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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 | 
