(6)
For human external cargo applications requiring use
of Category A rotorcraft, have
one-engine-inoperative hover performance data and
procedures in the flight manual for the weights,
altitudes, and temperatures for which external load
approval is requested.
(d)
The critically configured jettisonable external
loads must be shown by a combination of analysis,
ground tests, and flight tests to be both
transportable and releasable throughout the approved
operational envelope without hazard to the
rotorcraft during normal flight conditions. In
addition, these external loads—must be shown to be
releasable without hazard to the rotorcraft during
emergency flight conditions.
(e)
A placard or marking must be installed next to the
external-load attaching means clearly stating any
operational limitations and the maximum authorized
external load as demonstrated under 29.25 and this
section.
(f)
The fatigue evaluation of 29.571 of this part does
not apply to rotorcraft-load combinations to be used
for nonhuman external cargo except for the failure
of critical structural elements that would result in
a hazard to the rotorcraft. For rotorcraft-load
combinations to be used for human external cargo,
the fatigue evaluation of 29.571 of this part
applies to the entire quick release and personnel
carrying device structural systems and their
attachments.
There must be reference marks for leveling the
rotorcraft on the ground.
Ballast provisions must be designed and constructed
to prevent inadvertent shifting of ballast in
flight.
(a)
For the purpose of this part, the powerplant
installation includes each part of the rotorcraft
(other than the main and auxiliary rotor structures)
that—
(1)
Is necessary for propulsion;
(2)
Affects the control of the major propulsive units;
or
(3)
Affects the safety of the major propulsive units
between normal inspections or overhauls.
(b)
For each powerplant installation—
(1)
The installation must comply with—
(i)
The installation instructions provided under 33.5 of
this chapter; and
(ii)
The applicable provisions of this subpart.
(2)
Each component of the installation must be
constructed, arranged, and installed to ensure its
continued safe operation between normal inspections
or overhauls for the range of temperature and
altitude for which approval is requested.
(3)
Accessibility must be provided to allow any
inspection and maintenance necessary for continued
airworthiness; and
(4)
Electrical interconnections must be provided to
prevent differences of potential between major
components of the installation and the rest of the
rotorcraft.
(5)
Axial and radial expansion of turbine engines may
not affect the safety of the installation.
(6)
Design precautions must be taken to minimize the
possibility of incorrect assembly of components and
equipment essential to safe operation of the
rotorcraft, except where operation with the
incorrect assembly can be shown to be extremely
improbable.
(c)
For each powerplant and auxiliary power unit
installation, it must be established that no single
failure or malfunction or probable combination of
failures will jeopardize the safe operation of the
rotorcraft except that the failure of structural
elements need not be considered if the probability
of any such failure is extremely remote.
(d)
Each auxiliary power unit installation must meet the
applicable provisions of this subpart.
(a)
Engine type certification. Each engine must
have an approved type certificate. Reciprocating
engines for use in helicopters must be qualified in
accordance with 33.49(d) of this chapter or be
otherwise approved for the intended usage.
(b)
Category A; engine isolation. For each
category A rotorcraft, the powerplants must be
arranged and isolated from each other to allow
operation, in at least one configuration, so that
the failure or malfunction of any engine, or the
failure of any system that can affect any engine,
will not—
(1)
Prevent the continued safe operation of the
remaining engines; or
(2)
Require immediate action, other than normal pilot
action with primary flight controls, by any
crewmember to maintain safe operation.
(c)
Category A; control of engine rotation. For
each Category A rotorcraft, there must be a means
for stopping the rotation of any engine individually
in flight, except that, for turbine engine
installations, the means for stopping the engine
need be provided only where necessary for safety. In
addition—
(1)
Each component of the engine stopping system that is
located on the engine side of the firewall, and that
might be exposed to fire, must be at least fire
resistant; or
(2)
Duplicate means must be available for stopping the
engine and the controls must be where all are not
likely to be damaged at the same time in case of
fire.
(d)
Turbine engine installation. For turbine
engine installations—
(1)
Design precautions must be taken to minimize the
hazards to the rotorcraft in the event of an engine
rotor failure; and
(2)
The powerplant systems associated with engine
control devices, systems, and instrumentation must
be designed to give reasonable assurance that those
engine operating limitations that adversely affect
engine rotor structural integrity will not be
exceeded in service.
(e)
Restart capability. (1) A means to restart
any engine in flight must be provided.
(2)
Except for the in-flight shutdown of all engines,
engine restart capability must be demonstrated
throughout a flight envelope for the rotorcraft.
(3)
Following the in-flight shutdown of all engines,
in-flight engine restart capability must be
provided.
(a)
Each engine must be installed to prevent the harmful
vibration of any part of the engine or rotorcraft.
(b)
The addition of the rotor and the rotor drive system
to the engine may not subject the principal rotating
parts of the engine to excessive vibration stresses.
This must be shown by a vibration investigation.
For
cooling fans that are a part of a powerplant
installation the following apply:
(a)
Category A. For cooling fans installed in
Category A rotorcraft, it must be shown that a fan
blade failure will not prevent continued safe flight
either because of damage caused by the failed blade
or loss of cooling air.
(b)
Category B. For cooling fans installed in
category B rotorcraft, there must be means to
protect the rotorcraft and allow a safe landing if a
fan blade fails. It must be shown that—
(1)
The fan blade would be contained in the case of a
failure;
(2)
Each fan is located so that a fan blade failure will
not jeopardize safety; or
(3)
Each fan blade can withstand an ultimate load of 1.5
times the centrifugal force expected in service,
limited by either—
(i)
The highest rotational speeds achievable under
uncontrolled conditions; or
(ii)
An overspeed limiting device.
(c)
Fatigue evaluation. Unless a fatigue
evaluation under 29.571 is conducted, it must be
shown that cooling fan blades are not operating at
resonant conditions within the operating limits of
the rotorcraft.
(a)
General. The rotor drive system includes any
part necessary to transmit power from the engines to
the rotor hubs. This includes gear boxes, shafting,
universal joints, couplings, rotor brake assemblies,
clutches, supporting bearings for shafting, any
attendant accessory pads or drives, and any cooling
fans that are a part of, attached to, or mounted on
the rotor drive system.
(b)
Design assessment. A design assessment must
be performed to ensure that the rotor drive system
functions safely over the full range of conditions
for which certification is sought. The design
assessment must include a detailed failure analysis
to identify all failures that will prevent continued
safe flight or safe landing and must identify the
means to minimize the likelihood of their
occurrence.
(c)
Arrangement. Rotor drive systems must be
arranged as follows:
(1)
Each rotor drive system of multiengine rotorcraft
must be arranged so that each rotor necessary for
operation and control will continue to be driven by
the remaining engines if any engine fails.
(2)
For single-engine rotorcraft, each rotor drive
system must be so arranged that each rotor necessary
for control in autorotation will continue to be
driven by the main rotors after disengagement of the
engine from the main and auxiliary rotors.
(3)
Each rotor drive system must incorporate a unit for
each engine to automatically disengage that engine
from the main and auxiliary rotors if that engine
fails.
(4)
If a torque limiting device is used in the rotor
drive system, it must be located so as to allow
continued control of the rotorcraft when the device
is operating.
(5)
If the rotors must be phased for intermeshing, each
system must provide constant and positive phase
relationship under any operating condition.
(6)
If a rotor dephasing device is incorporated, there
must be means to keep the rotors locked in proper
phase before operation.
If
there is a means to control the rotation of the
rotor drive system independently of the engine, any
limitations on the use of that means must be
specified, and the control for that means must be
guarded to prevent inadvertent operation.
(a)
Endurance tests, general. Each rotor drive
system and rotor control mechanism must be tested,
as prescribed in paragraphs (b) through (n) and (p)
of this section, for at least 200 hours plus the
time required to meet the requirements of paragraphs
(b)(2), (b)(3), and (k) of this section. These tests
must be conducted as follows:
(1)
Ten-hour test cycles must be used, except that the
test cycle must be extended to include the OEI test
of paragraphs (b)(2) and (k), of this section if OEI
ratings are requested.
(2)
The tests must be conducted on the rotorcraft.
(3)
The test torque and rotational speed must be—
(i)
Determined by the powerplant limitations; and
(ii)
Absorbed by the rotors to be approved for the
rotorcraft.
(b)
Endurance tests; takeoff run. The takeoff run
must be conducted as follows:
(1)
Except as prescribed in paragraphs (b)(2) and (b)(3)
of this section, the takeoff torque run must consist
of 1 hour of alternate runs of 5 minutes at takeoff
torque and the maximum speed for use with takeoff
torque, and 5 minutes at as low an engine idle speed
as practicable. The engine must be declutched from
the rotor drive system, and the rotor brake, if
furnished and so intended, must be applied during
the first minute of the idle run. During the
remaining 4 minutes of the idle run, the clutch must
be engaged so that the engine drives the rotors at
the minimum practical r.p.m. The engine and the
rotor drive system must be accelerated at the
maximum rate. When declutching the engine, it must
be decelerated rapidly enough to allow the operation
of the overrunning clutch.
(2)
For helicopters for which the use of a 21/2-minute
OEI rating is requested, the takeoff run must be
conducted as prescribed in paragraph (b)(1) of this
section, except for the third and sixth runs for
which the takeoff torque and the maximum speed for
use with takeoff torque are prescribed in that
paragraph. For these runs, the following apply:
(i)
Each run must consist of at least one period of
21/2minutes with takeoff torque and the maximum
speed for use with takeoff torque on all engines.
(ii)
Each run must consist of at least one period, for
each engine in sequence, during which that engine
simulates a power failure and the remaining engines
are run at the 21/2-minute OEI torque and the
maximum speed for use with 21/2-minute OEI torque
for 21/2minutes.
(3)
For multiengine, turbine-powered rotorcraft for
which the use of 30-second/2-minute OEI power is
requested, the takeoff run must be conducted as
prescribed in paragraph (b)(1) of this section
except for the following:
(i)
Immediately following any one 5-minute power-on run
required by paragraph (b)(1) of this section,
simulate a failure for each power source in turn,
and apply the maximum torque and the maximum speed
for use with 30-second OEI power to the remaining
affected drive system power inputs for not less than
30 seconds. Each application of 30-second OEI power
must be followed by two applications of the maximum
torque and the maximum speed for use with the 2
minute OEI power for not less than 2 minutes each;
the second application must follow a period at
stabilized continuous or 30 minute OEI power
(whichever is requested by the applicant). At least
one run sequence must be conducted from a simulated
“flight idle” condition. When conducted on a bench
test, the test sequence must be conducted following
stabilization at take-off power.
(ii)
For the purpose of this paragraph, an affected power
input includes all parts of the rotor drive system
which can be adversely affected by the application
of higher or asymmetric torque and speed prescribed
by the test.
(iii) This test may be conducted on a representative
bench test facility when engine limitations either
preclude repeated use of this power or would result
in premature engine removals during the test. The
loads, the vibration frequency, and the methods of
application to the affected rotor drive system
components must be representative of rotorcraft
conditions. Test components must be those used to
show compliance with the remainder of this section.
(c)
Endurance tests; maximum continuous run.
Three hours of continuous operation at maximum
continuous torque and the maximum speed for use with
maximum continuous torque must be conducted as
follows:
(1)
The main rotor controls must be operated at a
minimum of 15 times each hour through the main rotor
pitch positions of maximum vertical thrust, maximum
forward thrust component, maximum aft thrust
component, maximum left thrust component, and
maximum right thrust component, except that the
control movements need not produce loads or blade
flapping motion exceeding the maximum loads of
motions encountered in flight.
(2)
The directional controls must be operated at a
minimum of 15 times each hour through the control
extremes of maximum right turning torque, neutral
torque as required by the power applied to the main
rotor, and maximum left turning torque.
(3)
Each maximum control position must be held for at
least 10 seconds, and the rate of change of control
position must be at least as rapid as that for
normal operation.
(d)
Endurance tests; 90 percent of maximum continuous
run. One hour of continuous operation at 90
percent of maximum continuous torque and the maximum
speed for use with 90 percent of maximum continuous
torque must be conducted.
(e)
Endurance tests; 80 percent of maximum continuous
run. One hour of continuous operation at 80
percent of maximum continuous torque and the minimum
speed for use with 80 percent of maximum continuous
torque must be conducted.
(f)
Endurance tests; 60 percent of maximum continuous
run. Two hours or, for helicopters for which the
use of either 30-minute OEI power or continuous OEI
power is requested, 1 hour of continuous operation
at 60 percent of maximum continuous torque and the
minimum speed for use with 60 percent of maximum
continuous torque must be conducted.
(g)
Endurance tests; engine malfunctioning run.
It must be determined whether malfunctioning of
components, such as the engine fuel or ignition
systems, or whether unequal engine power can cause
dynamic conditions detrimental to the drive system.
If so, a suitable number of hours of operation must
be accomplished under those conditions, 1 hour of
which must be included in each cycle, and the
remaining hours of which must be accomplished at the
end of the 20 cycles. If no detrimental condition
results, an additional hour of operation in
compliance with paragraph (b) of this section must
be conducted in accordance with the run schedule of
paragraph (b)(1) of this section without
consideration of paragraph (b)(2) of this section.
(h)
Endurance tests; overspeed run. One hour of
continuous operation must be conducted at maximum
continuous torque and the maximum power-on overspeed
expected in service, assuming that speed and torque
limiting devices, if any, function properly.
(i)
Endurance tests; rotor control positions.
When the rotor controls are not being cycled during
the tie-down tests, the rotor must be operated,
using the procedures prescribed in paragraph (c) of
this section, to produce each of the maximum thrust
positions for the following percentages of test time
(except that the control positions need not produce
loads or blade flapping motion exceeding the maximum
loads or motions encountered in flight):
(1)
For full vertical thrust, 20 percent.
(2)
For the forward thrust component, 50 percent.
(3)
For the right thrust component, 10 percent.
(4)
For the left thrust component, 10 percent.
(5)
For the aft thrust component, 10 percent.
(j)
Endurance tests, clutch and brake engagements.
A total of at least 400 clutch and brake
engagements, including the engagements of paragraph
(b) of this section, must be made during the takeoff
torque runs and, if necessary, at each change of
torque and speed throughout the test. In each clutch
engagement, the shaft on the driven side of the
clutch must be accelerated from rest. The clutch
engagements must be accomplished at the speed and by
the method prescribed by the applicant. During
deceleration after each clutch engagement, the
engines must be stopped rapidly enough to allow the
engines to be automatically disengaged from the
rotors and rotor drives. If a rotor brake is
installed for stopping the rotor, the clutch, during
brake engagements, must be disengaged above 40
percent of maximum continuous rotor speed and the
rotors allowed to decelerate to 40 percent of
maximum continuous rotor speed, at which time the
rotor brake must be applied. If the clutch design
does not allow stopping the rotors with the engine
running, or if no clutch is provided, the engine
must be stopped before each application of the rotor
brake, and then immediately be started after the
rotors stop.
(k)
Endurance tests; OEI power run —(1)
30-minute OEI power run. For rotorcraft for
which the use of 30-minute OEI power is requested, a
run at 30-minute OEI torque and the maximum speed
for use with 30-minute OEI torque must be conducted
as follows: For each engine, in sequence, that
engine must be inoperative and the remaining engines
must be run for a 30-minute period.
(2)
Continuous OEI power run. For rotorcraft for
which the use of continuous OEI power is requested,
a run at continuous OEI torque and the maximum speed
for use with continuous OEI torque must be conducted
as follows: For each engine, in sequence, that
engine must be inoperative and the remaining engines
must be run for 1 hour.
(3)
The number of periods prescribed in paragraph (k)(1)
or (k)(2) of this section may not be less than the
number of engines, nor may it be less than two.
(l)
[Reserved]
(m)
Any components that are affected by maneuvering and
gust loads must be investigated for the same flight
conditions as are the main rotors, and their service
lives must be determined by fatigue tests or by
other acceptable methods. In addition, a level of
safety equal to that of the main rotors must be
provided for—
(1)
Each component in the rotor drive system whose
failure would cause an uncontrolled landing;
(2)
Each component essential to the phasing of rotors on
multirotor rotorcraft, or that furnishes a driving
link for the essential control of rotors in
autorotation; and
(3)
Each component common to two or more engines on
multiengine rotorcraft.
(n)
Special tests. Each rotor drive system
designed to operate at two or more gear ratios must
be subjected to special testing for durations
necessary to substantiate the safety of the rotor
drive system.
(o)
Each part tested as prescribed in this section must
be in a serviceable condition at the end of the
tests. No intervening disassembly which might affect
test results may be conducted.
(p)
Endurance tests; operating lubricants. To be
approved for use in rotor drive and control systems,
lubricants must meet the specifications of
lubricants used during the tests prescribed by this
section. Additional or alternate lubricants may be
qualified by equivalent testing or by comparative
analysis of lubricant specifications and rotor drive
and control system characteristics. In addition—
(1)
At least three 10-hour cycles required by this
section must be conducted with transmission and
gearbox lubricant temperatures, at the location
prescribed for measurement, not lower than the
maximum operating temperature for which approval is
requested;
(2)
For pressure lubricated systems, at least three
10-hour cycles required by this section must be
conducted with the lubricant pressure, at the
location prescribed for measurement, not higher than
the minimum operating pressure for which approval is
requested; and
(3)
The test conditions of paragraphs (p)(1) and (p)(2)
of this section must be applied simultaneously and
must be extended to include operation at any
one-engine-inoperative rating for which approval is
requested.
(a)
Any additional dynamic, endurance, and operational
tests, and vibratory investigations necessary to
determine that the rotor drive mechanism is safe,
must be performed.
(b)
If turbine engine torque output to the transmission
can exceed the highest engine or transmission torque
limit, and that output is not directly controlled by
the pilot under normal operating conditions (such as
where the primary engine power control is
accomplished through the flight control), the
following test must be made:
(1)
Under conditions associated with all engines
operating, make 200 applications, for 10 seconds
each, of torque that is at least equal to the lesser
of—
(i)
The maximum torque used in meeting 29.923 plus 10
percent; or
(ii)
The maximum torque attainable under probable
operating conditions, assuming that torque limiting
devices, if any, function properly.
(2)
For multiengine rotorcraft under conditions
associated with each engine, in turn, becoming
inoperative, apply to the remaining transmission
torque inputs the maximum torque attainable under
probable operating conditions, assuming that torque
limiting devices, if any, function properly. Each
transmission input must be tested at this maximum
torque for at least fifteen minutes.
(c)
Lubrication system failure. For lubrication
systems required for proper operation of rotor drive
systems, the following apply:
(1)
Category A. Unless such failures are
extremely remote, it must be shown by test that any
failure which results in loss of lubricant in any
normal use lubrication system will not prevent
continued safe operation, although not necessarily
without damage, at a torque and rotational speed
prescribed by the applicant for continued flight,
for at least 30 minutes after perception by the
flight crew of the lubrication system failure or
loss of lubricant.
(2)
Category B. The requirements of Category A
apply except that the rotor drive system need only
be capable of operating under autorotative
conditions for at least 15 minutes.
(d)
Overspeed test. The rotor drive system must
be subjected to 50 overspeed runs, each 30 ±3
seconds in duration, at not less than either the
higher of the rotational speed to be expected from
an engine control device failure or 105 percent of
the maximum rotational speed, including transients,
to be expected in service. If speed and torque
limiting devices are installed, are independent of
the normal engine control, and are shown to be
reliable, their rotational speed limits need not be
exceeded. These runs must be conducted as follows:
(1)
Overspeed runs must be alternated with stabilizing
runs of from 1 to 5 minutes duration each at 60 to
80 percent of maximum continuous speed.
(2)
Acceleration and deceleration must be accomplished
in a period not longer than 10 seconds (except where
maximum engine acceleration rate will require more
than 10 seconds), and the time for changing speeds
may not be deducted from the specified time for the
overspeed runs.
(3)
Overspeed runs must be made with the rotors in the
flattest pitch for smooth operation.
(e)
The tests prescribed in paragraphs (b) and (d) of
this section must be conducted on the rotorcraft and
the torque must be absorbed by the rotors to be
installed, except that other ground or flight test
facilities with other appropriate methods of torque
absorption may be used if the conditions of support
and vibration closely simulate the conditions that
would exist during a test on the rotorcraft.
(f)
Each test prescribed by this section must be
conducted without intervening disassembly and,
except for the lubrication system failure test
required by paragraph (c) of this section, each part
tested must be in a serviceable condition at the
conclusion of the test.
(a)
The critical speeds of any shafting must be
determined by demonstration except that analytical
methods may be used if reliable methods of analysis
are available for the particular design.
(b)
If any critical speed lies within, or close to, the
operating ranges for idling, power-on, and
autorotative conditions, the stresses occurring at
that speed must be within safe limits. This must be
shown by tests.
(c)
If analytical methods are used and show that no
critical speed lies within the permissible operating
ranges, the margins between the calculated critical
speeds and the limits of the allowable operating
ranges must be adequate to allow for possible
variations between the computed and actual values.
Each
universal joint, slip joint, and other shafting
joints whose lubrication is necessary for operation
must have provision for lubrication.
(a)
Turbine engine operating characteristics must be
investigated in flight to determine that no adverse
characteristics (such asstall, surge, of flameout)
are present, to a hazardous degree, during normal
and emergency operation within the range of
operating limitations of the rotorcraft and of the
engine.
(b)
The turbine engine air inlet system may not, as a
result of airflow distortion during normal
operation, cause vibration harmful to the engine.
(c)
For governor-controlled engines, it must be shown
that there exists no hazardous torsional instability
of the drive system associated with critical
combinations of power, rotational speed, and control
displacement.
Fuel System
(a)
Each fuel system must be constructed and arranged to
ensure a flow of fuel at a rate and pressure
established for proper engine and auxiliary power
unit functioning under any likely operating
conditions, including the maneuvers for which
certification is requested and during which the
engine or auxiliary power unit is permitted to be in
operation.
(b)
Each fuel system must be arranged so that—
(1)
No engine or fuel pump can draw fuel from more than
one tank at a time; or
(2)
There are means to prevent introducing air into the
system.
(c)
Each fuel system for a turbine engine must be
capable of sustained operation throughout its flow
and pressure range with fuel initially saturated
with water at 80 degrees F. and having 0.75cc of
free water per gallon added and cooled to the most
critical condition for icing likely to be
encountered in operation.
Unless other means acceptable to the Administrator
are employed to minimize the hazard of fuel fires to
occupants following an otherwise survivable impact
(crash landing), the fuel systems must incorporate
the design features of this section. These systems
must be shown to be capable of sustaining the static
and dynamic deceleration loads of this section,
considered as ultimate loads acting alone, measured
at the system component's center of gravity without
structural damage to the system components, fuel
tanks, or their attachments that would leak fuel to
an ignition source.
(a)
Drop test requirements. Each tank, or the
most critical tank, must be drop-tested as follows:
(1)
The drop height must be at least 50 feet.
(2)
The drop impact surface must be non-deforming.
(3)
The tanks must be filled with water to 80 percent of
the normal, full capacity.
(4)
The tank must be enclosed in a surrounding structure
representative of the installation unless it can be
established that the surrounding structure is free
of projections or other design features likely to
contribute to upture of the tank.
(5)
The tank must drop freely and impact in a horizontal
position ±10°.
(6)
After the drop test, there must be no leakage.
(b)
Fuel tank load factors. Except for fuel tanks
located so that tank rupture with fuel release to
either significant ignition sources, such as
engines, heaters, and auxiliary power units, or
occupants is extremely remote, each fuel tank must
be designed and installed to retain its contents
under the following ultimate inertial load factors,
acting alone.
(1)
For fuel tanks in the cabin:
(i)
Upward—4g.
(ii)
Forward—16g.
(iii) Sideward—8g.
(iv)
Downward—20g.
(2)
For fuel tanks located above or behind the crew or
passenger compartment that, if loosened, could
injure an occupant in an emergency landing:
(i)
Upward—1.5g.
(ii)
Forward—8g.
(iii) Sideward—2g.
(iv)
Downward—4g.
(3)
For fuel tanks in other areas:
(i)
Upward—1.5g.
(ii)
Forward—4g.
(iii) Sideward—2g.
(iv)
Downward—4g.
(c)
Fuel line self-sealing breakaway couplings.
Self-sealing breakaway couplings must be installed
unless hazardous relative motion of fuel system
components to each other or to local rotorcraft
structure is demonstrated to be extremely improbable
or unless other means are provided. The couplings or
equivalent devices must be installed at all fuel
tank-to-fuel line connections, tank-to-tank
interconnects, and at other points in the fuel
system where local structural deformation could lead
to the release of fuel.
(1)
The design and construction of self-sealing
breakaway couplings must incorporate the following
design features:
(i)
The load necessary to separate a breakaway coupling
must be between 25 to 50 percent of the minimum
ultimate failure load (ultimate strength) of the
weakest component in the fluid-carrying line. The
separation load must in no case be less than 300
pounds, regardless of the size of the fluid line.
(ii)
A breakaway coupling must separate whenever its
ultimate load (as defined in paragraph (c)(1)(i) of
this section) is applied in the failure modes most
likely to occur.
(iii) All breakaway couplings must incorporate
design provisions to visually ascertain that the
coupling is locked together (leak-free) and is open
during normal installation and service.
(iv)
All breakaway couplings must incorporate design
provisions to prevent uncoupling or unintended
closing due to operational shocks, vibrations, or
accelerations.
(v)
No breakaway coupling design may allow the release
of fuel once the coupling has performed its intended
function.
(2)
All individual breakaway couplings, coupling fuel
feed systems, or equivalent means must be designed,
tested, installed, and maintained so inadvertent
fuel shutoff in flight is improbable in accordance
with 29.955(a) and must comply with the fatigue
evaluation requirements of 29.571 without leaking.
(3)
Alternate, equivalent means to the use of breakaway
couplings must not create a survivable
impact-induced load on the fuel line to which it is
installed greater than 25 to 50 percent of the
ultimate load (strength) of the weakest component in
the line and must comply with the fatigue
requirements of 29.571 without leaking.
(d)
Frangible or deformable structural attachments.
Unless hazardous relative motion of fuel tanks
and fuel system components to local rotorcraft
structure is demonstrated to be extremely improbable
in an otherwise survivable impact, frangible or
locally deformable attachments of fuel tanks and
fuel system components to local rotorcraft structure
must be used. The attachment of fuel tanks and fuel
system components to local rotorcraft structure,
whether frangible or locally deformable, must be
designed such that its separation or relative local
deformation will occur without rupture or local
tear-out of the fuel tank or fuel system component
that will cause fuel leakage. The ultimate strength
of frangible or deformable attachments must be as
follows:
(1)
The load required to separate a frangible attachment
from its support structure, or deform a locally
deformable attachment relative to its support
structure, must be between 25 and 50 percent of the
minimum ultimate load (ultimate strength) of the
weakest component in the attached system. In no case
may the load be less than 300 pounds.
(2)
A frangible or locally deformable attachment must
separate or locally deform as intended whenever its
ultimate load (as defined in paragraph (d)(1) of
this section) is applied in the modes most likely to
occur.
(3)
All frangible or locally deformable attachments must
comply with the fatigue requirements of 29.571.
(e)
Separation of fuel and ignition sources. To
provide maximum crash resistance, fuel must be
located as ACAR as practicable from all occupiable
areas and from all potential ignition sources.
(f)
Other basic mechanical design criteria. Fuel
tanks, fuel lines, electrical wires, and electrical
devices must be designed, constructed, and
installed, as ACAR as practicable, to be crash
resistant.
(g)
Rigid or semirigid fuel tanks. Rigid or
semirigid fuel tank or bladder walls must be impact
and tear resistant.
(a)
For category A rotorcraft—
(1)
The fuel system must meet the requirements of
29.903(b); and
(2)
Unless other provisions are made to meet paragraph
(a)(1) of this section, the fuel system must allow
fuel to be supplied to each engine through a system
independent of those parts of each system supplying
fuel to other engines.
(b)
Each fuel system for a multiengine category B
rotorcraft must meet the requirements of paragraph
(a)(2) of this section. However, separate fuel tanks
need not be provided for each engine.
The
fuel system must be designed and arranged to prevent
the ignition of fuel vapor within the system by—
(a)
Direct lightning strikes to areas having a high
probability of stroke attachment;
(b)
Swept lightning strokes to areas where swept strokes
are highly probable; and
(c)
Corona and streamering at fuel vent outlets.
(a)
General. The fuel system for each engine must
provide the engine with at least 100 percent of the
fuel required under all operating and maneuvering
conditions to be approved for the rotorcraft,
including, as applicable, the fuel required to
operate the engines under the test conditions
required by 29.927. Unless equivalent methods are
used, compliance must be shown by test during which
the following provisions are met, except that
combinations of conditions which are shown to be
improbable need not be considered.
(1)
The fuel pressure, corrected for accelerations (load
factors), must be within the limits specified by the
engine type certificate data sheet.
(2)
The fuel level in the tank may not exceed that
established as the unusable fuel supply for that
tank under 29.959, plus that necessary to conduct
the test.
(3)
The fuel head between the tank and the engine must
be critical with respect to rotorcraft flight
attitudes.
(4)
The fuel flow transmitter, if installed, and the
critical fuel pump (for pump-fed systems) must be
installed to produce (by actual or simulated
failure) the critical restriction to fuel flow to be
expected from component failure.
(5)
Critical values of engine rotational speed,
electrical power, or other sources of fuel pump
motive power must be applied.
(6)
Critical values of fuel properties which adversely
affect fuel flow are applied during demonstrations
of fuel flow capability.
(7)
The fuel filter required by 29.997 is blocked to the
degree necessary to simulate the accumulation of
fuel contamination required to activate the
indicator required by 29.1305(a)(17).
(b)
Fuel transfer system. If normal operation of
the fuel system requires fuel to be transferred to
another tank, the transfer must occur automatically
via a system which has been shown to maintain the
fuel level in the receiving tank within acceptable
limits during flight or surface operation of the
rotorcraft.
(c)
Multiple fuel tanks. If an engine can be
supplied with fuel from more than one tank, the fuel
system, in addition to having appropriate manual
switching capability, must be designed to prevent
interruption of fuel flow to that engine, without
attention by the flight crew, when any tank
supplying fuel to that engine is depleted of usable
fuel during normal operation and any other tank that
normally supplies fuel to that engine alone contains
usable fuel.
(a)
Where tank outlets are interconnected and allow fuel
to flow between them due to gravity or flight
accelerations, it must be impossible for fuel to
flow between tanks in quantities great enough to
cause overflow from the tank vent in any sustained
flight condition.
(b)
If fuel can be pumped from one tank to another in
flight—
(1)
The design of the vents and the fuel transfer system
must prevent structural damage to tanks from
overfilling; and
(2)
There must be means to warn the crew before overflow
through the vents occurs.
29.959 Unusable
fuel supply.
The
unusable fuel supply for each tank must be
established as not less than the quantity at which
the first evidence of malfunction occurs under the
most adverse fuel feed condition occurring under any
intended operations and flight maneuvers involving
that tank.
Each
suction lift fuel system and other fuel systems
conducive to vapor formation must be shown to
operate satisfactorily (within certification limits)
when using fuel at the most critical temperature for
vapor formation under critical operating conditions
including, if applicable, the engine operating
conditions defined by 29.927(b)(1) and (b)(2).
(a)
Each fuel tank must be able to withstand, without
failure, the vibration, inertia, fluid, and
structural loads to which it may be subjected in
operation.
(b)
Each flexible fuel tank bladder or liner must be
approved or shown to be suitable for the particular
application and must be puncture resistant. Puncture
resistance must be shown by meeting the TSO-C80,
paragraph 16.0, requirements using a minimum
puncture force of 370 pounds.
(c)
Each integral fuel tank must have facilities for
inspection and repair of its interior.
(d)
The maximum exposed surface temperature of all
components in the fuel tank must be less by a safe
margin than the lowest expected auto ignition
temperature of the fuel or fuel vapor in the tank.
Compliance with this requirement must be shown under
all operating conditions and under all normal or
malfunction conditions of all components inside the
tank.
(e)
Each fuel tank installed in personnel compartments
must be isolated by fume-proof and fuel-proof
enclosures that are drained and vented to the
exterior of the rotorcraft. The design and
construction of the enclosures must provide
necessary protection for the tank, must be crash
resistant during a survivable impact in accordance
with 29.952, and must be adequate to withstand loads
and abrasions to be expected in personnel
compartments.
(a)
Each fuel tank must be able to withstand the
applicable pressure tests in this section without
failure or leakage. If practicable, test pressures
may be applied in a manner simulating the pressure
distribution in service.
(b)
Each conventional metal tank, each nonmetallic tank
with walls that are not supported by the rotorcraft
structure, and each integral tank must be subjected
to a pressure of 3.5 p.s.i. unless the pressure
developed during maximum limit acceleration or
emergency deceleration with a full tank exceeds this
value, in which case a hydrostatic head, or
equivalent test, must be applied to duplicate the
acceleration loads as ACAR as possible. However, the
pressure need not exceed 3.5 p.s.i. on surfaces not
exposed to the acceleration loading.
(c)
Each nonmetallic tank with walls supported by the
rotorcraft structure must be subjected to the
following tests:
(1)
A pressure test of at least 2.0 p.s.i. This test may
be conducted on the tank alone in conjunction with
the test specified in paragraph (c)(2) of this
section.
(2)
A pressure test, with the tank mounted in the
rotorcraft structure, equal to the load developed by
the reaction of the contents, with the tank full,
during maximum limit acceleration or emergency
deceleration. However, the pressure need not exceed
2.0 p.s.i. on surfaces faces not exposed to the
acceleration loading.
(d)
Each tank with large unsupported or unstiffened flat
areas, or with other features whose failure or
deformation could cause leakage, must be subjected
to the following test or its equivalent:
(1)
Each complete tank assembly and its supports must be
vibration tested while mounted to simulate the
actual installation.
(2)
The tank assembly must be vibrated for 25 hours
while two-thirds full of any suitable fluid. The
amplitude of vibration may not be less than one
thirty-second of an inch, unless otherwise
substantiated.
(3)
The test frequency of vibration must be as follows:
(i)
If no frequency of vibration resulting from any
r.p.m. within the normal operating range of engine
or rotor system speeds is critical, the test
frequency of vibration, in number of cycles per
minute, must, unless a frequency based on a more
rational analysis is used, be the number obtained by
averaging the maximum and minimum power-on engine
speeds (r.p.m.) for reciprocating engine powered
rotorcraft or 2,000 c.p.m. for turbine engine
powered rotorcraft.
(ii)
If only one frequency of vibration resulting from
any r.p.m. within the normal operating range of
engine or rotor system speeds is critical, that
frequency of vibration must be the test frequency.
(iii) If more than one frequency of vibration
resulting from any r.p.m. within the normal
operating range of engine or rotor system speeds is
critical, the most critical of these frequencies
must be the test frequency.
(4)
Under paragraph (d)(3)(ii) and (iii), the time of
test must be adjusted to accomplish the same number
of vibration cycles as would be accomplished in 25
hours at the frequency specified in paragraph
(d)(3)(i) of this section.
(5)
During the test, the tank assembly must be rocked at
the rate of 16 to 20 complete cycles per minute
through an angle of 15 degrees on both sides of the
horizontal (30 degrees total), about the most
critical axis, for 25 hours. If motion about more
than one axis is likely to be critical, the tank
must be rocked about each critical axis for
121/2hours.
(a)
Each fuel tank must be supported so that tank loads
are not concentrated on unsupported tank surfaces.
In addition—
(1)
There must be pads, if necessary, to prevent chafing
between each tank and its supports;
(2)
The padding must be nonabsorbent or treated to
prevent the absorption of fuel;
(3)
If flexible tank liners are used, they must be
supported so that they are not required to withstand
fluid loads; and
(4)
Each interior surface of tank compartments must be
smooth and free of projections that could cause wear
of the liner, unless—
(i)
There are means for protection of the liner at those
points; or
(ii)
The construction of the liner itself provides such
protection.
(b)
Any spaces adjacent to tank surfaces must be
adequately ventilated to avoid accumulation of fuel
or fumes in those spaces due to minor leakage. If
the tank is in a sealed compartment, ventilation may
be limited to drain holes that prevent clogging and
that prevent excessive pressure resulting from
altitude changes. If flexible tank liners are
installed, the venting arrangement for the spaces
between the liner and its container must maintain
the proper relationship to tank vent pressures for
any expected flight condition.
(c)
The location of each tank must meet the requirements
of 29.1185(b) and (c).
(d)
No rotorcraft skin immediately adjacent to a major
air outlet from the engine compartment may act as
the wall of an integral tank.
Each
fuel tank or each group of fuel tanks with
interconnected vent systems must have an expansion
space of not less than 2 percent of the combined
tank capacity. It must be impossible to fill the
fuel tank expansion space inadvertently with the
rotorcraft in the normal ground attitude.
(a)
Each fuel tank must have a sump with a capacity of
not less than the greater of—
(1)
0.10 per cent of the tank capacity; or
(2)1/16gallon.
(b)
The capacity prescribed in paragraph (a) of this
section must be effective with the rotorcraft in any
normal attitude, and must be located so that the
sump contents cannot escape through the tank outlet
opening.
(c)
Each fuel tank must allow drainage of hazardous
quantities of water from each part of the tank to
the sump with the rotorcraft in any ground attitude
to be expected in service.
(d)
Each fuel tank sump must have a drain that allows
complete drainage of the sump on the ground.
(a)
Each fuel tank filler connection must prevent the
entrance of fuel into any part of the rotorcraft
other than the tank itself during normal operations
and must be crash resistant during a survivable
impact in accordance with 29.952(c). In addition—
(1)
Each filler must be marked as prescribed in
29.1557(c)(1);
(2)
Each recessed filler connection that can retain any
appreciable quantity of fuel must have a drain that
discharges clear of the entire rotorcraft; and
(3)
Each filler cap must provide a fuel-tight seal under
the fluid pressure expected in normal operation and
in a survivable impact.
(b)
Each filler cap or filler cap cover must warn when
the cap is not fully locked or seated on the filler
connection.
(a)
Fuel tank vents. Each fuel tank must be
vented from the top part of the expansion space so
that venting is effective under normal flight
conditions. In addition—
(1)
The vents must be arranged to avoid stoppage by dirt
or ice formation;
(2)
The vent arrangement must prevent siphoning of fuel
during normal operation;
(3)
The venting capacity and vent pressure levels must
maintain acceptable differences of pressure between
the interior and exterior of the tank, during—
(i)
Normal flight operation;
(ii)
Maximum rate of ascent and descent; and
(iii) Refueling and defueling (where applicable);
(4)
Airspaces of tanks with interconnected outlets must
be interconnected;
(5)
There may be no point in any vent line where
moisture can accumulate with the rotorcraft in the
ground attitude or the level flight attitude, unless
drainage is provided;
(6)
No vent or drainage provision may end at any point—
(i)
Where the discharge of fuel from the vent outlet
would constitute a fire hazard; or
(ii)
From which fumes could enter personnel compartments;
and
(7)
The venting system must be designed to minimize
spillage of fuel through the vents to an ignition
source in the event of a rollover during landing,
ground operations, or a survivable impact.
(b)
Carburetor vapor vents. Each carburetor with
vapor elimination connections must have a vent line
to lead vapors back to one of the fuel tanks. In
addition—
(1)
Each vent system must have means to avoid stoppage
by ice; and
(2)
If there is more than one fuel tank, and it is
necessary to use the tanks in a definite sequence,
each vapor vent return line must lead back to the
fuel tank used for takeoff and landing.
(a)
There must be a fuel strainer for the fuel tank
outlet or for the booster pump. This strainer must—
(1)
For reciprocating engine powered airplanes, have 8
to 16 meshes per inch; and
(2)
For turbine engine powered airplanes, prevent the
passage of any object that could restrict fuel flow
or damage any fuel system component.
(b)
The clear area of each fuel tank outlet strainer
must be at least five times the area of the outlet
line.
(c)
The diameter of each strainer must be at least that
of the fuel tank outlet.
(d)
Each finger strainer must be accessible for
inspection and cleaning.
(a)
Each fueling connection below the fuel level in each
tank must have means to prevent the escape of
hazardous quantities of fuel from that tank in case
of malfunction of the fuel entry valve.
(b)
For systems intended for pressure refueling, a means
in addition to the normal means for limiting the
tank content must be installed to prevent damage to
the tank in case of failure of the normal means.
(c)
The rotorcraft pressure fueling system (not fuel
tanks and fuel tank vents) must withstand an
ultimate load that is 2.0 times the load arising
from the maximum pressure, including surge, that is
likely to occur during fueling. The maximum surge
pressure must be established with any combination of
tank valves being either intentionally or
inadvertently closed.
(d)
The rotorcraft defueling system (not including fuel
tanks and fuel tank vents) must withstand an
ultimate load that is 2.0 times the load arising
from the maximum permissible defueling pressure
(positive or negative) at the rotorcraft fueling
connection.
(a)
Compliance with 29.955 must not be jeopardized by
failure of—
(1)
Any one pump except pumps that are approved and
installed as parts of a type certificated engine; or
(2)
Any component required for pump operation except the
engine served by that pump.
(b)
The following fuel pump installation requirements
apply:
(1)
When necessary to maintain the proper fuel pressure—
(i)
A connection must be provided to transmit the
carburetor air intake static pressure to the proper
fuel pump relief valve connection; and
(ii)
The gauge balance lines must be independently
connected to the carburetor inlet pressure to avoid
incorrect fuel pressure readings.
(2)
The installation of fuel pumps having seals or
diaphragms that may leak must have means for
draining leaking fuel.
(3)
Each drain line must discharge where it will not
create a fire hazard.
(a)
Each fuel line must be installed and supported to
prevent excessive vibration and to withstand loads
due to fuel pressure, valve actuation, and
accelerated flight conditions.
(b)
Each fuel line connected to components of the
rotorcraft between which relative motion could exist
must have provisions for flexibility.
(c)
Each flexible connection in fuel lines that may be
under pressure or subjected to axial loading must
use flexible hose assemblies.
(d)
Flexible hose must be approved.
(e)
No flexible hose that might be adversely affected by
high temperatures may be used where excessive
temperatures will exist during operation or after
engine shutdown.
In
addition to meeting the requirements of 29.1189,
each fuel valve must—
(a)
[Reserved]
(b)
Be supported so that no loads resulting from their
operation or from accelerated flight conditions are
transmitted to the lines attached to the valve.
There must be a fuel strainer or filter between the
fuel tank outlet and the inlet of the first fuel
system component which is susceptible to fuel
contamination, including but not limited to the fuel
metering device or an engine positive displacement
pump, whichever is nearer the fuel tank outlet. This
fuel strainer or filter must—
(a)
Be accessible for draining and cleaning and must
incorporate a screen or element which is easily
removable;
(b)
Have a sediment trap and drain, except that it need
not have a drain if the strainer or filter is easily
removable for drain purposes;
(c)
Be mounted so that its weight is not supported by
the connecting lines or by the inlet or outlet
connections of the strainer or filter inself, unless
adequate strength margins under all loading
conditions are provided in the lines and
connections; and
(d)
Provide a means to remove from the fuel any
contaminant which would jeopardize the flow of fuel
through rotorcraft or engine fuel system components
required for proper rotorcraft or engine fuel system
operation.
(a)
There must be at least one accessible drain at the
lowest point in each fuel system to completely drain
the system with the rotorcraft in any ground
attitude to be expected in service.
(b)
Each drain required by paragraph (a) of this section
including the drains prescribed in 29.971 must—
(1)
Discharge clear of all parts of the rotorcraft;
(2)
Have manual or automatic means to ensure positive
closure in the off position; and
(3)
Have a drain valve—
(i)
That is readily accessible and which can be easily
opened and closed; and
(ii)
That is either located or protected to prevent fuel
spillage in the event of a landing with landing gear
retracted.
If a
fuel jettisoning system is installed, the following
apply:
(a)
Fuel jettisoning must be safe during all flight
regimes for which jettisoning is to be authorized.
(b)
In showing compliance with paragraph (a) of this
section, it must be shown that—
(1)
The fuel jettisoning system and its operation are
free from fire hazard;
(2)
No hazard results from fuel or fuel vapors which
impinge on any part of the rotorcraft during fuel
jettisoning; and
(3)
Controllability of the rotorcraft remains
satisfactory throughout the fuel jettisoning
operation.
(c)
Means must be provided to automatically prevent
jettisoning fuel below the level required for an
all-engine climb at maximum continuous power from
sea level to 5,000 feet altitude and cruise
thereafter for 30 minutes at maximum range engine
power.
(d)
The controls for any fuel jettisoning system must be
designed to allow flight personnel (minimum crew) to
safely interrupt fuel jettisoning during any part of
the jettisoning operation.
(e)
The fuel jettisoning system must be designed to
comply with the powerplant installation requirements
of §29.901(c).
(f)
An auxiliary fuel jettisoning system which meets the
requirements of paragraphs (a), (b), (d), and (e) of
this section may be installed to jettison additional
fuel provided it has separate and independent
controls.
(a)
Each engine must have an independent oil system that
can supply it with an appropriate quantity of oil at
a temperature not above that safe for continuous
operation.
(b)
The usable oil capacity of each system may not be
less than the product of the endurance of the
rotorcraft under critical operating conditions and
the maximum allowable oil consumption of the engine
under the same conditions, plus a suitable margin to
ensure adequate circulation and cooling. Instead of
a rational analysis of endurance and consumption, a
usable oil capacity of one gallon for each 40
gallons of usable fuel may be used for reciprocating
engine installations.
(c)
Oil-fuel ratios lower than those prescribed in
paragraph (c) of this section may be used if they
are substantiated by data on the oil consumption of
the engine.
(d)
The ability of the engine and oil cooling provisions
to maintain the oil temperature at or below the
maximum established value must be shown under the
applicable requirements of 29.1041 through 29.1049.
(a)
Installation. Each oil tank installation must
meet the requirements of 29.967.
(b)
Expansion space. Oil tank expansion space
must be provided so that—
(1)
Each oil tank used with a reciprocating engine has
an expansion space of not less than the greater of
10 percent of the tank capacity or 0.5 gallon, and
each oil tank used with a turbine engine has an
expansion space of not less than 10 percent of the
tank capacity;
(2)
Each reserve oil tank not directly connected to any
engine has an expansion space of not less than two
percent of the tank capacity; and
(3)
It is impossible to fill the expansion space
inadvertently with the rotorcraft in the normal
ground attitude.
(c)
Filler connections. Each recessed oil tank
filler connection that can retain any appreciable
quantity of oil must have a drain that discharges
clear of the entire rotorcraft. In addition—
(1)
Each oil tank filler cap must provide an oil-tight
seal under the pressure expected in operation;
(2)
For category A rotorcraft, each oil tank filler cap
or filler cap cover must incorporate features that
provide a warning when caps are not fully locked or
seated on the filler connection; and
(3)
Each oil filler must be marked under 29.1557(c)(2).
(d)
Vent. Oil tanks must be vented as follows:
(1)
Each oil tank must be vented from the top part of
the expansion space to that venting is effective
under all normal flight conditions.
(2)
Oil tank vents must be arranged so that condensed
water vapor that might freeze and obstruct the line
cannot accumulate at any point;
(e)
Outlet. There must be means to prevent
entrance into the tank itself, or into the tank
outlet, of any object that might obstruct the flow
of oil through the system. No oil tank outlet may be
enclosed by a screen or guard that would reduce the
flow of oil below a safe value at any operating
temperature. There must be a shutoff valve at the
outlet of each oil tank used with a turbine engine
unless the external portion of the oil system
(including oil tank supports) is fireproof.
(f)
Flexible liners. Each flexible oil tank liner
must be approved or shown to be suitable for the
particular installation.
Each
oil tank must be designed and installed so that—
(a)
It can withstand, without failure, any vibration,
inertia, and fluid loads to which it may be
subjected in operation; and
(b)
It meets the requirements of 29.965, except that
instead of the pressure specified in 29.965(b)—
(1)
For pressurized tanks used with a turbine engine,
the test pressure may not be less than 5 p.s.i. plus
the maximum operating pressure of the tank; and
(2)
For all other tanks, the test pressure may not be
less than 5 p.s.i.
(a)
Each oil line must meet the requirements of 29.993.
(b)
Breather lines must be arranged so that—
(1)
Condensed water vapor that might freeze and obstruct
the line cannot accumulate at any point;
(2)
The breather discharge will not constitute a fire
hazard if foaming occurs, or cause emitted oil to
strike the pilot's windshield; and
(3)
The breather does not discharge into the engine air
induction system.
(a)
Each turbine engine installation must incorporate an
oil strainer or filter through which all of the
engine oil flows and which meets the following
requirements:
(1)
Each oil strainer or filter that has a bypass must
be constructed and installed so that oil will flow
at the normal rate through the rest of the system
with the strainer or filter completely blocked.
(2)
The oil strainer or filter must have the capacity
(with respect to operating limitations established
for the engine) to ensure that engine oil system
functioning is not impaired when the oil is
contaminated to a degree (with respect to particle
size and density) that is greater than that
established for the engine under Part 33 of this
chapter.
(3)
The oil strainer or filter, unless it is installed
at an oil tank outlet, must incorporate a means to
indicate contamination before it reaches the
capacity established in accordance with paragraph
(a)(2) of this section.
(4)
The bypass of a strainer or filter must be
constructed and installed so that the release of
collected contaminants is minimized by appropriate
location of the bypass to ensure that collected
contaminants are not in the bypass flow path.
(5)
An oil strainer or filter that has no bypass, except
one that is installed at an oil tank outlet, must
have a means to connect it to the warning system
required in 29.1305(a)(18).
(b)
Each oil strainer or filter in a powerplant
installation using reciprocating engines must be
constructed and installed so that oil will flow at
the normal rate through the rest of the system with
the strainer or filter element completely blocked.
A
drain (or drains) must be provided to allow safe
drainage of the oil system. Each drain must—
(a)
Be accessible; and
(b)
Have manual or automatic means for positive locking
in the closed position.
(a)
Each oil radiator must be able to withstand any
vibration, inertia, and oil pressure loads to which
it would be subjected in operation.
(b)
Each oil radiator air duct must be located, or
equipped, so that, in case of fire, and with the
airflow as it would be with and without the engine
operating, flames cannot directly strike the
radiator.
(a)
Each oil shutoff must meet the requirements of
29.1189.
(b)
The closing of oil shutoffs may not prevent
autorotation.
(c)
Each oil valve must have positive stops or suitable
index provisions in the “on” and “off” positions and
must be supported so that no loads resulting from
its operation or from accelerated flight conditions
are transmitted to the lines attached to the valve.
(a)
The oil system for components of the rotor drive
system that require continuous lubrication must be
sufficiently independent of the lubrication systems
of the engine(s) to ensure—
(1)
Operation with any engine inoperative; and
(2)
Safe autorotation.
(b)
Pressure lubrication systems for transmissions and
gearboxes must comply with the requirements of
29.1013, paragraphs (c), (d), and (f) only, 29.1015,
29.1017, 29.1021, 29.1023, and 29.1337(d). In
addition, the system must have—
(1)
An oil strainer or filter through which all the
lubricant flows, and must—
(i)
Be designed to remove from the lubricant any
contaminant which may damage transmission and drive
system components or impede the flow of lubricant to
a hazardous degree; and
(ii)
Be equipped with a bypass constructed and installed
so that—
(A)
The lubricant will flow at the normal rate through
the rest of the system with the strainer or filter
completely blocked; and
(B)
The release of collected contaminants is minimized
by appropriate location of the bypass to ensure that
collected contaminants are not in the bypass flow
path;
(iii) Be equipped with a means to indicate
collection of contaminants on the filter or strainer
at or before opening of the bypass;
(2)
For each lubricant tank or sump outlet supplying
lubrication to rotor drive systems and rotor drive
system components, a screen to prevent entrance into
the lubrication system of any object that might
obstruct the flow of lubricant from the outlet to
the filter required by paragraph (b)(1) of this
section. The requirements of paragraph (b)(1) of
this section do not apply to screens installed at
lubricant tank or sump outlets.
(c)
Splash type lubrication systems for rotor drive
system gearboxes must comply with 29.1021 and
29.1337(d).
(a)
The powerplant and auxiliary power unit cooling
provisions must be able to maintain the temperatures
of powerplant components, engine fluids, and
auxiliary power unit components and fluids within
the temperature limits established for these
components and fluids, under ground, water, and
flight operating conditions for which certification
is requested, and after normal engine or auxiliary
power unit shutdown, or both.
(b)
There must be cooling provisions to maintain the
fluid temperatures in any power transmission within
safe values under any critical surface (ground or
water) and flight operating conditions.
(c)
Except for ground-use-only auxiliary power units,
compliance with paragraphs (a) and (b) of this
section must be shown by flight tests in which the
temperatures of selected powerplant component and
auxiliary power unit component, engine, and
transmission fluids are obtained under the
conditions prescribed in those paragraphs.
(a)
General. For the tests prescribed in
29.1041(c), the following apply:
(1)
If the tests are conducted under conditions
deviating from the maximum ambient atmospheric
temperature specified in paragraph (b) of this
section, the recorded powerplant temperatures must
be corrected under paragraphs (c) and (d) of this
section, unless a more rational correction method is
applicable.
(2)
No corrected temperature determined under paragraph
(a)(1) of this section may exceed established
limits.
(3)
The fuel used during the cooling tests must be of
the minimum grade approved for the engines, and the
mixture settings must be those used in normal
operation.
(4)
The test procedures must be as prescribed in 29.1045
through 29.1049.
(5)
For the purposes of the cooling tests, a temperature
is “stabilized” when its rate of change is less than
2 °F per minute.
(b)
Maximum ambient atmospheric temperature. A
maximum ambient atmospheric temperature
corresponding to sea level conditions of at least
100 degrees F. must be established. The assumed
temperature lapse rate is 3.6 degrees F. per
thousand feet of altitude above sea level until a
temperature of −69.7 degrees F. is reached, above
which altitude the temperature is considered
constant at −69.7 degrees F. However, for
winterization installations, the applicant may
select a maximum ambient atmospheric temperature
corresponding to sea level conditions of less than
100 degrees F.
(c)
Correction factor (except cylinder barrels).
Unless a more rational correction applies,
temperatures of engine fluids and powerplant
components (except cylinder barrels) for which
temperature limits are established, must be
corrected by adding to them the difference between
the maximum ambient atmospheric temperature and the
temperature of the ambient air at the time of the
first occurrence of the maximum component or fluid
temperature recorded during the cooling test.
(d)
Correction factor for cylinder barrel
temperatures. Cylinder barrel temperatures must
be corrected by adding to them 0.7 times the
difference between the maximum ambient atmospheric
temperature and the temperature of the ambient air
at the time of the first occurrence of the maximum
cylinder barrel temperature recorded during the
cooling test.
(a)
Climb cooling tests must be conducted under this
section for—
(1)
Category A rotorcraft; and
(2)
Multiengine category B rotorcraft for which
certification is requested under the category A
powerplant installation requirements, and under the
requirements of 29.861(a) at the steady rate of
climb or descent established under 29.67(b).
(b)
The climb or descent cooling tests must be conducted
with the engine inoperative that produces the most
adverse cooling conditions for the remaining engines
and powerplant components.
(c)
Each operating engine must—
(1)
For helicopters for which the use of 30-minute OEI
power is requested, be at 30-minute OEI power for 30
minutes, and then at maximum continuous power (or at
full throttle when above the critical altitude);
(2)
For helicopters for which the use of continuous OEI
power is requested, be at continuous OEI power (or
at full throttle when above the critical altitude);
and
(3)
For other rotorcraft, be at maximum continuous power
(or at full throttle when above the critical
altitude).
(d)
After temperatures have stabilized in flight, the
climb must be—
(1)
Begun from an altitude not greater than the lower
of—
(i)
1,000 feet below the engine critcal altitude; and
(ii)
1,000 feet below the maximum altitude at which the
rate of climb is 150 f.p.m; and
(2)
Continued for at least five minutes after the
occurrence of the highest temperature recorded, or
until the rotorcraft reaches the maximum altitude
for which certification is requested.
(e)
For category B rotorcraft without a positive rate of
climb, the descent must begin at the
all-engine-critical altitude and end at the higher
of—
(1)
The maximum altitude at which level flight can be
maintained with one engine operative; and
(2)
Sea level.
(f)
The climb or descent must be conducted at an
airspeed representing a normal operational practice
for the configuration being tested. However, if the
cooling provisions are sensitive to rotorcraft
speed, the most critical airspeed must be used, but
need not exceed the speeds established under
29.67(a)(2) or 29.67(b). The climb cooling test may
be conducted in conjunction with the takeoff cooling
test of 29.1047.
(a)
Category A. For each category A rotorcraft,
cooling must be shown during takeoff and subsequent
climb as follows:
(1)
Each temperature must be stabilized while hovering
in ground effect with—
(i)
The power necessary for hovering;
(ii)
The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2)
After the temperatures have stabilized, a climb must
be started at the lowest practicable altitude and
must be conducted with one engine inoperative.
(3)
The operating engines must be at the greatest power
for which approval is sought (or at full throttle
when above the critical altitude) for the same
period as this power is used in determining the
takeoff climbout path under 29.59.
(4)
At the end of the time interval prescribed in
paragraph (b)(3) of this section, the power must be
changed to that used in meeting 29.67(a)(2) and the
climb must be continued for—
(i)
Thirty minutes, if 30-minute OEI power is used; or
(ii)
At least 5 minutes after the occurrence of the
highest temperature recorded, if continuous OEI
power or maximum continuous power is used.
(5)
The speeds must be those used in determining the
takeoff flight path under 29.59.
(b)
Category B. For each category B rotorcraft,
cooling must be shown during takeoff and subsequent
climb as follows:
(1)
Each temperature must be stabilized while hovering
in ground effect with—
(i)
The power necessary for hovering;
(ii)
The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2)
After the temperatures have stabilized, a climb must
be started at the lowest practicable altitude with
takeoff power.
(3)
Takeoff power must be used for the same time
interval as takeoff power is used in determining the
takeoff flight path under 29.63.
(4)
At the end of the time interval prescribed in
paragraph (a)(3) of this section, the power must be
reduced to maximum continuous power and the climb
must be continued for at least five minutes after
the occurence of the highest temperature recorded.
(5)
The cooling test must be conducted at an airspeed
corresponding to normal operating practice for the
configuration being tested. However, if the cooling
provisions are sensitive to rotorcraft speed, the
most critical airspeed must be used, but need not
exceed the speed for best rate of climb with maximum
continuous power.
The
hovering cooling provisions must be shown—
(a)
At maximum weight or at the greatest weight at which
the rotorcraft can hover (if less), at sea level,
with the power required to hover but not more than
maximum continuous power, in the ground effect in
still air, until at least five minutes after the
occurrence of the highest temperature recorded; and
(b)
With maximum continuous power, maximum weight, and
at the altitude resulting in zero rate of climb for
this configuration, until at least five minutes
after the occurrence of the highest temperature
recorded.
(a)
The air induction system for each engine and
auxiliary power unit must supply the air required by
that engine and auxiliary power unit under the
operating conditions for which certification is
requested.
(b)
Each engine and auxiliary power unit air induction
system must provide air for proper fuel metering and
mixture distribution with the induction system
valves in any position.
(c)
No air intake may open within the engine accessory
section or within other areas of any powerplant
compartment where emergence of backfire flame would
constitute a fire hazard.
(d)
Each reciprocating engine must have an alternate air
source.
(e)
Each alternate air intake must be located to prevent
the entrance of rain, ice, or other foreign matter.
(f)
For turbine engine powered rotorcraft and rotorcraft
incorporating auxiliary power units—
(1)
There must be means to prevent hazardous quantities
of fuel leakage or overflow from drains, vents, or
other components of flammable fluid systems from
entering the engine or auxiliary power unit intake
system; and
(2)
The air inlet ducts must be located or protected so
as to minimize the ingestion of foreign matter
during takeoff, landing, and taxiing.
(a)
Reciprocating engines. Each reciprocating
engine air induction system must have means to
prevent and eliminate icing. Unless this is done by
other means, it must be shown that, in air free of
visible moisture at a temperature of 30 °F., and
with the engines at 60 percent of maximum continuous
power—
(1)
Each rotorcraft with sea level engines using
conventional venturi carburetors has a pre-heater
that can provide a heat rise of 90 °F.;
(2)
Each rotorcraft with sea level engines using
carburetors tending to prevent icing has a
pre-heater that can provide a heat rise of 70 °F.;
(3)
Each rotorcraft with altitude engines using
conventional venturi carburetors has a pre-heater
that can provide a heat rise of 120 °F.; and
(4)
Each rotorcraft with altitude engines using
carburetors tending to prevent icing has a
pre-heater that can provide a heat rise of 100 °F.
(b)
Turbine engines. (1) It must be shown that
each turbine engine and its air inlet system can
operate throughout the flight power range of the
engine (including idling)—
(i)
Without accumulating ice on engine or inlet system
components that would adversely affect engine
operation or cause a serious loss of power under the
icing conditions specified in appendix C of this
Part; and
(ii)
In snow, both falling and blowing, without adverse
effect on engine operation, within the limitations
established for the rotorcraft.
(2)
Each turbine engine must idle for 30 minutes on the
ground, with the air bleed available for engine
icing protection at its critical condition, without
adverse effect, in an atmosphere that is at a
temperature between 15° and 30 °F (between −9° and
−1 °C) and has a liquid water content not less than
0.3 grams per cubic meter in the form of drops
having a mean effective diameter not less than 20
microns, followed by momentary operation at takeoff
power or thrust. During the 30 minutes of idle
operation, the engine may be run up periodically to
a moderate power or thrust setting in a manner
acceptable to the Administrator.
(c)
Supercharged reciprocating engines. For each
engine having a supercharger to pressurize the air
before it enters the carburetor, the heat rise in
the air caused by that supercharging at any altitude
may be utilized in determining compliance with
paragraph (a) of this section if the heat rise
utilized is that which will be available,
automatically, for the applicable altitude and
operation condition because of supercharging.
Each
carburetor air pre-heater must be designed and
constructed to—
(a)
Ensure ventilation of the pre-heater when the engine
is operated in cold air;
(b)
Allow inspection of the exhaust manifold parts that
it surrounds; and
(c)
Allow inspection of critical parts of the preheater
itself.
(a)
Each induction system duct upstream of the first
stage of the engine supercharger and of the
auxiliary power unit compressor must have a drain to
prevent the hazardous accumulation of fuel and
moisture in the ground attitude. No drain may
discharge where it might cause a fire hazard.
(b)
Each duct must be strong enough to prevent induction
system failure from normal backfire conditions.
(c)
Each duct connected to components between which
relative motion could exist must have means for
flexibility.
(d)
Each duct within any fire zone for which a
fire-extinguishing system is required must be at
least—
(1)
Fireproof, if it passes through any firewall; or
(2)
Fire resistant, for other ducts, except that ducts
for auxiliary power units must be fireproof within
the auxiliary power unit fire zone.
(e)
Each auxiliary power unit induction system duct must
be fireproof for a sufficient distance upstream of
the auxiliary power unit compartment to prevent hot
gas reverse flow from burning through auxiliary
power unit ducts and entering any other compartment
or area of the rotorcraft in which a hazard would be
created resulting from the entry of hot gases. The
materials used to form the remainder of the
induction system duct and plenum chamber of the
auxiliary power unit must be capable of resisting
the maximum heat conditions likely to occur.
(f)
Each auxiliary power unit induction system duct must
be constructed of materials that will not absorb or
trap hazardous quantities of flammable fluids that
could be ignited in the event of a surge or reverse
flow condition.
If
induction system screens are used—
(a)
Each screen must be upstream of the carburetor;
(b)
No screen may be in any part of the induction system
that is the only passage through which air can reach
the engine, unless it can be deiced by heated air;
(c)
No screen may be deiced by alcohol alone; and
(d)
It must be impossible for fuel to strike any screen.
Each
inter-cooler and after-cooler must be able to
withstand the vibration, inertia, and air pressure
loads to which it would be subjected in operation.
It
must be shown under 29.1043 that each installation
using two-stage superchargers has means to maintain
the air temperature, at the carburetor inlet, at or
below the maximum established value.
Exhaust System
For
powerplant and auxiliary power unit installations
the following apply:
(a)
Each exhaust system must ensure safe disposal of
exhaust gases without fire hazard or carbon monoxide
contamination in any personnel compartment.
(b)
Each exhaust system part with a surface hot enough
to ignite flammable fluids or vapors must be located
or shielded so that leakage from any system carrying
flammable fluids or vapors will not result in a fire
caused by impingement of the fluids or vapors on any
part of the exhaust system including shields for the
exhaust system.
(c)
Each component upon which hot exhaust gases could
impinge, or that could be subjected to high
temperatures from exhaust system parts, must be
fireproof. Each exhaust system component must be
separated by a fireproof shield from adjacent parts
of the rotorcraft that are outside the engine and
auxiliary power unit compartments.
(d)
No exhaust gases may discharge so as to cause a fire
hazard with respect to any flammable fluid vent or
drain.
(e)
No exhaust gases may discharge where they will cause
a glare seriously affecting pilot vision at night.
(f)
Each exhaust system component must be ventilated to
prevent points of excessively high temperature.
(g)
Each exhaust shroud must be ventilated or insulated
to avoid, during normal operation, a temperature
high enough to ignite any flammable fluids or vapors
outside the shroud.
(h)
If significant traps exist, each turbine engine
exhaust system must have drains discharging clear of
the rotorcraft, in any normal ground and flight
attitudes, to prevent fuel accumulation after the
failure of an attempted engine start.
(a)
Exhaust piping must be heat and corrosion resistant,
and must have provisions to prevent failure due to
expansion by operating temperatures.
(b)
Exhaust piping must be supported to withstand any
vibration and inertia loads to which it would be
subjected in operation.
(c)
Exhaust piping connected to components between which
relative motion could exist must have provisions for
flexibility.
For
reciprocating engine powered rotorcraft the
following apply:
(a)
Each exhaust heat exchanger must be constructed and
installed to withstand the vibration, inertia, and
other loads to which it would be subjected in
operation. In addition—
(1)
Each exchanger must be suitable for continued
operation at high temperatures and resistant to
corrosion from exhaust gases;
(2)
There must be means for inspecting the critical
parts of each exchanger;
(3)
Each exchanger must have cooling provisions wherever
it is subject to contact with exhaust gases; and
(4)
No exhaust heat exchanger or muff may have stagnant
areas or liquid traps that would increase the
probability of ignition of flammable fluids or
vapors that might be present in case of the failure
or malfunction of components carrying flammable
fluids.
(b)
If an exhaust heat exchanger is used for heating
ventilating air used by personnel—
(1)
There must be a secondary heat exchanger between the
primary exhaust gas heat exchanger and the
ventilating air system; or
(2)
Other means must be used to prevent harmful
contamination of the ventilating air.
(a)
Powerplant controls must be located and arranged
under 29.777 and marked under 29.1555.
(b)
Each control must be located so that it cannot be
inadvertently operated by persons entering, leaving,
or moving normally in the cockpit.
(c)
Each flexible powerplant control must be approved.
(d)
Each control must be able to maintain any set
position without—
(1)
Constant attention; or
(2)
Tendency to creep due to control loads or vibration.
(e)
Each control must be able to withstand operating
loads without excessive deflection.
(f)
Controls of powerplant valves required for safety
must have—
(1)
For manual valves, positive stops or in the case of
fuel valves suitable index provisions, in the open
and closed position; and
(2)
For power-assisted valves, a means to indicate to
the flight crew when the valve—
(i)
Is in the fully open or fully closed position; or
(ii)
Is moving between the fully open and fully closed
position.
Means must be provided on the flight deck for
starting, stopping, and emergency shutdown of each
installed auxiliary power unit.
(a)
There must be a separate power control for each
engine.
(b)
Power controls must be arranged to allow ready
synchronization of all engines by—
(1)
Separate control of each engine; and
(2)
Simultaneous control of all engines.
(c)
Each power control must provide a positive and
immediately responsive means of controlling its
engine.
(d)
Each fluid injection control other than fuel system
control must be in the corresponding power control.
However, the injection system pump may have a
separate control.
(e)
If a power control incorporates a fuel shutoff
feature, the control must have a means to prevent
the inadvertent movement of the control into the
shutoff position. The means must—
(1)
Have a positive lock or stop at the idle position;
and
(2)
Require a separate and distinct operation to place
the control in the shutoff position.
(f)
For rotorcraft to be certificated for a 30-second
OEI power rating, a means must be provided to
automatically activate and control the 30-second OEI
power and prevent any engine from exceeding the
installed engine limits associated with the
30-second OEI power rating approved for the
rotorcraft.
(a)
Ignition switches must control each ignition circuit
on each engine.
(b)
There must be means to quickly shut off all ignition
by the grouping of switches or by a master ignition
control.
(c)
Each group of ignition switches, except ignition
switches for turbine engines for which continuous
ignition is not required, and each master ignition
control must have a means to prevent its inadvertent
operation.
(a)
If there are mixture controls, each engine must have
a separate control, and the controls must be
arranged to allow—
(1)
Separate control of each engine; and
(2)
Simultaneous control of all engines.
(b)
Each intermediate position of the mixture controls
that corresponds to a normal operating setting must
be identifiable by feel and sight.
(a)
It must be impossible to apply the rotor brake
inadvertently in flight.
(b)
There must be means to warn the crew if the rotor
brake has not been completely released before
takeoff.
There must be a separate carburetor air temperature
control for each engine.
29.1159 Supercharger controls.
Each
supercharger control must be accessible to—
(a)
The pilots; or
(b)
(If there is a separate flight engineer station with
a control panel) the flight engineer.
(a)
Each engine mounted accessory must—
(1)
Be approved for mounting on the engine involved;
(2)
Use the provisions on the engine for mounting; and
(3)
Be sealed in such a way as to prevent contamination
of the engine oil system and the accessory system.
(b)
Electrical equipment subject to arcing or sparking
must be installed, to minimize the probability of
igniting flammable fluids or vapors.
(c)
If continued rotation of an engine-driven cabin
supercharger or any remote accessory driven by the
engine will be a hazard if they malfunction, there
must be means to prevent their hazardous rotation
without interfering with the continued operation of
the engine.
(d)
Unless other means are provided, torque limiting
means must be provided for accessory drives located
on any component of the transmission and rotor drive
system to prevent damage to these components from
excessive accessory load.
(a)
Each battery ignition system must be supplemented
with a generator that is automatically available as
an alternate source of electrical energy to allow
continued engine operation if any battery becomes
depleted.
(b)
The capacity of batteries and generators must be
large enough to meet the simultaneous demands of the
engine ignition system and the greatest demands of
any electrical system components that draw from the
same source.
(c)
The design of the engine ignition system must
account for—
(1)
The condition of an inoperative generator;
(2)
The condition of a completely depleted battery with
the generator running at its normal operating speed;
and
(3)
The condition of a completely depleted battery with
the generator operating at idling speed, if there is
only one battery.
(d)
Magneto ground wiring (for separate ignition
circuits) that lies on the engine side of any
firewall must be installed, located, or protected,
to minimize the probability of the simultaneous
failure of two or more wires as a result of
mechanical damage, electrical fault, or other cause.
(e)
No ground wire for any engine may be routed through
a fire zone of another engine unless each part of
that wire within that zone is fireproof.
(f)
Each ignition system must be independent of any
electrical circuit that is not used for assisting,
controlling, or analyzing the operation of that
system.
(g)
There must be means to warn appropriate crewmembers
if the malfunctioning of any part of the electrical
system is causing the continuous discharge of any
battery necessary for engine ignition.
29.1181 Designated fire zones: regions included.
(a)
Designated fire zones are—
(1)
The engine power section of reciprocating engines;
(2)
The engine accessory section of reciprocating
engines;
(3)
Any complete powerplant compartment in which there
is no isolation between the engine power section and
the engine accessory section, for reciprocating
engines;
(4)
Any auxiliary power unit compartment;
(5)
Any fuel-burning heater and other combustion
equipment installation described in 29.859;
(6)
The compressor and accessory sections of turbine
engines; and
(7)
The combustor, turbine, and tailpipe sections of
turbine engine installations except sections that do
not contain lines and components carrying flammable
fluids or gases and are isolated from the designated
fire zone prescribed in paragraph (a)(6) of this
section by a firewall that meets 29.1191.
(b)
Each designated fire zone must meet the requirements
of 29.1183 through 29.1203.
(a)
Except as provided in paragraph (b) of this section,
each line, fitting, and other component carrying
flammable fluid in any area subject to engine fire
conditions and each component which conveys or
contains flammable fluid in a designated fire zone
must be fire resistant, except that flammable fluid
tanks and supports in a designated fire zone must be
fireproof or be enclosed by a fireproof shield
unless damage by fire to any non-fireproof part will
not cause leakage or spillage of flammable fluid.
Components must be shielded or located so as to
safeguard against the ignition of leaking flammable
fluid. An integral oil sump of less than 25-quart
capacity on a reciprocating engine need not be
fireproof nor be enclosed by a fireproof shield.
(b)
Paragraph (a) of this section does not apply to—
(1)
Lines, fittings, and components which are already
approved as part of a type certificated engine; and
(2)
Vent and drain lines, and their fittings, whose
failure will not result in or add to, a fire hazard.
(a)
No tank or reservoir that is part of a system
containing flammable fluids or gases may be in a
designated fire zone unless the fluid contained, the
design of the system, the materials used in the tank
and its supports, the shutoff means, and the
connections, lines, and controls provide a degree of
safety equal to that which would exist if the tank
or reservoir were outside such a zone.
(b)
Each fuel tank must be isolated from the engines by
a firewall or shroud.
(c)
There must be at least one-half inch of clear
airspace between each tank or reservoir and each
firewall or shroud isolating a designated fire zone,
unless equivalent means are used to prevent heat
transfer from the fire zone to the flammable fluid.
(d)
Absorbent material close to flammable fluid system
components that might leak must be covered or
treated to prevent the absorption of hazardous
quantities of fluids.
(a)
There must be complete drainage of each part of each
designated fire zone to minimize the hazards
resulting from failure or malfunction of any
component containing flammable fluids. The drainage
means must be—
(1)
Effective under conditions expected to prevail when
drainage is needed; and
(2)
Arranged so that no discharged fluid will cause an
additional fire hazard.
(b)
Each designated fire zone must be ventilated to
prevent the accumulation of flammable vapors.
(c)
No ventilation opening may be where it would allow
the entry of flammable fluids, vapors, or flame from
other zones.
(d)
Ventilation means must be arranged so that no
discharged vapors will cause an additional fire
hazard.
(e)
For category A rotorcraft, there must be means to
allow the crew to shut off the sources of forced
ventilation in any fire zone (other than the engine
power section of the powerplant compartment) unless
the amount of extinguishing agent and the rate of
discharge are based on the maximum airflow through
that zone.
(a)
There must be means to shut off or otherwise prevent
hazardous quantities of fuel, oil, de-icing fluid,
and other flammable fluids from flowing into,
within, or through any designated fire zone, except
that this means need not be provided—
(1)
For lines, fittings, and components forming an
integral part of an engine;
(2)
For oil systems for turbine engine installations in
which all components of the system, including oil
tanks, are fireproof or located in areas not subject
to engine fire conditions; or
(3)
For engine oil systems in category B rotorcraft
using reciprocating engines of less than 500 cubic
inches displacement.
(b)
The closing of any fuel shutoff valve for any engine
may not make fuel unavailable to the remaining
engines.
(c)
For category A rotorcraft, no hazardous quantity of
flammable fluid may drain into any designated fire
zone after shutoff has been accomplished, nor may
the closing of any fuel shutoff valve for an engine
make fuel unavailable to the remaining engines.
(d)
The operation of any shutoff may not interfere with
the later emergency operation of any other
equipment, such as the means for declutching the
engine from the rotor drive.
(e)
Each shutoff valve and its control must be designed,
located, and protected to function properly under
any condition likely to result from fire in a
designated fire zone.
(f)
Except for ground-use-only auxiliary power unit
installations, there must be means to prevent
inadvertent operation of each shutoff and to make it
possible to reopen it in flight after it has been
closed.
(a)
Each engine, including the combustor, turbine, and
tailpipe sections of turbine engine installations,
must be isolated by a firewall, shroud, or
equivalent means, from personnel compartments,
structures, controls, rotor mechanisms, and other
parts that are—
(1)
Essential to controlled flight and landing; and
(2)
Not protected under 29.861.
(b)
Each auxiliary power unit, combustion heater, and
other combustion equipment to be used in flight,
must be isolated from the rest of the rotorcraft by
firewalls, shrouds, or equivalent means.
(c)
Each firewall or shroud must be constructed so that
no hazardous quantity of air, fluid, or flame can
pass from any engine compartment to other parts of
the rotorcraft.
(d)
Each opening in the firewall or shroud must be
sealed with close-fitting fireproof grommets,
bushings, or firewall fittings.
(e)
Each firewall and shroud must be fireproof and
protected against corrosion.
(f)
In meeting this section, account must be taken of
the probable path of a fire as affected by the
airflow in normal flight and in autorotation.
a)
Each cowling and engine compartment covering must be
constructed and supported so that it can resist the
vibration, inertia, and air loads to which it may be
subjected in operation.
(b)
Cowling must meet the drainage and ventilation
requirements of 29.1187.
(c)
On rotorcraft with a diaphragm isolating the engine
power section from the engine accessory section,
each part of the accessory section cowling subject
to flame in case of fire in the engine power section
of the powerplant must—
(1)
Be fireproof; and
(2)
Meet the requirements of 29.1191.
(d)
Each part of the cowling or engine compartment
covering subject to high temperatures due to its
nearness to exhaust system parts or exhaust gas
impingement must be fireproof.
(e)
Each rotorcraft must—
(1)
Be designated and constructed so that no fire
originating in any fire zone can enter, either
through openings or by burning through external
skin, any other zone or region where it would create
additional hazards;
(2)
Meet the requirements of paragraph (e)(1) of this
section with the landing gear retracted (if
applicable); and
(3)
Have fireproof skin in areas subject to flame if a
fire starts in or burns out of any designated fire
zone.
(f)
A means of retention for each openable or readily
removable panel, cowling, or engine or rotor drive
system covering must be provided to preclude
hazardous damage to rotors or critical control
components in the event of—
(1)
Structural or mechanical failure of the normal
retention means, unless such failure is extremely
improbable; or
(2)
Fire in a fire zone, if such fire could adversely
affect the normal means of retention.
All
surfaces aft of, and near, engine compartments and
designated fire zones, other than tail surfaces not
subject to heat, flames, or sparks emanating from a
designated fire zone or engine compartment, must be
at least fire resistant.
(a)
Each turbine engine powered rotorcraft and Category
A reciprocating engine powered rotorcraft, and each
Category B reciprocating engine powered rotorcraft
with engines of more than 1,500 cubic inches must
have a fire extinguishing system for the designated
fire zones. The fire extinguishing system for a
powerplant must be able to simultaneously protect
all zones of the powerplant compartment for which
protection is provided.
(b)
For multiengine powered rotorcraft, the fire
extinguishing system, the quantity of extinguishing
agent, and the rate of discharge must—
(1)
For each auxiliary power unit and combustion
equipment, provide at least one adequate discharge;
and
(2)
For each other designated fire zone, provide two
adequate discharges.
(c)
For single engine rotorcraft, the quantity of
extinguishing agent and the rate of discharge must
provide at least one adequate discharge for the
engine compartment.
(d)
It must be shown by either actual or simulated
flight tests that under critical airflow conditions
in flight the discharge of the extinguishing agent
in each designated fire zone will provide an agent
concentration capable of extinguishing fires in that
zone and of minimizing the probability of reignition.
(a)
Fire extinguishing agents must—
(1)
Be capable of extinguishing flames emanating from
any burning of fluids or other combustible materials
in the area protected by the fire extinguishing
system; and
(2)
Have thermal stability over the temperature range
likely to be experienced in the compartment in which
they are stored.
(b)
If any toxic extinguishing agent is used, it must be
shown by test that entry of harmful concentrations
of fluid or fluid vapors into any personnel
compartment (due to leakage during normal operation
of the rotorcraft, or discharge on the ground or in
flight) is prevented, even though a defect may exist
in the extinguishing system.
(a)
Each extinguishing agent container must have a
pressure relief to prevent bursting of the container
by excessive internal pressures.
(b)
The discharge end of each discharge line from a
pressure relief connection must be located so that
discharge of the fire extinguishing agent would not
damage the rotorcraft. The line must also be located
or protected to prevent clogging caused by ice or
other foreign matter.
(c)
There must be a means for each fire extinguishing
agent container to indicate that the container has
discharged or that the charging pressure is below
the established minimum necessary for proper
functioning.
(d)
The temperature of each container must be
maintained, under intended operating conditions, to
prevent the pressure in the container from—
(1)
Falling below that necessary to provide an adequate
rate of discharge; or
(2)
Rising high enough to cause premature discharge.
(a)
No materials in any fire extinguishing system may
react chemically with any extinguishing agent so as
to create a hazard.
(b)
Each system component in an engine compartment must
be fireproof.
(a)
For each turbine engine powered rotorcraft and
Category A reciprocating engine powered rotorcraft,
and for each Category B reciprocating engine powered
rotorcraft with engines of more than 900 cubic
inches displacement, there must be approved,
quick-acting fire detectors in designated fire zones
and in the combustor, turbine, and tailpipe sections
of turbine installations (whether or not such
sections are designated fire zones) in numbers and
locations ensuring prompt detection of fire in those
zones.
(b)
Each fire detector must be constructed and installed
to withstand any vibration, inertia, and other loads
to which it would be subjected in operation.
(c)
No fire detector may be affected by any oil, water,
other fluids, or fumes that might be present.
(d)
There must be means to allow crewmembers to check,
in flight, the functioning of each fire detector
system electrical circuit.
(e)
The writing and other components of each fire
detector system in an engine compartment must be at
least fire resistant.
(f)
No fire detector system component for any fire zone
may pass through another fire zone, unless—
(1)
It is protected against the possibility of false
warnings resulting from fires in zones through which
it passes; or
(2)
The zones involved are simultaneously protected by
the same detector and extinguishing systems.
Each
item of installed equipment must—
(a)
Be of a kind and design appropriate to its intended
function;
(b)
Be labeled as to its identification, function, or
operating limitations, or any applicable combination
of these factors;
(c)
Be installed according to limitations specified for
that equipment; and
(d)
Function properly when installed.
The
following are required flight and navigational
instruments:
(a)
An airspeed indicator. For Category A rotorcraft
with VNEless than a speed at which
unmistakable pilot cues provide overspeed warning, a
maximum allowable airspeed indicator must be
provided. If maximum allowable airspeed varies with
weight, altitude, temperature, or r.p.m., the
indicator must show that variation.
(b)
A sensitive altimeter.
(c)
A magnetic direction indicator.
(d)
A clock displaying hours, minutes, and seconds with
a sweep-second pointer or digital presentation.
(e)
A free-air temperature indicator.
(f)
A non-tumbling gyroscopic bank and pitch indicator.
(g)
A gyroscopic rate-of-turn indicator combined with an
integral slip-skid indicator (turn-and-bank
indicator) except that only a slip-skid indicator is
required on rotorcraft with a third attitude
instrument system that—
(1)
Is usable through flight attitudes of ±80 degrees of
pitch and ±120 degrees of roll;
(2)
Is powered from a source independent of the
electrical generating system;
(3)
Continues reliable operation for a minimum of 30
minutes after total failure of the electrical
generating system;
(4)
Operates independently of any other attitude
indicating system;
(5)
Is operative without selection after total failure
of the electrical generating system;
(6)
Is located on the instrument panel in a position
acceptable to the Administrator that will make it
plainly visible to and useable by any pilot at his
station; and
(7)
Is appropriately lighted during all phases of
operation.
(h)
A gyroscopic direction indicator.
(i)
A rate-of-climb (vertical speed) indicator.
(j)
For Category A rotorcraft, a speed warning device
when VNEis less than the speed at which
unmistakable overspeed warning is provided by other
pilot cues. The speed warning device must give
effective aural warning (differing distinctively
from aural warnings used for other purposes) to the
pilots whenever the indicated speed exceeds VNEplus
3 knots and must operate satisfactorily throughout
the approved range of altitudes and temperatures.
The
following are required powerplant instruments:
(a)
For each rotorcraft—
(1)
A carburetor air temperature indicator for each
reciprocating engine;
(2)
A cylinder head temperature indicator for each
air-cooled reciprocating engine, and a coolant
temperature indicator for each liquid-cooled
reciprocating engine;
(3)
A fuel quantity indicator for each fuel tank;
(4)
A low fuel warning device for each fuel tank which
feeds an engine. This device must—
(i)
Provide a warning to the crew when approximately 10
minutes of usable fuel remains in the tank; and
(ii)
Be independent of the normal fuel quantity
indicating system.
(5)
A manifold pressure indicator, for each
reciprocating engine of the altitude type;
(6)
An oil pressure indicator for each
pressure-lubricated gearbox.
(7)
An oil pressure warning device for each
pressure-lubricated gearbox to indicate when the oil
pressure falls below a safe value;
(8)
An oil quantity indicator for each oil tank and each
rotor drive gearbox, if lubricant is self-contained;
(9)
An oil temperature indicator for each engine;
(10)
An oil temperature warning device to indicate unsafe
oil temperatures in each main rotor drive gearbox,
including gearboxes necessary for rotor phasing;
(11)
A gas temperature indicator for each turbine engine;
(12)
A gas producer rotor tachometer for each turbine
engine;
(13)
A tachometer for each engine that, if combined with
the applicable instrument required by paragraph
(a)(14) of this section, indicates rotor r.p.m.
during autorotation.
(14)
At least one tachometer to indicate, as applicable—
(i)
The r.p.m. of the single main rotor;
(ii)
The common r.p.m. of any main rotors whose speeds
cannot vary appreciably with respect to each other;
and
(iii) The r.p.m. of each main rotor whose speed can
vary appreciably with respect to that of another
main rotor;
(15)
A free power turbine tachometer for each turbine
engine;
(16)
A means, for each turbine engine, to indicate power
for that engine;
(17)
For each turbine engine, an indicator to indicate
the functioning of the powerplant ice protection
system;
(18)
An indicator for the filter required by 29.997 to
indicate the occurrence of contamination of the
filter to the degree established in compliance with
29.955;
(19)
For each turbine engine, a warning means for the oil
strainer or filter required by 29.1019, if it has no
bypass, to warn the pilot of the occurrence of
contamination of the strainer or filter before it
reaches the capacity established in accordance with
29.1019(a)(2);
(20)
An indicator to indicate the functioning of any
selectable or controllable heater used to prevent
ice clogging of fuel system components;
(21)
An individual fuel pressure indicator for each
engine, unless the fuel system which supplies that
engine does not employ any pumps, filters, or other
components subject to degradation or failure which
may adversely affect fuel pressure at the engine;
(22)
A means to indicate to the flightcrew the failure of
any fuel pump installed to show compliance with
29.955;
(23)
Warning or caution devices to signal to the
flightcrew when ferromagnetic particles are detected
by the chip detector required by 29.1337(e); and
(24)
For auxiliary power units, an individual indicator,
warning or caution device, or other means to advise
the flightcrew that limits are being exceeded, if
exceeding these limits can be hazardous, for—
(i)
Gas temperature;
(ii)
Oil pressure; and
(iii) Rotor speed.
(25)
For rotorcraft for which a 30-second/2-minute OEI
power rating is requested, a means must be provided
to alert the pilot when the engine is at the
30-second and 2-minute OEI power levels, when the
event begins, and when the time interval expires.
(26)
For each turbine engine utilizing 30-second/2-minute
OEI power, a device or system must be provided for
use by ground personnel which—
(i)
Automatically records each usage and duration of
power at the 30-second and 2-minute OEI levels;
(ii)
Permits retrieval of the recorded data;
(iii) Can be reset only by ground maintenance
personnel; and
(iv)
Has a means to verify proper operation of the system
or device.
(b)
For category A rotorcraft—
(1)
An individual oil pressure indicator for each
engine, and either an independent warning device for
each engine or a master warning device for the
engines with means for isolating the individual
warning circuit from the master warning device;
(2)
An independent fuel pressure warning device for each
engine or a master warning device for all engines
with provision for isolating the individual warning
device from the master warning device; and
(3)
Fire warning indicators.
(c)
For category B rotorcraft—
(1)
An individual oil pressure indicator for each
engine; and
(2)
Fire warning indicators, when fire detection is
required.
The
following is required miscellaneous equipment:
(a)
An approved seat for each occupant.
(b)
A master switch arrangement for electrical circuits
other than ignition.
(c)
Hand fire extinguishers.
(d)
A windshield wiper or equivalent device for each
pilot station.
(e)
A two-way radio communication system.
(a)
The equipment, systems, and installations whose
functioning is required by this subchapter must be
designed and installed to ensure that they perform
their intended functions under any foreseeable
operating condition.
(b)
The rotorcraft systems and associated components,
considered separately and in relation to other
systems, must be designed so that—
(1)
For Category B rotorcraft, the equipment, systems,
and installations must be designed to prevent
hazards to the rotorcraft if they malfunction or
fail; or
(2)
For Category A rotorcraft—
(i)
The occurrence of any failure condition which would
prevent the continued safe flight and landing of the
rotorcraft is extremely improbable; and
(ii)
The occurrence of any other failure conditions which
would reduce the capability of the rotorcraft or the
ability of the crew to cope with adverse operating
conditions is improbable.
(c)
Warning information must be provided to alert the
crew to unsafe system operating conditions and to
enable them to take appropriate corrective action.
Systems, controls, and associated monitoring and
warning means must be designed to minimize crew
errors which could create additional hazards.
(d)
Compliance with the requirements of paragraph (b)(2)
of this section must be shown by analysis and, where
necessary, by appropriate ground, flight, or
simulator tests. The analysis must consider—
(1)
Possible modes of failure, including malfunctions
and damage from external sources;
(2)
The probability of multiple failures and undetected
failures;
(3)
The resulting effects on the rotorcraft and
occupants, considering the stage of flight and
operating conditions; and
(4)
The crew warning cues, corrective action required,
and the capability of detecting faults.
(e)
For Category A rotorcraft, each installation whose
functioning is required by this subchapter and which
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 system with the system
functioning normally.
(2)
Essential loads, after failure of any one prime
mover, power converter, or energy storage device.
(3)
Essential loads, after failure of—
(i)
Any one engine, on rotorcraft with two engines; and
(ii)
Any two engines, on rotorcraft with three or more
engines.
(f)
In determining compliance with paragraphs (e)(2) and
(3) 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 for controlled flight
need not be considered for the
two-engine-inoperative condition on rotorcraft with
three or more engines.
(g)
In showing compliance with paragraphs (a) and (b) of
this section with regard to the electrical system
and to equipment design and installation, critical
environmental conditions must be considered. For
electrical generation, distribution, and utilization
equipment required by or used in complying with this
subchapter, except equipment covered by Technical
Standard Orders containing environmental test
procedures, the ability to provide continuous, safe
service under foreseeable environmental conditions
may be shown by environmental tests, design
analysis, or reference to previous comparable
service experience on other aircraft.
(h)
In showing compliance with paragraphs (a) and (b) of
this section, the effects of lightning strikes on
the rotorcraft must be considered.
(a)
Except as provided in paragraph (d) of this section,
each electrical and electronic system that performs
a function whose failure would prevent the continued
safe flight and landing of the rotorcraft must be
designed and installed so that—
(1)
The function is not adversely affected during and
after the time the rotorcraft is exposed to HIRF
environment I, as described in appendix E to this
part;
(2)
The system automatically recovers normal operation
of that function, in a timely manner, after the
rotorcraft is exposed to HIRF environment I, as
described in appendix E to this part, unless this
conflicts with other operational or functional
requirements of that system;
(3)
The system is not adversely affected during and
after the time the rotorcraft is exposed to HIRF
environment II, as described in appendix E to this
part; and
(4)
Each function required during operation under visual
flight rules is not adversely affected during and
after the time the rotorcraft is exposed to HIRF
environment III, as described in appendix E to this
part.
(b)
Each electrical and electronic system that performs
a function whose failure would significantly reduce
the capability of the rotorcraft or the ability of
the flight crew to respond to an adverse operating
condition must be designed and installed so the
system is not adversely affected when the equipment
providing these functions is exposed to equipment
HIRF test level 1 or 2, as described in appendix E
to this part.
(c)
Each electrical and electronic system that performs
such a function whose failure would reduce the
capability of the rotorcraft or the ability of the
flight crew to respond to an adverse operating
condition must be designed and installed so the
system is not adversely affected when the equipment
providing these functions is exposed to equipment
HIRF test level 3, as described in appendix E to
this part.
(d)
An electrical or electronic system that performs a
function whose failure would prevent the continued
safe flight and landing of a rotorcraft may be
designed and installed without meeting the
provisions of paragraph (a) provided—
(1)
The system has previously been shown to comply with
special conditions for HIRF, prescribed under 21.16;
(2)
The HIRF immunity characteristics of the system have
not changed since compliance with the special
conditions was demonstrated; and
(3)
The data used to demonstrate compliance with the
special conditions is provided.
29.1321 Arrangement and visibility.
(a)
Each flight, navigation, and powerplant instrument
for use by any pilot must be easily visible to him
from his station with the minimum practicable
deviation from his normal position and line of
vision when he is looking forward along the flight
path.
(b)
Each instrument necessary for safe operation,
including the airspeed indicator, gyroscopic
direction indicator, gyroscopic bank-and-pitch
indicator, slip-skid indicator, altimeter,
rate-of-climb indicator, rotor tachometers, and the
indicator most representative of engine power, must
be grouped and centered as nearly as practicable
about the vertical plane of the pilot's forward
vision. In addition, for rotorcraft approved for IFR
flight—
(1)
The instrument that most effectively indicates
attitude must be on the panel in the top center
position;
(2)
The instrument that most effectively indicates
direction of flight must be adjacent to and directly
below the attitude instrument;
(3)
The instrument that most effectively indicates
airspeed must be adjacent to and to the left of the
attitude instrument; and
(4)
The instrument that most effectively indicates
altitude or is most frequently utilized in control
of altitude must be adjacent to and to the right of
the attitude instrument.
(c)
Other required powerplant instruments must be
closely grouped on the instrument panel.
(d)
Identical powerplant instruments for the engines
must be located so as to prevent any confusion as to
which engine each instrument relates.
(e)
Each powerplant instrument vital to safe operation
must be plainly visible to appropriate crewmembers.
(f)
Instrument panel vibration may not damage, or impair
the readability or accuracy of, any instrument.
(g)
If a visual indicator is provided to indicate
malfunction of an instrument, it must be effective
under all probable cockpit lighting conditions.
If
warning, caution or advisory lights are installed in
the cockpit they must, unless otherwise approved by
the Administrator, be—
(a)
Red, for warning lights (lights indicating a hazard
which may require immediate corrective action);
(b)
Amber, for caution lights (lights indicating the
possible need for future corrective action);
(c)
Green, for safe operation lights; and
(d)
Any other color, including white, for lights not
described in paragraphs (a) through (c) of this
section, provided the color differs sufficiently
from the colors prescribed in paragraphs (a) through
(c) of this section to avoid possible confusion.
For
each airspeed indicating system, the following
apply:
(a)
Each airspeed indicating instrument must be
calibrated to indicate true airspeed (at sea level
with a standard atmosphere) with a minimum
practicable instrument calibration error when the
corresponding pitot and static pressures are
applied.
(b)
Each system must be calibrated to determine system
error excluding airspeed instrument error. This
calibration must be determined—
(1)
In level flight at speeds of 20 knots and greater,
and over an appropriate range of speeds for flight
conditions of climb and autorotation; and
(2)
During takeoff, with repeatable and readable
indications that ensure—
(i)
Consistent realization of the field lengths
specified in the Rotorcraft Flight Manual; and
(ii)
Avoidance of the critical areas of the
height-velocity envelope as established under 29.87.
(c)
For Category A rotorcraft—
(1)
The indication must allow consistent definition of
the takeoff decision point; and
(2)
The system error, excluding the airspeed instrument
calibration error, may not exceed—
(i)
Three percent or 5 knots, whichever is greater, in
level flight at speeds above 80 percent of takeoff
safety speed; and
(ii)
Ten knots in climb at speeds from 10 knots below
takeoff safety speed to 10 knots above VY.
(d)
For Category B rotorcraft, the system error,
excluding the airspeed instrument calibration error,
may not exceed 3 percent or 5 knots, whichever is
greater, in level flight at speeds above 80 percent
of the climbout speed attained at 50 feet when
complying with 29.63.
(e)
Each system must be arranged, so for as practicable,
to prevent malfunction or serious error due to the
entry of moisture, dirt, or other substances.
(f)
Each system must have a heated pitot tube or an
equivalent means of preventing malfunction due to
icing.
(a)
Each instrument with static air case connections
must be vented to the outside atmosphere through an
appropriate piping system.
(b)
Each vent must be located where its orifices are
least affected by airflow variation, moisture, or
foreign matter.
(c)
Each static pressure port must be designed and
located in such manner that the correlation between
air pressure in the static pressure system and true
ambient atmospheric static pressure is not altered
when the rotorcraft encounters icing conditions. An
anti-icing means or an alternate source of static
pressure may be used in showing compliance with this
requirement. If the reading of the altimeter, when
on the alternate static pressure system, differs
from the reading of altimeter when on the primary
static system by more than 50 feet, a correction
card must be provided for the alternate static
system.
(d)
Except for the vent into the atmosphere, each system
must be airtight.
(e)
Each pressure altimeter must be approved and
calibrated to indicate pressure altitude in a
standard atmosphere with a minimum practicable
calibration error when the corresponding static
pressures are applied.
(f)
Each system must be designed and installed so that
an error in indicated pressure altitude, at sea
level, with a standard atmosphere, excluding
instrument calibration error, does not result in an
error of more than ±30 feet per 100 knots speed.
However, the error need not be less than ±30 feet.
(g)
Except as provided in paragraph (h) of this section,
if the static pressure system incorporates both a
primary and an alternate static pressure source, the
means for selecting one or the other source must be
designed so that—
(1)
When either source is selected, the other is blocked
off; and
(2)
Both sources cannot be blocked off simultaneously.
(h)
For unpressurized rotorcraft, paragraph (g)(1) of
this section does not apply if it can be
demonstrated that the static pressure system
calibration, when either static pressure source is
selected, is not changed by the other static
pressure source being open or blocked.
(a)
Each magnetic direction indicator must be installed
so that its accuracy is not excessively affected by
the rotorcraft's vibration or magnetic fields.
(b)
The compensated installation may not have a
deviation, in level flight, greater than 10 degrees
on any heading.
(a)
Each automatic pilot system must be designed so that
the automatic pilot can—
(1)
Be sufficiently overpowered by one pilot to allow
control of the rotorcraft; and
(2)
Be readily and positively disengaged by each pilot
to prevent it from interfering with the control of
the rotorcraft.
(b)
Unless there is automatic synchronization, each
system must have a means to readily indicate to the
pilot the alignment of the actuating device in
relation to the control system it operates.
(c)
Each manually operated control for the system's
operation must be readily accessible to the pilots.
(d)
The system must be designed and adjusted so that,
within the range of adjustment available to the
pilot, it cannot produce hazardous loads on the
rotorcraft, or create hazardous deviations in the
flight path, under any flight condition appropriate
to its use, either during normal operation or in the
event of a malfunction, assuming that corrective
action begins within a reasonable period of time.
(e)
If the automatic pilot integrates signals from
auxiliary controls or furnishes signals for
operation of other equipment, there must be positive
interlocks and sequencing of engagement to prevent
improper operation.
(f)
If the automatic pilot system can be coupled to
airborne navigation equipment, means must be
provided to indicate to the pilots the current mode
of operation. Selector switch position is not
acceptable as a means of indication.
For
category A rotorcraft—
(a)
Each required flight instrument using a power supply
must have—
(1)
Two independent sources of power;
(2)
A means of selecting either power source; and
(3)
A visual means integral with each instrument to
indicate when the power adequate to sustain proper
instrument performance is not being supplied. The
power must be measured at or near the point where it
enters the instrument. For electrical instruments,
the power is considered to be adequate when the
voltage is within the approved limits; and
(b)
The installation and power supply system must be
such that failure of any flight instrument connected
to one source, or of the energy supply from one
source, or a fault in any part of the power
distribution system does not interfere with the
proper supply of energy from any other source.
For
systems that operate the required flight instruments
which are located at each pilot's station, the
following apply:
(a)
Only the required flight instruments for the first
pilot may be connected to that operating system.
(b)
The equipment, systems, and installations must be
designed so that one display of the information
essential to the safety of flight which is provided
by the flight instruments remains available to a
pilot, without additional crewmember action, after
any single failure or combination of failures that
are not shown to be extremely improbable.
(c)
Additional instruments, systems, or equipment may
not be connected to the operating system for a
second pilot unless provisions are made to ensure
the continued normal functioning of the required
flight instruments in the event of any malfunction
of the additional instruments, systems, or equipment
which is not shown to be extremely improbable.
If a
flight director system is installed, means must be
provided to indicate to the flight crew its current
mode of operation. Selector switch position is not
acceptable as a means of indication.
(a)
Instruments and instrument lines. (1) Each
powerplant and auxiliary power unit instrument line
must meet the requirements of 29.993 and 29.1183.
(2)
Each line carrying flammable fluids under pressure
must—
(i)
Have restricting orifices or other safety devices at
the source of pressure to prevent the escape of
excessive fluid if the line fails; and
(ii)
Be installed and located so that the escape of
fluids would not create a hazard.
(3)
Each powerplant and auxiliary power unit instrument
that utilizes flammable fluids must be installed and
located so that the escape of fluid would not create
a hazard.
(b)
Fuel quantity indicator. There must be means
to indicate to the flight crew members the quantity,
in gallons or equivalent units, of usable fuel in
each tank during flight. In addition—
(1)
Each fuel quantity indicator must be calibrated to
read “zero” during level flight when the quantity of
fuel remaining in the tank is equal to the unusable
fuel supply determined under 29.959;
(2)
When two or more tanks are closely interconnected by
a gravity feed system and vented, and when it is
impossible to feed from each tank separately, at
least one fuel quantity indicator must be installed;
(3)
Tanks with interconnected outlets and airspaces may
be treated as one tank and need not have separate
indicators; and
(4)
Each exposed sight gauge used as a fuel quantity
indicator must be protected against damage.
(c)
Fuel flow meter system. If a fuel flow meter
system is installed, each metering component must
have a means for bypassing the fuel supply if
malfunction of that component severely restricts
fuel flow.
(d)
Oil quantity indicator. There must be a stick
gauge or equivalent means to indicate the quantity
of oil—
(1)
In each tank; and
(2)
In each transmission gearbox.
(e)
Rotor drive system transmissions and gearboxes
utilizing ferromagnetic materials must be equipped
with chip detectors designed to indicate the
presence of ferromagnetic particles resulting from
damage or excessive wear within the transmission or
gearbox. Each chip detector must—
(1)
Be designed to provide a signal to the indicator
required by 29.1305(a)(22); and
(2)
Be provided with a means to allow crewmembers to
check, in flight, the function of each detector
electrical circuit and signal.
(a)
Electrical system capacity. The required
generating capacity and the number and kind of power
sources must—
(1)
Be determined by an electrical load analysis; and
(2)
Meet the requirements of 29.1309.
(b)
Generating system. The generating system
includes electrical power sources, main power
busses, transmission cables, and associated control,
regulation, and protective devices. It must be
designed so that—
(1)
Power sources function properly when independent and
when connected in combination;
(2)
No failure or malfunction of any power source can
create a hazard or impair the ability of remaining
sources to supply essential loads;
(3)
The system voltage and frequency (as applicable) at
the terminals of essential load equipment can be
maintained within the limits for which the equipment
is designed, during any probable operating
condition;
(4)
System transients due to switching, fault clearing,
or other causes do not make essential loads
inoperative, and do not cause a smoke or fire
hazard;
(5)
There are means accessible in flight to appropriate
crewmembers for the individual and collective
disconnection of the electrical power sources from
the main bus; and
(6)
There are means to indicate to appropriate
crewmembers the generating system quantities
essential for the safe operation of the system, such
as the voltage and current supplied by each
generator.
(c)
External power. If provisions are made for
connecting external power to the rotorcraft, and
that external power can be electrically connected to
equipment other than that used for engine starting,
means must be provided to ensure that no external
power supply having a reverse polarity, or a reverse
phase sequence, can supply power to the rotorcraft's
electrical system.
(d)
Operation with the normal electrical power
generating system inoperative.
(1)
It must be shown by analysis, tests, or both, that
the rotorcraft can be operated safely in VFR
conditions for a period of not less than 5 minutes,
with the normal electrical power generating system
(electrical power sources excluding the battery)
inoperative, with critical type fuel (from the
standpoint of flameout and restart capability), and
with the rotorcraft initially at the maximum
certificated altitude. Parts of the electrical
system may remain on if—
(i)
A single malfunction, including a wire bundle or
junction box fire, cannot result in loss of the part
turned off and the part turned on;
(ii)
The parts turned on are electrically and
mechanically isolated from the parts turned off; and
(2)
Additional requirements for Category A Rotorcraft.
(i)
Unless it can be shown that the loss of the normal
electrical power generating system is extremely
improbable, an emergency electrical power system,
independent of the normal electrical power
generating system, must be provided, with sufficient
capacity to power all systems necessary for
continued safe flight and landing.
(ii)
Failures, including junction box, control panel, or
wire bundle fires, which would result in the loss of
the normal and emergency systems, must be shown to
be extremely improbable.
(iii) Systems necessary for immediate safety must
continue to operate following the loss of the normal
electrical power generating system, without the need
for flight crew action.
(a)
Electrical equipment, controls, and wiring must be
installed so that operation of any one unit or
system of units will not adversely affect the
simultaneous operation of any other electrical unit
or system essential to safe operation.
(b)
Cables must be grouped, routed, and spaced so that
damage to essential circuits will be minimized if
there are faults in heavy current-carrying cables.
(c)
Storage batteries must be designed and installed as
follows:
(1)
Safe cell temperatures and pressures must be
maintained during any probable charging and
discharging condition. No uncontrolled increase in
cell temperature may result when the battery is
recharged (after previous complete discharge)—
(i)
At maximum regulated voltage or power;
(ii)
During a flight of maximum duration; and
(iii) Under the most adverse cooling condition
likely in service.
(2)
Compliance with paragraph (a)(1) of this section
must be shown by test unless experience with similar
batteries and installations has shown that
maintaining safe cell temperatures and pressures
presents no problem.
(3)
No explosive or toxic gases emitted by any battery
in normal operation, or as the result of any
probable malfunction in the charging system or
battery installation, may accumulate in hazardous
quantities within the rotorcraft.
(4)
No corrosive fluids or gases that may escape from
the battery may damage surrounding structures or
adjacent essential equipment.
(5)
Each nickel cadmium battery installation capable of
being used to start an engine or auxiliary power
unit must have provisions to prevent any hazardous
effect on structure or essential systems that may be
caused by the maximum amount of heat the battery can
generate during a short circuit of the battery or of
its individual cells.
(6)
Nickel cadmium battery installations capable of
being used to start an engine or auxiliary power
unit must have—
(i)
A system to control the charging rate of the battery
automatically so as to prevent battery overheating;
(ii)
A battery temperature sensing and over-temperature
warning system with a means for disconnecting the
battery from its charging source in the event of an
over-temperature condition; or
(iii) A battery failure sensing and warning system
with a means for disconnecting the battery from its
charging source in the event of battery failure.
(a)
The distribution system includes the distribution
busses, their associated feeders, and each control
and protective device.
(b)
If two independent sources of electrical power for
particular equipment or systems are required by this
chapter, in the event of the failure of one power
source for such equipment or system, another power
source (including its separate feeder) must be
provided automatically or be manually selectable to
maintain equipment or system operation.
(a)
Automatic protective devices must be used to
minimize distress to the electrical system and
hazard to the rotorcraft system and hazard to the
rotorcraft in the event of wiring faults or serious
malfunction of the system or connected equipment.
(b)
The protective and control devices in the generating
system must be designed to de-energize and
disconnect faulty power sources and power
transmission equipment from their associated buses
with sufficient rapidity to provide protection from
hazardous overvoltage and other malfunctioning.
(c)
Each resettable circuit protective device must be
designed so that, when an overload or circuit fault
exists, it will open the circuit regardless of the
position of the operating control.
(d)
If the ability to reset a circuit breaker or replace
a fuse is essential to safety in flight, that
circuit breaker or fuse must be located and
identified so that it can be readily reset or
replaced in flight.
(e)
Each essential load must have individual circuit
protection. However, individual protection for each
circuit in an essential load system (such as each
position light circuit in a system) is not required.
(f)
If fuses are used, there must be spare fuses for use
in flight equal to at least 50 percent of the number
of fuses of each rating required for complete
circuit protection.
(g)
Automatic reset circuit breakers may be used as
integral protectors for electrical equipment
provided there is circuit protection for the cable
supplying power to the equipment.
(a)
Components of the electrical system must meet the
applicable fire and smoke protection provisions of
29.831 and 29.863.
(b)
Electrical cables, terminals, and equipment, in
designated fire zones, and that are used in
emergency procedures, must be at least fire
resistant.
(c)
Insulation on electrical wire and cable installed in
the rotorcraft must be self-extinguishing when
tested in accordance with Appendix F, Part I(a)(3),
of part 25 of this chapter.
(a)
When laboratory tests of the electrical system are
conducted—
(1)
The tests must be performed on a mock-up using the
same generating equipment used in the rotorcraft;
(2)
The equipment must simulate the electrical
characteristics of the distribution wiring and
connected loads to the extent necessary for valid
test results; and
(3)
Laboratory generator drives must simulate the prime
movers on the rotorcraft with respect to their
reaction to generator loading, including loading due
to faults.
(b)
For each flight condition that cannot be simulated
adequately in the laboratory or by ground tests on
the rotorcraft, flight tests must be made.
The
instrument lights must—
(a)
Make each instrument, switch, and other device for
which they are provided easily readable; and
(b)
Be installed so that—
(1)
Their direct rays are shielded from the pilot's
eyes; and
(2)
No objectionable reflections are visible to the
pilot.
(a)
Each required landing or hovering light must be
approved.
(b)
Each landing light must be installed so that—
(1)
No objectionable glare is visible to the pilot;
(2)
The pilot is not adversely affected by halation; and
(3)
It provides enough light for night operation,
including hovering and landing.
(c)
At least one separate switch must be provided, as
applicable—
(1)
For each separately installed landing light; and
(2)
For each group of landing lights installed at a
common location.
(a)
General. Each part of each position light
system must meet the applicable requirements of this
section and each system as a whole must meet the
requirements of 29.1387 through 29.1397.
(b)
Forward position lights. Forward position
lights must consist of a red and a green light
spaced laterally as far apart as practicable and
installed forward on the rotorcraft so that, with
the rotorcraft in the normal flying position, the
red light is on the left side, and the green light
is on the right side. Each light must be approved.
(c)
Rear position light. The rear position light
must be a white light mounted as far after as
practicable, and must be approved.
(d)
Circuit. The two forward position lights and
the rear position light must make a single circuit.
(e)
Light covers and color filters. Each light
cover or color filter must be at least flame
resistant and may not change color or shape or lose
any appreciable light transmission during normal
use.
(a)
Except as provided in paragraph (e) of this section,
each forward and rear position light must, as
installed, show unbroken light within the dihedral
angles described in this section.
(b)
Dihedral angle L (left) is formed by two
intersecting vertical planes, the first parallel to
the longitudinal axis of the rotorcraft, and the
other at 110 degrees to the left of the first, as
viewed when looking forward along the longitudinal
axis.
(c)
Dihedral angle R (right) is formed by two
intersecting vertical planes, the first parallel to
the longitudinal axis of the rotorcraft, and the
other at 110 degrees to the right of the first, as
viewed when looking forward along the longitudinal
axis.
(d)
Dihedral angle A (aft) is formed by two
intersecting vertical planes making angles of 70
degrees to the right and to the left, respectively,
to a vertical plane passing through the longitudinal
axis, as viewed when looking aft along the
longitudinal axis.
(e)
If the rear position light, when mounted as far aft
as practicable in accordance with 29.1385(c), cannot
show unbroken light within dihedral angle A (as
defined in paragraph (d) of this section), a solid
angle or angles of obstructed visibility totaling
not more than 0.04 steradians is allowable within
that dihedral angle, if such solid angle is within a
cone whose apex is at the rear position light and
whose elements make an angle of 30° with a vertical
line passing through the rear position light.
(a)
General. The intensities prescribed in this
section must be provided by new equipment with light
covers and color filters in place. Intensities must
be determined with the light source operating at a
steady value equal to the average luminous output of
the source at the normal operating voltage of the
rotorcraft. The light distribution and intensity of
each position light must meet the requirements of
paragraph (b) of this section.
(b)
Forward and rear position lights. The light
distribution and intensities of forward and rear
position lights must be expressed in terms of
minimum intensities in the horizontal plane, minimum
intensities in any vertical plane, and maximum
intensities in overlapping beams, within dihedral
angles, L, R, and A, and must meet the
following requirements:
(1)
Intensities in the horizontal plane. Each
intensity in the horizontal plane (the plane
containing the longitudinal axis of the rotorcraft
and perpendicular to the plane of symmetry of the
rotorcraft), must equal or exceed the values in
29.1391.
(2)
Intensities in any vertical plane. Each
intensity in any vertical plane (the plane
perpendicular to the horizontal plane) must equal or
exceed the appropriate value in 29.1393 where I
is the minimum intensity prescribed in 29.1391
for the corresponding angles in the horizontal
plane.
(3)
Intensities in overlaps between adjacent signals.
No intensity in any overlap between adjacent
signals may exceed the values in 29.1395, except
that higher intensities in overlaps may be used with
the use of main beam intensities substantially
greater than the minima specified in 29.1391 and
29.1393 if the overlap intensities in relation to
the main beam intensities do not adversely affect
signal clarity.
29.1391 Minimum
intensities in the horizontal plane of forward and
rear position lights.
Each
position light intensity must equal or exceed the
applicable values in the following table:
|
Dihedral angle (light included) |
Angle from right or left of longitudinal
axis, measured from dead ahead |
Intensity (candles) |
|
L and R (forward red and
green) |
0° to 10°
10° to 20°
20° to 110° |
40
30
5 |
|
A (rear white) |
110° to 180° |
20 |
Each
position light intensity must equal or exceed the
applicable values in the following table:
|
Angle above or below the horizontal plane |
Intensity, I |
|
0° |
1.00 |
|
0° to 5° |
.90 |
|
5° to 10° |
.80 |
|
10° to 15° |
.70 |
|
15° to 20° |
.50 |
|
20° to 30° |
.30 |
|
30° to 40° |
.10 |
|
40° to 90° |
.05 |
No
position light intensity may exceed the applicable
values in the following table, except as provided in
29.1389(b)(3).
|
Overlaps |
Maximum intensity |
|
Area A (candles) |
Area B (candles) |
|
Green in dihedral angle L |
10 |
1 |
|
Red in dihedral angle R |
10 |
1 |
|
Green in dihedral angle A |
5 |
1 |
|
Red in dihedral angle A |
5 |
1 |
|
Rear white in dihedral angle L |
5 |
1 |
|
Rear white in dihedral angle R |
5 |
1 |
Where—
(a)
Area A includes all directions in the adjacent
dihedral angle that pass through the light source
and intersect the common boundary plane at more than
10 degrees but less than 20 degrees; and
(b)
Area B includes all directions in the adjacent
dihedral angle that pass through the light source
and intersect the common boundary plane at more than
20 degrees.
Each
position light color must have the applicable
International Commission on Illumination
chromaticity coordinates as follows:
(a)
Aviation red—
y is not
greater than 0.335; and
z is not
greater than 0.002.
(b)
Aviation green—
x is not
greater than 0.440−0.320 y ;
x is not
greater than y −0.170; and
y is not less
than 0.390−0.170 x .
(c)
Aviation white—
x is not less
than 0.300 and not greater than 0.540;
y is not less
than x −0.040 or y c−0.010, whichever
is the smaller; and
y is not
greater than x +0.020 nor 0.636−0.400 x
;
Where Y eis the y coordinate of the
Planckian radiator for the value of x
considered.
(a)
Each riding light required for water operation must
be installed so that it can—
(1)
Show a white light for at least two miles at night
under clear atmospheric conditions; and
(2)
Show a maximum practicable unbroken light with the
rotorcraft on the water.
(b)
Externally hung lights may be used.
29.1401 Anti-collision light system.
(a)
General. If certification for night operation
is requested, the rotorcraft must have an
anti-collision light system that—
(1)
Consists of one or more approved anti-collision
lights located so that their emitted light will not
impair the crew's vision or detract from the
conspicuity of the position lights; and
(2)
Meets the requirements of paragraphs (b) through (f)
of this section.
(b)
Field of coverage. The system must consist of
enough lights to illuminate the vital areas around
the rotorcraft, considering the physical
configuration and flight characteristics of the
rotorcraft. The field of coverage must extend in
each direction within at least 30 degrees above and
30 degrees below the horizontal plane of the
rotorcraft, except that there may be solid angles of
obstructed visibility totaling not more than 0.5
steradians.
(c)
Flashing characteristics. The arrangement of
the system, that is, the number of light sources,
beam width, speed of rotation, and other
characteristics, must give an effective flash
frequency of not less than 40, nor more than 100,
cycles per minute. The effective flash frequency is
the frequency at which the rotorcraft's complete
anti-collision light system is observed from a
distance, and applies to each sector of light
including any overlaps that exist when the system
consists of more than one light source. In overlaps,
flash frequencies may exceed 100, but not 180,
cycles per minute.
(d)
Color. Each anti-collision light must be
aviation red and must meet the applicable
requirements of 29.1397.
(e)
Light intensity. The minimum light
intensities in any vertical plane, measured with the
red filter (if used) and expressed in terms of
“effective” intensities must meet the requirements
of paragraph (f) of this section. The following
relation must be assumed:
where:
I e=effective
intensity (candles).
I(t)
=instantaneous intensity as a function of time.
t 2−
t l=flash time interval (seconds).
Normally, the maximum value of effective intensity
is obtained when t 2and t 1are chosen
so that the effective intensity is equal to the
instantaneous intensity at t 2and
t 1.
(f)
Minimum effective intensities for anti-collision
light. Each anti-collision light effective
intensity must equal or exceed the applicable values
in the following table:
|
Angle above or below the horizontal plane |
Effective intensity (candles) |
|
0° to 5° |
150 |
|
5° to 10° |
90 |
|
10° to 20° |
30 |
|
20° to 30° |
15 |
(a)
Accessibility. Required safety equipment to
be used by the crew in an emergency, such as
automatic liferaft releases, must be readily
accessible.
(b)
Stowage provisions. Stowage provisions for
required emergency equipment must be furnished and
must—
(1)
Be arranged so that the equipment is directly
accessible and its location is obvious; and
(2)
Protect the safety equipment from inadvertent
damage.
(c) Emergency exit descent device.
The
stowage provisions for the emergency exit descent
device required by 29.809(f) must be at the exits
for which they are intended.
(d)
Liferafts. Liferafts must be stowed near
exits through which the rafts can be launched during
an unplanned ditching. Rafts automatically or
remotely released outside the rotorcraft must be
attached to the rotorcraft by the static line
prescribed in 29.1415.
(e)
Long-range signaling device. The stowage
provisions for the long-range signaling device
required by 29.1415 must be near an exit available
during an unplanned ditching.
(f)
Life preservers. Each life preserver must be
within easy reach of each occupant while seated.
(a)
If there are means to indicate to the passengers
when safety belts should be fastened, they must be
installed to be operated from either pilot seat.
(b)
Each safety belt must be equipped with a metal to
metal latching device.
(a)
Emergency flotation and signaling equipment required
by any operating rule of this chapter must meet the
requirements of this section.
(b)
Each liferaft and each life preserver must be
approved. In addition—
(1)
Provide not less than two rafts, of an approximately
equal rated capacity and buoyancy to accommodate the
occupants of the rotorcraft; and
(2)
Each raft must have a trailing line, and must have a
static line designed to hold the raft near the
rotorcraft but to release it if the rotorcraft
becomes totally submerged.
(c)
Approved survival equipment must be attached to each
liferaft.
(d)
There must be an approved survival type emergency
locator transmitter for use in one life raft.
(a)
To obtain certification for flight into icing
conditions, compliance with this section must be
shown.
(b)
It must be demonstrated that the rotorcraft can be
safely operated in the continuous maximum and
intermittent maximum icing conditions determined
under appendix C of this part within the rotorcraft
altitude envelope. An analysis must be performed to
establish, on the basis of the rotorcraft's
operational needs, the adequacy of the ice
protection system for the various components of the
rotorcraft.
(c)
In addition to the analysis and physical evaluation
prescribed in paragraph (b) of this section, the
effectiveness of the ice protection system and its
components must be shown by flight tests of the
rotorcraft or its components in measured natural
atmospheric icing conditions and by one or more of
the following tests as found necessary to determine
the adequacy of the ice protection system:
(1)
Laboratory dry air or simulated icing tests, or a
combination of both, of the components or models of
the components.
(2)
Flight dry air tests of the ice protection system as
a whole, or its individual components.
(3)
Flight tests of the rotorcraft or its components in
measured simulated icing conditions.
(d)
The ice protection provisions of this section are
considered to be applicable primarily to the
airframe. Powerplant installation requirements are
contained in Subpart E of this part.
(e)
A means must be identified or provided for
determining the formation of ice on critical parts
of the rotorcraft. Unless otherwise restricted, the
means must be available for nighttime as well as
daytime operation. The rotorcraft flight manual must
describe the means of determining ice formation and
must contain information necessary for safe
operation of the rotorcraft in icing conditions.
(a)
Radio communication and navigation equipment
installations must be free from hazards in
themselves, in their method of operation, and in
their effects on other components, under any
critical environmental conditions.
(b)
Radio communication and navigation equipment,
controls, and wiring must be installed so that
operation of any one unit or system of units will
not adversely affect the simultaneous operation of
any other radio or electronic unit, or system of
units, required by this chapter.
(a)
There must be means, in addition to the normal
pressure relief, to automatically relieve the
pressure in the discharge lines from the vacuum air
pump when the delivery temperature of the air
becomes unsafe.
(b)
Each vacuum air system line and fitting on the
discharge side of the pump that might contain
flammable vapors or fluids must meet the
requirements of 29.1183 if they are in a designated
fire zone.
(c)
Other vacuum air system components in designated
fire zones must be at least fire resistant.
29.1435 Hydraulic
systems.
(a)
Design. Each hydraulic system must be
designed as follows:
(1)
Each element of the hydraulic system must be
designed to withstand, without detrimental,
permanent deformation, any structural loads that may
be imposed simultaneously with the maximum operating
hydraulic loads.
(2)
Each element of the hydraulic system must be
designed to withstand pressures sufficiently greater
than those prescribed in paragraph (b) of this
section to show that the system will not rupture
under service conditions.
(3)
There must be means to indicate the pressure in each
main hydraulic power system.
(4)
There must be means to ensure that no pressure in
any part of the system will exceed a safe limit
above the maximum operating pressure of the system,
and to prevent excessive pressures resulting from
any fluid volumetric change in lines likely to
remain closed long enough for such a change to take
place. The possibility of detrimental transient
(surge) pressures during operation must be
considered.
(5)
Each hydraulic line, fitting, and component must be
installed and supported to prevent excessive
vibration and to withstand inertia loads. Each
element of the installation must be protected from
abrasion, corrosion, and mechanical damage.
(6)
Means for providing flexibility must be used to
connect points, in a hydraulic fluid line, between
which relative motion or differential vibration
exists.
(b)
Tests. Each element of the system must be
tested to a proof pressure of 1.5 times the maximum
pressure to which that element will be subjected in
normal operation, without failure, malfunction, or
detrimental deformation of any part of the system.
(c)
Fire protection. Each hydraulic system using
flammable hydraulic fluid must meet the applicable
requirements of 29.861, 29.1183, 29.1185, and
29.1189.
(a)
If one or more cargo or baggage compartments are to
be accessible in flight, protective breathing
equipment must be available for an appropriate
crewmember.
(b)
For protective breathing equipment required by
paragraph (a) of this section or by any operating
rule of this chapter—
(1)
That equipment must be designed to protect the crew
from smoke, carbon dioxide, and other harmful gases
while on flight deck duty;
(2)
That equipment must include—
(i)
Masks covering the eyes, nose, and mouth; or
(ii)
Masks covering the nose and mouth, plus accessory
equipment to protect the eyes; and
(3)
That equipment must supply protective oxygen of 10
minutes duration per crewmember at a pressure
altitude of 8,000 feet with a respiratory minute
volume of 30 liters per minute BTPD.
29.1457 Cockpit
voice recorders.
(a)
Each cockpit voice recorder required by the
operating rules of this chapter must be approved,
and must be installed so that it will record the
following:
(1)
Voice communications transmitted from or received in
the rotorcraft by radio.
(2)
Voice communications of flight crewmembers on the
flight deck.
(3)
Voice communications of flight crewmembers on the
flight deck, using the rotorcraft's interphone
system.
(4)
Voice or audio signals identifying navigation or
approach aids introduced into a headset or speaker.
(5)
Voice communications of flight crewmembers using the
passenger loudspeaker system, if there is such a
system, and if the fourth channel is available in
accordance with the requirements of paragraph
(c)(4)(ii) of this section.
(b)
The recording requirements of paragraph (a)(2) of
this section may be met—
(1)
By installing a cockpit-mounted area microphone,
located in the best position for recording voice
communications originating at the first and second
pilot stations and voice communications of other
crewmembers on the flight deck when directed to
those stations; or
(2)
By installing a continually energized or
voice-actuated lip microphone at the first and
second pilot stations.
The
microphone specified in this paragraph must be so
located and, if necessary, the preamplifiers and
filters of the recorder must be so adjusted or
supplemented, that the recorded communications are
intelligible when recorded under flight cockpit
noise conditions and played back. The level of
intelligibility must be approved by the
Administrator. Repeated aural or visual playback of
the record may be used in evaluating
intelligibility.
(c)
Each cockpit voice recorder must be installed so
that the part of the communication or audio signals
specified in paragraph (a) of this section obtained
from each of the following sources is recorded on a
separate channel:
(1)
For the first channel, from each microphone,
headset, or speaker used at the first pilot station.
(2)
For the second channel, from each microphone,
headset, or speaker used at the second pilot
station.
(3)
For the third channel, from the cockpit-mounted area
microphone, or the continually energized or
voice-actuated lip microphones at the first and
second pilot stations.
(4)
For the fourth channel, from—
(i)
Each microphone, headset, or speaker used at the
stations for the third and fourth crewmembers; or
(ii)
If the stations specified in paragraph (c)(4)(i) of
this section are not required or if the signal at
such a station is picked up by another channel, each
microphone on the flight deck that is used with the
passenger loudspeaker system if its signals are not
picked up by another channel.
(iii) Each microphone on the flight deck that is
used with the rotorcraft's loudspeaker system if its
signals are not picked up by another channel.
(d)
Each cockpit voice recorder must be installed so
that—
(1)
It receives its electric power from the bus that
provides the maximum reliability for operation of
the cockpit voice recorder without jeopardizing
service to essential or emergency loads;
(2)
There is an automatic means to simultaneously stop
the recorder and prevent each erasure feature from
functioning, within 10 minutes after crash impact;
and
(3)
There is an aural or visual means for preflight
checking of the recorder for proper operation.
(e)
The record container must be located and mounted to
minimize the probability of rupture of the container
as a result of crash impact and consequent heat
damage to the record from fire.
(f)
If the cockpit voice recorder has a bulk erasure
device, the installation must be designed to
minimize the probability of inadvertent operation
and actuation of the device during crash impact.
(g)
Each recorder container must be either bright orange
or bright yellow.
(a)
Each flight recorder required by the operating rules
of Subchapter G of this chapter must be installed so
that:
(1)
It is supplied with airspeed, altitude, and
directional data obtained from sources that meet the
accuracy requirements of 29.1323, 29.1325, and
29.1327 of this part, as applicable;
(2)
The vertical acceleration sensor is rigidly
attached, and located longitudinally within the
approved center of gravity limits of the rotorcraft;
(3)
It receives its electrical power from the bus that
provides the maximum reliability for operation of
the flight recorder without jeopardizing service to
essential or emergency loads;
(4)
There is an aural or visual means for perflight
checking of the recorder for proper recording of
data in the storage medium; and
(5)
Except for recorders powered solely by the
engine-drive electrical generator system, there is
an automatic means to simultaneously stop a recorder
that has a data erasure feature and prevent each
erasure feature from functioning, within 10 minutes
after any crash impact.
(b)
Each nonejectable recorder container must be located
and mounted so as to minimize the probability of
container rupture resulting from crash impact and
subsequent damage to the record from fire.
(c)
A correlation must be established between the flight
recorder readings of airspeed, altitude, and heading
and the corresponding readings (taking into account
correction factors) of the first pilot's
instruments. This correlation must cover the
airspeed range over which the aircraft is to be
operated, the range of altitude to which the
aircraft is limited, and 360 degrees of heading.
Correlation may be established on the ground as
appropriate.
(d)
Each recorder container must:
(1)
Be either bright orange or bright yellow;
(2)
Have a reflective tape affixed to its external
surface to facilitate its location under water; and
(3)
Have an underwater locating device, when required by
the operating rules of this chapter, on or adjacent
to the container which is secured in such a manner
that it is not likely to be separated during crash
impact.
(a)
Equipment containing high energy rotors must meet
paragraph (b), (c), or (d) of this section.
(b)
High energy rotors contained in equipment must be
able to withstand damage caused by malfunctions,
vibration, abnormal speeds, and abnormal
temperatures. In addition—
(1)
Auxiliary rotor cases must be able to contain damage
caused by the failure of high energy rotor blades;
and
(2)
Equipment control devices, systems, and
instrumentation must reasonably ensure that no
operating limitations affecting the integrity of
high energy rotors will be exceeded in service.
(c)
It must be shown by test that equipment containing
high energy rotors can contain any failure of a high
energy rotor that occurs at the highest speed
obtainable with the normal speed control devices
inoperative.
(d)
Equipment containing high energy rotors must be
located where rotor failure will neither endanger
the occupants nor adversely affect continued safe
flight.
(a)
Each operating limitation specified in 29.1503
through 29.1525 and other limitations and
information necessary for safe operation must be
established.
(b)
The operating limitations and other information
necessary for safe operation must be made available
to the crewmembers as prescribed in 29.1541 through
29.1589.
(a)
An operating speed range must be established.
(b)
When airspeed limitations are a function of weight,
weight distribution, altitude, rotor speed, power,
or other factors, airspeed limitations corresponding
with the critical combinations of these factors must
be established.
(a)
The never-exceed speed,VNE,must be established so
that it is—
(1)
Not less than 40 knots (CAS); and
(2)
Not more than the lesser of—
(i)
0.9 times the maximum forward speeds established
under 29.309;
(ii)
0.9 times the maximum speed shown under 29.251 and
29.629; or
(iii) 0.9 times the maximum speed substantiated for
advancing blade tip mach number effects under
critical altitude conditions.
(b)
VNEmay vary with altitude, r.p.m., temperature, and
weight, if—
(1)
No more than two of these variables (or no more than
two instruments integrating more than one of these
variables) are used at one time; and
(2)
The ranges of these variables (or of the indications
on instruments integrating more than one of these
variables) are large enough to allow an
operationally practical and safe variation of VNE.
(c)
For helicopters, a stabilized power-offVNEdenoted
asVNE(power-off) may be established at a speed less
thanVNEestablished pursuant to paragraph (a) of this
section, if the following conditions are met:
(1)VNE(power-off) is not less than a speed midway
between the power-onVNEand the speed used in meeting
the requirements of—
(i)
29.67(a)(3) for Category A helicopters;
(ii)
29.65(a) for Category B helicopters, except
multi-engine helicopters meeting the requirements of
29.67(b); and
(iii) 29.67(b) for multi-engine Category B
helicopters meeting the requirements of 29.67(b).
(2)VNE(power-off) is—
(i)
A constant airspeed;
(ii)
A constant amount less than power-on VNE; or
(iii) A constant airspeed for a portion of the
altitude range for which certification is requested,
and a constant amount less than power-on VNE for the
remainder of the altitude range.
(a)
Maximum power-off (autorotation). The maximum
power-off rotor speed must be established so that it
does not exceed 95 percent of the lesser of—
(1)
The maximum design r.p.m. determined under
29.309(b); and
(2)
The maximum r.p.m. shown during the type tests.
(b)
Minimum power-off. The minimum power-off
rotor speed must be established so that it is not
less than 105 percent of the greater of—
(1)
The minimum shown during the type tests; and
(2)
The minimum determined by design substantiation.
(c)
Minimum power-on. The minimum power-on rotor
speed must be established so that it is—
(1)
Not less than the greater of—
(i)
The minimum shown during the type tests; and
(ii)
The minimum determined by design substantiation; and
(2)
Not more than a value determined under 29.33 (a)(1)
and (c)(1).
For
Category a rotorcraft, if a range of heights exists
at any speed, including zero, within which it is not
possible to make a safe landing following power
failure, the range of heights and its variation with
forward speed must be established, together with any
other pertinent information, such as the kind of
landing surface.
The
weight and center of gravity limitations determined
under 29.25 and 29.27, respectively, must be
established as operating limitations.
(a)
General. The powerplant limitations
prescribed in this section must be established so
that they do not exceed the corresponding limits for
which the engines are type certificated.
(b)
Takeoff operation. The powerplant takeoff
operation must be limited by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value shown during the type tests;
(2)
The maximum allowable manifold pressure (for
reciprocating engines);
(3)
The maximum allowable turbine inlet or turbine
outlet gas temperature (for turbine engines);
(4)
The maximum allowable power or torque for each
engine, considering the power input limitations of
the transmission with all engines operating;
(5)
The maximum allowable power or torque for each
engine considering the power input limitations of
the transmission with one engine inoperative;
(6)
The time limit for the use of the power
corresponding to the limitations established in
paragraphs (b)(1) through (5) of this section; and
(7)
If the time limit established in paragraph (b)(6) of
this section exceeds 2 minutes—
(i)
The maximum allowable cylinder head or coolant
outlet temperature (for reciprocating engines); and
(ii)
The maximum allowable engine and transmission oil
temperatures.
(c)
Continuous operation. The continuous
operation must be limited by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value shown during the type tests;
(2)
The minimum rotational speed shown under the rotor
speed requirements in 29.1509(c).
(3)
The maximum allowable manifold pressure (for
reciprocating engines);
(4)
The maximum allowable turbine inlet or turbine
outlet gas temperature (for turbine engines);
(5)
The maximum allowable power or torque for each
engine, considering the power input limitations of
the transmission with all engines operating;
(6)
The maximum allowable power or torque for each
engine, considering the power input limitations of
the transmission with one engine inoperative; and
(7)
The maximum allowable temperatures for—
(i)
The cylinder head or coolant outlet (for
reciprocating engines);
(ii)
The engine oil; and
(iii) The transmission oil.
(d)
Fuel grade or designation. The minimum fuel
grade (for reciprocating engines) or fuel
designation (for turbine engines) must be
established so that it is not less than that
required for the operation of the engines within the
limitations in paragraphs (b) and (c) of this
section.
(e)
Ambient temperature. Ambient temperature
limitations (including limitations for winterization
installations if applicable) must be established as
the maximum ambient atmospheric temperature at which
compliance with the cooling provisions of 29.1041
through 29.1049 is shown.
(f)
Two and one-half minute OEI power operation.
Unless otherwise authorized, the use of 21/2-minute
OEI power must be limited to engine failure
operation of multiengine, turbine-powered rotorcraft
for not longer than 21/2minutes for any period in
which that power is used. The use of 21/2-minute OEI
power must also be limited by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value shown during the type tests;
(2)
The maximum allowable gas temperature;
(3)
The maximum allowable torque; and
(4)
The maximum allowable oil temperature.
(g)
Thirty-minute OEI power operation. Unless
otherwise authorized, the use of 30-minute OEI power
must be limited to multiengine, turbine-powered
rotorcraft for not longer than 30 minutes after
failure of an engine. The use of 30-minute OEI power
must also be limited by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value shown during the type tests;
(2)
The maximum allowable gas temperature;
(3)
The maximum allowable torque; and
(4)
The maximum allowable oil temperature.
(h)
Continuous OEI power operation. Unless
otherwise authorized, the use of continuous OEI
power must be limited to multiengine,
turbine-powered rotorcraft for continued flight
after failure of an engine. The use of continuous
OEI power must also be limited by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value shown during the type tests.
(2)
The maximum allowable gas temperature;
(3)
The maximum allowable torque; and
(4)
The maximum allowable oil temperature.
(i)
Rated 30-second OEI power operation. Rated
30-second OEI power is permitted only on
multiengine, turbine-powered rotorcraft, also
certificated for the use of rated 2-minute OEI
power, and can only be used for continued operation
of the remaining engine(s) after a failure or
precautionary shutdown of an engine. It must be
shown that following application of 30-second OEI
power, any damage will be readily detectable by the
applicable inspections and other related procedures
furnished in accordance with Section A29.4 of
appendix A of this part and Section A33.4 of
appendix A of part 33. The use of 30-second OEI
power must be limited to not more than 30 seconds
for any period in which that power is used, and by—
(1)
The maximum rotational speed which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value demonstrated during the type
tests;
(2)
The maximum allowable gas temperature; and
(3)
The maximum allowable torque.
(j)
Rated 2-minute OEI power operation. Rated
2-minute OEI power is permitted only on multiengine,
turbine-powered rotorcraft, also certificated for
the use of rated 30-second OEI power, and can only
be used for continued operation of the remaining
engine(s) after a failure or precautionary shutdown
of an engine. It must be shown that following
application of 2-minute OEI power, any damage will
be readily detectable by the applicable inspections
and other related procedures furnished in accordance
with Section A29.4 of appendix a of this part and
Section A33.4 of appendix A of part 33. The use of
2-minute OEI power must be limited to not more than
2 minutes for any period in which that power is
used, and by—
(1)
The maximum rotational speed, which may not be
greater than—
(i)
The maximum value determined by the rotor design; or
(ii)
The maximum value demonstrated during the type
tests;
(2)
The maximum allowable gas temperature; and
(3)
The maximum allowable torque.
If
an auxiliary power unit that meets the requirements
of TSO-C77 is installed in the rotorcraft, the
limitations established for that auxiliary power
unit under the TSO including the categories of
operation must be specified as operating limitations
for the rotorcraft.
The
minimum flight crew must be established so that it
is sufficient for safe operation, considering—
(a)
The workload on individual crewmembers;
(b)
The accessibility and ease of operation of necessary
controls by the appropriate crewmember; and
(c)
The kinds of operation authorized under 29.1525.
The
kinds of operations (such as VFR, IFR, day, night,
or icing) for which the rotorcraft is approved are
established by demonstrated compliance with the
applicable certification requirements and by the
installed equipment.
The
maximum altitude up to which operation is allowed,
as limited by flight, structural, powerplant,
functional, or equipment characteristics, must be
established.
The
applicant must prepare Instructions for Continued
Airworthiness in accordance with appendix A to this
part that are acceptable to the Administrator. The
instructions may be incomplete at type certification
if a program exists to ensure their completion prior
to delivery of the first rotorcraft or issuance of a
standard certificate of airworthiness, whichever
occurs later.
(a)
The rotorcraft must contain—
(1)
The markings and placards specified in 29.1545
through 29.1565; and
(2)
Any additional information, instrument markings, and
placards required for the safe operation of the
rotorcraft if it has unusual design, operating or
handling characteristics.
(b)
Each marking and placard prescribed in paragraph (a)
of this section—
(1)
Must be displayed in a conspicuous place; and
(2)
May not be easily erased, disfigured, or obscured.
For
each instrument—
(a)
When markings are on the cover glass of the
instrument there must be means to maintain the
correct alignment of the glass cover with the face
of the dial; and
(b)
Each arc and line must be wide enough, and located
to be clearly visible to the pilot.
(a)
Each airspeed indicator must be marked as specified
in paragraph (b) of this section, with the marks
located at the corresponding indicated airspeeds.
(b)
The following markings must be made:
(1)
A red radial line—
(i)
For rotorcraft other than helicopters, at VNE; and
(ii)
For helicopters, at a VNE(power-on).
(2)
A red, cross-hatched radial line at VNE(power-off)
for helicopters, if VNE(power-off) is less than
VNE(power-on).
(3)
For the caution range, a yellow arc.
(4)
For the safe operating range, a green arc.
(a)
A placard meeting the requirements of this section
must be installed on or near the magnetic direction
indicator.
(b)
The placard must show the calibration of the
instrument in level flight with the engines
operating.
(c)
The placard must state whether the calibration was
made with radio receivers on or off.
(d)
Each calibration reading must be in terms of
magnetic heading in not more than 45 degree
increments.
For
each required powerplant instrument, as appropriate
to the type of instruments—
(a)
Each maximum and, if applicable, minimum safe
operating limit must be marked with a red radial or
a red line;
(b)
Each normal operating range must be marked with a
green arc or green line, not extending beyond the
maximum and minimum safe limits;
(c)
Each takeoff and precautionary range must be marked
with a yellow arc or yellow line;
(d)
Each engine or propeller range that is restricted
because of excessive vibration stresses must be
marked with red arcs or red lines; and
(e)
Each OEI limit or approved operating range must be
marked to be clearly differentiated from the
markings of paragraphs (a) through (d) of this
section except that no marking is normally required
for the 30-second OEI limit.
Each
oil quantity indicator must be marked with enough
increments to indicate readily and accurately the
quantity of oil.
If
the unusable fuel supply for any tank exceeds one
gallon, or five percent of the tank capacity,
whichever is greater, a red arc must be marked on
its indicator extending from the calibrated zero
reading to the lowest reading obtainable in level
flight.
(a)
Each cockpit control, other than primary flight
controls or control whose function is obvious, must
be plainly marked as to its function and method of
operation.
(b)
For powerplant fuel controls—
(1)
Each fuel tank selector valve control must be marked
to indicate the position corresponding to each tank
and to each existing cross feed position;
(2)
If safe operation requires the use of any tanks in a
specific sequence, that sequence must be marked on,
or adjacent to, the selector for those tanks; and
(3)
Each valve control for any engine of a multiengine
rotorcraft must be marked to indicate the position
corresponding to each engine controlled.
(c)
Usable fuel capacity must be marked as follows:
(1)
For fuel systems having no selector controls, the
usable fuel capacity of the system must be indicated
at the fuel quantity indicator.
(2)
For fuel systems having selector controls, the
usable fuel capacity available at each selector
control position must be indicated near the selector
control.
(d)
For accessory, auxiliary, and emergency controls—
(1)
Each essential visual position indicator, such as
those showing rotor pitch or landing gear position,
must be marked so that each crewmember can determine
at any time the position of the unit to which it
relates; and
(2)
Each emergency control must be red and must be
marked as to method of operation.
(e)
For rotorcraft incorporating retractable landing
gear, the maximum landing gear operating speed must
be displayed in clear view of the pilot.
(a)
Baggage and cargo compartments, and ballast
location. Each baggage and cargo compartment,
and each ballast location must have a placard
stating any limitations on contents, including
weight, that are necessary under the loading
requirements.
(b)
Seats. If the maximum allowable weight to be
carried in a seat is less than 170 pounds, a placard
stating the lesser weight must be permanently
attached to the seat structure.
(c)
Fuel and oil filler openings. The following
apply:
(1)
Fuel filler openings must be marked at or near the
filler cover with—
(i)
The word “fuel”;
(ii)
For reciprocating engine powered rotorcraft, the
minimum fuel grade;
(iii) For turbine-engine-powered rotorcraft, the
permissible fuel designations, except that if
impractical, this information may be included in the
rotorcraft flight manual, and the fuel filler may be
marked with an appropriate reference to the flight
manual; and
(iv)
For pressure fueling systems, the maximum
permissible fueling supply pressure and the maximum
permissible defueling pressure.
(2)
Oil filler openings must be marked at or near the
filler cover with the word “oil”.
(d)
Emergency exit placards. Each placard and
operating control for each emergency exit must
differ in color from the surrounding fuselage
surface as prescribed in 29.811(h)(2). A placard
must be near each emergency exit control and must
clearly indicate the location of that exit and its
method of operation.
There must be a placard in clear view of the pilot
that specifies the kinds of operations (VFR, IFR,
day, night, or icing) for which the rotorcraft is
approved.
(a)
Each safety equipment control to be operated by the
crew in emergency, such as controls for automatic
liferaft releases, must be plainly marked as to its
method of operation.
(b)
Each location, such as a locker or compartment, that
carries any fire extinguishing, signaling, or other
life saving equipment, must be so marked.
(c)
Stowage provisions for required emergency equipment
must be conspicuously marked to identify the
contents and facilitate removal of the equipment.
(d)
Each liferaft must have obviously marked operating
instructions.
(e)
Approved survival equipment must be marked for
identification and method of operation.
Each
tail rotor must be marked so that its disc is
conspicuous under normal daylight ground conditions.
(a)
Furnishing information. A Rotorcraft Flight
Manual must be furnished with each rotorcraft, and
it must contain the following:
(1)
Information required by 29.1583 through 29.1589.
(2)
Other information that is necessary for safe
operation because of design, operating, or handling
characteristics.
(b)
Approved information. Each part of the manual
listed in 29.1583 through 29.1589 that is
appropriate to the rotorcraft, must be furnished,
verified, and approved, and must be segregated,
indentified, and clearly distinguished from each
unapproved part of that manual.
(c)
[Reserved]
(d)
Table of contents. Each Rotorcraft Flight
Manual must include a table of contents if the
complexity of the manual indicates a need for it.
(a)
Airspeed and rotor limitations. Information
necessary for the marking of airspeed and rotor
limitations on or near their respective indicators
must be furnished. The significance of each
limitation and of the color coding must be
explained.
(b)
Powerplant limitations. The following
information must be furnished:
(1)
Limitations required by 29.1521.
(2)
Explanation of the limitations, when appropriate.
(3)
Information necessary for marking the instruments
required by 29.1549 through 29.1553.
(c)
Weight and loading distribution. The weight
and center of gravity limits required by 29.25 and
29.27, respectively, must be furnished. If the
variety of possible loading conditions warrants,
instructions must be included to allow ready
observance of the limitations.
(d)
Flight crew. When a flight crew of more than
one is required, the number and functions of the
minimum flight crew determined under 29.1523 must be
furnished.
(e)
Kinds of operation. Each kind of operation
for which the rotorcraft and its equipment
installations are approved must be listed.
(f)
Limiting heights. Enough information must be
furnished to allow compliance with 29.1517.
(g)
Maximum allowable wind. For Category A
rotorcraft, the maximum allowable wind for safe
operation near the ground must be furnished.
(h)
Altitude. The altitude established under
29.1527 and an explanation of the limiting factors
must be furnished.
(i)
Ambient temperature. Maximum and minimum
ambient temperature limitations must be furnished.
(a)
The parts of the manual containing operating
procedures must have information concerning any
normal and emergency procedures, and other
information necessary for safe operation, including
the applicable procedures, such as those involving
minimum speeds, to be followed if an engine fails.
(b)
For multiengine rotorcraft, information identifying
each operating condition in which the fuel system
independence prescribed in 29.953 is necessary for
safety must be furnished, together with instructions
for placing the fuel system in a configuration used
to show compliance with that section.
(c)
For helicopters for which a VNE(power-off) is
established under 29.1505(c), information must be
furnished to explain the VNE(power-off) and the
procedures for reducing airspeed to not more than
the VNE(power-off) following failure of all engines.
(d)
For each rotorcraft showing compliance with 29.1353
(c)(6)(ii) or (c)(6)(iii), the operating procedures
for disconnecting the battery from its charging
source must be furnished.
(e)
If the unusable fuel supply in any tank exceeds 5
percent of the tank capacity, or 1 gallon, whichever
is greater, information must be furnished which
indicates that when the fuel quantity indicator
reads “zero” in level flight, any fuel remaining in
the fuel tank cannot be used safely in flight.
(f)
Information on the total quantity of usable fuel for
each fuel tank must be furnished.
(g)
For Category B rotorcraft, the airspeeds and
corresponding rotor speeds for minimum rate of
descent and best glide angle as prescribed in 29.71
must be provided.
Flight manual performance information which exceeds
any operating limitation may be shown only to the
extent necessary for presentation clarity or to
determine the effects of approved optional equipment
or procedures. When data beyond operating limits are
shown, the limits must be clearly indicated. The
following must be provided:
(a)
Category A. For each category A rotorcraft,
the Rotorcraft Flight Manual must contain a summary
of the performance data, including data necessary
for the application of any operating rule of this
chapter, together with descriptions of the
conditions, such as airspeeds, under which this data
was determined, and must contain—
(1)
The indicated airspeeds corresponding with those
determined for take-off, and the procedures to be
followed if the critical engine fails during
takeoff;
(2)
The airspeed calibrations;
(3)
The techniques, associated airspeeds, and rates of
descent for autorotative landings;
(4)
The rejected takeoff distance determined under 29.62
and the takeoff distance determined under 29.61;
(5)
The landing data determined under 29.81 and 29.85;
(6)
The steady gradient of climb for each weight,
altitude, and temperature for which takeoff data are
to be scheduled, along the takeoff path determined
in the flight conditions required in 29.67(a)(1) and
(a)(2):
(i)
In the flight conditions required in 29.67(a)(1)
between the end of the takeoff distance and the
point at which the rotorcraft is 200 feet above the
takeoff surface (or 200 feet above the lowest point
of the takeoff profile for elevated heliports);
(ii)
In the flight conditions required in 29.67(a)(2)
between the points at which the rotorcraft is 200
and 1000 feet above the takeoff surface (or 200 and
1000 feet above the lowest point of the takeoff
profile for elevated heliports); and
(7)
Out-of-ground effect hover performance determined
under 29.49 and the maximum safe wind demonstrated
under the ambient conditions for data presented.
(b)
Category B. For each category B rotorcraft,
the Rotorcraft Flight Manual must contain—
(1)
The takeoff distance and the climbout speed together
with the pertinent information defining the flight
path with respect to autorotative landing if an
engine fails, including the calculated effects of
altitude and temperature;
(2)
The steady rates of climb and hovering ceiling,
together with the corresponding airspeeds and other
pertinent information, including the calculated
effects of altitude and temperature;
(3)
The landing distance, appropriate airspeed, and type
of landing surface, together with all pertinent
information that might affect this distance,
including the effects of weight, altitude, and
temperature;
(4)
The maximum safe wind for operation near the ground;
(5)
The airspeed calibrations;
(6)
The height-speed envelope except for rotorcraft
incorporating this as an operating limitation;
(7)
Glide distance as a function of altitude when
auto-rotating at the speeds and conditions for
minimum rate of descent and best glide angle, as
determined in 29.71;
(8)
Out-of-ground effect hover performance determined
under 29.49 and the maximum safe wind demonstrated
under the ambient conditions for data presented; and
(9)
Any additional performance data necessary for the
application of any operating rule in this chapter.
There must be loading instructions for each possible
loading condition between the maximum and minimum
weights determined under 29.25 that can result in a
center of gravity beyond any extreme prescribed in
29.27, assuming any probable occupant weights.
a29.1 General
(a)
This appendix specifies requirements for the
preparation of Instructions for Continued
Airworthiness as required by 29.1529.
(b)
The Instructions for Continued Airworthiness for
each rotorcraft must include the Instructions for
Continued Airworthiness for each engine and rotor
(hereinafter designated “products”), for each
appliance required by this chapter, and any required
information relating to the interface of those
appliances and products with the rotorcraft. If
Instructions for Continued Airworthiness are not
supplied by the manufacturer of an appliance or
product installed in the rotorcraft, the
Instructions for Continued Airworthiness for the
rotorcraft must include the information essential to
the continued airworthiness of the rotorcraft.
(c)
The applicant must submit to the ACAA a program to
show how changes to the Instructions for Continued
Airworthiness made by the applicant or by the
manufacturers of products and appliances installed
in the rotorcraft will be distributed.
a29.2 Format
(a)
The Instructions for Continued Airworthiness must be
in the form of a manual or manuals as appropriate
for the quantity of data to be provided.
(b)
The format of the manual or manuals must provide for
a practical arrangement.
a29.3 Content
The
contents of the manual or manuals must be prepared
in the English language. The Instructions for
Continued Airworthiness must contain the following
manuals or sections, as appropriate, and
information:
(a)
Rotorcraft maintenance manual or section. (1)
Introduction information that includes an
explanation of the rotorcraft's features and data to
the extent necessary for maintenance or preventive
maintenance.
(2)
A description of the rotorcraft and its systems and
installations including its engines, rotors, and
appliances.
(3)
Basic control and operation information describing
how the rotorcraft components and systems are
controlled and how they operate, including any
special procedures and limitations that apply.
(4)
Servicing information that covers details regarding
servicing points, capacities of tanks, reservoirs,
types of fluids to be used, pressures applicable to
the various systems, location of access panels for
inspection and servicing, locations of lubrication
points, the lubricants to be used, equipment
required for servicing, tow instructions and
limitations, mooring, jacking, and leveling
information.
(b)
Maintenance Instructions. (1) Scheduling
information for each part of the rotorcraft and its
engines, auxiliary power units, rotors, accessories,
instruments, and equipment that provides the
recommended periods at which they should be cleaned,
inspected, adjusted, tested, and lubricated, and the
degree of inspection, the applicable wear
tolerances, and work recommended at these periods.
However, the applicant may refer to an accessory,
instrument, or equipment manufacturer as the source
of this information if the applicant shows that the
item has an exceptionally high degree of complexity
requiring specialized maintenance techniques, test
equipment, or expertise. The recommended overhaul
periods and necessary cross references to the
Airworthiness Limitations section of the manual must
also be included. In addition, the applicant must
include an inspection program that includes the
frequency and extent of the inspections necessary to
provide for the continued airworthiness of the
rotorcraft.
(2)
Troubleshooting information describing probable
malfunctions, how to recognize those malfunctions,
and the remedial action for those malfunctions.
(3)
Information describing the order and method of
removing and replacing products and parts with any
necessary precautions to be taken.
(4)
Other general procedural instructions including
procedures for system testing during ground running,
symmetry checks, weighing and determining the center
of gravity, lifting and shoring, and storage
limitations.
(c)
Diagrams of structural access plates and information
needed to gain access for inspections when access
plates are not provided.
(d)
Details for the application of special inspection
techniques including radiographic and ultrasonic
testing where such processes are specified.
(e)
Information needed to apply protective treatments to
the structure after inspection.
(f)
All data relative to structural fasteners such as
identification, discard recommendations, and torque
values.
(g)
A list of special tools needed.
a29.4 Airworthiness Limitations Section
The
Instructions for Continued Airworthiness must
contain a section titled Airworthiness Limitations
that is segregated and clearly distinguishable from
the rest of the document. This section must set
forth each mandatory replacement time, structural
inspection interval, and related structural
inspection procedure approved under 29.571. If the
Instructions for Continued Airworthiness consist of
multiple documents, the section required by this
paragraph must be included in the principal manual.
This section must contain a legible statement in a
prominent location that reads: “The Airworthiness
Limitations section is ACAA approved and specifies
maintenance required under 43.16 and 91.403 of the
African Civil Aviation Regulations unless an
alternative program has been ACAA approved.”
I.
General. A transport category helicopter may
not be type certificated for operation under the
instrument flight rules (IFR) of this chapter unless
it meets the design and installation requirements
contained in this appendix.
II.
Definitions. (a) VYImeans
instrument climb speed, utilized instead of VYfor
compliance with the climb requirements for
instrument flight.
(b)
VNEImeans instrument flight never exceed
speed, utilized instead of VNEfor
compliance with maximum limit speed requirements for
instrument flight.
(c)
VMINImeans instrument flight minimum
speed, utilized in complying with minimum limit
speed requirements for instrument flight.
III.
Trim. It must be possible to trim the cyclic,
collective, and directional control forces to zero
at all approved IFR airspeeds, power settings, and
configurations appropriate to the type.
IV.
Static longitudinal stability. (a)
General. The helicopter must possess positive
static longitudinal control force stability at
critical combinations of weight and center of
gravity at the conditions specified in paragraphs IV
(b) through (f) of this appendix. The stick force
must vary with speed so that any substantial speed
change results in a stick force clearly perceptible
to the pilot. The airspeed must return to within 10
percent of the trim speed when the control force is
slowly released for each trim condition specified in
paragraphs IV (b) through (f) of this appendix.
(b)
Climb. Stability must be shown in climb
thoughout the speed range 20 knots either side of
trim with—
(1)
The helicopter trimmed at VYI;
(2)
Landing gear retracted (if retractable); and
(3)
Power required for limit climb rate (at least 1,000
fpm) at VYIor maximum continuous power,
whichever is less.
(c)
Cruise. Stability must be shown throughout
the speed range from 0.7 to 1.1 VHor VNEI,
whichever is lower, not to exceed ±20 knots from
trim with—
(1)
The helicopter trimmed and power adjusted for level
flight at 0.9 VHor 0.9 VNEI,
whichever is lower; and
(2)
Landing gear retracted (if retractable).
(d)
Slow cruise. Stability must be shown
throughout the speed range from 0.9 VMINIto
1.3 VMINIor 20 knots above trim speed,
whichever is greater, with—
(1)
The helicopter trimmed and power adjusted for level
flight at 1.1 VMINI; and
(2)
Landing gear retracted (if retractable).
(e)
Descent. Stability must be shown throughout
the speed range 20 knots either side of trim with—
(1)
The helicopter trimmed at 0.8 VH or 0.8 VNEI
(or 0.8 VLE for the landing gear
extended case), whichever is lower;
(2)
Power required for 1,000 fpm descent at trim speed;
and
(3)
Landing gear extended and retracted, if applicable.
(f)
Approach. Stability must be shown throughout
the speed range from 0.7 times the minimum
recommended approach speed to 20 knots above the
maximum recommended approach speed with—
(1)
The helicopter trimmed at the recommended approach
speed or speeds;
(2)
Landing gear extended and retracted, if applicable;
and
(3)
Power required to maintain a 3° glide path and power
required to maintain the steepest approach gradient
for which approval is requested.
V.
Static lateral-directional stability. (a)
Static directional stability must be positive
throughout the approved ranges of airspeed, power,
and vertical speed. In straight, steady sideslips up
to ±10° from trim, directional control position must
increase in approximately constant proportion to
angle of sideslip. At greater angles up to the
maximum sideslip angle appropriate to the type,
increased directional control position must produce
increased angle of sideslip.
(b)
During sideslips up to ±10° from trim throughout the
approved ranges of airspeed, power, and vertical
speed there must be no negative dihedral stability
perceptible to the pilot through lateral control
motion or force. Longitudinal cycle movement with
sideslip must not be excessive.
VI.
Dynamic stability. (a) Any oscillation having
a period of less than 5 seconds must damp to 1/2
amplitude in not more than one cycle.
(b)
Any oscillation having a period of 5 seconds or more
but less than 10 seconds must damp to 1/2 amplitude
in not more than two cycles.
(c)
Any oscillation having a period of 10 seconds or
more but less than 20 seconds must be damped.
(d)
Any oscillation having a period of 20 seconds or
more may not achieve double amplitude in less than
20 seconds.
(e)
Any a periodic response may not achieve double
amplitude in less than 9 seconds.
VII.
Stability augmentation system (SAS). (a) If a
SAS is used, the reliability of the SAS must be
related to the effects of its failure. The
occurrence of any failure condition which would
prevent continued safe flight and landing must be
extremely improbable. For any failure condition of
the SAS which is not shown to be extremely
improbable—
(1)
The helicopter must be safely controllable and
capable of prolonged instrument flight without undue
pilot effort. Additional unrelated probable failures
affecting the control system must be considered; and
(2)
The flight characteristics requirements in Subpart B
of Part 29 must be met throughout a practical flight
envelope.
(b)
The SAS must be designed so that it cannot create a
hazardous deviation in flight path or produce
hazardous loads on the helicopter during normal
operation or in the event of malfunction or failure,
assuming corrective action begins within an
appropriate period of time. Where multiple systems
are installed, subsequent malfunction conditions
must be considered in sequence unless their
occurrence is shown to be improbable.
VIII. Equipment, systems, and installation.
The basic equipment and installation must comply
with Subpart F of Part 29 through Amendment 29–14,
with the following exceptions and additions:
(a)
Flight and navigation instruments. (1) A
magnetic gyro-stabilized direction indicator instead
of the gyroscopic direction indicator required by
29.1303(h); and
(2)
A standby attitude indicator which meets the
requirements of 29.1303(g)(1) through (7), instead
of a rate-of-turn indicator required by 29.1303(g).
If standby batteries are provided, they may be
charged from the aircraft electrical system if
adequate isolation is incorporated. The system must
be designed so that the standby batteries may not be
used for engine starting.
(b)
Miscellaneous requirements. (1) Instrument
systems and other systems essential for IFR flight
that could be adversely affected by icing must be
provided with adequate ice protection whether or not
the rotorcraft is certificated for operation in
icing conditions.
(2)
There must be means in the generating system to
automatically de-energize and disconnect from the
main bus any power source developing hazardous
overvoltage.
(3)
Each required flight instrument using a power supply
(electric, vacuum, etc.) must have a visual means
integral with the instrument to indicate the
adequacy of the power being supplied.
(4)
When multiple systems performing like functions are
required, each system must be grouped, routed, and
spaced so that physical separation between systems
is provided to ensure that a single malfunction will
not adversely affect more than one system.
(5)
For systems that operate the required flight
instruments at each pilot's station—
(i)
Only the required flight instruments for the first
pilot may be connected to that operating system;
(ii)
Additional instruments, systems, or equipment may
not be connected to an operating system for a second
pilot unless provisions are made to ensure the
continued normal functioning of the required
instruments in the event of any malfunction of the
additional instruments, systems, or equipment which
is not shown to be extremely improbable;
(iii) The equipment, systems, and installations must
be designed so that one display of the information
essential to the safety of flight which is provided
by the instruments will remain available to a pilot,
without additional crew-member action, after any
single failure or combination of failures that is
not shown to be extremely improbable; and
(iv)
For single-pilot configurations, instruments which
require a static source must be provided with a
means of selecting an alternate source and that
source must be calibrated.
(6)
In determining compliance with the requirements of
29.1351(d)(2), the supply of electrical power to all
systems necessary for flight under IFR must be
included in the evaluation.
(c)
Thunderstorm lights. In addition to the
instrument lights required by 29.1381(a),
thunderstorm lights which provide high intensity
white flood lighting to the basic flight instruments
must be provided. The thunderstorm lights must be
installed to meet the requirements of 29.1381(b).
IX.
Rotorcraft Flight Manual. A Rotorcraft Flight
Manual or Rotorcraft Flight Manual IFR Supplement
must be provided and must contain—
(a)
Limitations. The approved IFR flight
envelope, the IFR flight crew composition, the
revised kinds of operation, and the steepest IFR
precision approach gradient for which the helicopter
is approved;
(b)
Procedures. Required information for proper
operation of IFR systems and the recommended
procedures in the event of stability augmentation or
electrical system failures; and
(c)
Performance. If VYI differs from VY,
climb performance at VYI and with maximum
continuous power throughout the ranges of weight,
altitude, and temperature for which approval is
requested.
(a)
Continuous maximum icing. The maximum
continuous intensity of atmospheric icing conditions
(continuous maximum icing) is defined by the
variables of the cloud liquid water content, the
mean effective diameter of the cloud droplets, the
ambient air temperature, and the interrelationship
of these three variables as shown in Figure 1 of
this appendix. The limiting icing envelope in terms
of altitude and temperature is given in Figure 2 of
this appendix. The interrelationship of cloud liquid
water content with drop diameter and altitude is
determined from Figures 1 and 2. The cloud liquid
water content for continuous maximum icing
conditions of a horizontal extent, other than 17.4
nautical miles, is determined by the value of liquid
water content of Figure 1, multiplied by the
appropriate factor from Figure 3 of this appendix.
(b)
Intermittent maximum icing. The intermittent
maximum intensity of atmospheric icing conditions
(intermittent maximum icing) is defined by the
variables of the cloud liquid water content, the
mean effective diameter of the cloud droplets, the
ambient air temperature, and the interrelationship
of these three variables as shown in Figure 4 of
this appendix. The limiting icing envelope in terms
of altitude and temperature is given in Figure 5 of
this appendix. The interrelationship of cloud liquid
water content with drop diameter and altitude is
determined from Figures 4 and 5. The cloud liquid
water content for intermittent maximum icing
conditions of a horizontal extent, other than 2.6
nautical miles, is determined by the value of cloud
liquid water content of Figure 4 multiplied by the
appropriate factor in Figure 6 of this appendix.
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(a)
The demonstration must be conducted either during
the dark of the night or during daylight with the
dark of night simulated. If the demonstration is
conducted indoors during daylight hours, it must be
conducted inside a darkened hangar having doors and
windows covered. In addition, the doors and windows
of the rotorcraft must be covered if the hangar
illumination exceeds that of a moonless night.
Illumination on the floor or ground may be used, but
it must be kept low and shielded against shining
into the rotorcraft's windows or doors.
(b)
The rotorcraft must be in a normal attitude with
landing gear extended.
(c)
Safety equipment such as mats or inverted liferafts
may be placed on the floor or ground to protect
participants. No other equipment that is not part of
the rotorcraft's emergency evacuation equipment may
be used to aid the participants in reaching the
ground.
(d)
Except as provided in paragraph (a) of this
appendix, only the rotorcraft's emergency lighting
system may provide illumination.
(e)
All emergency equipment required for the planned
operation of the rotorcraft must be installed.
(f)
Each external door and exit and each internal door
or curtain must be in the takeoff configuration.
(g)
Each crewmember must be seated in the normally
assigned seat for takeoff and must remain in that
seat until receiving the signal for commencement of
the demonstration. For compliance with this section,
each crewmember must be—
(1)
A member of a regularly scheduled line crew; or
(2)
A person having knowledge of the operation of exits
and emergency equipment.
(h)
A representative passenger load of persons in normal
health must be used as follows:
(1)
At least 25 percent must be over 50 years of age,
with at least 40 percent of these being females.
(2)
The remaining, 75 percent or less, must be 50 years
of age or younger, with at least 30 percent of these
being females.
(3)
Three life-size dolls, not included as part of the
total passenger load, must be carried by passengers
to simulate live infants 2 years old or younger,
except for a total passenger load of fewer than 44
but more than 19, one doll must be carried. A doll
is not required for a 19 or fewer passenger load.
(4)
Crewmembers, mechanics, and training personnel who
maintain or operate the rotorcraft in the normal
course of their duties may not be used as
passengers.
(i)
No passenger may be assigned a specific seat except
as the Administrator may require. Except as required
by paragraph (1) of this appendix, no employee of
the applicant may be seated next to an emergency
exit, except as allowed by the Administrator.
(j)
Seat belts and shoulder harnesses (as required) must
be fastened.
(k)
Before the start of the demonstration, approximately
one-half of the total average amount of carry-on
baggage, blankets, pillows, and other similar
articles must be distributed at several locations in
the aisles and emergency exit access ways to create
minor obstructions.
(l)
No prior indication may be given to any crewmember
or passenger of the particular exits to be used in
the demonstration.
(m)
The applicant may not practice, rehearse, or
describe the demonstration for the participants nor
may any participant have taken part in this type of
demonstration within the preceding 6 months.
(n)
A pre-take-off passenger briefing may be given. The
passengers may also be advised to follow directions
of crewmembers, but not be instructed on the
procedures to be followed in the demonstration.
(o)
If safety equipment, as allowed by paragraph (c) of
this appendix, is provided, either all passenger and
cockpit windows must be blacked out or all emergency
exits must have safety equipment to prevent
disclosure of the available emergency exits.
(p)
Not more than 50 percent of the emergency exits in
the sides of the fuselage of a rotorcraft that meet
all of the requirements applicable to the required
emergency exits for that rotorcraft may be used for
demonstration. Exits that are not to be used for the
demonstration must have the exit handle deactivated
or must be indicated by red lights, red tape, or
other acceptable means placed outside the exits to
indicate fire or other reasons why they are
unusable. The exits to be used must be
representative of all the emergency exits on the
rotorcraft and must be designated by the applicant,
subject to approval by the Administrator. If
installed, at least one floor level exit (Type I;
29.807(a)(1)) must be used as required by 29.807(c).
(q)
All evacuees must leave the rotorcraft by a means
provided as part of the rotorcraft's equipment.
(r)
Approved procedures must be fully utilized during
the demonstration.
(s)
The evacuation time period is completed when the
last occupant has evacuated the rotorcraft and is on
the ground.
This
appendix specifies the HIRF environments and
equipment HIRF test levels for electrical and
electronic systems under 29.1317. The field strength
values for the HIRF environments and laboratory
equipment HIRF test levels are expressed in
root-mean-square units measured during the peak of
the modulation cycle.
(a)
HIRF environment I is specified in the following
table:
Table I.—HIRF Environment I
|
Frequency |
Field strength
(volts/meter) |
|
Peak |
Average |
|
10 kHz–2 MHz |
50 |
50 |
|
2 MHz–30 MHz |
100 |
100 |
|
30 MHz–100 MHz |
50 |
50 |
|
100 MHz–400 MHz |
100 |
100 |
|
400 MHz–700 MHz |
700 |
50 |
|
700 MHz–1 GHz |
700 |
100 |
|
1 GHz–2 GHz |
2,000 |
200 |
|
2 GHz–6 GHz |
3,000 |
200 |
|
6 GHz–8 GHz |
1,000 |
200 |
|
8 GHz–12 GHz |
3,000 |
300 |
|
12 GHz–18 GHz |
2,000 |
200 |
|
18 GHz–40 GHz |
600 |
200 |
In this table, the higher field strength applies at the
frequency band edges.
(b)
HIRF environment II is specified in the following
table:
Table II.—HIRF Environment II
|
Frequency |
Field strength
(volts/meter) |
|
Peak |
Average |
|
10 kHz–500 kHz |
20 |
20 |
|
500 kHz–2 MHz |
30 |
30 |
|
2 MHz–30 MHz |
100 |
100 |
|
30 MHz–100 MHz |
10 |
10 |
|
100 MHz–200 MHz |
30 |
10 |
|
200 MHz–400 MHz |
10 |
10 |
|
400 MHz–1 GHz |
700 |
40 |
|
1 GHz–2 GHz |
1,300 |
160 |
|
2 GHz–4 GHz |
3,000 |
120 |
|
4 GHz–6 GHz |
3,000 |
160 |
|
6 GHz–8 GHz |
400 |
170 |
|
8 GHz–12 GHz |
1,230 |
230 |
|
12 GHz–18 GHz |
730 |
190 |
|
18 GHz–40 GHz |
600 |
150 |
In this table, the higher field strength applies at the
frequency band edges.
(c)
HIRF environment III is specified in the following
table:
Table III.—HIRF Environment III
|
Frequency |
Field strength
(volts/meter) |
|
Peak |
Average |
|
10 kHz–100 kHz |
150 |
150 |
|
100 kHz–400 MHz |
200 |
200 |
|
400 MHz–700 MHz |
730 |
200 |
|
700 MHz–1 GHz |
1,400 |
240 |
|
1 GHz–2 GHz |
5,000 |
250 |
|
2 GHz–4 GHz |
6,000 |
490 |
|
4 GHz–6 GHz |
7,200 |
400 |
|
6 GHz–8 GHz |
1,100 |
170 |
|
8 GHz–12 GHz |
5,000 |
330 |
|
12 GHz–18 GHz |
2,000 |
330 |
|
18 GHz–40 GHz |
1,000 |
420 |
In this table, the higher field strength applies at the
frequency band edges.
(d)
Equipment HIRF Test Level 1 .
(1)
From 10 kilohertz (kHz) to 400 megahertz (MHz), use
conducted susceptibility tests with continuous wave
(CW) and 1 kHz square wave modulation with 90
percent depth or greater. The conducted
susceptibility current must start at a minimum of
0.6 milliamperes (mA) at 10 kHz, increasing 20
decibel (dB) per frequency decade to a minimum of 30
mA at 500 kHz.
(2)
From 500 kHz to 40 MHz, the conducted susceptibility
current must be at least 30 mA.
(3)
From 40 MHz to 400 MHz, use conducted susceptibility
tests, starting at a minimum of 30 mA at 40 MHz,
decreasing 20 dB per frequency decade to a minimum
of 3 mA at 400 MHz.
(4)
From 100 MHz to 400 MHz, use radiated susceptibility
tests at a minimum of 20 volts per meter (V/m) peak
with CW and 1 kHz square wave modulation with 90
percent depth or greater.
(5)
From 400 MHz to 8 gigahertz (GHz), use radiated
susceptibility tests at a minimum of 150 V/m peak
with pulse modulation of 4 percent duty cycle with a
1 kHz pulse repetition frequency. This signal must
be switched on and off at a rate of 1 Hz with a duty
cycle of 50 percent.
(e)
Equipment HIRF Test Level 2 . Equipment HIRF
test level 2 is HIRF environment II in table II of
this appendix reduced by acceptable aircraft
transfer function and attenuation curves. Testing
must cover the frequency band of 10 kHz to 8 GHz.
(f)
Equipment HIRF Test Level 3 .
(1)
From 10 kHz to 400 MHz, use conducted susceptibility
tests, starting at a minimum of 0.15 mA at 10 kHz,
increasing 20 dB per frequency decade to a minimum
of 7.5 mA at 500 kHz.
(2)
From 500 kHz to 40 MHz, use conducted susceptibility
tests at a minimum of 7.5 mA.
(3)
From 40 MHz to 400 MHz, use conducted susceptibility
tests, starting at a minimum of 7.5 mA at 40 MHz,
decreasing 20 dB per frequency decade to a minimum
of 0.75 mA at 400 MHz.
(4)
From 100 MHz to 8 GHz, use radiated susceptibility
tests at a minimum of 5 V/m.
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