Part V - Airworthiness Manual Chapter 525 - Transport Category Aeroplanes

Notice:

On August 31, 2013, the Canadian Aviation Regulations (CARS) will only be available on the Department of Justice website. Standards, incorporated by reference, will continue to be published on the Transport Canada website. Transport Canada will continue to make the index of the CARS and Standards available from its site for the convenience of its users.

Please be advised that there was no scheduled amendment for December 2012 (2012-2). The next amendment is planned for the end of 2013 (2013-1).

Canadian Aviation Regulations (CARs) 2012-1

Content last revised: 2010/06/01

Preamble

SUBCHAPTERS 

  • A (525.1-525.2), 
  • B (525.21-525.255), 
  • C (525.301-525.581), 
  • D (525.601-525.899), 
  • E (525.901-525.1207), 
  • F (525.1301-525.1461), 
  • G (525.1501-525.1587)
  • H (525.1701-525.1733)

APPENDICES 

A, B, C, D, E, F, G, H, I, J, L, M, N

(2001/06/01; no previous version)

SUBCHAPTER E POWERPLANT: GENERAL

525.901 Installation

(a) For the purpose of this chapter, the aeroplane powerplant installation includes each component that:

(1) Is necessary for propulsion;

(2) Affects the control of the major propulsive units; or

(3) Affects the safety of the major propulsive units between normal inspections or overhauls.

(b) For each powerplant:

(1) The installation must comply with:

(i) The installation instructions provided under 533.5 and 535.3 of this Manual; and
(amended 2010/01/29; previous version)

(ii) The applicable provisions of this subchapter;

(2) The components of the installation must be constructed, arranged, and installed so as to ensure their continued safe operation between normal inspections or overhauls;

(3) The installation must be accessible for necessary inspections and maintenance; and

(4) The major components of the installation must be electrically bonded to the other parts of the aeroplane.

(c) For each powerplant and auxiliary power unit installation, it must be established that no single failure or malfunction or probable combination of failures will jeopardise the safe operation of the aeroplane except that the failure of structural elements need not be considered if the probability of such failure is extremely remote.

(d) Each auxiliary power unit installation must meet the applicable provisions of this subchapter.

525.903 Engines

(a) Engine Type Certificate.

(1) Each engine must have a type certificate and must meet the applicable requirements of Chapter 516, Second Edition, subchapter B of this manual.

FAR:

(1) Each engine must have a type certificate and must meet the applicable requirements of Part 34 of this chapter.

(2) Each turbine engine shall comply with one of the following:
(amended 2003/12/11; previous version)

(i) Sections 533.76, 533.77 and 533.78 of chapter 533 in effect on 5 March 2001, or as subsequently amended; or
(amended 2003/12/11; previous version)

(ii) Sections 533.77 and 533.78 of chapter 533 in effect on 29 October 1998, or as subsequently amended before 5 March 2001; or
(amended 2003/12/11; previous version)

(iii) Either of:
(amended 2003/12/11; previous version)

(A) FAR 33.77 in effect on October 31, 1974, or as subsequently amended prior to 1 January 1986, unless that engine’s foreign object ingestion service history has resulted in an unsafe condition; or
(amended 2003/12/11; previous version)

(B) After 1 January 1986, section 533.77 of chapter 533, or as subsequently amended prior to 29 October 1998, unless that engine’s foreign object ingestion service history has resulted in an unsafe condition; or
(amended 2003/12/11; previous version)

FAR:

(iii) Comply with FAR 33.77 of this chapter in effect on October 31, 1974, or as subsequently amended prior to April 30, 1998, unless that engine’s foreign object ingestion service history has resulted in an unsafe condition; or
(amended 2003/12/11; previous version)

(iv) Be shown to have a foreign object ingestion service history in similar installation locations which has not resulted in any unsafe condition.
(amended 2003/12/11; previous version)

(b) Engine isolation. The powerplants must be arranged and isolated from each other to allow operation, in at least one configuration, so that the failure or malfunction of any engine, or of any system that can affect the engine, will not:

 

(1) Prevent the continued safe operation of the remaining engines; or

(2) Require immediate action by any crew member for continued safe operation.

(c) Control of engine rotation. There must be means for stopping the rotation of any engine individually in flight, except that, for turbine engine installations, the means for stopping the rotation of any engine need be provided only where continued rotation could jeopardise the safety of the aeroplane. Each component of the stopping system on the engine side of the firewall that might be exposed to fire must be at least fire-resistant. If hydraulic propeller feathering systems are used for this purpose, the feathering lines must be at least fire-resistant under the operating conditions that may be expected to exist during feathering.

(d) Turbine engine installations. For turbine engine installations:

(1) Design precautions must be taken to minimise the hazards to the aeroplane in the event of an engine rotor failure or of a fire originating within the engine which burns through the engine case.

(2) The powerplant systems associated with engine control devices, systems, and instrumentation, must be designed to give reasonable assurance that those engine operating limitations that adversely affect turbine rotor structural integrity will not be exceeded in service.

(e) Restart capability.

(1) Means to restart any engine in flight must be provided.

(2) An altitude and airspeed envelope must be established for in-flight engine restarting, and each engine must have a restart capability within that envelope.

(3) For turbine engine powered aeroplanes, if the minimum windmilling speed of the engines, following the in-flight shutdown of all engines, is insufficient to provide the necessary electrical power for engine ignition, a power source independent of the engine-driven electrical power generating system must be provided to permit in-flight engine ignition for restarting.

(f) Auxiliary Power Unit. Each auxiliary power unit must be approved or meet the requirements of the category for its intended use.

(Change 525-3 (91-11-01))
(Change 525-4 (92-08-01))
(Change 525-8)

525.904 Automatic Take-off Thrust Control System (ATTCS)

Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when any engine fails during the take-off must comply with the requirements of Appendix I of this chapter.

(Change 525-2 (89-01-01))

525.905 Propellers

(a) Each propeller must have a type certificate.

(b) Engine power and propeller shaft rotational speed may not exceed the limits for which the propeller is certificated.

(c) The propeller blade pitch control system must meet the requirements of 535.21, 535.23, 535.42 and 535.43 of this Manual.
(amended 2010/01/29; previous version)

(d) Design precautions must be taken to minimise the hazards to the aeroplane in the event a propeller blade fails or is released by a hub failure. The hazards which must be considered include damage to structure and vital systems due to impact of a failed or released blade and the unbalance created by such failure or release.

(Change 525-3 (91-11-01))

525.907  Propeller Vibration and Fatigue
(amended 2010/01/29; previous version)

This section does not apply to fixed-pitch wood propellers of conventional design.
(amended 2010/01/29; previous version)

(a) The applicant must determine the magnitude of the propeller vibration stresses or loads, including any stress peaks and resonant conditions, throughout the oeprational envelope of the aeroplane by either:
(amended 2010/01/29; previous version)

(1) Measurement of stresses or loads through direct testing or analysis based on direct testing of the propeller on the aeroplane and engine installation for which approval is sought; or
(amended 2010/01/29; previous version)

(2) Comparison of the propeller to similar propellers installed on similar aeroplane installations for which these measurements have been made.
(amended 2010/01/29; previous version)

(b) The applicant must demonstrate by tests, analysis based on tests, or previous experience on similar designs that the propeller does not experience harmful effects of flutter throughout the operational envelope of the aeroplane.
(amended 2010/01/29; previous version)

(c) The applicant must perform an evaluation of the propeller to show that failure due to fatigue will be avoided throughout the operational life of the propeller using the fatigue and structural data obtained in accordance with Airworthiness Manual Chapter 535 and the vibration data obtained from compliance with paragraph (a) of this section. For the purpose of this paragraph, the propeller includes the hub, blades, blade retention component and any other propeller component whose failure due to fatigue could be catastrophic to the aeroplane. This evaluation must include:
(amended 2010/01/29; previous version)

(1) The intended loading spectra including all reasonably foreseeable propeller vibration and cyclic load patterns, identified emergency conditions, allowable overspeeds and overtorques and the effects of temperatures and humidity expected in service.
(amended 2010/01/29; previous version)

(2) The effects of aeroplane and propeller operating and airworthiness limitations.
(amended 2010/01/29; previous version)

525.925 Propeller Clearance

Unless smaller clearances are substantiated, propeller clearances with the aeroplane at maximum weight, with the most adverse centre of gravity, and with the propeller in the most adverse pitch position, may not be less than the following:

(a) Ground clearance. There must be a clearance of at least seven inches (for each aeroplane with nose wheel landing gear) or nine inches (for each aeroplane with tail wheel landing gear) between each propeller and the ground with the landing gear statically deflected and in the level take-off, or taxiing attitude, whichever is most critical. In addition, there must be positive clearance between the propeller and the ground when in the level take-off attitude with the critical tire(s) completely deflated and the corresponding landing gear strut bottomed.

(b) Water clearance. There must be a clearance of at least 18 inches between each propeller and the water, unless compliance with 525.239(a) can be shown with a lesser clearance.

(c) Structural clearance. There must be:

(1) At least one inch radial clearance between the blade tips and the aeroplane structure, plus any additional radial clearance necessary to prevent harmful vibration;

(2) At least one-half inch longitudinal clearance between the propeller blades or cuffs and stationary parts of the aeroplane; and

(3) Positive clearance between other rotating parts of the propeller or spinner and stationary parts of the aeroplane.

(Change 525-3 (91-11-01))

525.929 Propeller De-icing

(a) For aeroplanes intended for use where icing may be expected, there must be a means to prevent or remove hazardous ice accumulation on propellers or on accessories where ice accumulation would jeopardise engine performance.

(b) If combustible fluid is used for propeller de-icing, 525.1181 through 525.1185 and 525.1189 apply.

525.933 Reversing Systems

(a) For turbojet reversing systems:

(1) Each system intended for ground operation only must be designed so that during any reversal in flight the engine will produce no more than flight idle thrust. In addition, it must be shown by analysis or test, or both, that:

(i) Each operable reverser can be restored to the forward thrust position; and

(ii) The aeroplane is capable of continued safe flight and landing under any possible position of the thrust reverser.

(2) Each system intended for in-flight use must be designed so that no unsafe condition will result during normal operation of the system, or from any failure (or reasonably likely combination of failures) of the reversing system, under any anticipated condition of operation of the aeroplane including ground operation. Failure of structural elements need not be considered if the probability of this kind of failure is extremely remote.

(3) Each system must have means to prevent the engine from producing more than idle thrust when the reversing system malfunctions, except that it may produce any greater forward thrust that is shown to allow directional control to be maintained, with aerodynamic means alone, under the most critical reversing condition expected in operation.

(b) For propeller reversing systems:

 

(1) Each system intended for ground operation only must be designed so that no single failure (or reasonably likely combination of failures) or malfunction of the system will result in unwanted reverse thrust under any expected operating condition. Failure of structural elements need not be considered if this kind of failure is extremely remote.

(2) Compliance with this section may be shown by failure analysis or testing, or both, for propeller systems that allow propeller blades to move from the flight low-pitch position to a position that is substantially less than that at the normal flight low-pitch position. The analysis may include or be supported by the analysis made to show compliance with the requirements of 535.21 of this manual for the propeller and associated installation components.

(Change 525-2 (89-01-01))
(Change 525-3 (91-11-01))

525.934 Turbojet Engine Thrust Reverser System Tests

Thrust reversers installed on turbo jet engines must meet the requirements of 533.97 of this manual.

525.937 Turbopropeller-drag Limiting Systems

Turbopropeller powered aeroplane propeller-drag limiting systems must be designed so that no single failure or malfunction of any of the systems during normal or emergency operation results in propeller drag in excess of that for which the aeroplane was designed under 525.367. Failure of structural elements of the drag limiting systems need not be considered if the probability of this kind of failure is extremely remote.

525.939 Turbine Engine Operating Characteristics

(a) Turbine engine operating characteristics must be investigated in flight to determine that no adverse characteristics (such as stall, surge, or flameout) are present, to a hazardous degree, during normal and emergency operation within the range of operating limitations of the aeroplane and of the engine.

(b) (Reserved)

(c) The turbine engine air inlet system may not, as a result of air flow distortion during normal operation, cause vibration harmful to the engine.

525.941 Inlet, Engine, and Exhaust Compatibility

For aeroplanes using variable inlet or exhaust system geometry, or both:

(a) The system comprised of the inlet, engine (including thrust augmentation systems, if incorporated), and exhaust must be shown to function properly under all operating conditions for which approval is sought, including all engine rotating speeds and power settings, and engine inlet and exhaust configurations;

(b) The dynamic effects of the operation of these (including consideration of probable malfunctions) upon the aerodynamic control of the aeroplane may not result in any condition that would require exceptional skill, alertness, or strength on the part of the pilot to avoid exceeding an operational or structural limitation of the aeroplane; and

(c) In showing compliance with paragraph (b) of this section, the pilot strength required shall not exceed the limits set forth in 525.143(d), subject to the conditions set forth in paragraphs (e) and (f) of 525.143.
(amended 2008/10/30; previous version)

525.943 Negative Acceleration

No hazardous malfunction of an engine, an auxiliary power unit approved for use in flight, or any component or system associated with the powerplant or auxiliary power unit may occur when the aeroplane is operated at the negative accelerations within the flight envelopes prescribed in 525.333. This must be shown for the greatest duration expected for the acceleration.

525.945 Thrust or Power Augmentation System

(a) General. Each fluid injection system shall provide a flow of fluid at the rate and pressure established for proper engine functioning under each intended operating condition. If the fluid can freeze, fluid freezing shall not damage the aeroplane or adversely affect aeroplane performance.
(amended 2005/06/03; previous version)

(b) Fluid tanks. Each augmentation system fluid tank shall meet the following requirements:
(amended 2005/06/03; previous version)

(1) Each tank shall be able to withstand without failure the vibration, inertia, fluid, and structural loads that it may be subjected to in operation.
(amended 2005/06/03; previous version)

(2) The tanks as mounted in the aeroplane shall be able to withstand without failure or leakage an internal pressure l.5 times the maximum operating pressure.
(amended 2005/06/03; previous version)

(3) If a vent is provided, the venting shall be effective under all normal flight conditions.
(amended 2005/06/03; previous version)

(4) (Reserved)

(5) Each tank shall have an expansion space of not less than 2 percent of the tank capacity. It shall be impossible to fill the expansion space inadvertently with the aeroplane in the normal ground attitude.
(amended 2005/06/03; previous version)

(c) Augmentation system drains shall be designed and located in accordance with 525.1455 if:
(amended 2005/06/03; previous version)

(1) The augmentation system fluid is subject to freezing; and

(2) The fluid may be drained in flight or during ground operation.

(d) The augmentation liquid tank capacity available for the use of each engine shall be large enough to allow operation of the aeroplane under the approved procedures for the use of liquid-augmented power. The computation of liquid consumption shall be based on the maximum approved rate appropriate for the desired engine output and shall include the effect of temperature on engine performance as well as any other factors that might vary the amount of liquid required.
(amended 2005/06/03; previous version)

(e) This section does not apply to fuel injection systems.

(Change 525-3 (91-11-01))

Fuel System

525.951 General

(a) Each fuel system must be constructed and arranged to ensure a flow of fuel at a rate and pressure established for proper engine and auxiliary power unit functioning under each likely operating condition, including any manoeuvre for which certification is requested and during which the engine or auxiliary power unit is permitted to be in operation.

(b) Each fuel system must be arranged so that any air which is introduced into the system will not result in:

(1) Power interruption for more than 20 seconds for reciprocating engines; or

(2) Flame-out for turbine engines.

(c) Each fuel system for a turbine engine must be capable of sustained operation throughout its flow and pressure range with fuel initially saturated with water at 80°F and having 0.75cc of free water per gallon added and cooled to the most critical condition for icing likely to be encountered in operation.

(d) Each fuel system for a turbine engine powered aeroplane must meet the applicable fuel venting requirements of Chapter 516, Second Edition, subchapter B of this manual.

(Change 525-3 (91-11-01))

525.952 Fuel System Analysis and Test

(a) Proper fuel system functioning under all probable operating conditions must be shown by analysis and those tests found necessary by the Minister. Tests, if required, must be made using the aeroplane fuel system or a test article that reproduces the operating characteristics of the portion of the fuel system to be tested.

(b) The likely failure of any heat exchanger using fuel as one of its fluids may not result in a hazardous condition.

525.953 Fuel System Independence

Each fuel system must meet the requirements of 525.903(b) by:

(a) Allowing the supply of fuel to each engine through a system independent of each part of the system supplying fuel to any other engine; or

(b) Any other acceptable method.

525.954 Fuel System Lightning Protection

The fuel system must be designed and arranged to prevent the ignition of fuel vapour 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.

525.955 Fuel Flow

(a) Each fuel system must provide at least 100 percent of the fuel flow required under each intended operating condition and manoeuvre. Compliance must be shown as follows:

(1) Fuel must be delivered to each engine at a pressure within the limits specified in the engine type certificate.

(2) The quantity of fuel in the tank may not exceed the amount established as the unusable fuel supply for that tank under the requirements of 525.959 plus that necessary to show compliance with this section.

(3) Each main pump must be used that is necessary for each operating condition and attitude for which compliance with this section is shown, and the appropriate emergency pump must be substituted for each main pump so used.

(4) If there is a fuel flowmeter, it must be blocked and the fuel must flow through the meter or its bypass.

(b) If an engine can be supplied with fuel from more than one tank, the fuel system must:

(1) For each reciprocating engine, supply the full fuel pressure to that engine in not more than 20 seconds after switching to any other fuel tank containing usable fuel when engine malfunctioning becomes apparent due to the depletion of the fuel supply in any tank from which the engine can be fed; and

(2) For each turbine engine, in addition to having appropriate manual switching capability, be designed to prevent interruption of fuel flow to that engine, without attention by the flight crew, 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.

525.957 Flow Between Interconnected Tanks

If fuel can be pumped from one tank to another in flight, the fuel tank vents and the fuel transfer system must be designed so that no structural damage to the tanks can occur because of overfilling.

525.959 Unusable Fuel Supply

The unusable fuel quantity for each fuel tank and its fuel system components must be established at not less than the quantity at which the first evidence of engine malfunction occurs under the most adverse fuel feed condition for all intended operations and flight manoeuvres involving fuel feeding from that tank. Fuel system component failures need not be considered.

525.961 Fuel System Hot Weather Operation

(a) The fuel system must perform satisfactorily in hot weather operation. This must be shown by showing that the fuel system from the tank outlets to each engine is pressurised, under all intended operations, so as to prevent vapour formation, or must be shown by climbing from the altitude of the airport elected by the applicant to the maximum altitude established as an operating limitation under 525.1527. If a climb test is elected, there may be no evidence of vapour lock or other malfunctioning during the climb test conducted under the following conditions:

(1) For reciprocating engine powered aeroplanes, the engines must operate at maximum continuous power, except that take-off power must be used for the altitudes from 1,000 feet below the critical altitude through the critical altitude. The time interval during which take-off power is used may not be less than the take-off time limitation.

(2) For turbine engine powered aeroplanes, the engines must operate at take-off power for the time interval selected for showing the take-off flight path, and at maximum continuous power for the rest of the climb.

(3) The weight of the aeroplane must be the weight with full fuel tanks, minimum crew, and the ballast necessary to maintain the centre of gravity within allowable limits.

(4) The climb airspeed may not exceed:

(i) For reciprocating engine powered aeroplanes, the maximum airspeed established for climbing from take-off to the maximum operating altitude with the aeroplane in the following configuration:

(A) Landing gear retracted.

(B) Wing flaps in the most favourable position.

(C) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition.

(D) Engine operating within the maximum continuous power limitations.

(E) Maximum take-off weight; and

(ii) For turbine engine powered aeroplanes, the maximum airspeed established for climbing from take-off to the maximum operating altitude.

(5) The fuel temperature must be at least 110°F.

(b) The test prescribed in paragraph (a) of this section may be performed in flight or on the ground under closely simulated flight conditions. If a flight test is performed in weather cold enough to interfere with the proper conduct of the test, the fuel tank surfaces, fuel lines, and other fuel system parts subject to cold air must be insulated to simulate, insofar as practicable, flight in hot weather.

525.963 Fuel Tanks: General

(a) Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads that it may be subjected to in operation.

(b) Flexible fuel tank liners must be approved or must be shown to be suitable for the particular application.

(c) Integral fuel tanks must have facilities for interior inspection and repair.

(d) Fuel tanks within the fuselage contour must be able to resist rupture, and to retain fuel, under the inertia forces prescribed for the emergency landing conditions in 525.561. In addition, these tanks must be in a protected position so that exposure of the tanks to scraping action with the ground is unlikely.

(e) Fuel tank access covers must comply with the following criteria in order to avoid loss of hazardous quantities of fuel:

(1) All covers located in an area where experience or analysis indicates a strike is likely must be shown by analysis or test to minimise penetration and deformation by tire fragments, low energy engine debris, or other likely debris.

(2) All covers must be fire resistant as defined in Chapter 500 of this manual.

(f) For pressurised fuel tanks, a means with fail-safe features must be provided to prevent the build-up of an excessive pressure difference between the inside and the outside of the tank.

(Change 525-3 (91-11-01))

525.965 Fuel Tank Tests

(a) It must be shown by tests that the fuel tanks, as mounted in the aeroplane, can withstand, without failure or leakage, the more critical of the pressures resulting from the conditions specified in subparagraphs (1) and (2) of this paragraph. In addition, it must be shown by either analysis or tests, that tank surfaces subjected to more critical pressures resulting from the condition of subparagraphs (3) and (4) of this paragraph, are able to withstand the following pressures:

(1) An internal pressure of 3.5 p.s.i.

(2) 125 percent of the maximum air pressure developed in the tank from ram effect.

(3) Fluid pressures developed during maximum limit accelerations, and deflections, of the aeroplane with a full tank.

(4) Fluid pressures developed during the most adverse combination of aeroplane roll and fuel load.

(b) Each metallic tank with large unsupported or unstiffened flat surfaces, whose failure or deformation could cause fuel leakage, must be able to withstand the following test, or its equivalent, without leakage or excessive deformation of the tank walls:

(1) Each complete tank assembly and its supports must be vibration tested while mounted to simulate the actual installation.

(2) Except as specified in subparagraph (4) of this paragraph, the tank assembly must be vibrated for 25 hours at an amplitude of not less than 1/32 of an inch (unless another amplitude is substantiated) while 2/3 filled with water or other suitable test fluid.

(3) The test frequency of vibration must be as follows:

(i) If no frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the test frequency of vibration must be 2,000 cycles per minute.

(ii) If only one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, that frequency of vibration must be the test frequency.

(iii) If more than one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the most critical of these frequencies must be the test frequency.

(4) Under subparagraphs (3)(ii) and (iii) of this paragraph, the time of test must be adjusted to accomplish the same number of vibration cycles that would be accomplished in 25 hours at the frequency specified in subparagraph (3)(i) of this paragraph.

(5) During the test, the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute, through an angle of 15° on both sides of the horizontal (30° total), about the most critical axis, for 25 hours. If motion about more than one axis is likely to be critical, the tank must be rocked about each critical axis for 12½ hours.

(c) Except where satisfactory operating experience with a similar tank in a similar installation is shown, non-metallic tanks must withstand the test specified in paragraph (b)(5) of this section, with fuel at a temperature of 110° F. During this test, a representative specimen of the tank must be installed in a supporting structure simulating the installation in the aeroplane.

(d) For pressurised fuel tanks, it must be shown by analysis or tests that the fuel tanks can withstand the maximum pressure likely to occur on the ground or in flight.

525.967 Fuel Tank Installations

(a) Each fuel tank must be supported so that tank loads (resulting from the weight of the fuel in the tanks) are not concentrated on unsupported tank surfaces. In addition:

(1) There must be pads, if necessary, to prevent chafing between the tank and its supports;

(2) Padding must be non-absorbent or treated to prevent the absorption of fluids;

(3) If a flexible tank liner is used, it must be supported so that it is not required to withstand fluid loads; and

(4) Each interior surface of the tank compartment must be smooth and free of projections that could cause wear of the liner unless:

(i) Provisions are made for protection of the liner at these points; or

(ii) The construction of the liner itself provides that protection.

(b) Spaces adjacent to tank surfaces must be ventilated to avoid fume accumulation due to minor leakage. If the tank is in a sealed compartment, ventilation may be limited to drain holes large enough to prevent excessive pressure resulting from altitude changes.

(c) The location of each tank must meet the requirements of 525.1185(a).

(d) No engine nacelle skin immediately behind a major air outlet from the engine compartment may act as the wall of an integral tank.

(e) Each fuel tank must be isolated from personnel compartments by a fumeproof and fuelproof enclosure.

525.969 Fuel Tank Expansion Space

Each fuel tank must have an expansion space of not less than 2 percent of the tank capacity. It must be impossible to fill the expansion space inadvertently with the aeroplane in the normal ground attitude. For pressure fuelling systems, compliance with this section may be shown with the means provided to comply with 525.979(b).

525.971 Fuel Tank Sump

(a) Each fuel tank must have a sump with an effective capacity, in the normal ground attitude, of not less than the greater of 0.10 percent of the tank capacity or one-sixteenth of a gallon unless operating limitations are established to ensure that the accumulation of water in service will not exceed the sump capacity.

(b) Each fuel tank must allow drainage of any hazardous quantity of water from any part of the tank to its sump with the aeroplane in the ground attitude.

(c) Each fuel tank sump must have an accessible drain that:

(1) Allows complete drainage of the sump on the ground;

(2) Discharges clear of each part of the aeroplane; and

(3) Has manual or automatic means for positive locking in the closed position.

525.973 Fuel Tank Filler Connection

Each fuel tank filler connection shall prevent the entrance of fuel into any part of the aeroplane other than the tank itself. In addition:
(amended 2005/06/03; previous version)

(a) (Reserved)

(b) Each recessed filler connection that can retain any appreciable quantity of fuel shall have a drain that discharges clear of each part of the aeroplane;
(amended 2005/06/03; previous version)

(c) Each filler cap shall provide a fuel-tight seal; and
(amended 2005/06/03; previous version)

(d) Each fuel filling point shall have a provision for electrically bonding the aeroplane to ground fuelling equipment.
(amended 2005/06/03; previous version)

(Change 525-3 (91-11-01))

525.975 Fuel Tank Vents and Carburettor Vapour Vents

(a) Fuel tank vents. Each fuel tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition. In addition:

(1) Each vent must be arranged to avoid stoppage by dirt or ice formation;

(2) The vent arrangement must prevent siphoning of fuel during normal operation;

(3) The venting capacity and vent pressure levels must maintain acceptable differences of pressure between the interior and exterior of the tank, during:

(i) Normal flight operation;

(ii) Maximum rate of ascent and descent; and

(iii) Refuelling and defuelling (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 aeroplane in the ground attitude or the level flight attitude, unless drainage is provided; and

(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.

(b) Carburettor vapour vents. Each carburettor with vapour elimination connections must have a vent line to lead vapours 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 vapour vent return line must lead back to the fuel tank used for take-off and landing.

525.977 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 aeroplanes, have 8 to 16 meshes per inch; and

(2) For turbine engine powered aeroplanes, prevent the passage of any object that could restrict fuel flow or damage any fuel system component.

(b) (Reserved)

(c) The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line.

(d) The diameter of each strainer must be at least that of the fuel tank outlet.

(e) Each finger strainer must be accessible for inspection and cleaning.

525.979 Pressure Fuelling System

For pressure fuelling systems, the following apply:

(a) Each pressure fuelling system fuel manifold connection must have means to prevent the escape of hazardous quantities of fuel from the system if the fuel entry valve fails.

(b) An automatic shut-off means must be provided to prevent the quantity of fuel in each tank from exceeding the maximum quantity approved for that tank. This means must:

(1) Allow checking for proper shut-off operation before each fuelling of the tank; and

(2) Provide indication, at each fuelling station, of failure of the shut-off means to stop fuel flow at the maximum quantity approved for that tank.

(c) A means must be provided to prevent damage to the fuel system in the event of failure of the automatic shut-off means prescribed in paragraph (b) of this section.

(d) The aeroplane pressure fuelling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum pressures, including surge, that is likely to occur during fuelling. The maximum surge pressure must be established with any combination of tank valves being either intentionally or inadvertently closed.

(e) The aeroplane defuelling 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 defuelling pressure (positive or negative) at the aeroplane fuelling connection.

(Change 525-3 (91-11-01))

525.981 Fuel Tank Explosion Prevention
(amended 2009/05/11; previous version)

(a) No ignition source may be present at each point in the fuel tank or fuel tank system where catastrophic failure could occur due to ignition of fuel or vapours. This shall be demonstrated by:

(1) Determining the highest temperature allowing a safe margin below the lowest expected auto-ignition temperature of the fuel in the fuel tanks.

(2) Demonstrating that no temperature at each place inside each fuel tank where fuel ignition is possible will exceed the temperature determined under paragraph (a)(1) of this section. This shall be verified under all probable operating, failure, and malfunction conditions of each component whose operation, failure, or malfunction could increase the temperature inside the tank.

(3) Demonstrating that an ignition source could not result from each single failure, from each single failure in combination with each latent failure condition not shown to be extremely remote, and from all combinations of failures not shown to be extremely improbable. The effects of manufacturing variability, aging, wear, corrosion, and likely damage shall be considered.

(b) Except as provided in paragraphs (b)(2) and (c) of this section, no fuel tank Fleet Average Flammability Exposure on an aeroplane may exceed three percent of the Flammability Exposure Evaluation Time (FEET) as defined in Appendix N of this part, or that of a fuel tank within the wing of the aeroplane model being evaluated, whichever is greater. If the wing is not a conventional unheated aluminum wing, the analysis must be based on an assumed Equivalent Conventional Unheated Aluminum Wing Tank.
(amended 2009/05/11; previous version)

(1) Fleet Average Flammability Exposure is determined in accordance with Appendix N of this part. The assessment must be done in accordance with the methods and procedures set forth in the Fuel Tank Flammability Assessment Method User’s Manual, dated May 2008, document number DOT/FAA/AR–05/8.
(amended 2009/05/11; previous version)

(2) Any fuel tank other than a main fuel tank on an aeroplane must meet the flammability exposure criteria of Appendix M to this part if any portion of the tank is located within the fuselage contour.
(amended 2009/05/11; previous version)

(3) As used in this paragraph,
(amended 2009/05/11; previous version)

(i) Equivalent Conventional Unheated Aluminum Wing Tank is an integral tank in an unheated semi-monocoque aluminum wing of a subsonic aeroplane that is equivalent in aerodynamic performance, structural capability, fuel tank capacity and tank configuration to the designed wing.
(amended 2009/05/11; previous version)

(ii) Fleet Average Flammability Exposure is defined in Appendix N to this part and means the percentage of time each fuel tank ullage is flammable for a fleet of an aeroplane type operating over the range of flight lengths.
(amended 2009/05/11; previous version)

(iii) Main Fuel Tank means a fuel tank that feeds fuel directly into one or more engines and holds required fuel reserves continually throughout each flight.
(amended 2009/05/11; previous version)

(c) Paragraph (b) of this section does not apply to a fuel tank if means are provided to mitigate the effects of an ignition of fuel vapours within fuel tank such that no damage caused by an ignition will prevent continued safe flight and landing.
(amended 2009/05/11; previous version)

(d) Critical design configuration control limitations (CDCCL), inspections, or other procedures must be established, as necessary, to prevent development of ignition sources within the fuel tank system pursuant to (a) of this section, to prevent increasing the flammability exposure of the tanks above that permitted under (b) of this section, and to prevent degradation of the performance and reliability of any means provided under paragraphs (a) or (c) of this section. These CDCCL, inspections and procedures must be included in the Airworthiness Limitations section of the Instructions for Continued Airworthiness required by 525.1529. Visible means to identify critical features of the design must be placed in areas of the aeroplane where maintenance actions, repairs, or alterations may compromise the critical design configuration limitations (e.g., colour-coding of wire to identify separation limitation). These visible means must also be identified as CDCCL.
(amended 2009/05/11; previous version)

Fuel System Components

525.991 Fuel Pumps

(a) Main pumps. Each fuel pump required for proper engine operation, or required to meet the fuel system requirements of this subchapter (other than those in paragraph (b) of this section), is a main pump. For each main pump, provision must be made to allow the bypass of each positive displacement fuel pump other than a fuel injection pump (a pump that supplies the proper flow and pressure for fuel injection when the injection is not accomplished in a carburettor) approved as part of the engine.

(b) Emergency pumps. There must be emergency pumps or another main pump to feed each engine immediately after failure of any main pump (other than a fuel injection pump approved as part of the engine).

525.993 Fuel System Lines and Fittings

(a) Each fuel line must be installed and supported to prevent excessive vibration and to withstand loads due to fuel pressure and accelerated flight conditions.

(b) Each fuel line connected to components of the aeroplane between which relative motion could exist must have provisions for flexibility.

(c) Each flexible connection in fuel lines that may be under pressure and subjected to axial loading must use flexible hose assemblies.

(d) Flexible hose must be approved or must be shown to be suitable for the particular application.

(e) No flexible hose that might be adversely affected by exposure to high temperatures may be used where excessive temperatures will exist during operation or after engine shut-down.

(f) Each fuel line within the fuselage must be designed and installed to allow a reasonable degree of deformation and stretching without leakage.

525.994 Fuel System Components

Fuel system components in an engine nacelle or in the fuselage must be protected from damage which could result in spillage of enough fuel to constitute a fire hazard as a result of a wheels-up landing on a paved runway.

525.995 Fuel Valves

In addition to the requirements of 525.1189 for shut-off means, each fuel valve must:

(a) (Reserved)

(b) Be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve.

525.997 Fuel Strainer or Filter

There must be a fuel strainer or filter between the fuel tank outlet and the inlet of either the fuel metering device or an engine driven positive displacement pump, whichever is nearer the fuel tank outlet. This fuel strainer or filter must:

(a) Be accessible for draining and cleaning and must incorporate a screen or element which is easily removable;

(b) Have a sediment trap and drain except that it need not have a drain if the strainer or filter is easily removable for drain purposes;

(c) Be mounted so that its weight is not supported by the connecting lines or by the inlet or outlet connections of the strainer or filter itself, unless adequate strength margins under all loading conditions are provided in the lines and connections; and

(d) Have the capacity (with respect to operating limitations established for the engine) to ensure that engine fuel system functioning is not impaired, with the fuel contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine in Chapter 533 of this manual.

525.999 Fuel System Drains

(a) Drainage of the fuel system must be accomplished by the use of fuel strainer and fuel tank sump drains.

(b) Each drain required by paragraph (a) of this section must:

(1) Discharge clear of all parts of the aeroplane:

(2) Have manual or automatic means for positive locking in the closed position; and

(3) Have a drain valve:

(i) That is readily accessible and which can be easily opened and closed; and

(ii) That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted.

525.1001 Fuel Jettisoning System

(a) A fuel jettisoning system must be installed on each aeroplane unless it is shown that the aeroplane meets the climb requirements of 525.119 and 525.121(d) at maximum take-off weight, less the actual or computed weight of fuel necessary for a 15 minute flight comprised of a take-off, go-around, and landing at the airport of departure with the aeroplane configuration, speed, power, and thrust the same as that used in meeting the applicable take-off, approach, and landing climb performance requirements of this chapter.

(b) If a fuel jettisoning system is required it must be capable of jettisoning enough fuel within 15 minutes, starting with the weight given in paragraph (a) of this section, to enable the aeroplane to meet the climb requirements of 525.119 and 525.121(d), assuming that the fuel is jettisoned under the conditions, except weight, found least favourable during the flight tests prescribed in paragraph (c) of this section.

(c) Fuel jettisoning shall be demonstrated beginning at maximum take-off weight with flaps and landing gear up and in:
(amended 2003/11/10; previous version)

(1) A power-off glide at 1.3 VSR1;
(amended 2003/11/10; previous version)

(2) A climb at the one-engine inoperative best rate-of-climb speed, with the critical engine inoperative and the remaining engines at maximum continuous power; and

(3) Level flight at 1.3 VSR1; if the results of the tests in the conditions specified in paragraphs (c)(1) and (2) of this section demonstrate that this condition could be critical.
(amended 2003/11/10; previous version)

(d) During the flight tests prescribed in paragraph (c) of this section, it must be shown that:

(1) The fuel jettisoning system and its operation are free from fire hazard;

(2) The fuel discharges clear of any part of the aeroplane;

(3) Fuel or fumes do not enter any parts of the aeroplane; and

(4) The jettisoning operation does not adversely affect the controllability of the aeroplane.

(e) For reciprocating engine powered aeroplanes, means must be provided to prevent jettisoning the fuel in the tanks used for take-off and landing below the level allowing 45 minutes flight at 75 percent maximum continuous power. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control.

(f) For turbine engine powered aeroplanes, means must be provided to prevent jettisoning the fuel in the tanks used for take-off and landing below the level allowing climb from sea level to 10,000 feet and thereafter allowing 45 minutes cruise at a speed for maximum range. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control.

(g) The fuel jettisoning valve must be designed to allow flight personnel to close the valve during any part of the jettisoning operation.

(h) Unless it is shown that using any means (including flaps, slots, and slats) for changing the airflow across or around the wings does not adversely affect fuel jettisoning, there must be a placard, adjacent to the jettisoning control, to warn flight crew members against jettisoning fuel while the means that change the airflow are being used.

(i) The fuel jettisoning system must be designed so that any reasonably probable single malfunction in the system will not result in a hazardous condition due to unsymmetrical jettisoning of, or inability to jettison, fuel.

Oil System

525.1011 General

(a) Each engine must have an independent oil system that can supply it with an appropriate quantity of oil at a temperature not above that safe for continuous operation.

(b) The usable oil capacity may not be less than the product of the endurance of the aeroplane under critical operating conditions and the approved maximum allowable oil consumption of the engine under the same conditions, plus a suitable margin to ensure system circulation. Instead of a rational analysis of aeroplane range for the purpose of computing oil requirements for reciprocating engine powered aeroplanes, the following fuel/oil ratios may be used:

(1) For aeroplanes without a reserve oil or oil transfer system, a fuel/oil ratio of 30:1 by volume.

(2) For aeroplanes with either a reserve oil or oil transfer system, a fuel/oil ratio of 40:1 by volume.

(c) Fuel/oil ratios higher than those prescribed in paragraphs (b)(1) and (2) of this section may be used if substantiated by data on actual engine oil consumption.

525.1013 Oil Tanks

(a) Installation. Each oil tank installation must meet the requirements of 525.967.

 

(b) Expansion space. Oil tank expansion space must be provided as follows:

(1) Each oil tank used with a reciprocating engine must have an expansion space of not less than the greater of 10 percent of the tank capacity or 0.5 gallon, and each oil tank used with a turbine engine must have an expansion space of not less than 10 percent of the tank capacity.

(2) Each reserve oil tank not directly connected to any engine may have an expansion space of not less than two percent of the tank capacity.

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

(c) Filler connection. Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of each part of the aeroplane. In addition each oil tank filler cap must provide an oil-tight seal.

(d) Vent. Oil tanks must be vented as follows:

(1) Each oil tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition.

(2) Oil tank vents must be arranged so that condensed water vapour that might freeze and obstruct the line cannot accumulate at any point.

(e) Outlet. There must be means to prevent entrance into the tank itself, or into the tank outlet, of any object that might obstruct the flow of oil through the system. No oil tank outlet may be enclosed by any screen or guard that would reduce the flow of oil below a safe value at any operating temperature. There must be a shut-off valve at the outlet of each oil tank used with a turbine engine, unless the external portion of the oil system (including the oil tank supports) is fireproof.

(f) Flexible oil tank liners. Each flexible oil tank liner must be approved or must be shown to be suitable for the particular application.

(Change 525-3 (91-11-01))

525.1015 Oil Tank Tests

Each oil tank must be designed and installed so that:

(a) It can withstand, without failure, each vibration, inertia, and fluid load that it may be subjected to in operation; and

(b) It meets the provisions of 525.965, except:

(1) The test pressure:

(i) For pressurised tanks used with a turbine engine, may not be less than 5 p.s.i. plus the maximum operating pressure of the tank instead of the pressure specified in 525.965(a); and

(ii) For all other tanks may not be less than 5 p.s.i. instead of the pressure specified in 525.965(a); and

(2) The test fluid must be oil at 250°F instead of the fluid specified in 525.965(c).

525.1017 Oil Lines and Fittings

(a) Each oil line must meet the requirements of 525.993 and each oil line and fitting in any designated fire zone must meet the requirements of 525.1183.

(b) Breather lines must be arranged so that:

(1) Condensed water vapour that might freeze and obstruct the line cannot accumulate at any point;

(2) The breather discharge does not constitute a fire hazard if foaming occurs or causes emitted oil to strike the pilot’s windshield; and

(3) The breather does not discharge into the engine air induction system.

525.1019 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 Chapter 533 of this manual.

(3) The oil strainer or filter, unless it is installed at an oil tank outlet, must incorporate an indicator that will indicate contamination before it reaches the capacity established in accordance with subparagraph (2) of this paragraph.

(4) The bypass of a strainer or filter must be constructed and installed so that the release of collected contaminants is minimised 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 525.1305(c)(7).

(b) Each oil strainer or filter in a powerplant installation using reciprocating engines must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter element completely blocked.

525.1021 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.

525.1023 Oil Radiators

(a) Each oil radiator must be able to withstand, without failure, any vibration, inertia, and oil pressure load to which it would be subjected in operation.

(b) Each oil radiator air duct must be located so that, in case of fire, flames coming from normal openings of the engine nacelle cannot impinge directly upon the radiator.

525.1025 Oil Valves

(a) Each oil shut-off must meet the requirements of 525.1189.

(b) The closing of oil shut-off means may not prevent propeller feathering.

(c) Each oil valve must have positive stops or suitable index provisions in the "on" and "off" positions and must be supported so that no loads resulting from its operation or from accelerated flight conditions are transmitted to the lines attached to the valve.

525.1027 Propeller Feathering System

(a) If the propeller feathering system depends on engine oil, there must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricating system other than the tank itself.

 

(b) The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump.

(c) The ability of the system to accomplish feathering with the trapped oil must be shown. This may be done on the ground using an auxiliary source of oil for lubricating the engine during operation.

(d) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system.

Cooling

525.1041 General

The powerplant and auxiliary power unit cooling provisions must be able to maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the temperature limits established for these components and fluids, under ground, water, and flight operating conditions, and after normal engine or auxiliary power unit shutdown, or both.

525.1043 Cooling Tests

(a) General. Compliance with 525.1041 must be shown by tests, under critical ground, water, and flight operating conditions. For these tests, the following apply:

(1) If the tests are conducted under conditions deviating from the maximum ambient atmospheric temperature, the recorded powerplant temperatures must be corrected under paragraphs (c) and (d) of this section.

(2) No corrected temperatures determined under subparagraph (1) of this paragraph may exceed established limits.

(3) For reciprocating engines, the fuel used during the cooling tests must be the minimum grade approved for the engines, and the mixture settings must be those normally used in the flight stages for which the cooling tests are conducted. The test procedures must be as prescribed in 525.1045.

(b) Maximum ambient atmospheric temperature. A maximum ambient atmospheric temperature corresponding to sea level conditions of at least 100 degrees F must be established. The assumed temperature lapse rate is 3.6 degrees F per thousand feet of altitude above sea level until a temperature of -69.7 degrees F is reached, above which altitude the temperature is considered constant at -69.7 degrees F. However, for winterisation installations, the applicant may select a maximum ambient atmospheric temperature corresponding to sea level conditions of less than 100 degrees F.

(c) Correction factor (except cylinder barrels). Unless a more rational correction applies, temperatures of engine fluids and powerplant components (except cylinder barrels) for which temperature limits are established, must be corrected by adding to them the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test.

(d) Correction factor for cylinder barrel temperatures. Unless a more rational correction applies, cylinder barrel temperatures must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum cylinder barrel temperature recorded during the cooling test.

525.1045 Cooling Test Procedures

(a) Compliance with 525.1041 must be shown for the take-off, climb, en route, and landing stages of flight that correspond to the applicable performance requirements. The cooling tests must be conducted with the aeroplane in the configuration, and operating under the conditions, that are critical relative to cooling during each stage of flight. For the cooling tests, a temperature is "stabilised" when its rate of change is less than two degrees F per minute.

(b) Temperatures must be stabilised under the conditions from which entry is made into each stage of flight being investigated, unless the entry condition normally is not one during which component and engine fluid temperatures would stabilise (in which case, operation through the full entry condition must be conducted before entry into the stage of flight being investigated in order to allow temperatures to reach their natural levels at the time of entry). The take-off cooling test must be preceded by a period during which the powerplant component and engine fluid temperatures are stabilised with the engines at ground idle.

 

(c) Cooling tests for each stage of flight must be continued until:

(1) The component and engine fluid temperatures stabilise;

(2) The stage of flight is completed; or

(3) An operating limitation is reached.

(d) For reciprocating engine powered aeroplanes, it may be assumed, for cooling test purposes, that the take-off stage of flight is complete when the aeroplane reaches an altitude of 1,500 feet above the take-off surface or reaches a point in the take-off where the transition from the take-off to the en route configuration is completed and a speed is reached at which compliance with 525.121(c) is shown, whichever point is at a higher altitude. The aeroplane must be in the following configuration:

(1) Landing gear retracted.

(2) Wing flaps in the most favourable position.

(3) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition.

(4) Critical engine inoperative and its propeller stopped.

(5) Remaining engines at the maximum continuous power available for the altitude.

(e) For hull seaplanes and amphibians, cooling must be shown during taxiing downwind for 10 minutes, at 5 knots above step speed.

Induction System

525.1091 Air Induction

(a) The air induction system for each engine and auxiliary power unit must supply:

(1) The air required by the engine and auxiliary power unit under each operating condition for which certification is requested; and

(2) The air for proper fuel metering and mixture distribution with the induction system valves in any position.

(b) Each reciprocating engine must have an alternate air source that prevents the entry of rain, ice, or any other foreign matter.

(c) Air intakes may not open within the cowling, unless:

(1) That part of the cowling is isolated from the engine accessory section by means of a fireproof diaphragm; or

(2) For reciprocating engines, there are means to prevent the emergence of backfire flames.

(d) For turbine engine powered aeroplanes and aeroplanes 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 aeroplane must be designed to prevent water or slush on the runway, taxiway, or other airport operating surfaces from being directed into the engine or auxiliary power unit air inlet ducts in hazardous quantities, and the air inlet ducts must be located or protected so as to minimise the ingestion of foreign matter during take-off, landing, and taxiing.

(e) If the engine induction system contains parts or components that could be damaged by foreign objects entering the air inlet, it must be shown by tests or, if appropriate, by analysis that the induction system design can withstand the foreign object ingestion test conditions of sections 533.76, 533.77 and paragraph 533.78(a)(1) of this manual without failure of parts or components that could create a hazard.
(amended 2001/03/05; previous version)

525.1093 Induction System De-icing and Anti-icing Provisions

(a) Reciprocating engines. Each reciprocating engine air induction system must have means to prevent and eliminate icing. Unless this is done by other means, it must be shown that, in air free of visible moisture at a temperature of 30°F, each aeroplane with altitude engines using:

(1) Conventional venturi carburettors have a preheater that can provide a heat rise of 120°F with the engine at 60 percent of maximum continuous power; or

(2) Carburettors tending to reduce the probability of ice formation has a preheater that can provide a heat rise of 100°F with the engine at 60 percent of maximum continuous power.

(b) Turbine engines.

(1) Each turbine engine and its air inlet system must operate throughout the flight power range of the engine (including idling), without the accumulation of ice on the engine, inlet system components, or airframe components that would adversely affect engine operation or cause a serious loss of power or thrust:

(i) Under the icing conditions specified in Appendix C; and

(ii) In falling and blowing snow within the limitations established for the aeroplane for such operation.

(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 metre in the form of drops having a mean effective diameter not less than 20 microns, followed by momentary operation at take-off 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 Minister.

(c) Supercharged reciprocating engines. For each engine having a supercharger to pressurise the air before it enters the carburettor, the heat rise in the air caused by that supercharging at any altitude may be utilised in determining compliance with paragraph (a) of this section if the heat rise utilised is that which will be available, automatically, for the applicable altitude and operating condition because of supercharging.

(Change 525-3 (91-11-01))

525.1101 Carburettor Air Preheater Design

Each carburettor 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.

525.1103 Induction System Ducts and Air Duct Systems

(a) Each induction system duct upstream of the first stage of the engine supercharger and of the auxiliary power unit compressor must have a drain to prevent the hazardous accumulation of fuel and moisture in the ground attitude. No drain may discharge where it might cause a fire hazard.

(b) Each induction system duct must be:

(1) Strong enough to prevent induction system failures resulting from normal backfire conditions; and

(2) Fire resistant if it is in any fire zone for which a fire-extinguishing system is required, except that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone.

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

(d) For turbine engine and auxiliary power unit bleed air duct systems, no hazard may result if a duct failure occurs at any point between the air duct source and the aeroplane unit served by the air.

(e) Each auxiliary power unit induction system duct must be fireproof for a sufficient distance upstream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through auxiliary power unit ducts and entering any other compartment or area of the aeroplane 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.

 

525.1105 Induction System Screens

If induction system screens are used:

(a) Each screen must be upstream of the carburettor;

(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 de-iced by heated air;

 

(c) No screen may be de-iced by alcohol alone; and

(d) It must be impossible for fuel to strike any screen.

525.1107 Inter-coolers and After-coolers

Every inter-cooler and after-cooler must be able to withstand any vibration, inertia, and air pressure load to which it would be subjected in operation.

Exhaust System

525.1121 General

For powerplant and auxiliary power unit installations the following apply:

(a) Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment. For test purposes, any acceptable carbon monoxide detection method may be used to show the absence of carbon monoxide.

(b) Each exhaust system part with a surface hot enough to ignite flammable fluids or vapours must be located or shielded so that leakage from any system carrying flammable fluids or vapours will not result in a fire caused by impingement of the fluids or vapours on any part of the exhaust system including shields for the exhaust system.

(c) Each component that hot exhaust gases could strike, or that could be subjected to high temperatures from exhaust system parts, must be fireproof. All exhaust system components must be separated by fireproof shields from adjacent parts of the aeroplane 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 vapours external to the shroud.

525.1123 Exhaust Piping

For powerplant and auxiliary power unit installations, the following apply:

(a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures.

(b) Piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation; and

(c) Piping connected to components between which relative motion could exist must have means for flexibility.

525.1125 Exhaust Heat Exchangers

For reciprocating engine powered aeroplanes, the following apply:

(a) Each exhaust heat exchanger must be constructed and installed to withstand each vibration, inertia, and other load to which it would be subjected in operation. In addition:

(1) Each exchanger must be suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;

(2) There must be means for the inspection of the critical parts of each exchanger;

(3) Each exchanger must have cooling provisions wherever it is subject to contact with exhaust gases; and

(4) No exhaust heat exchanger or muff may have any stagnant areas or liquid traps that would increase the probability of ignition of flammable fluids or vapours that might be present in case of the failure or malfunction of components carrying flammable fluids.

(b) If an exhaust heat exchanger is used for heating ventilating air:

(1) There must be a secondary heat exchanger between the primary exhaust gas heat exchanger and the ventilating air system; or

(2) Other means must be used to preclude the harmful contamination of the ventilating air.

525.1127 Exhaust Driven Turbo-Superchargers

(a) Each exhaust driven turbo-supercharger must be approved or shown to be suitable for the particular application. It must be installed and supported to ensure safe operation between normal inspections and overhauls. In addition, there must be provisions for expansion and flexibility between exhaust conduits and the turbine.

(b) There must be provisions for lubricating the turbine and for cooling turbine parts where temperatures are critical.

(c) If the normal turbo-supercharger control system malfunctions, the turbine speed may not exceed its maximum allowable value. Except for the waste gate operating components, the components provided for meeting this requirement must be independent of the normal turbo-supercharger controls.

Powerplant Controls and Accessories

525.1141 Powerplant Controls: General

Each powerplant control shall be located, arranged, and designed under 525.777 through 525.781 and marked under 525.1555. In addition, it shall meet the following requirements:
(amended 2005/06/03; previous version)

(a) Each control shall be located so that it cannot be inadvertently operated by persons entering, leaving, or moving normally in, the cockpit.
(amended 2005/06/03; previous version)

(b) Each flexible control shall be approved or shall be shown to be suitable for the particular application.
(amended 2005/06/03; previous version)

(c) Each control shall have sufficient strength and rigidity to withstand operating loads without failure and without excessive deflection.
(amended 2005/06/03; previous version)

(d) Each control shall be able to maintain any set position without constant attention by flight crew members and without creep due to control loads or vibration.
(amended 2005/06/03; previous version)

(e) The portion of each powerplant control located in a designated fire zone that is required to be operated in the event of fire shall be at least fire resistant.
(amended 2005/06/03; previous version)

(f) For powerplant valve controls located in the flight deck there shall be a means:
(amended 2005/06/03; previous version)

(1) For the flight crew to select each intended position or function of the valve; and
(amended 2005/06/03; previous version)

(2) To indicate to the flight crew:
(amended 2005/06/03; previous version)

(i) The selected position or function of the valve; and
(amended 2005/06/03; previous version)

(ii) When the valve has not responded as intended to the selected position or function.
(amended 2005/06/03; previous version)

(Change 525-3 (91-11-01))

525.1142 Auxiliary Power Unit Controls

Means must be provided on the flight deck for starting, stopping, and emergency shutdown of each installed auxiliary power unit.

525.1143 Engine Controls

(a) There must be a separate power or thrust control for each engine.

(b) Power and thrust controls must be arranged to allow:

(1) Separate control of each engine; and

(2) Simultaneous control of all engines.

(c) Each power and thrust control must provide a positive and immediately responsive means of controlling its engine.

(d) For each fluid injection (other than fuel) system and its controls not provided and approved as part of the engine, the applicant must show that the flow of the injection fluid is adequately controlled.

(e) If a power or thrust control incorporates a fuel shut-off feature, the control must have a means to prevent the inadvertent movement of the control into the shut-off 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 shut-off position.

525.1145 Ignition Switches

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

(b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control.

(c) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means to prevent its inadvertent operation.

525.1147 Mixture Controls

(a) If there are mixture controls, each engine must have a separate control. The controls must be grouped and arranged to allow:

(1) Separate control of each engine; and

(2) Simultaneous control of all engines.

(b) Each intermediate position of the mixture controls that corresponds to a normal operating setting must be identifiable by feel and sight.

(c) The mixture controls must be accessible to both pilots. However, if there is a separate flight engineer station with a control panel, the controls need be accessible only to the flight engineer.

525.1149 Propeller Speed and Pitch Controls

(a) There must be a separate propeller speed and pitch control for each propeller.

(b) The controls must be grouped and arranged to allow:

(1) Separate control of each propeller; and

(2) Simultaneous control of all propellers.

(c) The controls must allow synchronisation of all propellers.

(d) The propeller speed and pitch controls must be to the right of, and at least one inch below, the pilot’s throttle controls.

525.1153 Propeller Feathering Controls

(a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation.

(b) If feathering is accomplished by movement of the propeller pitch or speed control lever, there must be means to prevent the inadvertent movement of this lever to the feathering position during normal operation.

525.1155 Reverse Thrust and Propeller Pitch Settings Below the Flight Regime

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

525.1157 Carburettor Air Temperature Controls

There must be a separate carburettor air temperature control for each engine.

525.1159 Supercharger Controls

Each supercharger control must be accessible to the pilots or, if there is a separate flight engineer station with a control panel, to the flight engineer.

525.1161 Fuel Jettisoning System Controls

Each fuel jettisoning system control must have guards to prevent inadvertent operation. No control may be near any fire extinguisher control or other control used to combat fire.

525.1163 Powerplant Accessories

(a) Each engine-mounted accessory must:

(1) Be approved for mounting on the engine involved;

(2) Use the provisions on the engine for mounting; and

(3) Be sealed to prevent contamination of the engine oil system and the accessory system.

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

(c) If continued rotation of an engine-driven cabin supercharger or of any remote accessory driven by the engine is hazardous if malfunctioning occurs, there must be means to prevent rotation without interfering with the continued operation of the engine.

525.1165 Engine Ignition Systems

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

(b) The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw electrical energy from the same source.

(c) The design of the engine ignition system must account for:

(1) The condition of an inoperative generator;

(2) The condition of a completely depleted battery with the generator running at its normal operating speed; and

(3) The condition of a completely depleted battery with the generator operating at idling speed, if there is only one battery.

(d) Magneto ground wiring (for separate ignition circuits) that lies on the engine side of the fire wall, must be installed, located, or protected, to minimise the probability of simultaneous failure of two or more wires as a result of mechanical damage, electrical faults, or other cause.

(e) No ground wire for any engine may be routed through a fire zone of another engine unless each part of that wire within that zone is fireproof.

(f) Each ignition system must be independent of any electrical circuit not used for assisting, controlling, or analysing the operation of that system.

(g) There must be means to warn appropriate flight crew members if the malfunctioning of any part of the electrical system is causing the continuous discharge of any battery necessary for engine ignition.

(h) Each engine ignition system of a turbine powered aeroplane must be considered an essential electrical load.

(Change 525-3 (91-11-01))

525.1167 Accessory Gearboxes

For aeroplanes equipped with an accessory gearbox that is not certified as part of an engine:

(a) The engine with gearbox and connecting transmissions and shafts attached must be subjected to the tests specified in 533.49 or 533.87 of this manual as applicable.

(b) The accessory gearbox must meet the requirements of 533.25 and 533.53 or 533.91 of this manual as applicable; and

(c) Possible misalignments and torsional loadings of the gearbox, transmission, and shaft system, expected to result under normal operating conditions must be evaluated.

Powerplant Fire Protection

525.1181 Designated Fire Zones; Regions Included

(a) Designated fire zones are:

(1) The engine power section;

(2) The engine accessory section;

(3) Except for reciprocating engines, any complete powerplant compartments in which no isolation is provided between the engine power section and the engine accessory section;

(4) Any auxiliary power unit compartment;

(5) Any fuel-burning heater and other combustion equipment installation described in 525.859.

(6) The compressor and accessory sections of turbine engines; and

(7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases.

(b) Each designated fire zone shall meet the requirements of 525.863, 525.865, 525.867, 525.869 and 525.1185 through 525.1203.
(amended 2005/06/03; previous version)

(Change 525-3 (91-11-01))

525.1182  Nacelle Areas Behind Firewalls, and Engine Pod Attaching Structures Containing Flammable Fluid Lines

(a) Each nacelle area immediately behind the firewall, and each portion of any engine pod attaching structure containing flammable fluid lines, must meet each requirement of 525.1103(b), 525.1165(d) and (e), 525.1183, 525.1185(c), 525.1187, 525.1189, and 525.1195 through 525.1203, including those concerning designated fire zones. However, engine pod attaching structures need not contain fire detection or extinguishing means.

 

(b) For each area covered by paragraph (a) of this section that contains a retractable landing gear, compliance with that paragraph need only be shown with the landing gear retracted.

 

525.1183 Flammable Fluid-Carrying Components

(a) Except as provided in paragraph (b) of this section, each line, fitting, and other component carrying flammable fluid in any area subject to engine fire conditions, and each component which conveys or contains flammable fluid in a designated fire zone must be fire resistant, except that flammable fluid tanks and supports in a designated fire zone must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non-fireproof part will not cause leakage or spillage of flammable fluid. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid. An integral oil sump of less than 25 quart capacity on a reciprocating engine need not be fireproof nor be enclosed by a fireproof shield.

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

(1) Lines, fittings, and components which are already approved as part of a type certificated engine; and

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

(c) All components, including ducts, within a designated fire zone shall be fireproof if, when exposed to or damaged by fire, they could:
(amended 2001/05/07; no previous version)

(1) result in fire spreading to other regions of the aeroplane; or

(2) cause unintentional operation of, or inability to operate, essential services or equipment.

525.1185 Flammable Fluids

(a) Except for the integral oil sumps specified in 525.1183(a), no tank or reservoir that is a part of a system containing flammable fluids or gases may be in a designated fire zone unless the fluid contained, the design of the system, the materials used in the tank, the shut-off means, and all connections, lines, and control provide a degree of safety equal to that which would exist if the tank or reservoir were outside such a zone.

(b) There must be at least one-half inch of clear airspace between each tank or reservoir and each firewall or shroud isolating a designated fire zone.

(c) Absorbent materials close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids.

525.1187 Drainage and Ventilation of Fire Zones

(a) There must be complete drainage of each part of each designated fire zone to minimise the hazards resulting from failure or malfunctioning of any component containing flammable fluids. The drainage means must be:

(1) Effective under conditions expected to prevail when drainage is needed; and

(2) Arranged so that no discharged fluid will cause an additional fire hazard.

(b) Each designated fire zone must be ventilated to prevent the accumulation of flammable vapours.

(c) No ventilation opening may be where it would allow the entry of flammable fluids, vapours, or flame from other zones.

(d) Each ventilation means must be arranged so that no discharged vapours will cause an additional fire hazard.

(e) Unless the extinguishing agent capacity and rate of discharge are based on maximum air flow through a zone, there must be means to allow the crew to shut off sources of forced ventilation to any fire zone except the engine power section of the nacelle and the combustion heater ventilating air ducts.

525.1189 Shut-off Means

(a) Each engine installation and each fire zone specified in 525.1181(a)(4) and (5) must have a means to shut off or otherwise prevent hazardous quantities of fuel, oil, de-icer, and other flammable fluids, from flowing into, within, or through any designated fire zone, except that shut-off means are not required for:

(1) Lines, fittings, and components forming an integral part of an engine; and

(2) Oil systems for turbine engine installations in which all components of the system in a designated fire zone, including oil tanks, are fireproof or located in areas not subject to engine fire conditions.

(b) The closing of any fuel shut-off valve for any engine may not make fuel unavailable to the remaining engines.

(c) Operation of any shut-off may not interfere with the later emergency operation of other equipment, such as the means for feathering the propeller.

(d) Each flammable fluid shut-off means and control must be fireproof or must be located and protected so that any fire in a fire zone will not affect its operation.

(e) No hazardous quantity of flammable fluid may drain into any designated fire zone after shut-off.

(f) There must be means to guard against inadvertent operation of the shut-off means and to make it possible for the crew to reopen the shut-off means in flight after it has been closed.

(g) Each tank-to-engine shut-off valve must be located so that the operation of the valve will not be affected by powerplant or engine mount structural failure.

(h) Each shut-off valve must have a means to relieve excessive pressure accumulation unless a means for pressure relief is otherwise provided in the system.

525.1191 Firewalls

(a) Each engine, auxiliary power unit, fuel-burning heater, other combustion equipment intended for operation in flight, and the combustion, turbine, and tailpipe sections of turbine engines, must be isolated from the rest of the aeroplane by firewalls, shrouds, or equivalent means.

(b) Each firewall and shroud must be:

(1) Fireproof;

(2) Constructed so that no hazardous quantity of air, fluid, or flame can pass from the compartment to other parts of the aeroplane;

(3) Constructed so that each opening is sealed with close fitting fireproof grommets, bushings, or firewall fittings, and;

(4) Protected against corrosion.

525.1192 Engine Accessory Section Diaphragm

For reciprocating engines, the engine power section and all portions of the exhaust system must be isolated from the engine accessory compartment by a diaphragm that complies with the firewall requirements of 525.1191.

525.1193 Cowling and Nacelle Skin

(a) Each cowling must be constructed and supported so that it can resist any vibration, inertia, and air load to which it may be subjected in operation.

(b) Cowling must meet the drainage and ventilation requirements of 525.1187.

(c) On aeroplanes 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 525.1191.

(d) Each part of the cowling subject to high temperatures due to its nearness to exhaust system parts or exhaust gas impingement must be fireproof.

(e) Each aeroplane must:

(1) Be designed and constructed so that no fire originating in any fire zone can enter, either through openings or by burning through external skin, any other zone or region where it would create additional hazards;

(2) Meet subparagraph (1) of this paragraph with the landing gear retracted (if applicable); and

(3) Have fireproof skin in areas subject to flame if a fire starts in the engine power or accessory sections.

525.1195 Fire Extinguishing Systems

(a) Except for combustor, turbine, and tailpipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases for which it is shown that a fire originating in these sections can be controlled, there must be a fire extinguisher system serving each designated fire zone.

(b) The fire extinguishing system, the quantity of the extinguishing agent, the rate of discharge, and the discharge distribution must be adequate to extinguish fires. It must be shown by either actual or simulated flight tests that under critical airflow conditions in flight the discharge of the extinguishing agent in each designated fire zone specified in paragraph (a) of this section will provide an agent concentration capable of extinguishing fires in that zone and of minimising the probability of reignition. An individual "one-shot" system may be used for auxiliary power units, fuel burning heaters, and other combustion equipment. For each other designated fire zone, two discharges must be provided each of which produces adequate agent concentration.

(c) The fire-extinguishing system for a nacelle must be able to simultaneously protect each zone of the nacelle for which protection is provided.

525.1197 Fire Extinguishing Agents

(a) Fire extinguishing agents must:

(1) Be capable of extinguishing flames emanating from any burning of fluids or other combustible materials in the area protected by the fire extinguishing system; and

(2) Have thermal stability over the temperature range likely to be experienced in the compartment in which they are stored.

(b) If any toxic extinguishing agent is used, provisions must be made to prevent harmful concentrations of fluid or fluid vapours (from leakage during normal operation of the aeroplane or as a result of discharging the fire extinguisher on the ground or in flight) from entering any personnel compartment, even though a defect may exist in the extinguishing system. This must be shown by test except for built-in carbon dioxide fuselage compartment fire extinguishing systems for which:

(1) Five pounds or less of carbon dioxide will be discharged, under established fire control procedures, into any fuselage compartment; or

(2) There is a protective breathing equipment for each flight crew member on flight deck duty.

525.1199 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 aeroplane. The line must also be located or protected to prevent clogging caused by ice or other foreign matter.

(c) There must be a means for each fire extinguishing agent container to indicate that the container has discharged or that the charging pressure is below the established minimum necessary for proper functioning.

(d) The temperature of each container must be maintained, under intended operating conditions, to prevent the pressure in the container from:

(1) Falling below that necessary to provide an adequate rate of discharge; or

(2) Rising high enough to cause premature discharge.

(e) If a pyrotechnic capsule is used to discharge the extinguishing agent, each container must be installed so that temperature conditions will not cause hazardous deterioration of the pyrotechnic capsule.

525.1201 Fire Extinguishing System Materials

(a) No material in any fire extinguishing system may react chemically with any extinguishing agent so as to create a hazard.

(b) Each system component in an engine compartment must be fireproof.

525.1203 Fire Detector System

(a) There must be approved, quick acting fire or overheat detectors in each designated fire zone, and in the combustion, turbine, and tailpipe sections of turbine engine installations, in numbers and locations ensuring prompt detection of fire in those zones.

(b) Each fire detector system must be constructed and installed so that:

(1) It will withstand the vibration, inertia, and other loads to which it may be subjected in operation;

(2) There is a means to warn the crew in the event that the sensor or associated wiring within a designated fire zone is severed at one point, unless the system continues to function as a satisfactory detection system after the severing; and

(3) There is a means to warn the crew in the event of a short circuit in the sensor or associated wiring within a designated fire zone, unless the system continues to function as a satisfactory detection system after the short circuit.

(c) No fire or overheat detector may be affected by any oil, water, other fluids, or fumes that might be present.

(d) There must be means to allow the crew to check, in flight, the functioning of each fire or overheat detector electric circuit.

(e) Components of each fire or overheat detector system in a fire zone must be fire resistant.
(amended 2009/05/11; previous version)

(f) No fire or overheat detector system component for any fire zone may pass through another fire zone, unless:

(1) It is protected against the possibility of false warnings resulting from fires in zones through which it passes; or

(2) Each zone involved is simultaneously protected by the same detector and extinguishing system.

(g) Each fire detector system must be constructed so that when it is in the configuration for installation it will not exceed the alarm activation time approved for the detectors using the response time criteria specified in the appropriate Technical Standard Order for the detector.

(h) EWIS for each fire or overheat detector system in a fire zone must meet the requirements of 525.1731.
(amended 2009/05/11; no previous version)

525.1207 Compliance

Unless otherwise specified, compliance with the requirements of 525.1181 through 525.1203 must be shown by a full scale fire test or by one or more of the following methods:

(a) Tests of similar powerplant configurations;

(b) Tests of components;

(c) Service experience of aircraft with similar powerplant configurations;

(d) Analysis.