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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 - Instruments: Installation
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(a) Each flight, navigation, and powerplant instrument for use by any pilot must be easily visible to him from his station with the minimum practicable deviation from his normal position and line of vision when he is looking forward along the flight path.
(b) Each instrument necessary for safe operation, including the airspeed indicator, gyroscopic direction indicator, gyroscopic bank-and-pitch indicator, slip-skid indicator, altimeter, rate-of-climb indicator, rotor tachometers, and the indicator most representative of engine power, must be grouped and centered as nearly as practicable about the vertical plane of the pilot's forward vision. In addition, for rotorcraft approved for IFR flight—
(1) The instrument that most effectively indicates attitude must be on the panel in the top center position;
(2) The instrument that most effectively indicates direction of flight must be adjacent to and directly below the attitude instrument;
(3) The instrument that most effectively indicates airspeed must be adjacent to and to the left of the attitude instrument; and
(4) The instrument that most effectively indicates altitude or is most frequently utilized in control of altitude must be adjacent to and to the right of the attitude instrument.
(c) Other required powerplant instruments must be closely grouped on the instrument panel.
(d) Identical powerplant instruments for the engines must be located so as to prevent any confusion as to which engine each instrument relates.
(e) Each powerplant instrument vital to safe operation must be plainly visible to appropriate crewmembers.
(f) Instrument panel vibration may not damage, or impair the readability or accuracy of, any instrument.
(g) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.
(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. 29-14, 42 FR 36972, July 18, 1977; Amdt. 29-21, 48 FR 4391, Jan. 31, 1983]
If warning, caution or advisory lights are installed in the cockpit they must, unless otherwise approved by the Administrator, be—
(a) Red, for warning lights (lights indicating 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 possible confusion.
[Amdt. 29-12, 41 FR 55474, Dec. 20, 1976]
For each airspeed indicating system, the following apply:
(a) Each airspeed indicating instrument must be calibrated to indicate true airspeed (at sea level with a standard atmosphere) with a minimum practicable instrument calibration error when the corresponding pitot and static pressures are applied.
(b) Each system must be calibrated to determine system error excluding airspeed instrument error. This calibration must be determined—
(1) In level flight at speeds of 20 knots and greater, and over an appropriate range of speeds for flight conditions of climb and autorotation; and
(2) During takeoff, with repeatable and readable indications that ensure—
(i) Consistent realization of the field lengths specified in the Rotorcraft Flight Manual; and
(ii) Avoidance of the critical areas of the height-velocity envelope as established under § 29.87.
(c) For Category A rotorcraft—
(1) The indication must allow consistent definition of the takeoff decision point; and
(2) The system error, excluding the airspeed instrument calibration error, may not exceed—
(i) Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of takeoff safety speed; and
(ii) Ten knots in climb at speeds from 10 knots below takeoff safety speed to 10 knots above VY.
(d) For Category B rotorcraft, the system error, excluding the airspeed instrument calibration error, may not exceed 3 percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of the climbout speed attained at 50 feet when complying with § 29.63.
(e) Each system must be arranged, so far as practicable, to prevent malfunction or serious error due to the entry of moisture, dirt, or other substances.
(f) Each system must have a heated pitot tube or an equivalent means of preventing malfunction due to icing.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-3, 33 FR 970, Jan. 26, 1968; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984; Amdt. 29-39, 61 FR 21901, May 10, 1996; Amdt. 29-44, 64 FR 45338, Aug. 19, 1999]
(a) Each instrument with static air case connections must be vented to the outside atmosphere through an appropriate piping system.
(b) Each vent must be located where its orifices are least affected by airflow variation, moisture, or foreign matter.
(c) Each static pressure port must be designed and located in such manner that the correlation between air pressure in the static pressure system and true ambient atmospheric static pressure is not altered when the rotorcraft encounters icing conditions. An anti-icing means or an alternate source of static pressure may be used in showing compliance with this requirement. If the reading of the altimeter, when on the alternate static pressure system, differs from the reading of altimeter when on the primary static system by more than 50 feet, a correction card must be provided for the alternate static system.
(d) Except for the vent into the atmosphere, each system must be airtight.
(e) Each pressure altimeter must be approved and calibrated to indicate pressure altitude in a standard atmosphere with a minimum practicable calibration error when the corresponding static pressures are applied.
(f) Each system must be designed and installed so that an error in indicated pressure altitude, at sea level, with a standard atmosphere, excluding instrument calibration error, does not result in an error of more than ±30 feet per 100 knots speed. However, the error need not be less than ±30 feet.
(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 rotorcraft, paragraph (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.
(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. 29-14, 42 FR 36972, July 18, 1977; Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]
(a) Each magnetic direction indicator must be installed so that its accuracy is not excessively affected by the rotorcraft's vibration or magnetic fields.
(b) The compensated installation may not have a deviation, in level flight, greater than 10 degrees on any heading.
For the purpose of this subpart, an automatic pilot and flight guidance system may consist of an autopilot, flight director, or a component that interacts with stability augmentation or trim.
(a) Each automatic pilot and flight guidance system must be designed so that it:
(1) Can be overpowered by one pilot to allow control of the rotorcraft;
(2) Provides a means to disengage the system, or any malfunctioning component of the system, by each pilot to prevent it from interfering with the control of the rotorcraft; and
(3) Provides a means to indicate to the flight crew its current mode of operation. Selector switch position is not acceptable as a means of indication.
(b) Unless there is automatic synchronization, each system must have a means to readily indicate to the pilot the alignment of the actuating device in relation to the control system it operates.
(c) Each manually operated control for the system's operation must be readily accessible to the pilots.
(d) The system must be designed so that, within the range of adjustment available to the pilot, it cannot produce hazardous loads on the rotorcraft, or create hazardous deviations in the flight path, under any flight condition appropriate to its use or in the event of a malfunction, assuming that corrective action begins within a reasonable period of time.
(e) If the automatic pilot and flight guidance system integrates signals from auxiliary controls or furnishes signals for operation of other equipment, there must be a means to prevent improper operation.
(f) If the automatic pilot system can be coupled to airborne navigation equipment, means must be provided to indicate to the pilots the current mode of operation. Selector switch position is not acceptable as a means of indication.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR 44439, Nov. 6, 1984; Amdt. 29-24, 49 FR 47594, Dec. 6, 1984; Amdt. 29-42, 63 FR 43285, Aug. 12, 1998; Amdt. 29-59, 88 FR 8739, Feb. 10, 2023]
For category A rotorcraft—
(a) Each required flight instrument using a power supply must have—
(1) Two independent sources of power;
(2) A means of selecting either power source; and
(3) A visual means integral with each instrument to indicate when the power adequate to sustain proper instrument performance is not being supplied. The power must be measured at or near the point where it enters the instrument. For electrical instruments, the power is considered to be adequate when the voltage is within the approved limits; and
(b) The installation and power supply system must be such that failure of any flight instrument connected to one source, or of the energy supply from one source, or a fault in any part of the power distribution system does not interfere with the proper supply of energy from any other source.
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-24, 49 FR 44439, Nov. 6, 1984]
For systems that operate the required flight instruments which are located at each pilot's station, the following apply:
(a) For pneumatic systems, only the required flight instruments for the first pilot may be connected to that operating system.
(b) The equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the flight instruments remains available to a pilot, without additional crewmember action, after any single failure or combination of failures that are not shown to be extremely improbable.
(c) Additional instruments, systems, or equipment may not be connected to the operating system for a second pilot unless provisions are made to ensure the continued normal functioning of the required flight instruments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable.
[Amdt. 29-24, 49 FR 44439, Nov. 6, 1984, as amended by Amdt. 29-59, 88 FR 8740, Feb. 10, 2023]
(a) Instruments and instrument lines.
(1) Each powerplant and auxiliary power unit instrument line must meet the requirements of §§ 29.993 and 29.1183.
(2) Each line carrying flammable fluids under pressure must—
(i) Have restricting orifices or other safety devices at the source of pressure 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 crew members 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 remaining in the tank is equal to the unusable fuel supply determined under § 29.959;
(2) When two or more tanks are closely interconnected by a gravity feed system and vented, and when it is impossible to feed from each tank separately, at least one fuel quantity indicator must be installed;
(3) Tanks with interconnected outlets and airspaces may be treated as one tank and need not have separate indicators; and
(4) Each exposed sight gauge used as a fuel quantity indicator must be protected 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 severely restricts fuel flow.
(d) Oil quantity indicator. There must be a stick gauge or equivalent means to indicate the quantity of oil—
(1) In each tank; and
(2) In each transmission gearbox.
(e) Rotor drive system transmissions and gearboxes utilizing ferromagnetic materials must be equipped with chip detectors designed to indicate the presence of ferromagnetic particles resulting from damage or excessive wear within the transmission or gearbox. Each chip detector must—
(1) Be designed to provide a signal to the indicator required by § 29.1305(a)(22); and
(2) Be provided with a means to allow crewmembers to check, in flight, the function of each detector electrical circuit and signal.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655(c))
[Doc. No. 5084, 29 FR 16150, Dec. 3, 1964, as amended by Amdt. 29-13, 42 FR 15047, Mar. 17, 1977; Amdt. 29-26, 53 FR 34219, Sept. 2, 1988]