title-14-part-25-sec25-629

266

14 CFR Ch. I (1–1–24 Edition)

§ 25.623

may be reduced when an approved qual-
ity control procedure is established.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25–139, 79 FR 59429, Oct. 2,
2014]

§ 25.623

Bearing factors.

(a) Except as provided in paragraph

(b) of this section, each part that has
clearance (free fit), and that is subject
to pounding or vibration, must have a
bearing factor large enough to provide
for the effects of normal relative mo-
tion.

(b) No bearing factor need be used for

a part for which any larger special fac-
tor is prescribed.

§ 25.625

Fitting factors.

For each fitting (a part or terminal

used to join one structural member to
another), the following apply:

(a) For each fitting whose strength is

not proven by limit and ultimate load
tests in which actual stress conditions
are simulated in the fitting and sur-
rounding structures, a fitting factor of
at least 1.15 must be applied to each
part of—

(1) The fitting;
(2) The means of attachment; and
(3) The bearing on the joined mem-

bers.

(b) No fitting factor need be used—
(1) For joints made under approved

practices and based on comprehensive
test data (such as continuous joints in
metal plating, welded joints, and scarf
joints in wood); or

(2) With respect to any bearing sur-

face for which a larger special factor is
used.

must be treated as a fitting up to the
point at which the section properties
become typical of the member.

(d) For each seat, berth, safety belt,

and harness, the fitting factor specified
in § 25.785(f)(3) applies.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25–23, 35 FR 5674, Apr. 8,
1970; Amdt. 25–72, 55 FR 29776, July 20, 1990]

§ 25.629

Aeroelastic stability require-

ments.

(a)

General. The aeroelastic stability

evaluations required under this section
include flutter, divergence, control re-
versal and any undue loss of stability

and control as a result of structural de-
formation. The aeroelastic evaluation
must include whirl modes associated
with any propeller or rotating device
that contributes significant dynamic
forces. Compliance with this section
must be shown by analyses, wind tun-
nel tests, ground vibration tests, flight
tests, or other means found necessary
by the Administrator.

(b)

Aeroelastic stability envelopes. The

airplane must be designed to be free
from aeroelastic instability for all con-
figurations and design conditions with-
in the aeroelastic stability envelopes
as follows:

(1) For normal conditions without

failures, malfunctions, or adverse con-
ditions, all combinations of altitudes
and speeds encompassed by the V

versus altitude envelope enlarged at all
points by an increase of 15 percent in
equivalent airspeed at both constant
Mach number and constant altitude. In
addition, a proper margin of stability
must exist at all speeds up to V

and, there must be no large and rapid
reduction in stability as V

is ap-

proached. The enlarged envelope may
be limited to Mach 1.0 when M

is less

than 1.0 at all design altitudes, and

(2) For the conditions described in

§ 25.629(d) below, for all approved alti-
tudes, any airspeed up to the greater
airspeed defined by;

(i) The V

envelope determined by

§ 25.335(b); or,

(ii) An altitude-airspeed envelope de-

fined by a 15 percent increase in equiv-
alent airspeed above V

at constant al-

titude, from sea level to the altitude of
the intersection of 1.15 V

with the ex-

tension of the constant cruise Mach
number line, M

, then a linear vari-

ation in equivalent airspeed to M

+ .05

at the altitude of the lowest V

intersection; then, at higher altitudes,
up to the maximum flight altitude, the
boundary defined by a .05 Mach in-
crease in M

at constant altitude.

(c)

Balance weights. If concentrated

balance weights are used, their effec-
tiveness and strength, including sup-
porting structure, must be substan-
tiated.

(d)

Failures, malfunctions, and adverse

conditions. The failures, malfunctions,
and adverse conditions which must be

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§ 25.651

considered in showing compliance with
this section are:

(1) Any critical fuel loading condi-

tions, not shown to be extremely im-
probable, which may result from mis-
management of fuel.

(2) Any single failure in any flutter

damper system.

(3) For airplanes not approved for op-

eration in icing conditions, the max-
imum likely ice accumulation expected
as a result of an inadvertent encounter.

(4) Failure of any single element of

the structure supporting any engine,
independently mounted propeller shaft,
large auxiliary power unit, or large ex-
ternally mounted aerodynamic body
(such as an external fuel tank).

(5) For airplanes with engines that

have propellers or large rotating de-
vices capable of significant dynamic
forces, any single failure of the engine
structure that would reduce the rigid-
ity of the rotational axis.

(6) The absence of aerodynamic or gy-

roscopic forces resulting from the most
adverse combination of feathered pro-
pellers or other rotating devices capa-
ble of significant dynamic forces. In
addition, the effect of a single feath-
ered propeller or rotating device must
be coupled with the failures of para-
graphs (d)(4) and (d)(5) of this section.

(7) Any single propeller or rotating

device capable of significant dynamic
forces rotating at the highest likely
overspeed.

(8) Any damage or failure condition,

required or selected for investigation
by § 25.571. The single structural fail-
ures described in paragraphs (d)(4) and
(d)(5) of this section need not be consid-
ered in showing compliance with this
section if;

(i) The structural element could not

fail due to discrete source damage re-
sulting from the conditions described
in § 25.571(e), and

(ii) A damage tolerance investigation

in accordance with § 25.571(b) shows
that the maximum extent of damage
assumed for the purpose of residual
strength evaluation does not involve
complete failure of the structural ele-
ment.

(9) Any damage, failure, or malfunc-

tion considered under §§ 25.631, 25.671,
25.672, and 25.1309.

(10) Any other combination of fail-

ures, malfunctions, or adverse condi-
tions not shown to be extremely im-
probable.

(e)

Flight flutter testing. Full scale

flight flutter tests at speeds up to V

must be conducted for new type

designs and for modifications to a type
design unless the modifications have
been shown to have an insignificant ef-
fect on the aeroelastic stability. These
tests must demonstrate that the air-
plane has a proper margin of damping
at all speeds up to V

, and that

there is no large and rapid reduction in
damping as V

, is approached. If a

failure, malfunction, or adverse condi-
tion is simulated during flight test in
showing compliance with paragraph (d)
of this section, the maximum speed in-
vestigated need not exceed V

if it

is shown, by correlation of the flight
test data with other test data or anal-
yses, that the airplane is free from any
aeroelastic instability at all speeds
within the altitude-airspeed envelope
described in paragraph (b)(2) of this
section.

[Doc. No. 26007, 57 FR 28949, June 29, 1992]

§ 25.631

Bird strike damage.

The empennage structure must be de-

signed to assure capability of contin-
ued safe flight and landing of the air-
plane after impact with an 8-pound bird
when the velocity of the airplane (rel-
ative to the bird along the airplane’s
flight path) is equal to

at sea level,

selected under § 25.335(a). Compliance
with this section by provision of redun-
dant structure and protected location
of control system elements or protec-
tive devices such as splitter plates or
energy absorbing material is accept-
able. Where compliance is shown by
analysis, tests, or both, use of data on
airplanes having similar structural de-
sign is acceptable.

[Amdt. 25–23, 35 FR 5674, Apr. 8, 1970]

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