234
14 CFR Ch. I (1–1–24 Edition)
§ 25.331
F
LIGHT
M
ANEUVER AND
G
UST
C
ONDITIONS
§ 25.331
Symmetric maneuvering con-
ditions.
(a)
Procedure. For the analysis of the
maneuvering flight conditions specified
in paragraphs (b) and (c) of this sec-
tion, the following provisions apply:
(1) Where sudden displacement of a
control is specified, the assumed rate
of control surface displacement may
not be less than the rate that could be
applied by the pilot through the con-
trol system.
(2) In determining elevator angles
and chordwise load distribution in the
maneuvering conditions of paragraphs
(b) and (c) of this section, the effect of
corresponding pitching velocities must
be taken into account. The in-trim and
out-of-trim flight conditions specified
in § 25.255 must be considered.
(b)
Maneuvering balanced conditions.
Assuming the airplane to be in equi-
librium with zero pitching accelera-
tion, the maneuvering conditions A
through I on the maneuvering envelope
in § 25.333(b) must be investigated.
(c)
Maneuvering pitching conditions.
The following conditions must be in-
vestigated:
(1)
Maximum pitch control displacement
at V
A
. The airplane is assumed to be
flying in steady level flight (point A
1
,
§ 25.333(b)) and the cockpit pitch con-
trol is suddenly moved to obtain ex-
treme nose up pitching acceleration. In
defining the tail load, the response of
the airplane must be taken into ac-
count. Airplane loads that occur subse-
quent to the time when normal accel-
eration at the c.g. exceeds the positive
limit maneuvering load factor (at point
A
2
in § 25.333(b)), or the resulting
tailplane normal load reaches its max-
imum, whichever occurs first, need not
be considered.
(2)
Checked maneuver between V
A
and
V
D
. Nose-up checked pitching maneu-
vers must be analyzed in which the
positive limit load factor prescribed in
§ 25.337 is achieved. As a separate condi-
tion, nose-down checked pitching ma-
neuvers must be analyzed in which a
limit load factor of 0g is achieved. In
defining the airplane loads, the flight
deck pitch control motions described
in paragraphs (c)(2)(i) through (iv) of
this section must be used:
(i) The airplane is assumed to be fly-
ing in steady level flight at any speed
between V
A
and V
D
and the flight deck
pitch control is moved in accordance
with the following formula:
d
(t) =
d
1
sin(
w
t) for 0
≤
t
≤
t
max
Where—
d
1
= the maximum available displacement of
the flight deck pitch control in the ini-
tial direction, as limited by the control
system stops, control surface stops, or by
pilot effort in accordance with § 25.397(b);
d
(t) = the displacement of the flight deck
pitch control as a function of time. In
the initial direction,
d
(t) is limited to
d
1
.
In the reverse direction,
d
(t) may be
truncated at the maximum available dis-
placement of the flight deck pitch con-
trol as limited by the control system
stops, control surface stops, or by pilot
effort in accordance with 25.397(b);
t
max
= 3
π
/2
w
;
w
= the circular frequency (radians/second) of
the control deflection taken equal to the
undamped natural frequency of the short
period rigid mode of the airplane, with
active control system effects included
where appropriate; but not less than:
Where
V = the speed of the airplane at entry to the
maneuver.
V
A
= the design maneuvering speed pre-
scribed in § 25.335(c).
(ii) For nose-up pitching maneuvers,
the complete flight deck pitch control
displacement history may be scaled
down in amplitude to the extent nec-
essary to ensure that the positive limit
load factor prescribed in § 25.337 is not
exceeded. For nose-down pitching ma-
neuvers, the complete flight deck con-
trol displacement history may be
scaled down in amplitude to the extent
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