Federal Aviation Administration, DOT

Section 25.331

(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 Section 25.255 must be considered.
(b) Maneuvering balanced conditions.
Assuming the airplane to be in equilibrium with zero pitching acceleration, the maneuvering conditions A
through I on the maneuvering envelope
in Section 25.333(b) must be investigated.
(c) Maneuvering pitching conditions.
The following conditions must be investigated:
(1) Maximum pitch control displacement
at VA. The airplane is assumed to be
flying in steady level flight (point A1,
Section 25.333(b)) and the cockpit pitch control is suddenly moved to obtain extreme nose up pitching acceleration. In
defining the tail load, the response of
the airplane must be taken into account. Airplane loads that occur subsequent to the time when normal acceleration at the c.g. exceeds the positive
limit maneuvering load factor (at point
A2 in Section 25.333(b)), or the resulting
tailplane normal load reaches its maximum, whichever occurs first, need not
be considered.
(2) Checked maneuver between VA and
VD. Nose-up checked pitching maneuvers must be analyzed in which the

Where

(ii) For nose-up pitching maneuvers,
the complete flight deck pitch control
displacement history may be scaled
down in amplitude to the extent necessary to ensure that the positive limit
load factor prescribed in Section 25.337 is not
exceeded. For nose-down pitching maneuvers, the complete flight deck control displacement history may be
scaled down in amplitude to the extent

Where - 
d1 = the maximum available displacement of
the flight deck pitch control in the initial direction, as limited by the control
system stops, control surface stops, or by
pilot effort in accordance with Section 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 d1.
In the reverse direction, d(t) may be
truncated at the maximum available displacement of the flight deck pitch control as limited by the control system
stops, control surface stops, or by pilot
effort in accordance with 25.397(b);
tmax = 3/2w;
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:

necessary to ensure that the normal
acceleration at the center of gravity
does not go below 0g.
(iii) In addition, for cases where the
airplane response to the specified flight
deck pitch control motion does not
achieve the prescribed limit load factors, then the following flight deck
pitch control motion must be used:
d(t) = d1 sin(wt) for 0 - t - t1
d(t) = d1 for t1 - t - t2
d(t) = d1 sin(w[t + t1 Section t2]) for t2 - t -
tmax
Where - 
t1 = /2w

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V = the speed of the airplane at entry to the
maneuver.
VA = the design maneuvering speed prescribed in Section 25.335(c).

positive limit load factor prescribed in
Section 25.337 is achieved. As a separate condition, nose-down checked pitching maneuvers 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 flying in steady level flight at any speed
between VA and VD and the flight deck
pitch control is moved in accordance
with the following formula:
d(t) = d1 sin(wt) for 0 - t - tmax