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Subpart
A—General
27.1 Applicability.
(a)
This part prescribes airworthiness standards for the
issue of type certificates, and changes to those
certificates, for normal category rotorcraft with
maximum weights of 7,000 pounds or less and nine or
less passenger seats.
(b)
Each person who applies under Part 21 for such a
certificate or change must show compliance with the
applicable requirements of this part.
(c)
Multi-engine rotorcraft may be type certified as
Category A provided the requirements referenced in
appendix C of this part are met.
(a)
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.
(1)
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.
(2)
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.
(3)
The safety belt and shoulder harness must meet the
static and dynamic strength requirements, if
applicable, specified by the rotorcraft type
certification basis.
(4)
For purposes of this section, the date of
manufacture is either—
(i)
The date the inspection acceptance records, or
equivalent, reflect that the rotorcraft is complete
and meets the AFRO-CAA-Approved Type Design Data; or
(ii)
The date the foreign civil airworthiness authority
certifies that the rotorcraft is complete and issues
an original standard airworthiness certificate, or
equivalent, in that country.
(b)
For rotorcraft
(1)
The maximum passenger seat capacity may be increased
to eight or nine provided the applicant shows
compliance with all the airworthiness requirements
of this part.
(2)
The maximum weight may be increased to greater than
6,000 pounds provided—
(i)
The number of passenger seats is not increased above
the maximum number certificated on or
(ii) The applicant shows compliance with all of the
airworthiness requirements of this part in effect on
October 18, 1999.
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) must
be established so that it is—
(1)
Not more than—
(i)
The highest weight selected by the applicant;
(ii)
The design maximum (the highest weight at which
compliance with each applicable structural loading
condition of this part is shown); or
(iii) The highest weight at which compliance with
each applicable flight requirement of this part is
shown; and
(2)
Not less than the sum of—
(i)
The empty weight determined under 27.29; and
(ii)
The weight of usable fuel appropriate to the
intended operation with full payload;
(iii) The weight of full oil capacity; and
(iv)
For each seat, an occupant weight of 170 pounds or
any lower weight for which certification is
requested.
(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—
(1)
Not more than the sum of—
(i)
The empty weight determined under 27.29; and
(ii)
The weight of the minimum crew necessary to operate
the rotorcraft, assuming for each crewmember a
weight no more than 170 pounds, or any lower weight
selected by the applicant or included in the loading
instructions; and
(2)
Not less than—
(i)
The lowest weight selected by the applicant;
(ii)
The design minimum weight (the lowest weight at
which compliance with each applicable structural
loading condition of this part is shown); or
(iii) The lowest weight at which compliance with
each applicable flight requirement of this part is
shown.
(c)
Total weight with jettison able external load.
A total weight for the rotorcraft with a
jettison able 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 27.865 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
27.865(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 27.25. 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
roto-craft 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 auto-rotative
maneuver to be performed throughout the ranges of
airspeed and weight for which certification is
requested.
(b)
Normal main rotor high pitch limits (power on).
For rotocraft, 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 limits (power off).
It must be shown, with power off, that—
(1)
The normal main rotor low pitch limit provides
sufficient rotor speed, in any auto-rotative
condition, under the most critical combinations of
weight and airspeed; and
(2)
It is possible to prevent over-speeding 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.
(a)
Unless otherwise prescribed, the performance
requirements of this subpart must be met for still
air and a standard atmosphere.
(b)
The performance must correspond to the engine power
available under the particular ambient atmospheric
conditions, the particular flight condition, and the
relative humidity specified in paragraphs (d) or (e)
of this section, as appropriate.
(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
appropriate to the particular ambient atmospheric
conditions and the particular flight condition.
(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 degrees F. Between these two temperatures, the
relative humidity must vary linearly.
(f)
For turbine-engine-powered rotorcraft, a means must
be provided to permit the pilot to determine prior
to take-off that each engine is capable of
developing the power necessary to achieve the
applicable rotorcraft performance prescribed in this
subpart.
(a)
The take-off, with take-off power and r.p.m., and
with the extreme forward center of gravity—
(1)
May not require exceptional piloting skill or
exceptionally favorable conditions; and
(2)
Must be made in such a manner that a landing can be
made safely at any point along the flight path if an
engine fails.
(b)
Paragraph (a) of this section must be met throughout
the ranges of—
(1)
Altitude, from standard sea level conditions to the
maximum altitude capability of the rotorcraft, or
7,000 feet, whichever is less; and
(2)
Weight, from the maximum weight (at sea level) to
each lesser weight selected by the applicant for
each altitude covered by paragraph (b)(1) of this
section.
(a)
For rotorcraft other than helicopters—
(1)
The steady rate of climb, at V Y, must be
determined—
(i)
With maximum continuous power on each engine;
(ii)
With the landing gear retracted; and
(iii) For the weights, altitudes, and temperatures
for which certification is requested; and
(2)
The climb gradient, at the rate of climb determined
in accordance with paragraph (a)(1) of this section,
must be either—
(i)
At least 1:10 if the horizontal distance required to
take-off and climb over a 50-foot obstacle is
determined for each weight, altitude, and
temperature within the range for which certification
is requested; or
(ii)
At least 1:6 under standard sea level conditions.
(b)
Each helicopter must meet the following
requirements:
(1)
VY must be determined—
(i)
For standard sea level conditions;
(ii)
At maximum weight; and
(iii) With maximum continuous power on each engine.
(2)
The steady rate of climb must be determined—
(i)
At the climb speed selected by the applicant at or
below VNE;
(ii)
Within the range from sea level up to the maximum
altitude for which certification is requested;
(iii) For the weights and temperatures that
correspond to the altitude range set forth in
paragraph (b)(2)(ii) of this section and for which
certification is requested; and
(iv)
With maximum continuous power on each engine.
For
multiengine helicopters, the steady rate of climb
(or descent), at V y(or at the speed for
minimum rate of descent), must be determined with—
(a)
Maximum weight;
(b)
The critical engine inoperative and the remaining
engines at either—
(1)
Maximum continuous power and, for helicopters for
which certification for the use of 30-minute OEI
power is requested, at 30-minute OEI power; or
(2)
Continuous OEI power for helicopters for which
certification for the use of continuous OEI power is
requested.
27.71 Glide performance.
For
single-engine helicopters and multiengine
helicopters that do not meet the Category A engine
isolation requirements of Part 29 of this chapter,
the minimum rate of descent airspeed and the best
angle-of-glide airspeed must be determined in
autorotation at—
(a)
Maximum weight; and
(b)
Rotor speed(s) selected by the applicant.
(a)
For helicopters—
(1)
The hovering ceiling must be determined over the
ranges of weight, altitude, and temperature for
which certification is requested, with—
(i)
Take-off power;
(ii)
The landing gear extended; and
(iii) The helicopter in ground effect at a height
consistent with normal take-off procedures; and
(2)
The hovering ceiling determined under paragraph
(a)(1) of this section must be at least—
(i)
For reciprocating engine powered helicopters, 4,000
feet at maximum weight with a standard atmosphere;
or
(ii)
For turbine engine powered helicopters, 2,500 feet
pressure altitude at maximum weight at a temperature
of standard +40 degrees F.
(b)
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)
Take-off power; and
(2)
The landing gear extended.
(a)
The rotorcraft must be able to be landed with no
excessive vertical acceleration, no tendency to
bounce, nose over, ground loop, porpoise, or water
loop, and without exceptional piloting skill or
exceptionally favorable conditions, with—
(1)
Approach or glide speeds appropriate to the type of
rotorcraft and selected by the applicant;
(2)
The approach and landing made with—
(i)
Power off, for single-engine rotorcraft; and
(ii)
For multiengine rotocraft, one engine inoperative
and with each operating engine within approved
operating limitations; and
(3)
The approach and landing entered from steady
autorotation.
(b)
Multi-engine rotorcraft must be able to be landed
safely after complete power failure under normal
operating conditions.
(a)
If there is any combination of height and forward
speed (including hover) under which a safe landing
cannot be made under the applicable power failure
condition in paragraph (b) of this section, a
limiting height-speed envelope must be established
(including all pertinent information) for that
condition, throughout the ranges of—
(1)
Altitude, from standard sea level conditions to the
maximum altitude capability of the rotorcraft, or
7,000 feet, whichever is less; and
(2)
Weight, from the maximum weight (at sea level) to
the lesser weight selected by the applicant for each
altitude covered by paragraph (a)(1) of this
section. For helicopters, the weight at altitudes
above sea level may not be less than the maximum
weight or the highest weight allowing hovering out
of ground effect which is lower.
(b)
The applicable power failure conditions are—
(1)
For single-engine helicopters, full autorotation;
(2)
For multi-engine helicopters, one engine inoperative
(where engine isolation features insure continued
operation of the remaining engines), and the
remaining engines at the greatest power for which
certification is requested, and
(3)
For other rotocraft, conditions appropriate to the
type.
The rotorcraft must—
(a)
Except as specifically required in the applicable
section, meet the flight characteristics
requirements of this subpart—
(1)
At the altitudes and temperatures expected in
operation;
(2)
Under any critical loading condition within the
range of weights and centers of gravity for which
certification is requested;
(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 of Part 29 of this chapter;
(2)
Sudden, complete power failure for other rotorcraft;
and
(3)
Sudden, complete control system failures specified
in 27.695 of this part; and
(c)
Have any additional characteristic 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)
Take-off;
(ii)
Climb;
(iii) Level flight;
(iv)
Turning flight;
(v)
Glide;
(vi)
Landing (power on and power off); and
(vii) Recovery to power-on flight from a balked
auto-rotative approach.
(b)
The margin of cyclic control must allow satisfactory
roll and pitch control at VNE with—
(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
take-offs, sideward flight, and rearward flight),
with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Critical rotor r.p.m.; and
(4)
Altitude, from standard sea level conditions to the
maximum altitude capability of the rotorcraft or
7,000 feet, whichever is less.
(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 engine 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 27.1505(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.
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 take-offs
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 27.175 (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 27.175(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
V Y to 1.2 V Y, 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 H or 0.7 V NE, whichever is less, to
1.1 V H or 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 H or 0.9 V
NE, whichever is less;
(4)
The landing gear retracted; and
(5)
The rotorcraft trimmed at 0.9 V H or 0.9 V
NE, whichever is less.
(c)
Auto-rotation. Static longitudinal stability
must be shown in autorotation at airspeeds from 0.5
times the speed for minimum rate of descent to VNE,
or to 1.1 VNE(power-off) if VNE(power-off) is
established under 27.1505(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 and
direction of motion prescribed in 27.173 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 27.175 (a) and (b).
This must be shown by steadily increasing
directional control deflection for sideslip angles
up to ±10° from trim. Sufficient cues must accompany
sideslip to alert the pilot when approaching
sideslip limits.
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, take-off, 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.
Subpart C—Strength
Requirements
(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 test 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 weight.
(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 power plant
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 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);
omega= The
angular velocity of rotor (radians per second); and
R= The rotor
radius (ft).
The
rotorcraft must be designed to withstand, at each
critical airspeed including hovering, the loads
resulting from a vertical gust 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
27.397(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
this section, in unaccelerated flight with zero yaw,
at forward speeds from 0.6 VNE up to VNE
or 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
27.397(a);
(2)
Attain a resulting sideslip angle or 15°, whichever
is less, at the lesser speed of VNE or 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.
(a)
For turbine engines, the limit torque may not be
less than the highest of—
(1)
The mean torque for maximum continuous power
multiplied by 1.25;
(2)
The torque required by 27.923;
(3)
The torque required by 27.927; or
(4)
The torque imposed by sudden engine stoppage due to
malfunction or structural failure (such as
compressor jamming).
(b)
For reciprocating engines, the limit torque may not
be less than 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 27.395,
27.397, 27.399, 27.411, and 27.427.
(a)
The part of each control system from the pilot's
controls to the control stops must be designed to
withstand pilot forces of not less than—
(1)
The forces specified in 27.397; or
(2)
If the system prevents the pilot from applying the
limit pilot forces to the system, the maximum forces
that the system allows the pilot to apply, but not
less than 0.60 times the forces specified in 27.397.
(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 27.397.
(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 force 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 27.397, 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 rational analysis, the design loads
resulting from 0.60 of the specified limit pilot
forces are acceptable minimum design loads.
(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 27.397, 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 follows 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 designed
to withstand the loads that result when pilot forces
of 0.75 times those obtained under 27.395 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 but 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 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 27.725.
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 27.235, 27.479 through 27.485, and 27.493
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 all wheels contact 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 27.471.
(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)
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 27.479(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 27.479 (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 all 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 27.479(a)(1);
and
(2)
1.0 for the attitude specified in 27.479(a)(2); and
(b)
The structure must be designed to withstand at the
ground contact point of each wheel with brakes, a
drag load 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 27.471
through 27.475;
(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 auto-rotative
landings. In this attitude—
(1)
The appropriate ground loads specified in paragraphs
(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 paragraphs (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 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 27.471 through
27.475.
(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 27.725;
and
(ii)
An assumed rotor lift of not more than 1.5 times
that used in the limit drop tests prescribed in
27.725.
(4)
Compliance with paragraphs (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)
Side-loads 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 side-load of 25 percent of their value.
(3)
The total side-load must be applied equally between
the 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 side-loads; and
(ii)
Outward acting side-loads.
(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 27.473(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.
If
certification for float operation is requested, the
rotorcraft, with floats, must be designed to
withstand the following loading conditions (where
the limit load factor is determined under 27.473(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.
(a)
Each main rotor assembly (including rotor hubs and
blades) must be designed as prescribed in this
section.
(b)
[Reserved]
(c)
The main rotor structure must be designed to
withstand the following loads prescribed in 27.337
through 27.341:
(1)
Critical flight loads.
(2)
Limit loads occurring under normal conditions of
autorotation. For this condition, the rotor r.p.m.
must be selected to include the effects of altitude.
(d)
The main 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 main 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; and
(ii)
The limit engine torque specified in 27.361.
(2)
The limit torque must be distributed to the rotor
blades in a rational manner.
(a)
Each fuselage, landing gear, and rotor pylon
structure must be designed as prescribed in this
section. Resultant rotor forces may be represented
as a single force applied at the rotor hub
attachment point.
(b)
Each structure must be designed to withstand—
(1)
The critical loads prescribed in 27.337 through
27.341;
(2)
The applicable ground loads prescribed in 27.235,
27.471 through 27.485, 27.493, 27.497, 27.501,
27.505, and 27.521; and
(3)
The loads prescribed in 27.547 (d)(2) and (e).
(c)
Auxiliary rotor thrust, and the balancing air and
inertia loads occurring under accelerated flight
conditions, must be considered.
(d)
Each engine mount and adjacent fuselage structure
must be designed to withstand the loads occurring
under accelerated flight and landing conditions,
including engine torque.
(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 intended displacement of the
seat device.
(v)
Rearward—1.5g.
(c)
The supporting structure must be designed to
restrain, under any ultimate inertial load 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,
transmissions, 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 when those loads are applied to that area:
(i)
Upward—1.5g.
(ii)
Forward—4.0g.
(iii) Sideward—2.0g.
(iv)
Downward—4.0g.
(a)
The rotorcraft, although it may be damaged in an
emergency 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 the loads resulting from
the conditions prescribed in this section.
(b)
Each seat type design or other seating device
approved for crew or passenger occupancy during
take-off 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 upper torso 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 27.801(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, of 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 27.801(d). Maximum roll and pitch
angles determined from compliance with 27.801(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 perscribed 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.
(a)
General. Each portion of the flight structure
(the flight structure includes rotors, rotor drive
systems between the engines and the rotor hubs,
controls, fuselage, landing gear, and their related
primary attachments), the failure of which could be
catastrophic, must be identified and must be
evaluated under paragraph (b), (c), (d), or (e) of
this section. The following apply to each fatigue
evaluation:
(1)
The procedure for the evaluation must be approved.
(2)
The locations of probable failure must be
determined.
(3)
Inflight measurement must be included in determining
the following:
(i)
Loads or stresses in all critical conditions
throughout the range of limitations in 27.309,
except that maneuvering load factors need not exceed
the maximum values expected in operation.
(ii)
The effect of altitude upon these loads or stresses.
(4)
The loading spectra must be as severe as those
expected in operation including, but not limited to,
external cargo operations, if applicable, and
ground-air-ground cycles. The loading spectra must
be based on loads or stresses determined under
paragraph (a)(3) of this section.
(b)
Fatigue tolerance evaluation. It must be
shown that the fatigue tolerance of the structure
ensures that the probability of catastrophic fatigue
failure is extremely remote without establishing
replacement times, inspection intervals or other
procedures under section A27.4 of appendix A.
(c)
Replacement time evaluation. it must be shown
that the probability of catastrophic fatigue failure
is extremely remote within a replacement time
furnished under section A27.4 of appendix A.
(d)
Fail-safe evaluation. The following apply to
fail-safe evaluation:
(1)
It must be shown that all partial failures will
become readily detectable under inspection
procedures furnished under section A27.4 of appendix
A.
(2)
The interval between the time when any partial
failure becomes readily detectable under paragraph
(d)(1) of this section, and the time when any such
failure is expected to reduce the remaining strength
of the structure to limit or maximum attainable
loads (whichever is less), must be determined.
(3)
It must be shown that the interval determined under
paragraph (d)(2) of this section is long enough, in
relation to the inspection intervals and related
procedures furnished under section A27.4 of appendix
A, to provide a probability of detection great
enough to ensure that the probability of
catastrophic failure is extremely remote.
(e)
Combination of replacement time and failsafe
evaluations. A component may be evaluated under
a combination of paragraphs (c) and (d) of this
section. For such component it must be shown that
the probability of catastrophic failure is extremely
remote with an approved combination of replacement
time, inspection intervals, and related procedures
furnished under section A27.4 of appendix A.
(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
characteristics 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.
(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 non-friction
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 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 so as not to endanger
the rotorcraft.
(d)
The electrical 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 an 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 the 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 structure, 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, Pennsylvania 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 27.621 through
27.625 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 to which 27.621 through 27.625 apply,
the factor of safety prescribed in 27.303 must be
multiplied by a special factor equal to—
(1)
The applicable special factors prescribed in 27.621
through 27.625; 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 §27.305 at an ultimate
load corresponding to a casting factor of 1.25; and
(ii)
The deformation requirements of §27.305 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 27.561(b)(3)
multiplied by a fitting factor of 1.33.
Each
aerodynamic surface of the rotorcraft must be free
from flutter under each appropriate speed and power
condition.
Rotors
(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 rotors 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.
(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 27.241.
(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.
27.672 Stability
augmentation, automatic, and power-operated systems.
If
the functioning of stability augmentation or other
automatic or power-operated systems is necessary to
show compliance with the flight characteristics
requirements of this part, such systems must comply
with 27.671 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 movement of the flight
controls in the normal sense and deactivating the
failed system.
(c)
It must be shown that after any single failure of
the stability augmentation system or any other
automatic or power-operated system—
(1)
The 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 permit 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)
Give unmistakable warning to the pilot when the lock
is engaged; 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 the 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 than three thirty-seconds of an
inch 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 which must be used
are specified in Military Handbook MIL-HDBK-5C, Vol.
1 & Vol. 2, Metallic Materials and Elements for
Flight Vehicle Structures, (Sept. 15, 1976, as
amended through December 15, 1978). This
incorporation by reference was approved by the
Director of the Federal Register in accordance with
5 AFRO-CAA section 552(a) and 1 CFR part 51. Copies
may be obtained from the Naval Publications and
Forms Center, 5801 Tabor Avenue, Philadelphia,
Pennsylvania, 19120. Copies may be inspected at the
AFRO-CAA, Rotorcraft Standards Staff, 4400 Blue
Mount Road, Fort Worth, Texas, or at the National
Archives and Records Administration (NARA). For
information on the availability of this material at
NARA, call 202–741–6030, or go to: http://www.archives.gov/federal_register/code_of_federal_regulations/ibr_locations.html.
(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 3°.
(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 must 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
hydraulic 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 27.725 and
27.727, 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—
(1)
13 inches from the lowest point of the landing gear
to the ground; or
(2)
Any lesser height, not less than eight inches,
resulting in a drop contact velocity equal to the
greatest probable sinking speed likely to occur at
ground contact in normal power-off landings.
(b)
If considered, the rotor lift specified in 27.473(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
formula may be used instead of more rational
computations:
where:
W e=the
effective weight to be used in the drop test (lbs.);
W = W
M for 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
N for 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
T for tail-wheel units (lbs.), equal to whichever of
the following is critical:
(1)
The static weight on the tail-wheel with the
rotorcraft resting on all wheels; or
(2)
The vertical component of the ground reaction that
would occur at the tail-wheel, assuming that the
mass of the rotorcraft acts at the center of gravity
and exerts a force of l 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 =ration 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
travels (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
27.725(a).
(b)
Rotor lift, where considered in a manner similar to
that prescribed in 27.725(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 the
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 mechansim,
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 a 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
27.777 and 27.779.
(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.
(a)
Each landing gear wheel must have a tire—
(1)
That is a proper fit on the rim of the wheel; and
(2)
Of the proper rating.
(b)
The maximum static load rating of each tire must
equal or exceed the static ground reaction obtained
at its wheel, assuming—
(1)
The design maximum weight; and
(2)
The most unfavorable center of gravity.
(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.
27.735 Brakes.
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.
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.
(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 main floats 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 27.521(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 27.521. These loads may be
distributed along the length of the float.
For
each rotorcraft, with a hull and auxiliary floats,
that is to be approved for both taking off from and
landing on water, the hull and auxiliary floats must
have enough watertight compartments so that, with
any single compartment flooded, the buoyancy of the
hull and auxiliary floats (and wheel tires if used)
provides a margin of positive stability great enough
to minimize the probability of capsizing.
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 seat; and
(c)
The vibration and noise characteristics of cockpit
appurtenances may not interfere with safe operation.
(a)
Each pilot compartment must be free from glare and
reflections that could interfere with the pilot's
view, and designed so that—
(1)
Each pilot's view is sufficiently extensive, clear,
and undistorted for safe operation; and
(2)
Each pilot is protected from the elements so that
moderate rain conditions do not unduly impair his
view of the flight path in normal flight and while
landing.
(b)
If certification for night operation is requested,
compliance with paragraph (a) of this section must
be shown in night flight tests.
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 pilots'
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 left-hand 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 where persons
using it will not be endangered by the rotors,
propellers, engine intakes, and exhausts when
appropriate operating procedures are used. If
opening procedures are required, they must be marked
inside, on or adjacent to the door opening device.
(a)
Each seat, safety belt, harness, and adjacent part
of the rotorcraft at each station designated for
occupancy during take-off 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 static inertial
load factors specified in 27.561(b) and dynamic
conditions specified in 27.562.
(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 27.562(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 belts
and harnesses, 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
enable the occupants to steady themselves while
using the aisle in moderately rough air.
(e)
Each projecting object that could 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 occupant,
seat, and safety belt or harness corresponding with
the applicable flight and ground load conditions,
including the emergency landing conditions of
27.561(b). In addition—
(1)
Each pilot seat must be designed for the reactions
resulting from the application of the pilot forces
prescribed in 27.397; and
(2)
The inertial forces prescribed in 27.561(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
27.561(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 27.561, 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 seats to reduce occupant loads for the emergency
landing dynamic conditions of 27.562; otherwise, the
system must remain intact and must not interfere
with rapid evacuation of the rotorcraft.
(k)
For the 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 27.561(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 diaphram, 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
load 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 27.561(b). The
fitting factor required by 27.625(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 27.561.
(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 27.561,
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 27.561 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 27.807(d), 27.1411 and 27.1415.
(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 factor
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 life rafts required by
27.1415. If compliance with this provision is shown
by buoyancy and trim computations, appropriate
allowances must be made for probable structural
damage and leakage. If the 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 buoyancy volume.
(e)
Unless the effects of the collapse of external doors
and windows are accounted for in the investigation
of the probable behavior of the 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)
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
rotor-craft 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 flight crew emergency exit must not be
obstructed by water or flotation devices after an
emergency landing on water. This must be shown by
test, demonstration, or analysis.
(a)
Number and location. (1) There must be at
least one emergency exit on each side of the cabin
readily accessible to each passenger. One of these
exits must be usable in any probable attitude that
may result from a crash;
(2)
Doors intended for normal use may also serve as
emergency exits, provided that they meet the
requirements of this section; and
(3)
If emergency flotation devices are installed, there
must be an emergency exit accessible to each
passenger on each side of the cabin that is shown by
test, demonstration, or analysis to;
(i)
Be above the waterline; and
(ii)
Open without interference from flotation devices,
whether stowed or deployed.
(b)
Type and operation. Each emergency exit
prescribed by paragraph (a) of this section must—
(1)
Consist of a movable window or panel, or additional
external door, providing an unobstructed opening
that will admit a 19-by 26-inch ellipse;
(2)
Have simple and obvious methods of opening, from the
inside and from the outside, which do not require
exceptional effort;
(3)
Be arranged and marked so as to be readily located
and opened even in darkness; and
(4)
Be reasonably protected from jamming by fuselage
deformation.
(c)
Tests. The proper functioning of each
emergency exit must be shown by test.
(d)
Ditching emergency exits for passengers. If
certification with ditching provisions is requested,
the markings required by paragraph (b)(3) of this
section must be designed to remain visible if the
rotorcraft is capsized and the cabin is submerged.
(a)
The ventilating system for the pilot and passenger
compartments must be designed to prevent the
presence of excessive quantities of fuel fumes and
carbon monoxide.
(b)
The concentration of carbon monoxide may not exceed
one part in 20,000 parts of air during forward
flight or hovering in still air. If the
concentration exceeds this value under other
conditions, there must be suitable operating
restrictions.
Each
combustion heater must be approved.
For
each compartment to be used by the crew or
passengers—
(a)
The materials must be at least flame-resistant;
(b)
[Reserved]
(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.
(a)
Each cargo and baggage compartment must be
constructed of, or lined with, materials that are at
least—
(1)
Flame resistant, in the case of compartments that
are readily accessible to a crewmember in flight;
and
(2)
Fire resistant, in the case of other compartments.
(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.
(a)
General. For each heating system that
involves the passage of cabin air over, or close to,
the exhaust manifold, there must be means to prevent
carbon monoxide from entering any cabin or pilot
compartment.
(b)
Heat exchangers. Each heat exchanger must be—
(1)
Of suitable materials;
(2)
Adequately cooled under all conditions; and
(3)
Easily disassembled for inspection.
(c)
Combustion heater fire protection. Except for
heaters which incorporate designs to prevent hazards
in the event of fuel leakage in the heater fuel
system, fire within the ventilating air passage, or
any other heater malfunction, each heater zone must
incorporate the fire protection features of the
applicable requirements of 27.1183, 27.1185,
27.1189, 27.1191, and be provided with—
(1)
Approved, quick-acting fire detectors in numbers and
locations ensuring prompt detection of fire in the
heater region.
(2)
Fire extinguisher systems that provide at least one
adequate discharge to all areas of the heater
region.
(3)
Complete drainage of each part of each zone to
minimize the hazards resulting from failure or
malfunction of any component containing flammable
fluids. The drainage means must be—
(i)
Effective under conditions expected to prevail when
drainage is needed; and
(ii)
Arranged so that no discharged fluid will cause an
additional fire hazard.
(4)
Ventilation, arranged so that no discharged vapors
will cause an additional fire hazard.
(d)
Ventilating air ducts. Each ventilating air
duct passing through any heater region must be
fireproof.
(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.
(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.
(e)
Combustion air ducts. Each combustion air
duct must be fireproof for a distance great enough
to prevent damage from backfiring or reverse flame
propagation.
(1)
No combustion air duct may connect 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.
(2)
No combustion air duct may restrict the prompt
relief of any backfire that, if so restricted, could
cause heater failure.
(f)
Heater control: 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.
(g)
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 (g)(1) of this
section for any individual heater must—
(i)
Be independent of components serving any other
heater, the heat output of which 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, the heat output of which is essential for
safe operation, has been shut off by the automatic
means prescribed in paragraph (g)(1) of this
section.
(h)
Air intakes. Each combustion and ventilating
air intake must be located so that no flammable
fluids or vapors can enter the heater system—
(1)
During normal operation; or
(2)
As a result of the malfunction of any other
component.
(i)
Heater exhaust. Each heater exhaust system
must meet the requirements of 27.1121 and 27.1123.
(1)
Each exhaust shroud must be sealed so that no
flammable fluids or hazardous quantities of vapors
can reach the exhaust system through joints.
(2)
No exhaust system may restrict the prompt relief of
any backfire that, if so restricted, could cause
heater failure.
(j)
Heater fuel systems. Each heater fuel system
must meet the power-plant 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.
(k)
Drains. There must be means for safe drainage
of any fuel that might accumulate in the combustion
chamber or the heat exchanger.
(1)
Each part of any drain that operates at high
temperatures must be protected in the same manner as
heater exhausts.
(2)
Each drain must be protected against hazardous ice
accumulation under any operating condition.
27.861 Fire
protection of structure, controls, and other parts.
Each
part of the structure, controls, rotor mechanism,
and other parts essential to a controlled landing
that would be affected by power-plant fires must be
fireproof or protected so they can perform their
essential functions for at least 5 minutes under any
foreseeable power-plant 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 27.337 through 27.341, 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 27.337
through 27.341, 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, the ingress and egress
instructions;
(4)
Have equipment to allow direct intercommunication
among required crewmembers and external occupants;
and
(5)
Have the appropriate limitations and procedures
incorporated in the flight manual for conducting
human external cargo operations.
(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 27.25 and this
section.
(f)
The fatigue evaluation of 27.571 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 27.571 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)
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;
(2)
Accessibility must be provided to allow any
inspection and maintenance necessary for continued
airworthiness;
(3)
Electrical interconnections must be provided to
prevent differences of potential between major
components of the installation and the rest of the
rotorcraft;
(4)
Axial and radial expansion of turbine engines may
not affect the safety of the installation; and
(5)
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)
The installation must comply with—
(1)
The installation instructions provided under 33.5 of
this chapter; and
(2)
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)
Engine or drive system cooling fan blade
protection. (1) If an engine or rotor drive
system cooling fan is installed, there must be a
means to protect the rotorcraft and allow a safe
landing if a fan blade fails. This must be shown by
showing that—
(i)
The fan blades are contained in case of failure;
(ii)
Each fan is located so that a failure will not
jeopardize safety; or
(iii) Each fan blade can withstand an ultimate load
of 1.5 times the centrifugal force resulting from
operation limited by the following:
(A)
For fans driven directly by the engine—
(
1 ) The terminal engine r.p.m. under
uncontrolled conditions; or
(
2 ) An overspeed limiting device.
(B)
For fans driven by the rotor drive system, the
maximum rotor drive system rotational speed to be
expected in service, including transients.
(2)
Unless a fatigue evaluation under §27.571 is
conducted, it must be shown that cooling fan blades
are not operating at resonant conditions within the
operating limits of the rotorcraft.
(c)
Turbine engine installation. For turbine
engine installations, 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 turbine rotor structural
integrity will not be exceeded in service.
(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.
(c)
No part of the rotor drive system may be subjected
to excessive vibration stresses.
(a)
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.
(b)
Each rotor drive system must be arranged so 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.
(c)
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.
(d)
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.
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)
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.
(b)
Each rotor drive system and control mechanism must
be tested for not less than 100 hours. The test 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.
(c)
A 60-hour part of the test prescribed in paragraph
(b) of this section must be run at not less than
maximum continuous torque and the maximum speed for
use with maximum continuous torque. In this test,
the main rotor controls must be set in the position
that will give maximum longitudinal cyclic pitch
change to simulate forward flight. The auxiliary
rotor controls must be in the position for normal
operation under the conditions of the test.
(d)
A 30-hour or, for rotorcraft for which the use of
either 30-minute OEI power or continuous OEI power
is requested, a 25-hour part of the test prescribed
in paragraph (b) of this section must be run at not
less than 75 percent of maximum continuous torque
and the minimum speed for use with 75 percent of
maximum continuous torque. The main and auxiliary
rotor controls must be in the position for normal
operation under the conditions of the test.
(e)
A 10-hour part of the test prescribed in paragraph
(b) of this section must be run at not less than
take-off torque and the maximum speed for use with
take-off torque. The main and auxiliary rotor
controls must be in the normal position for vertical
ascent.
(1)
For multiengine rotorcraft for which the use of
21/2minute OEI power is requested, 12 runs during
the 10-hour test must be conducted as follows:
(i)
Each run must consist of at least one period of
21/2minutes with take-off torque and the maximum
speed for use with take-off 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 21/2minute OEI torque and the maximum
speed for use with 21/2minute OEI torque for
21/2minutes.
(2)
For multiengine turbine-powered rotorcraft for which
the use of 30-second and 2-minute OEI power is
requested, 10 runs must be conducted as follows:
(i)
Immediately following a take-off run of at least 5
minutes, each power source must simulate a failure,
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, followed by application of
the maximum torque and the maximum speed for use
with 2-minute OEI power for not less than 2 minutes.
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 removal 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.
(f)
The parts of the test prescribed in paragraphs (c)
and (d) of this section must be conducted in
intervals of not less than 30 minutes and may be
accomplished either on the ground or in flight. The
part of the test prescribed in paragraph (e) of this
section must be conducted in intervals of not less
than five minutes.
(g)
At intervals of not more than five hours during the
tests prescribed in paragraphs (c), (d), and (e) of
this section, the engine must be stopped rapidly
enough to allow the engine and rotor drive to be
automatically disengaged from the rotors.
(h)
Under the operating conditions specified in
paragraph (c) of this section, 500 complete cycles
of lateral control, 500 complete cycles of
longitudinal control of the main rotors, and 500
complete cycles of control of each auxiliary rotor
must be accomplished. A “complete cycle” involves
movement of the controls from the neutral position,
through both extreme positions, and back to the
neutral position, except that control movements need
not produce loads or flapping motions exceeding the
maximum loads or motions encountered in flight. The
cycling may be accomplished during the testing
prescribed in paragraph (c) of this section.
(i)
At least 200 start-up clutch engagements must be
accomplished—
(1)
So that the shaft on the driven side of the clutch
is accelerated; and
(2)
Using a speed and method selected by the applicant.
(j)
For multiengine rotorcraft for which the use of
30-minute OEI power is requested, five runs must be
made at 30-minute OEI torque and the maximum speed
for use with 30-minute OEI torque, in which each
engine, in sequence, is made inoperative and the
remaining engine(s) is run for a 30-minute period.
(k)
For multiengine rotorcraft for which the use of
continuous OEI power is requested, five runs must be
made at continuous OEI torque and the maximum speed
for use with continuous OEI torque, in which each
engine, in sequence, is made inoperative and the
remaining engine(s) is run for a 1-hour period.
(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
rating 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, or torque that is at least equal to the lesser
of—
(i)
The maximum torque used in meeting 27.923 plus 10
percent; or
(ii)
The maximum attainable torque output of the engines,
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 15 minutes.
(3)
The tests prescribed in this paragraph must be
conducted on the rotorcraft at the maximum
rotational speed intended for the power condition of
the test 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.
(c)
It must be shown by tests that the rotor drive
system is capable of operating under auto-rotative
conditions for 15 minutes after the loss of pressure
in the rotor drive primary oil system.
(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 auto-rotative
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 as stall, surge, or 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 functioning under any
likely operating condition, including the maneuvers
for which certification is requested.
(b)
Each fuel system must be arranged so that—
(1)
No 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 °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 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 tank 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 rupture 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 that
inadvertent fuel shutoff in flight is improbable in
accordance with §27.955(a) and must comply with the
fatigue evaluation requirements of 27.571 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 27.571 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 components
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 27.571.
(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 semi-rigid fuel tanks. Rigid or
semi-rigid fuel tank or bladder walls must be impact
and tear resistant.
(a)
Each fuel system for multiengine rotorcraft must
allow fuel to be supplied to each engine through a
system independent of those parts of each system
supplying fuel to other engines. However, separate
fuel tanks need not be provided for each engine.
(b)
If a single fuel tank is used on a multiengine
rotorcraft, the following must be provided:
(1)
Independent tank outlets for each engine, each
incorporating a shutoff valve at the tank. This
shutoff valve may also serve as the firewall shutoff
valve required by 27.995 if the line between the
valve and the engine compartment does not contain a
hazardous amount of fuel that can drain into the
engine compartment.
(2)
At least two vents arranged to minimize the
probability of both vents becoming obstructed
simultaneously.
(3)
Filler caps designed to minimize the probability of
incorrect installation or in-flight loss.
(4)
A fuel system in which those parts of the system
from each tank outlet to any engine are independent
of each part of each system supplying fuel to other
engines.
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; or
(c)
Corona and streamering at fuel vent outlets.
(a)
General. The fuel system for each engine must
be shown to provide the engine with at least 100
percent of the fuel required under each operating
and maneuvering condition to be approved for the
rotorcraft including, as applicable, the fuel
required to operate the engine(s) under the test
conditions required by 27.927. 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 critical
accelerations, 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 27.959, plus the minimum additional fuel
necessary to conduct the test.
(3)
The fuel head between the tank outlet and the engine
inlet must be critical with respect to rotorcraft
flight attitudes.
(4)
The critical fuel pump (for pump-fed systems) is
installed to produce (by actual or simulated
failure) the critical restriction to fuel flow to be
expected from pump failure.
(5)
Critical values of engine rotation 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 must be applied.
(7)
The fuel filter required by 27.997 must be blocked
to the degree necessary to simulate the accumulation
of fuel contamination required to activate the
indicator required by 27.1305(q).
(b)
Fuel transfer systems. If normal operation of
the fuel system requires fuel to be transferred to
an engine feed tank, the transfer must occur
automatically via a system which has been shown to
maintain the fuel level in the engine feed 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
systems must, in addition to having appropriate
manual switching capability, 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 the engine alone
contains usable fuel.
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 with
features conducive to vapor formation must be shown
by test to operate satisfactorily (within
certification limits) when using fuel at a
temperature of 110 °F under critical operating
conditions including, if applicable, the engine
operating conditions defined by 27.927 (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 fuel tank of 10 gallons or greater capacity
must have internal baffles, or must have external
support to resist surging.
(c)
Each fuel tank must be separated from the engine
compartment by a firewall. At least one-half inch of
clear airspace must be provided between the tank and
the firewall.
(d)
Spaces adjacent to the surfaces of fuel tanks must
be ventilated so that fumes cannot accumulate in the
tank compartment in case of leakage. If two or more
tanks have interconnected outlets, they must be
considered as one tank, and the airspaces in those
tanks must be interconnected to prevent the flow of
fuel from one tank to another as a result of a
difference in pressure between those airspaces.
(e)
The maximum exposed surface temperature of any
component in the fuel tank must be less, by a safe
margin as determined by the Administrator, 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 failure or malfunction
conditions of all components inside the tank.
(f)
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 27.952, and must be adequate to withstand loads
and abrasions to be expected in personnel
compartments.
(g)
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.
(h)
Each integral fuel tank must have provisions for
inspection and repair of its interior.
(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, nonmetallic tank with
walls that are not supported by the rotorcraft
structure, and 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 far as possible. However, the pressure need not
exceed 3.5 p.s.i. on surfaces not exposed to the
acceleration loading.
(c)
Each non-metallic 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 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 support 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 calculation 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 paragraphs (d)(3)(ii) and (iii) of this
section, 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 it is not necessary for them 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
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 27.1185 (a) and (c).
(d)
No rotorcraft skin immediately adjacent to a major
air outlet from the engine compartment may act as
the wall of the 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 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 drainable sump with an
effective capacity in any ground attitude to be
expected in service of 0.25 percent of the tank
capacity or 1/16 gallon, whichever is greater,
unless—
(1)
The fuel system has a sediment bowl or chamber that
is accessible for preflight drainage and has a
minimum capacity of 1 ounce for every 20 gallons of
fuel tank capacity; and
(2)
Each fuel tank drain is located so that in any
ground attitude to be expected in service, water
will drain from all parts of the tank to the
sediment bowl or chamber.
(b)
Each sump, sediment bowl, and sediment chamber drain
required by this section must comply with the drain
provisions of 27.999(b).
(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 27.952(c). In addition—
(1)
Each filler must be marked as prescribed in
27.1557(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)
Each fuel tank must be vented from the top part of
the expansion space so that venting is effective
under all normal flight conditions. Each vent must
minimize the probability of stoppage by dirt or ice.
(b)
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 operation, or a survivable impact.
(a)
There must be a fuel stainer for the fuel tank
outlet or for the booster pump. This strainer must—
(1)
For reciprocating engine powered rotorcraft, have 8
to 16 meshes per inch; and
(2)
For turbine engine powered rotorcraft, 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.
Compliance with 27.955 may not be jeopardized by
failure of—
(a)
Any one pump except pumps that are approved and
installed as parts of a type certificated engine; or
(b)
Any component required for pump operation except,
for engine driven pumps, the engine served by that
pump.
(a)
Each fuel line must be installed and supported to
prevent excessive vibration and to withstand loads
due to fuel pressure 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)
Flexible hose must be approved.
(d)
Each flexible connection in fuel lines that may be
under pressure or subjected to axial loading must
use flexible hose assemblies.
(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.
(a)
There must be a positive, quick-acting valve to shut
off fuel to each engine individually.
(b)
The control for this valve must be within easy reach
of appropriate crewmembers.
(c)
Where there is more than one source of fuel supply
there must be means for independent feeding from
each source.
(d)
No shutoff valve may be on the engine side of any
firewall.
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 itself, 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 fuel system 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
must—
(1)
Discharge clear of all parts of the rotorcraft;
(2)
Have manual or automatic means to assure 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.
27.1011 Engines: General.
(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 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.
(c)
The oil cooling provisions for each engine must be
able to maintain the oil inlet temperature to that
engine at or below the maximum established value.
This must be shown by flight tests.
Each
oil tank must be designed and installed so that—
(a)
It can withstand, without failure, each vibration,
inertia, fluid, and structural load expected in
operation;
(b)
[Reserved]
(c)
Where used with a reciprocating engine, it has an
expansion space of not less than the greater of 10
percent of the tank capacity or 0.5 gallon, and
where used with a turbine engine, it has an
expansion space of not less than 10 percent of the
tank capacity.
(d)
It is impossible to fill the tank expansion space
inadvertently with the rotorcraft in the normal
ground attitude;
(e)
Adequate venting is provided; and
(f)
There are means in the filler opening to prevent oil
overflow from entering the oil tank compartment.
Each
oil tank must be designed and installed so that it
can withstand, without leakage, an internal pressure
of 5 p.s.i., except that each pressurized oil tank
used with a turbine engine must be designed and
installed so that it can withstand, without leakage,
an internal pressure of 5 p.s.i., plus the maximum
operating pressure of the tank.
(a)
Each oil line must be supported to prevent excessive
vibration.
(b)
Each oil line connected to components of the
rotorcraft between which relative motion could exist
must have provisions for flexibility.
(c)
Flexible hose must be approved.
(d)
Each oil line must have an inside diameter of not
less than the inside diameter of the engine inlet or
outlet. No line may have splices between
connections.
(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 27.1305(r).
(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)
The lubrication system for components of the rotor
drive system that requires continuous lubrication
must be sufficiently independent of the lubrication
systems of the engine(s) to ensure lubrication
during autorotation.
(b)
Pressure lubrication systems for transmissions and
gearboxes must comply with the engine oil system
requirements of 27.1013 (except paragraph (c)),
27.1015, 27.1017, 27.1021, and 27.1337(d).
(c)
Each pressure lubrication system must have an oil
strainer or filter through which all of the
lubricant flows and must—
(1)
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;
(2)
Be equipped with a means to indicate collection of
contaminants on the filter or strainer at or before
opening of the bypass required by paragraph (c)(3)
of this section; and
(3)
Be equipped with a bypass constructed and installed
so that—
(i)
The lubricant will flow at the normal rate through
the rest of the system with the strainer or filter
completely blocked; and
(ii)
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.
(d)
For each lubricant tank or sump outlet supplying
lubrication to rotor drive systems and rotor drive
system components, a screen must be provided 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
(c) of this section. The requirements of paragraph
(c) do not apply to screens installed at lubricant
tank or sump outlets.
(e)
Splash-type lubrication systems for rotor drive
system gearboxes must comply with 27.1021 and
27.1337(d).
Cooling
27.1041 General.
(a)
Each powerplant cooling system must be able to
maintain the temperatures of powerplant components
within the limits established for these components
under critical surface (ground or water) and flight
operating conditions for which certification is
required and after normal shutdown. Powerplant
components to be considered include but may not be
limited to engines, rotor drive system components,
auxiliary power units, and the cooling or
lubricating fluids used with these components.
(b)
Compliance with paragraph (a) of this section must
be shown in tests conducted under the conditions
prescribed in that paragraph.
(a)
General. For the tests prescribed in
27.1041(b), 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)
For reciprocating engines, the fuel used during the
cooling tests must be of the minimum grade approved
for the engines, and the mixture settings must be
those normally used in the flight stages for which
the cooling tests are conducted.
(4)
The test procedures must be as prescribed in
27.1045.
(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 power-plant
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)
General. For each stage of flight, the
cooling tests must be conducted with the rotorcraft—
(1)
In the configuration most critical for cooling; and
(2)
Under the conditions most critical for cooling.
(b)
Temperature stabilization. For the purpose of
the cooling tests, a temperature is “stabilized”
when its rate of change is less than two degrees F.
per minute. The following component and engine fluid
temperature stabilization rules apply:
(1)
For each rotorcraft, and for each stage of flight—
(i)
The temperatures must be stabilized under the
conditions from which entry is made into the stage
of flight being investigated; or
(ii)
If the entry condition normally does not allow
temperatures to stabilize, operation through the
full entry condition must be conducted before entry
into the stage of flight being investigated in order
to allow the temperatures to attain their natural
levels at the time of entry.
(2)
For each helicopter during the take-off stage of
flight, the climb at take-off power must be preceded
by a period of hover during which the temperatures
are stabilized.
(c)
Duration of test. For each stage of flight
the tests must be continued until—
(1)
The temperatures stabilize or 5 minutes after the
occurrence of the highest temperature recorded, as
appropriate to the test condition;
(2)
That stage of flight is completed; or
(3)
An operating limitation is reached.
27.1091 Air
induction.
(a)
The air induction system for each engine must supply
the air required by that engine under the operating
conditions and maneuvers for which certification is
requested.
(b)
Each cold air induction system opening must be
outside the cowling if backfire flames can emerge.
(c)
If fuel can accumulate in any air induction system,
that system must have drains that discharge fuel—
(1)
Clear of the rotorcraft; and
(2)
Out of the path of exhaust flames.
(d)
For turbine engine powered rotorcraft—
(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 intake system; and
(2)
The air inlet ducts must be located or protected so
as to minimize the ingestion of foreign matter
during take-off, 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 degrees F.,
and with the engines at 75 percent of maximum
continuous power—
(1)
Each rotorcraft with sea level engines using
conventional venturi carburetors has a preheater
that can provide a heat rise of 90 degrees F.;
(2)
Each rotorcraft with sea level engines using
carburetors tending to prevent icing has a sheltered
alternate source of air, and that the preheat
supplied to the alternate air intake is not less
than that provided by the engine cooling air
downstream of the cylinders;
(3)
Each rotorcraft with altitude engines using
conventional venturi carburetors has a preheater
capable of providing a heat rise of 120 degrees F.;
and
(4)
Each rotorcraft with altitude engines using
carburetors tending to prevent icing has a preheater
that can provide a heat rise of—
(i)
100 degrees F.; or
(ii)
If a fluid deicing system is used, at least 40
degrees F.
(b)
Turbine engine. (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 Part 29
of this chapter; 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 gram per cubic meter in the form of drops having
a mean effective diameter not less than 20 microns,
followed by momentary operation at take-off 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 superchargers 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
operating condition because of supercharging.
For
each exhaust system—
(a)
There must be means for thermal expansion of
manifolds and pipes;
(b)
There must be means to prevent local hot spots;
(c)
Exhaust gases must discharge clear of the engine air
intake, fuel system components, and drains;
(d)
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;
(e)
Exhaust gases may not impair pilot vision at night
due to glare;
(f)
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;
(g)
Each exhaust heat exchanger must incorporate means
to prevent blockage of the exhaust port after any
internal heat exchanger failure.
(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 operations.
(c)
Exhaust piping connected to components between which
relative motion could exist must have provisions for
flexibility.
(a)
Powerplant controls must be located and arranged
under 27.777 and marked under 27.1555.
(b)
Each flexible powerplant control must be approved.
(c)
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.
(d)
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.
(e)
For turbine engine powered rotorcraft, no single
failure or malfunction, or probable combination
thereof, in any powerplant control system may cause
the failure of any powerplant function necessary for
safety.
(a)
There must be a separate power control for each
engine.
(b)
Power controls must be grouped and arranged to
allow—
(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)
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.
(e)
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)
There must be means to quickly shut off all ignition
by the grouping of switches or by a master ignition
control.
(b)
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.
If
there are mixture controls, each engine must have a
separate control and the controls must be arranged
to allow—
(a)
Separate control of each engine; and
(b)
Simultaneous control of all engines.
(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
take-off.
(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)
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)
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 must be fire resistant, except that
flammable fluid tanks and supports which are part of
and attached to the engine 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) 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.
(c)
Each flammable fluid drain and vent must discharge
clear of the induction system air inlet.
(a)
Each fuel tank must be isolated from the engines by
a firewall or shroud.
(b)
Each tank or reservoir, other than a fuel tank, that
is part of a system containing flammable fluids or
gases must be isolated from the engine by a firewall
or shroud, unless 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
isolated from the engines.
(c)
There must be at least one-half inch of clear
airspace between each tank and each firewall or
shroud isolating that tank, unless equivalent means
are used to prevent heat transfer from each engine
compartment to the flammable fluid.
(d)
Absorbent materials close to flammable fluid system
components that might leak must be covered or
treated to prevent the absorption of hazardous
quantities of fluids.
Each
compartment containing any part of the power plant
installation must have provision for ventilation and
drainage of flammable fluids. The drainage means
must be—
(a)
Effective under conditions expected to prevail when
drainage is needed, and
(b)
Arranged so that no discharged fluid will cause an
additional fire hazard.
(a)
There must be means to shut off each line carrying
flammable fluids into the engine compartment,
except—
(1)
Lines, fittings, and components forming an integral
part of an engine;
(2)
For oil systems for which all components of the
system, including oil tanks, are fireproof or
located in areas not subject to engine fire
conditions; and
(3)
For reciprocating engine installations only, engine
oil system lines in installation using engines of
less than 500 cu. in. displacement.
(b)
There must be means to guard against inadvertent
operation of each shut-off, and to make it possible
for the crew to reopen it in flight after it has
been closed.
(c)
Each shut-off valve and its control must be
designed, located, and protected to function
properly under any condition likely to result from
an engine fire.
(a)
Each engine, including the combustor, turbine, and
tailpipe sections of turbine engines 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 a controlled landing: and
(2)
Not protected under 27.861.
(b)
Each auxiliary power unit and combustion heater, and
any other combustion equipment to be used in flight,
must be isolated from the rest of the rotorcraft by
firewalls, shrouds, or equivalent means.
(c)
In meeting paragraphs (a) and (b) of this section,
account must be taken of the probable path of a fire
as affected by the airflow in normal flight and in
autorotation.
(d)
Each firewall and shroud must be constructed so that
no hazardous quantity of air, fluids, or flame can
pass from any engine compartment to other parts of
the rotorcraft.
(e)
Each opening in the firewall or shroud must be
sealed with close-fitting, fireproof grommets,
bushings, or firewall fittings.
(f)
Each firewall and shroud must be fireproof and
protected against corrosion.
(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)
There must be means for rapid and complete drainage
of each part of the cowling or engine compartment in
the normal ground and flight attitudes.
(c)
No drain may discharge where it might cause a fire
hazard.
(d)
Each cowling and engine compartment covering must be
at least fire resistant.
(e)
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.
(f)
A means of retaining 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 structural or mechanical
failure of the normal retention means, unless such
failure is extremely improbable.
27.1194 Other
surfaces.
All
surfaces aft of, and near, powerplant compartments,
other than tail surfaces not subject to heat,
flames, or sparks emanating from a powerplant
compartment, must be at least fire resistant.
Each
turbine engine powered rotorcraft must have approved
quick-acting fire detectors in numbers and locations
insuring prompt detection of fire in the engine
compartment which cannot be readily observed in
flight by the pilot in the cockpit.
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 the required flight and navigation
instruments:
(a)
An airspeed indicator.
(b)
An altimeter.
(c)
A magnetic direction indicator.
The
following are the required power-plant instruments:
(a)
A carburetor air temperature indicator, for each
engine having a pre-heater that can provide a heat
rise in excess of 60 °F.
(b)
A cylinder head temperature indicator, for each—
(1)
Air cooled engine;
(2)
Rotorcraft with cooling shutters; and
(3)
Rotorcraft for which compliance with 27.1043 is
shown in any condition other than the most critical
flight condition with respect to cooling.
(c)
A fuel pressure indicator, for each pump-fed engine.
(d)
A fuel quantity indicator, for each fuel tank.
(e)
A manifold pressure indicator, for each altitude
engine.
(f)
An oil temperature warning device to indicate when
the temperature exceeds a safe value in each main
rotor drive gearbox (including any gearboxes
essential to rotor phasing) having an oil system
independent of the engine oil system.
(g)
An oil pressure warning device to indicate when the
pressure falls below a safe value in each
pressure-lubricated main rotor drive gearbox
(including any gearboxes essential to rotor phasing)
having an oil system independent of the engine oil
system.
(h)
An oil pressure indicator for each engine.
(i)
An oil quantity indicator for each oil tank.
(j)
An oil temperature indicator for each engine.
(k)
At least one tachometer to indicate the r.p.m. of
each engine and, as applicable—
(1)
The r.p.m. of the single main rotor;
(2)
The common r.p.m. of any main rotors whose speeds
cannot vary appreciably with respect to each other;
or
(3)
The r.p.m. of each main rotor whose speed can vary
appreciably with respect to that of another main
rotor.
(l)
A low fuel warning device for each fuel tank which
feeds an engine. This device must—
(1)
Provide a warning to the flight-crew when
approximately 10 minutes of usable fuel remains in
the tank; and
(2)
Be independent of the normal fuel quantity
indicating system.
(m)
Means to indicate to the flight-crew the failure of
any fuel pump installed to show compliance with
§27.955.
(n)
A gas temperature indicator for each turbine engine.
(o)
Means to enable the pilot to determine the torque of
each turbo-shaft engine, if a torque limitation is
established for that engine under 27.1521(e).
(p)
For each turbine engine, an indicator to indicate
the functioning of the power-plant ice protection
system.
(q)
An indicator for the fuel filter required by 27.997
to indicate the occurrence of contamination of the
filter at the degree established by the applicant in
compliance with 27.955.
(r)
For each turbine engine, a warning means for the oil
strainer or filter required by 27.1019, 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
27.1019(a)(2).
(s)
An indicator to indicate the functioning of any
selectable or controllable heater used to prevent
ice clogging of fuel system components.
(t)
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 the 2-minute OEI power levels, when
the event begins, and when the time interval
expires.
(u)
For each turbine engine utilizing 30-second/2-minute
OEI power, a device or system must be provided for
use by ground personnel which—
(1)
Automatically records each usage and duration of
power at the 30-second and 2-minute OEI levels;
(2)
Permits retrieval of the recorded data;
(3)
Can be reset only by ground maintenance personnel;
and
(4)
Has a means to verify proper operation of the system
or device.
(v)
Warning or caution devices to signal to the flight
crew when ferromagnetic particles are detected by
the chip detector required by 27.1337(e).
The
following is the required miscellaneous equipment:
(a)
An approved seat for each occupant.
(b)
An approved safety belt for each occupant.
(c)
A master switch arrangement.
(d)
An adequate source of electrical energy, where
electrical energy is necessary for operation of the
rotorcraft.
(e)
Electrical protective devices.
(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 equipment, systems, and installations of a
multiengine rotorcraft must be designed to prevent
hazards to the rotorcraft in the event of a probable
malfunction or failure.
(c)
The equipment, systems, and installations of
single-engine rotorcraft must be designed to
minimize hazards to the rotorcraft in the event of a
probable malfunction or failure.
(d)
In showing compliance with paragraph (a), (b), or
(c) of this section, the effects of lightning
strikes on the rotorcraft must be considered in
accordance with 27.610.
27.1317 High-intensity Radiated Fields (HIRF)
Protection.
(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 D 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 D 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 D 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 D 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 D
to this part.
(c)
Each electrical and electronic system that performs
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 D 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.
27.1321 Arrangement and visibility.
(a)
Each flight, navigation, and power plant instrument
for use by any pilot must be easily visible to him.
(b)
For each multiengine rotorcraft, identical power
plant instruments must be located so as to prevent
confusion as to which engine each instrument
relates.
(c)
Instrument panel vibration may not damage, or impair
the readability or accuracy of, any instrument.
(d)
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.
(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)
The airspeed indicating system must be calibrated in
flight at forward speeds of 20 knots and over.
(c)
At each forward speed above 80 percent of the
climb-out speed, the airspeed indicator must
indicate true airspeed, at sea level with a standard
atmosphere, to within an allowable installation
error of not more than the greater of—
(1)
±3 percent of the calibrated airspeed; or
(2)
Five knots.
(a)
Each instrument with static air case connections
must be vented so that the influence of rotorcraft
speed, the opening and closing of windows, airflow
variation, and moisture or other foreign matter does
not seriously affect its accuracy.
(b)
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 the altimeter when on the
primary static system by more than 50 feet, a
correction card must be provided for the alternate
static system.
(c)
Except as provided in paragraph (d) 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.
(d)
For un pressurized rotorcraft, paragraph (c)(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)
Except as provided in paragraph (b) of this section—
(1)
Each magnetic direction indicator must be installed
so that its accuracy is not excessively affected by
the rotorcraft's vibration or magnetic fields; and
(2)
The compensated installation may not have a
deviation, in level flight, greater than 10 degrees
on any heading.
(b)
A magnetic non-stabilized direction indicator may
deviate more than 10 degrees due to the operation of
electrically powered systems such as electrically
heated windshields if either a magnetic stabilized
direction indicator, which does not have a deviation
in level flight greater than 10 degrees on any
heading, or a gyroscopic direction indicator, is
installed. Deviations of a magnetic non-stabilized
direction indicator of more than 10 degrees must be
placarded in accordance with 27.1547(e).
(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 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.
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.
27.1337 Powerplant
instruments.
(a)
Instruments and instrument lines. (1) Each
powerplant instrument line must meet the
requirements of 27.- 961 and 27.993.
(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 power-plant 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. Each fuel quantity
indicator must be installed to clearly indicate to
the flight crew the quantity of fuel in each tank in
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 27.959;
(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;
and
(3)
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 means
to indicate the quantity of oil in each tank—
(1)
On the ground (including during the filling of each
tank); and
(2)
In flight, if there is an oil transfer system or
reserve oil supply system.
(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. Chip detectors must—
(1)
Be designed to provide a signal to the device
required by 27.1305(v) and be provided with a means
to allow crewmembers to check, in flight, the
function of each detector electrical circuit and
signal.
(2)
[Reserved]
(a)
Electrical system capacity. Electrical
equipment must be adequate for its intended use. In
addition—
(1)
Electric power sources, their transmission cables,
and their associated control and protective devices
must be able to furnish the required power at the
proper voltage to each load circuit essential for
safe operation; and
(2)
Compliance with paragraph (a)(1) of this section
must be shown by an electrical load analysis, or by
electrical measurements that take into account the
electrical loads applied to the electrical system,
in probable combinations and for probable durations.
(b)
Function. For each electrical system, the
following apply:
(1)
Each system, when installed, must be—
(i)
Free from hazards in itself, in its method of
operation, and in its effects on other parts of the
rotorcraft; and
(ii)
Protected from fuel, oil, water, other detrimental
substances, and mechanical damage.
(2)
Electric power sources must function properly when
connected in combination or independently.
(3)
No failure or malfunction of any source may impair
the ability of any remaining source to supply load
circuits essential for safe operation.
(4)
Each electric power source control must allow the
independent operation of each source.
(c)
Generating system. There must be at least one
generator if the system supplies power to load
circuits essential for safe operation. In addition—
(1)
Each generator must be able to deliver its
continuous rated power;
(2)
Generator voltage control equipment must be able to
dependably regulate each generator output within
rated limits;
(3)
Each generator must have a reverse current cutout
designed to disconnect the generator from the
battery and from the other generators when enough
reverse current exists to damage that generator; and
(4)
Each generator must have an overvoltage control
designed and installed to prevent damage to the
electrical system, or to equipment supplied by the
electrical system, that could result if that
generator were to develop an overvoltage condition.
(d)
Instruments. There must be means to indicate
to appropriate crewmembers the electric power system
quantities essential for safe operation of the
system. In addition—
(1)
For direct current systems, an ammeter that can be
switched into each generator feeder may be used; and
(2)
If there is only one generator, the ammeter may be
in the battery feeder.
(e)
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.
(a)
Each storage battery must be designed and installed
as prescribed in this section.
(b)
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)—
(1)
At maximum regulated voltage or power;
(2)
During a flight of maximum duration; and
(3)
Under the most adverse cooling condition likely to
occur in service.
(c)
Compliance with paragraph (b) 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.
(d)
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.
(e)
No corrosive fluids or gases that may escape from
the battery may damage surrounding structures or
adjacent essential equipment.
(f)
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.
(g)
Nickel cadmium battery installations capable of
being used to start an engine or auxiliary power
unit must have—
(1)
A system to control the charging rate of the battery
automatically so as to prevent battery overheating;
(2)
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
(3)
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)
Protective devices, such as fuses or circuit
breakers, must be installed in each electrical
circuit other than—
(1)
The main circuits of starter motors; and
(2)
Circuits in which no hazard is presented by their
omission.
(b)
A protective device for a circuit essential to
flight safety may not be used to protect any other
circuit.
(c)
Each resettable circuit protective device (“trip
free” device in which the tripping mechanism cannot
be overridden by the operating control) must be
designed so that—
(1)
A manual operation is required to restore service
after trippling; and
(2)
If an overload or circuit fault exists, the device
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)
If fuses are used, there must be one spare of each
rating, or 50 percent spare fuses of each rating,
whichever is greater.
(a)
There must be a master switch arrangement to allow
ready disconnection of each electric power source
from the main bus. The point of disconnection must
be adjacent to the sources controlled by the switch.
(b)
Load circuits may be connected so that they remain
energized after the switch is opened, if they are
protected by circuit protective devices, rated at
five amperes or less, adjacent to the electric power
source.
(c)
The master switch or its controls must be installed
so that the switch is easily discernible and
accessible to a crewmember in flight.
(a)
Each electric connecting cable must be of adequate
capacity.
(b)
Each cable that would overheat in the event of
circuit overload or fault must be at least flame
resistant and may not emit dangerous quantities of
toxic fumes.
(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.
Each
switch must be—
(a)
Able to carry its rated current;
(b)
Accessible to the crew; and
(c)
Labeled as to operation and the circuit controlled.
27.1381 Instrument
lights.
The
instrument lights must—
(a)
Make each instrument, switch, and other devices 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.
27.1383 Landing
lights.
(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 27.1387 through 27.1397.
(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 ACAR aft 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 ACAR aft
as practicable in accordance with 25.1385(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
27.1391.
(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 27.1393, where I
is the minimum intensity prescribed in 27.1391
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 27.1395, except
that higher intensities in overlaps may be used with
main beam intensities substantially greater than the
minima specified in 27.1391 and 27.1393, if the
overlap intensities in relation to the main beam
intensities do not adversely affect signal clarity.
When the peak intensity of the forward position
lights is greater than 100 candles, the maximum
overlap intensities between them may exceed the
values in 27.1395 if the overlap intensity in Area A
is not more than 10 percent of peak position light
intensity and the overlap intensity in Area B is not
more than 2.5 percent of peak position light
intensity.
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) |
10° 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, l |
|
0° |
1.00 |
|
0° to 5° |
0.90 |
|
5° to 10° |
0.80 |
|
10° to 15° |
0.70 |
|
15° to 20° |
0.50 |
|
20° to 30° |
0.30 |
|
30° to 40° |
0.10 |
|
40° to 90° |
0.05 |
No
position light intensity may exceed the applicable
values in the following table, except as provided in
27.1389(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 cis 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 nautical 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.
(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 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 27.1397.
(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 1=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)
Required safety equipment to be used by the crew in
an emergency, such as flares and automatic life-raft
releases, must be readily accessible.
(b)
Stowage provisions for required safety 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 damage caused by
being subjected to the inertia loads specified in
27.561.
Each
safety belt must be equipped with a metal to metal
latching device.
(a)
Emergency flotation and signaling equipment required
by any operating rule in this chapter must meet the
requirements of this section.
(b)
Each raft and each life preserver must be approved
and must be installed so that it is readily
available to the crew and passengers. The storage
provisions for life preservers must accommodate one
life preserver for each occupant for which
certification for ditching is requested.
(c)
Each raft released automatically or by the pilot
must be attached to the rotorcraft by a line to keep
it alongside the rotorcraft. This line must be weak
enough to break before submerging the empty raft to
which it is attached.
(d)
Each signaling device must be free from hazard in
its operation and must be installed in an accessible
location.
(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 Part 29 of this chapter 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 indentified 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)
Design. Each hydraulic system and its
elements must withstand, without yielding, any
structural loads expected in addition to hydraulic
loads.
(b)
Tests. Each system must be substantiated by
proof pressure tests. When proof tested, no part of
any system may fail, malfunction, or experience a
permanent set. The proof load of each system must be
at least 1.5 times the maximum operating pressure of
that system.
(c)
Accumulators. No hydraulic accumulator or
pressurized reservoir may be installed on the engine
side of any firewall unless it is an integral part
of an engine.
27.1457 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 adjusted or
supplemented so 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 microphone 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 27.1323, 27.1325, and
27.1327 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 preflight
checking of the recorder for proper recording of
data in the storage medium;
(5)
Except for recorders powered solely by the
engine-driven 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; and
(b)
Each non-ejectable 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 they are 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 27.1503
through 27.1525 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 27.1541 through
27.1589.
(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 27.309;
(ii)
0.9 times the maximum speed shown under 27.251 and
27.629; or
(iii) 0.9 times the maximum speed substantiated for
advancing blade tip mach number effects.
(b)
VNE may 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-off VNE denoted
as VNE(power-off) may be established at a speed less
than VNE established 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-on VNE and the speed used in
meeting the requirements of—
(i)
27.65(b) for single engine helicopters; and
(ii)
27.67 for multiengine helicopters.
(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
§27.309(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 27.33(a)(1)
and (b)(1).
27.1519 Weight and
center of gravity.
The
weight and center of gravity limitations determined
under 27.25 and 27.27, 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)
Take-off operation. The powerplant take-off
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 time limit for the use of the power
corresponding to the limitations established in
paragraphs (b)(1) and (2) of this section;
(4)
If the time limit in paragraph (b)(3) of this
section exceeds two minutes, the maximum allowable
cylinder head, coolant outlet, or oil temperatures;
(5)
The gas temperature limits for turbine engines over
the range of operating and atmospheric conditions
for which certification is requested.
(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 §27.1509(c); and
(3)
The gas temperature limits for turbine engines over
the range of operating and atmospheric conditions
for which certification is requested.
(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)
Turbo-shaft engine torque. For rotorcraft
with main rotors driven by turbo-shaft engines, and
that do not have a torque limiting device in the
transmission system, the following apply:
(1)
A limit engine torque must be established if the
maximum torque that the engine can exert is greater
than—
(i)
The torque that the rotor drive system is designed
to transmit; or
(ii)
The torque that the main rotor assembly is designed
to withstand in showing compliance with 27.547(e).
(2)
The limit engine torque established under paragraph
(e)(1) of this section may not exceed either torque
specified in paragraph (e)(1)(i) or (ii) of this
section.
(f)
Ambient temperature. For turbine engines,
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 27.1041 through
27.1045 is shown.
(g)
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 after failure of an
engine. 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 demonstrated during the type tests;
(2)
The maximum allowable gas temperature; and
(3)
The maximum allowable torque.
(h)
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 demonstrated during the type
tests;
(2)
The maximum allowable gas temperature; and
(3)
The maximum allowable torque.
(i)
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 demonstrated during the type
tests;
(2)
The maximum allowable gas temperature; and
(3)
The maximum allowable torque.
(j)
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 A27.4 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.
(k)
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 A27.4 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.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–14, 43 FR 2325, Jan. 16, 1978;
Amdt. 27–23, 53 FR 34214, Sept. 2, 1988; Amdt.
27–29, 59 FR 47767, Sept. 16, 1994]
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 §27.1525.
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.
[Amdt.
27–21, 49 FR 44435, Nov. 6, 1984]
The
maximum altitude up to which operation is allowed,
as limited by flight, structural, powerplant,
functional, or equipment characteristics, must be
established.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt.
27–14, 43 FR 2325, Jan. 16, 1978]
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.
[Amdt.
27–18, 45 FR 60177, Sept. 11, 1980]
§ 27.1541 General.
(a)
The rotorcraft must contain—
(1)
The markings and placards specified in §§27.1545
through 27.1565, and
(2)
Any additional information, instrument markings, and
placards required for the safe operation of
rotorcraft with 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 rotocraft other than helicopters, at VNE; and
(ii)
For helicopters at 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.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–14, 43 FR 2325, Jan. 16, 1978;
43 FR 3900, Jan. 30, 1978; Amdt. 27–16, 43 FR 50599,
Oct. 30, 1978]
(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.
(e)
If a magnetic nonstabilized direction indicator can
have a deviation of more than 10 degrees caused by
the operation of electrical equipment, the placard
must state which electrical loads, or combination of
loads, would cause a deviation of more than 10
degrees when turned on.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–13, 42 FR 36972, July 18, 1977]
For
each required powerplant instrument, as appropriate
to the type of instrument—
(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 take-off 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.
[Amdt.
27–11, 41 FR 55470, Dec. 20, 1976, as amended by
Amdt. 27–23, 53 FR 34215, Sept. 2, 1988; Amdt.
27–29, 59 FR 47768, Sept. 16, 1994]
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 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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–11, 41 FR 55470, Dec. 20, 1976;
Amdt. 27–21, 49 FR 44435, Nov. 6, 1984]
(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; 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 be
red. A placard must be near each emergency exit
control and must clearly indicate the location of
that exit and its method of operation.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–11, 41 FR 55471, Dec. 20, 1976]
There must be a placard in clear view of the pilot
that specifies the kinds of operations (such as VFR,
IFR, day, night, or icing) for which the rotorcraft
is approved.
[Amdt.
27–21, 49 FR 44435, Nov. 6, 1984]
(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.
Each
tail rotor must be marked so that its disc is
conspicuous under normal daylight ground conditions.
[Amdt.
27–2, 33 FR 965, Jan. 26, 1968]
(a)
Furnishing information. A Rotorcraft Flight
Manual must be furnished with each rotorcraft, and
it must contain the following:
(1)
Information required by §§27.1583 through 27.1589.
(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 §§27.1583 through 27.1589, that is
appropriate to the rotorcraft, must be furnished,
verified, and approved, and must be segregated,
identified, 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.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt.
27–14, 43 FR 2325, Jan. 16, 1978]
(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 §27.1521.
(2)
Explanation of the limitations, when appropriate.
(3)
Information necessary for marking the instruments
required by §§27.1549 through 27.1553.
(c)
Weight and loading distribution. The weight
and center of gravity limits required by §§27.25 and
27.27, 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 §27.1523 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)
[Reserved]
(g)
Altitude. The altitude established under
§27.1527 and an explanation of the limiting factors
must be furnished.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–2, 33 FR 965, Jan. 26, 1968;
Amdt. 27–14, 43 FR 2325, Jan. 16, 1978; Amdt. 27–16,
43 FR 50599, Oct. 30, 1978]
(a)
Parts of the manual containing operating procedures
must have information concerning any normal and
emergency procedures and other information necessary
for safe operation, including take-off and landing
procedures and associated airspeeds. The manual must
contain any pertinent information including—
(1)
The kind of take-off surface used in the tests and
each appropriate climbout speed; and
(2)
The kind of landing surface used in the tests and
appropriate approach and glide airspeeds.
(b)
For multiengine rotorcraft, information identifying
each operating condition in which the fuel system
independence prescribed in §27.953 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 §27.1505(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 §27.1353
(g)(2) or (g)(3), the operating procedures for
disconnecting the battery from its charging source
must be furnished.
(e)
If the unusable fuel supply in any tank exceeds five
percent of the tank capacity, or one 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)
The airspeeds and rotor speeds for minimum rate of
descent and best glide angle as prescribed in §27.71
must be provided.
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Amdt.
27–1, 32 FR 6914, May 5, 1967, as amended by Amdt.
27–14, 43 FR 2326, Jan. 16, 1978; Amdt. 27–16, 43 FR
50599, Oct. 30, 1978; Amdt. 27–21, 49 FR 44435, Nov.
6, 1984]
(a)
The rotorcraft must be furnished with the following
information, determined in accordance with §§27.51
through 27.79 and 27.143(c):
(1)
Enough information to determine the limiting
height-speed envelope.
(2)
Information relative to—
(i)
The hovering ceilings and the steady rates of climb
and descent, as affected by any pertinent factors
such as airspeed, temperature, and altitude;
(ii)
The maximum safe wind for operation near the ground.
If there are combinations of weight, altitude, and
temperature for which performance information is
provided and at which the rotorcraft cannot land and
take-off safely with the maximum wind value, those
portions of the operating envelope and the
appropriate safe wind conditions shall be identified
in the flight manual;
(iii) For reciprocating engine-powered rotorcraft,
the maximum atmospheric temperature at which
compliance with the cooling provisions of §§27.1041
through 27.1045 is shown; and
(iv)
Glide distance as a function of altitude when
autorotating at the speeds and conditions for
minimum rate of descent and best glide as determined
in §27.71.
(b)
The Rotorcraft Flight Manual must contain—
(1)
In its performance information section any pertinent
information concerning the take-off weights and
altitudes used in compliance with §27.51; and
(i)
Any pertinent information concerning the take-off
procedure, including the kind of take-off surface
used in the tests and each appropriate climb- out
speed; and
(ii)
Any pertinent landing procedures, including the kind
of landing surface used in the tests and appropriate
approach and glide airspeeds; and
(2)
The horizontal take-off distance determined in
accordance with §27.65(a)(2)(i).
(Secs.
313(a), 601, 603, 604, and 605 of the AFRO-CAA
member StatesAct of 1958 (49 U.S.C. 1354(a), 1421,
1423, 1424, and 1425); and sec. 6(c) of the Dept. of
Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–14, 43 FR 2326, Jan. 16, 1978;
Amdt. 27–21, 49 FR 44435, Nov. 6, 1984]
There must be loading instructions for each possible
loading condition between the maximum and minimum
weights determined under §27.25 that can result in a
center of gravity beyond any extreme prescribed in
§27.27, assuming any probable occupant weights.
A27.1 General.
(a)
This appendix specifies requirements for the
preparation of Instructions for Continued
Airworthiness as required by §27.1529.
(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 AFRO-CAA 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.
A27.2 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.
A27.3 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 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 problem
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, discarded recommendations, and
torque values.
(g)
A list of special tools needed.
A27.4 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 §27.571. 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 AFRO-CAA approved and
specifies inspections and other maintenance required
under §§43.16 and 91.403 of the AFRO-CAA member
StatesRegulations unless an alternative program has
been AFRO-CAA approved.”
[Amdt.
27–17, 45 FR 60178, Sept. 11, 1980, as amended by
Amdt. 27–24, 54 FR 34329, Aug. 18, 1989]
I.
General. A normal 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 paragraph IV
(b) or (c) of this appendix, as appropriate. The
stick force must vary with speed so that any
substantial speed change results in a stick force
clearly perceptible to the pilot. For single-pilot
approval, 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
paragraph IV(b) of the this appendix.
(b)
For single-pilot approval:
(1)
Climb. Stability must be shown in climb
throughout the speed range 20 knots either side of
trim with—
(i)
The helicopter trimmed at VYI;
(ii)
Landing gear retracted (if retractable); and
(iii) Power required for limit climb rate (at least
1,000 fpm) at VYIor maximum continuous
power, whichever is less.
(2)
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—
(i)
The helicopter trimmed and power adjusted for level
flight at 0.9 VHor 0.9 VNEI,
whichever is lower; and
(ii)
Landing gear retracted (if retractable).
(3)
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—
(i)
the helicopter trimmed and power adjusted for level
flight at 1.1 VMINI; and
(ii)
Landing gear retracted (if retractable).
(4)
Descent. Stability must be shown throughout
the speed range 20 knots either side of trim with—
(i)
The helicopter trimmed at 0.8 VHor 0.8 VNEI(or
0.8 VLEfor the landing gear extended
case), whichever is lower;
(ii)
Power required for 1,000 fpm descent at trim speed;
and
(iii) Landing gear extended and retracted, if
applicable.
(5)
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—
(i)
The helicopter trimmed at the recommended approach
speed or speeds;
(ii)
Landing gear extended and retracted, if applicable;
and
(iii) Power required to maintain a 3° glide path and
power required to maintain the steepest approach
gradient for which approval is requested.
(c)
Helicopters approved for a minimum crew of two
pilots must comply with the provisions of paragraphs
IV(b)(2) and IV(b)(5) of this appendix.
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 cyclic movement with
sideslip must not be excessive.
VI.
Dynamic stability. (a) For single-pilot
approval—
(1)
Any oscillation having a period of less than 5
seconds must damp to1/2amplitude in not more than
one cycle.
(2)
Any oscillation having a period of 5 seconds or more
but less than 10 seconds must damp to1/2amplitude in
not more than two cycles.
(3)
Any oscillation having a period of 10 seconds or
more but less than 20 seconds must be damped.
(4)
Any oscillation having a period of 20 seconds or
more may not achieve double amplitude in less than
20 seconds.
(5)
Any aperiodic response may not achieve double
amplitude in less than 6 seconds.
(b)
For helicopters approved with a minimum crew of two
pilots—
(1)
Any oscillation having a period of less than 5
seconds must damp to1/2amplitude in not more than
two cycles.
(2)
Any oscillation having a period of 5 seconds or more
but less than 10 seconds must be damped.
(3)
Any oscillation having a period of 10 seconds or
more may not achieve double amplitude in less than
10 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 27 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 §§29.1303, 29.1431, and 29.1433 through
Amendment 29–14, with the following exceptions and
additions:
(a)
Flight and Navigation Instruments. (1) A
magnetic gyro-stablized direction indicator instead
of a 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).
For two-pilot configurations, one pilot's primary
indicator may be designated for this purpose. If
standby batteries are provided, they may be charged
from the aircraft electrical system if adequate
isolation is incorporated.
(b)
Miscellaneous requirements. (1) Instrument
systems and other systems essential for IFR flight
that could be adversely affected by icing must be
adequately protected when exposed to the continuous
and intermittent maximum icing conditions defined in
appendix C of Part 29 of this chapter, 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 crewmember 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.
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 flightcrew 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 VYIdiffers from VY,
climb performance at VYIand with maximum
continuous power throughout the ranges of weight,
altitude, and temperature for which approval is
requested.
[Amdt.
27–19, 48 FR 4389, Jan. 31, 1983]
C27.1 General.
A
small multiengine rotorcraft may not be type
certificated for Category A operation unless it
meets the design installation and performance
requirements contained in this appendix in addition
to the requirements of this part.
C27.2 Applicable part 29 sections. The following
sections of part 29 of this chapter must be met in
addition to the requirements of this part:
29.45(a) and (b)(2)—General.
29.49(a)—Performance at minimum operating speed.
29.51—Take-off data: General.
29.53—Take-off: Category A.
29.55—Take-off decision point: Category A.
29.59—Take-off Path: Category A.
29.60—Elevated heliport take-off path: Category A.
29.61—Take-off distance: Category A.
29.62—Rejected take-off: Category A.
29.64—Climb: General.
29.65(a)—Climb: AEO.
29.67(a)—Climb: OEI.
29.75—Landing: General.
29.77—Landing decision point: Category A.
29.79—Landing: Category A.
29.81—Landing distance (Ground level sites):
Category A.
29.85—Balked landing: Category A.
29.87(a)—Height-velocity envelope.
29.547(a) and (b)—Main and tail rotor structure.
29.861(a)—Fire protection of structure, controls,
and other parts.
29.901(c)—Powerplant: Installation.
29.903(b) (c) and (e)—Engines.
29.908(a)—Cooling fans.
29.917(b) and (c)(1)—Rotor drive system: Design.
29.927(c)(1)—Additional tests.
29.953(a)—Fuel system independence.
29.1027(a)—Transmission and gearboxes: General.
29.1045(a)(1), (b), (c), (d), and (f)—Climb cooling
test procedures.
29.1047(a)—Take-off cooling test procedures.
29.1181(a)—Designated fire zones: Regions included.
29.1187(e)—Drainage and ventilation of fire zones.
29.1189(c)—Shutoff means.
29.1191(a)(1)—Firewalls.
29.1193(e)—Cowling and engine compartment covering.
29.1195(a) and (d)—Fire extinguishing systems (one
shot).
29.1197—Fire extinguishing agents.
29.1199—Extinguishing agent containers.
29.1201—Fire extinguishing system materials.
29.1305(a) (6) and (b)—Powerplant instruments.
29.1309(b)(2) (i) and (d)—Equipment, systems, and
installations.
29.1323(c)(1)—Airspeed indicating system.
29.1331(b)—Instruments using a power supply.
29.1351(d)(2)—Electrical systems and equipment:
General (operation without normal electrical power).
29.1587(a)—Performance information.
Note: In complying with the paragraphs listed in
paragraph C27.2 above, relevant material in the AC
“Certification of Transport Category Rotorcraft”
should be used.
[Doc. No. 28008, 61 FR 21907, May 10, 1996]
This
appendix specifies the HIRF environments and
equipment HIRF test levels for electrical and
electronic systems under §27.1317. 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
decibels (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.
[Doc. No. AFRO-CAA–2006–23657, 72 FR 44027, Aug. 6,
2007]
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