title-14-part-25-sec25-255

231

Federal Aviation Administration, DOT

§ 25.255

yaw must be mild and readily control-
lable, using normal piloting tech-
niques. When the airplane is trimmed
at V

, the slope of the elevator

control force versus speed curve need
not be stable at speeds greater than
V

, but there must be a push force

at all speeds up to V

and there

must be no sudden or excessive reduc-
tion of elevator control force as V

is reached.

(4) Adequate roll capability to assure

a prompt recovery from a lateral upset
condition must be available at any
speed up to V

(5) With the airplane trimmed at

, extension of the speedbrakes

over the available range of movements
of the pilot’s control, at all speeds
above V

, but not so high that

would be exceeded during the

maneuver, must not result in:

(i) An excessive positive load factor

when the pilot does not take action to
counteract the effects of extension;

(ii) Buffeting that would impair the

pilot’s ability to read the instruments
or control the airplane for recovery; or

(iii) A nose down pitching moment,

unless it is small.

(b)

Maximum speed for stability charac-

teristics, V

. V

is the max-

imum speed at which the requirements
of §§ 25.143(g), 25.147(f), 25.175(b)(1),
25.177(a) through (c), and 25.181 must be
met with flaps and landing gear re-
tracted. Except as noted in § 25.253(c),
V

may not be less than a speed

midway between V

and V

except that, for altitudes where Mach
number is the limiting factor, M

need

not exceed the Mach number at which
effective speed warning occurs.

(c)

Maximum speed for stability charac-

teristics in icing conditions. The max-
imum speed for stability characteris-
tics with the most critical of the ice
accretions defined in Appendices C and
O of this part, as applicable, in accord-
ance with § 25.21(g), at which the re-
quirements of §§ 25.143(g), 25.147(f),
25.175(b)(1), 25.177(a) through (c), and
25.181 must be met, is the lower of:

(1) 300 knots CAS;
(2) V

; or

(3) A speed at which it is dem-

onstrated that the airframe will be free

of ice accretion due to the effects of in-
creased dynamic pressure.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25–23, 35 FR 5671, Apr. 8,
1970; Amdt. 25–54, 45 FR 60172, Sept. 11, 1980;
Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt.
25–84, 60 FR 30750, June 9, 1995; Amdt. 25–121,
72 FR 44668, Aug. 8, 2007; Amdt. 25–135, 76 FR
74654, Dec. 1, 2011; Amdt. 25–140,79 FR 65525,
Nov. 4, 2014]

§ 25.255

Out-of-trim characteristics.

(a) From an initial condition with

the airplane trimmed at cruise speeds
up to V

MO,

the airplane must have

satisfactory maneuvering stability and
controllability with the degree of out-
of-trim in both the airplane nose-up
and nose-down directions, which re-
sults from the greater of—

(1) A three-second movement of the

longitudinal trim system at its normal
rate for the particular flight condition
with no aerodynamic load (or an equiv-
alent degree of trim for airplanes that
do not have a power-operated trim sys-
tem), except as limited by stops in the
trim system, including those required
by § 25.655(b) for adjustable stabilizers;
or

(2) The maximum mistrim that can

be sustained by the autopilot while
maintaining level flight in the high
speed cruising condition.

(b) In the out-of-trim condition speci-

fied in paragraph (a) of this section,
when the normal acceleration is varied
from + 1 g to the positive and negative
values specified in paragraph (c) of this
section—

(1) The stick force vs. g curve must

have a positive slope at any speed up to
and including V

; and

(2) At speeds between V

and

the direction of the primary

longitudinal control force may not re-
verse.

(d) and (e) of this section, compliance
with the provisions of paragraph (a) of
this section must be demonstrated in
flight over the acceleration range—

(1)

1 g to + 2.5 g; or

(2) 0 g to 2.0 g, and extrapolating by

an acceptable method to

1 g and + 2.5

(d) If the procedure set forth in para-

graph (c)(2) of this section is used to
demonstrate compliance and marginal
conditions exist during flight test with

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232

14 CFR Ch. I (1–1–24 Edition)

§ 25.301

regard to reversal of primary longitu-
dinal control force, flight tests must be
accomplished from the normal accel-
eration at which a marginal condition
is found to exist to the applicable limit
specified in paragraph (b)(1) of this sec-
tion.

(e) During flight tests required by

paragraph (a) of this section, the limit
maneuvering load factors prescribed in
§§ 25.333(b) and 25.337, and the maneu-
vering load factors associated with
probable inadvertent excursions be-
yond the boundaries of the buffet onset
envelopes determined under § 25.251(e),
need not be exceeded. In addition, the
entry speeds for flight test demonstra-
tions at normal acceleration values
less than 1 g must be limited to the ex-
tent necessary to accomplish a recov-
ery without exceeding V

(f) In the out-of-trim condition speci-

fied in paragraph (a) of this section, it
must be possible from an overspeed
condition at V

to produce at

least 1.5 g for recovery by applying not
more than 125 pounds of longitudinal
control force using either the primary
longitudinal control alone or the pri-
mary longitudinal control and the lon-
gitudinal trim system. If the longitu-
dinal trim is used to assist in pro-
ducing the required load factor, it must
be shown at V

that the longitu-

dinal trim can be actuated in the air-
plane nose-up direction with the pri-
mary surface loaded to correspond to
the least of the following airplane
nose-up control forces:

(1) The maximum control forces ex-

pected in service as specified in §§ 25.301
and 25.397.

(2) The control force required to

produce 1.5 g.

(3) The control force corresponding to

buffeting or other phenomena of such
intensity that it is a strong deterrent
to further application of primary longi-
tudinal control force.

[Amdt. 25–42, 43 FR 2322, Jan. 16, 1978]

Subpart C—Structure

ENERAL

§ 25.301

Loads.

(a) Strength requirements are speci-

fied in terms of limit loads (the max-
imum loads to be expected in service)

and ultimate loads (limit loads multi-
plied 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 in-
ertia forces, considering each item of
mass in the airplane. These loads must
be distributed to conservatively ap-
proximate or closely represent actual
conditions. Methods used to determine
load intensities and distribution must
be validated by flight load measure-
ment unless the methods used for de-
termining those loading conditions are
shown to be reliable.

significantly change the distribution of
external or internal loads, this redis-
tribution must be taken into account.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25–23, 35 FR 5672, Apr. 8,
1970]

§ 25.303

Factor of safety.

Unless otherwise specified, a factor of

safety of 1.5 must be applied to the pre-
scribed limit load which are considered
external loads on the structure. When a
loading condition is prescribed in
terms of ultimate loads, a factor of
safety need not be applied unless other-
wise specified.

[Amdt. 25–23, 35 FR 5672, Apr. 8, 1970]

§ 25.305

Strength and deformation.

(a) The structure must be able to

support limit loads without detri-
mental 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
for at least 3 seconds. However, when
proof of strength is shown by dynamic
tests simulating actual load condi-
tions, the 3-second limit does not
apply. Static tests conducted to ulti-
mate load must include the ultimate
deflections and ultimate deformation
induced by the loading. When analyt-
ical methods are used to show compli-
ance with the ultimate load strength
requirements, it must be shown that—

(1) The effects of deformation are not

significant;