595. § 29.1201 FIRE EXTINGUISHING SYSTEM MATERIALS.
a. Explanation.
(1) Many different fire extinguishing agents are available for use in fire extinguishing systems. The choice of extinguishing agent should take into account the chemical reaction (if any) between the extinguishing agent and the materials utilized in the extinguishing system. If there are any incompatibilities, they should not create a hazard by creating volatile or toxic vapors or fumes which could feed a fire or cause injury to passengers, crew, or other personnel.
(2) The fire extinguishing components in an engine compartment must be fireproof to ensure operation in the event of a compartment fire.
b. Procedures.
(1) Compliance with the requirements of § 29.1201(a) can be demonstrated by analysis, test, or a combination of both.
(2) Certification data submitted by the applicant should contain a listing of the chemical ingredients of the extinguishing agent and the other materials in the extinguishing system. These data should also show that the chemical reaction (if any) of these materials, when combined, does not create a hazard.
(3) Where chemical compounds exist and the chemical reaction is not predictable when two different compounds are combined, actual tests may be necessary to determine the hazard potential.
(4) Analysis, test, or a combination of both may be used to demonstrate compliance with the fireproof requirement for all fire extinguishing components located within the engine compartment.
596. § 29.1203 (Amendment 29-40) FIRE DETECTOR SYSTEMS.
a. Explanation.
(1) Fire detection systems are required in turbine engine powered rotorcraft, Category A reciprocating engine powered rotorcraft and each Category B reciprocating engine powered rotorcraft where the engine displacement is greater than 900 cubic inches.
(2) This section specifies material, installation, and some operational requirements for fire detectors to ensure prompt detection of fire in the fire zones and other designated areas.
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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:
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 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:
(1) Radiation-sensing detectors.
(il) Rate-of-temperature-rise detectors.
(III) Overheat detectors.
(iv) Smoke detectors.
(v) CO detectors.
(vi) Combustible mixture detectors.
(vii) Fibre-optic detectors.
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(vili) 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 rascal sensing, fate of rich senses ere within a visible satice ovine artectors are detectors can be obtained in either "continuous" or "unit" type.
(4) The detector system should:
Indicate fire within 15 seconds after ignition, and show which engine compartment in which the fire is located.
(її) Remain on for the duration of the fire.
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.
A false fire detector indication could significantly increase crew workload, impair crew efficiency, or reduce safety margins and so is classified as a major failure condition. In consequence, such false fire detector indication should be shown to be improbable based on a probability assessment and service experience of the fire detector system.
If the probability of the fire detection system experiencing a false indication cannot be shown to be improbable, a secondary means of determining the validity of the fire indication should be provided.
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(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.
(її) 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.
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 fire resistant.
(vi) Each detector system should actuate a light 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 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 radiationsensing 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. The ventilation requirements of § 29.1187(e) must also be taken into consideration. Compliance with these recommendations allows the effective placement of a minimum amount of detectors, and still ensures 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 must be installed in designated fire zones, the combustor, turbine and tailpipe sections of turbine installations.
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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 outiets provides the best protection, and intermediate detector locations are of value only when specific hazards are anticipated.
(ill) 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 must have a detection system.
597.-616. RESERVED.
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SECTION 34. EQUIPMENT - GENERAL
617. § 29.1301 FUNCTION AND INSTALLATION.
a. Explanation. It should be emphasized that this rule applies to each item of installed equipment which includes optional equipment 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 776 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 29.1309 is another example of a general rule, and this discussion is appropriate when applying its provisions.
618. § 29.1303 (Amendment 29-24) FLIGHT AND NAVIGATION INSTRUMENTS.
a. Explanation. This rule lists the flight and navigation instruments that are required for VFR certification. Several additional rules to be consulted when determining the flight and navigation instrument installation design are § 29.1321, arrangement and visibility, § 29.1331, flight instrument power supplies, and § 29.1333, systems that operate the flight instruments at each pilot station. 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 29 requirements; however, satisfactory installation normally results. Other considerations may also be found by reviewing the requirements of 99 29.1323, 29.1327, 29.1335, 29.1381, 29.1543, 29.1545, 29.1541, and Part 29, Appendix B. Paragraphs VIII(a) and (b) of Appendix B include IFF operation considerations for flight and navigation instruments. In addition, if the maximum allowable airspeed is dependent on conditions such as weight and altitude, that information is normally provided on tables, graphs, placards, or other means in the cockpit and in the rotorcraft flight manual.
b. Procedures.
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(1) The following instruments are considered to be flight instruments:
(i) Airspeed indicator.
(ii) Sensitive altimeter.
(iii) Free-air temperature indicator.
(iv) Nontumbling gyroscopic bank and pitch indicator.
(V) Gyroscopic rate-of-turn indicator with an integral slip-skid indicator.
(vi) Rate-of-climb (vertical speed) indicator.
(2) The remaining instruments are navigation instruments.
(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 cockpit evaluation.
(4) Electronic Flight Instrument Systems (EFS).
(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 § 29.1303(1), (g), and (h). 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) and replace the horizontal situation indicator (HSI) and the attitude direction indicator (ADI). This discussion presumes that the EFIS for which approval is sought meets the general requirements of an EFIS for a transport category airplane with regard to color, symbology, operation, and so forth.
This paragraph along with some others in this document principally highlights the areas which are peculiar to the installation in a transport category rotorcraft. A discussion of the flight director function performed by the EFIS is given in Paragraph 640. A discussion of the location of the displays is contained in Paragraph 632. A discussion of the requirements for an EFIS in a transport category airplane is contained in AC 25-11, Transport Category Airplane Electronic Display Systems, dated September 16, 1987.
(i¡) Procedures.
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(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. Generally the components of the EFIS should meet the requirements of TSO C-113.
(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. This will include testing in accordance with the appropriate categories of the latest revision of RTCA Document DO-160, JEDEC Publication No. 64D (Protection from lonizing Radiation), and UL Document No. 1418 (Impact Implosion Test).
(2) Software. The embedded software should be qualified to an appropriate standard. The software level is contingent on the worst case criticality of the function it performs. As an example, the display of incorrect roll and/or pitch by an EFIS ADI instrument is a hazardous or "critical" malfunction. The software should be designed to provide adequate consideration for this factor (reference Paragraph 621).
A similar consideration is required for altitude and airspeed.
(B) System Installation Consideration.
(1) Display Chromaticity and Luminance. The chromaticity and luminance of the displays should be determined to be acceptable for all critical cockpit lighting conditions which are expected in service. An expanded discussion of these characteristics may be found in AC 25-11.
(2) 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 guidelines may be used to aid in determining when an in-flight temperature survey is warranted.
(a) Components which contain a CRT require the temperature survey outlined in 618b(4)(ii)(B)(1) (i).
(b) Equipment which does not contain a CRT, but is specified by the manufacturer to require forced air cooling (by an airframe mounted system), usually requires a temperature survey.
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(c) Equipment which does not contain a CRT and is not specified as requiring forced air cooling may usually have its critical thermal environment substantiated by laboratory testing.
(…) 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 avionic 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 pilots) 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.
(3) System Reliability.
Failure of the EFIS to perform a required function which results in the reversion to standby instruments or requires the use of abnormal procedures should be shown to be improbable (for Category A rotorcraft, reference § 29.1309(b)).
(i1) For IFR operations, Appendix B of Part 29, paragraph VIlI(b)(5)(ii), 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 pilots) should be extremely improbable. Also, any malfunction or failure of the EFIS which would result in the display of simultaneous incorrect display of this critical information should be extremely improbable. In view of the relatively new technology embodied in the EFIS, the conventional technology electromechanical standby attitude indicator, with its independent power supply, should be retained.
(4) Symbology and Function. 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.
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(5) Circuit Protective Devices. All circuit protective devices for the EFIS and related, required interface units should be placed in the cockpit where they can be reset by one of the pilots without leaving the pilot seat.
c. Standby Instruments. The EFIS which have been approved on transport category rotorcraft at this time have only presented the critical function of attitude display. A specific requirement for a standby attitude instrument is contained in Appendix b or Part 29. his requirement is usually satistied by an electromechanical panel-mounted gyro with an independent power source. This type of installation is what was envisioned by the authors of Appendix B. Because of the mature technology of this type of standby attitude indicator, certain aspects of the EFIS installation have not been an area for concern. If, however, a total commitment of critical display functions is made to the "glass" technology, such that the standby attitude instrument requirement is satisfied by a software based CRT system, then several major areas of concern will be raised. Among these are the electromagnetic vulnerability of the system (protection from the effects of lightning and high energy radio frequency fields) and software. The certifications of EFIS with an electromechanical standby attitude indicator have not considered loss of function critical from a software aspect. (reference Paragraph 621 for a discussion of software qualification.)
619. § 29.1305 (Amendment 29-10) POWERPLANT INSTRUMENTS.
a. Explanation. This section specifies those instruments which are required for reciprocating and turbine engine installations. It also provides instrumentation requirements for operating rotorcraft in Category A or Category B. These instruments will provide the pilot with essential data to determine operational status of critical components and select desired performance conditions.
b. FAR 29.1305(a)(4), (a)(6), and (a)(9) requires a warning device for low tuel, gearbox oil pressure, and transmission oil temperature. An indicator/gage is not acceptable for use as a warning device since the indicator/gage is not a primary instrument and therefore is not actively monitored.
c. There are advanced display systems that take advantage of microprocessor power by integrating the processing of several parameters. These systems have to date been referred to as Engine Caution Advisory Systems (ECAS) or as Integrated Instrument Display Systems (IIDS) and possibly other variations of these names.
These systems typically integrate propulsion instruments, fuel quantity indication, and caution and warning system into a single display system. In traditional designs the powerplant instruments, fuel quantity display, and the caution and warning system are independent from each other. The integration of these systems/indicators eliminates their independence from one another and increases the probability of loss of more than one indicator/system as a result of a single fault or malfunction. Redundant design is generally applied to compensate for the loss of independence.
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(1) This integration and resultant mitigation of independence can result in an increased opportunity for common mode failures. Approval of the compensating features is elevated in importance as it is this aspect that allows the concept to be acceptable and subsequently certifiable. The loss of all displayed information or erroneous information should be considered for determination of worst case criticality.
With this determination of criticality, the design can be evaluated to see that it meets the minimum associated level of design assurance. Additionally, due to space limitations, some systems employ "page over" features that may have some difficulty displaying the required information when needed and human factors aspects must be considered.
(2) The instrument display system must be investigated and found to be acceptable under both normal and emergency conditions, must perform its intended function under foreseeable operating conditions, and must be designed to minimize the hazards in the event of probable malfunction or failure.
(3) It must be shown that there is appropriate redundancy to provide adequate compensation for the loss of independence in the system. If a multi-page system is employed, it must be shown that needed information is displayed when required.
Specific issues that must be addressed to assure compliance with the minimum safety standards are as follows:
The level of most severe hazard must be determined.
Equivalent reliability and software design assurance to the determined criticality level must be shown.
(ili) Pretest capability must be provided for the warning and caution system to preclude an associated latent failure.
(iv) Human Factors (reference §§ 29.1321 and 29.1322).
d. Additional rules to be consulted when determining the powerplant instrument installation design are §§ 29.1321, Arrangement and visibility; 29.1337, Powerplant instruments; 29.1381, Instrument lights; 29.1543, Instrument markings: General;
29.1549, Powerplant instruments; 29.1551, Oil quantity indicator; and 29.1553, Fuel quantity indicator.
619A. § 29.1305 (Amendment 29-26) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-26 revises, edits, and adds new powerplant instrument requirements. Section 29.1305(a)(4) was revised to require a low fuel warning device for each tank that can be used to feed an engine. The amendment allows a longer time between warning actuation and fuel exhaustion and requires the low fuel warning device to be independent of the normal fuel quantity indication system.
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7/30/97/ AC 29-2B Section 29.1305(a)(17) was changed to extend its application to all rotorcraft (not just those with turbine engines) and to require an indication to the crew of the degree of filter blockage as it relates to the fuel flow requirements in § 29.955.
Section 29.1305(a)(19) was revised to require function indicators only for fuel heaters that can be selected or are controllable. A new paragraph (a)(20) was added to § 29.1305 that combined identical requirements for fuel pressure indicators in Paragraphs (b)(2) and (c)(2) and modified the applicability of these requirements to only those fuel systems with devices or components that could adversely affect fuel pressure at the engine, if they fail. It also eliminates the requirement for fuel pressure indicators in fuel systems, such as suction or gravity feed systems, which do not incorporate pumps or filters. A new § 29.1305(a)(21) was added that requires a warning device to indicate to the flight crew the failure of any fuel pump that is required to supply adequate fuel flow to the engine according to § 29.955; such indication is not required for fuel pumps which are demonstrated to be only necessary for engine starting.
Section 29.1305(a)(22) adds a requirement for a warning or caution device to alert the flight crew when particles are detected by the chip detector required by § 29.1337 (d). A new § 29.1305(a)(23) added a requirement for powerplant instruments or warning devices for auxiliary power units installed in rotorcraft.
b. Procedures. The requirement and purpose for each instrument is self-explanatory in the amendment. Other sections that should be considered when designing powerplant instruments are listed in Paragraph 619.
619B. 29.1305 (Amendment 29-34) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-34 added §§ 29.1305(a)(24) and 29.1305(a)(25) to provide for 30-second/2-minute OEl power ratings.
(1) Section 29.1305(a)(24) adds the requirement that a device or means be provided to alert the crew of the use of the 30-second and 2-minute OE power level.
The crew should be alerted when the 30-second or 2-minute interval begins and when the time interval ends. The amount of time spent at the 30-second or 2-minute OEl power levels is at the crew's discretion, unlike the other limits for 30-second OLl that are set by an automatic control required by Paragraph 29.1143. The purpose for providing the time interval alerts and automatically controlling the 30-second OEI limits is to free the crew from monitoring the engine instruments during critical phases of flight caused by the loss of an engine.
(2) Section 29.1305(a)(25) adds the requirement for a device to record the usage and the amount of time spent at the 30-second OEI power level and the amount of time spent at the 2-minute OEI power level. The information recorded by this device is for the use of the ground crew to determine if maintenance actions/inspections are to be conducted.
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