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.
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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.
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(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.
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Tmin = Minimum temperature to which the installed equipment has been tested in the laboratory.
Ttest min = Minimum installed equipment temperature recorded during tests.
Tor = The low reference outside air temperature. It varies with altitude starting at the lowest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000-foot altitude.
T1 mar = Temperature margin between the minimum equipment temperature substantiated in the laboratory and the minimum installed equipment temperature. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.
T1 mar = - (Tmin - (T test min + (Tort - Toat min))) NOTE: This equation assumes all temperatures are negative. It is necessary to place a (-) in front of the right side of the equation in order to convert the T1 mar value to the conventional positive answer for acceptance and a negative answer for rejection. Temperature in the zero to 32° F range can be handled by conversion to the centigrade scale.
Example #2: Assume the applicant is seeking a low temperature operating limit at sea level of -25° F. Assume the hydraulic control cylinder has been substantiated in the laboratory to operate at a cylinder temperature of -40° F.
The in-flight test data are:
Measured Altitude (ft. MSL) Cylinder Temp (Trest min) OAT = Toat max sea level -0° F -25° F 5,000 -9° F -45° F 10,000 -21° F -59° F 15,000 -32° F -65° F 20,000 -40° F -69° F (Tor must be calculated for each altitude test point) @ sea level, Tor = -25° F @ 5,000 ft., Tor = -25° F -5,000 ft., x 3.6°/1,000 ft. = -43° F @ 10,000 ft., Tor = -25° F - 10,000 ft. × 3.6°/1,000 ft. = -61° F @ 15,000 ft., Tor = -25° F - 15,000 ft. x 3.6°/1,000 ft. = -79° F* @ 20,000 ft., Tori = -25° F - 20,000 ft. x 3.6°/1,000 ft. = -97° F* 812 Par 621
7/30/97 • AC 27-1A According to § 27.1043(b), the lowest temperature to be considered is -69.7° F.
Then at sea level:
T1 mar = -(-40 -(0 + (-25 - (-25)))) = 40° F At 5,000 feet:
T1 mar = -(-40 - (-9 + (-43 - (-45)) = 33° F At 10,000 feet:
T1 mar = (40 - (-21 + (-61 - (-59))) = 17° F At 15,000 feet:
T1 mar = -(-40 - (-32 + (-69.7 - (-65)))) = 3.3° F At 20,000 feet:
T1 mar = (40 - (-40+ (-69.7 - (-69)))) =-0.7° F It can be seen that there is an acceptable margin at all altitudes up to and including 15,000 feet. However, at 20,000 feet, the margin is negative and the system fails.
(ii) Vibration tests. Normally, installed vibration tests are not necessary for equipment qualified in accordance with the latest revision of RTCA document DO-160. This paper categorizes vibration tests according to installed rotorcraft equipment location such as fuselage, engine compartment, instrument panel, equipment rack, etc. However, installed equipment vibration tests may be necessary when it appears the equipment location environment may exceed the laboratory-tested equipment vibration limits.
(il) Altitude tests. If the equipment has been laboratory tested to the maximum certified altitude of the rotorcraft, installed altitude tests are unnecessary.
The installed equipment must be either laboratory tested or tested in the rotorcraft to the maximum certified altitude of the rotorcraft.
(3) Lightning Protection.
(i) Background. During the original design and development of rotorcraft and the development of regulations concerning these aircraft, little attention was given to protection from the meteorological phenomenon of lightning. This was, in part, Par 621 813
because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design normal category rotorcraft are utilizing the same advanced technology systems and materials as airplanes. Because of this fact, a specific requirement has been added by Amendment 27-21 for the consideration of lightning strike protection of required systems, equipment, and installations. The addition of Paragraph (d) to § 27.1309 further defines the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.
(її) Procedures.
(A) Section 27.1309(d) requires when showing compliance to § 27.1309(a) and (b) that the effects of lightning strikes on the rotorcraft must be considered. The first step in demonstrating compliance is to perform a fault/failure analysis (F/FA) to identify those functions for which the loss of function or malfunction may result in a catastrophe to the rotorcraft. An F/FA should be conducted on each system whose failure to function properly would prevent the continued safe flight and landing of the rotorcraft. These systems should be designed and installed to ensure that they can perform their intended function during and after exposure to lightning.
Additionally, evaluation must be performed to identify each system whose failure to function properly would reduce the capability of the rotorcraft or the ability of the flight crew to cope with adverse operating conditions. These systems should be designed and installed to ensure that they can perform their intended function after exposure to lightning.
The lightning strike models to be used for system justification should be as described in SAE AE4L Committee Report AE4L-87-3, Revision B, dated January 1989 or later version. The recommended reference for performing such analysis is Society of Automotive Engineers, Aerospace Recommended Practice 926A.
(B) Detailed means of compliance should be agreed with the authorities taking into account the effects on the rotorcraft and minimum considerations are as follows:
(1) Any combination of analysis and testing should be agreed with the authority.
(2) For test results, an extrapolation of the threat current parameters of more than a factor of 10 is not likely to be acceptable without an additional safety factor being applied.
(3) For a proven analysis technique, a safety factor of at least 2 will be necessary.
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(C) Flight and engine controls are examples of "critical" functions and with these critical functions defined, an analysis and/or testing should be performed to show compliance with § 27.1309(a); i.e., equipment, systems, and installations performing those identified functions equipment, systems, and installations performing those identified functions should be designed and installed to ensure that they continue to perform their intended functions considering the conditions of the rotorcraft experiencing a worst-case lightning strike encounter. Section 27.610 contains some methods which may be utilized for less complex mechanical systems; however, a great deal of difficulty will be experienced in trying to use these criteria to demonstrate that a very complex avionic system complies with § 27.1309(a). A method as outlined in Paragraph 621b(4), "Lightning Strike Protection of Full Authority Digital Engine Controls,i is recommended. This method may be readily adapted to other avionic systems performing critical functions. Also, this identifies an acceptable quantification of the expected airborne environment. The next step involves expanding the F/FA to determine what probable malfunctions or failures of the installed equipment systems and installation may occur as a result of a lightning strike to the rotorcraft. For a twin-engine rotorcraft, these systems must be designed such that a failure or malfunction will not cause a hazard to the rotorcraft. For a single-engine rotorcraft, these equipment systems and installation should be designed to minimize hazardous effects of failures or malfunctions. Note that the analysis which determines what are probable malfunctions or failures should consider that the encounter with a worst-case lightning strike is a given event.
(4) Lightning Strike Protection of Full Authority Digital Engine Controls.
(i> Explanation.
(A) The following discussion is written specifically for full authority digital engine controls (FADEC) with an alternate technology backup fuel control installed on rotorcraft with Part 29 Category A engine isolation. The requirement for increased consideration of lightning strike encounter effects on avionic equipment and systems has been brought about by the increased use of avionics to perform functions, the failure or malfunction of which could result in a hazard to the rotorcraft. The susceptibility of current high technology avionic systems is increased by the use of large scale integration (LSI), very large scale integration (VLSI), and complementary metallic oxide silicon (CMOS) technologies which exhibit a greatly reduced tolerance to large amplitude, low energy electrical transients as compared to conventional bipolar technology, and the reduced physical protection and electromagnetic shielding afforded aircraft avionic systems by the advanced technology composite airframe materials.
Additionally, processor-based systems have the failure phenomenon of digital upset. A digital upset occurs when a system, perturbed by an electrical transient, ceases proper operation in accordance with its embedded software while suffering no apparent component or device damage.
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АС 27-1A 7/30/97 (B) Since elements of electrical/electronic engine subsystems are typically spread throughout much of the rotorcraft, transients caused by lightning are coupled into the subsystem interface cables and may damage the system or cause upset. Effective lightning protection must be designed and incorporated into these systems. Reliance upon redundancy as a means of protection against lightning effects is generally not adequate because lightning electromagnetic fields and structural IR voltages usually interact (to some extent) with all electrical wiring aboard a rotorcraft.
(C) The testing and analysis outlined in this discussion are methods by which the FAA/AUTHORITY may be assured that when the rotorcraft experiences "the foreseeable operating condition" of a worst-case lightning strike encounter that the electronically controlled engines will continue to "perform their intended function" and therefore be in compliance with § 27.1309 as installed.
(D) The definition of what constitutes a full authority engine control is not at this time clearly defined. However, it has been accepted in past certification that any control which relies upon the electronics for the function on which Civil Certification or Military Qualification is based (e.g., rotor speed governing) is a full authority control, regardless of the backup control mode provided. If engine certification or qualification can be achieved without the electronic control which is subsequently added to achieve improved operational efficiency in the aircraft, the control is "supervisory."
However, if the controls used in a multiengine rotorcraft have a common failure caused by a lightning strike which could result in simultaneous failures which would cause a reduction in power greater than the loss of one engine, this would also be considered "full authority."
NOTE: If OEI ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OEl ratings.
(її) Procedures. Although not a regulatory requirement, it is recommended that a formal written certification plan be used to assure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA/AUTHORITY because it identifies and defines an acceptable resolution to the critical issues early in the certification process. These are the usual steps to be followed when utilizing a certification plan:
(A) Prepare a certification plan which describes the analytical procedures and/or the qualification tests to be utilized to demonstrate protection effectiveness. Test plans should describe the rotorcraft and FADEC system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed tests).
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