gine must have an independent cooling system (including coolant tank) installed so that

(1) Each coolant tank is supported so that tank loads are distributed over a large part of the tank surface;

(2) There are pads between the tank and its supports to prevent chafing; and

(3) No air or vapor can be trapped in any part of the system, except the expansion tank, during filling or during operation.

Padding must be nonabsorbent or must be treated to prevent the absorption of flammable fluids.

(b) Coolant tank. The tank capacity must be at least one gallon, plus 10 percent of the cooling system capacity. In addition

(1) Each coolant tank must be able to withstand the vibration, inertia, and fluid loads to which it may be subjected in operation;

(2) Each coolant tank must have an expansion space of at least 10 percent of the total cooling system capacity; and

(3) It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude.

(c) Filler connection. Each coolant tank filler connection must be marked as specified in § 23.1557(c). In addition

(1) Spilled coolant must be prevented from entering the coolant tank compartment or any part of the airplane other than the tank itself; and

(2) Each recessed coolant filler connection must have a drain that discharges clear of the entire airplane.

(d) Lines and fittings. Each coolant system line and fitting must meet the requirements of § 23.993, except that the inside diameter of the engine coolant inlet and outlet lines may not be less than the diameter of the corresponding engine inlet and outlet connections.

(e) Radiators. Each coolant radiator must be able to withstand any vibration, inertia, and coolant pressure load to which it may normally be subjected. In addition

(1) Each radiator must be supported to allow expansion due to operating temperatures and prevent the transmittal of harmful vibration to the radiator; and

(2) If flammable coolant is used, the air intake duct to the coolant radiator must be located so that (in case of fire) flames from the nacelle cannot strike the radiator.

(f) Drains. There must be an acc sible drain that

(1) Drains the entire cooling syst (including the coolant tank, radiat and the engine) when the airplane is the normal ground attitude;

(2) Discharges clear of the ent airplane; and

(3) Has means to positively lock closed.

§ 23.1063 Coolant tank tests.

Each coolant tank must be test under § 23.965, except that

(a) The test required by § 23.965 (1) must be replaced with a similar t using the sum of the pressure develop during the maximum ultimate acceler tion with a full tank or a pressure of pounds per square inch, whichever greater, plus the maximum working pre sure of the system; and

(b) For a tank with a nonmetall liner the test fluid must be coola rather than fuel as specified in § 23.9 (d), and the slosh test on a specim liner must be conducted with the co ant at operating temperature. INDUCTION SYSTEM

(a) The air induction system for eac engine must supply the air required b that engine under the operating cond tions for which certification is requeste

(b) Each reciprocating engine insta lation must have at least two separa air intake sources and must meet t following:

(1) Primary air intakes may op within the cowling if that part of t cowling is isolated from the engine a cessory section by a fire-resistant di phragm or if there are means to preve the emergence of backfire flames.

(2) Each alternate air intake must located in a sheltered position and m not open within the cowling if the eme gence of backfire flames will result in hazard.

(3) The supplying of air to the engin through the alternate air intake syster may not result in a loss of excessiv power in addition to the power loss du to the rise in air temperature.

(c) For turbine engine power airplanes—

(1) There must be means to preve hazardous quantities of fuel leakage overflow from drains, vents, or oth

onents of flammable fluid systems entering the engine intake system;

The air inlet ducts must be loor protected so as to minimize the tion of foreign matter during takeinding, and taxiing.

tet No. 4080, 29 F.R. 17955, Dec. 18, 1964, ended by Amdt. No. 23-7, 34 F.R. 13095, 13, 1969]

1093 Induction system icing protection.

Each reciprocating engine air inon system must have means to preand eliminate icing. Unless this is by other means, it must be shown in air free of visible moisture at a erature of 30° F.—

Each airplane with sea level enusing conventional venturi carburehas a preheater that can provide a rise of 90° F. with the engines at 75 ent of maximum continuous power; Each airplane with altitude en$ using conventional venturi cartors has a preheater that can provide at rise of 120° F. with the engines at percent of maximum continuous r;

> Each airplane with altitude en$ using carburetors tending to preicing has a preheater that, with the nes at 60 percent of maximum conous power, can provide a heat rise

to indicate to appropriate flight crewmembers the functioning of the powerplant ice protection system.

[Amdt. No. 27-3, 34 F.R. 13095, Aug. 13, 1969] § 23.1095 Carburetor deicing fluid flow

rate.

(a) If a carburetor deicing fluid system is used, it must be able to simultaneously supply each engine with a rate of fluid flow, expressed in pounds per hour, of not less than 2.5 times the square root of the maximum continuous power of the engine.

(b) The fluid must be introduced into the air induction system

(1) Close to, and upstream of, the carburetor; and

(2) So that it is equally distributed over the entire cross section of the induction system air passages.

§ 23.1097 Carburetor deicing fluid system capacity.

(a) The capacity of each carburetor deicing fluid system

(1) May not be less than the greater of

(1) That required to provide fluid at the rate specified in § 23.1095 for a time equal to three percent of the maximum endurance of the airplane; or

(ii) 20 minutes at that flow rate; and (2) Need not exceed that required for two hours of operation.

(b) If the available preheat exceeds 50° F. but is less than 100° F., the capacity of the system may be decreased in proportion to the heat rise available in excess of 50° F.

of fuel or moisture in the normal ground and flight attitudes. No drain may discharge where it will cause a fire hazard.

(b) Each duct connected to components between which relative motion could exist must have means for flexibility.

[Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964, as amended by Amdt. No. 23-7, 34 F.R. 13095, Aug. 13, 1969]

§ 23.1105 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—

(1) The available preheat is at least 100° F.; and

For turbine engine bleed air systems, the following apply:

(a) No hazard may result if duct rupture or failure occurs anywhere between the engine port and the airplane unit served by the bleed air.

(b) The effect on airplane and engine performance of using maximum bleed air must be established.

(c) For bleed air systems used for direct cabin pressurization, no failure of the engine lubrication system may result in contamination of cabin air systems. [Amdt. No. 27–3, 34 F.R. 13095, Aug. 13, 1969] EXHAUST SYSTEM General.

§ 23.1121

(a) Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment.

(b) Unless suitable precautions are taken, no exhaust system part may be dangerously close to any system carrying flammable fluids or vapors, or under any such system that may leak.

(c) Each exhaust system component must be separated by fireproof shields from adjacent flammable parts of the airplane that are outside the engine compartment.

(d) No exhaust gases may disc dangerously near any fuel or oil s drain.

(e) No exhaust gases may be charged where they will cause a seriously affecting pilot vision at

(f) Each exhaust system comp must be ventilated to prevent poi excessively high temperature.

(g) If significant traps exists, turbine engine exhaust system must drains discharging clear of the air in any normal ground and flight att to prevent fuel accumulation afte failure of an attempted engine sta [Docket No. 4080, 29 F.R. 17955, Dec. 18 as amended by Amdt. No. 23-7, 34 F.R. Aug. 13, 1969]

§ 23.1123 Exhaust manifold.

(a) Each exhaust manifold mu fireproof and corrosion-resistant, must have means to prevent failur to expansion by operating tempera

(b) Each exhaust manifold mu supported to withstand the vibration inertia loads to which it may be subj in operation.

(c) Parts of the manifold conn to components between which rel motion could exist must have m for flexibility.

§ 23.1125 Exhaust heat exchangers.

(a) Each exhaust heat excha must be constructed and installe withstand the vibration, inertia, other loads that it may be subject in normal operation. In addition

(1) Each exchanger must be sui for continued operation at high ten atures and resistant to corrosion exhaust gases;

(2) There must be means for in tion of critical parts of each excha and

(3) Each exchanger must be v lated where it is subject to contact exhaust gases.

(b) Each heat exchanger used heating ventilating air must be structed so that exhaust gases may enter the ventilating air. POWERPLANT CONTROLS AND ACCESSO § 23.1141 Powerplant controls: gen (a) Powerplant controls must be cated and arranged under 23.777 marked under § 23.1555(a).

(b) Each flexible control must b an acceptable kind.

c) Each control must be able to intain any necessary position with

1) Constant attention by flight crew mbers; or

2) Tendency to creep due to control ds or vibration.

(d) Each control must be able to withnd operating loads without failure or cessive deflection.

(e) For turbine engine powered airnes, no single failure or malfunction, probable combination thereof, in any werplant control system may cause the lure of any powerplant function necsary for safety.

ocket No. 4080, 29 F.R. 17955, Dec. 18, 1964, amended by Amdt. No. 23-7, 34 F.R. 13095, 1g. 13, 1969]

23.1143 Engine power and thrust, and supercharger controls.

(a) There must be a separate power or rust control for each engine and a parate control for each supercharger at requires a control.

(b) Power, thrust, and supercharger ontrols must be arranged to allow(1) Separate control of each engine nd each supercharger; and

(2) Simultaneous control of all enines and all superchargers.

(c) Each power, thrust, or superharger control must give a positive and mmediate responsive means of controling its engine or supercharger.

(d) The power, thrust, or supercharger controls for each engine or supercharger must be independent of those for every other engine or upercharger.

Amdt. No. 23-7, 34 F.R. 13095, Aug. 13, 1969] | 23.1145 Ignition switches.

(a) Ignition switches must control ach ignition circuit on each engine.

(b) There must be means to quickly hut off all ignition on multiengine airlanes by the grouping of switches or by master ignition control.

(c) Each master ignition control must have means to prevent its inadvertent operation.

$23.1147 Mixture controls.

If there are mixture controls, each engine must have a separate control, and each mixture control must have guards or must be shaped or arranged to prevent confusion by feel with other controls. The controls must be grouped and arranged to allow

(a) Separate control of each engine; and

(b) Simultaneous control of all engines.

[Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964, as amended by Amdt. No. 23-7, 34 F.R. 13096, Aug. 13, 1969]

§ 23.1149 Propeller speed and pitch controls.

(a) If there are propeller speed or pitch controls, they must be grouped and arranged to allow

(1) Separate control of each propeller; and (2) Simultaneous control of all propellers.

(b) The controls must allow ready synchronization of all propellers on multiengine airplanes.

§ 23.1153 Propeller feathering controls.

If there are propeller feathering controls, each propeller must have a separate control. Each control must have means to prevent inadvertent operation. § 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime.

For turbine engine installations, each control for reverse thrust and for propeller pitch settings below the flight regime must have means to prevent its inadvertent operation. The means must have a positive lock or stop at the flight idle position and must require a separate and distinct operation by the crew to displace the control from the flight regime (forward thrust regime for turbojet powered airplanes).

[Amdt. 27-3, 34 F.R. 13096, Aug. 13, 1969] § 23.1157 Carburetor air temperature controls.

There must be a separate carburetor air temperature control for each engine. § 23.1163 Powerplant accessories.

(a) Each engine-driven accessory must

(1) Be satisfactory for mounting on the engine concerned; and

(2) Use the provisions on the engine for mounting.

(b) Electrical equipment subject to arcing or sparking must be installed to minimize the probability of contact with any flammable fluids or vapors that might be present in a free state.

§ 23.1165 Engine ignition systems.

(a) Each battery ignition system must

be supplemented by a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any battery becomes depleted.

(b) The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw 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) There must be means to warn appropriate crewmembers if malfunctioning of any part of the electrical system is causing the continuous discharge of any battery used for engine ignition. POWERPLANT FIRE PROTECTION

(a) Except as provided in paragraph (b) of this section, each line and fitting carrying flammable fluids in any area subject to engine fire conditions must meet the following requirements:

(1) The line and fitting must be at least fire resistant.

(2) Flexible hose assemblies (hose and end fittings) must be approved.

(b) Paragraph (a) of this section does not apply to—

(1) Lines and fittings already approved as part of a type certificated engine under Part 33 of this chapter; and

(2) Vent and drain lines, and their fittings, whose failure will not result in, or add to, a fire hazard.

[Docket No. 4080, 29 F.R. 17955, Dec. 18, 1964, as amended by Amdt. 23-5, 32 F.R. 6912, May 5, 1967]

(2) The closing of the fuel sh valve for any engine may not make fuel unavailable to the remaining eng that would be available to those eng with that valve open.

(3) Operation of any shutoff m may not interfere with the later e gency operation of other equipment as propeller feathering devices.

(4) Each shutoff must be outsid the engine compartment unless equal degree of safety is provided the shutoff inside the compartment.

(5) No hazardous amount of f mable fluid may drain into the en compartment after shutoff.

(6) There must be means to g against inadvertent operation of shutoff means, and to make it pos for the crew to reopen the shutoff m in flight after it has been closed.

(b) Turbine engine installations not have an engine oil system shutoff (1) The oil tank is integral with mounted on, the engine; and

(2) All oil system components ternal to the engine are fireproof. [Docket No. 4080, 29 F.R. 17955, Dec. 18, 1 as amended by Amdt. No. 23-7, 34 F.R. 13 Aug. 13, 1969]

(a) Each engine, auxiliary power u fuel burning heater, and other comb tion equipment intended for operation flight, must be isolated from the rest the airplane by firewalls, shrouds, equivalent means.

(b) Each firewall or shroud must constructed so that no hazardous qua tity of liquid, gas, or flame can pass fr the engine compartment to other pa of the airplane.

(c) Each opening in the firewall shroud must be sealed with close fitti fireproof grommets, bushings, or fi wall fittings.

(d) Fire-resistant seals may be us on single-engine airplanes and mu engine airplanes not subject to § 23.670 or (b) (1), if—

(1) Each engine has a volumetric d placement of 1,000 cubic inches or les and

(2) No opening in the firewall shroud will allow the passage of a hazar ous amount of flame without seals.

(e) Each firewall and shroud must fireproof and protected against corr sion.

(f) Compliance with the criteria f