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29.1 Applicability.
(a)
This part prescribes airworthiness standards for the
issue of type certificates, and changes to those
certificates, for transport category rotorcraft.
(b)
Transport category rotorcraft must be certificated
in accordance with either the Category A or Category
B requirements of this part. A multiengine
rotorcraft may be type certificated as both Category
A and Category B with appropriate and different
operating limitations for each category.
(c)
Rotorcraft with a maximum weight greater than 20,000
pounds and 10 or more passenger seats must be type
certificated as Category A rotorcraft.
(d)
Rotorcraft with a maximum weight greater than 20,000
pounds and nine or less passenger seats may be type
certificated as Category B rotorcraft provided the
Category A requirements of Subparts C, D, E, and F
of this part are met.
(e)
Rotorcraft with a maximum weight of 20,000 pounds or
less but with 10 or more passenger seats may be type
certificated as Category B rotorcraft provided the
Category A requirements of 29.67(a)(2), 29.87,
29.1517, and subparts C, D, E, and F of this part
are met.
(f)
Rotorcraft with a maximum weight of 20,000 pounds or
less and nine or less passenger seats may be type
certificated as Category B rotorcraft.
(g)
Each person who applies under Part 21 for a
certificate or change described in paragraphs (a)
through (f) of this section must show compliance
with the applicable requirements of this part.
For
each rotorcraft, each applicant must show that each
occupant's seat is equipped with a safety belt and
shoulder harness that meets the requirements of
paragraphs (a), (b), and (c) of this section.
(a)
Each occupant's seat must have a combined safety
belt and shoulder harness with a single-point
release. Each pilot's combined safety belt and
shoulder harness must allow each pilot, when seated
with safety belt and shoulder harness fastened, to
perform all functions necessary for flight
operations. There must be a means to secure belts
and harnesses, when not in use, to prevent
interference with the operation of the rotorcraft
and with rapid egress in an emergency.
(b)
Each occupant must be protected from serious head
injury by a safety belt plus a shoulder harness that
will prevent the head from contacting any injurious
object.
(c)
The safety belt and shoulder harness must meet the
static and dynamic strength requirements, if
applicable, specified by the rotorcraft type
certification basis.
(d)
For purposes of this section, the date of
manufacture is either—
(1)
The date the inspection acceptance records, or
equivalent, reflect that the rotorcraft is complete
and meets the AFRO-CAA-Approved Type Design Data; or
(2)
The date that the foreign civil airworthiness
authority certifies the rotorcraft is complete and
issues an original standard airworthiness
certificate, or equivalent, in that country.
General
29.21 Proof of
compliance.
Each
requirement of this subpart must be met at each
appropriate combination of weight and center of
gravity within the range of loading conditions for
which certification is requested. This must be
shown—
(a)
By tests upon a rotorcraft of the type for which
certification is requested, or by calculations based
on, and equal in accuracy to, the results of
testing; and
(b)
By systematic investigation of each required
combination of weight and center of gravity, if
compliance cannot be reasonably inferred from
combinations investigated.
(a)
Maximum weight. The maximum weight (the
highest weight at which compliance with each
applicable requirement of this part is shown) or, at
the option of the applicant, the highest weight for
each altitude and for each practicably separable
operating condition, such as takeoff, enroute
operation, and landing, must be established so that
it is not more than—
(1)
The highest weight selected by the applicant;
(2)
The design maximum weight (the highest weight at
which compliance with each applicable structural
loading condition of this part is shown); or
(3)
The highest weight at which compliance with each
applicable flight requirement of this part is shown.
(b)
Minimum weight. The minimum weight (the
lowest weight at which compliance with each
applicable requirement of this part is shown) must
be established so that it is not less than—
(1)
The lowest weight selected by the applicant;
(2)
The design minimum weight (the lowest weight at
which compliance with each structural loading
condition of this part is shown); or
(3)
The lowest weight at which compliance with each
applicable flight requirement of this part is shown.
(c)
Total weight with jettisonable external load.
A total weight for the rotorcraft with a
jettisonable external load attached that is greater
than the maximum weight established under paragraph
(a) of this section may be established for any
rotorcraft-load combination if—
(1)
The rotorcraft-load combination does not include
human external cargo,
(2)
Structural component approval for external load
operations under either 29.865 or under equivalent
operational standards is obtained,
(3)
The portion of the total weight that is greater than
the maximum weight established under paragraph (a)
of this section is made up only of the weight of all
or part of the jettisonable external load,
(4)
Structural components of the rotorcraft are shown to
comply with the applicable structural requirements
of this part under the increased loads and stresses
caused by the weight increase over that established
under paragraph (a) of this section, and
(5)
Operation of the rotorcraft at a total weight
greater than the maximum certificated weight
established under paragraph (a) of this section is
limited by appropriate operating limitations under
29.865 (a) and (d) of this part.
The
extreme forward and aft centers of gravity and,
where critical, the extreme lateral centers of
gravity must be established for each weight
established under 29.25. Such an extreme may not lie
beyond—
(a)
The extremes selected by the applicant;
(b)
The extremes within which the structure is proven;
or
(c)
The extremes within which compliance with the
applicable flight requirements is shown.
(a)
The empty weight and corresponding center of gravity
must be determined by weighing the rotorcraft
without the crew and payload, but with—
(1)
Fixed ballast;
(2)
Unusable fuel; and
(3)
Full operating fluids, including—
(i)
Oil;
(ii)
Hydraulic fluid; and
(iii) Other fluids required for normal operation of
rotorcraft systems, except water intended for
injection in the engines.
(b)
The condition of the rotorcraft at the time of
determining empty weight must be one that is well
defined and can be easily repeated, particularly
with respect to the weights of fuel, oil, coolant,
and installed equipment.
Removable ballast may be used in showing compliance
with the flight requirements of this subpart.
(a)
Main rotor speed limits. A range of main
rotor speeds must be established that—
(1)
With power on, provides adequate margin to
accommodate the variations in rotor speed occurring
in any appropriate maneuver, and is consistent with
the kind of governor or synchronizer used; and
(2)
With power off, allows each appropriate autorotative
maneuver to be performed throughout the ranges of
airspeed and weight for which certification is
requested.
(b)
Normal main rotor high pitch limit (power on).
For rotorcraft, except helicopters required to
have a main rotor low speed warning under paragraph
(e) of this section, it must be shown, with power on
and without exceeding approved engine maximum
limitations, that main rotor speeds substantially
less than the minimum approved main rotor speed will
not occur under any sustained flight condition. This
must be met by—
(1)
Appropriate setting of the main rotor high pitch
stop;
(2)
Inherent rotorcraft characteristics that make unsafe
low main rotor speeds unlikely; or
(3)
Adequate means to warn the pilot of unsafe main
rotor speeds.
(c)
Normal main rotor low pitch limit (power off).
It must be shown, with power off, that—
(1)
The normal main rotor low pitch limit provides
sufficient rotor speed, in any autorotative
condition, under the most critical combinations of
weight and airspeed; and
(2)
It is possible to prevent overspeeding of the rotor
without exceptional piloting skill.
(d)
Emergency high pitch. If the main rotor high
pitch stop is set to meet paragraph (b)(1) of this
section, and if that stop cannot be exceeded
inadvertently, additional pitch may be made
available for emergency use.
(e)
Main rotor low speed warning for helicopters.
For each single engine helicopter, and each
multiengine helicopter that does not have an
approved device that automatically increases power
on the operating engines when one engine fails,
there must be a main rotor low speed warning which
meets the following requirements:
(1)
The warning must be furnished to the pilot in all
flight conditions, including power-on and power-off
flight, when the speed of a main rotor approaches a
value that can jeopardize safe flight.
(2)
The warning may be furnished either through the
inherent aerodynamic qualities of the helicopter or
by a device.
(3)
The warning must be clear and distinct under all
conditions, and must be clearly distinguishable from
all other warnings. A visual device that requires
the attention of the crew within the cockpit is not
acceptable by itself.
(4)
If a warning device is used, the device must
automatically deactivate and reset when the
low-speed condition is corrected. If the device has
an audible warning, it must also be equipped with a
means for the pilot to manually silence the audible
warning before the low-speed condition is corrected.
29.45 General.
(a)
The performance prescribed in this subpart must be
determined—
(1)
With normal piloting skill and;
(2)
Without exceptionally favorable conditions.
(b)
Compliance with the performance requirements of this
subpart must be shown—
(1)
For still air at sea level with a standard
atmosphere and;
(2)
For the approved range of atmospheric variables.
(c)
The available power must correspond to engine power,
not exceeding the approved power, less—
(1)
Installation losses; and
(2)
The power absorbed by the accessories and services
at the values for which certification is requested
and approved.
(d)
For reciprocating engine-powered rotorcraft, the
performance, as affected by engine power, must be
based on a relative humidity of 80 percent in a
standard atmosphere.
(e)
For turbine engine-powered rotorcraft, the
performance, as affected by engine power, must be
based on a relative humidity of—
(1)
80 percent, at and below standard temperature; and
(2)
34 percent, at and above standard temperature plus
50 °F.
Between these two temperatures, the relative
humidity must vary linearly.
(f)
For turbine-engine-power rotorcraft, a means must be
provided to permit the pilot to determine prior to
takeoff that each engine is capable of developing
the power necessary to achieve the applicable
rotorcraft performance prescribed in this subpart.
(a)
For each Category A helicopter, the hovering
performance must be determined over the ranges of
weight, altitude, and temperature for which takeoff
data are scheduled—
(1)
With not more than takeoff power;
(2)
With the landing gear extended; and
(3)
At a height consistent with the procedure used in
establishing the takeoff, climbout, and rejected
takeoff paths.
(b)
For each Category B helicopter, the hovering
performance must be determined over the ranges of
weight, altitude, and temperature for which
certification is requested, with—
(1)
Takeoff power;
(2)
The landing gear extended; and
(3)
The helicopter in ground effect at a height
consistent with normal takeoff procedures.
(c)
For each helicopter, the out-of-ground effect
hovering performance must be determined over the
ranges of weight, altitude, and temperature for
which certification is requested with takeoff power.
(d)
For rotorcraft other than helicopters, the steady
rate of climb at the minimum operating speed must be
determined over the ranges of weight, altitude, and
temperature for which certification is requested
with—
(1)
Takeoff power; and
(2)
The landing gear extended.
(a)
The takeoff data required by 29.53, 29.55, 29.59,
29.60, 29.61, 29.62, 29.63, and 29.67 must be
determined—
(1)
At each weight, altitude, and temperature selected
by the applicant; and
(2)
With the operating engines within approved operating
limitations.
(b)
Takeoff data must—
(1)
Be determined on a smooth, dry, hard surface; and
(2)
Be corrected to assume a level takeoff surface.
(c)
No takeoff made to determine the data required by
this section may require exceptional piloting skill
or alertness, or exceptionally favorable conditions.
The
takeoff performance must be determined and scheduled
so that, if one engine fails at any time after the
start of takeoff, the rotorcraft can—
(a)
Return to, and stop safely on, the takeoff area; or
(b)
Continue the takeoff and climbout, and attain a
configuration and airspeed allowing compliance with
29.67(a)(2).
(a)
The TDP is the first point from which a continued
takeoff capability is assured under 29.59 and is the
last point in the takeoff path from which a rejected
takeoff is assured within the distance determined
under 29.62.
(b)
The TDP must be established in relation to the
takeoff path using no more than two parameters;
e.g., airspeed and height, to designate the TDP.
(c)
Determination of the TDP must include the pilot
recognition time interval following failure of the
critical engine.
(a)
The takeoff path extends from the point of
commencement of the takeoff procedure to a point at
which the rotorcraft is 1,000 feet above the takeoff
surface and compliance with 29.67(a)(2) is shown. In
addition—
(1)
The takeoff path must remain clear of the
height-velocity envelope established in accordance
with 29.87;
(2)
The rotorcraft must be flown to the engine failure
point; at which point, the critical engine must be
made inoperative and remain inoperative for the rest
of the takeoff;
(3)
After the critical engine is made inoperative, the
rotorcraft must continue to the takeoff decision
point, and then attain VTOSS;
(4)
Only primary controls may be used while attaining VTOSSand
while establishing a positive rate of climb.
Secondary controls that are located on the primary
controls may be used after a positive rate of climb
and VTOSSare established but in no case
less than 3 seconds after the critical engine is
made inoperative; and
(5)
After attaining VTOSSand a positive rate
of a climb, the landing gear may be retracted.
(b)
During the takeoff path determination made in
accordance with paragraph (a) of this section and
after attaining VTOSSand a positive rate
of climb, the climb must be continued at a speed as
close as practicable to, but not less than, VTOSSuntil
the rotorcraft is 200 feet above the takeoff
surface. During this interval, the climb performance
must meet or exceed that required by 29.67(a)(1).
(c)
During the continued takeoff, the rotorcraft shall
not descend below 15 feet above the takeoff surface
when the takeoff decision point is above 15 feet.
(d)
From 200 feet above the takeoff surface, the
rotorcraft takeoff path must be level or positive
until a height 1,000 feet above the takeoff surface
is attained with not less than the rate of climb
required by 29.67(a)(2). Any secondary or auxiliary
control may be used after attaining 200 feet above
the takeoff surface.
(e)
Takeoff distance will be determined in accordance
with 29.61.
(a)
The elevated heliport takeoff path extends from the
point of commencement of the takeoff procedure to a
point in the takeoff path at which the rotorcraft is
1,000 feet above the takeoff surface and compliance
with 29.67(a)(2) is shown. In addition—
(1)
The requirements of 29.59(a) must be met;
(2)
While attaining VTOSSand a positive rate
of climb, the rotorcraft may descend below the level
of the takeoff surface if, in so doing and when
clearing the elevated heliport edge, every part of
the rotorcraft clears all obstacles by at least 15
feet;
(3)
The vertical magnitude of any descent below the
takeoff surface must be determined; and
(4)
After attaining VTOSSand a positive rate
of climb, the landing gear may be retracted.
(b)
The scheduled takeoff weight must be such that the
climb requirements of 29.67 (a)(1) and (a)(2) will
be met.
(c)
Takeoff distance will be determined in accordance
with 29.61.
29.61 Takeoff distance: Category A
(a)
The normal takeoff distance is the horizontal
distance along the takeoff path from the start of
the takeoff to the point at which the rotorcraft
attains and remains at least 35 feet above the
takeoff surface, attains and maintains a speed of at
least VTOSS, and establishes a positive
rate of climb, assuming the critical engine failure
occurs at the engine failure point prior to the
takeoff decision point.
(b)
For elevated heliports, the takeoff distance is the
horizontal distance along the takeoff path from the
start of the takeoff to the point at which the
rotorcraft attains and maintains a speed of at least
VTOSSand establishes a positive rate of
climb, assuming the critical engine failure occurs
at the engine failure point prior to the takeoff
decision point.
The
rejected takeoff distance and procedures for each
condition where takeoff is approved will be
established with—
(a)
The takeoff path requirements of 29.59 and 29.60
being used up to the TDP where the critical engine
failure is recognized and the rotorcraft is landed
and brought to a complete stop on the takeoff
surface;
(b)
The remaining engines operating within approved
limits;
(c)
The landing gear remaining extended throughout the
entire rejected takeoff; and
(d)
The use of only the primary controls until the
rotorcraft is on the ground. Secondary controls
located on the primary control may not be used until
the rotorcraft is on the ground. Means other than
wheel brakes may be used to stop the rotorcraft if
the means are safe and reliable and consistent
results can be expected under normal operating
conditions.
The
horizontal distance required to take off and climb
over a 50-foot obstacle must be established with the
most unfavorable center of gravity. The takeoff may
be begun in any manner if—
(a)
The takeoff surface is defined;
(b)
Adequate safeguards are maintained to ensure proper
center of gravity and control positions; and
(c)
A landing can be made safely at any point along the
flight path if an engine fails.
Compliance with the requirements of 29.65 and 29.67
must be shown at each weight, altitude, and
temperature within the operational limits
established for the rotorcraft and with the most
unfavorable center of gravity for each
configuration. Cowl flaps, or other means of
controlling the engine-cooling air supply, will be
in the position that provides adequate cooling at
the temperatures and altitudes for which
certification is requested.
(a)
The steady rate of climb must be determined—
(1)
With maximum continuous power;
(2)
With the landing gear retracted; and
(3)
At Vyfor standard sea level conditions
and at speeds selected by the applicant for other
conditions.
(b)
For each Category B rotorcraft except helicopters,
the rate of climb determined under paragraph (a) of
this section must provide a steady climb gradient of
at least 1:6 under standard sea level conditions.
(a)
For Category A rotorcraft, in the critical takeoff
configuration existing along the takeoff path, the
following apply:
(1)
The steady rate of climb without ground effect, 200
feet above the takeoff surface, must be at least 100
feet per minute for each weight, altitude, and
temperature for which takeoff data are to be
scheduled with—
(i)
The critical engine inoperative and the remaining
engines within approved operating limitations,
except that for rotorcraft for which the use of
30-second/2-minute OEI power is requested, only the
2-minute OEI power may be used in showing compliance
with this paragraph;
(ii)
The landing gear extended; and
(iii) The takeoff safety speed selected by the
applicant.
(2)
The steady rate of climb without ground effect, 1000
feet above the takeoff surface, must be at least 150
feet per minute, for each weight, altitude, and
temperature for which takeoff data are to be
scheduled with—
(i)
The critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, or at 30-minute
OEI power for rotorcraft for which certification for
use of 30-minute OEI power is requested;
(ii)
The landing gear retracted; and
(iii) The speed selected by the applicant.
(3)
The steady rate of climb (or descent) in feet per
minute, at each altitude and temperature at which
the rotorcraft is expected to operate and at any
weight within the range of weights for which
certification is requested, must be determined with—
(i)
The critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, and at 30-minute
OEI power for rotorcraft for which certification for
the use of 30-minute OEI power is requested;
(ii)
The landing gear retracted; and
(iii) The speed selected by the applicant.
(b)
For multiengine Category B rotorcraft meeting the
Category A engine isolation requirements, the steady
rate of climb (or descent) must be determined at the
speed for best rate of climb (or minimum rate of
descent) at each altitude, temperature, and weight
at which the rotorcraft is expected to operate, with
the critical engine inoperative and the remaining
engines at maximum continuous power including
continuous OEI power, if approved, and at 30-minute
OEI power for rotorcraft for which certification for
the use of 30-minute OEI power is requested.
For
each category B helicopter, except multiengine
helicopters meeting the requirements of 29.67(b) and
the powerplant installation requirements of category
A, the steady angle of glide must be determined in
autorotation—
(a)
At the forward speed for minimum rate of descent as
selected by the applicant;
(b)
At the forward speed for best glide angle;
(c)
At maximum weight; and
(d)
At the rotor speed or speeds selected by the
applicant.
(a)
For each rotorcraft—
(1)
The corrected landing data must be determined for a
smooth, dry, hard, and level surface;
(2)
The approach and landing must not require
exceptional piloting skill or exceptionally
favorable conditions; and
(3)
The landing must be made without excessive vertical
acceleration or tendency to bounce, nose over,
ground loop, porpoise, or water loop.
(b)
The landing data required by 29.77, 29.79, 29.81,
29.83, and 29.85 must be determined—
(1)
At each weight, altitude, and temperature for which
landing data are approved;
(2)
With each operating engine within approved operating
limitations; and
(3)
With the most unfavorable center of gravity.
(a)
The LDP is the last point in the approach and
landing path from which a balked landing can be
accomplished in accordance with 29.85.
(b)
Determination of the LDP must include the pilot
recognition time interval following failure of the
critical engine.
(a)
For Category A rotorcraft—
(1)
The landing performance must be determined and
scheduled so that if the critical engine fails at
any point in the approach path, the rotorcraft can
either land and stop safely or climb out and attain
a rotorcraft configuration and speed allowing
compliance with the climb requirement of
29.67(a)(2);
(2)
The approach and landing paths must be established
with the critical engine inoperative so that the
transition between each stage can be made smoothly
and safely;
(3)
The approach and landing speeds must be selected by
the applicant and must be appropriate to the type of
rotorcraft; and
(4)
The approach and landing path must be established to
avoid the critical areas of the height-velocity
envelope determined in accordance with 29.87.
(b)
It must be possible to make a safe landing on a
prepared landing surface after complete power
failure occurring during normal cruise.
The
horizontal distance required to land and come to a
complete stop (or to a speed of approximately 3
knots for water landings) from a point 50 ft above
the landing surface must be determined from the
approach and landing paths established in accordance
with 29.79.
(a)
For each Category B rotorcraft, the horizontal
distance required to land and come to a complete
stop (or to a speed of approximately 3 knots for
water landings) from a point 50 feet above the
landing surface must be determined with—
(1)
Speeds appropriate to the type of rotorcraft and
chosen by the applicant to avoid the critical areas
of the height-velocity envelope established under
29.87; and
(2)
The approach and landing made with power on and
within approved limits.
(b)
Each multi-engined Category B rotorcraft that meets
the powerplant installation requirements for
Category A must meet the requirements of—
(1)
Sections 29.79 and 29.81; or
(2)
Paragraph (a) of this section.
(c)
It must be possible to make a safe landing on a
prepared landing surface if complete power failure
occurs during normal cruise.
For
Category A rotorcraft, the balked landing path with
the critical engine inoperative must be established
so that—
(a)
The transition from each stage of the maneuver to
the next stage can be made smoothly and safely;
(b)
From the LDP on the approach path selected by the
applicant, a safe climbout can be made at speeds
allowing compliance with the climb requirements of
29.67(a)(1) and (2); and
(c)
The rotorcraft does not descend below 15 feet above
the landing surface. For elevated heliport
operations, descent may be below the level of the
landing surface provided the deck edge clearance of
29.60 is maintained and the descent (loss of height)
below the landing surface is determined.
(a)
If there is any combination of height and forward
velocity (including hover) under which a safe
landing cannot be made after failure of the critical
engine and with the remaining engines (where
applicable) operating within approved limits, a
height-velocity envelope must be established for—
(1)
All combinations of pressure altitude and ambient
temperature for which takeoff and landing are
approved; and
(2)
Weight from the maximum weight (at sea level) to the
highest weight approved for takeoff and landing at
each altitude. For helicopters, this weight need not
exceed the highest weight allowing hovering
out-of-ground effect at each altitude.
(b)
For single-engine or multiengine rotorcraft that do
not meet the Category A engine isolation
requirements, the height-velocity envelope for
complete power failure must be established.
Flight
Characteristics
The
rotorcraft must—
(a)
Except as specifically required in the applicable
section, meet the flight characteristics
requirements of this subpart—
(1)
At the approved operating altitudes and
temperatures;
(2)
Under any critical loading condition within the
range of weights and centers of gravity for which
certification is requested; and
(3)
For power-on operations, under any condition of
speed, power, and rotor r.p.m. for which
certification is requested; and
(4)
For power-off operations, under any condition of
speed, and rotor r.p.m. for which certification is
requested that is attainable with the controls
rigged in accordance with the approved rigging
instructions and tolerances;
(b)
Be able to maintain any required flight condition
and make a smooth transition from any flight
condition to any other flight condition without
exceptional piloting skill, alertness, or strength,
and without danger of exceeding the limit load
factor under any operating condition probable for
the type, including—
(1)
Sudden failure of one engine, for multiengine
rotorcraft meeting Transport Category A engine
isolation requirements;
(2)
Sudden, complete power failure, for other
rotorcraft; and
(3)
Sudden, complete control system failures specified
in 29.695 of this part; and
(c)
Have any additional characteristics required for
night or instrument operation, if certification for
those kinds of operation is requested. Requirements
for helicopter instrument flight are contained in
appendix B of this part.
(a)
The rotorcraft must be safely controllable and
maneuverable—
(1)
During steady flight; and
(2)
During any maneuver appropriate to the type,
including—
(i)
Takeoff;
(ii)
Climb;
(iii) Level flight;
(iv)
Turning flight;
(v)
Glide; and
(vi)
Landing (power on and power off).
(b)
The margin of cyclic control must allow satisfactory
roll and pitch control at VNEwith—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Critical rotor r.p.m.; and
(4)
Power off (except for helicopters demonstrating
compliance with paragraph (e) of this section) and
power on.
(c)
A wind velocity of not less than 17 knots must be
established in which the rotorcraft can be operated
without loss of control on or near the ground in any
maneuver appropriate to the type (such as crosswind
takeoffs, sideward flight, and rearward flight),
with—
(1)
Critical weight;
(2)
Critical center of gravity; and
(3)
Critical rotor r.p.m.
(d)
The rotorcraft, after (1) failure of one engine, in
the case of multiengine rotorcraft that meet
Transport Category A engine isolation requirements,
or (2) complete power failure in the case of other
rotorcraft, must be controllable over the range of
speeds and altitudes for which certification is
requested when such power failure occurs with
maximum continuous power and critical weight. No
corrective action time delay for any condition
following power failure may be less than—
(i)
For the cruise condition, one second, or normal
pilot reaction time (whichever is greater); and
(ii)
For any other condition, normal pilot reaction time.
(e)
For helicopters for which a VNE(power-off) is
established under 29.1505(c), compliance must be
demonstrated with the following requirements with
critical weight, critical center of gravity, and
critical rotor r.p.m.:
(1)
The helicopter must be safely slowed to VNE(power-off),
without exceptional pilot skill after the last
operating engine is made inoperative at power-on VNE.
(2)
At a speed of 1.1 VNE(power-off), the margin of
cyclic control must allow satisfactory roll and
pitch control with power off.
(a)
Longitudinal, lateral, directional, and collective
controls may not exhibit excessive breakout force,
friction, or preload.
(b)
Control system forces and free play may not inhibit
a smooth, direct rotorcraft response to control
system input.
29.161 Trim
control.
The
trim control—
(a)
Must trim any steady longitudinal, lateral, and
collective control forces to zero in level flight at
any appropriate speed; and
(b)
May not introduce any undesirable discontinuities in
control force gradients.
The
rotorcraft must be able to be flown, without undue
pilot fatigue or strain, in any normal maneuver for
a period of time as long as that expected in normal
operation. At least three landings and takeoffs must
be made during this demonstration.
(a)
The longitudinal control must be designed so that a
rearward movement of the control is necessary to
obtain a speed less than the trim speed, and a
forward movement of the control is necessary to
obtain a speed more than the trim speed.
(b)
With the throttle and collective pitch held constant
during the maneuvers specified in 29.175 (a) through
(c), the slope of the control position versus speed
curve must be positive throughout the full range of
altitude for which certification is requested.
(c)
During the maneuver specified in 29.175(d), the
longitudinal control position versus speed curve may
have a negative slope within the specified speed
range if the negative motion is not greater than 10
percent of total control travel.
(a)
Climb. Static longitudinal stability must be
shown in the climb condition at speeds from 0.85 VY,
or 15 knots below VY, whichever is less,
to 1.2 VYor 15 knots above VY,
whichever is greater, with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Maximum continuous power;
(4)
The landing gear retracted; and
(5)
The rotorcraft trimmed at V Y.
(b)
Cruise. Static longitudinal stability must be
shown in the cruise condition at speeds from 0.7
V Hor 0.7 V NE,whichever is less, to 1.1
V Hor 1.1 V
NE,whichever is
less, with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power for level flight at 0.9 V Hor 0.9 V
NE,whichever is less;
(4)
The landing gear retracted, and
(5)
The rotorcraft trimmed at 0.9 V Hor 0.9 V
NE,whichever is less.
(c)
Autorotation. Static longitudinal stability
must be shown in autorotation at airspeeds from 0.5
times the speed for minimum rate of descent, or 0.5
times the maximum range glide speed for Category A
rotorcraft, to VNEor to 1.1 VNE(power-off)
if VNE(power-off) is established under
29.1505(c), and with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power off;
(4)
The landing gear—
(i)
Retracted; and
(ii)
Extended; and
(5)
The rotorcraft trimmed at appropriate speeds found
necessary by the Administrator to demonstrate
stability throughout the prescribed speed range.
(d)
Hovering. For helicopters, the longitudinal
cyclic control must operate with the sense,
direction of motion, and position as prescribed in
29.173 between the maximum approved rearward speed
and a forward speed of 17 knots with—
(1)
Critical weight;
(2)
Critical center of gravity;
(3)
Power required to maintain an approximate constant
height in ground effect;
(4)
The landing gear extended; and
(5)
The helicopter trimmed for hovering.
Static directional stability must be positive with
throttle and collective controls held constant at
the trim conditions specified in 29.175 (a), (b),
and (c). Sideslip angle must increase steadily with
directional control deflection for sideslip angles
up to ±10° from trim. Sufficient cues must accompany
sideslip to alert the pilot when approaching
sideslip limits.
Any
short-period oscillation occurring at any speed from
VYto VNEmust be positively
damped with the primary flight controls free and in
a fixed position.
The
rotorcraft must have satisfactory ground and water
handling characteristics, including freedom from
uncontrollable tendencies in any condition expected
in operation.
The
rotorcraft must be designed to withstand the loads
that would occur when the rotorcraft is taxied over
the roughest ground that may reasonably be expected
in normal operation.
If
certification for water operation is requested, no
spray characteristics during taxiing, takeoff, or
landing may obscure the vision of the pilot or
damage the rotors, propellers, or other parts of the
rotorcraft.
The
rotorcraft may have no dangerous tendency to
oscillate on the ground with the rotor turning.
Each
part of the rotorcraft must be free from excessive
vibration under each appropriate speed and power
condition.
(a)
Strength requirements are specified in terms of
limit loads (the maximum loads to be expected in
service) and ultimate loads (limit loads multiplied
by prescribed factors of safety). Unless otherwise
provided, prescribed loads are limit loads.
(b)
Unless otherwise provided, the specified air,
ground, and water loads must be placed in
equilibrium with inertia forces, considering each
item of mass in the rotorcraft. These loads must be
distributed to closely approximate or conservatively
represent actual conditions.
(c)
If deflections under load would significantly change
the distribution of external or internal loads, this
redistribution must be taken into account.
Unless otherwise provided, a factor of safety of 1.5
must be used. This factor applies to external and
inertia loads unless its application to the
resulting internal stresses is more conservative.
(a)
The structure must be able to support limit loads
without detrimental or permanent deformation. At any
load up to limit loads, the deformation may not
interfere with safe operation.
(b)
The structure must be able to support ultimate loads
without failure. This must be shown by—
(1)
Applying ultimate loads to the structure in a static
test for at least three seconds; or
(2)
Dynamic tests simulating actual load application.
(a)
Compliance with the strength and deformation
requirements of this subpart must be shown for each
critical loading condition accounting for the
environment to which the structure will be exposed
in operation. Structural analysis (static or
fatigue) may be used only if the structure conforms
to those structures for which experience has shown
this method to be reliable. In other cases,
substantiating load tests must be made.
(b)
Proof of compliance with the strength requirements
of this subpart must include—
(1)
Dynamic and endurance tests of rotors, rotor drives,
and rotor controls;
(2)
Limit load tests of the control system, including
control surfaces;
(3)
Operation tests of the control system;
(4)
Flight stress measurement tests;
(5)
Landing gear drop tests; and
(6)
Any additional tests required for new or unusual
design features.
The
following values and limitations must be established
to show compliance with the structural requirements
of this subpart:
(a)
The design maximum and design minimum weights.
(b)
The main rotor r.p.m. ranges, power on and power
off.
(c)
The maximum forward speeds for each main rotor r.p.m.
within the ranges determined under paragraph (b) of
this section.
(d)
The maximum rearward and sideward flight speeds.
(e)
The center of gravity limits corresponding to the
limitations determined under paragraphs (b), (c),
and (d) of this section.
(f)
The rotational speed ratios between each powerplant
and each connected rotating component.
(g)
The positive and negative limit maneuvering load
factors.
(a)
The flight load factor must be assumed to act normal
to the longi | 
