- AC 29-2C, Chg 4 5/1/2014 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 assure the generator is operating within its approved temperature limit of 295° 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:
TOAT 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.
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.
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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) sea level 0°F 5,000 -9°F 10,000 -21°F 15,000 -32°F 20,000 -40°F OAT = TOAT MAX -25°F -45°F -59°F -65°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., x 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. x 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 § 29.1043(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:
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AC 29-2C, Chg 4 5/1/2014 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.
(B) 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.
(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.
(iii) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC).
(A) Explanation.
(1 2 ) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on rotorcraft with 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, very large scale integration, and complementary metallic oxide silicon 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.
( ) 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 Page F - 34
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.
(3 4 1 2 3 ) 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 § 29.1309(a) as installed.
( ) 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 are used in a multiengine rotorcraft, a common failure caused by a lightning strike 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 should be limited to allow flight manual performance based on OEI ratings.
(B) 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:
( ) 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 test(s).
( ) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.
( ) Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup.
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AC 29-2C, Chg 4 5/1/2014
(4) Schedule FAA/AUTHORITY witnessing of the test.
(5) Submit a final test report describing all results and obtain FAA/AUTHORITY approval of the report.
(C) Definition of Environment. The SAE AE4L Committee report dated June 6, 1999, has been and 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 could 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 14 CFR regulations. It presents a 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. On September 7, 2011 , FAA AC 20-136B, “Aircraft Electrical and Electronic System Lightning Protection,” was issued, superseding previous versions of that AC. For new designs and applications, this revised definition of the lightning environment should be used.
(D) 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 that is required to allow the FADEC to continue to operate properly when the rotorcraft experiences a worst-case lightning strike encounter.
(1) 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.
(2) 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.
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(3) Full-Level, Complete Vehicle Testing. In accordance with traditional FAA/AUTHORITY 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 § 29.1309(a).
Because of the difficulties involved in utilizing this type of an approach, it is generally not used.
( ) Analytical Processes. A description should be given in the certification plan of 4 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.
(i) 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.
(ii) Unless the applicant has opted to follow the guidance in AC 20-136B as a means of showing compliance with § 29.1309(h), 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 S AE AE4L Committee Report No. AE4L-81-2 is one example of an acceptable reference to be used for the testing accomplished to determine these tolerances. The testing which is performed to determine the tolerance level of the control computer should include a consideration for the occurrence of a non-recoverable 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 an acceptable procedure for the engine manufacturer, while he is obtaining his 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 Page F - 37
AC 29-2C, Chg 4 5/1/2014 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.
(iii) 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 agreed upon with the cognizant certification office.
(iv) 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 must 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.
(5) Pass or Fail Criteria. The certification plan should address a pa ss or fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance:
(i) No immediate crew action must be required.
(ii) Automatic control of the engine cannot be lost for any appreciable period of time. The engine must 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. Obviously, any rapid, uncontrolled divergence is not acceptable.
(iii) 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.
(iv vi ) The resumption of engine control after an upset must be reasonably within the range that existed before the upset.
(v) No critical data can be lost.
( ) After the system recovers, if the performance of the system has been degraded in a noncritical manner which would reduce the capability of the rotorcraft or the ability of the pilot to cope with adverse operating conditions, then the crew must be alerted to this system degradation.
(E) System Installation Considerations. In most cases, the installation of the system components is a constituent part of the lightning protection. This is Page F - 38
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 be provided to the parties who service and operate the rotorcraft to allow them to take actions necessary to ensure the continued effectiveness of the system lightning protection.
(iv) Lightning Protection.
(A) 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, because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design transport category rotorcraft are utilizing the same advanced technology systems and materials as transport category airplanes. Because of this fact, a specific requirement was added by Amendment 29-24 for the consideration of lightning strike protection of required systems, equipment, and installations. The addition of paragraph (h) to § 29.1309 further defined the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.
(B) Procedures.
(1 2 ii 3 ) Section 29.1309(h) requires, when showing compliance with § 29.1309(a) and (b), that the effects of lightning strikes on the rotorcraft be considered.
Guidance to show compliance with § 29.1309(h) is available in a dedicated AC 20-136B.
( ) Detailed means of compliance should be agreed with the authorities taking into account the effects on the rotorcraft and minimum considerations are as follows:
(i) Any combination of analysis and testing should be agreed to by the authority.
( ) 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 agreed upon by the cognizant certification office.
( ) Flight and engine controls are examples of “critical” functions. With these critical functions defined, an analysis and testing should be performed to show compliance with § 29.1309(a) (i.e., equipment, systems, and installations performing critical functions should be designed and installed to ensure they continue to perform those intended functions subject to the conditions encountered in a worst-case lightning Page F - 39
AC 29-2C, Chg 4 5/1/2014 strike). Section 29.610 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 § 29.1309(a). These avionic systems usually only require protection from indirect effects of lightning and therefore a method such as that outlined in paragraph b.(9)(iii) above 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 fault/failure analysis to determine if the malfunctioning of several “essential” systems in relation to other systems would result in a hazard to a category B rotorcraft or preclude the continued safe flight and landing of a category A rotorcraft. If groups of functions are so identified, sufficient lightning protection should be provided to prevent a hazardous malfunction situation on category B rotorcraft or provide those conditions which prevent continued safe flight and landing on a category A rotorcraft are extremely improbable.
In performing this part of the analysis, attention should be given to the fact that many of the required equipment, systems, and installations may fail simultaneously with other required equipment, systems, and installations and result in a reduction of the capability of the rotorcraft but still not result in a catastrophe. An example of required equipment for which the simultaneous failure of all the required equipment is catastrophic is a failure which results in a total loss of attitude display for IFR certified rotorcraft operating in instrument meteorological conditions. Note that the analysis which is utilized to demonstrate that these failures are extremely improbable should have the encounter with a worst-case lightning strike as a given event (i.e., probability is unity). Additionally, for a category A rotorcraft, an autorotation is not considered continued safe flight and landing.
AC 29.1309A. § 29.1309 (Amendment 29-53) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Explanation Procedures . Amendment 29 -53 removed § 29.1309(h) and incorporated a dedicated rule § 29.1316 to better address Electrical and Electronic System Lightning Protection.
b. . If the certification basis includes § 29.1309 (h), applicants should utilize the previous section of this guidance (section 29.1309 of this AC) to show compliance with protection from lightning effects. If the certification basis includes § 29.1316, applicants should show compliance with protection from lightning effects by referring to the applicable section of this AC (section 27.1316, Electrical and Electronic System Lightning Protection). The paragraphs other than (iv) of section 29.1309 of this AC are still applicable. This should be coordinated with the appropriate certification office early in the discussions so the path to show compliance with electrical and electronic system lightning protection is clear to the applicant and authority.
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AC 29.1309B. § 29.1309 (AMENDMENT 29-59) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Explanation. Amendment 29-59 revised § 29.1309 to address advances in technology and eliminate the distinction between Category A and Category B rotorcraft. The rule also:
Revised section § 29.1309 (a) to clarify the requirements for performing a failure analysis.
Replaced section § 29.1309 (b) with language consistent with contemporary industry practices.
Revised the alerting requirements of section § 29.1309 (c) to remove the term “warning,” which has been interpreted as requiring a red level alert, when the intent was to notify the crew of all required annunciations.
Removed sections § 29.1309 (e), (f), and (g) dealing specifically with electrical systems, as they are covered by other regulatory sections.
(1) Applicability
(i) General Section 29, Subparts B, C, and D . Section 29.1309 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.
. If a specific part 29 requirement 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. This precedence does not preclude accomplishment of a system safety assessment, if necessary. For example, § 29.695 predefines a required level of redundancy and an implied level of 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 the design features required by Subparts B, C, and D of part 29 for aspects such as the performance, flight characteristics, structural loads, and structural strength. However, the requirements in § 29.1309 do apply to any equipment and system that the proposed design uses to comply with the requirements of Subparts B, C, and D (e.g., Health Usage Monitoring System certified for maintenance credit and Stability Augmentation System).
(iii) Section 29, Subpart E.
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(A) Section 29.1309 does not apply to the 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.
(B) Section 29.1309 does not apply to the Rotor Drive Systems (reference § 29.917(b)).
(iv) Section 29, Subpart F. Section 29.1309 does not apply to the loss of function of non-required equipment; however, it does apply to hazards introduced by the installation of that equipment on the rotorcraft and any adverse effects on other systems.
(2) Background.
(i) This guidance describes a method of showing compliance that utilizes the system safety assessment techniques available in the aerospace industry. The safety assessment process described in this guidance can be adopted and applied to equipment, systems, and installations ranging from simple systems to the most critical and most complex systems.
(ii) This guidance is not intended to impose an overly rigorous safety assessment process for those installations involving equipment or systems that are not considered to be complex.
Note: A system is considered to be complex when its operation, failure modes, or failure effects are difficult to comprehend without the aid of analytical methods.
b. Procedures.
(1) Definitions of Probability Classifications.
Note: the following terms were derived from Table 1 – “Failure Conditions Severity as Related to Probability Objectives and Assurance Levels” of SAE ARP 4761.
(i) FREQUENT. The occurrence of any frequent event may be no more often than 10-3 times during the operational life of each rotorcraft.
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(ii) PROBABLE. Probable events may be expected to occur on the order of between 10-3 to 10-5 times during the operational life of each rotorcraft.
(iii) REMOTE. Remote events are not 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 to 10 .
(v) EXTREMELY IMPROBABLE -9 . 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 10-7 -9 . 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 or less is assigned to this classification.
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.
(2) Definitions of Failure Condition Classifications
(i) NO EFFECT MINOR . Failure Conditions may be classified according to the severity of their effects into one of the following categories (reference Figure AC 29.1309-2).
. 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 that would not significantly reduce rotorcraft safety and that 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 that 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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