relatively new technology embodied in the EFIS, the conventional technology electromechanical standby attitude indicator, with its independent power supply, should be retained.
c. Standby Instruments. The EFIS which have been approved on normal 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 of Part 27. This requirement is usually satisfied by an electromechanical panel -mounted gyro with an independent power source. 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 AC 27.1309 for a discussion of software qualification).
AC 27.1305. § 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 (l) 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 fault or malfunction. Redundant design is generally applied to compensate for the loss of independence.
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(i) 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.
(ii) 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.
(iii) 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.
(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.
AC 27.1305A. § 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.
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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 AC 27.1305.
AC 27.1305B. §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 OEI 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 (i.e., torque, measured gas temperature, and gas generator speed) for 30-second OEI that are set by an automatic control required by § 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 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 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.
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 OEI power rating are exceeded. The 30- second OEI 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.
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c. Integrated Display Systems. This advisory material is to provide guidance for compliance to Part 27 and Part 29 regulations as they apply to integrated display systems. The integration aspects of these systems require some additional issues to be addressed during certification. The term “must” in this advisory material is used in the sense of ensuring the applicability of these particular methods of compliance when the acceptable means of compliance described herein is used. This advisory material establishes an acceptable means, but not the only means of certifying an integrated display system.
(1) Definitions.
(i) Integrity. The term “integrity” for the purpose of this advisory material includes the hardware quality requirements, including reliability; as well as the software level requirements, as defined in DO178B.
(ii) Criticality. The term “criticality” refers to the five levels of criticality addressed in FAA Advisory Circulars AC 27-1B and AC 29-2C.
(2) Related documents.
(i) Federal Aviation Regulations (FARs) paragraphs 21.21, 27.1301,
27.1309, 27.1305, and 27.1322
(ii) Standards - Latest revision of RTCA/DO178 and RTCA/DO-160;
(iii) SAE documents ARP4754 and ARP4761
(3) Background. A tendency to integrate functions/indications that have previously been independent is a result of technology advancement. Microprocessor driven systems have facilitated the ease of this integration. Integrated Instrument Display System (IIDS) or Engine Instrument and Caution Advisory System (EICAS) are examples of integrated display systems. IIDS, EICAS, or any other similar systems are defined as a combination of engine instruments ( previously independent indicators), fuel quantity indication, and caution/warning parameters, as a minimum, presented by a common display driven by a common processor.
(4) Discussion. This design philosophy does not result in the traditional requirement for individual display independence for failure/malfunction considerations.
This loss of independence means that a single failure could result in loss of most, if not all, instrument displays on the integrated display system. Redundancy of the integrated display system is often proposed to compensate for this lack of independence.
However, redundancy alone may not meet the integrity requirements since they are derived from the level of criticality associated with the loss or malfunction of instrument/parameter displays for flight operations that are dependent on these indications.
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(5) Certification Approach. A two step procedure should be used to determine the adequate safety level for this type of system. The first step is to determine the level of criticality associated with the total loss/malfunction of these functions/indications or loss/malfunction of the critical parts of the display. This can be achieved through the use of a functional hazard assessment (FHA). This criticality assessment must be a product of failure/malfunction of the indication system and the flight operation that would represent the worst case for loss of this information. The second step is to determine that the design integrity of the system is at least equal to the assessed criticality level determined in step one.
(6) Functional Hazard Assessment. The operational classifications to be considered when assessing the criticality are Cat A, Cat B, and IFR. The need for critical information varies with each of these different operational categories. An example would be the demand for OEI parameter information in the single engine Cat A operation. Another example is the loss of fuel quantity indication and fuel low level indication simultaneously in IFR flight conditions. The FHA should address not only loss of one type of indication, but combined loss of engine parameter indication, including total loss of display information, caution/warning, fuel quantity indication, and any other included display in combination with a particular flight operation. There are techniques to lessen the consequences of the failure/malfunction requirements for integrity, such as providing back-up displays for the information deemed critical for a particular operational consideration.
(7) Summary. The loss of all integrated display information for certain types of flight operations may have the highest level of criticality associated with it. The same may be true for malfunctions that result in misleading indications. These failures/malfunctions must be addressed by the commensurate design integrity level.
Lesser levels of criticality must also be addressed by the appropriate design integrity levels.
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AC 27-1B, Chg 9 AC 27.1305C. § 27.1305 (AMENDMENT 27-51) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 27-51 revised § 27.1305(e), (k), (n), and (o) to incorporate flexibility for the use of a synthesized power indicator (SPI).
b. Procedures. The purpose of the SPI display is to reduce pilot workload by providing an indicator showing a single parameter that is closest to the applicable engine or transmission limit, versus monitoring individual engine parameters. The SPI display for reciprocating engines presents a normalized indication of the current engine manifold pressure and engine RPM. For turbine engines, the SPI display presents normalized indications of the engine power or torque, gas temperature, and engine compressor speed output. The normalized value is presented to the flight crew on the primary flight displays, along with a caption indicating the nearest engine limiting parameter (engine manifold pressure, RPM, torque, gas temperature indicator, engine compressor speed, as applicable) that is being used for the SPI display calculation. The SPI display function is typically performed by the avionics suite using data provided by the engine FADEC.
Section 27.1305 establishes the required powerplant instruments and warning and caution annunciations that must be displayed within the view of the flying pilot. These have traditionally been independent instruments and are needed not only to measure the performance output of the engines, but they collectively allow the flying pilot to continuously monitor the condition and health of the engines.
The absence of continuously displayed engine manifold pressure and RPM for reciprocating engines or engine torque, gas temperature, and engine compressor speed parameters may affect the pilot’s ability to directly monitor the health and condition of each engine. In addition, the absence of individual gauges deprives the pilot of the ability to quickly assess trend or rate of change information and proximity to individual parameter limits, and to accurately compare engine-to-engine data. As a result, in addition to the requirements in §§ 27.1305 and 27.1549, the SPI display must also meet the requirements of § 27.1321.
The applicant’s description and justification of the SPI display should address the following:
(1) The accuracy of the engine power and other engine indications to safely conduct Category B and, when applicable, Category A operations (power setting, NR control, piloting a given rating with precision).
(2) The flight crew must be able to assess how close the indicated parameter is relative to the limit, including the transitions from one limiting factor to another (engine manifold pressure to engine RPM or engine torque to gas temperature indicator to engine compressor speed), and the ability to quickly and accurately compare engine-to- Page F - 11.1
AC 27-1B, Chg 9 engine data when engine limiting factors are different (Engine 1 Torque, Engine 2 Gas Temperature Indicator). See §§ 27.1321 and 27.1549.
(3) Engine design features or characteristics that forewarn the crew prior to the parameter reaching the operating limit (the redline) should be evaluated.
(4) The SPI display must provide a timing countdown when operating with time limited power settings and as required by § 27.1305(t).
(5) The minimum and maximum operating limits must be marked as required by § 27.1549.
(6) The SPI display should allow the pilot to identify indicator failures, engine failures, partial power loss conditions, and adverse trends, to meet the requirements in §§ 27.771(a), 27.1309, 27.1321, and 27.1549. The SPI design should also address all normal and abnormal operations including start/shutdown and OEI conditions per § 27.771(a).
(7) Since the SPI is a system comprised of instruments required by subchapter C of Title 14, its hazards are among those that must be assessed under the requirements of § 27.1309. When conducting this assessment, to meet the requirements in § 27.1301, the applicant must take into account that the crew will operate the Rotorcraft, monitoring the engine status and health, without continuously accessing the powerplant display page. In addition, the assessment should assume the SPI will remain the only engine instrument available in case of loss of one display.
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AC 27-1B 9/ 30/99
AC 27.1307 § 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.
(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.
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AC 27.1309. § 27.1309 (Amendment 27-21) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Types of Equipment. Section 27.1309 of the Title 14 CFR regulations covers, but is not limited to, electrical, pneumatic, and hydraulic power sources, distribution, and corresponding utilization systems.
b. Environmental Qualification.
(1) Laboratory Tests.
(i) Environmental Standards. In order to ensure that the components and systems under consideration will function 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.
(iii) For electrical and 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 and 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 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 and 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 Page F - 13
AC 27-1B, Chg 4 3/31/2014 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 or 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.
(i) Temperature Tests.
(A) Temperature Survey to Determine Proper Cooling of Electrical and Electronic Equipment.
(1) Equipment Requiring Cooling Test. As with any avionic equipment, good engineering judgment may deem that all components of the electrical or electronic equipment 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.
(i) Components which contain a CRT (e.g., EDS) require the temperature survey testing outlined in paragraph b.(2)(i)(B) below.
(ii) Equipment that does not contain a CRT, but is specified by the manufacturer to require forced air cooling (by an airframe mounted system), usually requires temperature survey testing.
(iii) Equipment that 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.
(B) Temperature Survey Testing. The temperature tests for the electrical or electronic equipment units should include a short-term test of approximately 30 minutes, which accounts for an aircraft that 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 that are part of the “basic” TC rotorcraft may be utilized to ventilate the pilot’s stations. Both of 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 guidance. If an airframe cooling system is necessary to display units within acceptable temperature limits, then the pilot(s) must be made aware of a failure Page F - 14
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.
(C) 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 or alternator cooling test procedures, refer to section 27.1351 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.
(D) The § 27.1043(b) maximum ambient temperature definition 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 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.
(E) 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.
(F) 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.
Ttest 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 Page F - 15
AC 27-1B, Chg 4 3/31/2014 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 seekin g 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:
Measured Altitude (ft, MSL) Cylinder Temp (TTEST MIN) 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 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. x 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 Page F - 16