Title 14—Aeronautics and Space--Volume 1
PART 29—AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY ROTORCRAFT
§ 29.952   Fuel system crash resistance.
Unless other means acceptable to the Administrator are employed to minimize the hazard of fuel fires to occupants following an otherwise survivable impact (crash landing), the fuel systems must incorporate the design features of this section. These systems must be shown to be capable of sustaining the static and dynamic deceleration loads of this section, considered as ultimate loads acting alone, measured at the system component's center of gravity without structural damage to the system components, fuel tanks, or their attachments that would leak fuel to an ignition source.
(a) Drop test requirements. Each tank, or the most critical tank, must be drop-tested as follows:
(1) The drop height must be at least 50 feet.
(2) The drop impact surface must be nondeforming.
(3) The tanks must be filled with water to 80 percent of the normal, full capacity.
(4) The tank must be enclosed in a surrounding structure representative of the installation unless it can be established that the surrounding structure is free of projections or other design features likely to contribute to upture of the tank.
(5) The tank must drop freely and impact in a horizontal position ±10°.
(6) After the drop test, there must be no leakage.
(b) Fuel tank load factors. Except for fuel tanks located so that tank rupture with fuel release to either significant ignition sources, such as engines, heaters, and auxiliary power units, or occupants is extremely remote, each fuel tank must be designed and installed to retain its contents under the following ultimate inertial load factors, acting alone.
(1) For fuel tanks in the cabin:
(i) Upward—4g.
(ii) Forward—16g.
(iii) Sideward—8g.
(iv) Downward—20g.
(2) For fuel tanks located above or behind the crew or passenger compartment that, if loosened, could injure an occupant in an emergency landing:
(i) Upward—1.5g.
(ii) Forward—8g.
(iii) Sideward—2g.
(iv) Downward—4g.
(3) For fuel tanks in other areas:
(i) Upward—1.5g.
(ii) Forward—4g.
(iii) Sideward—2g.
(iv) Downward—4g.
(c) Fuel line self-sealing breakaway couplings. Self-sealing breakaway couplings must be installed unless hazardous relative motion of fuel system components to each other or to local rotorcraft structure is demonstrated to be extremely improbable or unless other means are provided. The couplings or equivalent devices must be installed at all fuel tank-to-fuel line connections, tank-to-tank interconnects, and at other points in the fuel system where local structural deformation could lead to the release of fuel.
(1) The design and construction of self-sealing breakaway couplings must incorporate the following design features:
(i) The load necessary to separate a breakaway coupling must be between 25 to 50 percent of the minimum ultimate failure load (ultimate strength) of the weakest component in the fluid-carrying line. The separation load must in no case be less than 300 pounds, regardless of the size of the fluid line.
(ii) A breakaway coupling must separate whenever its ultimate load (as defined in paragraph (c)(1)(i) of this section) is applied in the failure modes most likely to occur.
(iii) All breakaway couplings must incorporate design provisions to visually ascertain that the coupling is locked together (leak-free) and is open during normal installation and service.
(iv) All breakaway couplings must incorporate design provisions to prevent uncoupling or unintended closing due to operational shocks, vibrations, or accelerations.
(v) No breakaway coupling design may allow the release of fuel once the coupling has performed its intended function.
(2) All individual breakaway couplings, coupling fuel feed systems, or equivalent means must be designed, tested, installed, and maintained so inadvertent fuel shutoff in flight is improbable in accordance with § 29.955(a) and must comply with the fatigue evaluation requirements of § 29.571 without leaking.
(3) Alternate, equivalent means to the use of breakaway couplings must not create a survivable impact-induced load on the fuel line to which it is installed greater than 25 to 50 percent of the ultimate load (strength) of the weakest component in the line and must comply with the fatigue requirements of § 29.571 without leaking.
(d) Frangible or deformable structural attachments. Unless hazardous relative motion of fuel tanks and fuel system components to local rotorcraft structure is demonstrated to be extremely improbable in an otherwise survivable impact, frangible or locally deformable attachments of fuel tanks and fuel system components to local rotorcraft structure must be used. The attachment of fuel tanks and fuel system components to local rotorcraft structure, whether frangible or locally deformable, must be designed such that its separation or relative local deformation will occur without rupture or local tear-out of the fuel tank or fuel system component that will cause fuel leakage. The ultimate strength of frangible or deformable attachments must be as follows:
(1) The load required to separate a frangible attachment from its support structure, or deform a locally deformable attachment relative to its support structure, must be between 25 and 50 percent of the minimum ultimate load (ultimate strength) of the weakest component in the attached system. In no case may the load be less than 300 pounds.
(2) A frangible or locally deformable attachment must separate or locally deform as intended whenever its ultimate load (as defined in paragraph (d)(1) of this section) is applied in the modes most likely to occur.
(3) All frangible or locally deformable attachments must comply with the fatigue requirements of § 29.571.
(e) Separation of fuel and ignition sources. To provide maximum crash resistance, fuel must be located as far as practicable from all occupiable areas and from all potential ignition sources.
(f) Other basic mechanical design criteria. Fuel tanks, fuel lines, electrical wires, and electrical devices must be designed, constructed, and installed, as far as practicable, to be crash resistant.
(g) Rigid or semirigid fuel tanks. Rigid or semirigid fuel tank or bladder walls must be impact and tear resistant.
§ 29.953   Fuel system independence.
(a) For category A rotorcraft—
(1) The fuel system must meet the requirements of § 29.903(b); and
(2) Unless other provisions are made to meet paragraph (a)(1) of this section, the fuel system must allow fuel to be supplied to each engine through a system independent of those parts of each system supplying fuel to other engines.
(b) Each fuel system for a multiengine category B rotorcraft must meet the requirements of paragraph (a)(2) of this section. However, separate fuel tanks need not be provided for each engine.
§ 29.954   Fuel system lightning protection.
The fuel system must be designed and arranged to prevent the ignition of fuel vapor within the system by—
(a) Direct lightning strikes to areas having a high probability of stroke attachment;
(b) Swept lightning strokes to areas where swept strokes are highly probable; and
(c) Corona and streamering at fuel vent outlets.
§ 29.955   Fuel flow.
(a) General. The fuel system for each engine must provide the engine with at least 100 percent of the fuel required under all operating and maneuvering conditions to be approved for the rotorcraft, including, as applicable, the fuel required to operate the engines under the test conditions required by § 29.927. Unless equivalent methods are used, compliance must be shown by test during which the following provisions are met, except that combinations of conditions which are shown to be improbable need not be considered.
(1) The fuel pressure, corrected for accelerations (load factors), must be within the limits specified by the engine type certificate data sheet.
(2) The fuel level in the tank may not exceed that established as the unusable fuel supply for that tank under § 29.959, plus that necessary to conduct the test.
(3) The fuel head between the tank and the engine must be critical with respect to rotorcraft flight attitudes.
(4) The fuel flow transmitter, if installed, and the critical fuel pump (for pump-fed systems) must be installed to produce (by actual or simulated failure) the critical restriction to fuel flow to be expected from component failure.
(5) Critical values of engine rotational speed, electrical power, or other sources of fuel pump motive power must be applied.
(6) Critical values of fuel properties which adversely affect fuel flow are applied during demonstrations of fuel flow capability.
(7) The fuel filter required by § 29.997 is blocked to the degree necessary to simulate the accumulation of fuel contamination required to activate the indicator required by § 29.1305(a)(18).
(b) Fuel transfer system. If normal operation of the fuel system requires fuel to be transferred to another tank, the transfer must occur automatically via a system which has been shown to maintain the fuel level in the receiving tank within acceptable limits during flight or surface operation of the rotorcraft.
(c) Multiple fuel tanks. If an engine can be supplied with fuel from more than one tank, the fuel system, in addition to having appropriate manual switching capability, must be designed to prevent interruption of fuel flow to that engine, without attention by the flightcrew, when any tank supplying fuel to that engine is depleted of usable fuel during normal operation and any other tank that normally supplies fuel to that engine alone contains usable fuel.
§ 29.957   Flow between interconnected tanks.
(a) Where tank outlets are interconnected and allow fuel to flow between them due to gravity or flight accelerations, it must be impossible for fuel to flow between tanks in quantities great enough to cause overflow from the tank vent in any sustained flight condition.
(b) If fuel can be pumped from one tank to another in flight—
(1) The design of the vents and the fuel transfer system must prevent structural damage to tanks from overfilling; and
(2) There must be means to warn the crew before overflow through the vents occurs.
§ 29.959   Unusable fuel supply.
The unusable fuel supply for each tank must be established as not less than the quantity at which the first evidence of malfunction occurs under the most adverse fuel feed condition occurring under any intended operations and flight maneuvers involving that tank.
§ 29.961   Fuel system hot weather operation.
Each suction lift fuel system and other fuel systems conducive to vapor formation must be shown to operate satisfactorily (within certification limits) when using fuel at the most critical temperature for vapor formation under critical operating conditions including, if applicable, the engine operating conditions defined by § 29.927(b)(1) and (b)(2).
§ 29.963   Fuel tanks: general.
(a) Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads to which it may be subjected in operation.
(b) Each flexible fuel tank bladder or liner must be approved or shown to be suitable for the particular application and must be puncture resistant. Puncture resistance must be shown by meeting the TSO-C80, paragraph 16.0, requirements using a minimum puncture force of 370 pounds.
(c) Each integral fuel tank must have facilities for inspection and repair of its interior.
(d) The maximum exposed surface temperature of all components in the fuel tank must be less by a safe margin than the lowest expected autoignition temperature of the fuel or fuel vapor in the tank. Compliance with this requirement must be shown under all operating conditions and under all normal or malfunction conditions of all components inside the tank.
(e) Each fuel tank installed in personnel compartments must be isolated by fume-proof and fuel-proof enclosures that are drained and vented to the exterior of the rotorcraft. The design and construction of the enclosures must provide necessary protection for the tank, must be crash resistant during a survivable impact in accordance with § 29.952, and must be adequate to withstand loads and abrasions to be expected in personnel compartments.
§ 29.965   Fuel tank tests.
(a) Each fuel tank must be able to withstand the applicable pressure tests in this section without failure or leakage. If practicable, test pressures may be applied in a manner simulating the pressure distribution in service.
(b) Each conventional metal tank, each nonmetallic tank with walls that are not supported by the rotorcraft structure, and each integral tank must be subjected to a pressure of 3.5 p.s.i. unless the pressure developed during maximum limit acceleration or emergency deceleration with a full tank exceeds this value, in which case a hydrostatic head, or equivalent test, must be applied to duplicate the acceleration loads as far as possible. However, the pressure need not exceed 3.5 p.s.i. on surfaces not exposed to the acceleration loading.
(c) Each nonmetallic tank with walls supported by the rotorcraft structure must be subjected to the following tests:
(1) A pressure test of at least 2.0 p.s.i. This test may be conducted on the tank alone in conjunction with the test specified in paragraph (c)(2) of this section.
(2) A pressure test, with the tank mounted in the rotorcraft structure, equal to the load developed by the reaction of the contents, with the tank full, during maximum limit acceleration or emergency deceleration. However, the pressure need not exceed 2.0 p.s.i. on surfaces faces not exposed to the acceleration loading.
(d) Each tank with large unsupported or unstiffened flat areas, or with other features whose failure or deformation could cause leakage, must be subjected to the following test or its equivalent:
(1) Each complete tank assembly and its supports must be vibration tested while mounted to simulate the actual installation.
(2) The tank assembly must be vibrated for 25 hours while two-thirds full of any suitable fluid. The amplitude of vibration may not be less than one thirty-second of an inch, unless otherwise substantiated.
(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 or rotor system speeds is critical, the test frequency of vibration, in number of cycles per minute, must, unless a frequency based on a more rational analysis is used, be the number obtained by averaging the maximum and minimum power-on engine speeds (r.p.m.) for reciprocating engine powered rotorcraft or 2,000 c.p.m. for turbine engine powered rotorcraft.
(ii) If only one frequency of vibration resulting from any r.p.m. within the normal operating range of engine or rotor system 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 or rotor system speeds is critical, the most critical of these frequencies must be the test frequency.
(4) Under paragraph (d)(3)(ii) and (iii), the time of test must be adjusted to accomplish the same number of vibration cycles as would be accomplished in 25 hours at the frequency specified in paragraph (d)(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 degrees on both sides of the horizontal (30 degrees 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.
§ 29.967   Fuel tank installation.
(a) Each fuel tank must be supported so that tank loads are not concentrated on unsupported tank surfaces. In addition—
(1) There must be pads, if necessary, to prevent chafing between each tank and its supports;
(2) The padding must be nonabsorbent or treated to prevent the absorption of fuel;
(3) If flexible tank liners are used, they must be supported so that they are not required to withstand fluid loads; and
(4) Each interior surface of tank compartments must be smooth and free of projections that could cause wear of the liner, unless—
(i) There are means for protection of the liner at those points; or
(ii) The construction of the liner itself provides such protection.
(b) Any spaces adjacent to tank surfaces must be adequately ventilated to avoid accumulation of fuel or fumes in those spaces due to minor leakage. If the tank is in a sealed compartment, ventilation may be limited to drain holes that prevent clogging and that prevent excessive pressure resulting from altitude changes. If flexible tank liners are installed, the venting arrangement for the spaces between the liner and its container must maintain the proper relationship to tank vent pressures for any expected flight condition.
(c) The location of each tank must meet the requirements of § 29.1185(b) and (c).
(d) No rotorcraft skin immediately adjacent to a major air outlet from the engine compartment may act as the wall of an integral tank.
§ 29.969   Fuel tank expansion space.
Each fuel tank or each group of fuel tanks with interconnected vent systems must have an expansion space of not less than 2 percent of the combined tank capacity. It must be impossible to fill the fuel tank expansion space inadvertently with the rotorcraft in the normal ground attitude.
§ 29.971   Fuel tank sump.
(a) Each fuel tank must have a sump with a capacity of not less than the greater of—
(1) 0.10 per cent of the tank capacity; or
(2) 1/16 gallon.
(b) The capacity prescribed in paragraph (a) of this section must be effective with the rotorcraft in any normal attitude, and must be located so that the sump contents cannot escape through the tank outlet opening.
(c) Each fuel tank must allow drainage of hazardous quantities of water from each part of the tank to the sump with the rotorcraft in any ground attitude to be expected in service.
(d) Each fuel tank sump must have a drain that allows complete drainage of the sump on the ground.
§ 29.973   Fuel tank filler connection.
(a) Each fuel tank filler connection must prevent the entrance of fuel into any part of the rotorcraft other than the tank itself during normal operations and must be crash resistant during a survivable impact in accordance with § 29.952(c). In addition—
(1) Each filler must be marked as prescribed in § 29.1557(c)(1);
(2) Each recessed filler connection that can retain any appreciable quantity of fuel must have a drain that discharges clear of the entire rotorcraft; and
(3) Each filler cap must provide a fuel-tight seal under the fluid pressure expected in normal operation and in a survivable impact.
(b) Each filler cap or filler cap cover must warn when the cap is not fully locked or seated on the filler connection.
§ 29.975   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 normal flight conditions. In addition—
(1) The vents 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 rotorcraft 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) The venting system must be designed to minimize spillage of fuel through the vents to an ignition source in the event of a rollover during landing, ground operations, or a survivable impact.
(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.
§ 29.977   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 rotorcraft, have 8 to 16 meshes per inch; and
(2) For turbine engine powered rotorcraft, prevent the passage of any object that could restrict fuel flow or damage any fuel system component.
(b) The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line.
(c) The diameter of each strainer must be at least that of the fuel tank outlet.
(d) Each finger strainer must be accessible for inspection and cleaning.
§ 29.979   Pressure refueling and fueling provisions below fuel level.
(a) Each fueling connection below the fuel level in each tank must have means to prevent the escape of hazardous quantities of fuel from that tank in case of malfunction of the fuel entry valve.
(b) For systems intended for pressure refueling, a means in addition to the normal means for limiting the tank content must be installed to prevent damage to the tank in case of failure of the normal means.
(c) The rotorcraft pressure fueling system (not fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum pressure, 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.
(d) The rotorcraft 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 rotorcraft fueling connection.
§ 29.991   Fuel pumps.
(a) Compliance with § 29.955 must not be jeopardized by failure of—
(1) Any one pump except pumps that are approved and installed as parts of a type certificated engine; or
(2) Any component required for pump operation except the engine served by that pump.
(b) The following fuel pump installation requirements apply:
(1) When necessary to maintain the proper fuel pressure—
(i) A connection must be provided to transmit the carburetor air intake static pressure to the proper fuel pump relief valve connection; and
(ii) The gauge balance lines must be independently connected to the carburetor inlet pressure to avoid incorrect fuel pressure readings.
(2) The installation of fuel pumps having seals or diaphragms that may leak must have means for draining leaking fuel.
(3) Each drain line must discharge where it will not create a fire hazard.
§ 29.993   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, valve actuation, and accelerated flight conditions.
(b) Each fuel line connected to components of the rotorcraft between which relative motion could exist must have provisions for flexibility.
(c) Each flexible connection in fuel lines that may be under pressure or subjected to axial loading must use flexible hose assemblies.
(d) Flexible hose must be approved.
(e) No flexible hose that might be adversely affected by high temperatures may be used where excessive temperatures will exist during operation or after engine shutdown.
§ 29.995   Fuel valves.
In addition to meeting the requirements of § 29.1189, 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.
§ 29.997   Fuel strainer or filter.
There must be a fuel strainer or filter between the fuel tank outlet and the inlet of the first fuel system component which is susceptible to fuel contamination, including but not limited to the fuel metering device or an engine 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 inself, unless adequate strengh margins under all loading conditions are provided in the lines and connections; and
(d) Provide a means to remove from the fuel any contaminant which would jeopardize the flow of fuel through rotorcraft or engine fuel system components required for proper rotorcraft or engine fuel system operation.
§ 29.999   Fuel system drains.
(a) There must be at least one accessible drain at the lowest point in each fuel system to completely drain the system with the rotorcraft in any ground attitude to be expected in service.
(b) Each drain required by paragraph (a) of this section including the drains prescribed in § 29.971 must—
(1) Discharge clear of all parts of the rotorcraft;
(2) Have manual or automatic means to ensure positive closure in the off 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.
§ 29.1001   Fuel jettisoning.
If a fuel jettisoning system is installed, the following apply:
(a) Fuel jettisoning must be safe during all flight regimes for which jettisoning is to be authorized.
(b) In showing compliance with paragraph (a) of this section, it must be shown that—
(1) The fuel jettisoning system and its operation are free from fire hazard;
(2) No hazard results from fuel or fuel vapors which impinge on any part of the rotorcraft during fuel jettisoning; and
(3) Controllability of the rotorcraft remains satisfactory throughout the fuel jettisoning operation.
(c) Means must be provided to automatically prevent jettisoning fuel below the level required for an all-engine climb at maximum continuous power from sea level to 5,000 feet altitude and cruise thereafter for 30 minutes at maximum range engine power.
(d) The controls for any fuel jettisoning system must be designed to allow flight personnel (minimum crew) to safely interrupt fuel jettisoning during any part of the jettisoning operation.
(e) The fuel jettisoning system must be designed to comply with the powerplant installation requirements of § 29.901(c).
(f) An auxiliary fuel jettisoning system which meets the requirements of paragraphs (a), (b), (d), and (e) of this section may be installed to jettison additional fuel provided it has separate and independent controls.
§ 29.1011   Engines: 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 of each system may not be less than the product of the endurance of the rotorcraft under critical operating conditions and the maximum allowable oil consumption of the engine under the same conditions, plus a suitable margin to ensure adequate circulation and cooling. Instead of a rational analysis of endurance and consumption, a usable oil capacity of one gallon for each 40 gallons of usable fuel may be used for reciprocating engine installations.
(c) Oil-fuel ratios lower than those prescribed in paragraph (c) of this section may be used if they are substantiated by data on the oil consumption of the engine.
(d) The ability of the engine and oil cooling provisions to maintain the oil temperature at or below the maximum established value must be shown under the applicable requirements of §§ 29.1041 through 29.1049.
§ 29.1013   Oil tanks.
(a) Installation. Each oil tank installation must meet the requirements of § 29.967.
(b) Expansion space. Oil tank expansion space must be provided so that—
(1) Each oil tank used with a reciprocating engine has 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 has an expansion space of not less than 10 percent of the tank capacity;
(2) Each reserve oil tank not directly connected to any engine has an expansion space of not less than two percent of the tank capacity; and
(3) It is impossible to fill the expansion space inadvertently with the rotorcraft in the normal ground attitude.
(c) Filler connections. Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of the entire rotorcraft. In addition—
(1) Each oil tank filler cap must provide an oil-tight seal under the pressure expected in operation;
(2) For category A rotorcraft, each oil tank filler cap or filler cap cover must incorporate features that provide a warning when caps are not fully locked or seated on the filler connection; and
(3) Each oil filler must be marked under § 29.1557(c)(2).
(d) Vent. Oil tanks must be vented as follows:
(1) Each oil tank must be vented from the top part of the expansion space to that venting is effective under all normal flight conditions.
(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 a 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 oil tank supports) is fireproof.
(f) Flexible liners. Each flexible oil tank liner must be approved or shown to be suitable for the particular installation.
§ 29.1015   Oil tank tests.
Each oil tank must be designed and installed so that—
(a) It can withstand, without failure, any vibration, inertia, and fluid loads to which it may be subjected in operation; and
(b) It meets the requirements of § 29.965, except that instead of the pressure specified in § 29.965(b)—
(1) For pressurized tanks used with a turbine engine, the test pressure may not be less than 5 p.s.i. plus the maximum operating pressure of the tank; and
(2) For all other tanks, the test pressure may not be less than 5 p.s.i.
§ 29.1017   Oil lines and fittings.
(a) Each oil line must meet the requirements of § 29.993.
(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 will not constitute a fire hazard if foaming occurs, or cause emitted oil to strike the pilot's windshield; and
(3) The breather does not discharge into the engine air induction system.
§ 29.1019   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 a means to 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 § 29.1305(a)(19).
(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.
§ 29.1021   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.
§ 29.1023   Oil radiators.
(a) Each oil radiator must be able to withstand any vibration, inertia, and oil pressure loads to which it would be subjected in operation.
(b) Each oil radiator air duct must be located, or equipped, so that, in case of fire, and with the airflow as it would be with and without the engine operating, flames cannot directly strike the radiator.
§ 29.1025   Oil valves.
(a) Each oil shutoff must meet the requirements of § 29.1189.
(b) The closing of oil shutoffs may not prevent autorotation.
(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.
§ 29.1027   Transmission and gearboxes: general.
(a) The oil system for components of the rotor drive system that require continuous lubrication must be sufficiently independent of the lubrication systems of the engine(s) to ensure—
(1) Operation with any engine inoperative; and
(2) Safe autorotation.
(b) Pressure lubrication systems for transmissions and gearboxes must comply with the requirements of §§ 29.1013, paragraphs (c), (d), and (f) only, 29.1015, 29.1017, 29.1021, 29.1023, and 29.1337(d). In addition, the system must have—
(1) An oil strainer or filter through which all the lubricant flows, and must—
(i) Be designed to remove from the lubricant any contaminant which may damage transmission and drive system components or impede the flow of lubricant to a hazardous degree; and
(ii) Be equipped with a bypass constructed and installed so that—
(A) The lubricant will flow at the normal rate through the rest of the system with the strainer or filter completely blocked; and
(B) The release of collected contaminants is minimized by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flowpath;
(iii) Be equipped with a means to indicate collection of contaminants on the filter or strainer at or before opening of the bypass;
(2) For each lubricant tank or sump outlet supplying lubrication to rotor drive systems and rotor drive system components, a screen to prevent entrance into the lubrication system of any object that might obstruct the flow of lubricant from the outlet to the filter required by paragraph (b)(1) of this section. The requirements of paragraph (b)(1) of this section do not apply to screens installed at lubricant tank or sump outlets.
(c) Splash type lubrication systems for rotor drive system gearboxes must comply with §§ 29.1021 and 29.1337(d).
§ 29.1041   General.
(a) 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 for which certification is requested, and after normal engine or auxiliary power unit shutdown, or both.
(b) There must be cooling provisions to maintain the fluid temperatures in any power transmission within safe values under any critical surface (ground or water) and flight operating conditions.
(c) Except for ground-use-only auxiliary power units, compliance with paragraphs (a) and (b) of this section must be shown by flight tests in which the temperatures of selected powerplant component and auxiliary power unit component, engine, and transmission fluids are obtained under the conditions prescribed in those paragraphs.
§ 29.1043   Cooling tests.
(a) General. For the tests prescribed in § 29.1041(c), the following apply:
(1) If the tests are conducted under conditions deviating from the maximum ambient atmospheric temperature specified in paragraph (b) of this section, the recorded powerplant temperatures must be corrected under paragraphs (c) and (d) of this section, unless a more rational correction method is applicable.
(2) No corrected temperature determined under paragraph (a)(1) of this section may exceed established limits.
(3) The fuel used during the cooling tests must be of the minimum grade approved for the engines, and the mixture settings must be those used in normal operation.
(4) The test procedures must be as prescribed in §§ 29.1045 through 29.1049.
(5) For the purposes of the cooling tests, a temperature is “stabilized” when its rate of change is less than 2 °F per minute.
(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. 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.
§ 29.1045   Climb cooling test procedures.
(a) Climb cooling tests must be conducted under this section for—
(1) Category A rotorcraft; and
(2) Multiengine category B rotorcraft for which certification is requested under the category A powerplant installation requirements, and under the requirements of § 29.861(a) at the steady rate of climb or descent established under § 29.67(b).
(b) The climb or descent cooling tests must be conducted with the engine inoperative that produces the most adverse cooling conditions for the remaining engines and powerplant components.
(c) Each operating engine must—
(1) For helicopters for which the use of 30-minute OEI power is requested, be at 30-minute OEI power for 30 minutes, and then at maximum continuous power (or at full throttle when above the critical altitude);
(2) For helicopters for which the use of continuous OEI power is requested, be at continuous OEI power (or at full throttle when above the critical altitude); and
(3) For other rotorcraft, be at maximum continuous power (or at full throttle when above the critical altitude).
(d) After temperatures have stabilized in flight, the climb must be—
(1) Begun from an altitude not greater than the lower of—
(i) 1,000 feet below the engine critcal altitude; and
(ii) 1,000 feet below the maximum altitude at which the rate of climb is 150 f.p.m; and
(2) Continued for at least five minutes after the occurrence of the highest temperature recorded, or until the rotorcraft reaches the maximum altitude for which certification is requested.
(e) For category B rotorcraft without a positive rate of climb, the descent must begin at the all-engine-critical altitude and end at the higher of—
(1) The maximum altitude at which level flight can be maintained with one engine operative; and
(2) Sea level.
(f) The climb or descent must be conducted at an airspeed representing a normal operational practice for the configuration being tested. However, if the cooling provisions are sensitive to rotorcraft speed, the most critical airspeed must be used, but need not exceed the speeds established under § 29.67(a)(2) or § 29.67(b). The climb cooling test may be conducted in conjunction with the takeoff cooling test of § 29.1047.
§ 29.1047   Takeoff cooling test procedures.
(a) Category A. For each category A rotorcraft, cooling must be shown during takeoff and subsequent climb as follows:
(1) Each temperature must be stabilized while hovering in ground effect with—
(i) The power necessary for hovering;
(ii) The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2) After the temperatures have stabilized, a climb must be started at the lowest practicable altitude and must be conducted with one engine inoperative.
(3) The operating engines must be at the greatest power for which approval is sought (or at full throttle when above the critical altitude) for the same period as this power is used in determining the takeoff climbout path under § 29.59.
(4) At the end of the time interval prescribed in paragraph (b)(3) of this section, the power must be changed to that used in meeting § 29.67(a)(2) and the climb must be continued for—
(i) Thirty minutes, if 30-minute OEI power is used; or
(ii) At least 5 minutes after the occurrence of the highest temperature recorded, if continuous OEI power or maximum continuous power is used.
(5) The speeds must be those used in determining the takeoff flight path under § 29.59.
(b) Category B. For each category B rotorcraft, cooling must be shown during takeoff and subsequent climb as follows:
(1) Each temperature must be stabilized while hovering in ground effect with—
(i) The power necessary for hovering;
(ii) The appropriate cowl flap and shutter settings; and
(iii) The maximum weight.
(2) After the temperatures have stabilized, a climb must be started at the lowest practicable altitude with takeoff power.
(3) Takeoff power must be used for the same time interval as takeoff power is used in determining the takeoff flight path under § 29.63.
(4) At the end of the time interval prescribed in paragraph (a)(3) of this section, the power must be reduced to maximum continuous power and the climb must be continued for at least five minutes after the occurance of the highest temperature recorded.
(5) The cooling test must be conducted at an airspeed corresponding to normal operating practice for the configuration being tested. However, if the cooling provisions are sensitive to rotorcraft speed, the most critical airspeed must be used, but need not exceed the speed for best rate of climb with maximum continuous power.
§ 29.1049   Hovering cooling test procedures.
The hovering cooling provisions must be shown—
(a) At maximum weight or at the greatest weight at which the rotorcraft can hover (if less), at sea level, with the power required to hover but not more than maximum continuous power, in the ground effect in still air, until at least five minutes after the occurrence of the highest temperature recorded; and
(b) With maximum continuous power, maximum weight, and at the altitude resulting in zero rate of climb for this configuration, until at least five minutes after the occurrence of the highest temperature recorded.
§ 29.1091   Air induction.
(a) The air induction system for each engine and auxiliary power unit must supply the air required by that engine and auxiliary power unit under the operating conditions for which certification is requested.
(b) Each engine and auxiliary power unit air induction system must provide air for proper fuel metering and mixture distribution with the induction system valves in any position.
(c) No air intake may open within the engine accessory section or within other areas of any powerplant compartment where emergence of backfire flame would constitute a fire hazard.
(d) Each reciprocating engine must have an alternate air source.
(e) Each alternate air intake must be located to prevent the entrance of rain, ice, or other foreign matter.
(f) For turbine engine powered rotorcraft and rotorcraft 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 air inlet ducts must be located or protected so as to minimize the ingestion of foreign matter during takeoff, landing, and taxiing.
§ 29.1093   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., and with the engines at 60 percent of maximum continuous power—
(1) Each rotorcraft with sea level engines using conventional venturi carburetors has a preheater that can provide a heat rise of 90 °F.;
(2) Each rotorcraft with sea level engines using carburetors tending to prevent icing has a preheater that can provide a heat rise of 70 °F.;
(3) Each rotorcraft with altitude engines using conventional venturi carburetors has a preheater that can provide a heat rise of 120 °F.; and
(4) Each rotorcraft with altitude engines using carburetors tending to prevent icing has a preheater that can provide a heat rise of 100 °F.
(b) Turbine engines. (1) It must be shown that each turbine engine and its air inlet system can operate throughout the flight power range of the engine (including idling)—
(i) Without accumulating ice on engine or inlet system components that would adversely affect engine operation or cause a serious loss of power under the icing conditions specified in appendix C of this Part; and
(ii) In snow, both falling and blowing, without adverse effect on engine operation, within the limitations established for the rotorcraft.
(2) Each turbine engine must idle for 30 minutes on the ground, with the air bleed available for engine icing protection at its critical condition, without adverse effect, in an atmosphere that is at a temperature between 15° and 30 °F (between −9° and −1 °C) and has a liquid water content not less than 0.3 grams per cubic meter in the form of drops having a mean effective diameter not less than 20 microns, followed by momentary operation at takeoff power or thrust. During the 30 minutes of idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Administrator.
(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 operation condition because of supercharging.
§ 29.1101   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.
§ 29.1103   Induction systems 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 duct must be strong enough to prevent induction system failure from normal backfire conditions.
(c) Each duct connected to components between which relative motion could exist must have means for flexibility.
(d) Each duct within any fire zone for which a fire-extinguishing system is required must be at least—
(1) Fireproof, if it passes through any firewall; or
(2) Fire resistant, for other ducts, except that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone.
(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 rotorcraft 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.
§ 29.1105   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.
§ 29.1107   Inter-coolers and after-coolers.
Each inter-cooler and after-cooler must be able to withstand the vibration, inertia, and air pressure loads to which it would be subjected in operation.
§ 29.1109   Carburetor air cooling.
It must be shown under § 29.1043 that each installation using two-stage superchargers has means to maintain the air temperature, at the carburetor inlet, at or below the maximum established value.
§ 29.1121   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.
(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 upon which hot exhaust gases could impinge, or that could be subjected to high temperatures from exhaust system parts, must be fireproof. Each exhaust system component must be separated by a fireproof shield from adjacent parts of the rotorcraft 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 outside the shroud.
(h) If significant traps exist, each turbine engine exhaust system must have drains discharging clear of the rotorcraft, in any normal ground and flight attitudes, to prevent fuel accumulation after the failure of an attempted engine start.
§ 29.1123   Exhaust piping.
(a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures.
(b) Exhaust piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation.
(c) Exhaust piping connected to components between which relative motion could exist must have provisions for flexibility.
§ 29.1125   Exhaust heat exchangers.
For reciprocating engine powered rotorcraft the following apply:
(a) Each exhaust heat exchanger must be constructed and installed to withstand the vibration, inertia, and other loads 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 inspecting 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 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 used by personnel—
(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 prevent harmful contamination of the ventilating air.
§ 29.1141   Powerplant controls: general.
(a) Powerplant controls must be located and arranged under § 29.777 and marked under § 29.1555.
(b) Each control must be located so that it cannot be inadvertently operated by persons entering, leaving, or moving normally in the cockpit.
(c) Each flexible powerplant control must be approved.
(d) Each control must be able to maintain any set position without—
(1) Constant attention; or
(2) Tendency to creep due to control loads or vibration.
(e) Each control must be able to withstand operating loads without excessive deflection.
(f) Controls of powerplant valves required for safety must have—
(1) For manual valves, positive stops or in the case of fuel valves suitable index provisions, in the open and closed position; and
(2) For power-assisted valves, a means to indicate to the flight crew when the valve—
(i) Is in the fully open or fully closed position; or
(ii) Is moving between the fully open and fully closed position.
§ 29.1142   Auxiliary power unit controls.
Means must be provided on the flight deck for starting, stopping, and emergency shutdown of each installed auxiliary power unit.
§ 29.1143   Engine controls.
(a) There must be a separate power control for each engine.
(b) Power controls must be arranged to allow ready synchronization of all engines by—
(1) Separate control of each engine; and
(2) Simultaneous control of all engines.
(c) Each power control must provide a positive and immediately responsive means of controlling its engine.
(d) Each fluid injection control other than fuel system control must be in the corresponding power control. However, the injection system pump may have a separate control.
(e) If a power 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.
(f) For rotorcraft to be certificated for a 30-second OEI power rating, a means must be provided to automatically activate and control the 30-second OEI power and prevent any engine from exceeding the installed engine limits associated with the 30-second OEI power rating approved for the rotorcraft.
§ 29.1145   Ignition switches.
(a) Ignition switches must control each 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.
§ 29.1147   Mixture controls.
(a) If there are mixture controls, each engine must have a separate control, and the controls must be 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.
§ 29.1151   Rotor brake controls.
(a) It must be impossible to apply the rotor brake inadvertently in flight.
(b) There must be means to warn the crew if the rotor brake has not been completely released before takeoff.
§ 29.1157   Carburetor air temperature controls.
There must be a separate carburetor air temperature control for each engine.
§ 29.1159   Supercharger controls.
Each supercharger control must be accessible to—
(a) The pilots; or
(b) (If there is a separate flight engineer station with a control panel) the flight engineer.
§ 29.1163   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 in such a way as 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 igniting flammable fluids or vapors.
(c) If continued rotation of an engine-driven cabin supercharger or any remote accessory driven by the engine will be a hazard if they malfunction, there must be means to prevent their hazardous rotation without interfering with the continued operation of the engine.
(d) Unless other means are provided, torque limiting means must be provided for accessory drives located on any component of the transmission and rotor drive system to prevent damage to these components from excessive accessory load.
§ 29.1165   Engine ignition systems.
(a) Each battery ignition system must be supplemented with 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 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 any firewall must be installed, located, or protected, to minimize the probability of the simultaneous failure of two or more wires as a result of mechanical damage, electrical fault, 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 that is not used for assisting, controlling, or analyzing the operation of that system.
(g) There must be means to warn appropriate crewmembers if the malfunctioning of any part of the electrical system is causing the continuous discharge of any battery necessary for engine ignition.
§ 29.1181   Designated fire zones: regions included.
(a) Designated fire zones are—
(1) The engine power section of reciprocating engines;
(2) The engine accessory section of reciprocating engines;
(3) Any complete powerplant compartment in which there is no isolation between the engine power section and the engine accessory section, for reciprocating engines;
(4) Any auxiliary power unit compartment;
(5) Any fuel-burning heater and other combustion equipment installation described in § 29.859;
(6) The compressor and accessory sections of turbine engines; and
(7) The combustor, turbine, and tailpipe sections of turbine engine installations except sections that do not contain lines and components carrying flammable fluids or gases and are isolated from the designated fire zone prescribed in paragraph (a)(6) of this section by a firewall that meets § 29.1191.
(b) Each designated fire zone must meet the requirements of §§ 29.1183 through 29.1203.
§ 29.1183   Lines, fittings, and 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 so as 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.
§ 29.1185   Flammable fluids.
(a) No tank or reservoir that is 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 and its supports, the shutoff means, and the connections, lines, and controls provide a degree of safety equal to that which would exist if the tank or reservoir were outside such a zone.
(b) Each fuel tank must be isolated from the engines by a firewall or shroud.
(c) 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, unless equivalent means are used to prevent heat transfer from the fire zone to the flammable fluid.
(d) Absorbent material close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids.
§ 29.1187   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 malfunction 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) Ventilation means must be arranged so that no discharged vapors will cause an additional fire hazard.
(e) For category A rotorcraft, there must be means to allow the crew to shut off the sources of forced ventilation in any fire zone (other than the engine power section of the powerplant compartment) unless the amount of extinguishing agent and the rate of discharge are based on the maximum airflow through that zone.
§ 29.1189   Shutoff means.
(a) There must be means to shut off or otherwise prevent hazardous quantities of fuel, oil, de-icing fluid, and other flammable fluids from flowing into, within, or through any designated fire zone, except that this means need not be provided—
(1) For lines, fittings, and components forming an integral part of an engine;
(2) For oil systems for turbine engine installations in which all components of the system, including oil tanks, are fireproof or located in areas not subject to engine fire conditions; or
(3) For engine oil systems in category B rotorcraft using reciprocating engines of less than 500 cubic inches displacement.
(b) The closing of any fuel shutoff valve for any engine may not make fuel unavailable to the remaining engines.
(c) For category A rotorcraft, no hazardous quantity of flammable fluid may drain into any designated fire zone after shutoff has been accomplished, nor may the closing of any fuel shutoff valve for an engine make fuel unavailable to the remaining engines.
(d) The operation of any shutoff may not interfere with the later emergency operation of any other equipment, such as the means for declutching the engine from the rotor drive.
(e) Each shutoff valve and its control must be designed, located, and protected to function properly under any condition likely to result from fire in a designated fire zone.
(f) Except for ground-use-only auxiliary power unit installations, there must be means to prevent inadvertent operation of each shutoff and to make it possible to reopen it in flight after it has been closed.
§ 29.1191   Firewalls.
(a) Each engine, including the combustor, turbine, and tailpipe sections of turbine engine installations, must be isolated by a firewall, shroud, or equivalent means, from personnel compartments, structures, controls, rotor mechanisms, and other parts that are—
(1) Essential to controlled flight and landing; and
(2) Not protected under § 29.861.
(b) Each auxiliary power unit, combustion heater, and other combustion equipment to be used in flight, must be isolated from the rest of the rotorcraft by firewalls, shrouds, or equivalent means.
(c) Each firewall or shroud must be constructed so that no hazardous quantity of air, fluid, or flame can pass from any engine compartment to other parts of the rotorcraft.
(d) Each opening in the firewall or shroud must be sealed with close-fitting fireproof grommets, bushings, or firewall fittings.
(e) Each firewall and shroud must be fireproof and protected against corrosion.
(f) In meeting this section, account must be taken of the probable path of a fire as affected by the airflow in normal flight and in autorotation.
§ 29.1193   Cowling and engine compartment covering.
(a) Each cowling and engine compartment covering must be constructed and supported so that it can resist the vibration, inertia, and air loads to which it may be subjected in operation.
(b) Cowling must meet the drainage and ventilation requirements of § 29.1187.
(c) On rotorcraft 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 § 29.1191.
(d) Each part of the cowling or engine compartment covering subject to high temperatures due to its nearness to exhaust system parts or exhaust gas impingement must be fireproof.
(e) Each rotorcraft must—
(1) Be designated 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 the requirements of 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 or burns out of any designated fire zone.
(f) A means of retention for each openable or readily removable panel, cowling, or engine or rotor drive system covering must be provided to preclude hazardous damage to rotors or critical control components in the event of—
(1) Structural or mechanical failure of the normal retention means, unless such failure is extremely improbable; or
(2) Fire in a fire zone, if such fire could adversely affect the normal means of retention.
§ 29.1194   Other surfaces.
All surfaces aft of, and near, engine compartments and designated fire zones, other than tail surfaces not subject to heat, flames, or sparks emanating from a designated fire zone or engine compartment, must be at least fire resistant.
§ 29.1195   Fire extinguishing systems.
(a) Each turbine engine powered rotorcraft and Category A reciprocating engine powered rotorcraft, and each Category B reciprocating engine powered rotorcraft with engines of more than 1,500 cubic inches must have a fire extinguishing system for the designated fire zones. The fire extinguishing system for a powerplant must be able to simultaneously protect all zones of the powerplant compartment for which protection is provided.
(b) For multiengine powered rotorcraft, the fire extinguishing system, the quantity of extinguishing agent, and the rate of discharge must—
(1) For each auxiliary power unit and combustion equipment, provide at least one adequate discharge; and
(2) For each other designated fire zone, provide two adequate discharges.
(c) For single engine rotorcraft, the quantity of extinguishing agent and the rate of discharge must provide at least one adequate discharge for the engine compartment.
(d) It must be shown by either actual or simulated flight tests that under critical airflow conditions in flight the discharge of the extinguishing agent in each designated fire zone will provide an agent concentration capable of extinguishing fires in that zone and of minimizing the probability of reignition.
§ 29.1197   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, it must be shown by test that entry of harmful concentrations of fluid or fluid vapors into any personnel compartment (due to leakage during normal operation of the rotorcraft, or discharge on the ground or in flight) is prevented, even though a defect may exist in the extinguishing system.
§ 29.1199   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 rotorcraft. 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.
§ 29.1201   Fire extinguishing system materials.
(a) No materials 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.
§ 29.1203   Fire detector systems.
(a) For each turbine engine powered rotorcraft and Category A reciprocating engine powered rotorcraft, and for each Category B reciprocating engine powered rotorcraft with engines of more than 900 cubic inches displacement, there must be approved, quick-acting fire detectors in designated fire zones and in the combustor, turbine, and tailpipe sections of turbine installations (whether or not such sections are designated fire zones) in numbers and locations ensuring prompt detection of fire in those zones.
(b) Each fire detector must be constructed and installed to withstand any vibration, inertia, and other loads to which it would be subjected in operation.
(c) No fire detector may be affected by any oil, water, other fluids, or fumes that might be present.
(d) There must be means to allow crewmembers to check, in flight, the functioning of each fire detector system electrical circuit.
(e) The writing and other components of each fire detector system in an engine compartment must be at least fire resistant.
(f) No fire 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) The zones involved are simultaneously protected by the same detector and extinguishing systems.
§ 29.1301   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;
(c) Be installed according to limitations specified for that equipment; and
(d) Function properly when installed.
§ 29.1303   Flight and navigation instruments.
The following are required flight and navigational instruments:
(a) An airspeed indicator. For Category A rotorcraft with V NE less than a speed at which unmistakable pilot cues provide overspeed warning, a maximum allowable airspeed indicator must be provided. If maximum allowable airspeed varies with weight, altitude, temperature, or r.p.m., the indicator must show that variation.
(b) A sensitive altimeter.
(c) A magnetic direction indicator.
(d) A clock displaying hours, minutes, and seconds with a sweep-second pointer or digital presentation.
(e) A free-air temperature indicator.
(f) A non-tumbling gyroscopic bank and pitch indicator.
(g) A gyroscopic rate-of-turn indicator combined with an integral slip-skid indicator (turn-and-bank indicator) except that only a slip-skid indicator is required on rotorcraft with a third attitude instrument system that—
(1) Is usable through flight attitudes of ±80 degrees of pitch and ±120 degrees of roll;
(2) Is powered from a source independent of the electrical generating system;
(3) Continues reliable operation for a minimum of 30 minutes after total failure of the electrical generating system;
(4) Operates independently of any other attitude indicating system;
(5) Is operative without selection after total failure of the electrical generating system;
(6) Is located on the instrument panel in a position acceptable to the Administrator that will make it plainly visible to and useable by any pilot at his station; and
(7) Is appropriately lighted during all phases of operation.
(h) A gyroscopic direction indicator.
(i) A rate-of-climb (vertical speed) indicator.
(j) For Category A rotorcraft, a speed warning device when V NE is less than the speed at which unmistakable overspeed warning is provided by other pilot cues. The speed warning device must give effective aural warning (differing distinctively from aural warnings used for other purposes) to the pilots whenever the indicated speed exceeds V NE plus 3 knots and must operate satisfactorily throughout the approved range of altitudes and temperatures.
§ 29.1305   Powerplant instruments.
The following are required powerplant instruments:
(a) For each rotorcraft—
(1) A carburetor air temperature indicator for each reciprocating engine;
(2) A cylinder head temperature indicator for each air-cooled reciprocating engine, and a coolant temperature indicator for each liquid-cooled reciprocating engine;
(3) A fuel quantity indicator for each fuel tank;
(4) A low fuel warning device for each fuel tank which feeds an engine. This device must—
(i) Provide a warning to the crew when approximately 10 minutes of usable fuel remains in the tank; and
(ii) Be independent of the normal fuel quantity indicating system.
(5) A means to indicate manifold pressure for each altitude engine;
(6) An oil pressure indicator for each pressure-lubricated gearbox.
(7) An oil pressure warning device for each pressure-lubricated gearbox to indicate when the oil pressure falls below a safe value;
(8) An oil quantity indicator for each oil tank and each rotor drive gearbox, if lubricant is self-contained;
(9) An oil temperature indicator for each engine;
(10) An oil temperature warning device to indicate unsafe oil temperatures in each main rotor drive gearbox, including gearboxes necessary for rotor phasing;
(11) A means to indicate the gas temperature for each turbine engine;
(12) A means to indicate the gas producer speed for each turbine engine;
(13) A tachometer for each engine that, if combined with the applicable instrument required by paragraph (a)(14) of this section, indicates rotor r.p.m. during autorotation.
(14) At least one tachometer to indicate, as applicable—
(i) The r.p.m. of the single main rotor;
(ii) The common r.p.m. of any main rotors whose speeds cannot vary appreciably with respect to each other; and
(iii) The r.p.m. of each main rotor whose speed can vary appreciably with respect to that of another main rotor;
(15) A free power turbine tachometer for each turbine engine;
(16) A means, for each turbine engine, to indicate power for that engine;
(17) For each turbine engine, an indicator to indicate the functioning of the powerplant ice protection system;
(18) An indicator for the filter required by § 29.997 to indicate the occurrence of contamination of the filter to the degree established in compliance with § 29.955;
(19) For each turbine engine, a warning means for the oil strainer or filter required by § 29.1019, if it has no bypass, to warn the pilot of the occurrence of contamination of the strainer or filter before it reaches the capacity established in accordance with § 29.1019(a)(2);
(20) An indicator to indicate the functioning of any selectable or controllable heater used to prevent ice clogging of fuel system components;
(21) An individual fuel pressure indicator for each engine, unless the fuel system which supplies that engine does not employ any pumps, filters, or other components subject to degradation or failure which may adversely affect fuel pressure at the engine;
(22) A means to indicate to the flightcrew the failure of any fuel pump installed to show compliance with § 29.955;
(23) Warning or caution devices to signal to the flightcrew when ferromagnetic particles are detected by the chip detector required by § 29.1337(e); and
(24) For auxiliary power units, an individual indicator, warning or caution device, or other means to advise the flightcrew that limits are being exceeded, if exceeding these limits can be hazardous, for—
(i) Gas temperature;
(ii) Oil pressure; and
(iii) Rotor speed.
(25) For rotorcraft for which a 30-second/2-minute OEI power rating is requested, a means must be provided to alert the pilot when the engine is at the 30-second and 2-minute OEI power levels, when the event begins, and when the time interval expires.
(26) For each turbine engine utilizing 30-second/2-minute OEI power, a device or system must be provided for use by ground personnel which—
(i) Automatically records each usage and duration of power at the 30-second and 2-minute OEI levels;
(ii) Permits retrieval of the recorded data;
(iii) Can be reset only by ground maintenance personnel; and
(iv) Has a means to verify proper operation of the system or device.
(b) For category A rotorcraft—
(1) An individual oil pressure indicator for each engine, and either an independent warning device for each engine or a master warning device for the engines with means for isolating the individual warning circuit from the master warning device;
(2) An independent fuel pressure warning device for each engine or a master warning device for all engines with provision for isolating the individual warning device from the master warning device; and
(3) Fire warning indicators.
(4) For each Category A rotorcraft for which OEI Training Mode is requested, a means must be provided to indicate to the pilot the simulation of an engine failure, the annunciation of that simulation, and a representation of the OEI power being provided.
(c) For category B rotorcraft—
(1) An individual oil pressure indicator for each engine; and
(2) Fire warning indicators, when fire detection is required.
§ 29.1307   Miscellaneous equipment.
The following is required miscellaneous equipment:
(a) An approved seat for each occupant.
(b) A master switch arrangement for electrical circuits other than ignition.
(c) Hand fire extinguishers.
(d) A windshield wiper or equivalent device for each pilot station.
(e) A two-way radio communication system.
§ 29.1309   Equipment, systems, and installations.
The equipment, systems, and installations whose functioning is required by this subchapter must be designed and installed to ensure that they perform their intended functions under any foreseeable operating condition. For any item of equipment or system whose failure has not been specifically addressed by another requirement in this chapter, the following requirements also apply:
(a) The design of each item of equipment, system, and installation must be analyzed separately and in relation to other rotorcraft systems and installations to determine and identify any failure that would affect the capability of the rotorcraft or the ability of the crew to perform their duties in all operating conditions.
(b) Each item of equipment, system, and installation must be designed and installed so that:
(1) The occurrence of any catastrophic failure condition is extremely improbable;
(2) The occurrence of any major failure condition is no more than improbable; and
(3) For the occurrence of any other failure condition in between major and catastrophic, the probability of the failure condition must be inversely proportional to its consequences.
(c) A means to alert the crew in the event of a failure must be provided when an unsafe system operating condition exists and to enable them to take corrective action. Systems, controls, and associated monitoring and crew alerting means must be designed to minimize crew errors that could create additional hazards.
(d) Compliance with the requirements of this section must be shown by analysis and, where necessary, by ground, flight, or simulator tests. The analysis must account for:
(1) Possible modes of failure, including malfunctions and misleading data and input from external sources;
(2) The effect of multiple failures and latent failures;
(3) The resulting effects on the rotorcraft and occupants, considering the stage of flight and operating conditions; and
(4) The crew alerting cues and the corrective action required.
§ 29.1316   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 rotorcraft, must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to lightning; and
(2) The system automatically recovers normal operation of that function in a timely manner after the rotorcraft is exposed to lightning.
(b) Each electrical and electronic system that performs a function, for which failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition, must be designed and installed so that the function recovers normal operation in a timely manner after the rotorcraft is exposed to lightning.
§ 29.1317   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 rotorcraft must be designed and installed so that—
(1) The function is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment I, as described in appendix E to this part;
(2) The system automatically recovers normal operation of that function, in a timely manner, after the rotorcraft is exposed to HIRF environment I, as described in appendix E to this part, unless this conflicts with other operational or functional requirements of that system;
(3) The system is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment II, as described in appendix E to this part; and
(4) Each function required during operation under visual flight rules is not adversely affected during and after the time the rotorcraft is exposed to HIRF environment III, as described in appendix E to this part.
(b) Each electrical and electronic system that performs a function whose failure would significantly reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 1 or 2, as described in appendix E to this part.
(c) Each electrical and electronic system that performs such a function whose failure would reduce the capability of the rotorcraft or the ability of the flightcrew to respond to an adverse operating condition must be designed and installed so the system is not adversely affected when the equipment providing these functions is exposed to equipment HIRF test level 3, as described in appendix E to this part.
(d) Before December 1, 2012, an electrical or electronic system that performs a function whose failure would prevent the continued safe flight and landing of a rotorcraft may be designed and installed without meeting the provisions of paragraph (a) provided—
(1) The system has previously been shown to comply with special conditions for HIRF, prescribed under § 21.16, issued before December 1, 2007;
(2) The HIRF immunity characteristics of the system have not changed since compliance with the special conditions was demonstrated; and
(3) The data used to demonstrate compliance with the special conditions is provided.
§ 29.1321   Arrangement and visibility.
(a) Each flight, navigation, and powerplant instrument for use by any pilot must be easily visible to him from his station with the minimum practicable deviation from his normal position and line of vision when he is looking forward along the flight path.
(b) Each instrument necessary for safe operation, including the airspeed indicator, gyroscopic direction indicator, gyroscopic bank-and-pitch indicator, slip-skid indicator, altimeter, rate-of-climb indicator, rotor tachometers, and the indicator most representative of engine power, must be grouped and centered as nearly as practicable about the vertical plane of the pilot's forward vision. In addition, for rotorcraft approved for IFR flight—
(1) The instrument that most effectively indicates attitude must be on the panel in the top center position;
(2) The instrument that most effectively indicates direction of flight must be adjacent to and directly below the attitude instrument;
(3) The instrument that most effectively indicates airspeed must be adjacent to and to the left of the attitude instrument; and
(4) The instrument that most effectively indicates altitude or is most frequently utilized in control of altitude must be adjacent to and to the right of the attitude instrument.
(c) Other required powerplant instruments must be closely grouped on the instrument panel.
(d) Identical powerplant instruments for the engines must be located so as to prevent any confusion as to which engine each instrument relates.
(e) Each powerplant instrument vital to safe operation must be plainly visible to appropriate crewmembers.
(f) Instrument panel vibration may not damage, or impair the readability or accuracy of, any instrument.
(g) If a visual indicator is provided to indicate malfunction of an instrument, it must be effective under all probable cockpit lighting conditions.
§ 29.1322   Warning, caution, and advisory lights.
If warning, caution or advisory lights are installed in the cockpit they must, unless otherwise approved by the Administrator, be—
(a) Red, for warning lights (lights indicating a hazard which may require immediate corrective action);
(b) Amber, for caution lights (lights indicating the possible need for future corrective action);
(c) Green, for safe operation lights; and
(d) Any other color, including white, for lights not described in paragraphs (a) through (c) of this section, provided the color differs sufficiently from the colors prescribed in paragraphs (a) through (c) of this section to avoid possible confusion.
§ 29.1323   Airspeed indicating system.
For each airspeed indicating system, the following apply:
(a) Each airspeed indicating instrument must be calibrated to indicate true airspeed (at sea level with a standard atmosphere) with a minimum practicable instrument calibration error when the corresponding pitot and static pressures are applied.
(b) Each system must be calibrated to determine system error excluding airspeed instrument error. This calibration must be determined—
(1) In level flight at speeds of 20 knots and greater, and over an appropriate range of speeds for flight conditions of climb and autorotation; and
(2) During takeoff, with repeatable and readable indications that ensure—
(i) Consistent realization of the field lengths specified in the Rotorcraft Flight Manual; and
(ii) Avoidance of the critical areas of the height-velocity envelope as established under § 29.87.
(c) For Category A rotorcraft—
(1) The indication must allow consistent definition of the takeoff decision point; and
(2) The system error, excluding the airspeed instrument calibration error, may not exceed—
(i) Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of takeoff safety speed; and
(ii) Ten knots in climb at speeds from 10 knots below takeoff safety speed to 10 knots above V Y .
(d) For Category B rotorcraft, the system error, excluding the airspeed instrument calibration error, may not exceed 3 percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of the climbout speed attained at 50 feet when complying with § 29.63.
(e) Each system must be arranged, so far as practicable, to prevent malfunction or serious error due to the entry of moisture, dirt, or other substances.
(f) Each system must have a heated pitot tube or an equivalent means of preventing malfunction due to icing.
§ 29.1325   Static pressure and pressure altimeter systems.
(a) Each instrument with static air case connections must be vented to the outside atmosphere through an appropriate piping system.
(b) Each vent must be located where its orifices are least affected by airflow variation, moisture, or foreign matter.
(c) Each static pressure port must be designed and located in such manner that the correlation between air pressure in the static pressure system and true ambient atmospheric static pressure is not altered when the rotorcraft encounters icing conditions. An anti-icing means or an alternate source of static pressure may be used in showing compliance with this requirement. If the reading of the altimeter, when on the alternate static pressure system, differs from the reading of altimeter when on the primary static system by more than 50 feet, a correction card must be provided for the alternate static system.
(d) Except for the vent into the atmosphere, each system must be airtight.
(e) Each pressure altimeter must be approved and calibrated to indicate pressure altitude in a standard atmosphere with a minimum practicable calibration error when the corresponding static pressures are applied.
(f) Each system must be designed and installed so that an error in indicated pressure altitude, at sea level, with a standard atmosphere, excluding instrument calibration error, does not result in an error of more than ±30 feet per 100 knots speed. However, the error need not be less than ±30 feet.
(g) Except as provided in paragraph (h) of this section, if the static pressure system incorporates both a primary and an alternate static pressure source, the means for selecting one or the other source must be designed so that—
(1) When either source is selected, the other is blocked off; and
(2) Both sources cannot be blocked off simultaneously.
(h) For unpressurized rotorcraft, paragraph (g)(1) of this section does not apply if it can be demonstrated that the static pressure system calibration, when either static pressure source is selected, is not changed by the other static pressure source being open or blocked.
§ 29.1327   Magnetic direction indicator.
(a) Each magnetic direction indicator must be installed so that its accuracy is not excessively affected by the rotorcraft's vibration or magnetic fields.
(b) The compensated installation may not have a deviation, in level flight, greater than 10 degrees on any heading.
§ 29.1329   Automatic pilot and flight guidance system.
For the purpose of this subpart, an automatic pilot and flight guidance system may consist of an autopilot, flight director, or a component that interacts with stability augmentation or trim.
(a) Each automatic pilot and flight guidance system must be designed so that it:
(1) Can be overpowered by one pilot to allow control of the rotorcraft;
(2) Provides a means to disengage the system, or any malfunctioning component of the system, by each pilot to prevent it from interfering with the control of the rotorcraft; and
(3) Provides a means to indicate to the flight crew its current mode of operation. Selector switch position is not acceptable as a means of indication.
(b) Unless there is automatic synchronization, each system must have a means to readily indicate to the pilot the alignment of the actuating device in relation to the control system it operates.
(c) Each manually operated control for the system's operation must be readily accessible to the pilots.
(d) The system must be designed so that, within the range of adjustment available to the pilot, it cannot produce hazardous loads on the rotorcraft, or create hazardous deviations in the flight path, under any flight condition appropriate to its use or in the event of a malfunction, assuming that corrective action begins within a reasonable period of time.
(e) If the automatic pilot and flight guidance system integrates signals from auxiliary controls or furnishes signals for operation of other equipment, there must be a means to prevent improper operation.
(f) If the automatic pilot system can be coupled to airborne navigation equipment, means must be provided to indicate to the pilots the current mode of operation. Selector switch position is not acceptable as a means of indication.
§ 29.1331   Instruments using a power supply.
For category A rotorcraft—
(a) Each required flight instrument using a power supply must have—
(1) Two independent sources of power;
(2) A means of selecting either power source; and
(3) A visual means integral with each instrument to indicate when the power adequate to sustain proper instrument performance is not being supplied. The power must be measured at or near the point where it enters the instrument. For electrical instruments, the power is considered to be adequate when the voltage is within the approved limits; and
(b) The installation and power supply system must be such that failure of any flight instrument connected to one source, or of the energy supply from one source, or a fault in any part of the power distribution system does not interfere with the proper supply of energy from any other source.
§ 29.1333   Instrument systems.
For systems that operate the required flight instruments which are located at each pilot's station, the following apply:
(a) For pneumatic systems, only the required flight instruments for the first pilot may be connected to that operating system.
(b) The equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the flight instruments remains available to a pilot, without additional crewmember action, after any single failure or combination of failures that are not shown to be extremely improbable.
(c) Additional instruments, systems, or equipment may not be connected to the operating system for a second pilot unless provisions are made to ensure the continued normal functioning of the required flight instruments in the event of any malfunction of the additional instruments, systems, or equipment which is not shown to be extremely improbable.
§ 29.1337   Powerplant instruments.
(a) Instruments and instrument lines. (1) Each powerplant and auxiliary power unit instrument line must meet the requirements of §§ 29.993 and 29.1183.
(2) Each line carrying flammable fluids under pressure must—
(i) Have restricting orifices or other safety devices at the source of pressure to prevent the escape of excessive fluid if the line fails; and
(ii) Be installed and located so that the escape of fluids would not create a hazard.
(3) Each powerplant and auxiliary power unit instrument that utilizes flammable fluids must be installed and located so that the escape of fluid would not create a hazard.
(b) Fuel quantity indicator. There must be means to indicate to the flight crew members the quantity, in gallons or equivalent units, of usable fuel in each tank during flight. In addition—
(1) Each fuel quantity indicator must be calibrated to read “zero” during level flight when the quantity of fuel remaining in the tank is equal to the unusable fuel supply determined under § 29.959;
(2) When two or more tanks are closely interconnected by a gravity feed system and vented, and when it is impossible to feed from each tank separately, at least one fuel quantity indicator must be installed;
(3) Tanks with interconnected outlets and airspaces may be treated as one tank and need not have separate indicators; and
(4) Each exposed sight gauge used as a fuel quantity indicator must be protected against damage.
(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each metering component must have a means for bypassing the fuel supply if malfunction of that component severely restricts fuel flow.
(d) Oil quantity indicator. There must be a stick gauge or equivalent means to indicate the quantity of oil—
(1) In each tank; and
(2) In each transmission gearbox.
(e) Rotor drive system transmissions and gearboxes utilizing ferromagnetic materials must be equipped with chip detectors designed to indicate the presence of ferromagnetic particles resulting from damage or excessive wear within the transmission or gearbox. Each chip detector must—
(1) Be designed to provide a signal to the indicator required by § 29.1305(a)(22); and
(2) Be provided with a means to allow crewmembers to check, in flight, the function of each detector electrical circuit and signal.
§ 29.1351   General.
(a) Electrical system capacity. The required generating capacity and the number and kind of power sources must—
(1) Be determined by an electrical load analysis; and
(2) Meet the requirements of § 29.1309.
(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 essential load equipment can be maintained within the limits for which the equipment is designed, during any probable operating condition;
(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 main bus; and
(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 rotorcraft, 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 rotorcraft's electrical system.
(d) Operation with the normal electrical power generating system inoperative.
(1) It must be shown by analysis, tests, or both, that the rotorcraft can be operated safely in VFR conditions for a period of not less than 5 minutes, with the normal electrical power generating system (electrical power sources excluding the battery) inoperative, with critical type fuel (from the standpoint of flameout and restart capability), and with the rotorcraft initially at the maximum certificated altitude. Parts of the electrical system may remain on if—
(i) A single malfunction, including a wire bundle or junction box fire, cannot result in loss of the part turned off and the part turned on;
(ii) The parts turned on are electrically and mechanically isolated from the parts turned off; and
(2) Additional requirements for Category A Rotorcraft.
(i) Unless it can be shown that the loss of the normal electrical power generating system is extremely improbable, an emergency electrical power system, independent of the normal electrical power generating system, must be provided, with sufficient capacity to power all systems necessary for continued safe flight and landing.
(ii) Failures, including junction box, control panel, or wire bundle fires, which would result in the loss of the normal and emergency systems, must be shown to be extremely improbable.
(iii) Systems necessary for immediate safety must continue to operate following the loss of the normal electrical power generating system, without the need for flight crew action.
(e) Electrical 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 electrical unit or system essential to safe operation.
(f) Cables must be grouped, routed, and spaced so that damage to essential circuits will be minimized if there are faults in heavy current-carrying cables.
§ 29.1353   Energy storage systems.
Energy storage systems must be designed and installed as follows:
(a) Energy storage systems must provide automatic protective features for any conditions that could prevent continued safe flight and landing.
(b) Energy storage systems must not emit any flammable, explosive, or toxic gases, smoke, or fluids that could accumulate in hazardous quantities within the rotorcraft.
(c) Corrosive fluids or gases that escape from the system must not damage surrounding structures, adjacent equipment, or systems necessary for continued safe flight and landing.
(d) The maximum amount of heat and pressure that can be generated during any operation or under any failure condition of the energy storage system or its individual components must not result in any hazardous effect on rotorcraft structure, equipment, or systems necessary for continued safe flight and landing.
(e) Energy storage system installations required for continued safe flight and landing of the rotorcraft must have monitoring features and a means to indicate to the pilot the status of all critical system parameters.
§ 29.1355   Distribution system.
(a) The distribution system includes the distribution busses, their associated feeders, and each control and protective device.
(b) 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 provided automatically or be manually selectable to maintain equipment or system operation.
§ 29.1357   Circuit protective devices.
(a) Automatic protective devices must be used to minimize distress to the electrical system and hazard to the rotorcraft system and hazard to the rotorcraft 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 buses with sufficient rapidity to provide protection from hazardous overvoltage 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 regardless of the position of the operating control.
(d) If the ability to reset a circuit breaker or replace a fuse is essential to safety in flight, that circuit breaker or fuse must be located and identified so that it can be readily reset or replaced in flight.
(e) Each essential load 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) If fuses are used, there must be spare fuses for use in flight equal to at least 50 percent of the number of fuses of each rating required for complete circuit protection.
(g) Automatic reset circuit breakers may be used as integral protectors for electrical equipment provided there is circuit protection for the cable supplying power to the equipment.
§ 29.1359   Electrical system fire and smoke protection.
(a) Components of the electrical system must meet the applicable fire and smoke protection provisions of §§ 29.831 and 29.863.
(b) Electrical cables, terminals, and equipment, in designated fire zones, and that are used in emergency procedures, must be at least fire resistant.
(c) Insulation on electrical wire and cable installed in the rotorcraft must be self-extinguishing when tested in accordance with Appendix F, Part I(a)(3), of part 25 of this chapter.
§ 29.1363   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 rotorcraft;
(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 prime movers on the rotorcraft 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 rotorcraft, flight tests must be made.
§ 29.1381   Instrument lights.
The instrument lights must—
(a) Make each instrument, switch, and other device for which they are provided easily readable; and
(b) Be installed so that—
(1) Their direct rays are shielded from the pilot's eyes; and
(2) No objectionable reflections are visible to the pilot.
§ 29.1383   Landing lights.
(a) Each required landing or hovering light must be approved.
(b) Each landing light 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 operation, including hovering and landing.
(c) At least one separate switch must be provided, as applicable—
(1) For each separately installed landing light; and
(2) For each group of landing lights installed at a common location.
§ 29.1385   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 §§ 29.1387 through 29.1397.
(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 rotorcraft so that, with the rotorcraft 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, and must be approved.
(d) Circuit. The two forward position lights and the rear position light must make a single circuit.
(e) 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.
§ 29.1387   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 rotorcraft, 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 rotorcraft, 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 § 29.1385(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.
§ 29.1389   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 rotorcraft. 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 rotorcraft and perpendicular to the plane of symmetry of the rotorcraft), must equal or exceed the values in § 29.1391.
(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 § 29.1393 where I is the minimum intensity prescribed in § 29.1391 for the corresponding angles in the horizontal plane.
(3) Intensities in overlaps between adjacent signals. No intensity in any overlap between adjacent signals may exceed the values in § 29.1395, except that higher intensities in overlaps may be used with the use of main beam intensities substantially greater than the minima specified in §§ 29.1391 and 29.1393 if the overlap intensities in relation to the main beam intensities do not adversely affect signal clarity.
§ 29.1391   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:
§ 29.1393   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:
§ 29.1395   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 § 29.1389(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.
§ 29.1397   Color specifications.
Each position light color must have the applicable International Commission on Illumination chromaticity coordinates as follows:
(a) Aviation red —
y is not greater than 0.335; and
z is not greater than 0.002.
(b) Aviation green —
x is not greater than 0.440−0.320 y ;
x is not greater than y −0.170; and
y is not less than 0.390−0.170 x.
(c) Aviation white —
x is not less than 0.300 and not greater than 0.540;
y is not less than x −0.040 or y c −0.010, whichever is the smaller; and
y is not greater than x + 0.020 nor 0.636−0.400 x ;
Where Y e is the y coordinate of the Planckian radiator for the value of x considered.
§ 29.1399   Riding light.
(a) Each riding light required for water operation must be installed so that it can—
(1) Show a white light for at least two miles at night under clear atmospheric conditions; and
(2) Show a maximum practicable unbroken light with the rotorcraft on the water.
(b) Externally hung lights may be used.
§ 29.1401   Anticollision light system.
(a) General. If certification for night operation is requested, the rotorcraft must have an anticollision light system that—
(1) Consists of one or more approved anticollision lights located so that their emitted 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 rotorcraft, considering the physical configuration and flight characteristics of the rotorcraft. The field of coverage must extend in each direction within at least 30 degrees above and 30 degrees below the horizontal plane of the rotorcraft, except that there may be solid angles of obstructed visibility totaling not more than 0.5 steradians.
(c) Flashing characteristics. The arrangement of the system, that is, the number of light sources, beam width, speed of rotation, and other characteristics, must give an effective flash frequency of not less than 40, nor more than 100, cycles per minute. The effective flash frequency is the frequency at which the rotorcraft'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 aviation red and must meet the applicable requirements of § 29.1397.
(e) Light intensity. The minimum light intensities in any vertical plane, measured with the red filter (if used) and expressed in terms of “effective” intensities must meet the requirements of paragraph (f) of this section. The following relation must be assumed:
(f) Minimum effective intensities for anticollision light. Each anticollision light effective intensity must equal or exceed the applicable values in the following table:
§ 29.1411   General.
(a) Accessibility. Required safety equipment to be used by the crew in an emergency, such as automatic liferaft releases, 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 device required by § 29.809(f) must be at the exits for which they are intended.
(d) Liferafts. Liferafts must be stowed near exits through which the rafts can be launched during an unplanned ditching. Rafts automatically or remotely released outside the rotorcraft must be attached to the rotorcraft by the static line prescribed in § 29.1415.
(e) Long-range signaling device. The stowage provisions for the long-range signaling device required by § 29.1415 must be near an exit available during an unplanned ditching.
(f) Life preservers. Each life preserver must be within easy reach of each occupant while seated.
§ 29.1413   Safety belts: passenger warning device.
(a) If there are means to indicate to the passengers when safety belts should be fastened, they must be installed to be operated from either pilot seat.
(b) Each safety belt must be equipped with a metal to metal latching device.
§ 29.1415   Ditching equipment.
(a) Emergency flotation and signaling equipment required by any operating rule of this chapter must meet the requirements of this section.
(b) Each liferaft and each life preserver must be approved. In addition—
(1) Provide not less than two rafts, of an approximately equal rated capacity and buoyancy to accommodate the occupants of the rotorcraft; and
(2) Each raft must have a trailing line, and must have a static line designed to hold the raft near the rotorcraft but to release it if the rotorcraft 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.
§ 29.1419   Ice protection.
(a) To obtain certification for flight into icing conditions, compliance with this section must be shown.
(b) It must be demonstrated that the rotorcraft can be safely operated in the continuous maximum and intermittent maximum icing conditions determined under appendix C of this part within the rotorcraft altitude envelope. An analysis must be performed to establish, on the basis of the rotorcraft's operational needs, the adequacy of the ice protection system for the various components of the rotorcraft.
(c) In addition to the analysis and physical evaluation prescribed in paragraph (b) of this section, the effectiveness of the ice protection system and its components must be shown by flight tests of the rotorcraft or its components in measured natural atmospheric icing conditions and by one or more of the following tests as found necessary to determine the adequacy of the ice protection system:
(1) Laboratory dry air or simulated icing tests, or a combination of both, of the components or models of the components.
(2) Flight dry air tests of the ice protection system as a whole, or its individual components.
(3) Flight tests of the rotorcraft or its components in measured simulated icing conditions.
(d) The ice protection provisions of this section are considered to be applicable primarily to the airframe. Powerplant installation requirements are contained in Subpart E of this part.
(e) A means must be identified or provided for determining the formation of ice on critical parts of the rotorcraft. Unless otherwise restricted, the means must be available for nighttime as well as daytime operation. The rotorcraft flight manual must describe the means of determining ice formation and must contain information necessary for safe operation of the rotorcraft in icing conditions.
§ 29.1431   Electronic equipment.
(a) Radio communication and navigation equipment installations must be free from hazards in themselves, in their method of operation, and in their effects on other components, under any critical environmental conditions.
(b) Radio communication and navigation 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.
§ 29.1433   Vacuum systems.
(a) 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.
(b) Each vacuum air system line and fitting on the discharge side of the pump that might contain flammable vapors or fluids must meet the requirements of § 29.1183 if they are in a designated fire zone.
(c) Other vacuum air system components in designated fire zones must be at least fire resistant.
§ 29.1435   Hydraulic systems.
(a) Design. Each hydraulic system must be designed as follows:
(1) Each element of the hydraulic system must be designed to withstand, without detrimental, permanent deformation, any structural loads that may be imposed simultaneously with the maximum operating hydraulic loads.
(2) Each element of the hydraulic system must be designed to withstand pressures sufficiently greater than those prescribed in paragraph (b) of this section to show that the system will not rupture under service conditions.
(3) There must be means to indicate the pressure in each main hydraulic power system.
(4) There must be means to ensure that no pressure in any part of the system will exceed a safe limit above the maximum operating pressure of the system, and to prevent excessive pressures resulting from any fluid volumetric change in lines likely to remain closed long enough for such a change to take place. The possibility of detrimental transient (surge) pressures during operation must be considered.
(5) Each hydraulic line, fitting, and component must be installed and supported to prevent excessive vibration and to withstand inertia loads. Each element of the installation must be protected from abrasion, corrosion, and mechanical damage.
(6) Means for providing flexibility must be used to connect points, in a hydraulic fluid line, between which relative motion or differential vibration exists.
(b) Tests. Each element of the system must be tested to a proof pressure of 1.5 times the maximum pressure to which that element will be subjected in normal operation, without failure, malfunction, or detrimental deformation of any part of the system.
(c) Fire protection. Each hydraulic system using flammable hydraulic fluid must meet the applicable requirements of §§ 29.861, 29.1183, 29.1185, and 29.1189.
§ 29.1439   Protective breathing equipment.
(a) If one or more cargo or baggage compartments are to be accessible in flight, protective breathing equipment must be available for an appropriate crewmember.
(b) For protective breathing equipment required by paragraph (a) of this section or by any operating rule of this chapter—
(1) That equipment must be designed to protect the crew from smoke, carbon dioxide, and other harmful gases while on flight deck duty;
(2) That equipment must include—
(i) Masks covering the eyes, nose, and mouth; or
(ii) Masks covering the nose and mouth, plus accessory equipment to protect the eyes; and
(3) That equipment must supply protective oxygen of 10 minutes duration per crewmember at a pressure altitude of 8,000 feet with a respiratory minute volume of 30 liters per minute BTPD.
§ 29.1457   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 rotorcraft by radio.
(2) Voice communications of flight crewmembers on the flight deck.
(3) Voice communications of flight crewmembers on the flight deck, using the rotorcraft'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 may be met—
(1) 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; or
(2) By installing a continually energized or voice-actuated lip microphone at the first and second pilot stations.
(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 microphone, headset, or speaker used at the first pilot station.
(2) For the second channel, from each microphone, headset, or speaker used at the second pilot station.
(3) For the third channel, from the cockpit-mounted area microphone, or the continually energized or voice-actuated lip microphones at the first and second pilot stations.
(4) For the fourth channel, from—
(i) Each microphone, headset, or speaker used at the stations for the third and fourth crewmembers; 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.
(iii) Each microphone on the flight deck that is used with the rotorcraft's loudspeaker system if its signals are not picked up by another channel.
(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 rotorcraft.
(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) Whether the cockpit voice recorder and digital flight data recorder are installed in separate boxes or in a combination unit, no single electrical failure external to the recorder may disable both the cockpit voice recorder and the digital flight data recorder; and
(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.
(e) The record container must be located and mounted to minimize the probability of rupture of the container as a result of crash impact and consequent heat damage to the record from fire.
(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 be either bright orange or bright yellow.
(h) When both a cockpit voice recorder and a flight data recorder are required by the operating rules, one combination unit may be installed, provided that all other requirements of this section and the requirements for flight data recorders under this part are met.
§ 29.1459   Flight data recorders.
(a) Each flight recorder required by the operating rules of Subchapter G 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 §§ 29.1323, 29.1325, and 29.1327 of this part, as applicable;
(2) The vertical acceleration sensor is rigidly attached, and located longitudinally within the approved center of gravity limits of the rotorcraft;
(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 rotorcraft.
(4) There is an aural or visual means for perflight checking of the recorder for proper recording of data in the storage medium;
(5) Except for recorders powered solely by the engine-drive 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 any crash impact; and
(6) Whether the cockpit voice recorder and digital flight data recorder are installed in separate boxes or in a combination unit, no single electrical failure external to the recorder may disable both the cockpit voice recorder and the digital flight data recorder.
(b) Each nonejectable recorder 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.
(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. This correlation must cover the airspeed range over which the aircraft is to be operated, the range of altitude to which the aircraft 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 a 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 it is not likely to be separated during crash impact.
(e) When both a cockpit voice recorder and a flight data recorder are required by the operating rules, one combination unit may be installed, provided that all other requirements of this section and the requirements for cockpit voice recorders under this part are met.
§ 29.1461   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.
§ 29.1501   General.
(a) Each operating limitation specified in §§ 29.1503 through 29.1525 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 §§ 29.1541 through 29.1589.
§ 29.1503   Airspeed limitations: general.
(a) An operating speed range must be established.
(b) When airspeed limitations are a function of weight, weight distribution, altitude, rotor speed, power, or other factors, airspeed limitations corresponding with the critical combinations of these factors must be established.
§ 29.1505   Never-exceed speed.
(a) The never-exceed speed, V NE, must be established so that it is—
(1) Not less than 40 knots (CAS); and
(2) Not more than the lesser of—
(i) 0.9 times the maximum forward speeds established under § 29.309;
(ii) 0.9 times the maximum speed shown under §§ 29.251 and 29.629; or
(iii) 0.9 times the maximum speed substantiated for advancing blade tip mach number effects under critical altitude conditions.
(b) V NE may vary with altitude, r.p.m., temperature, and weight, if—
(1) No more than two of these variables (or no more than two instruments integrating more than one of these variables) are used at one time; and
(2) The ranges of these variables (or of the indications on instruments integrating more than one of these variables) are large enough to allow an operationally practical and safe variation of V NE .
(c) For helicopters, a stabilized power-off V NE denoted as V NE (power-off) may be established at a speed less than V NE established pursuant to paragraph (a) of this section, if the following conditions are met:
(1) V NE (power-off) is not less than a speed midway between the power-on V NE and the speed used in meeting the requirements of—
(i) § 29.67(a)(3) for Category A helicopters;
(ii) § 29.65(a) for Category B helicopters, except multi-engine helicopters meeting the requirements of § 29.67(b); and
(iii) § 29.67(b) for multi-engine Category B helicopters meeting the requirements of § 29.67(b).
(2) V NE (power-off) is—
(i) A constant airspeed;
(ii) A constant amount less than power-on V NE; or
(iii) A constant airspeed for a portion of the altitude range for which certification is requested, and a constant amount less than power-on V NE for the remainder of the altitude range.
§ 29.1509   Rotor speed.
(a) Maximum power-off (autorotation). The maximum power-off rotor speed must be established so that it does not exceed 95 percent of the lesser of—
(1) The maximum design r.p.m. determined under § 29.309(b); and
(2) The maximum r.p.m. shown during the type tests.
(b) Minimum power-off. The minimum power-off rotor speed must be established so that it is not less than 105 percent of the greater of—
(1) The minimum shown during the type tests; and
(2) The minimum determined by design substantiation.
(c) Minimum power-on. The minimum power-on rotor speed must be established so that it is—
(1) Not less than the greater of—
(i) The minimum shown during the type tests; and
(ii) The minimum determined by design substantiation; and
(2) Not more than a value determined under § 29.33 (a)(1) and (c)(1).
§ 29.1517   Limiting height-velocity envelope.
For Category A rotorcraft, if a range of heights exists at any speed, including zero, within which it is not possible to make a safe landing following power failure, the range of heights and its variation with forward speed must be established, together with any other pertinent information, such as the kind of landing surface.
§ 29.1519   Weight and center of gravity.
The weight and center of gravity limitations determined under §§ 29.25 and 29.27, respectively, must be established as operating limitations.
§ 29.1521   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 are type certificated.
(b) Takeoff operation. The powerplant takeoff operation must be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable manifold pressure (for reciprocating engines);
(3) The maximum allowable turbine inlet or turbine outlet gas temperature (for turbine engines);
(4) The maximum allowable power or torque for each engine, considering the power input limitations of the transmission with all engines operating;
(5) The maximum allowable power or torque for each engine considering the power input limitations of the transmission with one engine inoperative;
(6) The time limit for the use of the power corresponding to the limitations established in paragraphs (b)(1) through (5) of this section; and
(7) If the time limit established in paragraph (b)(6) of this section exceeds 2 minutes—
(i) The maximum allowable cylinder head or coolant outlet temperature (for reciprocating engines); and
(ii) The maximum allowable engine and transmission oil temperatures.
(c) Continuous operation. The continuous operation must be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The minimum rotational speed shown under the rotor speed requirements in § 29.1509(c).
(3) The maximum allowable manifold pressure (for reciprocating engines);
(4) The maximum allowable turbine inlet or turbine outlet gas temperature (for turbine engines);
(5) The maximum allowable power or torque for each engine, considering the power input limitations of the transmission with all engines operating;
(6) The maximum allowable power or torque for each engine, considering the power input limitations of the transmission with one engine inoperative; and
(7) The maximum allowable temperatures for—
(i) The cylinder head or coolant outlet (for reciprocating engines);
(ii) The engine oil; and
(iii) The transmission oil.
(d) Fuel grade or designation. The minimum fuel grade (for reciprocating engines) or fuel designation (for turbine engines) must be established so that it is not less than that required for the operation of the engines within the limitations in paragraphs (b) and (c) of this section.
(e) Ambient temperature. Ambient temperature limitations (including limitations for winterization installations if applicable) must be established as the maximum ambient atmospheric temperature at which compliance with the cooling provisions of §§ 29.1041 through 29.1049 is shown.
(f) Two and one-half minute OEI power operation. Unless otherwise authorized, the use of 2 1/2 -minute OEI power must be limited to engine failure operation of multiengine, turbine-powered rotorcraft for not longer than 2 1/2 minutes for any period in which that power is used. The use of 2 1/2 -minute OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(g) Thirty-minute OEI power operation. Unless otherwise authorized, the use of 30-minute OEI power must be limited to multiengine, turbine-powered rotorcraft for not longer than 30 minutes after failure of an engine. The use of 30-minute OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests;
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(h) Continuous OEI power operation. Unless otherwise authorized, the use of continuous OEI power must be limited to multiengine, turbine-powered rotorcraft for continued flight after failure of an engine. The use of continuous OEI power must also be limited by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value shown during the type tests.
(2) The maximum allowable gas temperature;
(3) The maximum allowable torque; and
(4) The maximum allowable oil temperature.
(i) Rated 30-second OEI power operation. Rated 30-second OEI power is permitted only on multiengine, turbine-powered rotorcraft, also certificated for the use of rated 2-minute OEI power, and can only be used for continued operation of the remaining engine(s) after a failure or precautionary shutdown of an engine. It must be shown that following application of 30-second OEI power, any damage will be readily detectable by the applicable inspections and other related procedures furnished in accordance with Section A29.4 of appendix A of this part and Section A33.4 of appendix A of part 33. The use of 30-second OEI power must be limited to not more than 30 seconds for any period in which that power is used, and by—
(1) The maximum rotational speed which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
(j) Rated 2-minute OEI power operation. Rated 2-minute OEI power is permitted only on multiengine, turbine-powered rotorcraft, also certificated for the use of rated 30-second OEI power, and can only be used for continued operation of the remaining engine(s) after a failure or precautionary shutdown of an engine. It must be shown that following application of 2-minute OEI power, any damage will be readily detectable by the applicable inspections and other related procedures furnished in accordance with Section A29.4 of appendix a of this part and Section A33.4 of appendix A of part 33. The use of 2-minute OEI power must be limited to not more than 2 minutes for any period in which that power is used, and by—
(1) The maximum rotational speed, which may not be greater than—
(i) The maximum value determined by the rotor design; or
(ii) The maximum value demonstrated during the type tests;
(2) The maximum allowable gas temperature; and
(3) The maximum allowable torque.
§ 29.1522   Auxiliary power unit limitations.
If an auxiliary power unit that meets the requirements of TSO-C77 is installed in the rotorcraft, the limitations established for that auxiliary power unit under the TSO including the categories of operation must be specified as operating limitations for the rotorcraft.
§ 29.1523   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 kinds of operation authorized under § 29.1525.
§ 29.1525   Kinds of operations.
The kinds of operations (such as VFR, IFR, day, night, or icing) for which the rotorcraft is approved are established by demonstrated compliance with the applicable certification requirements and by the installed equipment.
§ 29.1527   Maximum operating altitude.
The maximum altitude up to which operation is allowed, as limited by flight, structural, powerplant, functional, or equipment characteristics, must be established.
§ 29.1529   Instructions for Continued Airworthiness.
The applicant must prepare Instructions for Continued Airworthiness in accordance with appendix A 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 rotorcraft or issuance of a standard certificate of airworthiness, whichever occurs later.
§ 29.1541   General.
(a) The rotorcraft must contain—
(1) The markings and placards specified in §§ 29.1545 through 29.1565; and
(2) Any additional information, instrument markings, and placards required for the safe operation of the rotorcraft if it has 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.
§ 29.1543   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 arc and line must be wide enough, and located to be clearly visible to the pilot.
§ 29.1545   Airspeed indicator.
(a) Each airspeed indicator must be marked as specified in paragraph (b) of this section, with the marks located at the corresponding indicated airspeeds.
(b) The following markings must be made:
(1) A red line:
(i) For rotorcraft other than helicopters, at V NE .
(ii) For helicopters, at V NE (power-on).
(iii) For helicopters, at V NE (power-off). If V NE (power-off) is less than V NE (power-on) and both are simultaneously displayed, the red line at V NE (power-off) must be clearly distinguishable from the red line at V NE (power-on).
(2) [Reserved]
(3) For the caution range, a yellow range.
(4) For the normal operating range, a green or unmarked range.
§ 29.1547   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.
§ 29.1549   Powerplant instruments.
For each required powerplant instrument, as appropriate to the type of instruments—
(a) Each maximum and, if applicable, minimum safe operating limit must be marked with a red line;
(b) Each normal operating range must be marked as a green or unmarked range;
(c) Each takeoff and precautionary range must be marked with a yellow range or yellow line;
(d) Each engine or rotor range that is restricted because of excessive vibration stresses must be marked with red ranges or red lines; and
(e) Each OEI limit or approved operating range must be marked to be clearly differentiated from the markings of paragraphs (a) through (d) of this section except that no marking is normally required for the 30-second OEI limit.
§ 29.1551   Oil quantity indicator.
Each oil quantity indicator must be marked with enough increments to indicate readily and accurately the quantity of oil.
§ 29.1553   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.
§ 29.1555   Control markings.
(a) Each cockpit control, other than primary flight controls or control whose function is obvious, must be plainly marked as to its function and method of operation.
(b) For powerplant fuel controls—
(1) Each fuel tank selector valve 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 any engine of a multiengine rotorcraft must be marked to indicate the position corresponding to each engine controlled.
(c) Usable fuel capacity must be marked as follows:
(1) For fuel systems having no selector controls, the usable fuel capacity of the system must be indicated at the fuel quantity indicator unless it is:
(i) Provided by another system or equipment readily accessible to the pilot; and
(ii) Contained in the limitations section of the rotorcraft flight manual.
(2) For fuel systems having selector controls, the usable fuel capacity available at each selector control position must be indicated near the selector control.
(d) For accessory, auxiliary, and emergency controls—
(1) Each essential visual position indicator, such as those showing rotor pitch or landing gear position, must be marked so that each crewmember can determine at any time the position of the unit to which it relates; and
(2) Each emergency control must be red and must be marked as to method of operation.
(e) For rotorcraft incorporating retractable landing gear, the maximum landing gear operating speed must be displayed in clear view of the pilot.
§ 29.1557   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.
(b) Seats. If the maximum allowable weight to be carried in a seat is less than 170 pounds, a placard stating the lesser weight must be permanently attached to the seat structure.
(c) Fuel and oil 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 rotorcraft, the minimum fuel grade;
(iii) For turbine-engine-powered rotorcraft, the permissible fuel designations, except that if impractical, this information may be included in the rotorcraft flight manual, and the fuel filler may be marked with an appropriate reference to the flight manual; 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”.
(d) Emergency exit placards. Each placard and operating control for each emergency exit must differ in color from the surrounding fuselage surface as prescribed in § 29.811(f)(2). A placard must be near each emergency exit control and must clearly indicate the location of that exit and its method of operation.
§ 29.1559   Limitations placard.
There must be a placard in clear view of the pilot that specifies the kinds of operations (VFR, IFR, day, night, or icing) for which the rotorcraft is approved.
§ 29.1561   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 so marked.
(c) Stowage provisions for required emergency equipment must be conspicuously marked to identify the contents and facilitate 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.
§ 29.1565   Tail rotor.
Each tail rotor must be marked so that its disc is conspicuous under normal daylight ground conditions.
§ 29.1581   General.
(a) Furnishing information. A Rotorcraft Flight Manual must be furnished with each rotorcraft, and it must contain the following:
(1) Information required by §§ 29.1583 through 29.1589.
(2) Other information that is necessary for safe operation because of design, operating, or handling characteristics.
(b) Approved information. Each part of the manual listed in §§ 29.1583 through 29.1589 that is appropriate to the rotorcraft, must be furnished, verified, and approved, and must be segregated, indentified, and clearly distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Table of contents. Each Rotorcraft Flight Manual must include a table of contents if the complexity of the manual indicates a need for it.
§ 29.1583   Operating limitations.
(a) Airspeed and rotor limitations. Information necessary for the marking of airspeed and rotor limitations on or near their respective indicators must be furnished. The significance of each limitation and of the color coding must be explained.
(b) Powerplant limitations. The following information must be furnished:
(1) Limitations required by § 29.1521.
(2) Explanation of the limitations, when appropriate.
(3) Information necessary for marking the instruments required by §§ 29.1549 through 29.1553.
(c) Weight and loading distribution. The weight and center of gravity limits required by §§ 29.25 and 29.27, respectively, must be furnished. If the variety of possible loading conditions warrants, instructions must be included to allow ready observance of the limitations.
(d) Flight crew. When a flight crew of more than one is required, the number and functions of the minimum flight crew determined under § 29.1523 must be furnished.
(e) Kinds of operation. Each kind of operation for which the rotorcraft and its equipment installations are approved must be listed.
(f) Limiting heights. Enough information must be furnished to allow compliance with § 29.1517.
(g) Maximum allowable wind. For Category A rotorcraft, the maximum allowable wind for safe operation near the ground must be furnished.
(h) Altitude. The altitude established under § 29.1527 and an explanation of the limiting factors must be furnished.
(i) Ambient temperature. Maximum and minimum ambient temperature limitations must be furnished.
§ 29.1585   Operating procedures.
(a) The parts of the manual containing operating procedures must have information concerning any normal and emergency procedures, and other information necessary for safe operation, including the applicable procedures, such as those involving minimum speeds, to be followed if an engine fails.
(b) For multiengine rotorcraft, information identifying each operating condition in which the fuel system independence prescribed in § 29.953 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.
(c) For helicopters for which a V NE (power-off) is established under § 29.1505(c), information must be furnished to explain the V NE (power-off) and the procedures for reducing airspeed to not more than the V NE (power-off) following failure of all engines.
(d) For each rotorcraft showing compliance with § 29.1353 (c)(6)(ii) or (c)(6)(iii), the operating procedures for disconnecting the battery from its charging source must be furnished.
(e) If the unusable fuel supply in any tank exceeds 5 percent of the tank capacity, or 1 gallon, whichever is greater, information must be furnished which 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.
(g) For Category B rotorcraft, the airspeeds and corresponding rotor speeds for minimum rate of descent and best glide angle as prescribed in § 29.71 must be provided.
§ 29.1587   Performance information.
Flight manual performance information which exceeds any operating limitation may be shown only to the extent necessary for presentation clarity or to determine the effects of approved optional equipment or procedures. When data beyond operating limits are shown, the limits must be clearly indicated. The following must be provided:
(a) Category A. For each category A rotorcraft, the Rotorcraft Flight Manual must contain a summary of the performance data, including data necessary for the application of any operating rule of this chapter, together with descriptions of the conditions, such as airspeeds, under which this data was determined, and must contain—
(1) The indicated airspeeds corresponding with those determined for takeoff, and the procedures to be followed if the critical engine fails during takeoff;
(2) The airspeed calibrations;
(3) The techniques, associated airspeeds, and rates of descent for autorotative landings;
(4) The rejected takeoff distance determined under § 29.62 and the takeoff distance determined under § 29.61;
(5) The landing data determined under § 29.81 and § 29.85;
(6) The steady gradient of climb for each weight, altitude, and temperature for which takeoff data are to be scheduled, along the takeoff path determined in the flight conditions required in § 29.67(a)(1) and (a)(2):
(i) In the flight conditions required in § 29.67(a)(1) between the end of the takeoff distance and the point at which the rotorcraft is 200 feet above the takeoff surface (or 200 feet above the lowest point of the takeoff profile for elevated heliports);
(ii) In the flight conditions required in § 29.67(a)(2) between the points at which the rotorcraft is 200 and 1000 feet above the takeoff surface (or 200 and 1000 feet above the lowest point of the takeoff profile for elevated heliports); and
(7) Out-of-ground effect hover performance determined under § 29.49 and the maximum weight for each altitude and temperature condition at which the rotorcraft can safely hover out-of-ground effect in winds of not less than 17 knots from all azimuths. These data must be clearly referenced to the appropriate hover charts.
(b) Category B. For each category B rotorcraft, the Rotorcraft Flight Manual must contain—
(1) The takeoff distance and the climbout speed together with the pertinent information defining the flight path with respect to autorotative landing if an engine fails, including the calculated effects of altitude and temperature;
(2) The steady rates of climb and in-ground-effect hovering ceiling, together with the corresponding airspeeds and other pertinent information, including the calculated effects of altitude and temperature;
(3) The landing distance, appropriate airspeed, and type of landing surface, together with all pertinent information that might affect this distance, including the effects of weight, altitude, and temperature;
(4) The maximum safe wind for operation near the ground;
(5) The airspeed calibrations;
(6) The height-velocity envelope except for rotorcraft incorporating this as an operating limitation;
(7) Glide distance as a function of altitude when autorotating at the speeds and conditions for minimum rate of descent and best glide angle, as determined in § 29.71;
(8) Out-of-ground effect hover performance determined under § 29.49 and the maximum safe wind demonstrated under the ambient conditions for data presented. In addition, the maximum weight for each altitude and temperature condition at which the rotorcraft can safely hover out-of-ground-effect in winds of not less than 17 knots from all azimuths. These data must be clearly referenced to the appropriate hover charts; and
(9) Any additional performance data necessary for the application of any operating rule in this chapter.
§ 29.1589   Loading information.
There must be loading instructions for each possible loading condition between the maximum and minimum weights determined under § 29.25 that can result in a center of gravity beyond any extreme prescribed in § 29.27, assuming any probable occupant weights.