517
Federal Aviation Administration, DOT
§ 27.725
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
U.S.C. section 552(a) and 1 CFR part 51.
Copies may be obtained from the Naval
Publications and Forms Center, 5801
Tabor Avenue, Philadelphia, Pennsyl-
vania, 19120. Copies may be inspected
at the National Archives and Records
Administration (NARA). For informa-
tion on the availability of this mate-
rial at NARA, call 202–741–6030, or go
to:
http://www.archives.gov/federal-reg-
ister/cfr/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 sys-
tem.
(9) Turnbuckles attached to parts
having angular motion must be in-
stalled to prevent binding throughout
the range of travel.
(10) There must be means for visual
inspection at each fairlead, pulley, ter-
minal, 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 rat-
ing of ball and roller bearings must not
be exceeded.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–11, 41 FR 55469, Dec. 20,
1976; Amdt. 27–26, 55 FR 8001, Mar. 6, 1990; 69
FR 18803, Apr. 9, 2004; Doc. No. FAA–2018–
0119, Amdt. 27–49, 83 FR 9170, Mar. 5, 2018]
§ 27.687
Spring devices.
(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.
§ 27.691
Autorotation control mecha-
nism.
Each main rotor blade pitch control
mechanism must allow rapid entry into
autorotation after power failure.
§ 27.695
Power boost and power-oper-
ated control system.
(a) If a power boost or power-oper-
ated control system is used, an alter-
nate 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 actu-
ators.
(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.
L
ANDING
G
EAR
§ 27.723
Shock absorption tests.
The landing inertia load factor and
the reserve energy absorption capacity
of the landing gear must be substan-
tiated by the tests prescribed in
§§ 27.725 and 27.727, respectively. These
tests must be conducted on the com-
plete rotorcraft or on units consisting
of wheel, tire, and shock absorber in
their proper relation.
§ 27.725
Limit drop test.
The limit drop test must be con-
ducted 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 con-
tact velocity equal to the greatest
probable sinking speed likely to occur
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14 CFR Ch. I (1–1–24 Edition)
§ 27.727
at ground contact in normal power-off
landings.
(b) If considered, the rotor lift speci-
fied in § 27.473(a) must be introduced
into the drop test by appropriate en-
ergy 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:
W
W
h
d
h
d
n
n
W
W
L
e
j
e
=
×
+ −
(
)
+
=
+
1 L
and
;
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 reac-
tion, taking into consideration the mo-
ment arm between the main wheel reac-
tion and the rotorcraft center of gravity.
W = W
N
for nose gear units (lbs.), equal to
the vertical component of the static re-
action 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 tailwheel units (lbs.), equal to
whichever of the following is critical:
(1) The static weight on the tailwheel with
the rotorcraft resting on all wheels; or
(2) The vertical component of the ground
reaction that would occur at the tailwheel,
assuming that the mass of the rotorcraft
acts at the center of gravity and exerts a
force of 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 rotor-
craft 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).
§ 27.727
Reserve energy absorption
drop test.
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.
[Doc. No. 5074, 29 FR 15695, Nov. 24, 1964, as
amended by Amdt. 27–26, 55 FR 8001, Mar. 6,
1990]
§ 27.729
Retracting mechanism.
For rotorcraft with retractable land-
ing gear, the following apply:
(a)
Loads. The landing gear, retract-
ing mechansim, wheel-well doors, and
supporting structure must be designed
for—
(1) The loads occurring in any ma-
neuvering condition with the gear re-
tracted;
(2) The combined friction, inertia,
and air loads occurring during retrac-
tion 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 en-
ergy.
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