AC 27.1316 § 27.1316 (Amendment 27-46) ELECTRICAL AND ELECTRONIC SYSTEM LIGHTNING PROTECTION.
a. Explanation.
(1) 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, because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design of rotorcraft utilizes the same advanced technology systems and materials as airplanes. Because of this, a specific requirement was added by Amendment 27-46 establishing standards for lightning strike protection of required systems, equipment, and installations. Section 27.1316 replaces paragraph (d) of § 27.1309, which previously defined the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft and systems.
Note: Aviation authorities in other countries may not have issued regulations or standards similar to § 27.1316 for lightning protection. As a result, authorities in other countries may use other means, such as special conditions, to establish requirements for lightning protection.
b. Procedures.
(1) Guidance for you to show compliance with the applicable sections of § 27.1316 is available in a dedicated AC 20-136B. AC 20-136B is available for download from the FAA Regulatory Guidance Library website at http://rgl.faa.gov/.
(2) EASA Interpretative Material is introduced by a CRI referring to EUROCAE documents (ED) ED-113, ED-81, and FAA AC 20-136B.
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AC 27-1B, Chg 4 3/31/2014 AC 27.1317. § 27.1317 (Amendment 27-42) HIGH INTENSITY RADIATED FIELDS (HIRF) PROTECTION.
a. Explanation.
(1) Regulations amendments in 2007 added requirements for the protection of aircraft electrical and electronic systems from the effects of High Intensity Radio Frequency (HIRF) environment. This effort was due to technological advances in airframe and electronic systems design and a concurrent increase in the levels of radiated power in the aircraft environment. These changes have raised vulnerability to the electromagnetic environment of the electrical and electronic systems, which perform critical and essential functions. Prior to this regulatory requirement, the issuance of special conditions for products involving advanced electrical and electronic systems provided an adequate level of safety. The new regulation included a five-year period of relief from the new testing requirements by allowing an applicant to show that the system continues to comply with previously issued HIRF special conditions. Beginning December 1, 2012, data used to show compliance as part of a previously issued special condition is no longer accepted as a means of showing compliance with paragraph (a) of § 27.1317. All systems will be required to show compliance data to the appropriate paragraphs and sections of the regulation, based on the HIRF safety analysis.
(2) EASA has not yet adopted regulatory requirements for HIRF protection.
EASA is using special conditions invoking the JAA INT POL 27/29-1 Issue 3, dated January 10, 2003. The requirements addressed in the JAA INT POL are similar to those of § 27.1317(a)(b)(c).
b. Procedures.
(1) Guidance for you to show compliance with the applicable sections of § 27.1317 is available in AC 20-158. This AC is available for download from the FAA Regulatory Guidance Library website: http://rgl.faa.gov/
(2) EASA Interpretative Material provided in EUROCAE ED-107 and AMJ 20.1317 is considered by the FAA to be technically equivalent to the guidance material in AC 20-158 and may be used as a means of compliance with § 27.1317.
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AC 27-1B, Chg 7 2/4/2016 [This page intentionally blank.] Page F - 33
AC 27.1321. § 27.1321 (Amendment 27-13) ARRANGEMENT AND VISIBILITY.
a. Background. Part 27 contains specific requirements for locating instruments to allow pilots to operate the rotorcraft safely within authorized limits and to indicate system conditions. The instruments should be arranged for use by any pilot and must be readily visible to the pilot. The design for the instrument location and arrangement, with respect to the pilot's seat, should accommodate 5th to the 95th percentile of male and female pilots. Pilots within this height range should be able to see and, where necessary, reach and operate all of the displays.
b. Procedures.
(1) When evaluating the location of the instruments for rotorcraft approved for VFR (IFR locations are discussed in section AC 27 Appendix B of this AC), the placement of flight, navigation, and powerplant instruments should be such that the pilots) can easily see and read the instruments when seated normally. Additionally, the instruments' location should minimize unnecessary pilot head turning.
(i) Primary field of view (FOV): Figure AC 27.1321-1 depicts the primary FOV for rotorcraft. Display high priority information and the primary flight information in the primary FOV.
LT 15 ° I C/L UP Horiz 15 +15 ° RT Horizontal Vertical -15 ° NOM DN Figure AC 27.1321-1 PRIMARY FIELDS OF VIEW
(ii) Secondary FOV: Figure AC 27.1321-2 depicts the secondary fields of view. Signals and information pilots scan as part of monitoring the aircraft and its performance are located in the secondary FOV. The priority of the information and signals when locating information in the secondary field of view should be considered.
The farther away from the primary field of view a signal is located, decreases the chance a pilot will notice it without a prompt. Additionally, the pilot's scan workload to see information increases the farther away from the pilot's primary field of view it is located. The secondary FOV is bound laterally to minimize the need for pilots to turn their heads to view the information. Examples of information typically located in the secondary FOV are:
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(A) Ancillary navigation information (moving maps, weather displays, etc).
(B) Secondary powerplant information, such as engine oil pressures and temperatures.
(C) caution advisory and warning system (CAWS) panel (unless a master caution or warning indicator is not located in the primary FOV, in which case the CAWS panel should be as close to the pilots primary FOV as possible).
(D) Autopilot (if installed).
(E) Navigation controls.
UP LT Horizontal Vertical 30° C/L Horiz 30° 15° 20° NOM RT DN Figure AC 27.1321-2 SECONDARY FIELDS OF VIEW
(2) On multiengine rotorcraft, the required powerplant instruments should be grouped so there will be no confusion regarding which engine an individual gage represents. This is usually accomplished by mounting the engine gages in a vertical grouping. Identical parameter gages are placed next to each other and positioned from left to right in the same position and sequence as the engine location in the airframe.
(3) An evaluation should be made to determine that vibration of the instrument panel does not exceed the tolerances of the instrument. The instrument manufacturer will usually provide data that indicate the level of vibration for which the instrument has been qualified. The flight test evaluation of the rotorcraft should explore and determine про рату ест полоно го that the vibration of the instrument panel does not affect the readability of the instrument. To meet these two criteria, it has been necessary in some installations to "shock mount" or otherwise isolate the instrument panel.
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AC 27-1B, Chg 7 2/4/2016
(4) The flight test evaluation should also determine that the flags or malfunction indicators of the instruments are readily visible in all combinations of lighting for approved kinds of operations.
AC 27.1322. § 27.1322 (Amendment 27-11) WARNING, CAUTION, AND ADVISORY LIGHTS.
a. Explanation.
[Section AC 27.1322 continued on Page F - 35.] Page F - 34.2
(1) Cockpit devices are color-coded to symbolically represent various functions and varying levels of importance for flightcrew operation. From early times, an attempt has been made to take full advantage of associations developed early in life as a result of continuous exposure to our daily environment.
(2) Military design specifications were the first to reference color-coding in cockpit design requirements. In the mid-1940s, the CAA initiated the first color-coding requirements for civil cockpit design. Color-coding standards for cockpit visual signals soon followed. MIL-STD-411, May 31, 1957, identified three separate categories of light signals:
Warning Light - indicates the existence of a hazardous condition which may require immediate corrective action.
(її) Caution Light - serves to alert the operator to an impending dangerous condition requiring attention but not necessarily immediate action.
(iii) Advisory Light - indicates safe or normal configuration, condition of performance, operation of essential equipment, or attracts attention for routine purposes.
(3) Examples of warning and caution signals are included in later versions of the military standard, and a few of those are shown below:
Warning Signals Caution Signals Cabin Pressure Failure Trim Failure Fire Fuel Low Fuel System Failure Generator Inoperative Landing Gear Unsafe Defrosting Failure
(4) Specific color designation for civil advisory lights was first addressed in Amendment 3 to the rotorcraft certification rules (Parts 27 and 29) on January 19, 1968, with adoption of new §§ 27.1322 and 29.1322.
(5) In a subsequent revision (Amendment 27-11), green lights were redesignated and additional colors introduced for flexibility in the requirement.
b. Procedures.
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(1) Red shall be reserved for annunciation of emergency conditions requiring immediate corrective action. Typical examples include fire, transmission oil pressure, engine failure, and battery overheat. The use of red for annunciators which do not require immediate action must be avoided. Use of red when it is not needed tends to lessen the impact of a red annunciator and the needed pilot association for immediate action. In evaluating cockpit annunciators for acceptability, the FAA should ensure all annunciators which require immediate action are red and that only those requiring such action are red. If a master warning light is provided, it should be red, and it should be powered by the same signal that powers any of the individual red warning signals. An aural warning may accompany visual warning signals to enhance pilot response. Care should be taken that any aural signal is sufficiently distinct from other aural warnings, such as low rotor RPM to prevent confusion and to ensure proper crew response. A means to deactivate and reset the master warning (visual and aural) is required.
Resetting the master warning must not deactivate any individual warning signal.
(2) Amber shall be reserved for indicating malfunction or failure conditions which do not require immediate crew action to ensure safe flight. Typical examples include door unlatched, inverter failure, generator failure, fuel filter clogged, and parking brake engaged. Amber should generally be utilized for malfunction and failure conditions which do not require immediate action. The key word here is "require."
Obviously, a pilot should perform corrective action for malfunction or failure conditions in a timely manner as soon as other cockpit priorities allow. The time increment associated with "immediate action" may vary with the system involved, the flight regime, and the aircraft; however, 15 seconds is a representative value in evaluating this term.
This by no means indicates that any red annunciator can be ignored for 15 seconds.
For red annunciators, some type of immediate pilot response is expected. If immediate pilot action is not required, the FAA should recommend the use of an amber designation. If a master caution light is provided in addition to a master warning light, the master caution annunciator should be amber, and should be powered by the same signal that powers any of the individual amber caution signals. Reset considerations for the master caution are the same as those detailed above for the master warning.
(3) Green signifies a safe operating condition and more specifically has come to signify landing gear extended and locked. Extensive use of green annunciators throughout the cockpit should generally be avoided due to possible confusion with the special use of green for landing gear. If green annunciators are physically and functionally removed from the landing gear operation, they may be found acceptable for a variety of "safe operating" applications. One such application is "all green for approach" used in autopilot, flight director, and other navigation system displays.
(4) Other colors may be utilized as advisory lights in accordance with § 27.1322(d). Red and amber must not be used as advisory lights due to the possibility of introducing confusion into the cockpit. Obviously, yellow and pink annunciators should be avoided due to their similarity to amber and red. White and blue have been successfully utilized as advisory segments in past civil designs.
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(5) The primary test for designation of color is:
Red - Is immediate action required?
(ii) Amber - Is pilot action (other than immediate) required?
(III) Green - Is safe operation indicated, and is the indication sufficiently distinct to prevent confusion with the landing gear down indication?
(iV) Other advisory lights - Is the meaning clear and distinct enough to prevent confusion with other annunciations? Do the colors which are utilized differ sufficiently from the colors specified in paragraphs b(1), (2), and (3) above?
(6) Annunciator lights should be visible during bright daylight conditions. This should include visibility in direct sunlight unless lights are located in such a manner that direct sunlight cannot impinge on them.
(7) If dimming capability is provided, all annunciators, including master warning and caution, may be dimmable so long as the annunciation is clearly discernable for night operation at the lower lighting level. Undimmed annunciations have been found unacceptable for night operation due to disruption of cockpit vision at the high intensity.
The dimming circuit should automatically revert to the high intensity setting when power is removed. Automatic dimming/brightening through the use of a photo cell is also acceptable, as are circuits which enable a dimming switch through a position light or other cockpit lighting controls.
(8) The use of flashing lights should be minimized. If a flashing feature is used, it should be controllable through pilot action so that flashing annunciation does not persist indefinitely. The indicator should be so designed that if it is energized and the flasher device fails, the light will illuminate and burn steadily.
(9) The activation of caution and warning lights should readily attract the attention of the appropriate crewmember while performing duties under both normal and high workload conditions.
c. Annunciator Panel Design
(1) Explanation.
The annunciator panel design should be reviewed for the presence of failure modes that can cause illumination of multiple panel segments.
Many test circuits that are diode isolated are vulnerable to this condition. A typical sequence begins with the shorting of a test circuit diode. This failure is undetectable and goes unnoticed until an actual failure condition occurs which Page F - 37
causes the associated panel segment to illuminate. At this time all panel segments connected to the test circuit will illuminate.
(iii) This configuration becomes a special problem when one or more of the panel segments are red. A red light calls for immediate action by the crew, and the crew does not have adequate information for immediate action when many false panel segments are illuminated.
(iv) If the design review indicates a problem, a redesign of the panel to eliminate the condition is considered to be the best solution and is highly encouraged.
(2) Procedures.
(1) An alternative to panel redesign might be the following:
(A) Review the annunciator panel design and note which segments are red.
(B) Determine if cross reference information is available in the cockpit to allow elimination from consideration of any of the red segments (e.g., red low fuel pressure light and low fuel pressure gauge). Normal operation of the gauge would be a reason to assume the light did not cause the problem.
(C) Where a cross reference is available, further design review of that function is not necessary; however, it may be appropriate to include procedural information in the emergency procedures section of the rotorcraft flight manual.
(D) If cross references are not available for red segments, additional isolation should be incorporated into the annunciator design for those functions.
(ii) If cross referencing is not practical, the following approach is encouraged.
(A) Review the annunciator panel design and note which segments are red.
(B) Determine if isolation diodes are checked during the application of battery or external power before starting the engines (e.g., red low oil pressure light). If isolation diode is shorted, all panel segments will light as soon as battery or external power is applied.
(C) When the isolation diode can be checked before starting engines, further design review is not necessary.
(D) If diodes are not automatically checked before starting, then additional isolation, should be considered.
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(3) Annunciator Panel Arrangement. The annunciator panels should be arranged in a logical manner to reduce the crew's time required to locate faults and to increase their efficiency in following aircraft flight manual procedures. For example, engine annunciators on multiengine rotorcraft should be physically located on the panel to coincide with engine location (left or right) so that properly operating engines are not inadvertently shut down due to crew confusion over which engine has malfunctioned.
AC 27.1323. § 27.1323 (Amendment 27-13) AIRSPEED INDICATING SYSTEM.
a. Explanation.
(1) The accuracy of all flight test data concerned with the velocity of the rotorcraft is dependent on the calibration of the airspeed indicating system. For this reason, the airspeed system position error should be determined very early in the program.
(2) Since air density varies with altitude, the speed reading will only be correct under standard sea level conditions. However, in an actual installation, the indicator reading, even under standard sea level conditions, may differ from the calibrated airspeed because the static system does not sense true static pressure. This error in detection of static pressure is called position error. It is caused by the pressure field built up around the rotorcraft in flight. This pressure field will vary in intensity with dynamic pressure making the position error a function of calibrated airspeed. Since airspeed information is presented to the crew in terms of indicated airspeed, it is necessary to determine the position error for the rotorcraft to be flown safely.
b. Procedures.
(1) There are different methods to determine position error such as trailing bomb, airspeed course, boom system, etc. Each method has its own advantages and disadvantages, but will yield satisfactory results if done correctly. The airspeed system should be calibrated throughout the airspeed range of the rotorcraft and under the various flight conditions of cruise, climb, and autorotation. In addition, the effects of gross weight and center of gravity should be investigated.
(2) It may also be necessary to recalibrate the system with a change in external configuration if such a change may affect the airflow near the pitot or static sources.
(3) Additional information regarding position error is included in paragraph AC 27 Appendix B b(10) and should be considered if pursuing an FR approval.
(4) Static system installation information is included in paragraph AC 27.1325.
Technical Standard Order (TSO) C16, Airspeed Tubes (Heated), gives minimum Page F - 39
performance standards for pitot tubes, and pitot tubes qualified to this TSO normally allow for a satisfactory aircraft installation.
(5) The calibration requirements of the standard seem to be self-explanatory and are not discussed further in this paragraph.
AC 27.1325. § 27.1325 (Amendment 27-13) STATIC PRESSURE SYSTEMS.
a. Explanation.
(1) This section, in conjunction with § 27.1323, provides minimum performance standards for static pressure systems. The standard provides some relief when considering the icing environmental condition in that it allows the use of an alternate static port to account for the icing condition.
(2) The standard for the consideration of environmental conditions is § 27.1309(a).
(3) The standard for consideration of malfunction conditions is § 27.1309(b).
(4) For rotorcraft that will be approved for IFR operation, the provisions of Appendix B VIII(b)(5) of Part 27 as discussed in paragraph AC 27 Appendix B should also be considered.
b. Procedures. The installation of the static system should consider the following:
(1) Static lines should be initially routed upward immediately behind the static pressure port. This procedure will minimize the entry of moisture into the system when operating in rain or washing the rotorcraft.
(2) Drain(s) should be located at low points in the system. Line routing and clamping should allow for all moisture that does enter the system to be routed to the
(3) If independent systems are provided, the placement of each system component should allow for maximum practicable separation of each system. As much as possible, one system should be on one side of the rotorcraft and the second system on the opposite side.
(4) Most static pressure ports that are provided for IFR operation are heated.
Before any tests are conducted, a program to qualify the heater on the port should normally be agreed upon through discussions between the FAA and the applicant. It is suggested that the requirements of TSO C16, Airspeed Tubes (Heated), be used as a guide for these discussions. If the ports are not to be heated, a comprehensive analysis should be prepared, and limited testing should be conducted to verify the analysis.
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