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 i n 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:
(i) 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.
(iii) 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:
(i) 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.
(viii) 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:
(i) 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.
(iii) 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:
(i) 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.
(iii) 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.
(vii) 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 ai r leaves the compartment.
(7) Fire detectors should be installed in designated fire zones, the combustor, turbine, and tailpipe sections of turbine installations.
(i) 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.
(ii) 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.
(iii) 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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AC 27-1B, Chg 1 2/12/03
a. Explanation. This rule lists the flight and navigation instruments that are required for VFR operation. Additional rules to be used when determining compliance for the flight and navigation instrument installation for rotorcraft are § 27.1321, Arrangement and Visibility, and Part 27, Appendix B, paragraphs VIII(a) and (b), for IFR operation considerations. Additional information regarding the different instruments can be found by referring to the Technical Standard Order (TSO) for each one. Compliance with the TSO requirements does not ensure compliance with the appropriate Part 27 installation requirements. Other considerations may also be found by reviewing the requirements of §§ 27.1323, 27.1327, 27.1335, 27.1381, 27.1543, 27.1545, and
27.1547.
b. Procedures.
(1) The following instruments are the required minimum flight and navigation instruments for VFR:
(i) Airspeed indicator.
(ii) Altimeter.
(iii) Magnetic Direction Indicator.
If additional instruments are required per other considerations, then evaluation will take place to determine appropriateness.
(2) Refer to Part 27, Appendix B, paragraphs VIII(a) and (b), for IFR operation considerations.
(3) If a speed warning device is required to be included as part of the rotorcraft design, it must meet the performance requirements given in the rule. In addition, the evaluation of the acceptability of the aural warning (i.e., this warning differs distinctively from aural warnings used for other purposes) should be accomplished by flight test personnel as part of their overall evaluation.
(4) Electronic Flight Instrument Systems (EFIS).
(i) Explanation. The increased use of microprocessor technology in avionic systems has resulted in the use of computer-generated graphics to replace conventional electromechanical instruments, which are used for the display of flight information required by § 27.1303. For IFR certified aircraft, the EFIS usually is used for the display of the magnetic gyro-stabilized direction indicator (slaved compass system). These computer -generated graphics are usually displayed on small multicolor-shadow-mask cathode ray tubes (CRT) or liquid crystal displays (LCD), and replace the horizontal situation indicator (HSI) and the attitude direction indicator (ADI).
This section presumes that the EFIS for which approval is sought meet the general requirements of an EFIS for a technical standard order or a small category airplane with Page F - 2
regard to color, symbology, operation, and so forth. This paragraph along with some others in this document principally highlights the areas that are peculiar to the installation in a normal category rotorcraft. A discussion of the flight director function performed by the EFIS is given in paragraph AC 27.1335. A discussion of the location of the displays is contained in paragraph AC 27.1321. A discussion of the recommendations for an EFIS is contained in the latest versions of both AC 25-11, Transport Category Airplane Electronic Display Systems and AMJ 25-11, Electronic Display Systems, and refer to AC-23.1311-1A, Installation of Electronic Displays in Part 23 Small Airplanes, dated March 12, 1999.
(ii) Procedures.
(A) System Components. The system components require qualification testing to determine that their design is acceptable, free from hazards, and suitable to their airborne environment. The components of the EFIS should meet the requirements of TSO C-113, C2d, C3d, C4c, C5e, C6d, and C10b (other TSOs for EFIS may apply).
(1) Environmental Qualification. The EFIS hardware must be shown to be suitable to its airborne environment. A desirable way to qualify the system component is to obtain approval to the appropriate TSO. If the equipment is not TSO approved, it should be shown via testing that it complies with the requirements of SAE Document AS-8034, the latest revision of RTCA Document DO-160, and other appropriate standards. This may include testing in accordance with the appropriate categories of the latest revision of RTCA Document DO-160, JEDEC Publication No. 64D (Protection from Ionizing Radiation), and UL Document No. 1418 (Impact Implosion Test).
(2) Software. The embedded software should be qualified in accordance with the latest revision of RTCA Document DO-178B. The software level is contingent on the worst case criticality of the function it performs. The software should be designed to provide adequate consideration for this factor (reference paragraph AC 27.1309). A similar consideration is required for altitude and airspeed.
(B) System Installation Considerations.
(1) Human Factors. Humans are very adaptable, but they adapt at varying rates with varying degrees of effectiveness and mental processing compensation. Thus, what some pilots might find acceptable and approvable, others would reject as being unusable and unsafe. Rotorcraft EFIS displays must be effective when used by pilots who cover the entire spectrum of variability. Relying on a requirement of “train to proficiency” may be unforeseeable, economically impracticable, or unachievable by some pilots without excessive mental workload as compensation.
(i) In order to minimize the potential for changes to first time EFIS installation approvals, the test program for EFIS displays should include sufficient flight and simulation time, using a representative population of pilots, to substantiate:
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AC 27-1B, Chg 1 2/12/03
(a) Reasonable training times and learning curves; (however, realistic assumptions should be made concerning the actual time for an average pilot to learn/adapt to a system. Also, these assumptions should reflect the fact that Part 91 pilots might not be trained, since training is not typically required.);
(b) Usability in an operational environment;
(c) Acceptable interpretation error rates equivalent to or less than conventional displays;
(d) Proper integration with other equipment that uses electronic display functions;
(e) Acceptability of all failure modes including those failures or combinations of failures not shown to be extremely improbable; and
(f) Compatibility with other displays and controls.
The manufacturers should provide human factors support/rationale for their decisions regarding new or unique features in their EFIS display. Evaluation pilots should verify that the data supports a conclusion that any new or unique features have no human factors traps or pitfalls, such as display perceptual or interpretive problems, for a representative pilot population.
(ii) It is desirable to have display evaluations conducted by more than one pilot, even for certification of EFIS displays that do not incorporate significant new features. For display designs that incorporate unproven features, evaluation by a greater number of pilots should be considered. To help the FAA/AUTHORITY certification team gain assurance of a sufficiently broad exposure base, the electronic display manufacturer or installer should develop a test program with the certificating office that gathers data from FAA/AUTHORITY test pilots, company test pilots, and customer pilots who will use the display. A reasonable amount of time for the pilot to adapt to a display feature can be allowed, but long adaptation times must receive careful consideration. Any attitude display format presented for FAA/AUTHORITY approval should be sufficiently intuitive in its design so that no training is required for basic manual rotorcraft control.
(iii) For those electronic display systems that have been previously approved (including display formats) and are to be installed in rotorcraft in which these systems have not been previously approved, a routine FAA/AUTHORITY certification review should be conducted. This program should emphasize the systems’ integration in the rotorcraft, taking into account the operational aspects, which may require further detailed systems failure analysis (where “system” means display, driving electronics, sensors, and sources of information).
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(iv) Simulation is an invaluable tool for display evaluation. Acceptable simulation ranges from a rudimentary bench test set up, where the display elements are viewed statically, to full flight training simulation with motion, external visual scene, and entire rotorcraft systems representation. For minor or simple changes to previously approved displays, one of these levels of simulation may be deemed adequate for display evaluation. For evaluation of display elements that relate directly to rotorcraft control (i.e., air data, attitude, power parameters, etc.), simulation should not be relied upon entirely. The dynamics of rotorcraft motion, coupled with the many added distractions and sensory demands made upon the pilot that are attendant to actual helicopter flight, have a profound effect on the pilot’s perception and usability of displays. Display designers, as well as FAA/AUTHORITY test pilots, should be aware that display formats evaluated and found acceptable in simulation may well (and frequently do) turn out to be unacceptable in actual flight.
(v) Prior to defining the characteristics (color, symbology, etc.) or standards to be used on a specific display, a flight deck design philosophy should be established. Additionally, displays should be consistent across all systems in the flight deck. Documentation of the usage for display philosophy would help establish the working basis for determining compliance.
(2) Symbology and Function.
(i) When assessing the acceptability of the EFIS, consideration should be given to the effect of the loss of one of the CRT color guns. This type of failure is especially a factor in determining the acceptability of the installation for single-pilot operation.
(ii) Symbols should be distinctive to minimize misinterpretation or confusion with other symbols utilized in the displays. The type and function of symbology should be clearly defined and appropriately classified for pilot understanding.
Symbols representing the same functions on more than one display should utilize the same shape and/or color-coding.
(3) Display Chromaticity and Luminance. The chromaticity and luminance of the displays should be determined to be acceptable for all cockpit lighting conditions which are expected in service. An expanded discussion of these characteristics may be found in AC 25-11 and AC 23.1311.
(4) Temperature Survey to Determine Proper Cooling of EFIS Components.
(i) Equipment Requiring Cooling Test. As with any avionic equipment, good engineering judgment may deem that all components of the EFIS should have an in-flight temperature survey performed to ascertain that the thermal environmental tolerance of the system components is not exceeded. Usually, the following general Page F - 5
AC 27-1B, Chg 1 2/12/03 guidelines may be used to aid in determining when an in-flight temperature survey is warranted.
(a) Components that contain a CRT, LCD, etc., require the temperature survey outlined in paragraph AC 2 7.1303b(4)(ii)(B)(4)(ii).
(b) Equipment which does not contain a CRT, LCD, etc., but is specified by the manufacturer to require forced air cooling (by an airframe-mounted system), usually requires a temperature survey.
(c) Equipment which does not contain a CRT, LCD, etc., and is not specified as requiring forced air cooling may usually have its critical thermal environment substantiated by laboratory testing.
(ii) Temperature Survey Testing. The temperature tests for the EFIS units should consist of a short-term test of approximately 30 minutes which accounts for an aircraft which has heat-soaked on the ramp. A factor of 25° F should be added to the maximum corrected temperature to account for “greenhouse effect.” A long-term test should be accomplished at various altitudes and limiting (low and high) airspeeds.
All avionics equipment should be turned on during this test, and the cockpit panel lights should be operated at full intensity. The environmental control unit (ECU) or air conditioning system should not be operating during these tests; however, any windows or vents which are part of the “basic” TC rotorcraft may be utilized to ventilate the pilot’s stations. Both these tests should be corrected to the maximum temperature for which the rotorcraft is certified and a standard lapse rate for altitude as specified in this AC. If an airframe cooling system is necessary to keep the display units within acceptable temperature limits, then the pilot(s) must be made aware of a failure or malfunction of this cooling system. Some type of cockpit visual annunciation with the capability to perform a preflight test is usually utilized to fulfill this requirement.
(5) System Reliability.
(i) Failure of the EFIS to perform an intended function which results in the reversion to standby instruments or requires the use of abnormal procedures should be shown to be improbable.
(ii) For IFR operations, Appendix B of Part 27, paragraph VIII(b)(5)(iii), requires that the equipment, systems, and installations must be designed so that one display of the information essential to the safety of flight which is provided by the instruments will remain available to a pilot, without additional crewmember action, after any single failure or combination of failures that is not shown to be extremely improbable. The display of attitude, altitude, or airspeed is individually “essential to the safety of flight,” and, therefore, the loss of all attitude display, all airspeed information, or all altitude information to the pilot(s) should be extremely improbable. Also, any malfunction or failure of the EFIS which would result in the simultaneous incorrect display of this critical information should be extremely improbable. In view of the Page F - 6