AC 27-1B, Chg 9 6/23/23 TH MAR = 275 - (275 + (100 - 90)) = 10° F At 5,000 feet:
TH MAR = 275 - (270 + (82 - 80)) = 23° F At 10,000 feet:
TH MAR = 275 - (285 + (64 - 60)) = 6° F At 15,000 feet:
TH MAR = 275 - (274 + (46 - 42)) = -3° F At 20,000 feet:
TH MAR = 275 - (270 + (28-20)) = -3° F Since TH MAR comes out negative at the 15,000- and 20,000-toot 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 275° F.
5. 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:
TAT MIN = Minimum outside air temperature at which temperature tests are conducted.
TMIN = Minimum temperature to which the installed equipment has been tested in the laboratory.
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TTEST MIN = Minimum installed equipment temperature recorded during tests.
ToRL = 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 - (TTEST MIN + (TORL - 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 0 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 (TTEST 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 (TORL must be calculated for each altitude test point) @ sea level, TORL = -25° F @ 5,000 ft., ToRL = -25° F -5,000 ft., × 3.6°/1,000 ft. = -43° F @ 10,000 ft., ToRL = -25° F - 10,000 ft. x 3.6°/1,000 ft. = -61° F @ 15,000 ft., ToRL = -25° F - 15,000 ft. × 3.6°/1,000 ft. = -79° F* @ 20,000 ft., ToRL = -25° F - 20,000 ft. x 3.6°/1,000 ft. = -97° F* *According to § 27.1043(b), the lowest temperature to be considered is -69.7° F.
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AC 27-1B, Chg 9 6/23/23 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 In this example, there is an acceptable margin at all altitudes up to and including 15,000 teet. However, at 20,000 feet, the margin is negative and the system fails.
(B) Vibration tests. Normally, installed vibration tests are not necessary for equipment qualified in accordance with the latest revision of RTCA DO-160. This document categorizes vibration tests according to installed rotorcraft equipment location such as fuselage, engine compartment, instrument panel, and equipment rack.
However, installed equipment vibration tests may be necessary when it appears the equipment location environment may exceed the laboratory-tested equipment vibration limits.
(C) 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. See introductory text of section §§ 27.1309 and 27.1309(d).
(10) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC).
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(ı) Explanation.
(A) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on part 27 rotorcraft that meet the category A engine isolation requirements. 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, very large scale integration, and complementary metallic oxide silicon technologies that 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.
(B) Since elements of electrical and 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 should 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 may be assured that when the rotorcraft experiences "the foreseeable operating conditioni 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 that 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 that 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 that could result in simultaneous failures that would cause a reduction in power greater than the loss of one engine, this would also be considered "full authority."
Note: If OE ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OE ratings.
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(ii) 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 because it identifies and defines an acceptable resolution to 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 that describes the analytical procedures and 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).
(B) Obtain FAA concurrence that the certification plan is adequate.
(C) Obtain FAA concurrence on detail part conformity of the test articles and installation conformity of applicable portions of the test setup.
(D) Schedule FAA witnessing of the test.
(E) Submit a final test report describing all results and obtain FAA approval of the report.
(ili) Definition of Environment. The SAE AE4L Committee report dated June 6, 1999, remains acceptable criteria to define the worst-case lightning strike that may be encountered by the rotorcraft in service. An additional explanation of the lightning environment can be found in FAA Report DOT/FAA/CT-89/22, Aircraft Lightning Protection Handbook. This handbook will assist aircraft design, manufacturing, and certification organizations in protecting aircraft against the direct and indirect effects of lightning strikes, in compliance with the FAA's regulations. It presents comprehensive test criteria to provide the essential information for the in-flight lightning protection of all types of fixed wing, rotary wing, and powered lift aircraft of conventional, composite, and mixed construction and their electrical and fuel systems.
The handbook contains chapters on the natural phenomenon of lightning, the interaction between the aircraft and the electrically charged atmosphere, the mechanism of the lightning strike, and the interaction with the airframe, wiring, and fuel system. Further chapters cover details of designing for optimum protection; the physics behind the voltages, currents, and electromagnetic fields developed by the strike; and shielding techniques and damage analysis. The handbook ends with discussion of test and analytical techniques for determining the adequacy of a given protection scheme. For new designs and applications, the revised definitions of the lightning environment in AC Page F - 30.22
20-136B, Aircraft Electrical and Electronic System Lightning Protection, dated September 7, 2011, should be used.
(iv) Certification Plan. The following subjects are not intended to provide a complete list of the items that should be included in the certification plan, but rather highlight some of the areas that should receive consideration. The certification plan should address the total protection required to allow the FADEC to continue to operate properly when the rotorcraft experiences a worst-case lightning strike encounter.
(A) Determination of Lightning Strike Attachments. Determine the locations on the rotorcraft where lightning strike attachment is likely to occur and the portions of the airframe through which currents may flow between attachments. The main and tail rotors are recognized as likely attachment points; however, consideration should be given to all possible attachment points. The swept stroke phenomenon may not exist for all lightning strike encounters due to the fact that the rotorcraft may be airborne with little or no airspeed.
(B) Establish the Lightning Environment. Establish the components of the total lightning event to be considered. These are the currents and voltages that are described in the definition of the environment.
(C) Full-Level, Complete Vehicle Testing. In accordance with traditional FAA policy, the demonstration that the FADEC installed in a complete type design rotorcraft will continue to operate properly when exposed to a worst-case lightning strike is sufficient to demonstrate compliance with § 27.1309(a). Because of the difficulties involved in utilizing this type of an approach, it is generally not used.
(D) Analytical Processes. A description should be given in the certification plan of the analytical process and certification tests to be utilized to demonstrate protection effectiveness. Typically, the certification plan will include a combination of analysis and tests. (Analytical techniques are most often utilized to predict the levels of lightning-induced transients in interconnecting wiring.) In most cases, successful analyses are based upon well-defined geometrical or electrical parameters such as structural dimensions and materials resistivity. When electrical characteristics of structural materials are not well established, development tests are often utilized to obtain this data, which is subsequently utilized in an analysis. In more complex structures or electrical and electronic system installations, it is sometimes difficult or impossible to define the problem in terms that can be analyzed. In these cases, development or verification testing is often relied upon. The purpose of the certification plan is to show how developmental tests, analyses, and verification tests are combined to demonstrate protection design adequacy. In certain cases, previously verified designs can be incorporated and their adequacy confirmed by reference to previous verifications. Such reference should also be incorporated in the certification plan.
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1. The verification testing should be conducted on a system that simulates as closely as possible the installed configuration. As few items as possible of actual hardware should be simulated.
2. Unless the applicant has opted to follow the guidance in AC 20-136B as a means of showing compliance with § 27.1309(d), then the following or other acceptable guidance should be followed. The use of various analytical processes usually requires that the system component tolerance is established. The SAE AE4L Committee Report No. AE4L-81-2 is one example of an acceptable reference to the testing that will let control hose tolerances. the testin determine these tolerances. The testing performed to determine the tolerance level of the control computer should include a consideration for the occurrence of a nonrecoverable digital upset. One method to provide this consideration is to have the unit powered and the processor operating normally under software control (usually this should be the exact software for which approval is sought) when the test is performed. If strike testing is used, then several shots should be made to develop enough data to provide a reasonable confidence level. It is acceptable for the engine manufacturer, while it is obtaining the engine type certificate, to accomplish this bench testing to determine the level of tolerance of the FADEC system components to lightning encounter indirect effects. This approach has the advantages that the bench tests are not necessarily required to be repeated when the engine is installed in a different airframe. This recommendation is not meant to add a requirement to the engine manufacturer but to propose a more efficient method of certification. If this tolerance was not determined by the engine manufacturer, the applicant installing the FADEC in a rotorcraft would be expected to furnish this data.
3. As with any analytical method, it is prudent to include a margin of safety to account for the uncertainties involved in the analytical and testing processes. Margins account for uncertainty in the compliance method. As confidence in the compliance method increases, the margin can decrease. A factor of two is an acceptable margin for systems with catastrophic Failure Conditions, if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be approved by the cognizant certification office.
4. When an analysis has no associated full-scale vehicle testing to confirm the analysis, the analysis should be very rigorous. Additionally, it should be expected in this situation that this analysis indicates a very large margin of protection. Many factors need to be considered in determining what constitutes an acceptably large margin. The specific additional margin required should be based on an assessment of the inherent uncertainty of a given analysis.
(E) Pass or Fail Criteria. The certification plan should address pass or fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance:
1. No immediate crew action must be needed.
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2. Automatic control of the engine cannot be lost for any appreciable period of time. The engine should not be allowed to be out of control for a period of time that will result in a hazard in a worst-case flight condition. Any rapid, uncontrolled divergence would not be acceptable.
3. No crew action should be required to reset the system. This is not to imply that the system cannot be designed with a manual reset, but the manual reset cannot be used to show compliance to recover from a digital upset.
4. The resumption of engine control after an upset should be reasonably within the range that existed before the upset.
5. No critical data can be lost.
6. After the system recovers, if the performance of the system has been degraded in a noncritical manner that would reduce the capability of the rotorcraft or the ability of the pilot to cope with adverse operating conditions, then an alert to the crew about this system degradation is needed.
(F) System Installation Considerations. In most cases, the installation of the system components is a constituent part of the lightning protection. This is particularly true in the use of shielding techniques. If these installation features are required for adequate lightning protection, consideration should be given to ensure that their effectiveness is not derogated in service. ICAs must be furnished to the aircraft owner on initial delivery, and thereafter made available to parties who service and operate the rotorcraft, to allow them to take actions necessary to ensure the continued effectiveness of the system lightning protection. See § 21.50(b).
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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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