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.
(il) 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.
(il) 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:
Radiation-sensing detectors.
(ii) Rate-of-temperature-rise detectors.
(IlI) Overheat detectors.
(iv) Smoke detectors.
(v) CO detectors.
(vi) Combustible mixture detectors.
(vil) 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 776 Par 592
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" 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.
(ill) 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.
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.
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АС 27-1A 7/30/97 (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 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.
(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 tuel system or exnaust system 778 Par 592
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.
593.-616. _ RESERVED.
Par 592 779 (thru 803)
SECTION 34. EQUIPMENT - GENERAL
617. § 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 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 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.
618. § 27.1303 FLIGHT AND NAVIGATION INSTRUMENTS.
This rule lists the flight and navigation instruments that are required for VFR operation.
Additional rules to be consulted when determining the flight and navigation instrument installation design are § 27.1321, arrangement and visibility, and Part 27, Appendix B, Paragraphs VIll(a) and (b), for IFR operation considerations. 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.
804 Par 617
619. § 27.1305 (Amendment 27-9) POWERPLANT INSTRUMENTS.
a. Explanation. This section specifies the instruments which are required for reciprocating and turbine engine installations. These instruments will provide the pilot with essential data to determine operational status of critical components and select desired performance conditions.
b. Procedures.
(1) FAR 27.1305(f), (g), and (I) requires a warning device for transmission oil temperature, gearbox oil pressure, and low fuel. 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.
(2) 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 tault or malfunction. Redundant design is generally applied to compensate for the loss of independence.
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.
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.
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.
Par 619 805
Specific issues that must be addressed to assure compliance with the minimum safety standards are as follows:
1. The level of most severe hazard must be determined.
2. Equivalent reliability and software design assurance to the determined criticality level must be shown.
(3) Additional sections to be consulted when determining the powerplant instrument installation design are §§ 27.1321, Arrangement and visibility; 27.1337, Powerplant instruments; 27.1381, Instrument lights; 27.1543, Instrument markings:
General; 27.1549, Powerplant instruments; 27.1551, Oil quantity indicator; and
27.1553, Fuel quantity indicator.
619A. § 27.1305 (Amendment 27-23) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 27-23 revises, edits, and adds new powerplant instrument requirements. Section 27.1305(l) 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.
Section 27.1305(q) 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 § 27.955. Section 27.1305(s) was revised to require function indicators only for fuel heaters that can be selected or are controllable.
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. 27.1305 (Amendment 27-29) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 27-29 added Sections 27.1305(t) and 27.1305(u) to provide for 30-second/2-minute OEl power ratings.
(1) Section 27.1305(t) adds the requirement that a device or means be provided to alert the crew of the use of the 30-second and 2-minute OEI 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 OEI power levels is at the crew's discretion, unlike the other limits (l.e., torque, measured gas temperature, and gas generator speed) for 30-second OEI 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 806 Par 619
from monitoring the engine instruments during critical phases of flight caused by the loss of an engine.
(2) Section 27.1305(u) 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 OEl 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.
b. Procedures. For the purpose of complying with FAR 27.1305(t) and
27.1305(u), the 2-minute OEI power level is considered to be achieved whenever one or more of the operating limitations applicable to the next lower OE power rating are exceeded. The 30-second OEl power level is considered to be achieved whenever one or more of the operating limitations applicable to the 2-minute OEI power rating are exceeded.
(1) A review of the method to meet the requirements of Section 27.1305(t) should be conducted by the FAA/AUTHORITY Flight Test pilot. A determination should be made as to whether the method used to alert the crew of 30-second or 2-minute OEI power usage can be recognized and understood by the crew.
(2) To meet the requirements of Section 27.1305(u), a device should be installed on the engine or the airframe to record the time and each usage of 30-second and 2-minute OEI power levels. The information on the time and usage of 30-second and 2-minute OEI power should be recoverable from the recording device by ground personnel. The device should not be capable of being reset in flight and should only be capable of being reset by ground personnel. Prior to each flight this device should be capable of being checked for proper operation and to determine if 30-second or 2-minute OEI power levels were used during the previous flight.
620. § 27.1307 MISCELLANEOUS EQUIPMENT.
a. Explanation. This rule provides a listing of several items of required miscellaneous equipment. Each item seems to be self-explanatory except the one requiring a master switch arrangement. The purpose of a master switch arrangement is to allow rapid removal of sources of electrical power from the rotorcraft in an emergency situation.
b. Procedures. When reviewing possible solutions to the master switch arrangement requirement, the following considerations should be included.
(1) System separation. Since wiring from each electrical system will be brought in close proximity to each other, extra care should be taken to maintain some separation. As examples, common connectors, common grounds, and common wire routing should be avoided.
Par 619B 807
(2) Installation of switches. A single switch should be avoided since it introduces the possibility of a single failure turning off the entire electrical system. One solution that is commonly used provides a close grouping of the switches such that the pilot can easily reach all switches and turn them all off with one action. This solution requires a cockpit evaluation to ensure the installation will be suitable for different hand sizes. Another solution involves a gang bar that can be moved with a single motion to turn off all sources. This solution has been found to be acceptable in several instances.
Other solutions should be evaluated on their own merits, and the primary emphasis should be on maintaining some minimum system separation and conducting a cockpit evaluation by flight test personnel.
621. § 27.1309 (Amendment 27-21) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Types of Equipment. This regulation covers, but is not limited to electrical, pneumatic, and hydraulic power sources, associated distribution, and corresponding utilization systems.
b. Environmental Qualification.
(1) Laboratory Tests.
Environmental Standards. In order to ensure that the components/systems under consideration will tunction properly when exposed to adverse environments, they should be tested in the laboratory under a simulated adverse environment. If a TSO exists and it is appropriate in environmental range and performance for an equipment installation, it is preferable that the equipment be TSO approved. If there is no applicable TSO or an existing TSO does not provide for a sufficiently adverse environment, the latest revision of Radio Technical Commission for Aeronautics (RTCA) Document DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.
(ii) Adverse environmental variables for all types of required and critical equipment include, but are not necessarily limited to, temperature, humidity, vibration, shock, altitude, overpressure, and power source transients.
(iïi) For electrical/electronic equipment, adverse environmental variables include all of (b)(1)(i) and (ii) above plus overvoltage and undervoltage. Electronic equipment should also be tested for electromagnetic interference (EMI). These tests should include both emission and susceptibility evaluations of both conducted and radiated EMI.
(iv) Explosion Tests. Those items of electrical/electronic equipment that are to be located in areas subject to flammable fluids and vapors, as a result of any single probable malfunction or failure including failure of couplings or lines, should be 808 Par 620
tested as an ignition source. These tests consist of normal operation of the equipment in a physically contained explosive atmosphere. The explosion test procedure in the latest revision of DO-160 will satisfy this requirement. If another standard is used that is at least as good as the latest revision of DO-160, it may also be accepted to satisfy this requirement.
(2) Installed Environmental Tests. After the environmental ratings of the components/systems have been established, it should be ensured that as installed, these ratings will not be exceeded. Normally, installed equipment need not be instrumented and tested in flight, and it is not necessary to instrument the compartment or rack where the equipment is installed. Satisfactory environment and equipment compatibility are ensured by selection of the proper environmental category of laboratory tests. The category is determined by the type of aircraft (reciprocating or turbine) and flight envelope (altitude and temperature). Exceptions to normal installations are: (a) Alternator/generator cooling, where radiated and conducted heat is almost always uncertain, also cooling air temperatures and flow rates are uncertain;
(b) Where flight tests reveal excessive instrument panel vibration. In this case, the panel should be instrumented, tested, and, if necessary, design improvements made;
and (c) Any other cases where good engineering judgment and application of sound engineering principles indicate a high likelihood that the installed environment is more severe than the equipment is capable of operating within.
(1) Temperature Tests.
(A) Temperature tests may be accomplished by instrumenting the installed equipment environment with a recorder that provides a permanent record of time, altitude, and temperature. The pertinent temperature should be recorded as the rotorcraft is operated throughout its altitude range, including ground operation. The maximum and minimum temperatures recorded should be corrected degree for degree to ensure the equipment under test remains within its temperature rating while the rotorcraft operates throughout its approved ambient temperature envelope. (For generator/alternator cooling test procedures, refer to Paragraph 778 of this AC.)
Section 27.1043(b) requires the maximum approved operating OAT to be at least 100° F for powerplant-mounted accessories such as starter generators, vacuum pumps, etc. Due to the impracticality of the 100° F hot day temperature limit, rotorcraft systems mounted on the powerplant are normally evaluated for at least 115° F hot day sea level conditions with corresponding 3.6° F/1,000-foot correction. The maximum hot day OAT at sea level must be specified in the rotorcraft flight manual. Section 27.1043(b) is the regulatory basis for the lapse rate of 3.6° F/1,000 feet. This lapse rate should be applied regardless of the hot day sea level temperature the applicant chooses to certify for operation.
(B) The § 27.1043(b) maximum ambient temperature detinition should not be confused with operating temperatures in closed areas. Closed equipment rack areas can easily reach temperatures of 140° F when sitting on the ramp Par 621 809
in the southern United States in midsummer. Normally, proper selection of the altitude temperature category in the latest revision of DO-160 will ensure compliance.
(C) In some cases, the equipment manufacturer furnishes temperature limits for internal critical parts. (For example, brushes, bearings, or field windings on DC generators.) In these cases it is better to record the critical component temperature rather than equipment or equipment environment temperature.
(D) The following will illustrate an acceptable high temperature evaluation method:
Toat max = Maximum outside air temperature at which temperature tests are conducted.
Tmax = Maximum temperature to which the installed equipment has been tested in the laboratory.
Trest max = Maximum installed equipment temperature recorded during tests.
Torh = The high reference outside air temperature. It varies with altitude starting at the highest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000 foot altitude. It can be no less than 100° F (reference § 27.1043(b));
however, it can be as high as the applicant wants.
Th mar = Temperature margin between the maximum equipment temperature substantiated in the laboratory and the maximum installed equipment temperature when the rotorcraft is operating in the highest available OAT and approximately corrected at the altitude under consideration. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.
Th mar = Tmax - (Ttest max + (Torh - Toat max)) Example #1: Assume the applicant is seeking approval for rotorcraft operation at the lowest acceptable OAT, at sea level, of 100° F and Tmax for Generator Brush = 295° F at maximum load current throughout the altitude range. In-flight test data are:
Altitude (ft, MSL) Measured Brush Temp (Liest max- OAT = Toat max sea level 275° F 90° F 5,000 270° F 80° F 10,000 285° F 60° F 15,000 294° F 42° F 20,000 290° F 20° F 810 Par 621
First, Torh must be calculated for each altitude test point.
@ sea level, Torh = 100° F @ 5,000 ft., Torh = 100° F -5,000 ft., x 3.6°/1,000 ft. = 82° F @ 10,000 ft., Torh = 100° F - 10,000 ft. x 3.6°/1,000 ft. = 64° F @ 15,000 ft., Torh = 100° F - 15,000 ft. × 3.6°/1,000 ft. = 46° F @ 20,000 ft., Torh = 100° F - 20,000 ft. x 3.6°/1,000 ft. = 28° F Then at sea level:
Th mar = 295 - (275 + (100 - 90)) = 10° F At 5,000 feet:
Th mar = 295 - (270 + (82 - 80)) = 23° F At 10,000 feet:
Th mar = 295 - (285 +(64 - 60)) = 6° F At 15,000 feet:
Th mar = 295 - (294 + (46 - 42)) = -3° F At 20,000 feet:
Th mar = 295 - (290 + (28-20)) = -3° F Since Th mar comes out negative at the 15,000- and 20,000-foot points, the generator fails. It will be necessary for the applicant to reduce the maximum load current, improve cooling, or otherwise change the design to ensure the generator is operating within its approved temperature limit of 295° F.
(E) In most cases, the equipment is laboratory tested to minimum temperatures as severe as that of the rotorcraft's maximum certified altitude on a minimum temperature day. Therefore, unless equipment minimum temperature is affected by refrigeration or other temperature reducing environments, actual installed instrumented minimum temperature tests are unnecessary. If low temperature evaluation is necessary for the installed equipment, the following is an acceptable method:
Toat min = Minimum outside air temperature at which temperature tests are conducted.
Par 621 811