b. Procedures.
(1) Combustible fluid supply lines which pass into, within, or through a firewall into the fire zone must incorporate shutoff valves. This requirement does not apply to lines, fittings, and components which were certified with and are part of the engine.
These requirements do not apply to oil systems for reciprocating engines with less than 500 cubic inches displacement or to any other installation where all components, including the oil tanks, are fireproof or are located in an area that will not be affected by an engine fire.
(2) Eight fluid ounces or less of a combustible fluid is not considered hazardous and no more than this amount should be present after activating the shutoff valve.
(3) Engine isolation is to be maintained when incorporating shutoff valves into engine tuel and lubrication lines. The design should ensure that when one engine is shut down or fails and the fuel and lubrication fluid shutoff valves are activated, the remaining good engine is not affected in any way, and the rotorcraft can continue safe flight to a landing. This should be demonstrated by test.
(4) Each shutoff valve located in a fire zone should be fireproof. If the shutoff valve is located outside of the fire zone, then it should be at least fire resistant or protected so that it will function under a worst case fire condition within a fire zone. This should be demonstrated by test.
(5) For primary propulsion engine installations, the flammable fluid shutoff should be protected from inadvertent operation. Where electrical shutoffs are used, the switches should be guarded or require double actions. If the shutoffs are mechanically activated, the design of the knob and the location of the lever should be such that inadvertent actuation cannot occur. It must be possible to reopen the shutoff valve after it has been closed and this should be demonstrated by test.
AC 27.1191. § 27.1191 (Amendment 27-2) FIREWALLS.
a. Explanation. This section states the certification requirements for the use of fireproof protective devices such as firewalls, shrouds, or equivalent. These devices are necessary to isolate each engine (including combustor, turbine, and tailpipe sections of turbine engines and auxiliary propulsion units (APU); each APU; each combustion heater; each unit of combustion equipment; or each high temperature device (or source) from personnel compartments and critical components (not already protected under § 27.861). The isolation of these fire zones is necessary to prevent the spread of fires, prevent or minimize thermal injuries and fatalities, and prevent damage to critical components that are essential to a controlled landing. Even though § 27.1191(b) implicitly excludes AP's, combustion heaters, and other combustion equipment that are not used in flight; they should be protected by fireproof enclosures, because of the requirements of the relevant parts of §§ 27.1183 through 27.1203. This is because, even if the device is rendered inoperative in flight, it typically contains residual heat, Page E - 128
fuel, fumes and potential ignition sources (i.e., "potential hazards"). Each fireproof protective device must, by regulation, meet the following criteria:
(1) Its design and location must take into account the probable fire path from each fire zone or source considering factors such as internal airflow, external airflow, and gravity.
(2) It must be constructed so that no hazardous quantity of air, fumes, fluids, or flame can propagate through it to unprotected parts of the rotorcraft.
(3) Its openings (e.g., shaftholes, lineholes, etc.) must be sealed with close fitting fireproof grommets, bushings, bearings, firewall, fittings, or equivalent that prevent burn through and leakage of hazardous fumes or fluids from the fire zone.
(4) It must be fireproof (see definition).
(5) It must be either corrosion resistant or otherwise safely protected from corrosion.
b. Definitions.
(1) Fireproof Protective Device. A fireproof protective device is a device such as a firewall, shroud, enclosure, or equivalent used to isolate a heat or potential fire source (severe fire) from personnel compartments and from critical aircraft components which are essential for a controlled landing.
(2) Fireproof. Fireproof is defined in § 1.1 "General Definitions."
(3) Controlled Landing. A landing which is survivable (i.e., does not fatally injure all occupants) but may produce an unairworthy, partially salvageable, or unsalvageable rotorcraft.
(4) Severe Fire. See Definition in paragraph AC 27.859.
c. Procedures. Fireproof protective devices are typically certified by analysis, tests, or a combination conducted during the certification process, including flight tests or simulated flight tests, as follows:
(1) Fireproof protective devices should be provided wherever a hazard exists which requires isolation from a severe fire to avoid fires in personnel compartments and to avoid thermal damage to critical components (such as structural elements, controls, rotor mechanisms, and system components) that are necessary for a controlled landing.
A thorough hazard analysis should be conducted during certification to identify, define and quantify in order of severity (i.e., maximum temperature, hot exposed area, etc.) all thermal hazards or zones that require fireproof protection in a given design. Engines (including the combustor, turbine, and tailpipe sections of turbine engines), APU's, Page E - 129
combustion heaters, and combustion devices are required by regulation to be isolated.
Other high temperature devices may also require isolation because of local hot spots (which occur during normal operations or from failure modes) that can thermally injure occupants or cause spontaneous combustion of surroundings. A hazard analysis should identify these potential problems and provide proper certification solutions.
(2) Fireproof protective devices should be able to withstand at least 2000 ‡ 150° F for at least 15 minutes (reference AC 20-135). The fireproof protective device should allow protected parts, subsystems or systems to perform their intended function for the duration of a severe fire (see definitions). For firewalls, examples of flat, geometry materials undergoing uniform heat fluxes with material gauges that automatically meet the certification requirements are given in figure AC 27.1191-1. If firewalls are utilized that involve other materials, significant geometric changes, or significantly non-uniform heat fluxes, then automatic compliance may not be assured.
In such cases the fireproof protective device should be analyzed using the severe fire definition and, in some cases, tested in accordance with AC 23-2 to ensure proper certification. For example, a curved protective surface may absorb a uniform incident heat flux unevenly and create a local hot spot that exceeds 2,150° degrees Fahrenheit that burns through in less than 15 minutes; whereas, a flat surface of equal thickness might not exceed 2,150° degrees Fahrenheit and would not burn through in less than 15 minutes. It should be noted that composite materials are not generally used for protective devices because of their inability to withstand high temperatures (i.e., exceedance of the glass transition temperature); however, some specially formulated composites have been previously certified as engine cowlings. Titanium is an acceptable material for fireproof protective devices such as firewalls. However, use of titanium should always be carefully considered and reviewed, because it can lose all structural ability and burn severely (self combust) above 1,050° F, under certain thermodynamic environments, and contribute to the fire instead of providing the intended fire protection. AC 33-4, "Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines" and MIL-HDBK-5D contain more detailed information on the unique thermal properties of titanium.
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FIGURE AC 27.1191-1
TABLE OF MATERIALS AND GAGES ACCEPTABLE
FOR FIREPROOF PROTECTIVE DEVICES WITH FLAT SURFACE GEOMETRIES(") MATERIAL 2) MINIMUM THICKNESS (3) Titanium Sheet .016 in Stainless Steel .015 in Mild Carbon Steel .018 in Terne Plate .018 in Monel Metal 018 in Firewall Fittings .018 in(4) (Steel or Copper Base) NOTES:
(1) Assumes essentially flat vertical or horizontal surfaces undergoing a uniform heat flux. Any significant variation in either geometry or heat flux distribution should be examined in detail for adequate gauge thicknesses on a case-by-case basis.
(2) Must have corrosion protection if not inherent in the material itself.
(3) The minimum thickness is for thermal containment only. Structural integrity considerations may require thickness increases. MIL-HDBK-5D contains material allowable versus temperature data for most common metallic materials.
(4) This is the minimum wall thickness measured at the smallest dimension (e.g., thread root or other location) of the part.
(5) Distortion of thin sheet materials and the subsequent gapping at lap joints or between rivets is difficult to predict; therefore, testing of the simulated installation is necessary to prove the integrity of the design. However, rivet pitches of 2 inches or less on non load-carrying titanium firewalls of .020 inch or steel firewalls of .018 inch are acceptable without further testing.
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(3) The probable path of a fire (as affected by internal and external air flow during normal flight and autorotation, gravity, flame propagation paths, or other considerations) should be taken into account when performing the hazard analysis of item (1). Such a review will ensure that fireproof protective devices are placed in the proper location for intercepting, blocking or containing a severe fire before occupants are injured and a controlled landing is prevented. If the probable path cannot be readily determined by inspection or analysis, testing using simulated air flows, rotorcraft attitudes, and dyed inert fluids or vapors can be used to aid in this determination.
(4) Each opening in a protective device should be sealed with close fitting sealing devices such as fireproof grommets, bushings, firewall fittings, rotating seals or equivalent that are at least as effective as the fireproof protective device itself. This is necessary to ensure that no local breakdowns in protection occur. For materials not listed as acceptable in item (1), analysis and testing should be required in accordance with FAA/AUTHORITY standards and the definition of a severe fire for proper substantiation.
(5) Each protective device should be fireproof in order to withstand a severe fire. Unless designs and materials have been previously FAA/AUTHORITY approved
(e.g., see Item 1), the protective device's design and material selection should be tested to ensure its fireproof thermal and structural integrity. A full-scale test of a structurally loaded article or a representative sample should be conducted to ensure proper compliance is achieved. Also, the continued sealing ability of the protective device in its deformed state due to a hard controlled landing should be considered during certification (e.g., use of ductile materials). The corrosion environment should be defined and appropriate protection provided. Phased inspections should be specified, if necessary, to ensure continued corrosion integrity. Certification tests for adequacy of corrosion protection should be conducted using sample plates or by other equivalent means, as required.
AC 27.1193. § 27.1193 COWLING AND ENGINE COMPARTMENT COVERING.
a. Explanation.
(1) Section 27.1193(a) requires the cowling and engine compartment coverings to structurally withstand loads experienced in flight.
(2) In order to prevent pooling of flammable fluids, § 27.1193(b) requires rapid and complete drainage from the cowling and engine compartment.
(3) Section 27.1193(c) requires the drain of paragraph (b) to purge the fluid in such a manner not to create a fire hazard.
(4) Section 27.1193(d) requires the cowling and engine compartment covering to be at least fire resistant and paragraph (e) requires them to be fireproof where they may experience high temperatures due to the exhaust system.
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b. Procedures.
(1) Compliance with § 27.1193(a) can be shown by analyzing the cowling and engine compartment covering and determining that no structural degradation will occur under the highest loads experienced on the ground or in flight.
(2) Compliance with § 27.1193(b) can be accomplished by ensuring that the drain will discharge positively with no traps and is a minimum of 0.25 inches in diameter.
(3) Compliance with § 27.1193(c) can be demonstrated by colored liquid flowing through the drain system while in flight. The dye should not impinge on any ignition source during any approved flight regime.
(4) Compliance with § 27.1193(d) can be accomplished by showing that the cowling and engine compartment covering is fire resistant. Fire resistant in this context means a material that has the capacity, under expected service conditions (load, vibration, airflow), to withstand the heat associated with fire at least as well as aluminum alloy in dimensions appropriate for the purpose.
(5) Compliance with § 27.1193(e) can be accomplished by showing that the cowling and engine compartment coverings retain adequate structural integrity when subjected to elevated temperatures that may be expected in service.
AC 27.1193A. § 27.1193 (Amendment 27-23) COWLING AND ENGINE COMPARTMENT COVERING.
a. Explanation. Amendment 27-23 adds a new § 27.1193(f) that requires redundant retention means for each panel, cowling, engine, or rotor drive system covering that can be opened or readily removed. Conventional fasteners for these devices are subject to frequent operation by maintenance personnel and have deteriorated, failed from wear or vibration, or been left unsecured after preflight inspections. Such a failure could be hazardous if a loose panel, cowling, or covering comes in contact with the rotors or critical controls.
b. Procedures. All of the policy material pertaining to this section remains in effect with the following additions:
(1) Compliance with § 27.1193(f) can be accomplished by simulating, or actually failing, one or more of the retention devices or by structural analysis. It should be shown that the cowling or cover will not open, strike, or be struck by the rotor or other critical component.
(2) Consideration should be given to minimize the possibility of latches being improperly closed that could result in a cowl coming open in flight.
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(3) The failure of one latching device should not cause the failure of another latching device. If a failure of a single retention device can contribute to multiple failures, these multiple failures should be considered.
(4) The consequences of "forgetting" to latch a cowl should be considered.
(5) The use of safety straps should be considered to minimize the impact of a latching device failure.
AC 27.1194. § 27.1194 (Amendment 27-2) OTHER SURFACES.
a. Explanation. This section states the fire resistance requirements for material surfaces near engine compartments and designated fire zones (other than tail surfaces not subject to heat, flames or sparks emanating from a designated fire zone or engine compartment).
b. Definition.
(1) Other Surface. Any airframe, system, or powerplant component aft of and near an engine compartment, a designated fire zone, or another heat source which would receive a heat flux as a result of a fire in the engine compartment or fire zone that would require the component to be fire resistant.
(2) Fire Resistant. In accordance with § 1.1, is defined as follows:
Sheet metal or structural members with the capacity to withstand the heat associated with the fire at least as well as aluminum alloy in dimensions appropriate for the purpose for which they are used.
(ii) Fluid carrying lines, fluid system parts, wiring, air ducts, fittings and powerplant controls with the capacity to perform their intended functions under the heat and other conditions resulting from a fire.
(3) Fire. A fire in either an engine compartment or a designated fire zone is assumed to occur that produces a heat flux on a system, airframe or powerplant component aft of or near the fire. The effect of each such fire on other surfaces must be considered on a case-by-case basis to determine the critical case. Unless a more rationale definition is furnished and approved during certification, the fire in any engine compartment or designated fire zone should be assumed, for purposes of analysis, to be a severe fire (see definition in paragraph AC 27.859).
c. Procedures.
(1) Other surfaces should be identified during certification by a design review and by a conservative, thorough hazard analysis based on an analytical estimate of the total heat flux (i.e., conduction, convection, and radiation in combination, as applicable) Page E - 134
using the definition of a severe fire and of the resultant "other surface" temperature based on a single fire occurring in each engine compartment and designated fire zone, on a case-by-case basis. Once the other surfaces are identified and their severe fire induced maximum temperatures determined, their configuration and material selection should be reviewed on a case-by-case basis to determine either that they are fire resistant, that they can be made fire resistant (within the limits of practicability), or that it is impracticable to make them fire resistant. If the non-fire resistant other surfaces can be readily made fire resistant they should be. If it is impracticable to make them fire resistant, then they should be relocated, insulated, or a combination in order to reduce the total incident heat flux (and, thus, lower their surface temperature) so that they no longer need be fire resistant. If insulation is used to shield a surface that is subjected to a significant temperature, it must be fire resistant.
(2) A partial validation of analytical heat flux models using the definition of a severe fire can sometimes be achieved during certification tests by using thermocouples or heat-sensitive stickers to measure in-flight temperature ranges and distributions on other surfaces from known thermal environments in engine compartments or other designated fire zones.
AC 27.1195. § 27.1195 (Amendment 27-5) FIRE DETECTOR SYSTEMS.
a. Explanation.
(1) This section requires quick-acting fire detectors to be installed on turbine powered rotorcraft, when the engine compartment cannot be readily observed in flight by the pilot in the cockpit.
(2) The number of detectors and locations must be sufficient to ensure prompt detection of fire in the engine compartment.
b. Procedures.
(1) The detector system should be designed for highest reliability to detect a fire and not to give a false alarm. It is desirable that it only responds to a fire and misinterpretation with a lesser hazard should not be possible. Engine overtemperature, harmless exhaust leakage, and bleed air leakage should not be indicated by a fire detector system. A fire detection system should be reserved for a condition requiring immediate measures such as engine shutdown or fire extinguishing. There are three general types of detector-procedure systems that are commonly used:
(i) A manual system utilizes warning lights to alert the pilot who then follows prescribed cockpit procedure as a countermeasure. A manual system is adequate for hazards in which a few seconds are not important.
(II) There is also a semi-automatic system. Occasionally a rotorcraft becomes so complex that the emergency procedure exceeds reasonable expectations Page E - 135
of the pilot. In such cases, psychology should be weighted against complexity, and "panic switches," combining multiple procedure functions, should be provided to simplify the mental demands on the pilot. Speed is gained by such designs for hazards which may need it.
(¡¡i) The detector of an automatic system automatically triggers the appropriate countermeasures and warns the pilot simultaneously. Such a system should be carefully evaluated to assure that the advantages outweigh the disadvantages and potential malfunctions.
(2) Fires, or dangerous fire conditions can be detected by means of various existing techniques. The following is a partial list of available detectors:
(1) Radiation-sensing detectors.
(ii) Rate-of-temperature-rise detectors.
(iii) Overheat detectors.
(iv) Smoke detectors.
(v) CO detectors.
(vi) Combustible mixture detectors.
(vii) Fibre-optic detectors.
(vill) Ultraviolet.
(ix) Observation of crew or passengers.
(3) In many rotorcraft it is desirable to have a detection system which incorporates several of these different types of detectors. Radiation-sensing detectors are most useful where the materials present will burn brightly soon after ignition, such as in the powerplant accessory section. Rate of rise detectors are well-suited to compartments of normally low ambient temperatures and low rates of temperature rise where a fire would produce a high temperature differential and rapid temperature rise. It should be noted that under certain circumstances, where a relatively slow temperature increase occurs over a considerable period of time, a fire can occur without detection by rate of rise detectors. Overheat detectors should be used wherever the hazard is evidenced by temperatures exceeding a predicted, set value. Smoke detectors may be suited to low air flow areas where materials may burn slowly, or smolder. Fibre-optic detectors can be used to visually observe the existence of flame or smoke. The three major detector types used for fast detection of fires are the radiation-sensing, rate-of-rise, and overheat detectors. Radiation-sensing detectors are basically "volume" Page E - 136
type which senses flame within a visible space. Overheat-fire detectors can be obtained in either "continuous" or "unit" type.
(4) The detector system should:
Indicate fire within 15 seconds after ignition, and show in which engine compartment the fire is located.
(ii) Remain on for the duration of the fire.
(ili) Indicate when the fire is out.
(iv) Indicate re-ignition of the fire.
(v) Not by itself precipitate or add to the potential of any other hazards.
(vi) Not cause false warnings under any flight or ground operating condition.
(5) Additional features of the detection system are as follows:
A means should be incorporated so that operation of the system can be tested from the cockpit.
(ii) Detector units should be of rugged construction, to resist maintenance handling, exposure to fuel, oil, dirt, water, cleaning agent, extreme temperatures, vibration, salt air, fungus, and altitude. Also, they should be light in weight, small, and compact, and readily adaptable to desired positions of mounting.
(i¡i) The detector system should operate on the rotorcraft electric system without inverters. The circuit should require minimum current unless indicating a fire or unless a monitoring system is in use.
(iv) Fixed temperature fire detectors should preferably be set at 100° F
(37.7° C) to 150° F (65.6° C) above maximum safe ambient temperature, or higher when in compartments where extremely high rate of rise is normally encountered.
(v) Detector system components located within fire zones should be fireproof.
(vi) Each detector system should actuate a warning device which indicates the location of the fire. If fire warning lights are used, they must be in the pilot's normal field of view.
(vil) Two or more engines should not be dependent upon any one detector circuit. The installation of common zone detection equipment prevents the detection Page E - 137
system from distinguishing between the engine installations, necessitating shutting down more than one engine.
(6) The sensing portion of the fire detection system should not extend outside of the coverage area into another fire zone. Detectors, with the exception of radiation-sensing detectors, should be located at points where the ventilation air leaves compartments. If a reverse-flow cooling system is used, detectors should be installed at locations which are outlets under both flight and ground operating conditions. Stagnant air spaces should be avoided and the number of ventilation air exits should be kept to a minimum. Compliance with these recommendations allow the effective placement of a minimum amount of detectors, and still ensure prompt detection of fire in those zones.
Radiation-sensing detectors should be located such that any flame within the compartment is immediately sensed. This may or may not be where the ventilation air leaves the compartment.
(7) Fire detectors should be installed in designated fire zones, the combustor, turbine, and tailpipe sections of turbine installations.
Engine Power Section (Combustor, Turbine, and Tailpipe): This zone is usually characterized by predictable hazard areas which facilitate proper detector location. It is recommended that coverage be provided for any ventilating air outlet as well as intermediate stations where leaking combustibles may be expected.
Compressor Compartment: This is usually a zone of relatively low air flow velocities, but wide geographical possibility for fires. When fire detectors other than radiation-sensing detectors are used, detection at air outlets provides the best protection, and intermediate detector locations are of value only when specific hazards are anticipated.
(i¡i) Accessory Bullet Nose: Where such a compartment is so equipped that it is a possible fire zone, its narrow confines permit sufficient coverage with one or more detectors at the outlets.
(iv) Heater Detector Location: An overheat detector should be placed in the hot air duct downstream of the heater. If the heater fuel system or exhaust system configuration is such that it is a fire hazard, the compartment surrounding the heater should also be examined as a possible fire zone.
(v) Auxiliary Power Unit Detector Location: The use of a combustion-driven auxiliary power unit creates another set of typical engine compartments defined and treated as above. Some units are so shrouded with fireproof material that these compartments exist only within the confines of the shroud. They are still, however, fire zones and should have a detection system.
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CHAPTER 2. PART 27
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
SUBPART F - EQUIPMENT
EQUIPMENT - GENERAL AC 27.1301. § 27.1301 FUNCTION AND INSTALLATION.
a. Explanation. It should be emphasized that this rule applies to each item of installed equipment including optional as well as required equipment.
b. Procedures.
(1) Information regarding installation limitations and proper functioning is normally available from the equipment manufacturers in their installation and operations manuals. In addition, some other paragraphs in this AC include criteria for evaluating proper functioning of particular systems. (An example is paragraph AC 27 MG 1 for avionic equipment.)
(2) This general rule is quite specific in that it applies to each item of installed equipment. It should be emphasized, however, that even though a general rule is relevant, a rule that gives specific functional requirements for a particular system will prevail over a general rule. Therefore, if a rule exists that defines specific system functioning requirements, its provisions should be used to evaluate the acceptability of the installed system and not the provisions of this general rule. It should also be understood that an interpretation of a general rule should not be used to lessen or increase the requirements of a specific rule. Section 27.1309 is another example of a general rule, and this discussion is appropriate when applying its provisions.
(3) For optional equipment, the emphasis on functioning is rather limited compared to that for required equipment. The conditions under which the optional equipment is evaluated should be recorded in the type inspection report. The major emphasis for this type of equipment should be to ensure it does not interfere with the operation of systems that are required for safe operation of the rotorcraft, and that the failure modes are acceptable and do not create any hazards.
AC 27.1303. § 27.1303 FLIGHT AND NAVIGATION INSTRUMENTS.
NOTE: The EFIS guidance material provided in this section is not directly related to a showing of compliance with § 27.1303. The material will be moved to a new AC 27 EFIS MG section at the next revision.
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