(5) Assessment Methods. Various methods for assessing the causes, severity, and probability of Failure Conditions are available to support experienced engineering and operational judgment. Some of these methods are structured. The various types of analysis are based on either inductive or deductive approaches. Probability assessments may be qualitative or quantitative. Descriptions of some types of analysis are provided below and in SAE ARP 4761.
(i) Design Appraisal. This is a qualitative appraisal of the integrity and safety of the system design.
(ii) . This is a qualitative appraisal of the integrity and safety of the inst Installation Appraisal allation. Any deviations from normal, industry-accepted installation practices such as clearances or tolerances, should be evaluated, especially when appraising modifications made after entry into service. Hazards introduced by the installation of the system should also be evaluated.
(iii) Failure Modes and Effects Analysis (FMEA).
(A) This is a structured, inductive, bottom-up analysis, which is used to evaluate the effects on the system and the rotorcraft of each possible element or component failure. When properly formatted, it will aid in identifying latent failures and the possible causes of each failure mode. SAE ARP 4761 provides methodology and detailed guidelines that may be used to perform this type of analysis. An FMEA could be a piece part FMEA or a functional FMEA. For modern microcircuit based line replaceable units (LRUs) and systems, an exhaustive piece part FMEA may not be practically feasible given current technology. In that context, an FMEA may be more functional than piece part oriented. A functional oriented FMEA can lead to uncertainties in the qualitative and quantitative aspects, which can be compensated for by more conservative assessment, such as:
1. Assuming all failure modes result in the Failure Conditions of interest,
2. Carefully choosing system architecture, or
3. Taking into account the experience lessons learned on the use of similar technology.
(B) Specific reliability numbers are not shown in § 29.1309. The necessary degree of reliability is a function of the criticality of the system under consideration. Acceptable sources of component failure rates include (1) military aircraft component failure or fault reports or handbooks, such as MIL-HDBK-217C; (2) operator Page F - 40.12
or manufacturer component malfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.
Fault Tree or Dependence Diagram Analysis Markov Analysis - Common Cause Analysis (CCA) . This is a structured, deductive,
(iv) top- down analysis that is used to identify the conditions, failures, and events that would cause each defined Failure Condition. It is a graphical methods of identifying the logical relationship between each particular Failure Condition and the primary element or component failures, other events, or combinations thereof that can cause it.
A failure modes and effects analysis may be used as the source document for those primary failures or other events.
(v) . A Markov model (chain) represents various system states and the relationships among them. The states can be either operational or non operational. The transitions from one state to another are a function of the failure and repair rates. Markov analysis can be used as a replacement for fault tree and dependence diagram analysis, but it often leads to a more complex representation, especially when the system has many states. It is recommended that Markov analysis be used when fault tree or dependence diagrams are not easily usable, namely to take into account complex transition states of systems that are difficult to represent and handle with classical fault tree or dependence diagram analysis.
(vi) . The acceptance of adequate probability of Failure Conditions is often derived from the assessment of multiple systems based on the assumption that failures are independent. Therefore, it is necessary to recognize that such independence may not exist in the practical sense and specific studies are necessary to ensure that independence can either be assured or deemed acceptable.
CCA identifies common causes for Failure Conditions. Once identified, the source of the common cause is either eliminated from the design, the common cause or the system design is made to be tolerant of the related fault or malfunction, or the common cause or system design combination is found to be acceptable due to the failure or malfunction condition category it represents. The potential for failures or malfunctions due to common causes is inherent in designs that provide multiple functions reliant on common hardware or common software. This is also true for systems that produce related functions and share a common installation area. The installation area may represent a threat from several sources such as Electro-Magnetic Interference (EMI), mechanical hazards, and environmental influences. The CCA is sub-divided into three areas of study:
Zonal Safety Analysis (ZSA). This analysis has the objective of ensuring that the (A) equipment installations within each zone of the rotorcraft are at an adequate safety standard with respect to design and installation standards, interference between systems, and maintenance errors. The ZSA examines the physical zone of the rotorcraft in which the system under consideration is installed to ensure that the surrounding equipment or appliance installations do not compromise the system Page F - 40.13
independence requirements. Mechanical failures that might generate fragments that could damage the system under evaluation are an example of the type of event that would be considered in a zonal analysis.
(B) Particular Risk Analysis. Particular risks are defined as those events or influences that are outside the systems concerned. Some examples are fire, leaking fluids, bird strike, tire burst, high intensity radiated fields exposure, lightning, and uncontained failure of high energy rotating machines. Each risk should be the subject of a specific study to examine and document the simultaneous or cascading effects or influences that may violate independence.
(C) Common Mode Analysis (CMA). This analysis is performed to confirm the assumed independence of the events that were considered in combination for a given Failure Condition. The effects of specification, design, implementation, installation, maintenance and manufacturing errors, environmental factors other than those already considered in the particular risk analysis, and failures of system components should be considered. The CMA considers many aspects of design; one is design dissimilarity for both hardware and software. Without dissimilarity for redundant applications, there is a possibility that a hardware or software failure or malfunction could occur in the same flight for all redundant subparts of a system with a Catastrophic or Hazardous Condition Category. Another major contribution from the CMA is the determination of which failure or malfunction combinations inputs to the Fault Tree Analysis (FTA) must be independent for events that are Catastrophic or Hazardous.
This analysis is iterative in nature as it will be employed early in design to identify possible common modes for failures or malfunctions and then used after the design is complete to determine if the FTA goals have been met.
(6) Documentation. All laboratory, ground, and flight tests and failure analyses, must be documented in sufficient detail to show compliance with § 29.1309 and must be included in the type design file. Section 21.31(a) provides the regulatory basis for requiring this documentation. If the applicant elects to use a numerical reliability and probability analysis, it must also be documented in sufficient detail.
(7) . RTCA DO-178C, Software Considerations in Airborne Systems and Equipme Software nt Certification, dated December 13, 2011, is the latest standard and may be used for qualification and subsequent approval of airborne software. See AC 20- 115C, Airborne Software Assurance, for guidance on using DO-178C and earlier standards.
(8) Airborne Electronic Hardware (AEH). For airborne complex custom micro coded electronic components assessed as level A, B, or C in the safety assessment, RTCA DO-254, Design Assurance Guidance for Airborne Electronic Hardware, dated April 19, 2000, provides certification guidance for qualification and subsequent approval of these components. Per DO-254, the component will need to be identified as either “simple” or “complex” and tagged to one of the five failure condition classifications.
Based on these identifiers, DO-254 provides guidance to show compliance with Page F - 40.14
§ 29.1309. Although RTCA DO-254 applies specifically to complex custom micro-coded components, applicants are highly encouraged to apply DO-254 to other hardware components up to the LRU and systems level.
(9) Environmental Qualification.
(i) Laboratory Tests.
(A) Environmental Standards. In order to ensure that the components and systems under consideration will function properly when exposed to adverse environments, they should be tested in the laboratory under a simulated adverse environment. If a TSO exists and is appropriate in environmental range and performance for an equipment installation, it is preferable the equipment be TSO approved. If there is no applicable TSO or an existing TSO does not provide for a sufficiently adverse environment, the latest revision of RTCA DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.
(B) Explosion Tests. Those items of electrical and electronic equipment that are to be located in areas subject to flammable fluids and vapors, as a result of any single probable malfunction or failure, including failure of couplings or lines should be tested as an ignition source. These tests consist of normal operation of the equipment in a physically contained explosive atmosphere. The explosion test procedure in the latest revision of DO-160 will satisfy this requirement. Section 29.863 of this AC provides further guidance on safety from explosion. Another standard that is equivalent to the latest revision of DO-160, it may also be accepted to satisfy this requirement.
(ii) Installed Environmental Tests. After the environmental ratings of the components and systems have been established, it should be ensured that as installed, these ratings will not be exceeded. Normally, installed equipment need not be instrumented and tested in flight nor is it necessary to instrument the compartment or rack where the equipment is installed. Satisfactory environment and equipment compatibility are ensured by selection of the proper environmental category of laboratory tests. The category is determined by the type of aircraft (reciprocating or turbine) and flight envelope (altitude and temperature). Exceptions to normal installations are (a) Alternator or generator cooling, where radiated and conducted heat is almost always uncertain, and cooling air temperatures and flow rates are uncertain;
(b) Where flight tests reveal excessive instrument panel vibration. In this case, the panel should be instrumented, tested, and, if necessary, design improvements made;
and (c) Any other case where good engineering judgment and application of sound engineering principles indicate a high likelihood that the installed environment is more severe than the equipment is capable of operating within.
Page F - 40.15
(A) Temperature Tests.
Temperature tests may be accomplished by instrumenting the installed equipment environment with a recorder that provides a permanent record of time, altitude, and temperature. The pertinent temperature should be recorded as the rotorcraft is operated throughout its altitude range, including ground operation. The maximum and minimum temperatures recorded should be corrected degree for degree to ensure the equipment under test remains within its temperature rating while the rotorcraft operates throughout its approved ambient temperature envelope. (For generator or alternator cooling test procedures, refer to section 29.1351 of this AC.)
Section 29.1043(b) requires the maximum approved operating OAT be at least 100° F for powerplant-mounted accessories such as starter generators and vacuum pumps.
Due to the impracticality of the 100° F hot day temperature limit, rotorcraft systems mounted on the powerplant are normally evaluated for at least 115° F hot day sea level conditions with corresponding 3.6° F/1,000-foot correction. The maximum hot day OAT at sea level must be specified in the rotorcraft flight manual per §§ 29.1309 and
29.1581(a)(2). Section 29.1043(b) is the regulatory basis for the lapse rate of 3.6° F/1,000 feet. This lapse rate should be applied regardless of the hot day sea level temperature the applicant chooses to certify for operation.
The § 29.1043(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 ensure compliance.
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.
The following illustrates an acceptable high temperature evaluation method:
TOAT MAX = Maximum outside air temperature at which temperature tests are conducted.
TMAX = Maximum temperature to which the installed equipment has been tested in the laboratory.
TTEST MAX = Maximum installed equipment temperature recorded during tests.
TORH = The high reference outside air temperature. It varies with altitude starting at the highest sea level temperature at which rotorcraft operation is to be Page F - 40.16
approved and decreases at3.6° 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:
Measured Altitude (ft, MSL) Cylinder Temp (TTEST MIN) 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, TORH must be calculated for each altitude test point.
@ sea level, TORH = 100° F @ 5,000 ft., TORH = 100° F -5,000 ft., x 3.6°/1,000 ft. = 82° F @ 10,000 ft., TORH = 100° F - 10,000 ft. x 3.6°/1,000 ft. = 64° F @ 15,000 ft., TORH = 100° F - 15,000 ft. x 3.6°/1,000 ft. = 46° F @ 20,000 ft., TORH = 100° F - 20,000 ft. x 3.6°/1,000 ft. = 28° F Then at sea level:
TH MAR = 295 - (275 + (100 - 90)) = 10° F At 5,000 feet:
TH MAR = 295 - (270 + (82 - 80)) = 23° F Page F - 40.17
At 10,000 feet:
TH MAR = 295 - (285 + (64 - 60)) = 6° F At 15,000 feet:
TH MAR = 295 - (294 + (46 - 42)) = -3° F At 20,000 feet:
TH MAR = 295 - (290 + (28 -20)) = -3° F Since TH MAR comes out negative at the 15,000- and 20,000-foot points, the generator fails. It will be necessary for the applicant to reduce the maximum load current, improve cooling, or otherwise change the design to ensure the generator is operating within its approved temperature limit of 295° F.
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.
Page F - 40.18
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 Page F - 40.19
(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:
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 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 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.
Page F - 40.20
(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 section of §§ 29.1309 and 29.1309(d).
(10) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC).
(i) Explanation.
(A) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on part 29 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 de signed 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 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 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 Page F - 40.21
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 OEI ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OEI ratings.
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
(ii) 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 test(s).
(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.
(iii) 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 Page F - 40.22
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 20-136B, Aircraft Electrical and Electronic System Lightning Protection, dated September 7, 2011, should be used.
Certification Plan. The following subjects are not intended to provide a compl
(iv) ete 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.
(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 § 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 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 Page F - 40.23