background image

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 

VerDate Sep<11>2014 

09:06 Jun 28, 2024

Jkt 262046

PO 00000

Frm 00355

Fmt 8010

Sfmt 8010

Y:\SGML\262046.XXX

262046

jspears on DSK121TN23PROD with CFR

background image

346 

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 

VerDate Sep<11>2014 

09:06 Jun 28, 2024

Jkt 262046

PO 00000

Frm 00356

Fmt 8010

Sfmt 8010

Y:\SGML\262046.XXX

262046

jspears on DSK121TN23PROD with CFR