subpart f—equipment - gpo · ertia, and other loads to which it may be subjected in operation;...

492

14 CFR Ch. I (1–1–11 Edition) § 25.1203

(b) Each system component in an en-gine compartment must be fireproof.

§ 25.1203 Fire detector system. (a) There must be approved, quick

acting fire or overheat detectors in each designated fire zone, and in the combustion, turbine, and tailpipe sec-tions of turbine engine installations, in numbers and locations ensuring prompt detection of fire in those zones.

(b) Each fire detector system must be constructed and installed so that—

(1) It will withstand the vibration, in-ertia, and other loads to which it may be subjected in operation;

(2) There is a means to warn the crew in the event that the sensor or associ-ated wiring within a designated fire zone is severed at one point, unless the system continues to function as a sat-isfactory detection system after the severing; and

(3) There is a means to warn the crew in the event of a short circuit in the sensor or associated wiring within a designated fire zone, unless the system continues to function as a satisfactory detection system after the short cir-cuit.

(c) No fire or overheat detector may be affected by any oil, water, other fluids or fumes that might be present.

(d) There must be means to allow the crew to check, in flight, the func-tioning of each fire or overheat detec-tor electric circuit.

(e) Components of each fire or over-heat detector system in a fire zone must be fire-resistant.

(f) No fire or overheat detector sys-tem component for any fire zone may pass through another fire zone, un-less—

(1) It is protected against the possi-bility of false warnings resulting from fires in zones through which it passes; or

(2) Each zone involved is simulta-neously protected by the same detector and extinguishing system.

(g) Each fire detector system must be constructed so that when it is in the configuration for installation it will not exceed the alarm activation time approved for the detectors using the re-sponse time criteria specified in the ap-propriate Technical Standard Order for the detector.

(h) EWIS for each fire or overheat de-tector system in a fire zone must meet the requirements of § 25.1731.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5678, Apr. 8, 1970; Amdt. 25–26, 36 FR 5493, Mar. 24, 1971; Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1207 Compliance. Unless otherwise specified, compli-

ance with the requirements of §§ 25.1181 through 25.1203 must be shown by a full scale fire test or by one or more of the following methods:

(a) Tests of similar powerplant con-figurations;

(b) Tests of components; (c) Service experience of aircraft

with similar powerplant configura-tions;

(d) Analysis.

[Amdt. 25–46, 43 FR 50598, Oct. 30, 1978]

Subpart F—Equipment

GENERAL

§ 25.1301 Function and installation. (a) Each item of installed equipment

must— (1) Be of a kind and design appro-

priate to its intended function; (2) Be labeled as to its identification,

function, or operating limitations, or any applicable combination of these factors;

(3) Be installed according to limita-tions specified for that equipment; and

(4) Function properly when installed. (b) EWIS must meet the require-

ments of subpart H of this part.

[Dockt. No. 5066, Amdt. 1–6, 29 FR 18333, Dec. 24, 1964, as amended by Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1303 Flight and navigation instru-ments.

(a) The following flight and naviga-tion instruments must be installed so that the instrument is visible from each pilot station:

(1) A free air temperature indicator or an air-temperature indicator which provides indications that are convert-ible to free-air temperature.

(2) A clock displaying hours, min-utes, and seconds with a sweep-second pointer or digital presentation.

VerDate Mar<15>2010 14:10 Mar 01, 2011 Jkt 223043 PO 00000 Frm 00502 Fmt 8010 Sfmt 8010 Y:\SGML\223043.XXX 223043ww

oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

493

Federal Aviation Administration, DOT § 25.1305

(3) A direction indicator (non-stabilized magnetic compass).

(b) The following flight and naviga-tion instruments must be installed at each pilot station:

(1) An airspeed indicator. If airspeed limitations vary with altitude, the in-dicator must have a maximum allow-able airspeed indicator showing the variation of VMO with altitude.

(2) An altimeter (sensitive). (3) A rate-of-climb indicator (vertical

speed). (4) A gyroscopic rate-of-turn indi-

cator combined with an integral slip- skid indicator (turn-and-bank indi-cator) except that only a slip-skid indi-cator is required on large airplanes with a third attitude instrument sys-tem useable through flight attitudes of 360° of pitch and roll and installed in accordance with § 121.305(k) of this title.

(5) A bank and pitch indicator (gyro-scopically stabilized).

(6) A direction indicator (gyroscop-ically stabilized, magnetic or non-magnetic).

(c) The following flight and naviga-tion instruments are required as pre-scribed in this paragraph:

(1) A speed warning device is required for turbine engine powered airplanes and for airplanes with VMO/MMO great-er than 0.8 VDF/MDF or 0.8 V D/MD. The speed warning device must give effec-tive aural warning (differing distinc-tively from aural warnings used for other purposes) to the pilots, whenever the speed exceeds VMO plus 6 knots or MMO +0.01. The upper limit of the pro-duction tolerance for the warning de-vice may not exceed the prescribed warning speed.

(2) A machmeter is required at each pilot station for airplanes with com-pressibility limitations not otherwise indicated to the pilot by the airspeed indicating system required under para-graph (b)(1) of this section.

[Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25–24, 35 FR 7108, May 6, 1970; Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–90, 62 FR 13253, Mar. 19, 1997]

§ 25.1305 Powerplant instruments.

The following are required power-plant instruments:

(a) For all airplanes. (1) A fuel pres-sure warning means for each engine, or a master warning means for all engines with provision for isolating the indi-vidual warning means from the master warning means.

(2) A fuel quantity indicator for each fuel tank.

(3) An oil quantity indicator for each oil tank.

(4) An oil pressure indicator for each independent pressure oil system of each engine.

(5) An oil pressure warning means for each engine, or a master warning means for all engines with provision for isolating the individual warning means from the master warning means.

(6) An oil temperature indicator for each engine.

(7) Fire-warning devices that provide visual and audible warning.

(8) An augmentation liquid quantity indicator (appropriate for the manner in which the liquid is to be used in op-eration) for each tank.

(b) For reciprocating engine-powered airplanes. In addition to the powerplant instruments required by paragraph (a) of this section, the following power-plant instruments are required:

(1) A carburetor air temperature indi-cator for each engine.

(2) A cylinder head temperature indi-cator for each air-cooled engine.

(3) A manifold pressure indicator for each engine.

(4) A fuel pressure indicator (to indi-cate the pressure at which the fuel is supplied) for each engine.

(5) A fuel flowmeter, or fuel mixture indicator, for each engine without an automatic altitude mixture control.

(6) A tachometer for each engine. (7) A device that indicates, to the

flight crew (during flight), any change in the power output, for each engine with—

(i) An automatic propeller feathering system, whose operation is initiated by a power output measuring system; or

(ii) A total engine piston displace-ment of 2,000 cubic inches or more.

(8) A means to indicate to the pilot when the propeller is in reverse pitch, for each reversing propeller.

(c) For turbine engine-powered air-planes. In addition to the powerplant instruments required by paragraph (a)


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

494

14 CFR Ch. I (1–1–11 Edition) § 25.1307

of this section, the following power-plant instruments are required:

(1) A gas temperature indicator for each engine.

(2) A fuel flowmeter indicator for each engine.

(3) A tachometer (to indicate the speed of the rotors with established limiting speeds) for each engine.

(4) A means to indicate, to the flight crew, the operation of each engine starter that can be operated continu-ously but that is neither designed for continuous operation nor designed to prevent hazard if it failed.

(5) An indicator to indicate the func-tioning of the powerplant ice protec-tion system for each engine.

(6) An indicator for the fuel strainer or filter required by § 25.997 to indicate the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with § 25.997(d).

(7) A warning means for the oil strainer or filter required by § 25.1019, if it has no bypass, to warn the pilot of the occurrence of contamination of the strainer or filter screen before it reaches the capacity established in ac-cordance with § 25.1019(a)(2).

(8) An indicator to indicate the prop-er functioning of any heater used to prevent ice clogging of fuel system components.

(d) For turbojet engine powered air-planes. In addition to the powerplant instruments required by paragraphs (a) and (c) of this section, the following powerplant instruments are required:

(1) An indicator to indicate thrust, or a parameter that is directly related to thrust, to the pilot. The indication must be based on the direct measure-ment of thrust or of parameters that are directly related to thrust. The indi-cator must indicate a change in thrust resulting from any engine malfunction, damage, or deterioration.

(2) A position indicating means to in-dicate to the flightcrew when the thrust reversing device—

(i) Is not in the selected position, and (ii) Is in the reverse thrust position,

for each engine using a thrust revers-ing device.

(3) An indicator to indicate rotor sys-tem unbalance.

(e) For turbopropeller-powered air-planes. In addition to the powerplant instruments required by paragraphs (a) and (c) of this section, the following powerplant instruments are required:

(1) A torque indicator for each en-gine.

(2) Position indicating means to indi-cate to the flight crew when the pro-peller blade angle is below the flight low pitch position, for each propeller.

(f) For airplanes equipped with fluid systems (other than fuel) for thrust or power augmentation, an approved means must be provided to indicate the proper functioning of that system to the flight crew.

[Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25–35, 39 FR 1831, Jan. 15, 1974; Amdt. 25–36, 39 FR 35461, Oct. 1, 1974; Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–115, 69 FR 40527, July 2, 2004]

§ 25.1307 Miscellaneous equipment.

The following is required miscella-neous equipment:

(a) [Reserved] (b) Two or more independent sources

of electrical energy. (c) Electrical protective devices, as

prescribed in this part. (d) Two systems for two-way radio

communications, with controls for each accessible from each pilot station, designed and installed so that failure of one system will not preclude operation of the other system. The use of a com-mon antenna system is acceptable if adequate reliability is shown.

(e) Two systems for radio navigation, with controls for each accessible from each pilot station, designed and in-stalled so that failure of one system will not preclude operation of the other system. The use of a common antenna system is acceptable if adequate reli-ability is shown.

[Amdt. 25–23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, 55 FR 29785, July 20, 1990]


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

495


§ 25.1309 Equipment, systems, and in-stallations.

(a) The equipment, systems, and in-stallations whose functioning is re-quired by this subchapter, must be de-signed to ensure that they perform their intended functions under any foreseeable operating condition.

(b) The airplane systems and associ-ated components, considered sepa-rately and in relation to other systems, must be designed so that—

(1) The occurrence of any failure con-dition which would prevent the contin-ued safe flight and landing of the air-plane is extremely improbable, and

(2) The occurrence of any other fail-ure conditions which would reduce the capability of the airplane or the ability of the crew to cope with adverse oper-ating conditions is improbable.

(c) Warning information must be pro-vided to alert the crew to unsafe sys-tem operating conditions, and to en-able them to take appropriate correc-tive action. Systems, controls, and as-sociated monitoring and warning means must be designed to minimize crew errors which could create addi-tional hazards.

(d) Compliance with the require-ments of paragraph (b) of this section must be shown by analysis, and where necessary, by appropriate ground, flight, or simulator tests. The analysis must consider—

(1) Possible modes of failure, includ-ing malfunctions and damage from ex-ternal sources.

(2) The probability of multiple fail-ures and undetected failures.

(3) The resulting effects on the air-plane and occupants, considering the stage of flight and operating condi-tions, and

(4) The crew warning cues, corrective action required, and the capability of detecting faults.

(e) In showing compliance with para-graphs (a) and (b) of this section with regard to the electrical system and equipment design and installation, critical environmental conditions must be considered. For electrical genera-tion, distribution, and utilization equipment required by or used in com-plying with this chapter, except equip-ment covered by Technical Standard Orders containing environmental test

procedures, the ability to provide con-tinuous, safe service under foreseeable environmental conditions may be shown by environmental tests, design analysis, or reference to previous com-parable service experience on other air-craft.

(f) EWIS must be assessed in accord-ance with the requirements of § 25.1709.

[Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1310 Power source capacity and distribution.

(a) Each installation whose func-tioning is required for type certifi-cation or under operating rules and that requires a power supply is an ‘‘es-sential load’’ on the power supply. The power sources and the system must be able to supply the following power loads in probable operating combina-tions and for probable durations:

(1) Loads connected to the system with the system functioning normally.

(2) Essential loads, after failure of any one prime mover, power converter, or energy storage device.

(3) Essential loads after failure of— (i) Any one engine on two-engine air-

planes; and (ii) Any two engines on airplanes

with three or more engines. (4) Essential loads for which an alter-

nate source of power is required, after any failure or malfunction in any one power supply system, distribution sys-tem, or other utilization system.

(b) In determining compliance with paragraphs (a)(2) and (3) of this section, the power loads may be assumed to be reduced under a monitoring procedure consistent with safety in the kinds of operation authorized. Loads not re-quired in controlled flight need not be considered for the two-engine-inoper-ative condition on airplanes with three or more engines.

[Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1316 System lightning protection. (a) For functions whose failure would

contribute to or cause a condition that would prevent the continued safe flight and landing of the airplane, each elec-trical and electronic system that per-forms these functions must be designed


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

496

14 CFR Ch. I (1–1–11 Edition) § 25.1317

and installed to ensure that the oper-ation and operational capabilities of the systems to perform these functions are not adversely affected when the airplane is exposed to lightning.

(b) For functions whose failure would contribute to or cause a condition that would reduce the capability of the air-plane or the ability of the flightcrew to cope with adverse operating conditions, each electrical and electronic system that performs these functions must be designed and installed to ensure that these functions can be recovered in a timely manner after the airplane is ex-posed to lightning.

(c) Compliance with the lightning protection criteria prescribed in para-graphs (a) and (b) of this section must be shown for exposure to a severe light-ning environment. The applicant must design for and verify that aircraft elec-trical/electronic systems are protected against the effects of lightning by:

(1) Determining the lightning strike zones for the airplane;

(2) Establishing the external light-ning environment for the zones;

(3) Establishing the internal environ-ment;

(4) Identifying all the electrical and electronic systems that are subject to the requirements of this section, and their locations on or within the air-plane;

(5) Establishing the susceptibility of the systems to the internal and exter-nal lightning environment;

(6) Designing protection; and (7) Verifying that the protection is

adequate.

[Doc. No. 25912, 59 FR 22116, Apr. 28, 1994]

§ 25.1317 High-intensity Radiated Fields (HIRF) Protection.

(a) Except as provided in paragraph (d) of this section, each electrical and electronic system that performs a func-tion whose failure would prevent the continued safe flight and landing of the airplane must be designed and installed so that—

(1) The function is not adversely af-fected during and after the time the airplane is exposed to HIRF environ-ment I, as described in appendix L to this part;

(2) The system automatically recov-ers normal operation of that function,

in a timely manner, after the airplane is exposed to HIRF environment I, as described in appendix L to this part, unless the system’s recovery conflicts with other operational or functional requirements of the system; and

(3) The system is not adversely af-fected during and after the time the airplane is exposed to HIRF environ-ment II, as described in appendix L to this part.

(b) Each electrical and electronic system that performs a function whose failure would significantly reduce the capability of the airplane or the ability of the flightcrew to respond to an ad-verse operating condition must be de-signed and installed so the system is not adversely affected when the equip-ment providing these functions is ex-posed to equipment HIRF test level 1 or 2, as described in appendix L to this part.

(c) Each electrical and electronic sys-tem that performs a function whose failure would reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse op-erating condition must be designed and installed so the system is not adversely affected when the equipment providing the function is exposed to equipment HIRF test level 3, as described in ap-pendix L to this part.

(d) Before December 1, 2012, an elec-trical or electronic system that per-forms a function whose failure would prevent the continued safe flight and landing of an airplane may be designed and installed without meeting the pro-visions of paragraph (a) provided—

(1) The system has previously been shown to comply with special condi-tions for HIRF, prescribed under § 21.16, issued before December 1, 2007;

(2) The HIRF immunity characteris-tics of the system have not changed since compliance with the special con-ditions was demonstrated; and

(3) The data used to demonstrate compliance with the special conditions is provided.

[Doc. No. FAA–2006–23657, 72 FR 44025, Aug. 6, 2007]


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

497


INSTRUMENTS: INSTALLATION

§ 25.1321 Arrangement and visibility. (a) Each flight, navigation, and pow-

erplant instrument for use by any pilot must be plainly visible to him from his station with the minimum practicable deviation from his normal position and line of vision when he is looking for-ward along the flight path.

(b) The flight instruments required by § 25.1303 must be grouped on the in-strument panel and centered as nearly as practicable about the vertical plane of the pilot’s forward vision. In addi-tion—

(1) The instrument that most effec-tively indicates attitude must be on the panel in the top center position;

(2) The instrument that most effec-tively indicates airspeed must be adja-cent to and directly to the left of the instrument in the top center position:

(3) The instrument that most effec-tively indicates altitude must be adja-cent to and directly to the right of the instrument in the top center position; and

(4) The instrument that most effec-tively indicates direction of flight must be adjacent to and directly below the instrument in the top center posi-tion.

(c) Required powerplant instruments must be closely grouped on the instru-ment panel. In addition—

(1) The location of identical power-plant instruments for the engines must prevent confusion as to which engine each instrument relates; and

(2) Powerplant instruments vital to the safe operation of the airplane must be plainly visible to the appropriate crewmembers.

(d) Instrument panel vibration may not damage or impair the accuracy of any instrument.

(e) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.

[Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977]

§ 25.1322 Warning, caution, and advi-sory lights.

If warning, caution or advisory lights are installed in the cockpit, they must,

unless otherwise approved by the Ad-ministrator, be—

(a) Red, for warning lights (lights in-dicating a hazard which may require immediate corrective action);

(b) Amber, for caution lights (lights indicating the possible need for future corrective action);

(c) Green, for safe operation lights; and

(d) Any other color, including white, for lights not described in paragraphs (a) through (c) of this section, provided the color differs sufficiently from the colors prescribed in paragraphs (a) through (c) of this section to avoid pos-sible confusion.

[Amdt. 25–38, 41 FR 55467, Dec. 20, 1976]

EFFECTIVE DATE NOTE: At 75 FR 67209, Nov. 2, 2010, § 25.1322 was revised, effective Jan. 3, 2011. For the convenience of the user, the re-vised text is set forth as follows:

§ 25.1322 Flightcrew alerting. (a) Flightcrew alerts must: (1) Provide the flightcrew with the infor-

mation needed to: (i) Identify non-normal operation or air-

plane system conditions, and (ii) Determine the appropriate actions, if

any. (2) Be readily and easily detectable and in-

telligible by the flightcrew under all foresee-able operating conditions, including condi-tions where multiple alerts are provided.

(3) Be removed when the alerting condition no longer exists.

(b) Alerts must conform to the following prioritization hierarchy based on the ur-gency of flightcrew awareness and response.

(1) Warning: For conditions that require immediate flightcrew awareness and imme-diate flightcrew response.

(2) Caution: For conditions that require immediate flightcrew awareness and subse-quent flightcrew response.

(3) Advisory: For conditions that require flightcrew awareness and may require subse-quent flightcrew response.

(c) Warning and caution alerts must: (1) Be prioritized within each category,

when necessary. (2) Provide timely attention-getting cues

through at least two different senses by a combination of aural, visual, or tactile indi-cations.

(3) Permit each occurrence of the atten-tion-getting cues required by paragraph (c)(2) of this section to be acknowledged and suppressed, unless they are required to be continuous.

(d) The alert function must be designed to minimize the effects of false and nuisance alerts. In particular, it must be designed to:


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

498

14 CFR Ch. I (1–1–11 Edition) § 25.1323

(1) Prevent the presentation of an alert that is inappropriate or unnecessary.

(2) Provide a means to suppress an atten-tion-getting component of an alert caused by a failure of the alerting function that inter-feres with the flightcrew’s ability to safely operate the airplane. This means must not be readily available to the flightcrew so that it could be operated inadvertently or by ha-bitual reflexive action. When an alert is sup-pressed, there must be a clear and unmistak-able annunciation to the flightcrew that the alert has been suppressed.

(e) Visual alert indications must: (1) Conform to the following color conven-

tion: (i) Red for warning alert indications. (ii) Amber or yellow for caution alert indi-

cations. (iii) Any color except red or green for advi-

sory alert indications. (2) Use visual coding techniques, together

with other alerting function elements on the flight deck, to distinguish between warning, caution, and advisory alert indications, if they are presented on monochromatic dis-plays that are not capable of conforming to the color convention in paragraph (e)(1) of this section.

(f) Use of the colors red, amber, and yellow on the flight deck for functions other than flightcrew alerting must be limited and must not adversely affect flightcrew alerting.

§ 25.1323 Airspeed indicating system. For each airspeed indicating system,

the following apply: (a) Each airspeed indicating instru-

ment must be approved and must be calibrated to indicate true airspeed (at sea level with a standard atmosphere) with a minimum practicable instru-ment calibration error when the cor-responding pitot and static pressures are applied.

(b) Each system must be calibrated to determine the system error (that is, the relation between IAS and CAS) in flight and during the accelerated take-off ground run. The ground run calibra-tion must be determined—

(1) From 0.8 of the minimum value of V1 to the maximum value of V2, consid-ering the approved ranges of altitude and weight; and

(2) With the flaps and power settings corresponding to the values determined in the establishment of the takeoff path under § 25.111 assuming that the critical engine fails at the minimum value of V1.

(c) The airspeed error of the installa-tion, excluding the airspeed indicator

instrument calibration error, may not exceed three percent or five knots, whichever is greater, throughout the speed range, from—

(1) VMO to 1.23 VSR1, with flaps re-tracted; and

(2) 1.23 VSR0 to VFE with flaps in the landing position.

(d) From 1.23 VSR to the speed at which stall warning begins, the IAS must change perceptibly with CAS and in the same sense, and at speeds below stall warning speed the IAS must not change in an incorrect sense.

(e) From VMO to VMO + 2/3 (VDF ¥

VMO), the IAS must change perceptibly with CAS and in the same sense, and at higher speeds up to VDF the IAS must not change in an incorrect sense.

(f) There must be no indication of airspeed that would cause undue dif-ficulty to the pilot during the takeoff between the initiation of rotation and the achievement of a steady climbing condition.

(g) The effects of airspeed indicating system lag may not introduce signifi-cant takeoff indicated airspeed bias, or significant errors in takeoff or accel-erate-stop distances.

(h) Each system must be arranged, so far as practicable, to prevent malfunc-tion or serious error due to the entry of moisture, dirt, or other substances.

(i) Each system must have a heated pitot tube or an equivalent means of preventing malfunction due to icing.

(j) Where duplicate airspeed indica-tors are required, their respective pitot tubes must be far enough apart to avoid damage to both tubes in a colli-sion with a bird.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25–109, 67 FR 76656, Dec. 12, 2002]

§ 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 in such manner that the static pressure system performance is least affected by airflow variation, or by moisture or other foreign matter, and that the correlation between air pressure in the static pressure system


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

499


and true ambient atmospheric static pressure is not changed when the air-plane is exposed to the continuous and intermittent maximum icing condi-tions defined in appendix C of this part.

(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 VSR0 with flaps extended and 1.7 VSR1 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 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]

§ 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]


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

500

14 CFR Ch. I (1–1–11 Edition) § 25.1327

§ 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 from providing guidance or control to an unsafe speed.

(i) The flight guidance system func-tions, controls, indications, and alerts must be designed to minimize flightcrew errors and confusion con-cerning the behavior and operation of the flight guidance system. Means must be provided to indicate the cur-rent mode of operation, including any armed modes, transitions, and rever-sions. Selector switch position is not an acceptable means of indication. The controls and indications must be grouped and presented in a logical and consistent manner. The indications must be visible to each pilot under all expected lighting conditions.

(j) Following disengagement of the autopilot, a warning (visual and audi-tory) must be provided to each pilot and be timely and distinct from all other cockpit warnings.

(k) Following disengagement of the autothrust function, a caution must be provided to each pilot.

(l) The autopilot may not create a potential hazard when the flightcrew applies an override force to the flight controls.

(m) During autothrust operation, it must be possible for the flightcrew to move the thrust levers without requir-ing excessive force. The autothrust may not create a potential hazard when the flightcrew applies an override force to the thrust levers.

(n) For purposes of this section, a transient is a disturbance in the con-trol or flight path of the airplane that is not consistent with response to flightcrew inputs or environmental conditions.

(1) A minor transient would not sig-nificantly reduce safety margins and


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

501


would involve flightcrew actions that are well within their capabilities. A minor transient may involve a slight increase in flightcrew workload or some physical discomfort to passengers or cabin crew.

(2) A significant transient may lead to a significant reduction in safety margins, an increase in flightcrew workload, discomfort to the flightcrew, or physical distress to the passengers or cabin crew, possibly including non- fatal injuries. Significant transients do not require, in order to remain within or recover to the normal flight enve-lope, any of the following:

(i) Exceptional piloting skill, alert-ness, or strength.

(ii) Forces applied by the pilot which are greater than those specified in § 25.143(c).

(iii) Accelerations or attitudes in the airplane that might result in further hazard to secured or non-secured occu-pants.

[Doc. No. FAA–2004–18775, 71 FR 18191, Apr. 11, 2006]

§ 25.1331 Instruments using a power supply.

(a) For each instrument required by § 25.1303(b) that uses a power supply, the following apply:

(1) Each instrument must have a vis-ual means integral with, the instru-ment, to indicate when power adequate to sustain proper instrument perform-ance is not being supplied. The power must be measured at or near the point where it enters the instruments. For electric instruments, the power is con-sidered to be adequate when the volt-age is within approved limits.

(2) Each instrument must, in the event of the failure of one power source, be supplied by another power source. This may be accomplished automatically or by manual means.

(3) If an instrument presenting navi-gation data receives information from sources external to that instrument and loss of that information would render the presented data unreliable, the instrument must incorporate a vis-ual means to warn the crew, when such loss of information occurs, that the presented data should not be relied upon.

(b) As used in this section, ‘‘instru-ment’’ includes devices that are phys-ically contained in one unit, and de-vices that are composed of two or more physically separate units or compo-nents connected together (such as a re-mote indicating gyroscopic direction indicator that includes a magnetic sensing element, a gyroscopic unit, an amplifier and an indicator connected together).

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977]

§ 25.1333 Instrument systems. For systems that operate the instru-

ments required by § 25.1303(b) which are located at each pilot’s station—

(a) Means must be provided to con-nect the required instruments at the first pilot’s station to operating sys-tems which are independent of the op-erating systems at other flight crew stations, or other equipment;

(b) The equipment, systems, and in-stallations must be designed so that one display of the information essen-tial to the safety of flight which is pro-vided by the instruments, including at-titude, direction, airspeed, and altitude will remain available to the pilots, without additional crewmember ac-tion, after any single failure or com-bination of failures that is not shown to be extremely improbable; and

(c) Additional instruments, systems, or equipment may not be connected to the operating systems for the required instruments, unless provisions are made to ensure the continued normal functioning of the required instru-ments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable.

[Amdt. 25–23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977]

§ 25.1337 Powerplant instruments. (a) Instruments and instrument lines.

(1) Each powerplant and auxiliary power unit instrument line must meet the requirements of §§ 25.993 and 25.1183.

(2) Each line carrying flammable fluids under pressure must—

(i) Have restricting orifices or other safety devices at the source of pressure


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

502

14 CFR Ch. I (1–1–11 Edition) § 25.1351

to prevent the escape of excessive fluid if the line fails; and

(ii) Be installed and located so that the escape of fluids would not create a hazard.

(3) Each powerplant and auxiliary power unit instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard.

(b) Fuel quantity indicator. There must be means to indicate to the flight crewmembers, the quantity, in gallons or equivalent units, of usable fuel in each tank during flight. In addition—

(1) Each fuel quantity indicator must be calibrated to read ‘‘zero’’ during level flight when the quantity of fuel re-maining in the tank is equal to the un-usable fuel supply determined under § 25.959;

(2) Tanks with interconnected outlets and airspaces may be treated as one tank and need not have separate indi-cators; and

(3) Each exposed sight gauge, used as a fuel quantity indicator, must be pro-tected against damage.

(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each metering component must have a means for bypassing the fuel supply if malfunction of that component se-verely restricts fuel flow.

(d) Oil quantity indicator. There must be a stick gauge or equivalent means to indicate the quantity of oil in each tank. If an oil transfer or reserve oil supply system is installed, there must be a means to indicate to the flight crew, in flight, the quantity of oil in each tank.

(e) Turbopropeller blade position indi-cator. Required turbopropeller blade position indicators must begin indi-cating before the blade moves more than eight degrees below the flight low pitch stop. The source of indication must directly sense the blade position.

(f) Fuel pressure indicator. There must be means to measure fuel pressure, in each system supplying reciprocating engines, at a point downstream of any fuel pump except fuel injection pumps. In addition—

(1) If necessary for the maintenance of proper fuel delivery pressure, there must be a connection to transmit the carburetor air intake static pressure to

the proper pump relief valve connec-tion; and

(2) If a connection is required under paragraph (f)(1) of this section, the gauge balance lines must be independ-ently connected to the carburetor inlet pressure to avoid erroneous readings.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–40, 42 FR 15044, Mar. 17, 1977]

ELECTRICAL SYSTEMS AND EQUIPMENT

§ 25.1351 General.

(a) Electrical system capacity. The re-quired generating capacity, and num-ber and kinds of power sources must—

(1) Be determined by an electrical load analysis; and

(2) Meet the requirements of § 25.1309. (b) Generating system. The generating

system includes electrical power sources, main power busses, trans-mission cables, and associated control, regulation, and protective devices. It must be designed so that—

(1) Power sources function properly when independent and when connected in combination;

(2) No failure or malfunction of any power source can create a hazard or impair the ability of remaining sources to supply essential loads;

(3) The system voltage and frequency (as applicable) at the terminals of all essential load equipment can be main-tained within the limits for which the equipment is designed, during any probable operating condition; and

(4) System transients due to switch-ing, fault clearing, or other causes do not make essential loads inoperative, and do not cause a smoke or fire haz-ard.

(5) There are means accessible, in flight, to appropriate crewmembers for the individual and collective dis-connection of the electrical power sources from the system.

(6) There are means to indicate to ap-propriate crewmembers the generating system quantities essential for the safe operation of the system, such as the voltage and current supplied by each generator.

(c) External power. If provisions are made for connecting external power to the airplane, and that external power


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

503


can be electrically connected to equip-ment other than that used for engine starting, means must be provided to ensure that no external power supply having a reverse polarity, or a reverse phase sequence, can supply power to the airplane’s electrical system.

(d) Operation without normal electrical power. It must be shown by analysis, tests, or both, that the airplane can be operated safely in VFR conditions, for a period of not less than five minutes, with the normal electrical power (elec-trical power sources excluding the bat-tery) inoperative, with critical type fuel (from the standpoint of flameout and restart capability), and with the airplane initially at the maximum cer-tificated altitude. Parts of the elec-trical system may remain on if—

(1) A single malfunction, including a wire bundle or junction box fire, can-not result in loss of both the part turned off and the part turned on; and

(2) The parts turned on are elec-trically and mechanically isolated from the parts turned off.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–72, 55 FR 29785, July 20, 1990]

§ 25.1353 Electrical equipment and in-stallations.

(a) Electrical equipment and controls must be installed so that operation of any one unit or system of units will not adversely affect the simultaneous operation of any other electrical unit or system essential to safe operation. Any electrical interference likely to be present in the airplane must not result in hazardous effects on the airplane or its systems.

(b) Storage batteries must be de-signed and installed as follows:

(1) Safe cell temperatures and pres-sures must be maintained during any probable charging or discharging con-dition. No uncontrolled increase in cell temperature may result when the bat-tery is recharged (after previous com-plete discharge)—

(i) At maximum regulated voltage or power;

(ii) During a flight of maximum dura-tion; and

(iii) Under the most adverse cooling condition likely to occur in service.

(2) Compliance with paragraph (b)(1) of this section must be shown by test unless experience with similar bat-teries and installations has shown that maintaining safe cell temperatures and pressures presents no problem.

(3) No explosive or toxic gases emit-ted by any battery in normal oper-ation, or as the result of any probable malfunction in the charging system or battery installation, may accumulate in hazardous quantities within the air-plane.

(4) No corrosive fluids or gases that may escape from the battery may dam-age surrounding airplane structures or adjacent essential equipment.

(5) Each nickel cadmium battery in-stallation must have provisions to pre-vent any hazardous effect on structure or essential systems that may be caused by the maximum amount of heat the battery can generate during a short circuit of the battery or of indi-vidual cells.

(6) Nickel cadmium battery installa-tions must have—

(i) A system to control the charging rate of the battery automatically so as to prevent battery overheating;

(ii) A battery temperature sensing and over-temperature warning system with a means for disconnecting the battery from its charging source in the event of an over-temperature condi-tion; or

(iii) A battery failure sensing and warning system with a means for dis-connecting the battery from its charg-ing source in the event of battery fail-ure.

(c) Electrical bonding must provide an adequate electrical return path under both normal and fault condi-tions, on airplanes having grounded electrical systems.

[Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1355 Distribution system. (a) The distribution system includes

the distribution busses, their associ-ated feeders, and each control and pro-tective device.

(b) [Reserved] (c) If two independent sources of elec-

trical power for particular equipment or systems are required by this chap-ter, in the event of the failure of one power source for such equipment or


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

504

14 CFR Ch. I (1–1–11 Edition) § 25.1357

system, another power source (includ-ing its separate feeder) must be auto-matically provided or be manually se-lectable to maintain equipment or sys-tem operation.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–23, 35 FR 5679, Apr. 8, 1970; Amdt. 25–38, 41 FR 55468, Dec. 20, 1976]

§ 25.1357 Circuit protective devices.

(a) Automatic protective devices must be used to minimize distress to the electrical system and hazard to the airplane in the event of wiring faults or serious malfunction of the system or connected equipment.

(b) The protective and control de-vices in the generating system must be designed to de-energize and disconnect faulty power sources and power trans-mission equipment from their associ-ated busses with sufficient rapidity to provide protection from hazardous over-voltage and other malfunctioning.

(c) Each resettable circuit protective device must be designed so that, when an overload or circuit fault exists, it will open the circuit irrespective of the position of the operating control.

(d) If the ability to reset a circuit breaker or replace a fuse is essential to safety in flight, that circuit breaker or fuse must be located and identified so that it can be readily reset or replaced in flight. Where fuses are used, there must be spare fuses for use in flight equal to at least 50% of the number of fuses of each rating required for com-plete circuit protection.

(e) Each circuit for essential loads must have individual circuit protec-tion. However, individual protection for each circuit in an essential load system (such as each position light cir-cuit in a system) is not required.

(f) For airplane systems for which the ability to remove or reset power during normal operations is necessary, the system must be designed so that circuit breakers are not the primary means to remove or reset system power unless specifically designed for use as a switch.

(g) Automatic reset circuit breakers may be used as integral protectors for electrical equipment (such as thermal

cut-outs) if there is circuit protection to protect the cable to the equipment.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§ 25.1360 Precautions against injury.

(a) Shock. The electrical system must be designed to minimize risk of electric shock to crew, passengers, and servicing personnel and to mainte-nance personnel using normal pre-cautions.

(b) Burns. The temperature of any part that may be handled by a crew-member during normal operations must not cause dangerous inadvertent movement by the crewmember or in-jury to the crewmember.

[Amdt. 25–123, 72 FR 63406, Nov. 8, 2007]

§ 25.1362 Electrical supplies for emer-gency conditions.

A suitable electrical supply must be provided to those services required for emergency procedures after an emer-gency landing or ditching. The circuits for these services must be designed, protected, and installed so that the risk of the services being rendered inef-fective under these emergency condi-tions is minimized.

[Amdt. 25–123, 72 FR 63406, Nov. 8, 2007]

§ 25.1363 Electrical system tests.

(a) When laboratory tests of the elec-trical system are conducted—

(1) The tests must be performed on a mock-up using the same generating equipment used in the airplane;

(2) The equipment must simulate the electrical characteristics of the dis-tribution wiring and connected loads to the extent necessary for valid test re-sults; and

(3) Laboratory generator drives must simulate the actual prime movers on the airplane with respect to their reac-tion to generator loading, including loading due to faults.

(b) For each flight condition that cannot be simulated adequately in the laboratory or by ground tests on the airplane, flight tests must be made.


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

505


§ 25.1365 Electrical appliances, motors, and transformers.

(a) Domestic appliances must be de-signed and installed so that in the event of failures of the electrical sup-ply or control system, the require-ments of § 25.1309(b), (c), and (d) will be satisfied. Domestic appliances are items such as cooktops, ovens, coffee makers, water heaters, refrigerators, and toilet flush systems that are placed on the airplane to provide serv-ice amenities to passengers.

(b) Galleys and cooking appliances must be installed in a way that mini-mizes risk of overheat or fire.

(c) Domestic appliances, particularly those in galley areas, must be installed or protected so as to prevent damage or contamination of other equipment or systems from fluids or vapors which may be present during normal oper-ation or as a result of spillage, if such damage or contamination could create a hazardous condition.

(d) Unless compliance with § 25.1309(b) is provided by the circuit protective device required by § 25.1357(a), electric motors and transformers, including those installed in domestic systems, must have a suitable thermal protec-tion device to prevent overheating under normal operation and failure conditions, if overheating could create a smoke or fire hazard.

[Amdt. 25–123, 72 FR 63406, Nov. 8, 2007]

LIGHTS

§ 25.1381 Instrument lights.

(a) The instrument lights must— (1) Provide sufficient illumination to

make each instrument, switch and other device necessary for safe oper-ation easily readable unless sufficient illumination is available from another source; and

(2) Be installed so that— (i) Their direct rays are shielded from

the pilot’s eyes; and (ii) No objectionable reflections are

visible to the pilot. (b) Unless undimmed instrument

lights are satisfactory under each ex-pected flight condition, there must be a

means to control the intensity of illu-mination.


§ 25.1383 Landing lights. (a) Each landing light must be ap-

proved, and must be installed so that— (1) No objectionable glare is visible

to the pilot; (2) The pilot is not adversely affected

by halation; and (3) It provides enough light for night

landing. (b) Except when one switch is used

for the lights of a multiple light instal-lation at one location, there must be a separate switch for each light.

(c) There must be a means to indicate to the pilots when the landing lights are extended.

§ 25.1385 Position light system installa-tion.

(a) General. Each part of each posi-tion light system must meet the appli-cable requirements of this section and each system as a whole must meet the requirements of §§ 25.1387 through 25.1397.

(b) Forward position lights. Forward position lights must consist of a red and a green light spaced laterally as far apart as practicable and installed forward on the airplane so that, with the airplane in the normal flying posi-tion, the red light is on the left side and the green light is on the right side. Each light must be approved.

(c) Rear position light. The rear posi-tion light must be a white light mount-ed as far aft as practicable on the tail or on each wing tip, and must be ap-proved.

(d) Light covers and color filters. Each light cover or color filter must be at least flame resistant and may not change color or shape or lose any ap-preciable light transmission during normal use.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976]

§ 25.1387 Position light system dihe-dral angles.

(a) Except as provided in paragraph (e) of this section, each forward and


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

506

14 CFR Ch. I (1–1–11 Edition) § 25.1389

rear position light must, as installed, show unbroken light within the dihe-dral angles described in this section.

(b) Dihedral angle L (left) is formed by two intersecting vertical planes, the first parallel to the longitudinal axis of the airplane, and the other at 110 de-grees to the left of the first, as viewed when looking forward along the longi-tudinal axis.

(c) Dihedral angle R (right) is formed by two intersecting vertical planes, the first parallel to the longitudinal axis of the airplane, and the other at 110 de-grees to the right of the first, as viewed when looking forward along the longi-tudinal axis.

(d) Dihedral angle A (aft) is formed by two intersecting vertical planes making angles of 70 degrees to the right and to the left, respectively, to a vertical plane passing through the lon-gitudinal axis, as viewed when looking aft along the longitudinal axis.

(e) If the rear position light, when mounted as far aft as practicable in ac-cordance with § 25.1385(c), cannot show unbroken light within dihedral angle A (as defined in paragraph (d) of this sec-tion), a solid angle or angles of ob-structed visibility totaling not more than 0.04 steradians is allowable within that dihedral angle, if such solid angle is within a cone whose apex is at the rear position light and whose elements make an angle of 30° with a vertical line passing through the rear position light.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–30, 36 FR 21278, Nov. 5, 1971]

§ 25.1389 Position light distribution and intensities.

(a) General. The intensities prescribed in this section must be provided by new equipment with light covers and color filters in place. Intensities must be de-termined with the light source oper-ating at a steady value equal to the av-erage luminous output of the source at the normal operating voltage of the airplane. The light distribution and in-tensity of each position light must meet the requirements of paragraph (b) of this section.

(b) Forward and rear position lights. The light distribution and intensities of forward and rear position lights

must be expressed in terms of min-imum intensities in the horizontal plane, minimum intensities in any vertical plane, and maximum inten-sities in overlapping beams, within di-hedral angles L, R, and A, and must meet the following requirements:

(1) Intensities in the horizontal plane. Each intensity in the horizontal plane (the plane containing the longitudinal axis of the airplane and perpendicular to the plane of symmetry of the air-plane) must equal or exceed the values in § 25.1391.

(2) Intensities in any vertical plane. Each intensity in any vertical plane (the plane perpendicular to the hori-zontal plane) must equal or exceed the appropriate value in § 25.1393, where I is the minimum intensity prescribed in § 25.1391 for the corresponding angles in the horizontal plane.

(3) Intensities in overlaps between adja-cent signals. No intensity in any over-lap between adjacent signals may ex-ceed the values given in § 25.1395, except that higher intensities in overlaps may be used with main beam intensities substantially greater than the minima specified in §§ 25.1391 and 25.1393 if the overlap intensities in relation to the main beam intensities do not adversely affect signal clarity. When the peak in-tensity of the forward position lights is more than 100 candles, the maximum overlap intensities between them may exceed the values given in § 25.1395 if the overlap intensity in Area A is not more than 10 percent of peak position light intensity and the overlap inten-sity in Area B is not greater than 2.5 percent of peak position light inten-sity.

§ 25.1391 Minimum intensities in the horizontal plane of forward and rear position lights.

Each position light intensity must equal or exceed the applicable values in the following table:

Dihedral angle (light in-cluded)

Angle from right or left of longitu-dinal axis, meas-ured from dead

ahead

Intensity (candles)

L and R (forward red and green).

0° to 10° ..............10° to 20° ............20° to 110° ..........

40 30 5

A (rear white) ..................... 110° to 180° ........ 20


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

507


§ 25.1393 Minimum intensities in any vertical plane of forward and rear position lights.

Each position light intensity must equal or exceed the applicable values in the following table:

Angle above or below the horizontal plane Intensity, l

0° ......................................................................... 1.00 0° to 5° ................................................................ 0.90 5° to 10° .............................................................. 0.80 10° to 15° ............................................................ 0.70 15° to 20° ............................................................ 0.50 20° to 30° ............................................................ 0.30 30° to 40° ............................................................ 0.10 40° to 90° ............................................................ 0.05

§ 25.1395 Maximum intensities in over-lapping beams of forward and rear position lights.

No position light intensity may ex-ceed the applicable values in the fol-lowing table, except as provided in § 25.1389(b)(3).

Overlaps

Maximum intensity

Area A (candles)

Area B (candles)

Green in dihedral angle L ............. 10 1 Red in dihedral angle R ................ 10 1 Green in dihedral angle A ............. 5 1 Red in dihedral angle A ................ 5 1 Rear white in dihedral angle L ...... 5 1 Rear white in dihedral angle R ..... 5 1

Where— (a) Area A includes all directions in

the adjacent dihedral angle that pass through the light source and intersect the common boundary plane at more than 10 degrees but less than 20 de-grees; and

(b) Area B includes all directions in the adjacent dihedral angle that pass through the light source and intersect the common boundary plane at more than 20 degrees.

§ 25.1397 Color specifications. Each position light color must have

the applicable International Commis-sion on Illumination chromaticity co-ordinates as follows:

(a) Aviation red—

y is not greater than 0.335; and z is not greater than 0.002.

(b) Aviation green—

x is not greater than 0.440¥0.320y ; x is not greater than y¥0.170; and y is not less than 0.390¥0.170x.

(c) Aviation white—

x is not less than 0.300 and not greater than 0.540;

y is not less than x¥0.040; or y0¥0.010, which-ever is the smaller; and

y is not greater than x+0.020 nor 0.636¥0.400x; Where y0 is the y coordinate of the Planckian radiator for the value of x considered.


§ 25.1399 Riding light.

(a) Each riding (anchor) light re-quired for a seaplane or amphibian must be installed so that it can—

(1) Show a white light for at least 2 nautical miles at night under clear at-mospheric conditions; and

(2) Show the maximum unbroken light practicable when the airplane is moored or drifting on the water.

(b) Externally hung lights may be used.

§ 25.1401 Anticollision light system.

(a) General. The airplane must have an anticollision light system that—

(1) Consists of one or more approved anticollision lights located so that their light will not impair the crew’s vision or detract from the conspicuity of the position lights; and

(2) Meets the requirements of para-graphs (b) through (f) of this section.

(b) Field of coverage. The system must consist of enough lights to illuminate the vital areas around the airplane considering the physical configuration and flight characteristics of the air-plane. The field of coverage must ex-tend in each direction within at least 75 degrees above and 75 degrees below the horizontal plane of the airplane, except that a solid angle or angles of obstructed visibility totaling not more than 0.03 steradians is allowable within a solid angle equal to 0.15 steradians centered about the longitudinal axis in the rearward direction.

(c) Flashing characteristics. The ar-rangement of the system, that is, the number of light sources, beam width, speed of rotation, and other character-istics, must give an effective flash fre-quency of not less than 40, nor more than 100 cycles per minute. The effec-tive flash frequency is the frequency at


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

508

14 CFR Ch. I (1–1–11 Edition) § 25.1403

which the airplane’s complete anti-collision light system is observed from a distance, and applies to each sector of light including any overlaps that exist when the system consists of more than one light source. In overlaps, flash frequencies may exceed 100, but not 180 cycles per minute.

(d) Color. Each anticollision light must be either aviation red or aviation white and must meet the applicable re-quirements of § 25.1397.

(e) Light intensity. The minimum light intensities in all vertical planes, measured with the red filter (if used) and expressed in terms of ‘‘effective’’ in-tensities, must meet the requirements of paragraph (f) of this section. The fol-lowing relation must be assumed:

II(t)dt

t tet

t

2 1

1

2

=+ −( )

∫0 2.

where: Ie=effective intensity (candles). I(t)=instantaneous intensity as a function of

time. t2—t1=flash time interval (seconds).

Normally, the maximum value of effec-tive intensity is obtained when t2 and t1 are chosen so that the effective inten-sity is equal to the instantaneous in-tensity at t2 and t1.

(f) Minimum effective intensities for anticollision lights. Each anticollision light effective intensity must equal or exceed the applicable values in the fol-lowing table.

Angle above or below the horizontal plane Effective intensity (candles)

0° to 5° ................................................................ 400 5° to 10° .............................................................. 240 10° to 20° ............................................................ 80 20° to 30° ............................................................ 40 30° to 75° ............................................................ 20

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–27, 36 FR 12972, July 10, 1971; Amdt. 25–41, 42 FR 36970, July 18, 1977]

§ 25.1403 Wing icing detection lights. Unless operations at night in known

or forecast icing conditions are prohib-ited by an operating limitation, a means must be provided for illu-minating or otherwise determining the formation of ice on the parts of the

wings that are critical from the stand-point of ice accumulation. Any illu-mination that is used must be of a type that will not cause glare or reflection that would handicap crewmembers in the performance of their duties.

[Amdt. 25–38, 41 FR 55468, Dec. 20, 1976]

SAFETY EQUIPMENT

§ 25.1411 General.

(a) Accessibility. Required safety equipment to be used by the crew in an emergency must be readily accessible.

(b) Stowage provisions. Stowage provi-sions for required emergency equip-ment must be furnished and must—

(1) Be arranged so that the equip-ment is directly accessible and its loca-tion is obvious; and

(2) Protect the safety equipment from inadvertent damage.

(c) Emergency exit descent device. The stowage provisions for the emergency exit descent devices required by § 25.810(a) must be at each exit for which they are intended.

(d) Liferafts. (1) The stowage provi-sions for the liferafts described in § 25.1415 must accommodate enough rafts for the maximum number of occu-pants for which certification for ditch-ing is requested.

(2) Liferafts must be stowed near exits through which the rafts can be launched during an unplanned ditch-ing.

(3) Rafts automatically or remotely released outside the airplane must be attached to the airplane by means of the static line prescribed in § 25.1415.

(4) The stowage provisions for each portable liferaft must allow rapid de-tachment and removal of the raft for use at other than the intended exits.

(e) Long-range signaling device. The stowage provisions for the long-range signaling device required by § 25.1415 must be near an exit available during an unplanned ditching.

(f) Life preserver stowage provisions. The stowage provisions for life pre-servers described in § 25.1415 must ac-commodate one life preserver for each occupant for which certification for ditching is requested. Each life pre-server must be within easy reach of each seated occupant.

VerDate Mar<15>2010 14:10 Mar 01, 2011 Jkt 223043 PO 00000 Frm 00518 Fmt 8010 Sfmt 8010 Y:\SGML\223043.XXX 223043 EC

28S

E91

.049

</M

AT

H>

ww

oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

509


(g) Life line stowage provisions. If cer-tification for ditching under § 25.801 is requested, there must be provisions to store life lines. These provisions must—

(1) Allow one life line to be attached to each side of the fuselage; and

(2) Be arranged to allow the life lines to be used to enable the occupants to stay on the wing after ditching.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972; Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. 25–53, 45 FR 41593, June 19, 1980; Amdt. 25–70, 54 FR 43925, Oct. 27, 1989; Amdt. 25–79, 58 FR 45229, Aug. 26, 1993; Amdt. 25–116, 69 FR 62789, Oct. 27, 2004]

§ 25.1415 Ditching equipment.

(a) Ditching equipment used in air-planes to be certificated for ditching under § 25.801, and required by the oper-ating rules of this chapter, must meet the requirements of this section.

(b) Each liferaft and each life pre-server must be approved. In addition—

(1) Unless excess rafts of enough ca-pacity are provided, the buoyancy and seating capacity beyond the rated ca-pacity of the rafts must accommodate all occupants of the airplane in the event of a loss of one raft of the largest rated capacity; and

(2) Each raft must have a trailing line, and must have a static line de-signed to hold the raft near the air-plane but to release it if the airplane becomes totally submerged.

(c) Approved survival equipment must be attached to each liferaft.

(d) There must be an approved sur-vival type emergency locator trans-mitter for use in one life raft.

(e) For airplanes not certificated for ditching under § 25.801 and not having approved life preservers, there must be an approved flotation means for each occupant. This means must be within easy reach of each seated occupant and must be readily removable from the airplane.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–29, 36 FR 18722, Sept. 21, 1971; Amdt 25–50, 45 FR 38348, June 9, 1980; Amdt. 25–72, 55 FR 29785, July 20, 1990; Amdt. 25–82, 59 FR 32057, June 21, 1994]

§ 25.1419 Ice protection.

If the applicant seeks certification for flight in icing conditions, the air-plane must be able to safely operate in the continuous maximum and inter-mittent maximum icing conditions of appendix C. To establish this—

(a) An analysis must be performed to establish that the ice protection for the various components of the airplane is adequate, taking into account the various airplane operational configura-tions; and

(b) To verify the ice protection anal-ysis, to check for icing anomalies, and to demonstrate that the ice protection system and its components are effec-tive, the airplane or its components must be flight tested in the various operational configurations, in meas-ured natural atmospheric icing condi-tions and, as found necessary, by one or more of the following means:

(1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the com-ponents.

(2) Flight dry air tests of the ice pro-tection system as a whole, or of its in-dividual components.

(3) Flight tests of the airplane or its components in measured simulated icing conditions.

(c) Caution information, such as an amber caution light or equivalent, must be provided to alert the flightcrew when the anti-ice or de-ice system is not functioning normally.

(d) For turbine engine powered air-planes, the ice protection provisions of this section are considered to be appli-cable primarily to the airframe. For the powerplant installation, certain ad-ditional provisions of subpart E of this part may be found applicable. (e) One of the following methods of icing detec-tion and activation of the airframe ice protection system must be provided:

(1) A primary ice detection system that automatically activates or alerts the flightcrew to activate the airframe ice protection system;

(2) A definition of visual cues for rec-ognition of the first sign of ice accre-tion on a specified surface combined with an advisory ice detection system that alerts the flightcrew to activate the airframe ice protection system; or


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

510

14 CFR Ch. I (1–1–11 Edition) § 25.1421

(3) Identification of conditions con-ducive to airframe icing as defined by an appropriate static or total air tem-perature and visible moisture for use by the flightcrew to activate the air-frame ice protection system.

(f) Unless the applicant shows that the airframe ice protection system need not be operated during specific phases of flight, the requirements of paragraph (e) of this section are appli-cable to all phases of flight.

(g) After the initial activation of the airframe ice protection system—

(1) The ice protection system must be designed to operate continuously;

(2) The airplane must be equipped with a system that automatically cy-cles the ice protection system; or

(3) An ice detection system must be provided to alert the flightcrew each time the ice protection system must be cycled.

(h) Procedures for operation of the ice protection system, including acti-vation and deactivation, must be estab-lished and documented in the Airplane Flight Manual.

[Amdt. 25–72, 55 FR 29785, July 20, 1990, as amended by Amdt. 25–121, 72 FR 44669, Aug. 8, 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009]

§ 25.1421 Megaphones.

If a megaphone is installed, a re-straining means must be provided that is capable of restraining the mega-phone when it is subjected to the ulti-mate inertia forces specified in § 25.561(b)(3).

[Amdt. 25–41, 42 FR 36970, July 18, 1977]

§ 25.1423 Public address system.

A public address system required by this chapter must—

(a) Be powerable when the aircraft is in flight or stopped on the ground, after the shutdown or failure of all en-gines and auxiliary power units, or the disconnection or failure of all power sources dependent on their continued operation, for—

(1) A time duration of at least 10 min-utes, including an aggregate time dura-tion of at least 5 minutes of announce-ments made by flight and cabin crew-members, considering all other loads which may remain powered by the

same source when all other power sources are inoperative; and

(2) An additional time duration in its standby state appropriate or required for any other loads that are powered by the same source and that are essential to safety of flight or required during emergency conditions.

(b) Be capable of operation within 3 seconds from the time a microphone is removed from its stowage.

(c) Be intelligible at all passenger seats, lavatories, and flight attendant seats and work stations.

(d) Be designed so that no unused, unstowed microphone will render the system inoperative.

(e) Be capable of functioning inde-pendently of any required crewmember interphone system.

(f) Be accessible for immediate use from each of two flight crewmember stations in the pilot compartment.

(g) For each required floor-level pas-senger emergency exit which has an ad-jacent flight attendant seat, have a microphone which is readily accessible to the seated flight attendant, except that one microphone may serve more than one exit, provided the proximity of the exits allows unassisted verbal communication between seated flight attendants.

[Doc. No. 26003, 58 FR 45229, Aug. 26, 1993, as amended by Amdt. 25–115, 69 FR 40527, July 2, 2004]

MISCELLANEOUS EQUIPMENT

§ 25.1431 Electronic equipment. (a) In showing compliance with

§ 25.1309 (a) and (b) with respect to radio and electronic equipment and their installations, critical environ-mental conditions must be considered.

(b) Radio and electronic equipment must be supplied with power under the requirements of § 25.1355(c).

(c) Radio and electronic equipment, controls, and wiring must be installed so that operation of any one unit or system of units will not adversely af-fect the simultaneous operation of any other radio or electronic unit, or sys-tem of units, required by this chapter.

(d) Electronic equipment must be de-signed and installed such that it does not cause essential loads to become in-operative as a result of electrical


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

511


power supply transients or transients from other causes.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–113, 69 FR 12530, Mar. 16, 2004]

§ 25.1433 Vacuum systems. There must be means, in addition to

the normal pressure relief, to auto-matically relieve the pressure in the discharge lines from the vacuum air pump when the delivery temperature of the air becomes unsafe.


§ 25.1435 Hydraulic systems. (a) Element design. Each element of

the hydraulic system must be designed to:

(1) Withstand the proof pressure without permanent deformation that would prevent it from performing its intended functions, and the ultimate pressure without rupture. The proof and ultimate pressures are defined in terms of the design operating pressure (DOP) as follows:

Element Proof (xDOP)

Ultimate (xDOP)

1. Tubes and fittings. ......................... 1.5 3.0 2. Pressure vessels containing gas:

High pressure (e.g., accumulators) 3.0 4.0 Low pressure (e.g., reservoirs) ...... 1.5 3.0

3. Hoses ............................................ 2.0 4.0 4. All other elements ......................... 1.5 2.0

(2) Withstand, without deformation that would prevent it from performing its intended function, the design oper-ating pressure in combination with limit structural loads that may be im-posed;

(3) Withstand, without rupture, the design operating pressure multiplied by a factor of 1.5 in combination with ulti-mate structural load that can reason-ably occur simultaneously;

(4) Withstand the fatigue effects of all cyclic pressures, including tran-sients, and associated externally in-duced loads, taking into account the consequences of element failure; and

(5) Perform as intended under all en-vironmental conditions for which the airplane is certificated.

(b) System design. Each hydraulic sys-tem must:

(1) Have means located at a flightcrew station to indicate appro-priate system parameters, if

(i) It performs a function necessary for continued safe flight and landing; or

(ii) In the event of hydraulic system malfunction, corrective action by the crew to ensure continued safe flight and landing is necessary;

(2) Have means to ensure that system pressures, including transient pres-sures and pressures from fluid volu-metric changes in elements that are likely to remain closed long enough for such changes to occur, are within the design capabilities of each element, such that they meet the requirements defined in § 25.1435(a)(1) through (a)(5);

(3) Have means to minimize the re-lease of harmful or hazardous con-centrations of hydraulic fluid or vapors into the crew and passenger compart-ments during flight;

(4) Meet the applicable requirements of §§ 25.863, 25.1183, 25.1185, and 25.1189 if a flammable hydraulic fluid is used; and

(5) Be designed to use any suitable hydraulic fluid specified by the air-plane manufacturer, which must be identified by appropriate markings as required by § 25.1541.

(c) Tests. Tests must be conducted on the hydraulic system(s), and/or sub-system(s) and elements, except that analysis may be used in place of or to supplement testing, where the analysis is shown to be reliable and appropriate. All internal and external influences must be taken into account to an ex-tent necessary to evaluate their ef-fects, and to assure reliable system and element functioning and integration. Failure or unacceptable deficiency of an element or system must be cor-rected and be sufficiently retested, where necessary.

(1) The system(s), subsystem(s), or element(s) must be subjected to per-formance, fatigue, and endurance tests representative of airplane ground and flight operations.

(2) The complete system must be tested to determine proper functional performance and relation to the other systems, including simulation of rel-evant failure conditions, and to sup-port or validate element design.


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

512

14 CFR Ch. I (1–1–11 Edition) § 25.1438

(3) The complete hydraulic system(s) must be functionally tested on the air-plane in normal operation over the range of motion of all associated user systems. The test must be conducted at the system relief pressure or 1.25 times the DOP if a system pressure relief de-vice is not part of the system design. Clearances between hydraulic system elements and other systems or struc-tural elements must remain adequate and there must be no detrimental ef-fects.

[Doc. No. 28617, 66 FR 27402, May 16, 2001]

§ 25.1438 Pressurization and pneu-matic systems.

(a) Pressurization system elements must be burst pressure tested to 2.0 times, and proof pressure tested to 1.5 times, the maximum normal operating pressure.

(b) Pneumatic system elements must be burst pressure tested to 3.0 times, and proof pressure tested to 1.5 times, the maximum normal operating pres-sure.

(c) An analysis, or a combination of analysis and test, may be substituted for any test required by paragraph (a) or (b) of this section if the Adminis-trator finds it equivalent to the re-quired test.

[Amdt. 25–41, 42 FR 36971, July 18, 1977]

§ 25.1439 Protective breathing equip-ment.

(a) Fixed (stationary, or built in) pro-tective breathing equipment must be installed for the use of the flightcrew, and at least one portable protective breathing equipment shall be located at or near the flight deck for use by a flight crewmember. In addition, port-able protective breathing equipment must be installed for the use of appro-priate crewmembers for fighting fires in compartments accessible in flight other than the flight deck. This in-cludes isolated compartments and upper and lower lobe galleys, in which crewmember occupancy is permitted during flight. Equipment must be in-stalled for the maximum number of crewmembers expected to be in the area during any operation.

(b) For protective breathing equip-ment required by paragraph (a) of this

section or by the applicable Operating Regulations:

(1) The equipment must be designed to protect the appropriate crewmember from smoke, carbon dioxide, and other harmful gases while on flight deck duty or while combating fires.

(2) The equipment must include— (i) Masks covering the eyes, nose and

mouth, or (ii) Masks covering the nose and

mouth, plus accessory equipment to cover the eyes.

(3) Equipment, including portable equipment, must allow communication with other crewmembers while in use. Equipment available at flightcrew as-signed duty stations must also enable the flightcrew to use radio equipment.

(4) The part of the equipment pro-tecting the eyes shall not cause any ap-preciable adverse effect on vision and must allow corrective glasses to be worn.

(5) The equipment must supply pro-tective oxygen of 15 minutes duration per crewmember at a pressure altitude of 8,000 feet with a respiratory minute volume of 30 liters per minute BTPD. The equipment and system must be de-signed to prevent any inward leakage to the inside of the device and prevent any outward leakage causing signifi-cant increase in the oxygen content of the local ambient atmosphere. If a de-mand oxygen system is used, a supply of 300 liters of free oxygen at 70 °F. and 760 mm. Hg. pressure is considered to be of 15-minute duration at the pre-scribed altitude and minute volume. If a continuous flow open circuit protec-tive breathing system is used, a flow rate of 60 liters per minute at 8,000 feet (45 liters per minute at sea level) and a supply of 600 liters of free oxygen at 70 °F. and 760 mm. Hg. pressure is consid-ered to be of 15-minute duration at the prescribed altitude and minute volume. Continuous flow systems must not in-crease the ambient oxygen content of the local atmosphere above that of de-mand systems. BTPD refers to body temperature conditions (that is, 37 °C., at ambient pressure, dry).

(6) The equipment must meet the re-quirements of § 25.1441.

[Doc. No. FAA–2002–13859, 69 FR 40528, July 2, 2004]


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

513


§ 25.1441 Oxygen equipment and sup-ply.

(a) If certification with supplemental oxygen equipment is requested, the equipment must meet the requirements of this section and §§ 25.1443 through 25.1453.

(b) The oxygen system must be free from hazards in itself, in its method of operation, and in its effect upon other components.

(c) There must be a means to allow the crew to readily determine, during flight, the quantity of oxygen available in each source of supply.

(d) The oxygen flow rate and the oxy-gen equipment for airplanes for which certification for operation above 40,000 feet is requested must be approved.

§ 25.1443 Minimum mass flow of sup-plemental oxygen.

(a) If continuous flow equipment is installed for use by flight crew-members, the minimum mass flow of supplemental oxygen required for each crewmember may not be less than the flow required to maintain, during in-spiration, a mean tracheal oxygen par-tial pressure of 149 mm. Hg. when breathing 15 liters per minute, BTPS, and with a maximum tidal volume of 700 cc. with a constant time interval between respirations.

(b) If demand equipment is installed for use by flight crewmembers, the minimum mass flow of supplemental oxygen required for each crewmember may not be less than the flow required to maintain, during inspiration, a mean tracheal oxygen partial pressure of 122 mm. Hg., up to and including a cabin pressure altitude of 35,000 feet, and 95 percent oxygen between cabin pressure altitudes of 35,000 and 40,000 feet, when breathing 20 liters per minute BTPS. In addition, there must be means to allow the crew to use undi-luted oxygen at their discretion.

(c) For passengers and cabin attend-ants, the minimum mass flow of sup-plemental oxygen required for each person at various cabin pressure alti-tudes may not be less than the flow re-quired to maintain, during inspiration and while using the oxygen equipment (including masks) provided, the fol-lowing mean tracheal oxygen partial pressures:

(1) At cabin pressure altitudes above 10,000 feet up to and including 18,500 feet, a mean tracheal oxygen partial pressure of 100 mm. Hg. when breathing 15 liters per minute, BTPS, and with a tidal volume of 700 cc. with a constant time interval between respirations.

(2) At cabin pressure altitudes above 18,500 feet up to and including 40,000 feet, a mean tracheal oxygen partial pressure of 83.8 mm. Hg. when breath-ing 30 liters per minute, BTPS, and with a tidal volume of 1,100 cc. with a constant time interval between res-pirations.

(d) If first-aid oxygen equipment is installed, the minimum mass flow of oxygen to each user may not be less than four liters per minute, STPD. However, there may be a means to de-crease this flow to not less than two li-ters per minute, STPD, at any cabin al-titude. The quantity of oxygen re-quired is based upon an average flow rate of three liters per minute per per-son for whom first-aid oxygen is re-quired.

(e) If portable oxygen equipment is installed for use by crewmembers, the minimum mass flow of supplemental oxygen is the same as specified in para-graph (a) or (b) of this section, which-ever is applicable.

§ 25.1445 Equipment standards for the oxygen distributing system.

(a) When oxygen is supplied to both crew and passengers, the distribution system must be designed for either—

(1) A source of supply for the flight crew on duty and a separate source for the passengers and other crewmembers; or

(2) A common source of supply with means to separately reserve the min-imum supply required by the flight crew on duty.

(b) Portable walk-around oxygen units of the continuous flow, diluter- demand, and straight demand kinds may be used to meet the crew or pas-senger breathing requirements.

§ 25.1447 Equipment standards for ox-ygen dispensing units.

If oxygen dispensing units are in-stalled, the following apply:

(a) There must be an individual dis-pensing unit for each occupant for


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

514

14 CFR Ch. I (1–1–11 Edition) § 25.1449

whom supplemental oxygen is to be supplied. Units must be designed to cover the nose and mouth and must be equipped with a suitable means to re-tain the unit in position on the face. Flight crew masks for supplemental oxygen must have provisions for the use of communication equipment.

(b) If certification for operation up to and including 25,000 feet is requested, an oxygen supply terminal and unit of oxygen dispensing equipment for the immediate use of oxygen by each crew-member must be within easy reach of that crewmember. For any other occu-pants, the supply terminals and dis-pensing equipment must be located to allow the use of oxygen as required by the operating rules in this chapter.

(c) If certification for operation above 25,000 feet is requested, there must be oxygen dispensing equipment meeting the following requirements:

(1) There must be an oxygen dis-pensing unit connected to oxygen sup-ply terminals immediately available to each occupant, wherever seated, and at least two oxygen dispensing units con-nected to oxygen terminals in each lav-atory. The total number of dispensing units and outlets in the cabin must ex-ceed the number of seats by at least 10 percent. The extra units must be as uniformly distributed throughout the cabin as practicable. If certification for operation above 30,000 feet is requested, the dispensing units providing the re-quired oxygen flow must be automati-cally presented to the occupants before the cabin pressure altitude exceeds 15,000 feet. The crew must be provided with a manual means of making the dispensing units immediately available in the event of failure of the automatic system.

(2) Each flight crewmember on flight deck duty must be provided with a quick-donning type oxygen dispensing unit connected to an oxygen supply terminal. This dispensing unit must be immediately available to the flight crewmember when seated at his sta-tion, and installed so that it:

(i) Can be placed on the face from its ready position, properly secured, sealed, and supplying oxygen upon de-mand, with one hand, within five sec-onds and without disturbing eyeglasses

or causing delay in proceeding with emergency duties; and

(ii) Allows, while in place, the per-formance of normal communication functions.

(3) The oxygen dispensing equipment for the flight crewmembers must be:

(i) The diluter demand or pressure de-mand (pressure demand mask with a diluter demand pressure breathing reg-ulator) type, or other approved oxygen equipment shown to provide the same degree of protection, for airplanes to be operated above 25,000 feet.

(ii) The pressure demand (pressure demand mask with a diluter demand pressure breathing regulator) type with mask-mounted regulator, or other ap-proved oxygen equipment shown to provide the same degree of protection, for airplanes operated at altitudes where decompressions that are not ex-tremely improbable may expose the flightcrew to cabin pressure altitudes in excess of 34,000 feet.

(4) Portable oxygen equipment must be immediately available for each cabin attendant. The portable oxygen equipment must have the oxygen dis-pensing unit connected to the portable oxygen supply.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–87, 61 FR 28696, June 5, 1996; Amdt. 25–116, 69 FR 62789, Oct. 27, 2004]

§ 25.1449 Means for determining use of oxygen.

There must be a means to allow the crew to determine whether oxygen is being delivered to the dispensing equip-ment.

§ 25.1450 Chemical oxygen generators. (a) For the purpose of this section, a

chemical oxygen generator is defined as a device which produces oxygen by chemical reaction.

(b) Each chemical oxygen generator must be designed and installed in ac-cordance with the following require-ments:

(1) Surface temperature developed by the generator during operation may not create a hazard to the airplane or to its occupants.

(2) Means must be provided to relieve any internal pressure that may be haz-ardous.


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

515


(c) In addition to meeting the re-quirements in paragraph (b) of this sec-tion, each portable chemical oxygen generator that is capable of sustained operation by successive replacement of a generator element must be placarded to show—

(1) The rate of oxygen flow, in liters per minute;

(2) The duration of oxygen flow, in minutes, for the replaceable generator element; and

(3) A warning that the replaceable generator element may be hot, unless the element construction is such that the surface temperature cannot exceed 100 degrees F.

[Amdt. 25–41, 42 FR 36971, July 18, 1977]

§ 25.1453 Protection of oxygen equip-ment from rupture.

Oxygen pressure tanks, and lines be-tween tanks and the shutoff means, must be—

(a) Protected from unsafe tempera-tures; and

(b) Located where the probability and hazards of rupture in a crash landing are minimized.

§ 25.1455 Draining of fluids subject to freezing.

If fluids subject to freezing may be drained overboard in flight or during ground operation, the drains must be designed and located to prevent the formation of hazardous quantities of ice on the airplane as a result of the drainage.

[Amdt. 25–23, 35 FR 5680, Apr. 8, 1970]

§ 25.1457 Cockpit voice recorders. (a) Each cockpit voice recorder re-

quired by the operating rules of this chapter must be approved and must be installed so that it will record the fol-lowing:

(1) Voice communications trans-mitted from or received in the airplane by radio.

(2) Voice communications of flight crewmembers on the flight deck.

(3) Voice communications of flight crewmembers on the flight deck, using the airplane’s interphone system.

(4) Voice or audio signals identifying navigation or approach aids introduced into a headset or speaker.

(5) Voice communications of flight crewmembers using the passenger loud-speaker system, if there is such a sys-tem and if the fourth channel is avail-able in accordance with the require-ments of paragraph (c)(4)(ii) of this sec-tion.

(6) If datalink communication equip-ment is installed, all datalink commu-nications, using an approved data mes-sage set. Datalink messages must be recorded as the output signal from the communications unit that translates the signal into usable data.

(b) The recording requirements of paragraph (a)(2) of this section must be met by installing a cockpit-mounted area microphone, located in the best position for recording voice commu-nications originating at the first and second pilot stations and voice commu-nications of other crewmembers on the flight deck when directed to those sta-tions. The microphone must be so lo-cated and, if necessary, the pre-amplifiers and filters of the recorder must be so adjusted or supplemented, that the intelligibility of the recorded communications is as high as prac-ticable when recorded under flight cockpit noise conditions and played back. Repeated aural or visual play-back of the record may be used in eval-uating intelligibility.

(c) Each cockpit voice recorder must be installed so that the part of the communication or audio signals speci-fied in paragraph (a) of this section ob-tained from each of the following sources is recorded on a separate chan-nel:

(1) For the first channel, from each boom, mask, or hand-held microphone, headset, or speaker used at the first pilot station.

(2) For the second channel from each boom, mask, or hand-held microphone, headset, or speaker used at the second pilot station.

(3) For the third channel—from the cockpit-mounted area microphone.

(4) For the fourth channel, from— (i) Each boom, mask, or hand-held

microphone, headset, or speaker used at the station for the third and fourth crew members; or

(ii) If the stations specified in para-graph (c)(4)(i) of this section are not re-quired or if the signal at such a station


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

516

14 CFR Ch. I (1–1–11 Edition) § 25.1459

is picked up by another channel, each microphone on the flight deck that is used with the passenger loudspeaker system, if its signals are not picked up by another channel.

(5) As far as is practicable all sounds received by the microphone listed in paragraphs (c)(1), (2), and (4) of this section must be recorded without interruption irrespective of the posi-tion of the interphone-transmitter key switch. The design shall ensure that sidetone for the flight crew is produced only when the interphone, public ad-dress system, or radio transmitters are in use.

(d) Each cockpit voice recorder must be installed so that—

(1)(i) It receives its electrical power from the bus that provides the max-imum reliability for operation of the cockpit voice recorder without jeopard-izing service to essential or emergency loads.

(ii) It remains powered for as long as possible without jeopardizing emer-gency operation of the airplane.

(2) There is an automatic means to simultaneously stop the recorder and prevent each erasure feature from func-tioning, within 10 minutes after crash impact;

(3) There is an aural or visual means for preflight checking of the recorder for proper operation;

(4) Any single electrical failure exter-nal to the recorder does not disable both the cockpit voice recorder and the flight data recorder;

(5) It has an independent power source—

(i) That provides 10 ± 1 minutes of electrical power to operate both the cockpit voice recorder and cockpit- mounted area microphone;

(ii) That is located as close as prac-ticable to the cockpit voice recorder; and

(iii) To which the cockpit voice re-corder and cockpit-mounted area microphone are switched automati-cally in the event that all other power to the cockpit voice recorder is inter-rupted either by normal shutdown or by any other loss of power to the elec-trical power bus; and

(6) It is in a separate container from the flight data recorder when both are required. If used to comply with only

the cockpit voice recorder require-ments, a combination unit may be in-stalled.

(e) The recorder container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and con-sequent heat damage to the recorder from fire.

(1) Except as provided in paragraph (e)(2) of this section, the recorder con-tainer must be located as far aft as practicable, but need not be outside of the pressurized compartment, and may not be located where aft-mounted en-gines may crush the container during impact.

(2) If two separate combination dig-ital flight data recorder and cockpit voice recorder units are installed in-stead of one cockpit voice recorder and one digital flight data recorder, the combination unit that is installed to comply with the cockpit voice recorder requirements may be located near the cockpit.

(f) If the cockpit voice recorder has a bulk erasure device, the installation must be designed to minimize the prob-ability of inadvertent operation and ac-tuation of the device during crash im-pact.

(g) Each recorder container must— (1) Be either bright orange or bright

yellow; (2) Have reflective tape affixed to its

external surface to facilitate its loca-tion under water; and

(3) Have an underwater locating de-vice, when required by the operating rules of this chapter, on or adjacent to the container which is secured in such manner that they are not likely to be separated during crash impact.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–2, 30 FR 3932, Mar. 26, 1965; Amdt. 25–16, 32 FR 13914, Oct. 6, 1967; Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–65, 53 FR 26143, July 11, 1988; Amdt. No. 25– 124, 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009]

§ 25.1459 Flight data recorders. (a) Each flight recorder required by

the operating rules of this chapter must be installed so that—

(1) It is supplied with airspeed, alti-tude, and directional data obtained from sources that meet the accuracy


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

517


requirements of §§ 25.1323, 25.1325, and 25.1327, as appropriate;

(2) The vertical acceleration sensor is rigidly attached, and located longitu-dinally either within the approved cen-ter of gravity limits of the airplane, or at a distance forward or aft of these limits that does not exceed 25 percent of the airplane’s mean aerodynamic chord;

(3)(i) It receives its electrical power from the bus that provides the max-imum reliability for operation of the flight data recorder without jeopard-izing service to essential or emergency loads.

(ii) It remains powered for as long as possible without jeopardizing emer-gency operation of the airplane.

(4) There is an aural or visual means for preflight checking of the recorder for proper recording of data in the stor-age medium;

(5) Except for recorders powered sole-ly by the engine-driven electrical gen-erator system, there is an automatic means to simultaneously stop a re-corder that has a data erasure feature and prevent each erasure feature from functioning, within 10 minutes after crash impact;

(6) There is a means to record data from which the time of each radio transmission either to or from ATC can be determined;

(7) Any single electrical failure exter-nal to the recorder does not disable both the cockpit voice recorder and the flight data recorder; and

(8) It is in a separate container from the cockpit voice recorder when both are required. If used to comply with only the flight data recorder require-ments, a combination unit may be in-stalled. If a combination unit is in-stalled as a cockpit voice recorder to comply with § 25.1457(e)(2), a combina-tion unit must be used to comply with this flight data recorder requirement.

(b) Each nonejectable record con-tainer must be located and mounted so as to minimize the probability of con-tainer rupture resulting from crash im-pact and subsequent damage to the record from fire. In meeting this re-quirement the record container must be located as far aft as practicable, but need not be aft of the pressurized com-partment, and may not be where aft-

mounted engines may crush the con-tainer upon impact.

(c) A correlation must be established between the flight recorder readings of airspeed, altitude, and heading and the corresponding readings (taking into ac-count correction factors) of the first pi-lot’s instruments. The correlation must cover the airspeed range over which the airplane is to be operated, the range of altitude to which the air-plane is limited, and 360 degrees of heading. Correlation may be estab-lished on the ground as appropriate.

(d) Each recorder container must— (1) Be either bright orange or bright

yellow; (2) Have reflective tape affixed to its

external surface to facilitate its loca-tion under water; and

(3) Have an underwater locating de-vice, when required by the operating rules of this chapter, on or adjacent to the container which is secured in such a manner that they are not likely to be separated during crash impact.

(e) Any novel or unique design or operational characteristics of the air-craft shall be evaluated to determine if any dedicated parameters must be re-corded on flight recorders in addition to or in place of existing requirements.

[Amdt. 25–8, 31 FR 127, Jan. 6, 1966, as amend-ed by Amdt. 25–25, 35 FR 13192, Aug. 19, 1970; Amdt. 25–37, 40 FR 2577, Jan. 14, 1975; Amdt. 25–41, 42 FR 36971, July 18, 1977; Amdt. 25–65, 53 FR 26144, July 11, 1988; Amdt. No. 25–124, 73 FR 12563, Mar. 7, 2008; 74 FR 32800, July 9, 2009]

§ 25.1461 Equipment containing high energy rotors.

(a) Equipment containing high en-ergy rotors must meet paragraph (b), (c), or (d) of this section.

(b) High energy rotors contained in equipment must be able to withstand damage caused by malfunctions, vibra-tion, abnormal speeds, and abnormal temperatures. In addition—

(1) Auxiliary rotor cases must be able to contain damage caused by the fail-ure of high energy rotor blades; and

(2) Equipment control devices, sys-tems, and instrumentation must rea-sonably ensure that no operating limi-tations affecting the integrity of high energy rotors will be exceeded in serv-ice.


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

518

14 CFR Ch. I (1–1–11 Edition) § 25.1501

(c) It must be shown by test that equipment containing high energy ro-tors can contain any failure of a high energy rotor that occurs at the highest speed obtainable with the normal speed control devices inoperative.

(d) Equipment containing high en-ergy rotors must be located where rotor failure will neither endanger the occupants nor adversely affect contin-ued safe flight.

[Amdt. 25–41, 42 FR 36971, July 18, 1977]

Subpart G—Operating Limitations and Information

§ 25.1501 General. (a) Each operating limitation speci-

fied in §§ 25.1503 through 25.1533 and other limitations and information nec-essary for safe operation must be es-tablished.

(b) The operating limitations and other information necessary for safe operation must be made available to the crewmembers as prescribed in §§ 25.1541 through 25.1587.

[Amdt. 25–42, 43 FR 2323, Jan. 16, 1978]

OPERATING LIMITATIONS

§ 25.1503 Airspeed limitations: general. When airspeed limitations are a func-

tion of weight, weight distribution, al-titude, or Mach number, limitations corresponding to each critical com-bination of these factors must be estab-lished.

§ 25.1505 Maximum operating limit speed.

The maximum operating limit speed (VMO/MMO airspeed or Mach Number, whichever is critical at a particular al-titude) is a speed that may not be de-liberately exceeded in any regime of flight (climb, cruise, or descent), unless a higher speed is authorized for flight test or pilot training operations. VMO/ MMO must be established so that it is not greater than the design cruising speed VC and so that it is sufficiently below VD/MD or VDF/MDF, to make it highly improbable that the latter speeds will be inadvertently exceeded in operations. The speed margin be-tween VMO/MMO and VD/MD or VDFM/DF may not be less than that determined

under § 25.335(b) or found necessary dur-ing the flight tests conducted under § 25.253.

[Amdt. 25–23, 35 FR 5680, Apr. 8, 1970]

§ 25.1507 Maneuvering speed.

The maneuvering speed must be es-tablished so that it does not exceed the design maneuvering speed VA deter-mined under § 25.335(c).

§ 25.1511 Flap extended speed.

The established flap extended speed VFE must be established so that it does not exceed the design flap speed VF chosen under §§ 25.335(e) and 25.345, for the corresponding flap positions and engine powers.

§ 25.1513 Minimum control speed.

The minimum control speed VMC de-termined under § 25.149 must be estab-lished as an operating limitation.

§ 25.1515 Landing gear speeds.

(a) The established landing gear oper-ating speed or speeds, VLO, may not ex-ceed the speed at which it is safe both to extend and to retract the landing gear, as determined under § 25.729 or by flight characteristics. If the extension speed is not the same as the retraction speed, the two speeds must be des-ignated as VLO(EXT) and VLO(RET), re-spectively.

(b) The established landing gear ex-tended speed VLE may not exceed the speed at which it is safe to fly with the landing gear secured in the fully ex-tended position, and that determined under § 25.729.

[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976]

§ 25.1516 Other speed limitations.

Any other limitation associated with speed must be established.

[Doc. No. 2000–8511, 66 FR 34024, June 26, 2001]

§ 25.1517 Rough air speed, VRA.

A rough air speed, VRA, for use as the recommended turbulence penetration airspeed in § 25.1585(a)(8), must be es-tablished, which—


oods

2 on

DS

K1D

XX

6B1P

RO

D w

ith C

FR

subpart f—equipment - gpo · ertia, and other loads to which it may be subjected in operation;...

Documents