345
Federal Aviation Administration, DOT
§ 25.1325
(a) For airplanes certificated in ac-
cordance with § 25.1420(a)(1), the icing
conditions that the airplane is certified
to safely exit following detection.
(b) For airplanes certificated in ac-
cordance with § 25.1420(a)(2), the icing
conditions that the airplane is certified
to safely operate in and the icing con-
ditions that the airplane is certified to
safely exit following detection.
(c) For airplanes certificated in ac-
cordance with § 25.1420(a)(3) and for air-
planes not subject to § 25.1420, all icing
conditions.
[Amdt. 25–140, 79 FR 65527, Nov. 4, 2014]
§ 25.1325
Static pressure systems.
(a) Each instrument with static air
case connections must be vented to the
outside atmosphere through an appro-
priate piping system.
(b) Each static port must be designed
and located so that:
(1) The static pressure system per-
formance is least affected by airflow
variation, or by moisture or other for-
eign matter; and
(2) The correlation between air pres-
sure in the static pressure system and
true ambient atmospheric static pres-
sure is not changed when the airplane
is exposed to the icing conditions de-
fined in Appendix C of this part, and
the following icing conditions specified
in Appendix O of this part:
(i) For airplanes certificated in ac-
cordance with § 25.1420(a)(1), the icing
conditions that the airplane is certified
to safely exit following detection.
(ii) For airplanes certificated in ac-
cordance with § 25.1420(a)(2), the icing
conditions that the airplane is certified
to safely operate in and the icing con-
ditions that the airplane is certified to
safely exit following detection.
(iii) For airplanes certificated in ac-
cordance with § 25.1420(a)(3) and for air-
planes not subject to § 25.1420, all icing
conditions.
(c) The design and installation of the
static pressure system must be such
that—
(1) Positive drainage of moisture is
provided; chafing of the tubing and ex-
cessive distortion or restriction at
bends in the tubing is avoided; and the
materials used are durable, suitable for
the purpose intended, and protected
against corrosion; and
(2) It is airtight except for the port
into the atmosphere. A proof test must
be conducted to demonstrate the integ-
rity of the static pressure system in
the following manner:
(i)
Unpressurized airplanes. Evacuate
the static pressure system to a pres-
sure differential of approximately 1
inch of mercury or to a reading on the
altimeter, 1,000 feet above the airplane
elevation at the time of the test. With-
out additional pumping for a period of
1 minute, the loss of indicated altitude
must not exceed 100 feet on the altim-
eter.
(ii)
Pressurized airplanes. Evacuate
the static pressure system until a pres-
sure differential equivalent to the max-
imum cabin pressure differential for
which the airplane is type certificated
is achieved. Without additional pump-
ing for a period of 1 minute, the loss of
indicated altitude must not exceed 2
percent of the equivalent altitude of
the maximum cabin differential pres-
sure or 100 feet, whichever is greater.
(d) Each pressure altimeter must be
approved and must be calibrated to in-
dicate pressure altitude in a standard
atmosphere, with a minimum prac-
ticable calibration error when the cor-
responding static pressures are applied.
(e) Each system must be designed and
installed so that the error in indicated
pressure altitude, at sea level, with a
standard atmosphere, excluding instru-
ment calibration error, does not result
in an error of more than
±
30 feet per 100
knots speed for the appropriate con-
figuration in the speed range between
1.23
V
SR0
with flaps extended and 1.7
V
SR1
with flaps retracted. However, the
error need not be less than
±
30 feet.
(f) If an altimeter system is fitted
with a device that provides corrections
to the altimeter indication, the device
must be designed and installed in such
manner that it can be bypassed when it
malfunctions, unless an alternate al-
timeter system is provided. Each cor-
rection device must be fitted with a
means for indicating the occurrence of
reasonably probable malfunctions, in-
cluding power failure, to the flight
crew. The indicating means must be ef-
fective for any cockpit lighting condi-
tion likely to occur.
(g) Except as provided in paragraph
(h) of this section, if the static pressure
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14 CFR Ch. I (1–1–24 Edition)
§ 25.1326
system incorporates both a primary
and an alternate static pressure source,
the means for selecting one or the
other source must be designed so
that—
(1) When either source is selected, the
other is blocked off; and
(2) Both sources cannot be blocked
off simultaneously.
(h) For unpressurized airplanes, para-
graph (g)(1) of this section does not
apply if it can be demonstrated that
the static pressure system calibration,
when either static pressure source is
selected, is not changed by the other
static pressure source being open or
blocked.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25–5, 30 FR 8261, June 29,
1965; Amdt. 25–12, 32 FR 7587, May 24, 1967;
Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt.
25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25–
140, 79 FR 65527, Nov. 4, 2014]
§ 25.1326
Pitot heat indication systems.
If a flight instrument pitot heating
system is installed, an indication sys-
tem must be provided to indicate to
the flight crew when that pitot heating
system is not operating. The indication
system must comply with the following
requirements:
(a) The indication provided must in-
corporate an amber light that is in
clear view of a flight crewmember.
(b) The indication provided must be
designed to alert the flight crew if ei-
ther of the following conditions exist:
(1) The pitot heating system is
switched ‘‘off’’.
(2) The pitot heating system is
switched ‘‘on’’ and any pitot tube heat-
ing element is inoperative.
[Amdt. 25–43, 43 FR 10339, Mar. 13, 1978]
§ 25.1327
Magnetic direction indicator.
(a) Each magnetic direction indicator
must be installed so that its accuracy
is not excessively affected by the air-
plane’s vibration or magnetic fields.
(b) The compensated installation
may not have a deviation, in level
flight, greater than 10 degrees on any
heading.
§ 25.1329
Flight guidance system.
(a) Quick disengagement controls for
the autopilot and autothrust functions
must be provided for each pilot. The
autopilot quick disengagement con-
trols must be located on both control
wheels (or equivalent). The autothrust
quick disengagement controls must be
located on the thrust control levers.
Quick disengagement controls must be
readily accessible to each pilot while
operating the control wheel (or equiva-
lent) and thrust control levers.
(b) The effects of a failure of the sys-
tem to disengage the autopilot or
autothrust functions when manually
commanded by the pilot must be as-
sessed in accordance with the require-
ments of § 25.1309.
(c) Engagement or switching of the
flight guidance system, a mode, or a
sensor may not cause a transient re-
sponse of the airplane’s flight path any
greater than a minor transient, as de-
fined in paragraph (n)(1) of this section.
(d) Under normal conditions, the dis-
engagement of any automatic control
function of a flight guidance system
may not cause a transient response of
the airplane’s flight path any greater
than a minor transient.
(e) Under rare normal and non-nor-
mal conditions, disengagement of any
automatic control function of a flight
guidance system may not result in a
transient any greater than a signifi-
cant transient, as defined in paragraph
(n)(2) of this section.
(f) The function and direction of mo-
tion of each command reference con-
trol, such as heading select or vertical
speed, must be plainly indicated on, or
adjacent to, each control if necessary
to prevent inappropriate use or confu-
sion.
(g) Under any condition of flight ap-
propriate to its use, the flight guidance
system may not produce hazardous
loads on the airplane, nor create haz-
ardous deviations in the flight path.
This applies to both fault-free oper-
ation and in the event of a malfunc-
tion, and assumes that the pilot begins
corrective action within a reasonable
period of time.
(h) When the flight guidance system
is in use, a means must be provided to
avoid excursions beyond an acceptable
margin from the speed range of the
normal flight envelope. If the airplane
experiences an excursion outside this
range, a means must be provided to
prevent the flight guidance system
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