(B) The § 29.1043 Paragraph (b) maximum ambient temperature definition 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 in the southern United States in midsummer. Normally, proper selection of the altitude temperature category in the latest revision of DO-160 will assure 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.
MAX = Maximum temperature to which the installed equipment has been tested in the laboratory.
TTEST MAX = Maximum installed equipment temperature recorded during tests.
ToR = 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 § 29.1043(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:
Par 621 949
Measured Altitude (ft. MSL) Brush Temp (ITEST 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 First, ToR must be calculated for each altitude test point.
@ sea level, TORH = 100° F @ 5,000 ft., ToR = 100° F -5,000 ft., × 3.6°/1,000 ft. = 82° F @ 10,000 ft., ToRn = 100° F - 10,000 ft. x 3.6/1,000 ft. = 64° F @ 15,000 ft., ToR = 100° F - 15,000 ft. x 3.6/1,000 ft. = 46° F @ 20,000 ft., ToR = 100° F - 20,000 ft. × 3.6°/1,000 ft. = 28° F Then at sea level:
TH MAR = 295 - (275 + (100 - 90)) 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 assure 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 950 Par 621
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.
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 (ITEST 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 Par 621 951
(Tor 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. × 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:
TI MAR = (40-(0 + (-25-(-25)))) = 40° F At 5,000 feet:
TI MAR =-(-40- (-9 + (-43- (-45))) = 33° F At 10,000 feet:
TI MAR = (40- (-21 + (-61- (-59)))) = 17° F At 15,000 feet:
TI MAR = (40 - (-32 + (-69.7 - (-65))) = 3.3° F At 20,000 feet:
TI 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.
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.
Altitude tests. If the equipment has been laboratory tested to the maximum certified altitude of the rotorcraft, installed altitude tests are unnecessary.
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7/30/97 АС 29-2B The installed equipment must be either laboratory tested or tested in the rotorcraft to the maximum certified altitude of the rotorcraft.
(3) Lightning Strike Protection of Full Authority Digital Engine Controls.
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 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.
(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 § 29.1309(a) 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 Par 621 953
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.
(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/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).
(B) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.
(C) Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup.
(D) Schedule FAA/AUTHORITY witnessing of the test.
(E) Submit a final test report describing all results and obtain FAA/AUTHORITY approval of the report.
(iii) Definition of Environment. The SAE AE4L Committee report dated June 20, 1978, is an acceptable criteria to define the worst-case lightning strike which may be encountered by the rotorcraft in service. An additional explanation of the lightning environment may 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 Federal Aviation Regulations.
It presents a comprehensive test criteria to provide the essential information for the in-flight lightning protection of all types of fixed/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 954 Par 621
7/30/97 АС 29-2B 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 March 5, 1990, FAA Advisory Circular AC 20-136, "Protection of Aircraft Electrical/Electronic Systems Against the Indirect Effects of Lightning,i was issued.
Appendix 3 of that document contains an updated quantification of the severe natural lightning environment. Additionally, an AIR-100 policy letter, dated August 25, 1993, addressing multiple burst lightning strikes, should be considered. It is recommended that for new designs and applications after March 5, 1990, this revised definition of the lightning be used.
(iv) Certification Plan. The following subjects are not intended to provide a complete list of the items which should be included in the certification plan, but rather highlight some of the areas which should receive consideration. The certification plan should address the total protection which is 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 which are described in the definition of the environment.
(C) 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.
(D) Analytical Processes. A description should be given in the certification plan of the analytical process and/or 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 Par 621 955
parameters such as structural dimensions and materials resistivities. 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 and/or electrical/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.
(1) The verification testing should be conducted on a system which simulates as closely as possible the installed configuration. As few items as possible of actual hardware should be simulated.
(2) 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 the recommended 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 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 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 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) For complete airframe verification testing, a minimum level of at least 4KA peak and a current rise time of 2KA/microsecond are recommended. It is often difficult to obtain valid results at lower levels due to poor signal-to-noise ratios.
When complete vehicle testing is accomplished at some lower level, or through some alternate test technique such as low level swept CW testing, consideration should be given to nonlinear airframe response, diffusion effects, and alterations in current paths caused by arcing and flashover.
(4) 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.
956 Par 621
A level of 6 dB is recommended for those analyses which are confirmed by the use of reduced level, full-scale vehicle testing. This safety margin is the difference between the airframe installed system responses and the system component tolerance, not an adjustment to the quantification of the atmospheric environment. (The airframe system response to the worst-case lightning event should be at least 6 dB less than the FADEC system computer and components tolerance level. Number of dB is defined as 20 LOG10 (V1N2) and 20 LOG10 (11/l½) where V, and I, are the determined tolerance levels of the system components and V½ and I are the extrapolated airframe response.)
(5) 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. Approximately an additional 25 dB of protection has been deemed acceptable for a reasonably rigorous analysis performed on an airframe for which the response characteristics are known.
(E) Pass/Fail Criteria. The certification plan should address a pass/fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance:
No immediate crew action must be required.
(2) 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 which will result in a hazard in a worst-case flight condition. Obviously, any rapid, uncontrolled divergence is not 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 must be reasonably within the range which 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 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.
(v) System Installation Considerations. In most cases, the installation of the system components is a constituent part of the lightning protection. This is Par 621 957
АС 29-2B 7/30/97 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 derogate in service. Information should be made available 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.
(4) Lightning Protection.
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 has been 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 defines the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.
(її) Procedures.
(A) 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.
(1) 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.
(2) 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.
(3) The lightning strike models to be used for system justification should be as described in SAE AE4L Committee Report AE4L-87-3, Rev. B, dated January 1989 (or later version). The recommended reference for performing such analysis is Society of Automotive Engineers, Aerospace Recommended Practice 926A.
958 Par 621
(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 50 is not recommended for full scale low level pulse testing, due to the difficulty of obtaining valid results with poor signal-to-noise ratios.
(3) For a proven analysis technique, a safety factor of at least 2 will be necessary.
(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 § 29.1309(a); i.e., 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 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 thus identified usually only require protection from indirect effects of lightning. If it is determined such is the case, then a method as outlined in Paragraph 621b(3), Lightning Strike Protection of Full Authority Digital Engine Controls, 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 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.
Par 621 959
for a Category A rotorcraft, an autorotation is not considered continued safe flight and landing.
c. Failure Analyses.
(1) Power and distribution systems should be analyzed to show compliance with § 29.1309.
One acceptable procedure for documenting the analysis is contained in Society of Automotive Engineers (SAL) Fault/Failure Analysis Procedure ARP 926A, revised November 15, 1979.
(II) As a minimum, any analysis should consider the effect of failures of components and systems on the capability of the rotorcraft to perform its intended function without hazard.
(¡i) The analysis should consider the indication of failure. Those latent failures that occur without indication should be considered in all possible sequences and combinations of additional failures until a positive indication of failure is provided.
(iv) The analysis should consider failure of indirectly related parts of installations which could induce failure in the system being analyzed. For example: the effect of hydraulic fluid sprayed on electrical components as a result of a ruptured hydraulic line. Another example is the result of a ruptured bleed air line and its effect on hydraulic, fuel, or electrical lines/cables.
(V) The Type Inspection Authorization (TIA) should call for specific simulated failures, evaluation of failure detection, failure warning, and performance of the remaining system on the ground and in-flight to verify the critical aspects of the failure analysis. The applicant should provide a proposed detailed test procedure for incorporation in the TIA to accomplish this verification. The applicant's proposed tests simulating in-flight failures should be carefully reviewed by both the systems engineer and flight test pilot to assure the flight test crew will not be subjected to hazardous flight.
Where practicable those simulated failures that would be hazardous in flight should be evaluated by ground tests. Analyzed and tested systems (where functioning is required) exhibiting hazards or failing to perform their intended functions under any foreseeable operating conditions must be redesigned to comply with § 29.1309.
(2) Utilization systems that are required or critical as to performance of intended function or result in rotorcraft hazard upon failure should also be analyzed for failures by the procedures of Paragraphs c(1)(a) through c(1)(d) above. Examples of systems which may be critical are autopilots, hydraulic control systems, navigation and flight instruments on IFR approved rotorcraft, and bleed air systems.
960 Par 621