Title 14—Aeronautics and Space--Volume 1
PART 25—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES
§ 25.961   Fuel system hot weather operation.
(a) The fuel system must perform satisfactorily in hot weather operation. This must be shown by showing that the fuel system from the tank outlets to each engine is pressurized, under all intended operations, so as to prevent vapor formation, or must be shown by climbing from the altitude of the airport elected by the applicant to the maximum altitude established as an operating limitation under § 25.1527. If a climb test is elected, there may be no evidence of vapor lock or other malfunctioning during the climb test conducted under the following conditions:
(1) For reciprocating engine powered airplanes, the engines must operate at maximum continuous power, except that takeoff power must be used for the altitudes from 1,000 feet below the critical altitude through the critical altitude. The time interval during which takeoff power is used may not be less than the takeoff time limitation.
(2) For turbine engine powered airplanes, the engines must operate at takeoff power for the time interval selected for showing the takeoff flight path, and at maximum continuous power for the rest of the climb.
(3) The weight of the airplane must be the weight with full fuel tanks, minimum crew, and the ballast necessary to maintain the center of gravity within allowable limits.
(4) The climb airspeed may not exceed—
(i) For reciprocating engine powered airplanes, the maximum airspeed established for climbing from takeoff to the maximum operating altitude with the airplane in the following configuration:
(A) Landing gear retracted.
(B) Wing flaps in the most favorable position.
(C) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition.
(D) Engine operating within the maximum continuous power limitations.
(E) Maximum takeoff weight; and
(ii) For turbine engine powered airplanes, the maximum airspeed established for climbing from takeoff to the maximum operating altitude.
(5) The fuel temperature must be at least 110 °F.
(b) The test prescribed in paragraph (a) of this section may be performed in flight or on the ground under closely simulated flight conditions. If a flight test is performed in weather cold enough to interfere with the proper conduct of the test, the fuel tank surfaces, fuel lines, and other fuel system parts subject to cold air must be insulated to simulate, insofar as practicable, flight in hot weather.
§ 25.963   Fuel tanks: general.
(a) Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads that it may be subjected to in operation.
(b) Flexible fuel tank liners must be approved or must be shown to be suitable for the particular application.
(c) Integral fuel tanks must have facilities for interior inspection and repair.
(d) Fuel tanks must, so far as it is practicable, be designed, located, and installed so that no fuel is released in or near the fuselage, or near the engines, in quantities that would constitute a fire hazard in otherwise survivable emergency landing conditions, and—
(1) Fuel tanks must be able to resist rupture and retain fuel under ultimate hydrostatic design conditions in which the pressure P within the tank varies in accordance with the formula:
(2) For those parts of wing fuel tanks near the fuselage or near the engines, the greater of the fuel pressures resulting from paragraphs (d)(2)(i) or (d)(2)(ii) of this section must be used:
(i) The fuel pressures resulting from paragraph (d)(1) of this section, and
(ii) The lesser of the two following conditions:
(A) Fuel pressures resulting from the accelerations specified in § 25.561(b)(3) considering the fuel tank full of fuel at maximum fuel density. Fuel pressures based on the 9.0g forward acceleration may be calculated using the fuel static head equal to the streamwise local chord of the tank. For inboard and outboard conditions, an acceleration of 1.5g may be used in lieu of 3.0g as specified in § 25.561(b)(3).
(B) Fuel pressures resulting from the accelerations as specified in § 25.561(b)(3) considering a fuel volume beyond 85 percent of the maximum permissible volume in each tank using the static head associated with the 85 percent fuel level. A typical density of the appropriate fuel may be used. For inboard and outboard conditions, an acceleration of 1.5g may be used in lieu of 3.0g as specified in § 25.561(b)(3).
(3) Fuel tank internal barriers and baffles may be considered as solid boundaries if shown to be effective in limiting fuel flow.
(4) For each fuel tank and surrounding airframe structure, the effects of crushing and scraping actions with the ground must not cause the spillage of enough fuel, or generate temperatures that would constitute a fire hazard under the conditions specified in § 25.721(b).
(5) Fuel tank installations must be such that the tanks will not rupture as a result of the landing gear or an engine pylon or engine mount tearing away as specified in § 25.721(a) and (c).
(e) Fuel tank access covers must comply with the following criteria in order to avoid loss of hazardous quantities of fuel:
(1) All covers located in an area where experience or analysis indicates a strike is likely must be shown by analysis or tests to minimize penetration and deformation by tire fragments, low energy engine debris, or other likely debris.
(2) All covers must be fire resistant as defined in part 1 of this chapter.
(f) For pressurized fuel tanks, a means with fail-safe features must be provided to prevent the buildup of an excessive pressure difference between the inside and the outside of the tank.
§ 25.965   Fuel tank tests.
(a) It must be shown by tests that the fuel tanks, as mounted in the airplane, can withstand, without failure or leakage, the more critical of the pressures resulting from the conditions specified in paragraphs (a)(1) and (2) of this section. In addition, it must be shown by either analysis or tests, that tank surfaces subjected to more critical pressures resulting from the condition of paragraphs (a)(3) and (4) of this section, are able to withstand the following pressures:
(1) An internal pressure of 3.5 psi.
(2) 125 percent of the maximum air pressure developed in the tank from ram effect.
(3) Fluid pressures developed during maximum limit accelerations, and deflections, of the airplane with a full tank.
(4) Fluid pressures developed during the most adverse combination of airplane roll and fuel load.
(b) Each metallic tank with large unsupported or unstiffened flat surfaces, whose failure or deformation could cause fuel leakage, must be able to withstand the following test, or its equivalent, without leakage or excessive deformation of the tank walls:
(1) Each complete tank assembly and its supports must be vibration tested while mounted to simulate the actual installation.
(2) Except as specified in paragraph (b)(4) of this section, the tank assembly must be vibrated for 25 hours at an amplitude of not less than 1/32 of an inch (unless another amplitude is substantiated) while 2/3 filled with water or other suitable test fluid.
(3) The test frequency of vibration must be as follows:
(i) If no frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the test frequency of vibration must be 2,000 cycles per minute.
(ii) If only one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, that frequency of vibration must be the test frequency.
(iii) If more than one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the most critical of these frequencies must be the test frequency.
(4) Under paragraphs (b)(3)(ii) and (iii) of this section, the time of test must be adjusted to accomplish the same number of vibration cycles that would be accomplished in 25 hours at the frequency specified in paragraph (b)(3)(i) of this section.
(5) During the test, the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute, through an angle of 15° on both sides of the horizontal (30° total), about the most critical axis, for 25 hours. If motion about more than one axis is likely to be critical, the tank must be rocked about each critical axis for 12 1/2 hours.
(c) Except where satisfactory operating experience with a similar tank in a similar installation is shown, nonmetallic tanks must withstand the test specified in paragraph (b)(5) of this section, with fuel at a temperature of 110 °F. During this test, a representative specimen of the tank must be installed in a supporting structure simulating the installation in the airplane.
(d) For pressurized fuel tanks, it must be shown by analysis or tests that the fuel tanks can withstand the maximum pressure likely to occur on the ground or in flight.
§ 25.967   Fuel tank installations.
(a) Each fuel tank must be supported so that tank loads (resulting from the weight of the fuel in the tanks) are not concentrated on unsupported tank surfaces. In addition—
(1) There must be pads, if necessary, to prevent chafing between the tank and its supports;
(2) Padding must be nonabsorbent or treated to prevent the absorption of fluids;
(3) If a flexible tank liner is used, it must be supported so that it is not required to withstand fluid loads; and
(4) Each interior surface of the tank compartment must be smooth and free of projections that could cause wear of the liner unless—
(i) Provisions are made for protection of the liner at these points; or
(ii) The construction of the liner itself provides that protection.
(b) Spaces adjacent to tank surfaces must be ventilated to avoid fume accumulation due to minor leakage. If the tank is in a sealed compartment, ventilation may be limited to drain holes large enough to prevent excessive pressure resulting from altitude changes.
(c) The location of each tank must meet the requirements of § 25.1185(a).
(d) No engine nacelle skin immediately behind a major air outlet from the engine compartment may act as the wall of an integral tank.
(e) Each fuel tank must be isolated from personnel compartments by a fumeproof and fuelproof enclosure.
§ 25.969   Fuel tank expansion space.
Each fuel tank must have an expansion space of not less than 2 percent of the tank capacity. It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude. For pressure fueling systems, compliance with this section may be shown with the means provided to comply with § 25.979(b).
§ 25.971   Fuel tank sump.
(a) Each fuel tank must have a sump with an effective capacity, in the normal ground attitude, of not less than the greater of 0.10 percent of the tank capacity or one-sixteenth of a gallon unless operating limitations are established to ensure that the accumulation of water in service will not exceed the sump capacity.
(b) Each fuel tank must allow drainage of any hazardous quantity of water from any part of the tank to its sump with the airplane in the ground attitude.
(c) Each fuel tank sump must have an accessible drain that—
(1) Allows complete drainage of the sump on the ground;
(2) Discharges clear of each part of the airplane; and
(3) Has manual or automatic means for positive locking in the closed position.
§ 25.973   Fuel tank filler connection.
Each fuel tank filler connection must prevent the entrance of fuel into any part of the airplane other than the tank itself. In addition—
(a) [Reserved]
(b) Each recessed filler connection that can retain any appreciable quantity of fuel must have a drain that discharges clear of each part of the airplane;
(c) Each filler cap must provide a fuel-tight seal; and
(d) Each fuel filling point must have a provision for electrically bonding the airplane to ground fueling equipment.
§ 25.975   Fuel tank vents and carburetor vapor vents.
(a) Fuel tank vents. Each fuel tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition. In addition—
(1) Each vent must be arranged to avoid stoppage by dirt or ice formation;
(2) The vent arrangement must prevent siphoning of fuel during normal operation;
(3) The venting capacity and vent pressure levels must maintain acceptable differences of pressure between the interior and exterior of the tank, during—
(i) Normal flight operation;
(ii) Maximum rate of ascent and descent; and
(iii) Refueling and defueling (where applicable);
(4) Airspaces of tanks with interconnected outlets must be interconnected;
(5) There may be no point in any vent line where moisture can accumulate with the airplane in the ground attitude or the level flight attitude, unless drainage is provided;
(6) No vent or drainage provision may end at any point—
(i) Where the discharge of fuel from the vent outlet would constitute a fire hazard; or
(ii) From which fumes could enter personnel compartments; and
(7) Each fuel tank vent system must prevent explosions, for a minimum of 2 minutes and 30 seconds, caused by propagation of flames from outside the tank through the fuel tank vents into fuel tank vapor spaces when any fuel tank vent is continuously exposed to flame.
(b) Carburetor vapor vents. Each carburetor with vapor elimination connections must have a vent line to lead vapors back to one of the fuel tanks. In addition—
(1) Each vent system must have means to avoid stoppage by ice; and
(2) If there is more than one fuel tank, and it is necessary to use the tanks in a definite sequence, each vapor vent return line must lead back to the fuel tank used for takeoff and landing.
§ 25.977   Fuel tank outlet.
(a) There must be a fuel strainer for the fuel tank outlet or for the booster pump. This strainer must—
(1) For reciprocating engine powered airplanes, have 8 to 16 meshes per inch; and
(2) For turbine engine powered airplanes, prevent the passage of any object that could restrict fuel flow or damage any fuel system component.
(b) [Reserved]
(c) The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line.
(d) The diameter of each strainer must be at least that of the fuel tank outlet.
(e) Each finger strainer must be accessible for inspection and cleaning.
§ 25.979   Pressure fueling system.
For pressure fueling systems, the following apply:
(a) Each pressure fueling system fuel manifold connection must have means to prevent the escape of hazardous quantities of fuel from the system if the fuel entry valve fails.
(b) An automatic shutoff means must be provided to prevent the quantity of fuel in each tank from exceeding the maximum quantity approved for that tank. This means must—
(1) Allow checking for proper shutoff operation before each fueling of the tank; and
(2) Provide indication at each fueling station of failure of the shutoff means to stop the fuel flow at the maximum quantity approved for that tank.
(c) A means must be provided to prevent damage to the fuel system in the event of failure of the automatic shutoff means prescribed in paragraph (b) of this section.
(d) The airplane pressure fueling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum pressures, including surge, that is likely to occur during fueling. The maximum surge pressure must be established with any combination of tank valves being either intentionally or inadvertently closed.
(e) The airplane defueling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum permissible defueling pressure (positive or negative) at the airplane fueling connection.
§ 25.981   Fuel tank explosion prevention.
(a) No ignition source may be present at each point in the fuel tank or fuel tank system where catastrophic failure could occur due to ignition of fuel or vapors. This must be shown by:
(1) Determining the highest temperature allowing a safe margin below the lowest expected autoignition temperature of the fuel in the fuel tanks.
(2) Demonstrating that no temperature at each place inside each fuel tank where fuel ignition is possible will exceed the temperature determined under paragraph (a)(1) of this section. This must be verified under all probable operating, failure, and malfunction conditions of each component whose operation, failure, or malfunction could increase the temperature inside the tank.
(3) Except for ignition sources due to lightning addressed by § 25.954, demonstrating that an ignition source could not result from each single failure, from each single failure in combination with each latent failure condition not shown to be extremely remote, and from all combinations of failures not shown to be extremely improbable, taking into account the effects of manufacturing variability, aging, wear, corrosion, and likely damage.
(b) Except as provided in paragraphs (b)(2) and (c) of this section, no fuel tank Fleet Average Flammability Exposure on an airplane may exceed three percent of the Flammability Exposure Evaluation Time (FEET) as defined in Appendix N of this part, or that of a fuel tank within the wing of the airplane model being evaluated, whichever is greater. If the wing is not a conventional unheated aluminum wing, the analysis must be based on an assumed Equivalent Conventional Unheated Aluminum Wing Tank.
(1) Fleet Average Flammability Exposure is determined in accordance with Appendix N of this part. The assessment must be done in accordance with the methods and procedures set forth in the Fuel Tank Flammability Assessment Method User's Manual, dated May 2008, document number DOT/FAA/AR-05/8 (incorporated by reference, see § 25.5).
(2) Any fuel tank other than a main fuel tank on an airplane must meet the flammability exposure criteria of Appendix M to this part if any portion of the tank is located within the fuselage contour.
(3) As used in this paragraph,
(i) Equivalent Conventional Unheated Aluminum Wing Tank is an integral tank in an unheated semi-monocoque aluminum wing of a subsonic airplane that is equivalent in aerodynamic performance, structural capability, fuel tank capacity and tank configuration to the designed wing.
(ii) Fleet Average Flammability Exposure is defined in Appendix N to this part and means the percentage of time each fuel tank ullage is flammable for a fleet of an airplane type operating over the range of flight lengths.
(iii) Main Fuel Tank means a fuel tank that feeds fuel directly into one or more engines and holds required fuel reserves continually throughout each flight.
(c) Paragraph (b) of this section does not apply to a fuel tank if means are provided to mitigate the effects of an ignition of fuel vapors within that fuel tank such that no damage caused by an ignition will prevent continued safe flight and landing.
(d) To protect design features that prevent catastrophic ignition sources within the fuel tank or fuel tank system according to paragraph (a) of this section, and to prevent increasing the flammability exposure of the tanks above that permitted in paragraph (b) of this section, the type design must include critical design configuration control limitations (CDCCLs) identifying those features and providing instructions on how to protect them. To ensure the continued effectiveness of those features, and prevent degradation of the performance and reliability of any means provided according to paragraphs (a), (b), or (c) of this section, the type design must also include necessary inspection and test procedures, intervals between repetitive inspections and tests, and mandatory replacement times for those features. The applicant must include information required by this paragraph in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by § 25.1529. The type design must also include visible means of identifying critical features of the design in areas of the airplane where foreseeable maintenance actions, repairs, or alterations may compromise the CDCCLs.
§ 25.991   Fuel pumps.
(a) Main pumps. Each fuel pump required for proper engine operation, or required to meet the fuel system requirements of this subpart (other than those in paragraph (b) of this section, is a main pump. For each main pump, provision must be made to allow the bypass of each positive displacement fuel pump other than a fuel injection pump (a pump that supplies the proper flow and pressure for fuel injection when the injection is not accomplished in a carburetor) approved as part of the engine.
(b) Emergency pumps. There must be emergency pumps or another main pump to feed each engine immediately after failure of any main pump (other than a fuel injection pump approved as part of the engine).
§ 25.993   Fuel system lines and fittings.
(a) Each fuel line must be installed and supported to prevent excessive vibration and to withstand loads due to fuel pressure and accelerated flight conditions.
(b) Each fuel line connected to components of the airplane between which relative motion could exist must have provisions for flexibility.
(c) Each flexible connection in fuel lines that may be under pressure and subjected to axial loading must use flexible hose assemblies.
(d) Flexible hose must be approved or must be shown to be suitable for the particular application.
(e) No flexible hose that might be adversely affected by exposure to high temperatures may be used where excessive temperatures will exist during operation or after engine shut-down.
(f) Each fuel line within the fuselage must be designed and installed to allow a reasonable degree of deformation and stretching without leakage.
§ 25.994   Fuel system components.
Fuel system components in an engine nacelle or in the fuselage must be protected from damage that could result in spillage of enough fuel to constitute a fire hazard as a result of a wheels-up landing on a paved runway under each of the conditions prescribed in § 25.721(b).
§ 25.995   Fuel valves.
In addition to the requirements of § 25.1189 for shutoff means, each fuel valve must—
(a) [Reserved]
(b) Be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve.
§ 25.997   Fuel strainer or filter.
There must be a fuel strainer or filter between the fuel tank outlet and the inlet of either the fuel metering device or an engine driven positive displacement pump, whichever is nearer the fuel tank outlet. This fuel strainer or filter must—
(a) Be accessible for draining and cleaning and must incorporate a screen or element which is easily removable;
(b) Have a sediment trap and drain except that it need not have a drain if the strainer or filter is easily removable for drain purposes;
(c) Be mounted so that its weight is not supported by the connecting lines or by the inlet or outlet connections of the strainer or filter itself, unless adequate strength margins under all loading conditions are provided in the lines and connections; and
(d) Have the capacity (with respect to operating limitations established for the engine) to ensure that engine fuel system functioning is not impaired, with the fuel contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine in Part 33 of this chapter.
§ 25.999   Fuel system drains.
(a) Drainage of the fuel system must be accomplished by the use of fuel strainer and fuel tank sump drains.
(b) Each drain required by paragraph (a) of this section must—
(1) Discharge clear of all parts of the airplane;
(2) Have manual or automatic means for positive locking in the closed position; and
(3) Have a drain valve—
(i) That is readily accessible and which can be easily opened and closed; and
(ii) That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted.
§ 25.1001   Fuel jettisoning system.
(a) A fuel jettisoning system must be installed on each airplane unless it is shown that the airplane meets the climb requirements of §§ 25.119 and 25.121(d) at maximum takeoff weight, less the actual or computed weight of fuel necessary for a 15-minute flight comprised of a takeoff, go-around, and landing at the airport of departure with the airplane configuration, speed, power, and thrust the same as that used in meeting the applicable takeoff, approach, and landing climb performance requirements of this part.
(b) If a fuel jettisoning system is required it must be capable of jettisoning enough fuel within 15 minutes, starting with the weight given in paragraph (a) of this section, to enable the airplane to meet the climb requirements of §§ 25.119 and 25.121(d), assuming that the fuel is jettisoned under the conditions, except weight, found least favorable during the flight tests prescribed in paragraph (c) of this section.
(c) Fuel jettisoning must be demonstrated beginning at maximum takeoff weight with flaps and landing gear up and in—
(1) A power-off glide at 1.3 V SR1 ;
(2) A climb at the one-engine inoperative best rate-of-climb speed, with the critical engine inoperative and the remaining engines at maximum continuous power; and
(3) Level flight at 1.3 V SR1 ; if the results of the tests in the conditions specified in paragraphs (c)(1) and (2) of this section show that this condition could be critical.
(d) During the flight tests prescribed in paragraph (c) of this section, it must be shown that—
(1) The fuel jettisoning system and its operation are free from fire hazard;
(2) The fuel discharges clear of any part of the airplane;
(3) Fuel or fumes do not enter any parts of the airplane; and
(4) The jettisoning operation does not adversely affect the controllability of the airplane.
(e) For reciprocating engine powered airplanes, means must be provided to prevent jettisoning the fuel in the tanks used for takeoff and landing below the level allowing 45 minutes flight at 75 percent maximum continuous power. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control.
(f) For turbine engine powered airplanes, means must be provided to prevent jettisoning the fuel in the tanks used for takeoff and landing below the level allowing climb from sea level to 10,000 feet and thereafter allowing 45 minutes cruise at a speed for maximum range. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control.
(g) The fuel jettisoning valve must be designed to allow flight personnel to close the valve during any part of the jettisoning operation.
(h) Unless it is shown that using any means (including flaps, slots, and slats) for changing the airflow across or around the wings does not adversely affect fuel jettisoning, there must be a placard, adjacent to the jettisoning control, to warn flight crewmembers against jettisoning fuel while the means that change the airflow are being used.
(i) The fuel jettisoning system must be designed so that any reasonably probable single malfunction in the system will not result in a hazardous condition due to unsymmetrical jettisoning of, or inability to jettison, fuel.
§ 25.1011   General.
(a) Each engine must have an independent oil system that can supply it with an appropriate quantity of oil at a temperature not above that safe for continuous operation.
(b) The usable oil capacity may not be less than the product of the endurance of the airplane under critical operating conditions and the approved maximum allowable oil consumption of the engine under the same conditions, plus a suitable margin to ensure system circulation. Instead of a rational analysis of airplane range for the purpose of computing oil requirements for reciprocating engine powered airplanes, the following fuel/oil ratios may be used:
(1) For airplanes without a reserve oil or oil transfer system, a fuel/oil ratio of 30:1 by volume.
(2) For airplanes with either a reserve oil or oil transfer system, a fuel/oil ratio of 40:1 by volume.
(c) Fuel/oil ratios higher than those prescribed in paragraphs (b)(1) and (2) of this section may be used if substantiated by data on actual engine oil consumption.
§ 25.1013   Oil tanks.
(a) Installation. Each oil tank installation must meet the requirements of § 25.967.
(b) Expansion space. Oil tank expansion space must be provided as follows:
(1) Each oil tank used with a reciprocating engine must have an expansion space of not less than the greater of 10 percent of the tank capacity or 0.5 gallon, and each oil tank used with a turbine engine must have an expansion space of not less than 10 percent of the tank capacity.
(2) Each reserve oil tank not directly connected to any engine may have an expansion space of not less than two percent of the tank capacity.
(3) It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude.
(c) Filler connection. Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of each part of the airplane. In addition, each oil tank filler cap must provide an oil-tight seal.
(d) Vent. Oil tanks must be vented as follows:
(1) Each oil tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition.
(2) Oil tank vents must be arranged so that condensed water vapor that might freeze and obstruct the line cannot accumulate at any point.
(e) Outlet. There must be means to prevent entrance into the tank itself, or into the tank outlet, of any object that might obstruct the flow of oil through the system. No oil tank outlet may be enclosed by any screen or guard that would reduce the flow of oil below a safe value at any operating temperature. There must be a shutoff valve at the outlet of each oil tank used with a turbine engine, unless the external portion of the oil system (including the oil tank supports) is fireproof.
(f) Flexible oil tank liners. Each flexible oil tank liner must be approved or must be shown to be suitable for the particular application.
§ 25.1015   Oil tank tests.
Each oil tank must be designed and installed so that—
(a) It can withstand, without failure, each vibration, inertia, and fluid load that it may be subjected to in operation; and
(b) It meets the provisions of § 25.965, except—
(1) The test pressure—
(i) For pressurized tanks used with a turbine engine, may not be less than 5 p.s.i. plus the maximum operating pressure of the tank instead of the pressure specified in § 25.965(a); and
(ii) For all other tanks may not be less than 5 p.s.i. instead of the pressure specified in § 25.965(a); and
(2) The test fluid must be oil at 250 °F. instead of the fluid specified in § 25.965(c).
§ 25.1017   Oil lines and fittings.
(a) Each oil line must meet the requirements of § 25.993 and each oil line and fitting in any designated fire zone must meet the requirements of § 25.1183.
(b) Breather lines must be arranged so that—
(1) Condensed water vapor that might freeze and obstruct the line cannot accumulate at any point;
(2) The breather discharge does not constitute a fire hazard if foaming occurs or causes emitted oil to strike the pilot's windshield; and
(3) The breather does not discharge into the engine air induction system.
§ 25.1019   Oil strainer or filter.
(a) Each turbine engine installation must incorporate an oil strainer or filter through which all of the engine oil flows and which meets the following requirements:
(1) Each oil strainer or filter that has a bypass must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter completely blocked.
(2) The oil strainer or filter must have the capacity (with respect to operating limitations established for the engine) to ensure that engine oil system functioning is not impaired when the oil is contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine under Part 33 of this chapter.
(3) The oil strainer or filter, unless it is installed at an oil tank outlet, must incorporate an indicator that will indicate contamination before it reaches the capacity established in accordance with paragraph (a)(2) of this section.
(4) The bypass of a strainer or filter must be constructed and installed so that the release of collected contaminants is minimized by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flow path.
(5) An oil strainer or filter that has no bypass, except one that is installed at an oil tank outlet, must have a means to connect it to the warning system required in § 25.1305(c)(7).
(b) Each oil strainer or filter in a powerplant installation using reciprocating engines must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter element completely blocked.
§ 25.1021   Oil system drains.
A drain (or drains) must be provided to allow safe drainage of the oil system. Each drain must—
(a) Be accessible; and
(b) Have manual or automatic means for positive locking in the closed position.
§ 25.1023   Oil radiators.
(a) Each oil radiator must be able to withstand, without failure, any vibration, inertia, and oil pressure load to which it would be subjected in operation.
(b) Each oil radiator air duct must be located so that, in case of fire, flames coming from normal openings of the engine nacelle cannot impinge directly upon the radiator.
§ 25.1025   Oil valves.
(a) Each oil shutoff must meet the requirements of § 25.1189.
(b) The closing of oil shutoff means may not prevent propeller feathering.
(c) Each oil valve must have positive stops or suitable index provisions in the “on” and “off” positions and must be supported so that no loads resulting from its operation or from accelerated flight conditions are transmitted to the lines attached to the valve.
§ 25.1027   Propeller feathering system.
(a) If the propeller feathering system depends on engine oil, there must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricating system other than the tank itself.
(b) The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump.
(c) The ability of the system to accomplish feathering with the trapped oil must be shown. This may be done on the ground using an auxiliary source of oil for lubricating the engine during operation.
(d) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system.
§ 25.1041   General.
The powerplant and auxiliary power unit cooling provisions must be able to maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the temperature limits established for these components and fluids, under ground, water, and flight operating conditions, and after normal engine or auxiliary power unit shutdown, or both.
§ 25.1043   Cooling tests.
(a) General. Compliance with § 25.1041 must be shown by tests, under critical ground, water, and flight operating conditions. For these tests, the following apply:
(1) If the tests are conducted under conditions deviating from the maximum ambient atmospheric temperature, the recorded powerplant temperatures must be corrected under paragraphs (c) and (d) of this section.
(2) No corrected temperatures determined under paragraph (a)(1) of this section may exceed established limits.
(3) For reciprocating engines, the fuel used during the cooling tests must be the minimum grade approved for the engines, and the mixture settings must be those normally used in the flight stages for which the cooling tests are conducted. The test procedures must be as prescribed in § 25.1045.
(b) Maximum ambient atmospheric temperature. A maximum ambient atmospheric temperature corresponding to sea level conditions of at least 100 degrees F must be established. The assumed temperature lapse rate is 3.6 degrees F per thousand feet of altitude above sea level until a temperature of −69.7 degrees F is reached, above which altitude the temperature is considered constant at −69.7 degrees F. However, for winterization installations, the applicant may select a maximum ambient atmospheric temperature corresponding to sea level conditions of less than 100 degrees F.
(c) Correction factor (except cylinder barrels). Unless a more rational correction applies, temperatures of engine fluids and powerplant components (except cylinder barrels) for which temperature limits are established, must be corrected by adding to them the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test.
(d) Correction factor for cylinder barrel temperatures. Unless a more rational correction applies, cylinder barrel temperatures must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum cylinder barrel temperature recorded during the cooling test.
§ 25.1045   Cooling test procedures.
(a) Compliance with § 25.1041 must be shown for the takeoff, climb, en route, and landing stages of flight that correspond to the applicable performance requirements. The cooling tests must be conducted with the airplane in the configuration, and operating under the conditions, that are critical relative to cooling during each stage of flight. For the cooling tests, a temperature is “stabilized” when its rate of change is less than two degrees F. per minute.
(b) Temperatures must be stabilized under the conditions from which entry is made into each stage of flight being investigated, unless the entry condition normally is not one during which component and the engine fluid temperatures would stabilize (in which case, operation through the full entry condition must be conducted before entry into the stage of flight being investigated in order to allow temperatures to reach their natural levels at the time of entry). The takeoff cooling test must be preceded by a period during which the powerplant component and engine fluid temperatures are stabilized with the engines at ground idle.
(c) Cooling tests for each stage of flight must be continued until—
(1) The component and engine fluid temperatures stabilize;
(2) The stage of flight is completed; or
(3) An operating limitation is reached.
(d) For reciprocating engine powered airplanes, it may be assumed, for cooling test purposes, that the takeoff stage of flight is complete when the airplane reaches an altitude of 1,500 feet above the takeoff surface or reaches a point in the takeoff where the transition from the takeoff to the en route configuration is completed and a speed is reached at which compliance with § 25.121(c) is shown, whichever point is at a higher altitude. The airplane must be in the following configuration:
(1) Landing gear retracted.
(2) Wing flaps in the most favorable position.
(3) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition.
(4) Critical engine inoperative and its propeller stopped.
(5) Remaining engines at the maximum continuous power available for the altitude.
(e) For hull seaplanes and amphibians, cooling must be shown during taxiing downwind for 10 minutes, at five knots above step speed.
§ 25.1091   Air induction.
(a) The air induction system for each engine and auxiliary power unit must supply—
(1) The air required by that engine and auxiliary power unit under each operating condition for which certification is requested; and
(2) The air for proper fuel metering and mixture distribution with the induction system valves in any position.
(b) Each reciprocating engine must have an alternate air source that prevents the entry of rain, ice, or any other foreign matter.
(c) Air intakes may not open within the cowling, unless—
(1) That part of the cowling is isolated from the engine accessory section by means of a fireproof diaphragm; or
(2) For reciprocating engines, there are means to prevent the emergence of backfire flames.
(d) For turbine engine powered airplanes and airplanes incorporating auxiliary power units—
(1) There must be means to prevent hazardous quantities of fuel leakage or overflow from drains, vents, or other components of flammable fluid systems from entering the engine or auxiliary power unit intake system; and
(2) The airplane must be designed to prevent water or slush on the runway, taxiway, or other airport operating surfaces from being directed into the engine or auxiliary power unit air inlet ducts in hazardous quantities, and the air inlet ducts must be located or protected so as to minimize the ingestion of foreign matter during takeoff, landing, and taxiing.
(e) If the engine induction system contains parts or components that could be damaged by foreign objects entering the air inlet, it must be shown by tests or, if appropriate, by analysis that the induction system design can withstand the foreign object ingestion test conditions of §§ 33.76, 33.77 and 33.78(a)(1) of this chapter without failure of parts or components that could create a hazard.
§ 25.1093   Induction system icing protection.
(a) Reciprocating engines. Each reciprocating engine air induction system must have means to prevent and eliminate icing. Unless this is done by other means, it must be shown that, in air free of visible moisture at a temperature of 30 F., each airplane with altitude engines using—
(1) Conventional venturi carburetors have a preheater that can provide a heat rise of 120 F. with the engine at 60 percent of maximum continuous power; or
(2) Carburetors tending to reduce the probability of ice formation has a preheater that can provide a heat rise of 100 °F. with the engine at 60 percent of maximum continuous power.
(b) Turbine engines. Except as provided in paragraph (b)(3) of this section, each engine, with all icing protection systems operating, must:
(1) Operate throughout its flight power range, including the minimum descent idling speeds, in the icing conditions defined in Appendices C and O of this part, and Appendix D of part 33 of this chapter, and in falling and blowing snow within the limitations established for the airplane for such operation, without the accumulation of ice on the engine, inlet system components, or airframe components that would do any of the following:
(i) Adversely affect installed engine operation or cause a sustained loss of power or thrust; or an unacceptable increase in gas path operating temperature; or an airframe/engine incompatibility; or
(ii) Result in unacceptable temporary power loss or engine damage; or
(iii) Cause a stall, surge, or flameout or loss of engine controllability (for example, rollback).
(2) Operate at ground idle speed for a minimum of 30 minutes on the ground in the following icing conditions shown in Table 1 of this section, unless replaced by similar test conditions that are more critical. These conditions must be demonstrated with the available air bleed for icing protection at its critical condition, without adverse effect, followed by an acceleration to takeoff power or thrust in accordance with the procedures defined in the airplane flight manual. During the idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Administrator. Analysis may be used to show ambient temperatures below the tested temperature are less critical. The applicant must document the engine run-up procedure (including the maximum time interval between run-ups from idle, run-up power setting, and duration at power), the associated minimum ambient temperature, and the maximum time interval. These conditions must be used in the analysis that establishes the airplane operating limitations in accordance with § 25.1521.
(3) For the purposes of this section, the icing conditions defined in appendix O of this part, including the conditions specified in Condition 3 of Table 1 of this section, are not applicable to airplanes with a maximum takeoff weight equal to or greater than 60,000 pounds.
Table 1—Icing Conditions for Ground Tests
(c) Supercharged reciprocating engines. For each engine having a supercharger to pressurize the air before it enters the carburetor, the heat rise in the air caused by that supercharging at any altitude may be utilized in determining compliance with paragraph (a) of this section if the heat rise utilized is that which will be available, automatically, for the applicable altitude and operating condition because of supercharging.
§ 25.1101   Carburetor air preheater design.
Each carburetor air preheater must be designed and constructed to—
(a) Ensure ventilation of the preheater when the engine is operated in cold air;
(b) Allow inspection of the exhaust manifold parts that it surrounds; and
(c) Allow inspection of critical parts of the preheater itself.
§ 25.1103   Induction system ducts and air duct systems.
(a) Each induction system duct upstream of the first stage of the engine supercharger and of the auxiliary power unit compressor must have a drain to prevent the hazardous accumulation of fuel and moisture in the ground attitude. No drain may discharge where it might cause a fire hazard.
(b) Each induction system duct must be—
(1) Strong enough to prevent induction system failures resulting from normal backfire conditions; and
(2) Fire-resistant if it is in any fire zone for which a fire-extinguishing system is required, except that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone.
(c) Each duct connected to components between which relative motion could exist must have means for flexibility.
(d) For turbine engine and auxiliary power unit bleed air duct systems, no hazard may result if a duct failure occurs at any point between the air duct source and the airplane unit served by the air.
(e) Each auxiliary power unit induction system duct must be fireproof for a sufficient distance upstream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through auxiliary power unit ducts and entering any other compartment or area of the airplane in which a hazard would be created resulting from the entry of hot gases. The materials used to form the remainder of the induction system duct and plenum chamber of the auxiliary power unit must be capable of resisting the maximum heat conditions likely to occur.
(f) Each auxiliary power unit induction system duct must be constructed of materials that will not absorb or trap hazardous quantities of flammable fluids that could be ignited in the event of a surge or reverse flow condition.
§ 25.1105   Induction system screens.
If induction system screens are used—
(a) Each screen must be upstream of the carburetor;
(b) No screen may be in any part of the induction system that is the only passage through which air can reach the engine, unless it can be deiced by heated air;
(c) No screen may be deiced by alcohol alone; and
(d) It must be impossible for fuel to strike any screen.
§ 25.1107   Inter-coolers and after-coolers.
Each inter-cooler and after-cooler must be able to withstand any vibration, inertia, and air pressure load to which it would be subjected in operation.
§ 25.1121   General.
For powerplant and auxiliary power unit installations the following apply:
(a) Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment. For test purposes, any acceptable carbon monoxide detection method may be used to show the absence of carbon monoxide.
(b) Each exhaust system part with a surface hot enough to ignite flammable fluids or vapors must be located or shielded so that leakage from any system carrying flammable fluids or vapors will not result in a fire caused by impingement of the fluids or vapors on any part of the exhaust system including shields for the exhaust system.
(c) Each component that hot exhaust gases could strike, or that could be subjected to high temperatures from exhaust system parts, must be fireproof. All exhaust system components must be separated by fireproof shields from adjacent parts of the airplane that are outside the engine and auxiliary power unit compartments.
(d) No exhaust gases may discharge so as to cause a fire hazard with respect to any flammable fluid vent or drain.
(e) No exhaust gases may discharge where they will cause a glare seriously affecting pilot vision at night.
(f) Each exhaust system component must be ventilated to prevent points of excessively high temperature.
(g) Each exhaust shroud must be ventilated or insulated to avoid, during normal operation, a temperature high enough to ignite any flammable fluids or vapors external to the shroud.
§ 25.1123   Exhaust piping.
For powerplant and auxiliary power unit installations, the following apply:
(a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures.
(b) Piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation; and
(c) Piping connected to components between which relative motion could exist must have means for flexibility.
§ 25.1125   Exhaust heat exchangers.
For reciprocating engine powered airplanes, the following apply:
(a) Each exhaust heat exchanger must be constructed and installed to withstand each vibration, inertia, and other load to which it would be subjected in operation. In addition—
(1) Each exchanger must be suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;
(2) There must be means for the inspection of the critical parts of each exchanger;
(3) Each exchanger must have cooling provisions wherever it is subject to contact with exhaust gases; and
(4) No exhaust heat exchanger or muff may have any stagnant areas or liquid traps that would increase the probability of ignition of flammable fluids or vapors that might be present in case of the failure or malfunction of components carrying flammable fluids.
(b) If an exhaust heat exchanger is used for heating ventilating air—
(1) There must be a secondary heat exchanger between the primary exhaust gas heat exchanger and the ventilating air system; or
(2) Other means must be used to preclude the harmful contamination of the ventilating air.
§ 25.1127   Exhaust driven turbo-superchargers.
(a) Each exhaust driven turbo-supercharger must be approved or shown to be suitable for the particular application. It must be installed and supported to ensure safe operation between normal inspections and overhauls. In addition, there must be provisions for expansion and flexibility between exhaust conduits and the turbine.
(b) There must be provisions for lubricating the turbine and for cooling turbine parts where temperatures are critical.
(c) If the normal turbo-supercharger control system malfunctions, the turbine speed may not exceed its maximum allowable value. Except for the waste gate operating components, the components provided for meeting this requirement must be independent of the normal turbo-supercharger controls.
§ 25.1141   Powerplant controls: general.
Each powerplant control must be located, arranged, and designed under §§ 25.777 through 25.781 and marked under § 25.1555. In addition, it must meet the following requirements:
(a) Each control must be located so that it cannot be inadvertently operated by persons entering, leaving, or moving normally in, the cockpit.
(b) Each flexible control must be approved or must be shown to be suitable for the particular application.
(c) Each control must have sufficient strength and rigidity to withstand operating loads without failure and without excessive deflection.
(d) Each control must be able to maintain any set position without constant attention by flight crewmembers and without creep due to control loads or vibration.
(e) The portion of each powerplant control located in a designated fire zone that is required to be operated in the event of fire must be at least fire resistant.
(f) For powerplant valve controls located in the flight deck there must be a means:
(1) For the flightcrew to select each intended position or function of the valve; and
(2) To indicate to the flightcrew:
(i) The selected position or function of the valve; and
(ii) When the valve has not responded as intended to the selected position or function.
§ 25.1142   Auxiliary power unit controls.
Means must be provided on the flight deck for starting, stopping, and emergency shutdown of each installed auxiliary power unit.
§ 25.1143   Engine controls.
(a) There must be a separate power or thrust control for each engine.
(b) Power and thrust controls must be arranged to allow—
(1) Separate control of each engine; and
(2) Simultaneous control of all engines.
(c) Each power and thrust control must provide a positive and immediately responsive means of controlling its engine.
(d) For each fluid injection (other than fuel) system and its controls not provided and approved as part of the engine, the applicant must show that the flow of the injection fluid is adequately controlled.
(e) If a power or thrust control incorporates a fuel shutoff feature, the control must have a means to prevent the inadvertent movement of the control into the shutoff position. The means must—
(1) Have a positive lock or stop at the idle position; and
(2) Require a separate and distinct operation to place the control in the shutoff position.
§ 25.1145   Ignition switches.
(a) Ignition switches must control each engine ignition circuit on each engine.
(b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control.
(c) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means to prevent its inadvertent operation.
§ 25.1147   Mixture controls.
(a) If there are mixture controls, each engine must have a separate control. The controls must be grouped and arranged to allow—
(1) Separate control of each engine; and
(2) Simultaneous control of all engines.
(b) Each intermediate position of the mixture controls that corresponds to a normal operating setting must be identifiable by feel and sight.
(c) The mixture controls must be accessible to both pilots. However, if there is a separate flight engineer station with a control panel, the controls need be accessible only to the flight engineer.
§ 25.1149   Propeller speed and pitch controls.
(a) There must be a separate propeller speed and pitch control for each propeller.
(b) The controls must be grouped and arranged to allow—
(1) Separate control of each propeller; and
(2) Simultaneous control of all propellers.
(c) The controls must allow synchronization of all propellers.
(d) The propeller speed and pitch controls must be to the right of, and at least one inch below, the pilot's throttle controls.
§ 25.1153   Propeller feathering controls.
(a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation.
(b) If feathering is accomplished by movement of the propeller pitch or speed control lever, there must be means to prevent the inadvertent movement of this lever to the feathering position during normal operation.
§ 25.1155   Reverse thrust and propeller pitch settings below the flight regime.
Each control for reverse thrust and for propeller pitch settings below the flight regime must have means to prevent its inadvertent operation. The means must have a positive lock or stop at the flight idle position and must require a separate and distinct operation by the crew to displace the control from the flight regime (forward thrust regime for turbojet powered airplanes).
§ 25.1157   Carburetor air temperature controls.
There must be a separate carburetor air temperature control for each engine.
§ 25.1159   Supercharger controls.
Each supercharger control must be accessible to the pilots or, if there is a separate flight engineer station with a control panel, to the flight engineer.
§ 25.1161   Fuel jettisoning system controls.
Each fuel jettisoning system control must have guards to prevent inadvertent operation. No control may be near any fire extinguisher control or other control used to combat fire.
§ 25.1163   Powerplant accessories.
(a) Each engine mounted accessory must—
(1) Be approved for mounting on the engine involved;
(2) Use the provisions on the engine for mounting; and
(3) Be sealed to prevent contamination of the engine oil system and the accessory system.
(b) Electrical equipment subject to arcing or sparking must be installed to minimize the probability of contact with any flammable fluids or vapors that might be present in a free state.
(c) If continued rotation of an engine-driven cabin supercharger or of any remote accessory driven by the engine is hazardous if malfunctioning occurs, there must be means to prevent rotation without interfering with the continued operation of the engine.
§ 25.1165   Engine ignition systems.
(a) Each battery ignition system must be supplemented by a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any battery becomes depleted.
(b) The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw electrical energy from the same source.
(c) The design of the engine ignition system must account for—
(1) The condition of an inoperative generator;
(2) The condition of a completely depleted battery with the generator running at its normal operating speed; and
(3) The condition of a completely depleted battery with the generator operating at idling speed, if there is only one battery.
(d) Magneto ground wiring (for separate ignition circuits) that lies on the engine side of the fire wall, must be installed, located, or protected, to minimize the probability of simultaneous failure of two or more wires as a result of mechanical damage, electrical faults, or other cause.
(e) No ground wire for any engine may be routed through a fire zone of another engine unless each part of that wire within that zone is fireproof.
(f) Each ignition system must be independent of any electrical circuit, not used for assisting, controlling, or analyzing the operation of that system.
(g) There must be means to warn appropriate flight crewmembers if the malfunctioning of any part of the electrical system is causing the continuous discharge of any battery necessary for engine ignition.
(h) Each engine ignition system of a turbine powered airplane must be considered an essential electrical load.
§ 25.1167   Accessory gearboxes.
For airplanes equipped with an accessory gearbox that is not certificated as part of an engine—
(a) The engine with gearbox and connecting transmissions and shafts attached must be subjected to the tests specified in § 33.49 or § 33.87 of this chapter, as applicable;
(b) The accessory gearbox must meet the requirements of §§ 33.25 and 33.53 or 33.91 of this chapter, as applicable; and
(c) Possible misalignments and torsional loadings of the gearbox, transmission, and shaft system, expected to result under normal operating conditions must be evaluated.
§ 25.1181   Designated fire zones; regions included.
(a) Designated fire zones are—
(1) The engine power section;
(2) The engine accessory section;
(3) Except for reciprocating engines, any complete powerplant compartment in which no isolation is provided between the engine power section and the engine accessory section;
(4) Any auxiliary power unit compartment;
(5) Any fuel-burning heater and other combustion equipment installation described in § 25.859;
(6) The compressor and accessory sections of turbine engines; and
(7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases.
(b) Each designated fire zone must meet the requirements of §§ 25.863, 25.865, 25.867, 25.869, and 25.1185 through 25.1203.
§ 25.1182   Nacelle areas behind firewalls, and engine pod attaching structures containing flammable fluid lines.
(a) Each nacelle area immediately behind the firewall, and each portion of any engine pod attaching structure containing flammable fluid lines, must meet each requirement of §§ 25.1103(b), 25.1165 (d) and (e), 25.1183, 25.1185(c), 25.1187, 25.1189, and 25.1195 through 25.1203, including those concerning designated fire zones. However, engine pod attaching structures need not contain fire detection or extinguishing means.
(b) For each area covered by paragraph (a) of this section that contains a retractable landing gear, compliance with that paragraph need only be shown with the landing gear retracted.
§ 25.1183   Flammable fluid-carrying components.
(a) Except as provided in paragraph (b) of this section, each line, fitting, and other component carrying flammable fluid in any area subject to engine fire conditions, and each component which conveys or contains flammable fluid in a designated fire zone must be fire resistant, except that flammable fluid tanks and supports in a designated fire zone must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non-fireproof part will not cause leakage or spillage of flammable fluid. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid. An integral oil sump of less than 25-quart capacity on a reciprocating engine need not be fireproof nor be enclosed by a fireproof shield.
(b) Paragraph (a) of this section does not apply to—
(1) Lines, fittings, and components which are already approved as part of a type certificated engine; and
(2) Vent and drain lines, and their fittings, whose failure will not result in, or add to, a fire hazard.
(c) All components, including ducts, within a designated fire zone must be fireproof if, when exposed to or damaged by fire, they could—
(1) Result in fire spreading to other regions of the airplane; or
(2) Cause unintentional operation of, or inability to operate, essential services or equipment.
§ 25.1185   Flammable fluids.
(a) Except for the integral oil sumps specified in § 25.1183(a), no tank or reservoir that is a part of a system containing flammable fluids or gases may be in a designated fire zone unless the fluid contained, the design of the system, the materials used in the tank, the shut-off means, and all connections, lines, and control provide a degree of safety equal to that which would exist if the tank or reservoir were outside such a zone.
(b) There must be at least one-half inch of clear airspace between each tank or reservoir and each firewall or shroud isolating a designated fire zone.
(c) Absorbent materials close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids.
§ 25.1187   Drainage and ventilation of fire zones.
(a) There must be complete drainage of each part of each designated fire zone to minimize the hazards resulting from failure or malfunctioning of any component containing flammable fluids. The drainage means must be—
(1) Effective under conditions expected to prevail when drainage is needed; and
(2) Arranged so that no discharged fluid will cause an additional fire hazard.
(b) Each designated fire zone must be ventilated to prevent the accumulation of flammable vapors.
(c) No ventilation opening may be where it would allow the entry of flammable fluids, vapors, or flame from other zones.
(d) Each ventilation means must be arranged so that no discharged vapors will cause an additional fire hazard.
(e) Unless the extinguishing agent capacity and rate of discharge are based on maximum air flow through a zone, there must be means to allow the crew to shut off sources of forced ventilation to any fire zone except the engine power section of the nacelle and the combustion heater ventilating air ducts.
§ 25.1189   Shutoff means.
(a) Each engine installation and each fire zone specified in § 25.1181(a)(4) and (5) must have a means to shut off or otherwise prevent hazardous quantities of fuel, oil, deicer, and other flammable fluids, from flowing into, within, or through any designated fire zone, except that shutoff means are not required for—
(1) Lines, fittings, and components forming an integral part of an engine; and
(2) Oil systems for turbine engine installations in which all components of the system in a designated fire zone, including oil tanks, are fireproof or located in areas not subject to engine fire conditions.
(b) The closing of any fuel shutoff valve for any engine may not make fuel unavailable to the remaining engines.
(c) Operation of any shutoff may not interfere with the later emergency operation of other equipment, such as the means for feathering the propeller.
(d) Each flammable fluid shutoff means and control must be fireproof or must be located and protected so that any fire in a fire zone will not affect its operation.
(e) No hazardous quantity of flammable fluid may drain into any designated fire zone after shutoff.
(f) There must be means to guard against inadvertent operation of the shutoff means and to make it possible for the crew to reopen the shutoff means in flight after it has been closed.
(g) Each tank-to-engine shutoff valve must be located so that the operation of the valve will not be affected by powerplant or engine mount structural failure.
(h) Each shutoff valve must have a means to relieve excessive pressure accumulation unless a means for pressure relief is otherwise provided in the system.
§ 25.1191   Firewalls.
(a) Each engine, auxiliary power unit, fuel-burning heater, other combustion equipment intended for operation in flight, and the combustion, turbine, and tailpipe sections of turbine engines, must be isolated from the rest of the airplane by firewalls, shrouds, or equivalent means.
(b) Each firewall and shroud must be—
(1) Fireproof;
(2) Constructed so that no hazardous quantity of air, fluid, or flame can pass from the compartment to other parts of the airplane;
(3) Constructed so that each opening is sealed with close fitting fireproof grommets, bushings, or firewall fittings; and
(4) Protected against corrosion.
§ 25.1192   Engine accessory section diaphragm.
For reciprocating engines, the engine power section and all portions of the exhaust system must be isolated from the engine accessory compartment by a diaphragm that complies with the firewall requirements of § 25.1191.
§ 25.1193   Cowling and nacelle skin.
(a) Each cowling must be constructed and supported so that it can resist any vibration, inertia, and air load to which it may be subjected in operation.
(b) Cowling must meet the drainage and ventilation requirements of § 25.1187.
(c) On airplanes with a diaphragm isolating the engine power section from the engine accessory section, each part of the accessory section cowling subject to flame in case of fire in the engine power section of the powerplant must—
(1) Be fireproof; and
(2) Meet the requirements of § 25.1191.
(d) Each part of the cowling subject to high temperatures due to its nearness to exhaust system parts or exhaust gas impingement must be fireproof.
(e) Each airplane must—
(1) Be designed and constructed so that no fire originating in any fire zone can enter, either through openings or by burning through external skin, any other zone or region where it would create additional hazards;
(2) Meet paragraph (e)(1) of this section with the landing gear retracted (if applicable); and
(3) Have fireproof skin in areas subject to flame if a fire starts in the engine power or accessory sections.
§ 25.1195   Fire extinguishing systems.
(a) Except for combustor, turbine, and tail pipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases for which it is shown that a fire originating in these sections can be controlled, there must be a fire extinguisher system serving each designated fire zone.
(b) The fire extinguishing system, the quantity of the extinguishing agent, the rate of discharge, and the discharge distribution must be adequate to extinguish fires. It must be shown by either actual or simulated flights tests that under critical airflow conditions in flight the discharge of the extinguishing agent in each designated fire zone specified in paragraph (a) of this section will provide an agent concentration capable of extinguishing fires in that zone and of minimizing the probability of reignition. An individual “one-shot” system may be used for auxiliary power units, fuel burning heaters, and other combustion equipment. For each other designated fire zone, two discharges must be provided each of which produces adequate agent concentration.
(c) The fire extinguishing system for a nacelle must be able to simultaneously protect each zone of the nacelle for which protection is provided.
§ 25.1197   Fire extinguishing agents.
(a) Fire extinguishing agents must—
(1) Be capable of extinguishing flames emanating from any burning of fluids or other combustible materials in the area protected by the fire extinguishing system; and
(2) Have thermal stability over the temperature range likely to be experienced in the compartment in which they are stored.
(b) If any toxic extinguishing agent is used, provisions must be made to prevent harmful concentrations of fluid or fluid vapors (from leakage during normal operation of the airplane or as a result of discharging the fire extinguisher on the ground or in flight) from entering any personnel compartment, even though a defect may exist in the extinguishing system. This must be shown by test except for built-in carbon dioxide fuselage compartment fire extinguishing systems for which—
(1) Five pounds or less of carbon dioxide will be discharged, under established fire control procedures, into any fuselage compartment; or
(2) There is protective breathing equipment for each flight crewmember on flight deck duty.
§ 25.1199   Extinguishing agent containers.
(a) Each extinguishing agent container must have a pressure relief to prevent bursting of the container by excessive internal pressures.
(b) The discharge end of each discharge line from a pressure relief connection must be located so that discharge of the fire extinguishing agent would not damage the airplane. The line must also be located or protected to prevent clogging caused by ice or other foreign matter.
(c) There must be a means for each fire extinguishing agent container to indicate that the container has discharged or that the charging pressure is below the established minimum necessary for proper functioning.
(d) The temperature of each container must be maintained, under intended operating conditions, to prevent the pressure in the container from—
(1) Falling below that necessary to provide an adequate rate of discharge; or
(2) Rising high enough to cause premature discharge.
(e) If a pyrotechnic capsule is used to discharge the extinguishing agent, each container must be installed so that temperature conditions will not cause hazardous deterioration of the pyrotechnic capsule.
§ 25.1201   Fire extinguishing system materials.
(a) No material in any fire extinguishing system may react chemically with any extinguishing agent so as to create a hazard.
(b) Each system component in an engine compartment must be fireproof.
§ 25.1203   Fire detector system.
(a) There must be approved, quick acting fire or overheat detectors in each designated fire zone, and in the combustion, turbine, and tailpipe sections of turbine engine installations, in numbers and locations ensuring prompt detection of fire in those zones.
(b) Each fire detector system must be constructed and installed so that—
(1) It will withstand the vibration, inertia, and other loads to which it may be subjected in operation;
(2) There is a means to warn the crew in the event that the sensor or associated wiring within a designated fire zone is severed at one point, unless the system continues to function as a satisfactory detection system after the severing; and
(3) There is a means to warn the crew in the event of a short circuit in the sensor or associated wiring within a designated fire zone, unless the system continues to function as a satisfactory detection system after the short circuit.
(c) No fire or overheat detector may be affected by any oil, water, other fluids or fumes that might be present.
(d) There must be means to allow the crew to check, in flight, the functioning of each fire or overheat detector electric circuit.
(e) Components of each fire or overheat detector system in a fire zone must be fire-resistant.
(f) No fire or overheat detector system component for any fire zone may pass through another fire zone, unless—
(1) It is protected against the possibility of false warnings resulting from fires in zones through which it passes; or
(2) Each zone involved is simultaneously protected by the same detector and extinguishing system.
(g) Each fire detector system must be constructed so that when it is in the configuration for installation it will not exceed the alarm activation time approved for the detectors using the response time criteria specified in the appropriate Technical Standard Order for the detector.
(h) EWIS for each fire or overheat detector system in a fire zone must meet the requirements of § 25.1731.
§ 25.1207   Compliance.
Unless otherwise specified, compliance with the requirements of §§ 25.1181 through 25.1203 must be shown by a full scale fire test or by one or more of the following methods:
(a) Tests of similar powerplant configurations;
(b) Tests of components;
(c) Service experience of aircraft with similar powerplant configurations;
(d) Analysis.
§ 25.1301   Function and installation.
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; and
(c) Be installed according to limitations specified for that equipment.
§ 25.1302   Installed systems and equipment for use by the flightcrew.
This section applies to installed systems and equipment intended for flightcrew members' use in operating the airplane from their normally seated positions on the flight deck. The applicant must show that these systems and installed equipment, individually and in combination with other such systems and equipment, are designed so that qualified flightcrew members trained in their use can safely perform all of the tasks associated with the systems' and equipment's intended functions. Such installed equipment and systems must meet the following requirements:
(a) Flight deck controls must be installed to allow accomplishment of all the tasks required to safely perform the equipment's intended function, and information must be provided to the flightcrew that is necessary to accomplish the defined tasks.
(b) Flight deck controls and information intended for the flightcrew's use must:
(1) Be provided in a clear and unambiguous manner at a resolution and precision appropriate to the task;
(2) Be accessible and usable by the flightcrew in a manner consistent with the urgency, frequency, and duration of their tasks; and
(3) Enable flightcrew awareness, if awareness is required for safe operation, of the effects on the airplane or systems resulting from flightcrew actions.
(c) Operationally-relevant behavior of the installed equipment must be:
(1) Predictable and unambiguous; and
(2) Designed to enable the flightcrew to intervene in a manner appropriate to the task.
(d) To the extent practicable, installed equipment must incorporate means to enable the flightcrew to manage errors resulting from the kinds of flightcrew interactions with the equipment that can be reasonably expected in service. This paragraph does not apply to any of the following:
(1) Skill-related errors associated with manual control of the airplane;
(2) Errors that result from decisions, actions, or omissions committed with malicious intent;
(3) Errors arising from a crewmember's reckless decisions, actions, or omissions reflecting a substantial disregard for safety; and
(4) Errors resulting from acts or threats of violence, including actions taken under duress.
§ 25.1303   Flight and navigation instruments.
(a) The following flight and navigation instruments must be installed so that the instrument is visible from each pilot station:
(1) A free air temperature indicator or an air-temperature indicator which provides indications that are convertible to free-air temperature.
(2) A clock displaying hours, minutes, and seconds with a sweep-second pointer or digital presentation.
(3) A direction indicator (nonstabilized magnetic compass).
(b) The following flight and navigation instruments must be installed at each pilot station:
(1) An airspeed indicator. If airspeed limitations vary with altitude, the indicator must have a maximum allowable airspeed indicator showing the variation of V MO with altitude.
(2) An altimeter (sensitive).
(3) A rate-of-climb indicator (vertical speed).
(4) A gyroscopic rate-of-turn indicator combined with an integral slip-skid indicator (turn-and-bank indicator) except that only a slip-skid indicator is required on large airplanes with a third attitude instrument system useable through flight attitudes of 360° of pitch and roll and installed in accordance with § 121.305(k) of this title.
(5) A bank and pitch indicator (gyroscopically stabilized).
(6) A direction indicator (gyroscopically stabilized, magnetic or nonmagnetic).
(c) The following flight and navigation instruments are required as prescribed in this paragraph:
(1) A speed warning device is required for turbine engine powered airplanes and for airplanes with V MO /M MO greater than 0.8 V DF /M DF or 0.8 V D /M D . 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 V MO plus 6 knots or M MO + 0.01. The upper limit of the production tolerance for the warning device may not exceed the prescribed warning speed.
(2) A machmeter is required at each pilot station for airplanes with compressibility limitations not otherwise indicated to the pilot by the airspeed indicating system required under paragraph (b)(1) of this section.
§ 25.1305   Powerplant instruments.
The following are required powerplant instruments:
(a) For all airplanes. (1) A fuel pressure warning means for each engine, or a master warning means for all engines with provision for isolating the individual warning means from the master warning means.
(2) A fuel quantity indicator for each fuel tank.
(3) An oil quantity indicator for each oil tank.
(4) An oil pressure indicator for each independent pressure oil system of each engine.
(5) An oil pressure warning means for each engine, or a master warning means for all engines with provision for isolating the individual warning means from the master warning means.
(6) An oil temperature indicator for each engine.
(7) Fire-warning devices that provide visual and audible warning.
(8) An augmentation liquid quantity indicator (appropriate for the manner in which the liquid is to be used in operation) for each tank.
(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) A carburetor air temperature indicator for each engine.
(2) A cylinder head temperature indicator for each air-cooled engine.
(3) A manifold pressure indicator for each engine.
(4) A fuel pressure indicator (to indicate the pressure at which the fuel is supplied) for each engine.
(5) A fuel flowmeter, or fuel mixture indicator, for each engine without an automatic altitude mixture control.
(6) A tachometer for each engine.
(7) A device that indicates, to the flight crew (during flight), any change in the power output, for each engine with—
(i) An automatic propeller feathering system, whose operation is initiated by a power output measuring system; or
(ii) A total engine piston displacement of 2,000 cubic inches or more.
(8) A means to indicate to the pilot when the propeller is in reverse pitch, for each reversing propeller.
(c) For turbine engine-powered 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 tachometer (to indicate the speed of the rotors with established limiting speeds) for each engine.
(4) A means to indicate, to the flight crew, the operation of each engine starter that can be operated continuously but that is neither designed for continuous operation nor designed to prevent hazard if it failed.
(5) An indicator to indicate the functioning of the powerplant ice protection system for each engine.
(6) An indicator for the fuel strainer or filter required by § 25.997 to indicate the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with § 25.997(d).
(7) A warning means for the oil strainer or filter required by § 25.1019, if it has no bypass, to warn the pilot of the occurrence of contamination of the strainer or filter screen before it reaches the capacity established in accordance with § 25.1019(a)(2).
(8) An indicator to indicate the proper functioning of any heater used to prevent ice clogging of fuel system components.
(d) For turbojet 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) An indicator to indicate thrust, or a parameter that is directly related to thrust, to the pilot. The indication must be based on the direct measurement of thrust or of parameters that are directly related to thrust. The indicator must indicate a change in thrust resulting from any engine malfunction, damage, or deterioration.
(2) A position indicating means to indicate to the flightcrew when the thrust reversing device—
(i) Is not in the selected position, and
(ii) Is in the reverse thrust position, for each engine using a thrust reversing device.
(3) An indicator to indicate rotor system unbalance.
(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) Position indicating means to indicate to the flight crew when the propeller blade angle is below the flight low pitch position, for each propeller.
(f) For airplanes equipped with fluid systems (other than fuel) for thrust or power augmentation, an approved means must be provided to indicate the proper functioning of that system to the flight crew.
§ 25.1307   Miscellaneous equipment.
The following is required miscellaneous equipment:
(a) [Reserved]
(b) Two or more independent sources of electrical energy.
(c) Electrical protective devices, as prescribed in this part.
(d) Two systems for two-way radio communications, with controls for each accessible from each pilot station, designed and installed so that failure of one system will not preclude operation of the other system. The use of a common antenna system is acceptable if adequate reliability is shown.
(e) Two systems for radio navigation, with controls for each accessible from each pilot station, designed and installed so that failure of one system will not preclude operation of the other system. The use of a common antenna system is acceptable if adequate reliability is shown.
§ 25.1309   Equipment, systems, and installations.
The requirements of this section, except as identified below, apply to any equipment or system as installed on the airplane. Although this section does not apply to the performance and flight characteristic requirements of subpart B of this part, or to the structural requirements of subparts C and D of this part, it does apply to any system on which compliance with any of those requirements is dependent. Section 25.1309(b) does not apply to the flight control jam conditions addressed by § 25.671(c)(3); single failures in the brake system addressed by § 25.735(b)(1); the failure conditions addressed by §§ 25.810(a)(1)(v) and 25.812; uncontained engine rotor failure, engine case rupture, or engine case burn-through failures addressed by §§ 25.903(d)(1) and 25.1193 and part 33 of this chapter; and propeller debris release failures addressed by § 25.905(d) and part 35 of this chapter.
(a) The airplane's equipment and systems must be designed and installed so that:
(1) The equipment and systems required for type certification or by operating rules, or whose improper functioning would reduce safety, perform as intended under the airplane operating and environmental conditions; and
(2) Other equipment and systems, functioning normally or abnormally, do not adversely affect the safety of the airplane or its occupants or the proper functioning of the equipment and systems addressed by paragraph (a)(1) of this section.
(b) The airplane systems and associated components, evaluated separately and in relation to other systems, must be designed and installed so that they meet all of the following requirements:
(1) Each catastrophic failure condition—
(i) Must be extremely improbable; and
(ii) Must not result from a single failure.
(2) Each hazardous failure condition must be extremely remote.
(3) Each major failure condition must be remote.
(4) Each significant latent failure must be eliminated as far as practical, or, if not practical to eliminate, the latency of the significant latent failure must be minimized. However, the requirements of the previous sentence do not apply if the associated system meets the requirements of paragraphs (b)(1) and (b)(2) of this section, assuming the significant latent failure has occurred.
(5) For each catastrophic failure condition that results from two failures, either of which could be latent for more than one flight, the applicant must show that—
(i) It is impractical to provide additional fault tolerance; and
(ii) Given the occurrence of any single latent failure, the residual average probability of the catastrophic failure condition due to all subsequent active failures is remote; and
(iii) The sum of the probabilities of the latent failures that are combined with each active failure does not exceed 1/1000.
(c) The airplane and systems must provide information concerning unsafe system operating conditions to the flightcrew to enable them to take appropriate corrective action in a timely manner. Systems and controls, including information, indications, and annunciations, must be designed to minimize flightcrew errors that could create additional hazards.
(d) [Reserved]
(e) The applicant must establish certification maintenance requirements as necessary to prevent the development of the failure conditions described in paragraph (b) of this section. These requirements must be included in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by § 25.1529.
§ 25.1310   Power source capacity and distribution.
(a) Each installation whose functioning is required for type certification or under operating rules 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 system with the system functioning normally.
(2) Essential loads, after failure of any one prime mover, power converter, or energy storage device.
(3) Essential loads after failure of—
(i) Any one engine on two-engine airplanes; and
(ii) Any two engines on airplanes with three or more engines.
(4) Essential loads for which an alternate source of power is required, after any failure or malfunction in any one power supply system, distribution system, or other utilization system.
(b) In determining compliance with paragraphs (a)(2) and (3) of this section, the power loads may be assumed to be reduced under a monitoring procedure consistent with safety in the kinds of operation authorized. Loads not required in controlled flight need not be considered for the two-engine-inoperative condition on airplanes with three or more engines.
§ 25.1316   Electrical and electronic system lightning protection.
(a) Each electrical and electronic system that performs a function, for which failure would prevent the continued safe flight and landing of the airplane, must be designed and installed so that—
(1) The function is not adversely affected during and after the time the airplane is exposed to lightning; and
(2) The system automatically recovers normal operation of that function in a timely manner after the airplane is exposed to lightning.
(b) Each electrical and electronic system that performs a function, for which failure would reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse operating condition, must be designed and installed so that the function recovers normal operation in a timely manner after the airplane is exposed to lightning.
§ 25.1317   High-intensity Radiated Fields (HIRF) Protection.
(a) Except as provided in paragraph (d) of this section, each electrical and electronic system that performs a function whose failure would prevent the continued safe flight and landing of the airplane must be designed and installed so that—
(1) The function is not adversely affected during and after the time the airplane is exposed to HIRF environment I, as described in appendix L to this part;
(2) The system automatically recovers normal operation of that function, in a timely manner, after the airplane is exposed to HIRF environment I, as described in appendix L to this part, unless the system's recovery conflicts with other operational or functional requirements of the system; and
(3) The system is not adversely affected during and after the time the airplane is exposed to HIRF environment II, as described in appendix L to this part.
(b) Each electrical and electronic system that performs a function whose failure would significantly reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 1 or 2, as described in appendix L to this part.
(c) Each electrical and electronic system that performs a function whose failure would reduce the capability of the airplane or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing the function is exposed to equipment HIRF test level 3, as described in appendix L to this part.
(d) Before December 1, 2012, an electrical or electronic system that performs a function whose failure would prevent the continued safe flight and landing of an airplane may be designed and installed without meeting the provisions of paragraph (a) provided—
(1) The system has previously been shown to comply with special conditions for HIRF, prescribed under § 21.16, issued before December 1, 2007;
(2) The HIRF immunity characteristics of the system have not changed since compliance with the special conditions was demonstrated; and
(3) The data used to demonstrate compliance with the special conditions is provided.
§ 25.1321   Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any pilot must be plainly visible to him from his station with the minimum practicable deviation from his normal position and line of vision when he is looking forward along the flight path.
(b) The flight instruments required by § 25.1303 must be grouped on the instrument panel and centered as nearly as practicable about the vertical plane of the pilot's forward vision. In addition—
(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 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; and
(4) The instrument that most effectively indicates direction of flight must be adjacent to and directly below the instrument in the top center position.
(c) Required powerplant instruments must be closely grouped on the instrument panel. In addition—
(1) The location of identical powerplant instruments for the engines must prevent confusion as to which engine each instrument relates; and
(2) Powerplant instruments vital to the safe operation of the airplane must be plainly visible to the appropriate crewmembers.
(d) Instrument panel vibration may not damage or impair the accuracy of any instrument.
(e) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.
§ 25.1322   Flightcrew alerting.
(a) Flightcrew alerts must:
(1) Provide the flightcrew with the information needed to:
(i) Identify non-normal operation or airplane system conditions, and
(ii) Determine the appropriate actions, if any.
(2) Be readily and easily detectable and intelligible by the flightcrew under all foreseeable operating conditions, including conditions where multiple alerts are provided.
(3) Be removed when the alerting condition no longer exists.
(b) Alerts must conform to the following prioritization hierarchy based on the urgency of flightcrew awareness and response.
(1) Warning: For conditions that require immediate flightcrew awareness and immediate flightcrew response.
(2) Caution: For conditions that require immediate flightcrew awareness and subsequent flightcrew response.
(3) Advisory: For conditions that require flightcrew awareness and may require subsequent flightcrew response.
(c) Warning and caution alerts must:
(1) Be prioritized within each category, when necessary.
(2) Provide timely attention-getting cues through at least two different senses by a combination of aural, visual, or tactile indications.
(3) Permit each occurrence of the attention-getting cues required by paragraph (c)(2) of this section to be acknowledged and suppressed, unless they are required to be continuous.
(d) The alert function must be designed to minimize the effects of false and nuisance alerts. In particular, it must be designed to:
(1) Prevent the presentation of an alert that is inappropriate or unnecessary.
(2) Provide a means to suppress an attention-getting component of an alert caused by a failure of the alerting function that interferes with the flightcrew's ability to safely operate the airplane. This means must not be readily available to the flightcrew so that it could be operated inadvertently or by habitual reflexive action. When an alert is suppressed, there must be a clear and unmistakable annunciation to the flightcrew that the alert has been suppressed.
(e) Visual alert indications must:
(1) Conform to the following color convention:
(i) Red for warning alert indications.
(ii) Amber or yellow for caution alert indications.
(iii) Any color except red or green for advisory alert indications.
(2) Use visual coding techniques, together with other alerting function elements on the flight deck, to distinguish between warning, caution, and advisory alert indications, if they are presented on monochromatic displays that are not capable of conforming to the color convention in paragraph (e)(1) of this section.
(f) Use of the colors red, amber, and yellow on the flight deck for functions other than flightcrew alerting must be limited and must not adversely affect flightcrew alerting.
§ 25.1323   Airspeed indicating system.
For each airspeed indicating system, the following apply:
(a) Each airspeed indicating instrument must be approved and 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 the system error (that is, the relation between IAS and CAS) in flight and during the accelerated takeoff ground run. The ground run calibration must be determined—
(1) From 0.8 of the minimum value of V 1 to the maximum value of V 2 , considering the approved ranges of altitude and weight; and
(2) With the flaps and power settings corresponding to the values determined in the establishment of the takeoff path under § 25.111 assuming that the critical engine fails at the minimum value of V 1 .
(c) The airspeed error of the installation, excluding the airspeed indicator instrument calibration error, may not exceed three percent or five knots, whichever is greater, throughout the speed range, from—
(1) V MO to 1.23 V SR 1 , with flaps retracted; and
(2) 1.23 V SR 0 to V FE with flaps in the landing position.
(d) From 1.23 V SR to the speed at which stall warning begins, the IAS must change perceptibly with CAS and in the same sense, and at speeds below stall warning speed the IAS must not change in an incorrect sense.
(e) From V MO to V MO + 2/3 (V DF − V MO ), the IAS must change perceptibly with CAS and in the same sense, and at higher speeds up to V DF the IAS must not change in an incorrect sense.
(f) There must be no indication of airspeed that would cause undue difficulty to the pilot during the takeoff between the initiation of rotation and the achievement of a steady climbing condition.
(g) The effects of airspeed indicating system lag may not introduce significant takeoff indicated airspeed bias, or significant errors in takeoff or accelerate-stop distances.
(h) 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.
(i) Each system must have a heated pitot tube or an equivalent means of preventing malfunction in the heavy rain conditions defined in Table 1 of this section; mixed phase and ice crystal conditions as defined in part 33, Appendix D, of this chapter; the icing conditions defined in Appendix C of this part; and the following icing conditions specified in Appendix O of this part:
(1) For airplanes certificated in accordance with § 25.1420(a)(1), the icing conditions that the airplane is certified to safely exit following detection.
(2) For airplanes certificated in accordance with § 25.1420(a)(2), the icing conditions that the airplane is certified to safely operate in and the icing conditions that the airplane is certified to safely exit following detection.
(3) For airplanes certificated in accordance with § 25.1420(a)(3) and for airplanes not subject to § 25.1420, all icing conditions.
Table 1—Heavy Rain Conditions for Airspeed Indicating System Tests
(j) 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.
§ 25.1324   Angle of attack system.
Each angle of attack system sensor must be heated or have an equivalent means of preventing malfunction in the heavy rain conditions defined in Table 1 of § 25.1323, the mixed phase and ice crystal conditions as defined in part 33, Appendix D, of this chapter, the icing conditions defined in Appendix C of this part, and the following icing conditions specified in Appendix O of this part:
(a) For airplanes certificated in accordance with § 25.1420(a)(1), the icing conditions that the airplane is certified to safely exit following detection.
(b) For airplanes certificated in accordance with § 25.1420(a)(2), the icing conditions that the airplane is certified to safely operate in and the icing conditions that the airplane is certified to safely exit following detection.
(c) For airplanes certificated in accordance with § 25.1420(a)(3) and for airplanes not subject to § 25.1420, all icing conditions.
§ 25.1325   Static pressure systems.
(a) Each instrument with static air case connections must be vented to the outside atmosphere through an appropriate piping system.
(b) Each static port must be designed and located so that:
(1) The static pressure system performance is least affected by airflow variation, or by moisture or other foreign matter; and
(2) The correlation between air pressure in the static pressure system and true ambient atmospheric static pressure is not changed when the airplane is exposed to the icing conditions defined in Appendix C of this part, and the following icing conditions specified in Appendix O of this part:
(i) For airplanes certificated in accordance with § 25.1420(a)(1), the icing conditions that the airplane is certified to safely exit following detection.
(ii) For airplanes certificated in accordance with § 25.1420(a)(2), the icing conditions that the airplane is certified to safely operate in and the icing conditions that the airplane is certified to safely exit following detection.
(iii) For airplanes certificated in accordance with § 25.1420(a)(3) and for airplanes not subject to § 25.1420, all icing conditions.
(c) The design and installation of the static pressure system must be such that—
(1) Positive drainage of moisture is provided; chafing of the tubing and excessive distortion or restriction at bends in the tubing is avoided; and the materials used are durable, suitable for the purpose intended, and protected against corrosion; and
(2) It is airtight except for the port into the atmosphere. A proof test must be conducted to demonstrate the 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 airplane 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.
(d) Each pressure altimeter must be approved and must be calibrated to indicate pressure altitude in a standard atmosphere, with a minimum practicable calibration error when the corresponding static pressures are applied.
(e) Each system must be designed and installed so that the error in indicated pressure altitude, at sea level, with a standard atmosphere, excluding instrument calibration error, does not result in an error of more than ±30 feet per 100 knots speed for the appropriate configuration in the speed range between 1.23 V SR 0 with flaps extended and 1.7 V SR 1 with flaps retracted. However, the error need not be less than ±30 feet.
(f) If an altimeter system is fitted with a device that provides corrections to the altimeter indication, the device must be designed and installed in such manner that it can be bypassed when it malfunctions, unless an alternate altimeter system is provided. Each correction device must be fitted with a means for indicating the occurrence of reasonably probable malfunctions, including power failure, to the flight crew. The indicating means must be effective for any cockpit lighting condition likely to occur.
(g) Except as provided in paragraph (h) of this section, if the static pressure system incorporates both a primary and an alternate static pressure source, the means for selecting one or the other source must be designed so that—
(1) When either source is selected, the other is blocked off; and
(2) Both sources cannot be blocked off simultaneously.
(h) For unpressurized airplanes, 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.
§ 25.1326   Pitot heat indication systems.
If a flight instrument pitot heating system is installed, 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 flight crewmember.
(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.
§ 25.1327   Magnetic direction indicator.
(a) Each magnetic direction indicator must be installed so that its accuracy is not excessively affected by the airplane's vibration or magnetic fields.
(b) The compensated installation may not have a deviation, in level flight, greater than 10 degrees on any heading.
§ 25.1329   Flight guidance system.
(a) Quick disengagement controls for the autopilot and autothrust functions must be provided for each pilot. The autopilot quick disengagement controls must be located on both control wheels (or equivalent). The autothrust quick disengagement controls must be located on the thrust control levers. Quick disengagement controls must be readily accessible to each pilot while operating the control wheel (or equivalent) and thrust control levers.
(b) The effects of a failure of the system to disengage the autopilot or autothrust functions when manually commanded by the pilot must be assessed in accordance with the requirements of § 25.1309.
(c) Engagement or switching of the flight guidance system, a mode, or a sensor may not cause a transient response of the airplane's flight path any greater than a minor transient, as defined in paragraph (n)(1) of this section.
(d) Under normal conditions, the disengagement of any automatic control function of a flight guidance system may not cause a transient response of the airplane's flight path any greater than a minor transient.
(e) Under rare normal and non-normal conditions, disengagement of any automatic control function of a flight guidance system may not result in a transient any greater than a significant transient, as defined in paragraph (n)(2) of this section.
(f) The function and direction of motion of each command reference control, such as heading select or vertical speed, must be plainly indicated on, or adjacent to, each control if necessary to prevent inappropriate use or confusion.
(g) Under any condition of flight appropriate to its use, the flight guidance system may not produce hazardous loads on the airplane, nor create hazardous deviations in the flight path. This applies to both fault-free operation and in the event of a malfunction, and assumes that the pilot begins corrective action within a reasonable period of time.
(h) When the flight guidance system is in use, a means must be provided to avoid excursions beyond an acceptable margin from the speed range of the normal flight envelope. If the airplane experiences an excursion outside this range, a means must be provided to prevent the flight guidance system from providing guidance or control to an unsafe speed.
(i) The flight guidance system functions, controls, indications, and alerts must be designed to minimize flightcrew errors and confusion concerning the behavior and operation of the flight guidance system. Means must be provided to indicate the current mode of operation, including any armed modes, transitions, and reversions. Selector switch position is not an acceptable means of indication. The controls and indications must be grouped and presented in a logical and consistent manner. The indications must be visible to each pilot under all expected lighting conditions.
(j) Following disengagement of the autopilot, a warning (visual and auditory) must be provided to each pilot and be timely and distinct from all other cockpit warnings.
(k) Following disengagement of the autothrust function, a caution must be provided to each pilot.
(l) The autopilot may not create a potential hazard when the flightcrew applies an override force to the flight controls.
(m) During autothrust operation, it must be possible for the flightcrew to move the thrust levers without requiring excessive force. The autothrust may not create a potential hazard when the flightcrew applies an override force to the thrust levers.
(n) For purposes of this section, a transient is a disturbance in the control or flight path of the airplane that is not consistent with response to flightcrew inputs or environmental conditions.
(1) A minor transient would not significantly reduce safety margins and would involve flightcrew actions that are well within their capabilities. A minor transient may involve a slight increase in flightcrew workload or some physical discomfort to passengers or cabin crew.
(2) A significant transient may lead to a significant reduction in safety margins, an increase in flightcrew workload, discomfort to the flightcrew, or physical distress to the passengers or cabin crew, possibly including non-fatal injuries. Significant transients do not require, in order to remain within or recover to the normal flight envelope, any of the following:
(i) Exceptional piloting skill, alertness, or strength.
(ii) Forces applied by the pilot which are greater than those specified in § 25.143(c).
(iii) Accelerations or attitudes in the airplane that might result in further hazard to secured or non-secured occupants.
§ 25.1331   Instruments using a power supply.
(a) For each instrument required by § 25.1303(b) that uses a power supply, the following apply:
(1) Each instrument must have a visual means integral with, the instrument, to indicate when 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 instruments. For electric instruments, the power is considered to be adequate when the voltage is within approved limits.
(2) Each instrument must, in the event of the failure of one power source, be supplied by another power source. This may be accomplished automatically or by manual means.
(3) If an instrument presenting navigation data receives information from sources external to that instrument and loss of that information would render the presented data unreliable, the instrument must incorporate a visual means to warn the crew, when such loss of information occurs, that the presented data should not be relied upon.
(b) As used in this section, “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).
§ 25.1333   Instrument systems.
For systems that operate the instruments required by § 25.1303(b) which are located at each pilot's station—
(a) Means must be provided to connect the required instruments at the first pilot's station to operating systems which are independent of the operating systems at other flight crew stations, or other equipment;
(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 instruments, including attitude, direction, airspeed, and altitude will remain available to the pilots, without additional crewmember action, after any single failure or combination of failures that is not shown to be extremely improbable; and
(c) Additional instruments, systems, or equipment may not be connected to the operating systems for the required instruments, unless provisions are made to ensure the continued normal functioning of the required instruments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable.
§ 25.1337   Powerplant instruments.
(a) Instruments and instrument lines. (1) Each powerplant and auxiliary power unit instrument line must meet the requirements of §§ 25.993 and 25.1183.
(2) Each line carrying flammable fluids under pressure must—
(i) Have restricting orifices or other safety devices at the source of pressure to prevent the escape of excessive fluid if the line fails; and
(ii) Be installed and located so that the escape of fluids would not create a hazard.
(3) Each powerplant and auxiliary power unit instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard.
(b) Fuel quantity indicator. There must be means to indicate to the flight crewmembers, the quantity, in gallons or equivalent units, of usable fuel in each tank during flight. In addition—
(1) Each fuel quantity indicator must be calibrated to read “zero” during level flight when the quantity of fuel remaining in the tank is equal to the unusable fuel supply determined under § 25.959;
(2) Tanks with interconnected outlets and airspaces may be treated as one tank and need not have separate indicators; and
(3) 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 in each tank. If an oil transfer or reserve oil supply system is installed, there must be a means to indicate to the flight crew, in flight, the quantity of oil in each tank.
(e) Turbopropeller blade position indicator. Required turbopropeller blade position indicators must begin indicating before the blade moves more than eight degrees below the flight low pitch stop. The source of indication must directly sense the blade position.
(f) Fuel pressure indicator. There must be means to measure fuel pressure, in each system supplying reciprocating engines, at a point downstream of any fuel pump except fuel injection pumps. In addition—
(1) If necessary for the maintenance of proper fuel delivery pressure, there must be a connection to transmit the carburetor air intake static pressure to the proper pump relief valve connection; and
(2) If a connection is required under paragraph (f)(1) of this section, the gauge balance lines must be independently connected to the carburetor inlet pressure to avoid erroneous readings.
§ 25.1351   General.
(a) Electrical system capacity. The required generating capacity, and number and kinds of power sources must—
(1) Be determined by an electrical load analysis; and
(2) Meet the requirements of § 25.1309.
(b) Generating system. The generating system includes electrical power sources, main power busses, transmission cables, and associated control, regulation, and protective devices. It must be designed so that—
(1) Power sources function properly when independent and when connected in combination;
(2) No failure or malfunction of any power source can create a hazard or impair the ability of remaining sources to supply essential loads;
(3) The system voltage and frequency (as applicable) at the terminals of all essential load equipment can be maintained within the limits for which the equipment is designed, during any probable operating condition; and
(4) System transients due to switching, fault clearing, or other causes do not make essential loads inoperative, and do not cause a smoke or fire hazard.
(5) There are means accessible, in flight, to appropriate crewmembers for the individual and collective disconnection of the electrical power sources from the system.
(6) There are means to indicate to appropriate crewmembers the generating system quantities essential for the safe operation of the system, such as the voltage and current supplied by each generator.
(c) External power. If provisions are made for connecting external power to the airplane, and that external power 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.
(d) Operation without normal electrical power. It must be shown by analysis, tests, or both, that the airplane can be operated safely in VFR conditions, for a period of not less than five minutes, with the normal electrical power (electrical power sources excluding the battery) 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 both 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.
§ 25.1353   Electrical equipment and installations.
(a) Electrical equipment and controls must be installed so that operation of any one unit or system of units will not adversely affect the simultaneous operation of any other electrical unit or system essential to safe operation. Any electrical interference likely to be present in the airplane must not result in hazardous effects on the airplane or its systems.
(b) Storage batteries must be designed and installed as follows:
(1) Safe cell temperatures and pressures must be maintained during any probable charging or discharging condition. No uncontrolled increase in cell temperature may result when the battery is recharged (after previous complete discharge)—
(i) At maximum regulated voltage or power;
(ii) During a flight of maximum duration; and
(iii) Under the most adverse cooling condition likely to occur in service.
(2) Compliance with paragraph (b)(1) of this section must be shown by test unless experience with similar batteries and installations has shown that maintaining safe cell temperatures and pressures presents no problem.
(3) 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.
(4) No corrosive fluids or gases that may escape from the battery may damage surrounding airplane structures or adjacent essential equipment.
(5) Each nickel cadmium battery installation 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 individual cells.
(6) Nickel cadmium battery installations must have—
(i) A system to control the charging rate of the battery automatically so as to prevent battery overheating;
(ii) A battery temperature sensing and over-temperature warning system with a means for disconnecting the battery from its charging source in the event of an over-temperature condition; or
(iii) A battery failure sensing and warning system with a means for disconnecting the battery from its charging source in the event of battery failure.
(c) Electrical bonding must provide an adequate electrical return path under both normal and fault conditions, on airplanes having grounded electrical systems.
§ 25.1355   Distribution system.
(a) The distribution system includes the distribution busses, their associated feeders, and each control and protective device.
(b) [Reserved]
(c) If two independent sources of electrical power for particular equipment or systems are required by this chapter, in the event of the failure of one power source for such equipment or system, another power source (including its separate feeder) must be automatically provided or be manually selectable to maintain equipment or system operation.
§ 25.1357   Circuit protective devices.
(a) Automatic protective devices must be used to minimize distress to the electrical system and hazard to the airplane in the event of wiring faults or serious malfunction of the system or connected equipment.
(b) The protective and control devices in the generating system must be designed to de-energize and disconnect faulty power sources and power transmission equipment from their associated busses with sufficient rapidity to provide protection from hazardous over-voltage and other malfunctioning.
(c) Each resettable circuit protective device must be designed so that, when an overload or circuit fault exists, it will open the circuit irrespective of the position of the operating control.
(d) If the ability to reset a circuit breaker or replace a fuse is essential to safety in flight, that circuit breaker or fuse must be located and identified so that it can be readily reset or replaced in flight. Where fuses are used, there must be spare fuses for use in flight equal to at least 50% of the number of fuses of each rating required for complete circuit protection.
(e) Each circuit for essential loads must have individual circuit protection. However, individual protection for each circuit in an essential load system (such as each position light circuit in a system) is not required.
(f) For airplane systems for which the ability to remove or reset power during normal operations is necessary, the system must be designed so that circuit breakers are not the primary means to remove or reset system power unless specifically designed for use as a switch.
(g) Automatic reset circuit breakers may be used as integral protectors for electrical equipment (such as thermal cut-outs) if there is circuit protection to protect the cable to the equipment.
§ 25.1360   Precautions against injury.
(a) Shock. The electrical system must be designed to minimize risk of electric shock to crew, passengers, and servicing personnel and to maintenance personnel using normal precautions.
(b) Burns. The temperature of any part that may be handled by a crewmember during normal operations must not cause dangerous inadvertent movement by the crewmember or injury to the crewmember.
§ 25.1362   Electrical supplies for emergency conditions.
A suitable electrical supply must be provided to those services required for emergency procedures after an emergency landing or ditching. The circuits for these services must be designed, protected, and installed so that the risk of the services being rendered ineffective under these emergency conditions is minimized.
§ 25.1363   Electrical system tests.
(a) When laboratory tests of the electrical system are conducted—
(1) The tests must be performed on a mock-up using the same generating equipment used in the airplane;
(2) The equipment must simulate the electrical characteristics of the distribution wiring and connected loads to the extent necessary for valid test results; and
(3) Laboratory generator drives must simulate the actual prime movers on the airplane with respect to their reaction to generator loading, including loading due to faults.
(b) For each flight condition that cannot be simulated adequately in the laboratory or by ground tests on the airplane, flight tests must be made.
§ 25.1365   Electrical appliances, motors, and transformers.
(a) An applicant must show that, in the event of a failure of the electrical supply or control system, the design and installation of domestic appliances meet the requirements of § 25.1309(b) and (c). Domestic appliances are items such as cooktops, ovens, coffee makers, water heaters, refrigerators, and toilet flush systems that are placed on the airplane to provide service amenities to passengers.
(b) Galleys and cooking appliances must be installed in a way that minimizes risk of overheat or fire.
(c) Domestic appliances, particularly those in galley areas, must be installed or protected so as to prevent damage or contamination of other equipment or systems from fluids or vapors which may be present during normal operation or as a result of spillage, if such damage or contamination could create a hazardous condition.
(d) Unless compliance with § 25.1309(b) is provided by the circuit protective device required by § 25.1357(a), electric motors and transformers, including those installed in domestic systems, must have a suitable thermal protection device to prevent overheating under normal operation and failure conditions, if overheating could create a smoke or fire hazard.
§ 25.1381   Instrument lights.
(a) The instrument lights must—
(1) Provide sufficient illumination to make each instrument, switch and other device necessary for safe operation easily readable unless sufficient illumination is available from another source; and
(2) Be installed so that—
(i) Their direct rays are shielded from the pilot's eyes; and
(ii) No objectionable reflections are visible to the pilot.
(b) Unless undimmed instrument lights are satisfactory under each expected flight condition, there must be a means to control the intensity of illumination.
§ 25.1383   Landing lights.
(a) Each landing light must be approved, and must be installed so that—
(1) No objectionable glare is visible to the pilot;
(2) The pilot is not adversely affected by halation; and
(3) It provides enough light for night landing.
(b) Except when one switch is used for the lights of a multiple light installation at one location, there must be a separate switch for each light.
(c) There must be a means to indicate to the pilots when the landing lights are extended.
§ 25.1385   Position light system 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 §§ 25.1387 through 25.1397.
(b) Forward position lights. Forward position lights must consist of a red and a green light spaced laterally as far apart as practicable and installed forward on the airplane so that, with the airplane in the normal flying position, the red light is on the left side and the green light is on the right side. Each light must be approved.
(c) Rear position light. The rear position light must be a white light mounted as far aft as practicable on the tail or on each wing tip, and must be approved.
(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.
§ 25.1387   Position light system dihedral angles.
(a) Except as provided in paragraph (e) of this section, each forward and rear 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 § 25.1385(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.
§ 25.1389   Position light distribution and intensities.
(a) General. The intensities prescribed in this section must be provided by new equipment with light covers and color filters in place. Intensities must be 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) Forward and rear position lights. The light distribution and intensities of forward and rear 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 § 25.1391.
(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 § 25.1393, where I is the minimum intensity prescribed in § 25.1391 for the corresponding angles in the horizontal plane.
(3) Intensities in overlaps between adjacent signals. No intensity in any overlap between adjacent signals may exceed the values given in § 25.1395, except that higher intensities in overlaps may be used with main beam intensities substantially greater than the minima specified in §§ 25.1391 and 25.1393 if the overlap intensities in relation to the main beam intensities do not adversely affect signal clarity. When the peak intensity of the forward position lights is more than 100 candles, the maximum overlap intensities between them may exceed the values given in § 25.1395 if the overlap intensity in Area A is not more than 10 percent of peak position light intensity and the overlap intensity in Area B is not greater than 2.5 percent of peak position light intensity.
§ 25.1391   Minimum intensities in the horizontal plane of forward and rear position lights.
Each position light intensity must equal or exceed the applicable values in the following table:
§ 25.1393   Minimum intensities in any vertical plane of forward and rear position lights.
Each position light intensity must equal or exceed the applicable values in the following table:
§ 25.1395   Maximum intensities in overlapping beams of forward and rear position lights.
No position light intensity may exceed the applicable values in the following table, except as provided in § 25.1389(b)(3).
(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.
§ 25.1397   Color specifications.
Each position light color must have the applicable International Commission on Illumination chromaticity coordinates as follows:
(a) Aviation red —
(b) Aviation green —
(c) Aviation white —
§ 25.1399   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 2 nautical 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.
§ 25.1401   Anticollision light system.
(a) General. The airplane must have an anticollision light system that—
(1) Consists of one or more approved anticollision lights located so that their light will not impair the crew's vision or detract from the conspicuity of the position lights; and
(2) Meets the requirements of 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 a solid angle or angles of obstructed visibility totaling not more than 0.03 steradians is allowable within a solid angle equal to 0.15 steradians centered about the longitudinal axis in the rearward direction.
(c) Flashing characteristics. The 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 § 25.1397.
(e) Light intensity. The minimum light intensities in all vertical planes, measured with the red filter (if used) and expressed in terms of “effective” intensities, must meet the requirements of paragraph (f) of this section. The following relation must be assumed:
(f) Minimum effective intensities for anticollision lights. Each anticollision light effective intensity must equal or exceed the applicable values in the following table.
§ 25.1403   Wing icing detection lights.
Unless operations at night in known or forecast icing conditions are prohibited by an operating limitation, a means must be provided for illuminating or otherwise determining the formation of ice on the parts of the wings that are critical from the standpoint of ice accumulation. 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.
§ 25.1411   General.
(a) Accessibility. Required safety equipment to be used by the crew in an emergency must be readily accessible.
(b) Stowage provisions. Stowage provisions for required emergency 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 inadvertent damage.
(c) Emergency exit descent device. The stowage provisions for the emergency exit descent devices required by § 25.810(a) must be at each exit for which they are intended.
(d) Liferafts. (1) The stowage provisions for the liferafts described in § 25.1415 must accommodate enough rafts for the maximum number of occupants for which certification for ditching is requested.
(2) Liferafts must be stowed near exits through which the rafts can be launched during an unplanned ditching.
(3) Rafts automatically or remotely released outside the airplane must be attached to the airplane by means of the static line prescribed in § 25.1415.
(4) The stowage provisions for each portable liferaft must allow rapid detachment and removal of the raft for use at other than the intended exits.
(e) Long-range signaling device. The stowage provisions for the long-range signaling device required by § 25.1415 must be near an exit available during an unplanned ditching.
(f) Life preserver stowage provisions. The stowage provisions for life preservers described in § 25.1415 must accommodate one life preserver for each occupant for which certification for ditching is requested. Each life preserver must be within easy reach of each seated occupant.
(g) Life line stowage provisions. If certification for ditching under § 25.801 is requested, there must be provisions to store life lines. These provisions must—
(1) Allow one life line to be attached to each side of the fuselage; and
(2) Be arranged to allow the life lines to be used to enable the occupants to stay on the wing after ditching.
§ 25.1415   Ditching equipment.
(a) Ditching equipment used in airplanes to be certificated for ditching under § 25.801, and required by the operating rules of this chapter, must meet the requirements of this section.
(b) Each liferaft and each life preserver must be approved. In addition—
(1) Unless excess rafts of enough capacity are provided, the buoyancy and seating capacity beyond the rated capacity of the rafts must accommodate all occupants of the airplane in the event of a loss of one raft of the largest rated capacity; and
(2) Each raft must have a trailing line, and must have a static line designed to hold the raft near the airplane but to release it if the airplane becomes totally submerged.
(c) Approved survival equipment must be attached to each liferaft.
(d) There must be an approved survival type emergency locator transmitter for use in one life raft.
(e) For airplanes not certificated for ditching under § 25.801 and not having approved life preservers, there must be an approved flotation means for each occupant. This means must be within easy reach of each seated occupant and must be readily removable from the airplane.
§ 25.1419   Ice protection.
If the applicant seeks certification for flight in icing conditions, the airplane must be able to safely operate in the continuous maximum and intermittent maximum icing conditions of appendix C. To establish this—
(a) An analysis must be performed to establish that the ice protection for the various components of the airplane is adequate, taking into account the various airplane operational configurations; and
(b) To verify the ice protection analysis, to check for icing anomalies, and to demonstrate that the ice protection system and its components are effective, the airplane or its components must be flight tested in the various operational configurations, in measured natural atmospheric icing conditions and, as found necessary, by one or more of the following means:
(1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components.
(2) Flight dry air tests of the ice protection system as a whole, or of its individual components.
(3) Flight tests of the airplane or its components in measured simulated icing conditions.
(c) Caution information, such as an amber caution light or equivalent, must be provided to alert the flightcrew when the anti-ice or de-ice system is not functioning normally.
(d) For turbine engine powered airplanes, the ice protection provisions of this section are considered to be applicable primarily to the airframe. For the powerplant installation, certain additional provisions of subpart E of this part may be found applicable.
(e) One of the following methods of icing detection and activation of the airframe ice protection system must be provided:
(1) A primary ice detection system that automatically activates or alerts the flightcrew to activate the airframe ice protection system;
(2) A definition of visual cues for recognition of the first sign of ice accretion on a specified surface combined with an advisory ice detection system that alerts the flightcrew to activate the airframe ice protection system; or
(3) Identification of conditions conducive to airframe icing as defined by an appropriate static or total air temperature and visible moisture for use by the flightcrew to activate the airframe ice protection system.
(f) Unless the applicant shows that the airframe ice protection system need not be operated during specific phases of flight, the requirements of paragraph (e) of this section are applicable to all phases of flight.
(g) After the initial activation of the airframe ice protection system—
(1) The ice protection system must be designed to operate continuously;
(2) The airplane must be equipped with a system that automatically cycles the ice protection system; or
(3) An ice detection system must be provided to alert the flightcrew each time the ice protection system must be cycled.
(h) Procedures for operation of the ice protection system, including activation and deactivation, must be established and documented in the Airplane Flight Manual.
§ 25.1420   Supercooled large drop icing conditions.
(a) If certification for flight in icing conditions is sought, in addition to the requirements of § 25.1419, an airplane with a maximum takeoff weight less than 60,000 pounds or with reversible flight controls must be capable of operating in accordance with paragraphs (a)(1), (2), or (3), of this section.
(1) Operating safely after encountering the icing conditions defined in Appendix O of this part:
(i) The airplane must have a means to detect that it is operating in Appendix O icing conditions; and
(ii) Following detection of Appendix O icing conditions, the airplane must be capable of operating safely while exiting all icing conditions.
(2) Operating safely in a portion of the icing conditions defined in Appendix O of this part as selected by the applicant:
(i) The airplane must have a means to detect that it is operating in conditions that exceed the selected portion of Appendix O icing conditions; and
(ii) Following detection, the airplane must be capable of operating safely while exiting all icing conditions.
(3) Operating safely in the icing conditions defined in Appendix O of this part.
(b) To establish that the airplane can operate safely as required in paragraph (a) of this section, an applicant must show through analysis that the ice protection for the various components of the airplane is adequate, taking into account the various airplane operational configurations. To verify the analysis, one, or more as found necessary, of the following methods must be used:
(1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components.
(2) Laboratory dry air or simulated icing tests, or a combination of both, of models of the airplane.
(3) Flight tests of the airplane or its components in simulated icing conditions, measured as necessary to support the analysis.
(4) Flight tests of the airplane with simulated ice shapes.
(5) Flight tests of the airplane in natural icing conditions, measured as necessary to support the analysis.
(c) For an airplane certified in accordance with paragraph (a)(2) or (3) of this section, the requirements of § 25.1419(e), (f), (g), and (h) must be met for the icing conditions defined in Appendix O of this part in which the airplane is certified to operate.
(d) For the purposes of this section, the following definitions apply:
(1) Reversible Flight Controls. Flight controls in the normal operating configuration that have force or motion originating at the airplane's control surface (for example, through aerodynamic loads, static imbalance, or trim or servo tab inputs) that is transmitted back to flight deck controls. This term refers to flight deck controls connected to the pitch, roll, or yaw control surfaces by direct mechanical linkages, cables, or push-pull rods in such a way that pilot effort produces motion or force about the hinge line.
(2) Simulated Icing Test. Testing conducted in simulated icing conditions, such as in an icing tunnel or behind an icing tanker.
(3) Simulated Ice Shape. Ice shape fabricated from wood, epoxy, or other materials by any construction technique.
§ 25.1421   Megaphones.
If a megaphone is installed, a restraining means must be provided that is capable of restraining the megaphone when it is subjected to the ultimate inertia forces specified in § 25.561(b)(3).
§ 25.1423   Public address system.
A public address system required by this chapter must—
(a) Be powerable when the aircraft is in flight or stopped on the ground, after the shutdown or failure of all engines and auxiliary power units, or the disconnection or failure of all power sources dependent on their continued operation, for—
(1) A time duration of at least 10 minutes, including an aggregate time duration of at least 5 minutes of announcements made by flight and cabin crewmembers, considering all other loads which may remain powered by the same source when all other power sources are inoperative; and
(2) An additional time duration in its standby state appropriate or required for any other loads that are powered by the same source and that are essential to safety of flight or required during emergency conditions.
(b) Be capable of operation within 3 seconds from the time a microphone is removed from its stowage.
(c) Be intelligible at all passenger seats, lavatories, and flight attendant seats and work stations.
(d) Be designed so that no unused, unstowed microphone will render the system inoperative.
(e) Be capable of functioning independently of any required crewmember interphone system.
(f) Be accessible for immediate use from each of two flight crewmember stations in the pilot compartment.
(g) For each required floor-level passenger emergency exit which has an adjacent flight attendant seat, have a microphone which is readily accessible to the seated flight attendant, except that one microphone may serve more than one exit, provided the proximity of the exits allows unassisted verbal communication between seated flight attendants.
§ 25.1431   Electronic equipment.
(a) In showing compliance with § 25.1309 (a) and (b) with respect to radio and electronic equipment and their installations, critical environmental conditions must be considered.
(b) Radio and electronic equipment must be supplied with power under the requirements of § 25.1355(c).
(c) Radio and electronic equipment, controls, and wiring must be installed so that operation of any one unit or system of units will not adversely affect the simultaneous operation of any other radio or electronic unit, or system of units, required by this chapter.
(d) Electronic equipment must be designed and installed such that it does not cause essential loads to become inoperative as a result of electrical power supply transients or transients from other causes.
§ 25.1433   Vacuum systems.
There must be means, in addition to the normal pressure relief, to automatically relieve the pressure in the discharge lines from the vacuum air pump when the delivery temperature of the air becomes unsafe.
§ 25.1435   Hydraulic systems.
(a) Element design. Each element of the hydraulic system must be designed to:
(1) Withstand the proof pressure without permanent deformation that would prevent it from performing its intended functions, and the ultimate pressure without rupture. The proof and ultimate pressures are defined in terms of the design operating pressure (DOP) as follows:
(2) Withstand, without deformation that would prevent it from performing its intended function, the design operating pressure in combination with limit structural loads that may be imposed;
(3) Withstand, without rupture, the design operating pressure multiplied by a factor of 1.5 in combination with ultimate structural load that can reasonably occur simultaneously;
(4) Withstand the fatigue effects of all cyclic pressures, including transients, and associated externally induced loads, taking into account the consequences of element failure; and
(5) Perform as intended under all environmental conditions for which the airplane is certificated.
(b) System design. Each hydraulic system must:
(1) Have means located at a flightcrew station to indicate appropriate system parameters, if
(i) It performs a function necessary for continued safe flight and landing; or
(ii) In the event of hydraulic system malfunction, corrective action by the crew to ensure continued safe flight and landing is necessary;
(2) Have means to ensure that system pressures, including transient pressures and pressures from fluid volumetric changes in elements that are likely to remain closed long enough for such changes to occur, are within the design capabilities of each element, such that they meet the requirements defined in § 25.1435(a)(1) through (a)(5);
(3) Have means to minimize the release of harmful or hazardous concentrations of hydraulic fluid or vapors into the crew and passenger compartments during flight;
(4) Meet the applicable requirements of §§ 25.863, 25.1183, 25.1185, and 25.1189 if a flammable hydraulic fluid is used; and
(5) Be designed to use any suitable hydraulic fluid specified by the airplane manufacturer, which must be identified by appropriate markings as required by § 25.1541.
(c) Tests. Tests must be conducted on the hydraulic system(s), and/or subsystem(s) and elements, except that analysis may be used in place of or to supplement testing, where the analysis is shown to be reliable and appropriate. All internal and external influences must be taken into account to an extent necessary to evaluate their effects, and to assure reliable system and element functioning and integration. Failure or unacceptable deficiency of an element or system must be corrected and be sufficiently retested, where necessary.
(1) The system(s), subsystem(s), or element(s) must be subjected to performance, fatigue, and endurance tests representative of airplane ground and flight operations.
(2) The complete system must be tested to determine proper functional performance and relation to the other systems, including simulation of relevant failure conditions, and to support or validate element design.
(3) The complete hydraulic system(s) must be functionally tested on the airplane in normal operation over the range of motion of all associated user systems. The test must be conducted at the system relief pressure or 1.25 times the DOP if a system pressure relief device is not part of the system design. Clearances between hydraulic system elements and other systems or structural elements must remain adequate and there must be no detrimental effects.
§ 25.1438   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.
§ 25.1439   Protective breathing equipment.
(a) Fixed (stationary, or built in) protective breathing equipment must be installed for the use of the flightcrew, and at least one portable protective breathing equipment shall be located at or near the flight deck for use by a flight crewmember. In addition, portable protective breathing equipment must be installed for the use of appropriate crewmembers for fighting fires in compartments accessible in flight other than the flight deck. This includes isolated compartments and upper and lower lobe galleys, in which crewmember occupancy is permitted during flight. Equipment must be installed for the maximum number of crewmembers expected to be in the area during any operation.
(b) For protective breathing equipment required by paragraph (a) of this section or by the applicable Operating Regulations:
(1) The equipment must be designed to protect the appropriate crewmember from smoke, carbon dioxide, and other harmful gases while on flight deck duty or while combating fires.
(2) The equipment must include—
(i) Masks covering the eyes, nose and mouth, or
(ii) Masks covering the nose and mouth, plus accessory equipment to cover the eyes.
(3) Equipment, including portable equipment, must allow communication with other crewmembers while in use. Equipment available at flightcrew assigned duty stations must also enable the flightcrew to use radio equipment.
(4) The part of the equipment protecting the eyes shall not cause any appreciable adverse effect on vision and must allow corrective glasses to be worn.
(5) The equipment must supply protective oxygen of 15 minutes duration per crewmember at a pressure altitude of 8,000 feet with a respiratory minute volume of 30 liters per minute BTPD. The equipment and system must be designed to prevent any inward leakage to the inside of the device and prevent any outward leakage causing significant increase in the oxygen content of the local ambient atmosphere. If a demand oxygen system is used, a supply of 300 liters of free oxygen at 70 °F. and 760 mm. Hg. pressure is considered to be of 15-minute duration at the prescribed altitude and minute volume. If a continuous flow open circuit protective breathing system is used, a flow rate of 60 liters per minute at 8,000 feet (45 liters per minute at sea level) and a supply of 600 liters of free oxygen at 70 °F. and 760 mm. Hg. pressure is considered to be of 15-minute duration at the prescribed altitude and minute volume. Continuous flow systems must not increase the ambient oxygen content of the local atmosphere above that of demand systems. BTPD refers to body temperature conditions (that is, 37 °C., at ambient pressure, dry).
(6) The equipment must meet the requirements of § 25.1441.
§ 25.1441   Oxygen equipment and supply.
(a) If certification with supplemental oxygen equipment is requested, the equipment must meet the requirements of this section and §§ 25.1443 through 25.1453.
(b) The oxygen system must be free from hazards in itself, in its method of operation, and in its effect upon other components.
(c) There must be a means to allow the crew to readily determine, during flight, the quantity of oxygen available in each source of supply.
(d) The oxygen flow rate and the oxygen equipment for airplanes for which certification for operation above 40,000 feet is requested must be approved.
§ 25.1443   Minimum mass flow of supplemental oxygen.
(a) If continuous flow equipment is installed for use by flight crewmembers, the minimum mass flow of supplemental oxygen required for each crewmember may not be less than the flow required to maintain, during inspiration, a mean tracheal oxygen partial pressure of 149 mm. Hg. when breathing 15 liters per minute, BTPS, and with a maximum tidal volume of 700 cc. with a constant time interval between respirations.
(b) If demand equipment is installed for use by flight crewmembers, the minimum mass flow of supplemental oxygen required for each crewmember may not be less than the flow required to maintain, during inspiration, a mean tracheal oxygen partial pressure of 122 mm. Hg., up to and including a cabin pressure altitude of 35,000 feet, and 95 percent oxygen between cabin pressure altitudes of 35,000 and 40,000 feet, when breathing 20 liters per minute BTPS. In addition, there must be means to allow the crew to use undiluted oxygen at their discretion.
(c) For passengers and cabin attendants, the minimum mass flow of supplemental oxygen required for each person 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:
(1) At cabin pressure altitudes above 10,000 feet up to and including 18,500 feet, a mean tracheal oxygen partial pressure of 100 mm. Hg. when breathing 15 liters per minute, BTPS, and with a tidal volume of 700 cc. with a constant time interval between respirations.
(2) At cabin pressure altitudes above 18,500 feet up to and including 40,000 feet, a mean tracheal oxygen partial pressure of 83.8 mm. Hg. when breathing 30 liters per minute, BTPS, and with a tidal volume of 1,100 cc. with a constant time interval between respirations.
(d) If first-aid oxygen equipment is installed, the minimum mass flow of oxygen to each user may not be less than four liters per minute, STPD. However, there may be a means to decrease this flow to not less than two liters per minute, STPD, at any cabin altitude. The quantity of oxygen required is based upon an average flow rate of three liters per minute per person for whom first-aid oxygen is required.
(e) If portable oxygen equipment is installed for use by crewmembers, the minimum mass flow of supplemental oxygen is the same as specified in paragraph (a) or (b) of this section, whichever is applicable.
§ 25.1445   Equipment standards for the oxygen distributing system.
(a) When oxygen is supplied to both crew and passengers, the distribution system must be designed for either—
(1) A source of supply for the flight crew on duty and a separate source for the passengers and other crewmembers; or
(2) A common source of supply with means to separately reserve the minimum supply required by the flight crew on duty.
(b) Portable walk-around oxygen units of the continuous flow, diluter-demand, and straight demand kinds may be used to meet the crew or passenger breathing requirements.
§ 25.1447   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. Units must be designed to cover the nose and mouth and must be equipped with a suitable means to retain the unit in position on the face. Flight crew masks for supplemental oxygen must have provisions for the use of communication equipment.
(b) If certification for operation up to and including 25,000 feet is requested, an oxygen supply terminal and unit of oxygen dispensing equipment for the immediate use of oxygen by each crewmember must be within easy reach of that crewmember. For any other occupants, the supply terminals and dispensing equipment must be located to allow the use of oxygen as required by the operating rules in this chapter.
(c) If certification for operation above 25,000 feet is requested, there must be oxygen dispensing equipment meeting the following requirements:
(1) There must be an oxygen dispensing unit connected to oxygen supply terminals immediately available to each occupant wherever seated, and at least two oxygen dispensing units connected to oxygen terminals in each lavatory. The total number of dispensing units and outlets in the cabin must exceed the number of seats by at least 10 percent. The extra units must be as uniformly distributed throughout the cabin as practicable. Except as provided in paragraph (c)(5) of this section, if certification for operation above 30,000 feet is requested, the dispensing units providing the required oxygen flow must be automatically presented to the occupants before the cabin pressure altitude exceeds 15,000 feet. The crewmembers must be provided with a manual means of making the dispensing units immediately available in the event of failure of the automatic system.
(2) Each flight crewmember on flight deck duty must be provided with a quick-donning type oxygen dispensing unit connected to an oxygen supply terminal. This dispensing unit must be immediately available to the flight crewmember when seated at his station, and installed so that it:
(i) Can be placed on the face from its ready position, properly secured, sealed, and supplying oxygen upon demand, with one hand, within five seconds and without disturbing eyeglasses or causing delay in proceeding with emergency duties; and
(ii) Allows, while in place, the performance of normal communication functions.
(3) The oxygen dispensing equipment for the flight crewmembers must be:
(i) The diluter demand or pressure demand (pressure demand mask with a diluter demand pressure breathing regulator) type, or other approved oxygen equipment shown to provide the same degree of protection, for airplanes to be operated above 25,000 feet.
(ii) The pressure demand (pressure demand mask with a diluter demand pressure breathing regulator) type with mask-mounted regulator, or other approved oxygen equipment shown to provide the same degree of protection, for airplanes operated at altitudes where decompressions that are not extremely improbable may expose the flightcrew to cabin pressure altitudes in excess of 34,000 feet.
(4) Portable oxygen equipment must be immediately available for each cabin attendant. The portable oxygen equipment must have the oxygen dispensing unit connected to the portable oxygen supply.
(5) When operating into or out of airports with elevations above 13,000 feet, the dispensing units providing the required oxygen flow must be automatically presented to the occupants at cabin pressure altitudes no higher than 2,000 feet above the airplane's maximum takeoff and landing altitude.
§ 25.1449   Means for determining use of oxygen.
There must be a means to allow the crew to determine whether oxygen is being delivered to the dispensing equipment.
§ 25.1450   Chemical oxygen generators.
(a) For the purpose of this section, a chemical oxygen generator is defined as a device which produces oxygen by chemical reaction.
(b) Each chemical oxygen generator must be designed and installed in 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.
(3) Except as provided in SFAR 109, each chemical oxygen generator installation must meet the requirements of § 25.795(d).
(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 degrees F.
§ 25.1453   Protection of oxygen equipment from rupture.
Oxygen pressure tanks, and lines between tanks and the shutoff means, must be—
(a) Protected from unsafe temperatures; and
(b) Located where the probability and hazards of rupture in a crash landing are minimized.
§ 25.1455   Draining of fluids subject to freezing.
If fluids subject to freezing may be drained overboard in flight or during ground operation, the drains must be designed and located to prevent the formation of hazardous quantities of ice on the airplane as a result of the drainage.
§ 25.1457   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.
(6) If datalink communication equipment is installed, all datalink communications, using an approved data message set. Datalink messages must be recorded as the output signal from the communications unit that translates the signal into usable data.
(b) The recording requirements of paragraph (a)(2) of this section must be met by installing a cockpit-mounted area microphone, located in the best position for recording voice 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, 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 hand-held microphone, headset, or speaker used at the first pilot station.
(2) For the second channel from each boom, mask, or hand-held microphone, headset, or speaker used at the second pilot station.
(3) For the third channel—from the cockpit-mounted area microphone.
(4) For the fourth channel, from—
(i) Each boom, mask, or hand-held microphone, headset, or speaker used at the station for the third and fourth crew members; or
(ii) If the stations specified in 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) 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)(i) It receives its electrical power from the bus that provides the maximum reliability for operation of the cockpit voice recorder without jeopardizing service to essential or emergency loads.
(ii) It remains powered for as long as possible without jeopardizing emergency operation of the airplane.
(2) There is an automatic means to simultaneously stop the recorder and prevent each erasure feature from functioning, within 10 minutes after crash impact;
(3) There is an aural or visual means for preflight checking of the recorder for proper operation;
(4) Any single electrical failure external to the recorder does not disable both the cockpit voice recorder and the flight data recorder;
(5) It has an independent power source—
(i) That provides 10 ±1 minutes of electrical power to operate both the cockpit voice recorder and cockpit-mounted area microphone;
(ii) That is located as close as practicable to the cockpit voice recorder; and
(iii) To which the cockpit voice recorder and cockpit-mounted area microphone are switched automatically in the event that all other power to the cockpit voice recorder is interrupted either by normal shutdown or by any other loss of power to the electrical power bus; and
(6) It is in a separate container from the flight data recorder when both are required. If used to comply with only the cockpit voice recorder requirements, a combination unit may be installed.
(e) The recorder container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and consequent heat damage to the recorder from fire.
(1) Except as provided in paragraph (e)(2) of this section, the recorder container must be located as far aft as practicable, but need not be outside of the pressurized compartment, and may not be located where aft-mounted engines may crush the container during impact.
(2) If two separate combination digital flight data recorder and cockpit voice recorder units are installed instead of one cockpit voice recorder and one digital flight data recorder, the combination unit that is installed to comply with the cockpit voice recorder requirements may be located near the cockpit.
(f) If the cockpit voice recorder has a bulk erasure device, the installation must be designed to minimize the 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.
§ 25.1459   Flight data recorders.
(a) Each flight recorder required by the operating rules of this chapter must be installed so that—
(1) It is supplied with airspeed, altitude, and directional data obtained from sources that meet the accuracy requirements of §§ 25.1323, 25.1325, and 25.1327, 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)(i) It receives its electrical power from the bus that provides the maximum reliability for operation of the flight data recorder without jeopardizing service to essential or emergency loads.
(ii) It remains powered for as long as possible without jeopardizing emergency operation of the airplane.
(4) There is an aural or visual means for preflight checking of the recorder for proper recording of data in the 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;
(6) There is a means to record data from which the time of each radio transmission either to or from ATC can be determined;
(7) Any single electrical failure external to the recorder does not disable both the cockpit voice recorder and the flight data recorder; and
(8) It is in a separate container from the cockpit voice recorder when both are required. If used to comply with only the flight data recorder requirements, a combination unit may be installed. If a combination unit is installed as a cockpit voice recorder to comply with § 25.1457(e)(2), a combination unit must be used to comply with this flight data recorder requirement.
(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.
§ 25.1461   Equipment containing high energy rotors.
(a) Equipment containing high energy rotors must meet paragraph (b), (c), or (d) of this section.
(b) High energy rotors contained in equipment must be able to withstand damage caused by malfunctions, 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.
§ 25.1501   General.
(a) Each operating limitation specified in §§ 25.1503 through 25.1533 and other limitations and information necessary for safe operation must be established.
(b) The operating limitations and other information necessary for safe operation must be made available to the crewmembers as prescribed in §§ 25.1541 through 25.1587.
§ 25.1503   Airspeed limitations: general.
When airspeed limitations are a function of weight, weight distribution, altitude, or Mach number, limitations corresponding to each critical combination of these factors must be established.
§ 25.1505   Maximum operating limit speed.
The maximum operating limit speed ( V MO / M MO airspeed or Mach Number, whichever is critical at a particular altitude) is a speed that may not be deliberately exceeded in any regime of flight (climb, cruise, or descent), unless a higher speed is authorized for flight test or pilot training operations. V MO / M MO must be established so that it is not greater than the design cruising speed V C and so that it is sufficiently below V D / M D or V DF / M DF, to make it highly improbable that the latter speeds will be inadvertently exceeded in operations. The speed margin between V MO / M MO and V D / M D or V DF M/ DF may not be less than that determined under § 25.335(b) or found necessary during the flight tests conducted under § 25.253.
§ 25.1507   Maneuvering speed.
The maneuvering speed must be established so that it does not exceed the design maneuvering speed V A determined under § 25.335(c).
§ 25.1511   Flap extended speed.
The established flap extended speed V FE must be established so that it does not exceed the design flap speed V F chosen under §§ 25.335(e) and 25.345, for the corresponding flap positions and engine powers.
§ 25.1513   Minimum control speed.
The minimum control speed V MC determined under § 25.149 must be established as an operating limitation.
§ 25.1515   Landing gear speeds.
(a) The established landing gear operating speed or speeds, V LO, may not exceed the speed at which it is safe both to extend and to retract the landing gear, as determined under § 25.729 or by flight characteristics. If the extension speed is not the same as the retraction speed, the two speeds must be designated as V LO(EXT) and V LO(RET), respectively.
(b) The established landing gear extended speed V LE may not exceed the speed at which it is safe to fly with the landing gear secured in the fully extended position, and that determined under § 25.729.
§ 25.1516   Other speed limitations.
Any other limitation associated with speed must be established.
§ 25.1517   Rough air speed, VRA.
(a) A rough air speed, V RA, for use as the recommended turbulence penetration airspeed, and a rough air Mach number, M RA , for use as the recommended turbulence penetration Mach number, must be established. V RA /M RA must be sufficiently less than V MO /M MO to ensure that likely speed variation during rough air encounters will not cause the overspeed warning to operate too frequently.
(b) At altitudes where V MO is not limited by Mach number, in the absence of a rational investigation substantiating the use of other values, V RA must be less than V MO minus 35 KTAS.
(c) At altitudes where V MO is limited by Mach number, M RA may be chosen to provide an optimum margin between low and high speed buffet boundaries.
§ 25.1519   Weight, center of gravity, and weight distribution.
The airplane weight, center of gravity, and weight distribution limitations determined under §§ 25.23 through 25.27 must be established as operating limitations.
§ 25.1521   Powerplant limitations.
(a) General. The powerplant limitations prescribed in this section must be established so that they do not exceed the corresponding limits for which the engines or propellers are type certificated and do not exceed the values on which compliance with any other requirement of this part is based.
(b) Reciprocating engine installations. Operating limitations relating to the following must be established for reciprocating engine installations:
(1) Horsepower or torque, r.p.m., manifold pressure, and time at critical pressure altitude and sea level pressure altitude for—
(i) Maximum continuous power (relating to unsupercharged operation or to operation in each supercharger mode as applicable); and
(ii) Takeoff power (relating to unsupercharged operation or to operation in each supercharger mode as applicable).
(2) Fuel grade or specification.
(3) Cylinder head and oil temperatures.
(4) Any other parameter for which a limitation has been established as part of the engine type certificate except that a limitation need not be established for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation.
(c) Turbine engine installations. Operating limitations relating to the following must be established for turbine engine installations:
(1) Horsepower, torque or thrust, r.p.m., gas temperature, and time for—
(i) Maximum continuous power or thrust (relating to augmented or unaugmented operation as applicable).
(ii) Takeoff power or thrust (relating to augmented or unaugmented operation as applicable).
(2) Fuel designation or specification.
(3) Maximum time interval between engine run-ups from idle, run-up power setting and duration at power for ground operation in icing conditions, as defined in § 25.1093(b)(2).
(4) Any other parameter for which a limitation has been established as part of the engine type certificate except that a limitation need not be established for a parameter that cannot be exceeded during normal operation due to the design of the installation or to another established limitation.
(d) Ambient temperature. An ambient temperature limitation (including limitations for winterization installations, if applicable) must be established as the maximum ambient atmospheric temperature established in accordance with § 25.1043(b).
§ 25.1522   Auxiliary power unit limitations.
If an auxiliary power unit is installed in the airplane, limitations established for the auxiliary power unit, including categories of operation, must be specified as operating limitations for the airplane.
§ 25.1523   Minimum flight crew.
The minimum flight crew must be established so that it is sufficient for safe operation, considering—
(a) The workload on individual crewmembers;
(b) The accessibility and ease of operation of necessary controls by the appropriate crewmember; and
(c) The kind of operation authorized under § 25.1525.
§ 25.1525   Kinds of operation.
The kinds of operation to which the airplane is limited are established by the category in which it is eligible for certification and by the installed equipment.
§ 25.1527   Ambient air temperature and operating altitude.
The extremes of the ambient air temperature and operating altitude for which operation is allowed, as limited by flight, structural, powerplant, functional, or equipment characteristics, must be established.
§ 25.1529   Instructions for Continued Airworthiness.
The applicant must prepare Instructions for Continued Airworthiness in accordance with appendix H to this part that are acceptable to the Administrator. The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first airplane or issuance of a standard certificate of airworthiness, whichever occurs later.
§ 25.1531   Maneuvering flight load factors.
Load factor limitations, not exceeding the positive limit load factors determined from the maneuvering diagram in § 25.333(b), must be established.
§ 25.1533   Additional operating limitations.
(a) Additional operating limitations must be established as follows:
(1) The maximum takeoff weights must be established as the weights at which compliance is shown with the applicable provisions of this part (including the takeoff climb provisions of § 25.121(a) through (c), for altitudes and ambient temperatures).
(2) The maximum landing weights must be established as the weights at which compliance is shown with the applicable provisions of this part (including the landing and approach climb provisions of §§ 25.119 and 25.121(d) for altitudes and ambient temperatures).
(3) The minimum takeoff distances must be established as the distances at which compliance is shown with the applicable provisions of this part (including the provisions of §§ 25.109 and 25.113, for weights, altitudes, temperatures, wind components, runway surface conditions (dry and wet), and runway gradients) for smooth, hard-surfaced runways. Additionally, at the option of the applicant, wet runway takeoff distances may be established for runway surfaces that have been grooved or treated with a porous friction course, and may be approved for use on runways where such surfaces have been designed constructed, and maintained in a manner acceptable to the Administrator.
(b) The extremes for variable factors (such as altitude, temperature, wind, and runway gradients) are those at which compliance with the applicable provisions of this part is shown.
(c) For airplanes certified in accordance with § 25.1420(a)(1) or (2), an operating limitation must be established to:
(1) Prohibit intentional flight, including takeoff and landing, into icing conditions defined in Appendix O of this part for which the airplane has not been certified to safely operate; and
(2) Require exiting all icing conditions if icing conditions defined in Appendix O of this part are encountered for which the airplane has not been certified to safely operate.
§ 25.1535   ETOPS approval.
Except as provided in § 25.3, each applicant seeking ETOPS type design approval must comply with the provisions of Appendix K of this part.
§ 25.1541   General.
(a) The airplane must contain—
(1) The specified markings and placards; and
(2) Any additional information, instrument markings, and placards required for the safe operation if there are unusual design, operating, or handling characteristics.
(b) Each marking and placard prescribed in paragraph (a) of this section—
(1) Must be displayed in a conspicuous place; and
(2) May not be easily erased, disfigured, or obscured.
§ 25.1543   Instrument markings: general.
For each instrument—
(a) When markings are on the cover glass of the instrument, there must be means to maintain the correct alignment of the glass cover with the face of the dial; and
(b) Each instrument marking must be clearly visible to the appropriate crewmember.
§ 25.1545   Airspeed limitation information.
The airspeed limitations required by § 25.1583 (a) must be easily read and understood by the flight crew.
§ 25.1547   Magnetic direction indicator.
(a) A placard meeting the requirements of this section must be installed on, or near, the magnetic direction indicator.
(b) The placard must show the calibration of the instrument in level flight with the engines operating.
(c) The placard must state whether the calibration was made with radio receivers on or off.
(d) Each calibration reading must be in terms of magnetic heading in not more than 45 degree increments.
§ 25.1549   Powerplant and auxiliary power unit instruments.
For each required powerplant and auxiliary power unit instrument, as appropriate to the type of instrument—
(a) Each maximum and, if applicable, minimum safe operating limit must be marked with a red radial or a red line;
(b) Each normal operating range must be marked with a green arc or green line, not extending beyond the maximum and minimum safe limits;
(c) Each takeoff and precautionary range must be marked with a yellow arc or a yellow line; and
(d) Each engine, auxiliary power unit, or propeller speed range that is restricted because of excessive vibration stresses must be marked with red arcs or red lines.
§ 25.1551   Oil quantity indication.
Each oil quantity indicating means must be marked to indicate the quantity of oil readily and accurately.
§ 25.1553   Fuel quantity indicator.
If the unusable fuel supply for any tank exceeds one gallon, or five percent of the tank capacity, whichever is greater, a red arc must be marked on its indicator extending from the calibrated zero reading to the lowest reading obtainable in level flight.
§ 25.1555   Control markings.
(a) Each cockpit control, other than primary flight controls and controls whose function is obvious, must be plainly marked as to its function and method of operation.
(b) Each aerodynamic control must be marked under the requirements of §§ 25.677 and 25.699.
(c) For powerplant fuel controls—
(1) Each fuel tank selector control must be marked to indicate the position corresponding to each tank and to each existing cross feed position;
(2) If safe operation requires the use of any tanks in a specific sequence, that sequence must be marked on, or adjacent to, the selector for those tanks; and
(3) Each valve control for each engine must be marked to indicate the position corresponding to each engine controlled.
(d) For accessory, auxiliary, and emergency controls—
(1) Each emergency control (including each fuel jettisoning and fluid shutoff must be colored red; and
(2) Each visual indicator required by § 25.729(e) must be marked so that the pilot can determine at any time when the wheels are locked in either extreme position, if retractable landing gear is used.
§ 25.1557   Miscellaneous markings and placards.
(a) Baggage and cargo compartments and ballast location. Each baggage and cargo compartment, and each ballast location must have a placard stating any limitations on contents, including weight, that are necessary under the loading requirements. However, underseat compartments designed for the storage of carry-on articles weighing not more than 20 pounds need not have a loading limitation placard.
(b) Powerplant fluid filler openings. The following apply:
(1) Fuel filler openings must be marked at or near the filler cover with—
(i) The word “fuel”;
(ii) For reciprocating engine powered airplanes, the minimum fuel grade;
(iii) For turbine engine powered airplanes, the permissible fuel designations; and
(iv) For pressure fueling systems, the maximum permissible fueling supply pressure and the maximum permissible defueling pressure.
(2) Oil filler openings must be marked at or near the filler cover with the word “oil”.
(3) Augmentation fluid filler openings must be marked at or near the filler cover to identify the required fluid.
(c) Emergency exit placards. Each emergency exit placard must meet the requirements of § 25.811.
(d) Doors. Each door that must be used in order to reach any required emergency exit must have a suitable placard stating that the door is to be latched in the open position during takeoff and landing.
§ 25.1561   Safety equipment.
(a) Each safety equipment control to be operated by the crew in emergency, such as controls for automatic liferaft releases, must be plainly marked as to its method of operation.
(b) Each location, such as a locker or compartment, that carries any fire extinguishing, signaling, or other life saving equipment must be marked accordingly.
(c) Stowage provisions for required emergency equipment must be conspicuously marked to identify the contents and facilitate the easy removal of the equipment.
(d) Each liferaft must have obviously marked operating instructions.
(e) Approved survival equipment must be marked for identification and method of operation.
§ 25.1563   Airspeed placard.
A placard showing the maximum airspeeds for flap extension for the takeoff, approach, and landing positions must be installed in clear view of each pilot.
§ 25.1581   General.
(a) Furnishing information. An Airplane Flight Manual must be furnished with each airplane, and it must contain the following:
(1) Information required by §§ 25.1583 through 25.1587.
(2) Other information that is necessary for safe operation because of design, operating, or handling characteristics.
(3) Any limitation, procedure, or other information established as a condition of compliance with the applicable noise standards of part 36 of this chapter.
(b) Approved information. Each part of the manual listed in §§ 25.1583 through 25.1587, that is appropriate to the airplane, must be furnished, verified, and approved, and must be segregated, identified, and clearly distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Each Airplane Flight Manual must include a table of contents if the complexity of the manual indicates a need for it.
§ 25.1583   Operating limitations.
(a) Airspeed limitations. The following airspeed limitations and any other airspeed limitations necessary for safe operation must be furnished:
(1) The maximum operating limit speed V MO / M MO and a statement that this speed limit may not be deliberately exceeded in any regime of flight (climb, cruise, or descent) unless a higher speed is authorized for flight test or pilot training.
(2) If an airspeed limitation is based upon compressibility effects, a statement to this effect and information as to any symptoms, the probable behavior of the airplane, and the recommended recovery procedures.
(3) The maneuvering speed established under § 25.1507 and statements, as applicable to the particular design, explaining that:
(i) Full application of pitch, roll, or yaw controls should be confined to speeds below the maneuvering speed; and
(ii) Rapid and large alternating control inputs, especially in combination with large changes in pitch, roll, or yaw, and full control inputs in more than one axis at the same time, should be avoided as they may result in structural failures at any speed, including below the maneuvering speed.
(4) The flap extended speed V FE and the pertinent flap positions and engine powers.
(5) The landing gear operating speed or speeds, and a statement explaining the speeds as defined in § 25.1515(a).
(6) The landing gear extended speed V LE, if greater than V LO, and a statement that this is the maximum speed at which the airplane can be safely flown with the landing gear extended.
(b) Powerplant limitations. The following information must be furnished:
(1) Limitations required by § 25.1521 and § 25.1522.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by §§ 25.1549 through 25.1553.
(c) Weight and loading distribution. The weight and center of gravity limitations established under § 25.1519 must be furnished in the Airplane Flight Manual. All of the following information, including the weight distribution limitations established under § 25.1519, must be presented either in the Airplane Flight Manual or in a separate weight and balance control and loading document that is incorporated by reference in the Airplane Flight Manual:
(1) The condition of the airplane and the items included in the empty weight as defined in accordance with § 25.29.
(2) Loading instructions necessary to ensure loading of the airplane within the weight and center of gravity limits, and to maintain the loading within these limits in flight.
(3) If certification for more than one center of gravity range is requested, the appropriate limitations, with regard to weight and loading procedures, for each separate center of gravity range.
(d) Flight crew. The number and functions of the minimum flight crew determined under § 25.1523 must be furnished.
(e) Kinds of operation. The kinds of operation approved under § 25.1525 must be furnished.
(f) Ambient air temperatures and operating altitudes. The extremes of the ambient air temperatures and operating altitudes established under § 25.1527 must be furnished.
(g) [Reserved]
(h) Additional operating limitations. The operating limitations established under § 25.1533 must be furnished.
(i) Maneuvering flight load factors. The positive maneuvering limit load factors for which the structure is proven, described in terms of accelerations, must be furnished.
§ 25.1585   Operating procedures.
(a) Operating procedures must be furnished for—
(1) Normal procedures peculiar to the particular type or model encountered in connection with routine operations;
(2) Non-normal procedures for malfunction cases and failure conditions involving the use of special systems or the alternative use of regular systems; and
(3) Emergency procedures for foreseeable but unusual situations in which immediate and precise action by the crew may be expected to substantially reduce the risk of catastrophe.
(b) Information or procedures not directly related to airworthiness or not under the control of the crew, must not be included, nor must any procedure that is accepted as basic airmanship.
(c) Information identifying each operating condition in which the fuel system independence prescribed in § 25.953 is necessary for safety must be furnished, together with instructions for placing the fuel system in a configuration used to show compliance with that section.
(d) The buffet onset envelopes, determined under § 25.251 must be furnished. The buffet onset envelopes presented may reflect the center of gravity at which the airplane is normally loaded during cruise if corrections for the effect of different center of gravity locations are furnished.
(e) Information must be furnished that indicates that when the fuel quantity indicator reads “zero” in level flight, any fuel remaining in the fuel tank cannot be used safely in flight.
(f) Information on the total quantity of usable fuel for each fuel tank must be furnished.
§ 25.1587   Performance information.
(a) Each Airplane Flight Manual must contain information to permit conversion of the indicated temperature to free air temperature if other than a free air temperature indicator is used to comply with the requirements of § 25.1303(a)(1).
(b) Each Airplane Flight Manual must contain the performance information computed under the applicable provisions of this part (including §§ 25.115, 25.123, and 25.125 for the weights, altitudes, temperatures, wind components, and runway gradients, as applicable) within the operational limits of the airplane, and must contain the following:
(1) In each case, the conditions of power, configuration, and speeds, and the procedures for handling the airplane and any system having a significant effect on the performance information.
(2) V SR determined in accordance with § 25.103.
(3) The following performance information (determined by extrapolation and computed for the range of weights between the maximum landing weight and the maximum takeoff weight):
(i) Climb in the landing configuration.
(ii) Climb in the approach configuration.
(iii) Landing distance.
(4) Procedures established under § 25.101(f) and (g) that are related to the limitations and information required by § 25.1533 and by this paragraph (b) in the form of guidance material, including any relevant limitations or information.
(5) An explanation of significant or unusual flight or ground handling characteristics of the airplane.
(6) Corrections to indicated values of airspeed, altitude, and outside air temperature.
(7) An explanation of operational landing runway length factors included in the presentation of the landing distance, if appropriate.