14 C.F.R. Subpart F—Equipment
Title 14 - Aeronautics and Space
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. [Amdt. 23–20, 42 FR 36968, July 18, 1977] The following are the minimum required flight and navigation instruments: (a) An airspeed indicator. (b) An altimeter. (c) A direction indicator (nonstabilized magnetic compass). (d) For reciprocating engine-powered airplanes of more than 6,000 pounds maximum weight and turbine engine powered airplanes, a free air temperature indicator or an air-temperature indicator which provides indications that are convertible to free-air. (e) A speed warning device for— (1) Turbine engine powered airplanes; and (2) Other airplanes for which The speed warning device must give effective aural warning (differing distinctively from aural warnings used for other purposes) to the pilots whenever the speed exceeds (f) When an attitude display is installed, the instrument design must not provide any means, accessible to the flightcrew, of adjusting the relative positions of the attitude reference symbol and the horizon line beyond that necessary for parallax correction. (g) In addition, for commuter category airplanes: (1) If airspeed limitations vary with altitude, the airspeed indicator must have a maximum allowable airspeed indicator showing the variation of VMO with altitude. (2) The altimeter must be a sensitive type. (3) Having a passenger seating configuration of 10 or more, excluding the pilot's seats and that are approved for IFR operations, a third attitude instrument must be provided that: (i) Is powered from a source independent of the electrical generating system; (ii) Continues reliable operation for a minimum of 30 minutes after total failure of the electrical generating system; (iii) Operates independently of any other attitude indicating system; (iv) Is operative without selection after total failure of the electrical generating system; (v) Is located on the instrument panel in a position acceptable to the Administrator that will make it plainly visible to and usable by any pilot at the pilot's station; and (vi) Is appropriately lighted during all phases of operation. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–43, 58 FR 18975, Apr. 9, 1993; Amdt. 23–49, 61 FR 5168, Feb. 9, 1996] The following are required powerplant instruments: (a) For all airplanes. (1) A fuel quantity indicator for each fuel tank, installed in accordance with §23.1337(b). (2) An oil pressure indicator for each engine. (3) An oil temperature indicator for each engine. (4) An oil quantity measuring device for each oil tank which meets the requirements of §23.1337(d). (5) A fire warning means for those airplanes required to comply with §23.1203. (b) For reciprocating engine-powered airplanes. In addition to the powerplant instruments required by paragraph (a) of this section, the following powerplant instruments are required: (1) An induction system air temperature indicator for each engine equipped with a preheater and having induction air temperature limitations that can be exceeded with preheat. (2) A tachometer indicator for each engine. (3) A cylinder head temperature indicator for— (i) Each air-cooled engine with cowl flaps; (ii) [Reserved] (iii) Each commuter category airplane. (4) For each pump-fed engine, a means: (i) That continuously indicates, to the pilot, the fuel pressure or fuel flow; or (ii) That continuously monitors the fuel system and warns the pilot of any fuel flow trend that could lead to engine failure. (5) A manifold pressure indicator for each altitude engine and for each engine with a controllable propeller. (6) For each turbocharger installation: (i) If limitations are established for either carburetor (or manifold) air inlet temperature or exhaust gas or turbocharger turbine inlet temperature, indicators must be furnished for each temperature for which the limitation is established unless it is shown that the limitation will not be exceeded in all intended operations. (ii) If its oil system is separate from the engine oil system, oil pressure and oil temperature indicators must be provided. (7) A coolant temperature indicator for each liquid-cooled engine. (c) For turbine engine-powered airplanes. In addition to the powerplant instruments required by paragraph (a) of this section, the following powerplant instruments are required: (1) A gas temperature indicator for each engine. (2) A fuel flowmeter indicator for each engine. (3) A fuel low pressure warning means for each engine. (4) A fuel low level warning means for any fuel tank that should not be depleted of fuel in normal operations. (5) A tachometer indicator (to indicate the speed of the rotors with established limiting speeds) for each engine. (6) An oil low pressure warning means for each engine. (7) An indicating means to indicate the functioning of the powerplant ice protection system for each engine. (8) For each engine, an indicating means for the fuel strainer or filter required by §23.997 to indicate the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with §23.997(d). (9) For each engine, a warning means for the oil strainer or filter required by §23.1019, 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 accordance with §23.1019(a)(5). (10) An indicating means to indicate the functioning of any heater used to prevent ice clogging of fuel system components. (d) For turbojet/turbofan engine-powered airplanes. In addition to the powerplant instruments required by paragraphs (a) and (c) of this section, the following powerplant instruments are required: (1) For each engine, an indicator to indicate thrust or to indicate a parameter that can be related to thrust, including a free air temperature indicator if needed for this purpose. (2) For each engine, a position indicating means to indicate to the flight crew when the thrust reverser, if installed, is in the reverse thrust position. (e) For turbopropeller-powered airplanes. 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 engine. (2) A position indicating means to indicate to the flight crew when the propeller blade angle is below the flight low pitch position, for each propeller, unless it can be shown that such occurrence is highly improbable. [Doc. No. 26344, 58 FR 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993; Amdt. 23–51, 61 FR 5138, Feb. 9, 1996; Amdt. 23–52, 61 FR 13644, Mar. 27, 1996] The equipment necessary for an airplane to operate at the maximum operating altitude and in the kinds of operation and meteorological conditions for which certification is requested and is approved in accordance with §23.1559 must be included in the type design. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–23, 43 FR 50593, Oct. 30, 1978; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; Amdt. 23–49, 61 FR 5168, Feb. 9, 1996] (a) Each item of equipment, each system, and each installation: (1) When performing its intended function, may not adversely affect the response, operation, or accuracy of any— (i) Equipment essential to safe operation; or (ii) Other equipment unless there is a means to inform the pilot of the effect. (2) In a single-engine airplane, must be designed to minimize hazards to the airplane in the event of a probable malfunction or failure. (3) In a multiengine airplane, must be designed to prevent hazards to the airplane in the event of a probable malfunction or failure. (4) In a commuter category airplane, must be designed to safeguard against hazards to the airplane in the event of their malfunction or failure. (b) The design of each item of equipment, each system, and each installation must be examined separately and in relationship to other airplane systems and installations to determine if the airplane is dependent upon its function for continued safe flight and landing and, for airplanes not limited to VFR conditions, if failure of a system would significantly reduce the capability of the airplane or the ability of the crew to cope with adverse operating conditions. Each item of equipment, each system, and each installation identified by this examination as one upon which the airplane is dependent for proper functioning to ensure continued safe flight and landing, or whose failure would significantly reduce the capability of the airplane or the ability of the crew to cope with adverse operating conditions, must be designed to comply with the following additional requirements: (1) It must perform its intended function under any foreseeable operating condition. (2) When systems and associated components are considered separately and in relation to other systems— (i) The occurrence of any failure condition that would prevent the continued safe flight and landing of the airplane must be extremely improbable; and (ii) The occurrence of any other failure condition that would significantly reduce the capability of the airplane or the ability of the crew to cope with adverse operating conditions must be improbable. (3) Warning information must be provided to alert the crew to unsafe system operating conditions and to enable them to take appropriate corrective action. Systems, controls, and associated monitoring and warning means must be designed to minimize crew errors that could create additional hazards. (4) Compliance with the requirements of paragraph (b)(2) of this section may be shown by analysis and, where necessary, by appropriate ground, flight, or simulator tests. The analysis must consider— (i) Possible modes of failure, including malfunctions and damage from external sources; (ii) The probability of multiple failures, and the probability of undetected faults.; (iii) The resulting effects on the airplane and occupants, considering the stage of flight and operating conditions; and (iv) The crew warning cues, corrective action required, and the crew's capability of determining faults. (c) Each item of equipment, each system, and each installation whose functioning is required by this chapter and that requires a power supply is an “essential load” on the power supply. The power sources and the system must be able to supply the following power loads in probable operating combinations and for probable durations: (1) Loads connected to the power distribution system with the system functioning normally. (2) Essential loads after failure of— (i) Any one engine on two-engine airplanes; or (ii) Any two engines on an airplane with three or more engines; or (iii) Any power converter or energy storage device. (3) Essential loads for which an alternate source of power is required, as applicable, by the operating rules of this chapter, after any failure or malfunction in any one power supply system, distribution system, or other utilization system. (d) In determining compliance with paragraph (c)(2) of this section, the power loads may be assumed to be reduced under a monitoring procedure consistent with safety in the kinds of operations authorized. Loads not required in controlled flight need not be considered for the two-engine-inoperative condition on airplanes with three or more engines. (e) In showing compliance with this section with regard to the electrical power system and to equipment design and installation, critical environmental and atmospheric conditions, including radio frequency energy and the effects (both direct and indirect) of lightning strikes, must be considered. For electrical generation, distribution, and utilization equipment required by or used in complying with this chapter, the ability to provide continuous, safe service under forseeable environmental conditions may be shown by environmental tests, design analysis, or reference to previous comparable service experience on other airplanes. (f) As used in this section, “system” refers to all pneumatic systems, fluid systems, electrical systems, mechanical systems, and powerplant systems included in the airplane design, except for the following: (1) Powerplant systems provided as part of the certificated engine. (2) The flight structure (such a wing, empennage, control surfaces and their systems, the fuselage, engine mounting, and landing gear and their related primary attachments) whose requirements are specific in subparts C and D of this part. [Amdt. 23–41, 55 FR 43309, Oct. 26, 1990; 55 FR 47028, Nov. 8, 1990, as amended by Amdt. 23–49, 61 FR 5168, Feb. 9, 1996]
Title 14: Aeronautics and Space
PART 23—AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES
Subpart F—Equipment
General
§ 23.1301 Function and installation.
§ 23.1303 Flight and navigation instruments.
§ 23.1305 Powerplant instruments.
§ 23.1307 Miscellaneous equipment.
§ 23.1309 Equipment, systems, and installations.
Instruments: Installation
§ 23.1311 Electronic display instrument systems.
(a) Electronic display indicators, including those with features that make isolation and independence between powerplant instrument systems impractical, must:
(1) Meet the arrangement and visibility requirements of §23.1321.
(2) Be easily legible under all lighting conditions encountered in the cockpit, including direct sunlight, considering the expected electronic display brightness level at the end of an electronic display indictor's useful life. Specific limitations on display system useful life must be contained in the Instructions for Continued Airworthiness required by §23.1529.
(3) Not inhibit the primary display of attitude, airspeed, altitude, or powerplant parameters needed by any pilot to set power within established limitations, in any normal mode of operation.
(4) Not inhibit the primary display of engine parameters needed by any pilot to properly set or monitor powerplant limitations during the engine starting mode of operation.
(5) Have an independent magnetic direction indicator and either an independent secondary mechanical altimeter, airspeed indicator, and attitude instrument or individual electronic display indicators for the altitude, airspeed, and attitude that are independent from the airplane's primary electrical power system. These secondary instruments may be installed in panel positions that are displaced from the primary positions specified by §23.1321(d), but must be located where they meet the pilot's visibility requirements of §23.1321(a).
(6) Incorporate sensory cues for the pilot that are equivalent to those in the instrument being replaced by the electronic display indicators.
(7) Incorporate visual displays of instrument markings, required by §§23.1541 through 23.1553, or visual displays that alert the pilot to abnormal operational values or approaches to established limitation values, for each parameter required to be displayed by this part.
(b) The electronic display indicators, including their systems and installations, and considering other airplane systems, must be designed so that one display of information essential for continued safe flight and landing will remain available to the crew, without need for immediate action by any pilot for continued safe operation, after any single failure or probable combination of failures.
(c) As used in this section, “instrument” includes devices that are physically contained in one unit, and devices that are composed of two or more physically separate units or components connected together (such as a remote indicating gyroscopic direction indicator that includes a magnetic sensing element, a gyroscopic unit, an amplifier, and an indicator connected together). As used in this section, “primary” display refers to the display of a parameter that is located in the instrument panel such that the pilot looks at it first when wanting to view that parameter.
[Doc. No. 27806, 61 FR 5168, Feb. 9, 1996]
§ 23.1321 Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any required pilot during takeoff, initial climb, final approach, and landing must be located so that any pilot seated at the controls can monitor the airplane's flight path and these instruments with minimum head and eye movement. The powerplant instruments for these flight conditions are those needed to set power within powerplant limitations.
(b) For each multiengine airplane, identical powerplant instruments must be located so as to prevent confusion as to which engine each instrument relates.
(c) Instrument panel vibration may not damage, or impair the accuracy of, any instrument.
(d) For each airplane, the flight instruments required by §23.1303, and, as applicable, by the operating rules of this chapter, must be grouped on the instrument panel and centered as nearly as practicable about the vertical plane of each required pilot's forward vision. In addition:
(1) The instrument that most effectively indicates the attitude must be on the panel in the top center position;
(2) The instrument that most effectively indicates airspeed must be adjacent to and directly to the left of the instrument in the top center position;
(3) The instrument that most effectively indicates altitude must be adjacent to and directly to the right of the instrument in the top center position;
(4) The instrument that most effectively indicates direction of flight, other than the magnetic direction indicator required by §23.1303(c), must be adjacent to and directly below the instrument in the top center position; and
(5) Electronic display indicators may be used for compliance with paragraphs (d)(1) through (d)(4) of this section when such displays comply with requirements in §23.1311.
(e) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; Amdt. 23–20, 42 FR 36968, July 18, 1977; Amdt. 23–41, 55 FR 43310, Oct. 26, 1990; 55 FR 46888, Nov. 7, 1990; Amdt. 23–49, 61 FR 5168, Feb. 9, 1996]
§ 23.1322 Warning, caution, and advisory lights.
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.
(e) Effective under all probable cockpit lighting conditions.
[Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23–43, 58 FR 18976, Apr. 9, 1993]
§ 23.1323 Airspeed indicating system.
(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 airspeed system must be calibrated in flight to determine the system error. The system error, including position error, but excluding the airspeed indicator instrument calibration error, may not exceed three percent of the calibrated airspeed or five knots, whichever is greater, throughout the following speed ranges:
(1) 1.3 VS1 to VMO/MMO or VNE, whichever is appropriate with flaps retracted.
(2) 1.3 VS1 to VFE with flaps extended.
(c) The design and installation of each airspeed indicating system must provide positive drainage of moisture from the pitot static plumbing.
(d) If certification for instrument flight rules or flight in icing conditions is requested, each airspeed system must have a heated pitot tube or an equivalent means of preventing malfunction due to icing.
(e) In addition, for commuter category airplanes, the airspeed indicating system must be calibrated to determine the system error during the accelerate-takeoff ground run. The ground run calibration must be obtained between 0.8 of the minimum value of V
(f) For commuter category airplanes, where duplicate airspeed indicators are required, their respective pitot tubes must be far enough apart to avoid damage to both tubes in a collision with a bird.
[Amdt. 23–20, 42 FR 36968, July 18, 1977, as amended by Amdt. 23–34, 52 FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23–42, 56 FR 354, Jan. 3, 1991; Amdt. 23–49, 61 FR 5168, Feb. 9, 1996]
§ 23.1325 Static pressure system.
(a) Each instrument provided with static pressure case connections must be so vented that the influence of airplane speed, the opening and closing of windows, airflow variations, moisture, or other foreign matter will least affect the accuracy of the instruments except as noted in paragraph (b)(3) of this section.
(b) If a static pressure system is necessary for the functioning of instruments, systems, or devices, it must comply with the provisions of paragraphs (b)(1) through (3) of this section.
(1) The design and installation of a static pressure system must be such that—
(i) Positive drainage of moisture is provided;
(ii) Chafing of the tubing, and excessive distortion or restriction at bends in the tubing, is avoided; and
(iii) The materials used are durable, suitable for the purpose intended, and protected against corrosion.
(2) A proof test must be conducted to demonstrate the integrity of the static pressure system in the following manner:
(i) Unpressurized airplanes. Evacuate the static pressure system to a pressure differential of approximately 1 inch of mercury or to a reading on the altimeter, 1,000 feet above the aircraft elevation at the time of the test. Without additional pumping for a period of 1 minute, the loss of indicated altitude must not exceed 100 feet on the altimeter.
(ii) Pressurized airplanes. Evacuate the static pressure system until a pressure differential equivalent to the maximum cabin pressure differential for which the airplane is type certificated is achieved. Without additional pumping 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 pressure or 100 feet, whichever is greater.
(3) If a static pressure system is provided for any instrument, device, or system required by the operating rules of this chapter, each static pressure port must be designed or located in such a manner that the correlation between air pressure in the static pressure system and true ambient atmospheric static pressure is not altered when the airplane encounters icing conditions. An antiicing 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 the altimeter when on the primary static system by more than 50 feet, a correction card must be provided for the alternate static system.
(c) Except as provided in paragraph (d) 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.
(d) For unpressurized airplanes, paragraph (c)(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.
(e) Each static pressure system must be calibrated in flight to determine the system error. The system error, in indicated pressure altitude, at sea-level, with a standard atmosphere, excluding instrument calibration error, may not exceed ±30 feet per 100 knot speed for the appropriate configuration in the speed range between 1.3 VS0 with flaps extended, and 1.8 VS1 with flaps retracted. However, the error need not be less than 30 feet.
(f) [Reserved]
(g) For airplanes prohibited from flight in instrument meteorological or icing conditions, in accordance with §23.1559(b) of this part, paragraph (b)(3) of this section does not apply.
[Amdt. 23–1, 30 FR 8261, June 29, 1965, as amended by Amdt. 23–6, 32 FR 7586, May 24, 1967; 32 FR 13505, Sept. 27, 1967; 32 FR 13714, Sept. 30, 1967; Amdt. 23–20, 42 FR 36968, July 18, 1977; Amdt. 23–34, 52 FR 1834, Jan. 15, 1987; Amdt. 23–42, 56 FR 354, Jan. 3, 1991; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996; Amdt. 23–50, 61 FR 5192, Feb. 9, 1996]
§ 23.1326 Pitot heat indication systems.
If a flight instrument pitot heating system is installed to meet the requirements specified in §23.1323(d), an indication system 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 incorporate an amber light that is in clear view of a flightcrew member.
(b) The indication provided must be designed to alert the flight crew if either 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 heating element is inoperative.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1327 Magnetic direction indicator.
(a) Except as provided in paragraph (b) of this section—
(1) Each magnetic direction indicator must be installed so that its accuracy is not excessively affected by the airplane's vibration or magnetic fields; and
(2) The compensated installation may not have a deviation in level flight, greater than ten degrees on any heading.
(b) A magnetic nonstabilized direction indicator may deviate more than ten degrees due to the operation of electrically powered systems such as electrically heated windshields if either a magnetic stabilized direction indicator, which does not have a deviation in level flight greater than ten degrees on any heading, or a gyroscopic direction indicator, is installed. Deviations of a magnetic nonstabilized direction indicator of more than 10 degrees must be placarded in accordance with §23.1547(e).
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1329 Automatic pilot system.
If an automatic pilot system is installed, it must meet the following:
(a) Each system must be designed so that the automatic pilot can—
(1) Be quickly and positively disengaged by the pilots to prevent it from interfering with their control of the airplane; or
(2) Be sufficiently overpowered by one pilot to let him control the airplane.
(b) If the provisions of paragraph (a)(1) of this section are applied, the quick release (emergency) control must be located on the control wheel (both control wheels if the airplane can be operated from either pilot seat) on the side opposite the throttles, or on the stick control, (both stick controls, if the airplane can be operated from either pilot seat) such that it can be operated without moving the hand from its normal position on the control.
(c) 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.
(d) Each manually operated control for the system operation must be readily accessible to the pilot. Each control must operate in the same plane and sense of motion as specified in §23.779 for cockpit controls. The direction of motion must be plainly indicated on or near each control.
(e) Each system must be designed and adjusted so that, within the range of adjustment available to the pilot, it cannot produce hazardous loads on the airplane or create hazardous deviations in the flight path, under any flight condition appropriate to its use, either during normal operation or in the event of a malfunction, assuming that corrective action begins within a reasonable period of time.
(f) Each system must be designed so that a single malfunction will not produce a hardover signal in more than one control axis. If the automatic pilot integrates signals from auxiliary controls or furnishes signals for operation of other equipment, positive interlocks and sequencing of engagement to prevent improper operation are required.
(g) There must be protection against adverse interaction of integrated components, resulting from a malfunction.
(h) If the automatic pilot system can be coupled to airborne navigation equipment, means must be provided to indicate to the flight crew the current mode of operation. Selector switch position is not acceptable as a means of indication.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–23, 43 FR 50593, Oct. 30, 1978; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
§ 23.1331 Instruments using a power source.
For each instrument that uses a power source, the following apply:
(a) Each instrument must have an integral visual power annunciator or separate power indicator to indicate when power is not adequate to sustain proper instrument performance. If a separate indicator is used, it must be located so that the pilot using the instruments can monitor the indicator with minimum head and eye movement. The power must be sensed at or near the point where it enters the instrument. For electric and vacuum/pressure instruments, the power is considered to be adequate when the voltage or the vacuum/pressure, respectively, is within approved limits.
(b) The installation and power supply systems must be designed so that—
(1) The failure of one instrument will not interfere with the proper supply of energy to the remaining instrument; and
(2) The failure of the energy supply from one source will not interfere with the proper supply of energy from any other source.
(c) There must be at least two independent sources of power (not driven by the same engine on multiengine airplanes), and a manual or an automatic means to select each power source.
[Doc. No. 26344, 58 FR 18976, Apr. 9, 1993]
§ 23.1335 Flight director systems.
If a flight director system is installed, means must be provided to indicate to the flight crew its current mode of operation. Selector switch position is not acceptable as a means of indication.
[Amdt. 23–20, 42 FR 36969, July 18, 1977]
§ 23.1337 Powerplant instruments installation.
(a) Instruments and instrument lines. (1) Each powerplant and auxiliary power unit instrument line must meet the requirements of §23.993.
(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 indication. There must be a means to indicate to the flightcrew members the quantity of usable fuel in each tank during flight. An indicator calibrated in appropriate units and clearly marked to indicate those units must be used. 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 §23.959(a);
(2) Each exposed sight gauge used as a fuel quantity indicator must be protected against damage;
(3) Each sight gauge that forms a trap in which water can collect and freeze must have means to allow drainage on the ground;
(4) There must be a means to indicate the amount of usable fuel in each tank when the airplane is on the ground (such as by a stick gauge);
(5) Tanks with interconnected outlets and airspaces may be considered as one tank and need not have separate indicators; and
(6) No fuel quantity indicator is required for an auxiliary tank that is used only to transfer fuel to other tanks if the relative size of the tank, the rate of fuel transfer, and operating instructions are adequate to—
(i) Guard against overflow; and
(ii) Give the flight crewmembers prompt warning if transfer is not proceeding as planned.
(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each metering component must have a means to by-pass the fuel supply if malfunctioning of that component severely restricts fuel flow.
(d) Oil quantity indicator. There must be a means to indicate the quantity of oil in each tank—
(1) On the ground (such as by a stick gauge); and
(2) In flight, to the flight crew members, if there is an oil transfer system or a reserve oil supply system.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; Amdt. 23–51, 61 FR 5138, Feb. 9, 1996; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
Electrical Systems and Equipment
§ 23.1351 General.
(a) Electrical system capacity. Each electrical system must be adequate for the intended use. In addition—
(1) Electric power sources, their transmission cables, and their associated control and protective devices, must be able to furnish the required power at the proper voltage to each load circuit essential for safe operation; and
(2) Compliance with paragraph (a)(1) of this section must be shown as follows—
(i) For normal, utility, and acrobatic category airplanes, by an electrical load analysis or by electrical measurements that account for the electrical loads applied to the electrical system in probable combinations and for probable durations; and
(ii) For commuter category airplanes, by an electrical load analysis that accounts for the electrical loads applied to the electrical system in probable combinations and for probable durations.
(b) Function. For each electrical system, the following apply:
(1) Each system, when installed, must be—
(i) Free from hazards in itself, in its method of operation, and in its effects on other parts of the airplane;
(ii) Protected from fuel, oil, water, other detrimental substances, and mechanical damage; and
(iii) So designed that the risk of electrical shock to crew, passengers, and ground personnel is reduced to a minimum.
(2) Electric power sources must function properly when connected in combination or independently.
(3) No failure or malfunction of any electric power source may impair the ability of any remaining source to supply load circuits essential for safe operation.
(4) In addition, for commuter category airplanes, the following apply:
(i) Each system must be designed so that essential load circuits can be supplied in the event of reasonably probable faults or open circuits including faults in heavy current carrying cables;
(ii) A means must be accessible in flight to the flight crewmembers for the individual and collective disconnection of the electrical power sources from the system;
(iii) The system must be designed so that voltage and frequency, if applicable, at the terminals of all essential load equipment can be maintained within the limits for which the equipment is designed during any probable operating conditions;
(iv) If two independent sources of electrical power for particular equipment or systems are required, their electrical energy supply must be ensured by means such as duplicate electrical equipment, throwover switching, or multichannel or loop circuits separately routed; and
(v) For the purpose of complying with paragraph (b)(5) of this section, the distribution system includes the distribution busses, their associated feeders, and each control and protective device.
(c) Generating system. There must be at least one generator/alternator if the electrical system supplies power to load circuits essential for safe operation. In addition—
(1) Each generator/alternator must be able to deliver its continuous rated power, or such power as is limited by its regulation system.
(2) Generator/alternator voltage control equipment must be able to dependably regulate the generator/alternator output within rated limits.
(3) Automatic means must be provided to prevent damage to any generator/alternator and adverse effects on the airplane electrical system due to reverse current. A means must also be provided to disconnect each generator/alternator from the battery and other generators/alternators.
(4) There must be a means to give immediate warning to the flight crew of a failure of any generator/alternator.
(5) Each generator/alternator must have an overvoltage control designed and installed to prevent damage to the electrical system, or to equipment supplied by the electrical system that could result if that generator/alternator were to develop an overvoltage condition.
(d) Instruments. A means must exist to indicate to appropriate flight crewmembers the electric power system quantities essential for safe operation.
(1) For normal, utility, and acrobatic category airplanes with direct current systems, an ammeter that can be switched into each generator feeder may be used and, if only one generator exists, the ammeter may be in the battery feeder.
(2) For commuter category airplanes, the essential electric power system quantities include the voltage and current supplied by each generator.
(e) Fire resistance. Electrical equipment must be so designed and installed that in the event of a fire in the engine compartment, during which the surface of the firewall adjacent to the fire is heated to 2,000 °F for 5 minutes or to a lesser temperature substantiated by the applicant, the equipment essential to continued safe operation and located behind the firewall will function satisfactorily and will not create an additional fire hazard.
(f) External power. If provisions are made for connecting external power to the airplane, and that external power can be electrically connected to equipment 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. The external power connection must be located so that its use will not result in a hazard to the airplane or ground personnel.
(g) 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 (electrical power sources excluding the battery and any other standby electrical sources) inoperative, with critical type fuel (from the standpoint of flameout and restart capability), and with the airplane initially at the maximum certificated altitude. Parts of the electrical system may remain on if—
(1) A single malfunction, including a wire bundle or junction box fire, cannot result in loss of the part turned off and the part turned on; and
(2) The parts turned on are electrically and mechanically isolated from the parts turned off.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–34, 52 FR 1834, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
§ 23.1353 Storage battery design and installation.
(a) Each storage battery must be designed and installed as prescribed in this section.
(b) Safe cell temperatures and pressures must be maintained during any probable charging and discharging condition. No uncontrolled increase in cell temperature may result when the battery is recharged (after previous complete discharge)—
(1) At maximum regulated voltage or power;
(2) During a flight of maximum duration; and
(3) Under the most adverse cooling condition likely to occur in service.
(c) Compliance with paragraph (b) of this section must be shown by tests unless experience with similar batteries and installations has shown that maintaining safe cell temperatures and pressures presents no problem.
(d) No explosive or toxic gases emitted by any battery in normal operation, or as the result of any probable malfunction in the charging system or battery installation, may accumulate in hazardous quantities within the airplane.
(e) No corrosive fluids or gases that may escape from the battery may damage surrounding structures or adjacent essential equipment.
(f) Each nickel cadmium battery installation capable of being used to start an engine or auxiliary power unit must have provisions to prevent 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 its individual cells.
(g) Nickel cadmium battery installations capable of being used to start an engine or auxiliary power unit must have—
(1) A system to control the charging rate of the battery automatically so as to prevent battery overheating;
(2) 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 condition; or
(3) A battery failure sensing and warning system with a means for disconnecting the battery from its charging source in the event of battery failure.
(h) In the event of a complete loss of the primary electrical power generating system, the battery must be capable of providing at least 30 minutes of electrical power to those loads that are essential to continued safe flight and landing. The 30 minute time period includes the time needed for the pilots to recognize the loss of generated power and take appropriate load shedding action.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–21, 43 FR 2319, Jan. 16, 1978; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
§ 23.1357 Circuit protective devices.
(a) Protective devices, such as fuses or circuit breakers, must be installed in all electrical circuits other than—
(1) Main circuits of starter motors used during starting only; and
(2) Circuits in which no hazard is presented by their omission.
(b) A protective device for a circuit essential to flight safety may not be used to protect any other circuit.
(c) Each resettable circuit protective device (“trip free” device in which the tripping mechanism cannot be overridden by the operating control) must be designed so that—
(1) A manual operation is required to restore service after tripping; and
(2) If an overload or circuit fault exists, the device will open the circuit regardless 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 so located and identified that it can be readily reset or replaced in flight.
(e) For fuses identified as replaceable in flight—
(1) There must be one spare of each rating or 50 percent spare fuses of each rating, whichever is greater; and
(2) The spare fuse(s) must be readily accessible to any required pilot.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–20, 42 FR 36969, July 18, 1977]; Amdt. 23–43, 58 FR 18976, Apr. 9, 1993
§ 23.1359 Electrical system fire protection.
(a) Each component of the electrical system must meet the applicable fire protection requirements of §§23.863 and 23.1182.
(b) Electrical cables, terminals, and equipment in designated fire zones that are used during emergency procedures must be fire-resistant.
(c) Insulation on electrical wire and electrical cable must be self-extinguishing when tested at an angle of 60 degrees in accordance with the applicable portions of appendix F of this part, or other approved equivalent methods. The average burn length must not exceed 3 inches (76 mm) and the average flame time after removal of the flame source must not exceed 30 seconds. Drippings from the test specimen must not continue to flame for more than an average of 3 seconds after falling.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1361 Master switch arrangement.
(a) There must be a master switch arrangement to allow ready disconnection of each electric power source from power distribution systems, except as provided in paragraph (b) of this section. The point of disconnection must be adjacent to the sources controlled by the switch arrangement. If separate switches are incorporated into the master switch arrangement, a means must be provided for the switch arrangement to be operated by one hand with a single movement.
(b) Load circuits may be connected so that they remain energized when the master switch is open, if the circuits are isolated, or physically shielded, to prevent their igniting flammable fluids or vapors that might be liberated by the leakage or rupture of any flammable fluid system; and
(1) The circuits are required for continued operation of the engine; or
(2) The circuits are protected by circuit protective devices with a rating of five amperes or less adjacent to the electric power source.
(3) In addition, two or more circuits installed in accordance with the requirements of paragraph (b)(2) of this section must not be used to supply a load of more than five amperes.
(c) The master switch or its controls must be so installed that the switch is easily discernible and accessible to a crewmember.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
§ 23.1365 Electric cables and equipment.
(a) Each electric connecting cable must be of adequate capacity.
(b) Any equipment that is associated with any electrical cable installation and that would overheat in the event of circuit overload or fault must be flame resistant. That equipment and the electrical cables must not emit dangerous quantities of toxic fumes.
(c) Main power cables (including generator cables) in the fuselage must be designed to allow a reasonable degree of deformation and stretching without failure and must—
(1) Be separated from flammable fluid lines; or
(2) Be shrouded by means of electrically insulated flexible conduit, or equivalent, which is in addition to the normal cable insulation.
(d) Means of identification must be provided for electrical cables, terminals, and connectors.
(e) Electrical cables must be installed such that the risk of mechanical damage and/or damage cased by fluids vapors, or sources of heat, is minimized.
(f) Where a cable cannot be protected by a circuit protection device or other overload protection, it must not cause a fire hazard under fault conditions.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996]
§ 23.1367 Switches.
Each switch must be—
(a) Able to carry its rated current;
(b) Constructed with enough distance or insulating material between current carrying parts and the housing so that vibration in flight will not cause shorting;
(c) Accessible to appropriate flight crewmembers; and
(d) Labeled as to operation and the circuit controlled.
Lights
§ 23.1381 Instrument lights.
The instrument lights must—
(a) Make each instrument and control easily readable and discernible;
(b) Be installed so that their direct rays, and rays reflected from the windshield or other surface, are shielded from the pilot's eyes; and
(c) Have enough distance or insulating material between current carrying parts and the housing so that vibration in flight will not cause shorting.
A cabin dome light is not an instrument light.
§ 23.1383 Taxi and landing lights.
Each taxi and landing light must be designed and installed so that:
(a) No dangerous glare is visible to the pilots.
(b) The pilot is not seriously affected by halation.
(c) It provides enough light for night operations.
(d) It does not cause a fire hazard in any configuration.
[Doc. No. 27806, 61 FR 5169, Feb. 9, 1996]
§ 23.1385 Position light system installation.
(a) General. Each part of each position light system must meet the applicable requirements of this section and each system as a whole must meet the requirements of §§23.1387 through 23.1397.
(b) Left and right position lights. Left and right position lights must consist of a red and a green light spaced laterally as far apart as practicable and installed on the airplane such that, with the airplane in the normal flying position, the red light is on the left side and the green light is on the right side.
(c) Rear position light. The rear position light must be a white light mounted as far aft as practicable on the tail or on each wing tip.
(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 appreciable light transmission during normal use.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993]
§ 23.1387 Position light system dihedral angles.
(a) Except as provided in paragraph (e) of this section, each position light must, as installed, show unbroken light within the dihedral 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 degrees to the left of the first, as viewed when looking forward along the longitudinal 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 degrees to the right of the first, as viewed when looking forward along the longitudinal 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 longitudinal axis, as viewed when looking aft along the longitudinal axis.
(e) If the rear position light, when mounted as far aft as practicable in accordance with §23.1385(c), cannot show unbroken light within dihedral angle A (as defined in paragraph (d) of this section), a solid angle or angles of obstructed 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. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–12, 36 FR 21278, Nov. 5, 1971; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993]
§ 23.1389 Position light distribution and intensities.
(a) General. The intensities prescribed in this section must be provided by new equipment with each light cover and color filter in place. Intensities must be determined with the light source operating at a steady value equal to the average luminous output of the source at the normal operating voltage of the airplane. The light distribution and intensity of each position light must meet the requirements of paragraph (b) of this section.
(b) Position lights. The light distribution and intensities of position lights must be expressed in terms of minimum intensities in the horizontal plane, minimum intensities in any vertical plane, and maximum intensities in overlapping beams, within dihedral 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 airplane) must equal or exceed the values in §23.1391.
(2) Intensities in any vertical plane. Each intensity in any vertical plane (the plane perpendicular to the horizontal plane) must equal or exceed the appropriate value in §23.1393, where I is the minimum intensity prescribed in §23.1391 for the corresponding angles in the horizontal plane.
(3) Intensities in overlaps between adjacent signals. No intensity in any overlap between adjacent signals may exceed the values in §23.1395, except that higher intensities in overlaps may be used with main beam intensities substantially greater than the minima specified in §§23.1391 and 23.1393, if the overlap intensities in relation to the main beam intensities do not adversely affect signal clarity. When the peak intensity of the left and right position lights is more than 100 candles, the maximum overlap intensities between them may exceed the values in §23.1395 if the overlap intensity in Area A is not more than 10 percent of peak position light intensity and the overlap intensity in Area B is not more than 2.5 percent of peak position light intensity.
(c) Rear position light installation. A single rear position light may be installed in a position displaced laterally from the plane of symmetry of an airplane if—
(1) The axis of the maximum cone of illumination is parallel to the flight path in level flight; and
(2) There is no obstruction aft of the light and between planes 70 degrees to the right and left of the axis of maximum illumination.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18977, Apr. 9, 1993]
§ 23.1391 Minimum intensities in the horizontal plane of position lights.
Each position light intensity must equal or exceed the applicable values in the following table:
------------------------------------------------------------------------ Angle from right or left of longitudinal Intensity Dihedral angle (light included) axis, measured from (candles) dead ahead------------------------------------------------------------------------L and R (red and green)............. 0° to 10°..... 40 10° to 20°.... 30 20° to 110°... 5A (rear white)...................... 110° to 180°.. 20------------------------------------------------------------------------
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[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18977, Apr. 9, 1993] § 23.1393 Minimum intensities in any vertical plane of position lights.Each position light intensity must equal or exceed the applicable values in the following table: |
------------------------------------------------------------------------ Intensity, Angle above or below the horizontal plane l------------------------------------------------------------------------0°...................................................... 1.000° to 5°............................................ 0.905° to 10°........................................... 0.8010° to 15°.......................................... 0.7015° to 20°.......................................... 0.5020° to 30°.......................................... 0.3030° to 40°.......................................... 0.1040° to 90°.......................................... 0.05------------------------------------------------------------------------
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[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18977, Apr. 9, 1993] § 23.1395 Maximum intensities in overlapping beams of position lights.No position light intensity may exceed the applicable values in the following equal or exceed the applicable values in §23.1389(b)(3): |
------------------------------------------------------------------------ Maximum intensity ----------------------- Overlaps Area A Area B (candles) (candles)------------------------------------------------------------------------Green in dihedral angle L....................... 10 1Red in dihedral angle R......................... 10 1Green in dihedral angle A....................... 5 1Red in dihedral angle A......................... 5 1Rear white in dihedral angle L.................. 5 1Rear white in dihedral angle R.................. 5 1------------------------------------------------------------------------
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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 degrees; 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. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18977, Apr. 9, 1993] § 23.1397 Color specifications.Each position light color must have the applicable International Commission on Illumination chromaticity coordinates 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, whichever 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. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, amended by Amdt. 23–11, 36 FR 12971, July 10, 1971] § 23.1399 Riding light.(a) Each riding (anchor) light required for a seaplane or amphibian, must be installed so that it can— (1) Show a white light for at least two miles at night under clear atmospheric 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. § 23.1401 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 flight crewmembers' vision or detract from the conspicuity of the position lights; and (2) Meets the requirements of paragraphs (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 airplane. The field of coverage must extend in each direction within at least 75 degrees above and 75 degrees below the horizontal plane of the airplane, except that there may be solid angles of obstructed visibility totaling not more than 0.5 steradians. (c) Flashing characteristics. The arrangement of the system, that is, the number of light sources, beam width, speed of rotation, and other characteristics, must give an effective flash frequency of not less than 40, nor more than 100, cycles per minute. The effective flash frequency is the frequency at which the airplane's complete anticollision 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 requirements of §23.1397. (e) Light intensity. The minimum light intensities in any vertical plane, measured with the red filter (if used) and expressed in terms of “effective” intensities, must meet the requirements of paragraph (f) of this section. The following relation must be assumed:
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 effective intensity is obtained when t2 and t1 are chosen so that the effective intensity is equal to the instantaneous intensity 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 following table. |
------------------------------------------------------------------------ Effective Angle above or below the horizontal plane intensity (candles)------------------------------------------------------------------------0° to 5°........................................... 4005° to 10°.......................................... 24010° to 20°......................................... 8020° to 30°......................................... 4030° to 75°......................................... 20------------------------------------------------------------------------
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[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–11, 36 FR 12972, July 10, 1971; Amdt. 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–49, 61 FR 5169, Feb. 9, 1996] Safety Equipment§ 23.1411 General.(a) Required safety equipment to be used by the flight crew in an emergency, such as automatic liferaft releases, must be readily accessible. (b) Stowage provisions for required safety equipment must be furnished and must— (1) Be arranged so that the equipment is directly accessible and its location is obvious; and (2) Protect the safety equipment from damage caused by being subjected to the inertia loads resulting from the ultimate static load factors specified in §23.561(b)(3) of this part. [Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23–36, 53 FR 30815, Aug. 15, 1988] § 23.1415 Ditching equipment.(a) Emergency flotation and signaling equipment required by any operating rule in this chapter must be installed so that it is readily available to the crew and passengers. (b) Each raft and each life preserver must be approved. (c) Each raft released automatically or by the pilot must be attached to the airplane by a line to keep it alongside the airplane. This line must be weak enough to break before submerging the empty raft to which it is attached. (d) Each signaling device required by any operating rule in this chapter, must be accessible, function satisfactorily, and must be free of any hazard in its operation. § 23.1416 Pneumatic de-icer boot system.If certification with ice protection provisions is desired and a pneumatic de-icer boot system is installed— (a) The system must meet the requirements specified in §23.1419. (b) The system and its components must be designed to perform their intended function under any normal system operating temperature or pressure, and (c) Means to indicate to the flight crew that the pneumatic de-icer boot system is receiving adequate pressure and is functioning normally must be provided. [Amdt. 23–23, 43 FR 50593, Oct. 30, 1978] § 23.1419 Ice protection.If certification with ice protection provisions is desired, compliance with the requirements of this section and other applicable sections of this part must be shown: (a) An analysis must be performed to establish, on the basis of the airplane's operational needs, the adequacy of the ice protection system for the various components of the airplane. In addition, tests of the ice protection system must be conducted to demonstrate that the airplane is capable of operating safely in continuous maximum and intermittent maximum icing conditions, as described in appendix C of part 25 of this chapter. As used in this section, “Capable of operating safely,” means that airplane performance, controllability, maneuverability, and stability must not be less than that required in part 23, subpart B. (b) Except as provided by paragraph (c) of this section, in addition to the analysis and physical evaluation prescribed in paragraph (a) of this section, the effectiveness of the ice protection system and its components must be shown by flight tests of the airplane or its components in measured natural atmospheric icing conditions and by one or more of the following tests, as found necessary to determine the adequacy of the ice protection system— (1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components. (2) Flight dry air tests of the ice protection system as a whole, or its individual components. (3) Flight test of the airplane or its components in measured simulated icing conditions. (c) If certification with ice protection has been accomplished on prior type certificated airplanes whose designs include components that are thermodynamically and aerodynamically equivalent to those used on a new airplane design, certification of these equivalent components may be accomplished by reference to previously accomplished tests, required in §23.1419 (a) and (b), provided that the applicant accounts for any differences in installation of these components. (d) A means must be identified or provided for determining the formation of ice on the critical parts of the airplane. Adequate lighting must be provided for the use of this means during night operation. Also, when monitoring of the external surfaces of the airplane by the flight crew is required for operation of the ice protection equipment, external lighting must be provided that is adequate to enable the monitoring to be done at night. Any illumination 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. The Airplane Flight Manual or other approved manual material must describe the means of determining ice formation and must contain information for the safe operation of the airplane in icing conditions. [Doc. No. 26344, 58 FR 18977, Apr. 9, 1993] Miscellaneous Equipment§ 23.1431 Electronic equipment.(a) In showing compliance with §23.1309(b)(1) and (2) with respect to radio and electronic equipment and their installations, critical environmental conditions must be considered. (b) Radio and electronic equipment, controls, and wiring must be installed so that operation of any unit or system of units will not adversely affect the simultaneous operation of any other radio or electronic unit, or system of units, required by this chapter. (c) For those airplanes required to have more than one flightcrew member, or whose operation will require more than one flightcrew member, the cockpit must be evaluated to determine if the flightcrew members, when seated at their duty station, can converse without difficulty under the actual cockpit noise conditions when the airplane is being operated. If the airplane design includes provision for the use of communication headsets, the evaluation must also consider conditions where headsets are being used. If the evaluation shows conditions under which it will be difficult to converse, an intercommunication system must be provided. (d) If installed communication equipment includes transmitter “off-on” switching, that switching means must be designed to return from the “transmit” to the “off” position when it is released and ensure that the transmitter will return to the off (non transmitting) state. (e) If provisions for the use of communication headsets are provided, it must be demonstrated that the flightcrew members will receive all aural warnings under the actual cockpit noise conditions when the airplane is being operated when any headset is being used. [Doc. No. 26344, 58 FR 18977, Apr. 9, 1993, as amended by Amdt. 23–49, 61 FR 5169, Feb. 9, 1996] § 23.1435 Hydraulic systems.(a) Design. Each hydraulic system must be designed as follows: (1) Each hydraulic system and its elements must withstand, without yielding, the structural loads expected in addition to hydraulic loads. (2) A means to indicate the pressure in each hydraulic system which supplies two or more primary functions must be provided to the flight crew. (3) There must be means to ensure that the pressure, including transient (surge) pressure, in any part of the system will not exceed the safe limit above design operating pressure and to prevent excessive pressure resulting from fluid volumetric changes in all lines which are likely to remain closed long enough for such changes to occur. (4) The minimum design burst pressure must be 2.5 times the operating pressure. (b) Tests. Each system must be substantiated by proof pressure tests. When proof tested, no part of any system may fail, malfunction, or experience a permanent set. The proof load of each system must be at least 1.5 times the maximum operating pressure of that system. (c) Accumulators. A hydraulic accumulator or reservoir may be installed on the engine side of any firewall if— (1) It is an integral part of an engine or propeller system, or (2) The reservoir is nonpressurized and the total capacity of all such nonpressurized reservoirs is one quart or less. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–14, 38 FR 31824, Nov. 19, 1973; Amdt. 23–43, 58 FR 18977, Apr. 9, 1993; Amdt. 23–49, 61 FR 5170, Feb. 9, 1996] § 23.1437 Accessories for multiengine airplanes.For multiengine airplanes, engine-driven accessories essential to safe operation must be distributed among two or more engines so that the failure of any one engine will not impair safe operation through the malfunctioning of these accessories. § 23.1438 Pressurization and pneumatic 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 pressure. (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 Administrator finds it equivalent to the required test. [Amdt. 23–20, 42 FR 36969, July 18, 1977] § 23.1441 Oxygen equipment and supply.(a) If certification with supplemental oxygen equipment is requested, or the airplane is approved for operations at or above altitudes where oxygen is required to be used by the operating rules, oxygen equipment must be provided that meets the requirements of this section and §§23.1443 through 23.1449. Portable oxygen equipment may be used to meet the requirements of this part if the portable equipment is shown to comply with the applicable requirements, is identified in the airplane type design, and its stowage provisions are found to be in compliance with the requirements of §23.561. (b) The oxygen system must be free from hazards in itself, in its method of operation, and its effect upon other components. (c) There must be a means to allow the crew to readily determine, during the flight, the quantity of oxygen available in each source of supply. (d) Each required flight crewmember must be provided with— (1) Demand oxygen equipment if the airplane is to be certificated for operation above 25,000 feet. (2) Pressure demand oxygen equipment if the airplane is to be certificated for operation above 40,000 feet. (e) There must be a means, readily available to the crew in flight, to turn on and to shut off the oxygen supply at the high pressure source. This shutoff requirement does not apply to chemical oxygen generators. [Amdt. 23–9, 35 FR 6386, Apr. 21, 1970, as amended by Amdt. 23–43, 58 FR 18978, Apr. 9, 1993] § 23.1443 Minimum mass flow of supplemental oxygen.(a) If continuous flow oxygen equipment is installed, an applicant must show compliance with the requirements of either paragraphs (a)(1) and (a)(2) or paragraph (a)(3) of this section: (1) For each passenger, the minimum mass flow of supplemental oxygen required at various cabin pressure altitudes may not be less than the flow required to maintain, during inspiration and while using the oxygen equipment (including masks) provided, the following mean tracheal oxygen partial pressures: (i) 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, Body Temperature, Pressure, Saturated (BTPS) and with a tidal volume of 700 cc. with a constant time interval between respirations. (ii) 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 breathing 30 liters per minute, BTPS, and with a tidal volume of 1,100 cc. with a constant time interval between respirations. (2) For each flight crewmember, the minimum mass flow may not be less than the flow required to maintain, during inspiration, a mean tracheal oxygen partial 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. (3) The minimum mass flow of supplemental oxygen supplied for each user must be at a rate not less than that shown in the following figure for each altitude up to and including the maximum operating altitude of the airplane. 
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(b) If demand equipment is installed for use by flight crewmembers, the minimum mass flow of supplemental oxygen required for each flight 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 undiluted oxygen at their discretion. (c) If first-aid oxygen equipment is installed, the minimum mass flow of oxygen to each user may not be less than 4 liters per minute, STPD. However, there may be a means to decrease this flow to not less than 2 liters per minute, STPD, at any cabin altitude. The quantity of oxygen required is based upon an average flow rate of 3 liters per minute per person for whom first-aid oxygen is required. (d) As used in this section: (1) BTPS means Body Temperature, and Pressure, Saturated (which is, 37 °C, and the ambient pressure to which the body is exposed, minus 47 mm. Hg, which is the tracheal pressure displaced by water vapor pressure when the breathed air becomes saturated with water vapor at 37 °C). (2) STPD means Standard, Temperature, and Pressure, Dry (which is, 0 °C at 760 mm. Hg with no water vapor). [Doc. No. 26344, 58 FR 18978, Apr. 9, 1993] § 23.1445 Oxygen distribution system.(a) Except for flexible lines from oxygen outlets to the dispensing units, or where shown to be otherwise suitable to the installation, nonmetallic tubing must not be used for any oxygen line that is normally pressurized during flight. (b) Nonmetallic oxygen distribution lines must not be routed where they may be subjected to elevated temperatures, electrical arcing, and released flammable fluids that might result from any probable failure. [Doc. No. 26344, 58 FR 18978, Apr. 9, 1993] § 23.1447 Equipment standards for oxygen dispensing units.If oxygen dispensing units are installed, the following apply: (a) There must be an individual dispensing unit for each occupant for whom supplemental oxygen is to be supplied. Each dispensing unit must: (1) Provide for effective utilization of the oxygen being delivered to the unit. (2) Be capable of being readily placed into position on the face of the user. (3) Be equipped with a suitable means to retain the unit in position on the face. (4) If radio equipment is installed, the flightcrew oxygen dispensing units must be designed to allow the use of that equipment and to allow communication with any other required crew member while at their assigned duty station. (b) If certification for operation up to and including 18,000 feet (MSL) is requested, each oxygen dispensing unit must: (1) Cover the nose and mouth of the user; or (2) Be a nasal cannula, in which case one oxygen dispensing unit covering both the nose and mouth of the user must be available. In addition, each nasal cannula or its connecting tubing must have permanently affixed— (i) A visible warning against smoking while in use; (ii) An illustration of the correct method of donning; and (iii) A visible warning against use with nasal obstructions or head colds with resultant nasal congestion. (c) If certification for operation above 18,000 feet (MSL) is requested, each oxygen dispensing unit must cover the nose and mouth of the user. (d) For a pressurized airplane designed to operate at flight altitudes above 25,000 feet (MSL), the dispensing units must meet the following: (1) The dispensing units for passengers must be connected to an oxygen supply terminal and be immediately available to each occupant wherever seated. (2) The dispensing units for crewmembers must be automatically presented to each crewmember before the cabin pressure altitude exceeds 15,000 feet, or the units must be of the quick-donning type, connected to an oxygen supply terminal that is immediately available to crewmembers at their station. (e) If certification for operation above 30,000 feet is requested, the dispensing units for passengers must be automatically presented to each occupant before the cabin pressure altitude exceeds 15,000 feet. (f) If an automatic dispensing unit (hose and mask, or other unit) system is installed, the crew must be provided with a manual means to make the dispensing units immediately available in the event of failure of the automatic system. [Amdt. 23–9, 35 FR 6387, Apr. 21, 1970, as amended by Amdt. 23–20, 42 FR 36969, July 18, 1977; Amdt. 23–30, 49 FR 7340, Feb. 28, 1984; Amdt. 23–43, 58 FR 18978, Apr. 9, 1993; Amdt. 23–49, 61 FR 5170, Feb. 9, 1996] § 23.1449 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 equipment. [Amdt. 23–9, 35 FR 6387, Apr. 21, 1970] § 23.1450 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 accordance with the following requirements: (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 hazardous. (c) In addition to meeting the requirements in paragraph (b) of this section, 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 °F. [Amdt. 23–20, 42 FR 36969, July 18, 1977] § 23.1451 Fire protection for oxygen equipment.Oxygen equipment and lines must: (a) Not be installed in any designed fire zones. (b) Be protected from heat that may be generated in, or escape from, any designated fire zone. (c) Be installed so that escaping oxygen cannot come in contact with and cause ignition of grease, fluid, or vapor accumulations that are present in normal operation or that may result from the failure or malfunction of any other system. [Doc. No. 27806, 61 FR 5170, Feb. 9, 1996] § 23.1453 Protection of oxygen equipment from rupture.(a) Each element of the oxygen system must have sufficient strength to withstand the maximum pressure and temperature, in combination with any externally applied loads arising from consideration of limit structural loads, that may be acting on that part of the system. (b) Oxygen pressure sources and the lines between the source and the shutoff means must be: (1) Protected from unsafe temperatures; and (2) Located where the probability and hazard of rupture in a crash landing are minimized. [Doc. No. 27806, 61 FR 5170, Feb. 9, 1996] § 23.1457 Cockpit voice recorders.(a) Each cockpit voice recorder required by the operating rules of this chapter must be approved and must be installed so that it will record the following: (1) Voice communications transmitted 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 loudspeaker system, if there is such a system and if the fourth channel is available in accordance with the requirements of paragraph (c)(4)(ii) of this section. (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 communications originating at the first and second pilot stations and voice communications of other crewmembers on the flight deck when directed to those stations. The microphone must be so located and, if necessary, the preamplifiers and filters of the recorder must be so adjusted or supplemented, so that the intelligibility of the recorded communications is as high as practicable when recorded under flight cockpit noise conditions and played back. Repeated aural or visual playback of the record may be used in evaluating intelligibility. (c) Each cockpit voice recorder must be installed so that the part of the communication or audio signals specified in paragraph (a) of this section obtained from each of the following sources is recorded on a separate channel: (1) For the first channel, from each boom, mask, or handheld microphone, headset, or speaker used at the first pilot station. (2) For the second channel from each boom, mask, or handheld 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 handheld microphone, headset, or speaker used at the station for the third and fourth crewmembers. (ii) If the stations specified in paragraph (c)(4)(i) of this section are not required or if the signal at such a station 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) And that 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 position of the interphone-transmitter key switch. The design shall ensure that sidetone for the flight crew is produced only when the interphone, public address system, or radio transmitters are in use. (d) Each cockpit voice recorder must be installed so that: (1) It receives its electric power from the bus that provides the maximum reliability for operation of the cockpit voice recorder without jeopardizing service to essential or emergency loads. (2) There is an automatic means to simultaneously stop the recorder and prevent each erasure feature from functioning, within 10 minutes after crash impact; and (3) There is an aural or visual means for preflight checking of the recorder for proper operation. (e) The record container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and consequent heat damage to the record from fire. In meeting this requirement, the record container must be as far aft as practicable, but may not be where aft mounted engines may crush the container during impact. However, it need not be outside of the pressurized compartment. (f) If the cockpit voice recorder has a bulk erasure device, the installation must be designed to minimize the probability of inadvertent operation and actuation of the device during crash impact. (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 location under water; and (3) Have an underwater locating device, 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. [Amdt. 23–35, 53 FR 26142, July 11, 1988] § 23.1459 Flight recorders.(a) Each flight recorder required by the operating rules of this chapter must be installed so that: (1) It is supplied with airspeed, altitude, and directional data obtained from sources that meet the accuracy requirements of §§23.1323, 23.1325, and 23.1327, as appropriate; (2) The vertical acceleration sensor is rigidly attached, and located longitudinally either within the approved center 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) It receives its electrical power power from the bus that provides the maximum reliability for operation of the flight recorder without jeopardizing service to essential or emergency loads; (4) There is an aural or visual means for preflight checking of the recorder for proper recording of data in the storage medium. (5) Except for recorders powered solely by the engine-driven electrical generator system, there is an automatic means to simultaneously stop a recorder that has a data erasure feature and prevent each erasure feature from functioning, within 10 minutes after crash impact; and (b) Each nonejectable record container must be located and mounted so as to minimize the probability of container rupture resulting from crash impact and subsequent damage to the record from fire. In meeting this requirement the record container must be located as far aft as practicable, but need not be aft of the pressurized compartment, and may not be where aft-mounted engines may crush the container upon impact. (c) A correlation must be established between the flight recorder readings of airspeed, altitude, and heading and the corresponding readings (taking into account correction factors) of the first pilot's instruments. The correlation must cover the airspeed range over which the airplane is to be operated, the range of altitude to which the airplane is limited, and 360 degrees of heading. Correlation may be established 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 location under water; and (3) Have an underwater locating device, 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 aircraft shall be evaluated to determine if any dedicated parameters must be recorded on flight recorders in addition to or in place of existing requirements. [Amdt. 23–35, 53 FR 26143, July 11, 1988] § 23.1461 Equipment containing high energy rotors.(a) Equipment, such as Auxiliary Power Units (APU) and constant speed drive units, containing high energy rotors must meet paragraphs (b), (c), or (d) of this section. (b) High energy rotors contained in equipment must be able to withstand damage caused by malfunctions, vibration, abnormal speeds, and abnormal temperatures. In addition— (1) Auxiliary rotor cases must be able to contain damage caused by the failure of high energy rotor blades; and (2) Equipment control devices, systems, and instrumentation must reasonably ensure that no operating limitations affecting the integrity of high energy rotors will be exceeded in service. (c) It must be shown by test that equipment containing high energy rotors 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 energy rotors must be located where rotor failure will neither endanger the occupants nor adversely affect continued safe flight. [Amdt. 23–20, 42 FR 36969, July 18, 1977, as amended by Amdt. 23–49, 61 FR 5170, Feb. 9, 1996] Browse Previous | Browse Next |
