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Docket No. 5084, 29 FR 16150, Dec. 3, 1964, unless otherwise noted.
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Subscribe to: 14 CFR Part 29 Subpart F - General
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Each item of installed equipment must—
(a) Be of a kind and design appropriate to its intended function;
(b) Be labeled as to its identification, function, or operating limitations, or any applicable combination of these factors;
(c) Be installed according to limitations specified for that equipment; and
(d) Function properly when installed.
The following are required flight and navigational instruments:
(a) An airspeed indicator. For Category A rotorcraft with VNE less than a speed at which unmistakable pilot cues provide overspeed warning, a maximum allowable airspeed indicator must be provided. If maximum allowable airspeed varies with weight, altitude, temperature, or r.p.m., the indicator must show that variation.
(b) A sensitive altimeter.
(c) A magnetic direction indicator.
(d) A clock displaying hours, minutes, and seconds with a sweep-second pointer or digital presentation.
(e) A free-air temperature indicator.
(f) A non-tumbling gyroscopic bank and pitch indicator.
(g) A gyroscopic rate-of-turn indicator combined with an integral slip-skid indicator (turn-and-bank indicator) except that only a slip-skid indicator is required on rotorcraft with a third attitude instrument system that—
(1) Is usable through flight attitudes of ±80 degrees of pitch and ±120 degrees of roll;
(2) Is powered from a source independent of the electrical generating system;
(3) Continues reliable operation for a minimum of 30 minutes after total failure of the electrical generating system;
(4) Operates independently of any other attitude indicating system;
(5) Is operative without selection after total failure of the electrical generating system;
(6) Is located on the instrument panel in a position acceptable to the Administrator that will make it plainly visible to and useable by any pilot at his station; and
(7) Is appropriately lighted during all phases of operation.
(h) A gyroscopic direction indicator.
(i) A rate-of-climb (vertical speed) indicator.
(j) For Category A rotorcraft, a speed warning device when VNE is less than the speed at which unmistakable overspeed warning is provided by other pilot cues. The speed warning device must give effective aural warning (differing distinctively from aural warnings used for other purposes) to the pilots whenever the indicated speed exceeds VNE plus 3 knots and must operate satisfactorily throughout the approved range of altitudes and temperatures.
(Secs. 313(a), 601, 603, 604, and 605 of the Federal Aviation Act of 1958 (49 U.S.C. 1354(a), 1421, 1423, 1424, and 1425); and sec. 6(c), Dept. of Transportation Act (49 U.S.C. 1655(c)))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-12, 41 FR 55474, Dec. 20, 1976; Amdt. 29-14, 42 FR 36972, July 18, 1977; Amdt. 29-24, 49 FR 44438, Nov. 6, 1984; 70 FR 2012, Jan. 12, 2005]
The following are required powerplant instruments:
(a) For each rotorcraft—
(1) A carburetor air temperature indicator for each reciprocating engine;
(2) A cylinder head temperature indicator for each air-cooled reciprocating engine, and a coolant temperature indicator for each liquid-cooled reciprocating engine;
(3) A fuel quantity indicator for each fuel tank;
(4) A low fuel warning device for each fuel tank which feeds an engine. This device must—
(i) Provide a warning to the crew when approximately 10 minutes of usable fuel remains in the tank; and
(ii) Be independent of the normal fuel quantity indicating system.
(5) A means to indicate manifold pressure for each altitude engine;
(6) An oil pressure indicator for each pressure-lubricated gearbox.
(7) An oil pressure warning device for each pressure-lubricated gearbox to indicate when the oil pressure falls below a safe value;
(8) An oil quantity indicator for each oil tank and each rotor drive gearbox, if lubricant is self-contained;
(9) An oil temperature indicator for each engine;
(10) An oil temperature warning device to indicate unsafe oil temperatures in each main rotor drive gearbox, including gearboxes necessary for rotor phasing;
(11) A means to indicate the gas temperature for each turbine engine;
(12) A means to indicate the gas producer speed for each turbine engine;
(13) A tachometer for each engine that, if combined with the applicable instrument required by paragraph (a)(14) of this section, indicates rotor r.p.m. during autorotation.
(14) At least one tachometer to indicate, as applicable—
(i) The r.p.m. of the single main rotor;
(ii) The common r.p.m. of any main rotors whose speeds cannot vary appreciably with respect to each other; and
(iii) The r.p.m. of each main rotor whose speed can vary appreciably with respect to that of another main rotor;
(15) A free power turbine tachometer for each turbine engine;
(16) A means, for each turbine engine, to indicate power for that engine;
(17) For each turbine engine, an indicator to indicate the functioning of the powerplant ice protection system;
(18) An indicator for the filter required by § 29.997 to indicate the occurrence of contamination of the filter to the degree established in compliance with § 29.955;
(19) For each turbine engine, a warning means for the oil strainer or filter required by § 29.1019, if it has no bypass, to warn the pilot of the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with § 29.1019(a)(2);
(20) An indicator to indicate the functioning of any selectable or controllable heater used to prevent ice clogging of fuel system components;
(21) An individual fuel pressure indicator for each engine, unless the fuel system which supplies that engine does not employ any pumps, filters, or other components subject to degradation or failure which may adversely affect fuel pressure at the engine;
(22) A means to indicate to the flightcrew the failure of any fuel pump installed to show compliance with § 29.955;
(23) Warning or caution devices to signal to the flightcrew when ferromagnetic particles are detected by the chip detector required by § 29.1337(e); and
(24) For auxiliary power units, an individual indicator, warning or caution device, or other means to advise the flightcrew that limits are being exceeded, if exceeding these limits can be hazardous, for—
(i) Gas temperature;
(ii) Oil pressure; and
(iii) Rotor speed.
(25) For rotorcraft for which a 30-second/2-minute OEI power rating is requested, a means must be provided to alert the pilot when the engine is at the 30-second and 2-minute OEI power levels, when the event begins, and when the time interval expires.
(26) For each turbine engine utilizing 30-second/2-minute OEI power, a device or system must be provided for use by ground personnel which—
(i) Automatically records each usage and duration of power at the 30-second and 2-minute OEI levels;
(ii) Permits retrieval of the recorded data;
(iii) Can be reset only by ground maintenance personnel; and
(iv) Has a means to verify proper operation of the system or device.
(b) For category A rotorcraft—
(1) An individual oil pressure indicator for each engine, and either an independent warning device for each engine or a master warning device for the engines with means for isolating the individual warning circuit from the master warning device;
(2) An independent fuel pressure warning device for each engine or a master warning device for all engines with provision for isolating the individual warning device from the master warning device; and
(3) Fire warning indicators.
(4) For each Category A rotorcraft for which OEI Training Mode is requested, a means must be provided to indicate to the pilot the simulation of an engine failure, the annunciation of that simulation, and a representation of the OEI power being provided.
(c) For category B rotorcraft—
(1) An individual oil pressure indicator for each engine; and
(2) Fire warning indicators, when fire detection is required.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR 970, Jan. 26, 1968; Amdt. 29-10, 39 FR 35463, Oct. 1, 1974; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988; Amdt. 29-34, 59 FR 47768, Sept. 16, 1994; Amdt. 29-40, 61 FR 21908, May 10, 1996; 61 FR 43952, Aug. 27, 1996; Amdt. 29-59, 88 FR 8739, Feb. 10, 2023]
The following is required miscellaneous equipment:
(a) An approved seat for each occupant.
(b) A master switch arrangement for electrical circuits other than ignition.
(c) Hand fire extinguishers.
(d) A windshield wiper or equivalent device for each pilot station.
(e) A two-way radio communication system.
[Amdt. 29-12, 41 FR 55473, Dec. 20, 1976]
The equipment, systems, and installations whose functioning is required by this subchapter must be designed and installed to ensure that they perform their intended functions under any foreseeable operating condition. For any item of equipment or system whose failure has not been specifically addressed by another requirement in this chapter, the following requirements also apply:
(a) The design of each item of equipment, system, and installation must be analyzed separately and in relation to other rotorcraft systems and installations to determine and identify any failure that would affect the capability of the rotorcraft or the ability of the crew to perform their duties in all operating conditions.
(b) Each item of equipment, system, and installation must be designed and installed so that:
(1) The occurrence of any catastrophic failure condition is extremely improbable;
(2) The occurrence of any major failure condition is no more than improbable; and
(3) For the occurrence of any other failure condition in between major and catastrophic, the probability of the failure condition must be inversely proportional to its consequences.
(c) A means to alert the crew in the event of a failure must be provided when an unsafe system operating condition exists and to enable them to take corrective action. Systems, controls, and associated monitoring and crew alerting means must be designed to minimize crew errors that could create additional hazards.
(d) Compliance with the requirements of this section must be shown by analysis and, where necessary, by ground, flight, or simulator tests. The analysis must account for:
(1) Possible modes of failure, including malfunctions and misleading data and input from external sources;
(2) The effect of multiple failures and latent failures;
(3) The resulting effects on the rotorcraft and occupants, considering the stage of flight and operating conditions; and
(4) The crew alerting cues and the corrective action required.
[Amdt. 29-59, 88 FR 8739, Feb. 10, 2023]
(a) Each electrical and electronic system that performs a function, for which failure would prevent the continued safe flight and landing of the rotorcraft, must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to lightning; and
(2) The system automatically recovers normal operation of that function in a timely manner after the rotorcraft is exposed to lightning.
(b) Each electrical and electronic system that performs a function, for which failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition, must be designed and installed so that the function recovers normal operation in a timely manner after the rotorcraft is exposed to lightning.
[Doc. No. FAA-2010-0224, Amdt. 29-53, 76 FR 33136, June 8, 2011]
(a) Except as provided in paragraph (d) of this section, each electrical and electronic system that performs a function whose failure would prevent the continued safe flight and landing of the rotorcraft must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment I, as described in appendix E to this part;
(2) The system automatically recovers normal operation of that function, in a timely manner, after the rotorcraft is exposed to HIRF environment I, as described in appendix E to this part, unless this conflicts with other operational or functional requirements of that system;
(3) The system is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment II, as described in appendix E to this part; and
(4) Each function required during operation under visual flight rules is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment III, as described in appendix E to this part.
(b) Each electrical and electronic system that performs a function whose failure would significantly reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 1 or 2, as described in appendix E to this part.
(c) Each electrical and electronic system that performs such a function whose failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 3, as described in appendix E to this part.
(d) Before December 1, 2012, an electrical or electronic system that performs a function whose failure would prevent the continued safe flight and landing of a rotorcraft may be designed and installed without meeting the provisions of paragraph (a) provided—
(1) The system has previously been shown to comply with special conditions for HIRF, prescribed under § 21.16, issued before December 1, 2007;
(2) The HIRF immunity characteristics of the system have not changed since compliance with the special conditions was demonstrated; and
(3) The data used to demonstrate compliance with the special conditions is provided.
[Doc. No. FAA-2006-23657, 72 FR 44027, Aug. 6, 2007]