AC 29.1305. § 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 fuel, 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.
(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 [Section AC 29.1305 continued on next page] Page F - 7
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:
(i) The level of most severe hazard must be determined.
(ii) Equivalent reliability and software design assurance to the determined criticality level must be shown.
(iii) 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.
AC 29.1305A. § 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.
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 Page F - 8
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 AC 29.1305.
AC 29.1305B. 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 OEI 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 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 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 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 dev ice 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 29.1305(a)(24) and
29.1305(a)(25), 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 § 29.1305(a)(24) should be conducted by flight test personnel. A determination should be made as to whether Page F - 9
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 29.1305(a)(25), 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.
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, 29.1301,
29.1309, 29.1305, and 29.1322
(ii) Standards - Latest revision of RTCA/DO178 and RTCA/DO-160; SAE documents
(iii) 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), Page F - 10
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.
(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.
Summary 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 onl y 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) . 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 29-2C, Chg 4 5/1/2014 AC 29.1305C. § 29.1305 (Amendment 29- 40) POWERPLANT INSTRUMENTS.
a. Explanation Procedures . Amendm ent 29-40 added section 29.1305(a)(6) to require an oil pressure indicator for each pressure-lubricated gearbox. Paragraphs (a)(6) through
(a)(25), prior to this amendment, have been redesignated as paragraphs (a)(7) through
(a)(26).
b. . I n addition to providing an oil pressure indicator for each pressure-lubricated gearbox, the guidance material of the previous AC 29.1305 paragraphs continues to apply.
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AC 29.1305D. § 29.1305 (AMENDMENT 29-59) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-59 revised § 29.1305 to include the use of technology such as a synthesized power indicator (SPI), and added § 29.1305(b)(4) to address One Engine Inoperative (OEI) Training Mode.
(1) Revised §§ 29.1305(a)(5), (a)(11), and (a)(12) to incorporate flexibility for the use of a synthesized power indicator (SPI).
(2) Added § 29.1305(b)(4) to provide a requirement for when an applicant requests approval for OEI Training Mode. OEI Training Mode is intended to simulate an engine failure, provide an annunciation of that simulation, and indicate the actual power being provided by the engines to the flight crew.
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 29.1305 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 §§ 29.1305 and 29.1549, the SPI display must also meet the requirements of § 29.1321.
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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-toengine data when engine limiting factors are different (Engine 1 Torque, Engine 2 Gas Temperature Indicator). See §§ 29.1321 and 29.1549.
(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 § 29.1305(a)(25).
(5) The minimum and maximum operating limits must be marked as required by § 29.1549.
(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 §§ 29.771(a), 29.1309, 29.1321, and 29.1549. The SPI design should also address all normal and abnormal operations including start/shutdown and OEI conditions per § 29.771(a).
(7) Since the SPI is a system comprised of instruments requirement by subchapter C of Title 14, its hazards are among those that must assessed under the requirements of § 29.1309. When conducting this assessment, to meet the requirements in § 29.1301, 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.
For guidance on § 29.1305(b)(4) OEI training mode, see Miscellaneous Guidance 22 (MG 22).
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AC 29-2C, Chg 4 5/1/2014 AC 29.1307. § 29.1307 (Amendment 29-1 2) MISCELLANEOUS EQUIPMENT.
a. Explanation. Thi s rule provides a listing of several items of required miscellaneous equipment. Each item is self-explanatory, except for the one requiring a master switch arrangement for electrical circuits other than ignition. The purpose of a master switch arrangement is to allow rapid removal of all bus loads from sources of electrical power in an emergency situation. Requirements for radio communications are discussed more in AC 29.1431.
b. Procedures. When reviewing possible solutions to the master switch arrangement requirement, the following considerations should be included.
(1) Installation o f switches 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) . The 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 29.1309. § 29.1309 (Amendment 29-40) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Explanation.
(1) Applicability. Section 29.1309 of the Title 14 CFR regulations is intended as a general requirement that is applicable to any equipment or system as installed, in addition to specific systems requirements, except as indicated below.
(i) General. If a specific part 29 requirement exists which predefines systems safety aspects (e.g., redundancy level or criticality) for a specific type of equipment, system, or installation, then the specific part 29 requirement will take precedence (see section 29.1301b.(2) of this AC). This precedence does not preclude accomplishment of a system safety assessment, if necessary. For example, § 29.695 is a rule that predefines a required level of redundancy and an implied system reliability. However, a system safety assessment approach may still be required to show that the implied system reliability is met and to address assessment of the failure modes.
(ii) . Section 29.1309 does not apply to Subparts B, C, and D for aspects such as the performance, flight characteristics, structural loads Section 29, Subparts B, C, and D , and structural strength requirements. However, it does apply to any equipment and system on which compliance with the requirements of Subparts B, C, and D is based (e.g., Health Usage Monitoring System certified for maintenance credit and Stability Augmentation System).
(iii) Section 29, Subpart E.
(A) Section 29.1309 does not apply to the uninstalled type-certificated engine. However, it does apply to the equipment and systems associated with the engine installation (e.g., electrical power generation, engine displays, transducers, etc.)
on the rotorcraft (reference § 29.901).
(B) Section 29.1309 does not apply to the Rotor Drive Systems (reference § 29.917(b)).
(iv) Section 29, Subpart F.
(A) Section 29.1309 does not apply to stowed safety equipment such as life rafts, life preservers, and emergency floatation equipment. It also does not apply to safety belts, rotorcraft seats, and hand held fire extinguishers. However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and inadvertent deployment of emergency floatation equipment) approved as part of the type design.
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AC 29-2C, Chg 4 5/1/2014 (B) Section 29.1309 does not apply to the functional aspects of aircraft non-safety related equipment such as entertainment systems, hoists, Forward Looking Infrared systems (FLIR), or emergency medical equipment such as defibrillators, etc.
However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and failure of the electrical system fault protection scheme) approved as part of the type design.
(C) Section 29.1309 does not apply to the lighting characteristics (e.g., light intensity, color, and coverage) of the position lights, anti-collision lights, and riding lights. However, it does apply to hazards to the rotorcraft, its occupants, and flight crew introduced by the installation or presence of this type of equipment or system (e.g., Electromagnetic-Interference considerations, fire hazards, and pilot visibility impairment due to glare) approved as part of the type design.
(2) Background.
(i) This guidance is not intended to change the existing reliability requirements of the current §§ 29.1309(b) and (c). Neither is this guidance intended to impose an overly rigorous safety assessment process for those installations involving equipment or systems that are not considered to be complex or do not perform functions whose failure or malfunction result in higher level failure condition categories.
(ii) However, this section does provide guidance that describes a method of showing compliance that utilizes the latest system safety assessment techniques available in the aerospace industry. The safety assessment process described in this guidance can be easily adopted and applied with various degrees to equipment or systems, and installations ranging from simple systems that perform functions whose failure or malfunction have no safety impact, to the most critical and most complex systems.
b. Procedures.
(1) Definitions of Probability Classifications.
(i) FREQUENT. (This is the lower part of the range 10-5 or greater previously applied to the term “PROBABLE”.) Frequent events may be expected to occur often during the operational life of each rotorcraft that are based on a probability on the order of 10-3 or greater.
(ii) REASONABLY PROBABLE. (This is the upper part of the range 10 or greater previously applied to the term “PROBABLE”.) Reasonably probable even -5 -3 to 10 .
ts may be expected to occur several times during the operational life of each rotorcraft that are based on a probability on the order of between 10 -5 Page F - 14
(iii) REMOTE. (The term “REMOTE” is not related to the structural use of the term.) (This is the lower part of the range 10 -9 to 10-5 previously applied to the term “IMPROBABLE”.) Remote events are expected to occur a few times during the total operational life of a random single rotorcraft of a particular type, but may occur several times during the total operational life of a number of rotorcraft of a particular type, that are based on a probability on the order of between 10-5 to 10-7 .
(iv) EXTREMELY REMOTE -9 -5 -7 to 10 .
(v) EX
TREMELY IMPROBABLE -9 or less is assigned to this classification.
(2) Definitions of Failure Condition Classifications
NO EFFECT MINOR . (T he term “REMOTE” is not related to the structural use of the term.) (This is the upper part of the range 10 to 10 previously applied to the term “IMPROBABLE”.) Extremely remote events are not expected to occur during the total operational life of a random single rotorcraft of a particular type, but may occur a few times during the total operational life of all rotorcraft of a particular type, that are based on a probability on the order of between 10 -9 . ( Remains the same for both the three or the five classifications.) Extremely improbable events are so unlikely that they need not be considered to ever occur, unless engineering judgment would require their consideration. A probability on the order of 10 Note: The five probability terms defined in paragraph b.(1) above are intended to relate to an identified failure condition resulting from or contributed to by the improper operation or loss of a function or functions. These terms do not define the reliability of specific components or systems.
. Failure Conditions may be classified according to the severity of their effects into one of the following categories (reference Figure AC 29.1309-2).
(i) . Failure Conditions that would have no effect on safety; for example, Failure Conditions that would not affect the operational capability of the rotorcraft or increase crew workload, however, could result in an inconvenience to the occupants, excluding the flight crew.
(ii) . Failure conditions which would not significantly reduce rotorcraft safety, and which would involve crew actions that are well within their capabilities.
Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as, routine flight plan changes, or some physical discomfort to occupants.
(iii) MAJOR. Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew work load or in conditions impairing crew efficiency, physical distress to occupants, possibly including injuries, or physical discomfort to the flight crew.
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