AC 29-2C, Chg 4 5/1/2014 may use less than the entire safety assessment process, but require more than a simple statement to show compliance.
(iii) Elements of the safety assessment process are as follows:
• Functional Hazard Assessment (FHA)
• Preliminary System Safety Assessment (PSSA)
•• Fault Tree Analysis (FTA)
• System Safety Assessment (SSA)
•• Failure Mode and Effect Analysis (FMEA)
• Common Cause Analysis (CCA)
•• Zonal Safety Analysis (ZSA)
•• Particular Risk Assessment (PRA)
•• Common Cause Analysis (CMA) (A) Define the system and its interfaces, and identify the functions that the system is to perform. Determine whether or not the system is complex, similar to systems used on other aircraft, or conventional. Where multiple systems and functions are to be evaluated, consider the relationships between multiple safety assessments.
(B) Identify and classify Failure Conditions. All relevant applicant - engineering organizations, such as systems, structures, propulsion, and flight test, should be involved in this process. This identification and classification may be done by conducting a Functional Hazard Assessment (FHA) which is a top down assessment.
The FHA starts at the rotorcraft operational level to identify possible hazards to the rotorcraft, its occupants, and flight crew. When classifying a function, consider for example, the loss of function, erroneous or misleading, or degradation of the function as possible hazards. The hazards are classified using the five failure condition categories found in Figure AC 29.1309-2. The identified failure conditions must be evaluated with respect to the intended operations and combinations of operations for which certification of the rotorcraft is sought (e.g., VFR, IFR, CAT A, IFR plus CAT A, etc.). The product of the FHA is a determination of the failure or malfunction condition category for each system function and identification of the source of the function. Lower level FHAs that focus on the system level hazards reveal detailed information about the specific part of a system that provides the function under evaluation. This more detailed information may narrow the system part to be considered for compliance with enhanced integrity or for additional analysis. However, it must be recognized that there is a limit to the Page F - 20
number of levels that can be relevant in this process. Too many levels can result in the assessment becoming design-dependent rather than the FHA driving the design, as intended.
(C) The FHA is usually based on one of the following methods, as appropriate:
( ) If the system is not complex and its relevant attributes are similar to those of systems us 1 2 ed on other aircraft, the identification and classification may be derived from design, installation, and service experience appraisals that can be shown to be sufficiently similar.
( ) If the system is complex, it is necessary to systematically postulate the effects on the safety of the aircraft and its occupants resulting from any possible failures, considered both individually and in combination with other failures or events.
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.
(D) Choose the means to be used to determine compliance with Section 29.1309. The depth and scope of the analysis depends on the types of functions performed by the system, the severity of the associated Failure Conditions, and whether or not the system is complex (see Figure AC 29.1309-5). Once the failure conditions are determined to be Major Failure Conditions, experienced engineering and operational judgment, design and installation appraisals, and similarity consideration for previously installed systems may be acceptable. This may be sufficient in itself or in conjunction with qualitative analyses or selectively used quantitative analyses. For Catastrophic and Hazardous or Severe-Major Failure Conditions, a thorough safety assessment is necessary. The applicant should obtain early concurrence of the certification authority on the choice of an acceptable means of compliance.
(E) If an analysis is used as an acceptable means of showing compliance, it should typically include the following information:
(1 2 3 ) A statement of the functions, boundaries, and interfaces of the system.
( ) A list of the parts and equipment of which the system is comprised, including their performance specifications or design standards and development assurance levels if applicable. This list may reference other documents (e.g., Technical Standard Orders (TSOs), manufacturer's or military specifications, etc.).
( ) The conclusions, including a statement of the Failure Conditions and their classifications and probabilities (expressed qualitatively or quantitatively, as appropriate) that show compliance with the requirements of § 29.1309.
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- AC 29-2C, Chg 4 5/1/2014
(4) A description that establishes correctness and completeness and traces the work leading to the conclusions. This description should include the basis for the classification of each Failure Condition (e.g., analysis or ground, flight, or simulator tests). It should also include a description of precautions taken against common-cause failures, provide any data such as component failure rates and their sources and applicability, support any assumptions made, and identify any required flight crew or ground crew actions, including any Candidate Certification Maintenance Requirements (CCMRs) which are also referred to as Certification Check Requirements (CCRs).
Certification Maintenance Requirements (CMRs) are not the preferred method for showing compliance to the rules; therefore, they must be reviewed by the FAA/AUTHORITY for approval.
(F) Certification Maintenance Requirements (CMRs) may be needed to help show compliance with § 29.1309 for significant latent failures. Rational methods, which usually involve quantitative analyses or relevant service experience data, should be used to determine CMR intervals. These intervals should have reasonable tolerances so that CMRs can be performed concurrently with other maintenance, inspection, or check procedures not required by design for compliance with § 29.1309.
Such tolerances are acceptable because the uncertainty described in the “Note” below is accounted for as discussed in this section. If CMRs are used, they and their intervals and tolerances, and any post-certification changes, or procedures provided in the type design for a rotorcraft owner or operator to make such changes, should be approved by the FAA/AUTHORITY having cognizance over the type design that relates to the system and its installation.
Note: It is recognized that, for various reasons, component failure rate data are not precise enough to enable accurate estimates of the probabilities of Failure Conditions. This results in some degree of uncertainty, as indicated by the expression “on the order of” in the descriptions of the quantitative probability terms that are provided in paragraph b.(1) above. When calculating the probability of each Failure Condition, this uncertainty should be accounted for in a way that does not compromise safety.
(G) Assess the analyses and conclusions of multiple safety assessments to ensure compliance with the requirements for all rotorcraft level Failure Conditions.
(H) Prepare compliance statements, maintenance requirements, and flight manual requirements.
(I) When assessing the acceptability of a failure condition using a quantitative analysis, the given numerical range should normally be interpreted to be the allowable risk for an hour of flight time based on a flight of mean duration for the rotorcraft type. However, when assessing a function which is used only at a specific time during a flight, the probability of the failure condition should be calculated for the specific period and expressed as the risk for the flight condition, takeoff, landing, etc., as appropriate. This is only true for those systems that cannot fail undetected when they Page F - 22
are not being used. The probability of failure for those systems must be calculated from that value based on the extended period of not only the active operation, but also the exposure time for the latent failure.
Note: If a quantitative analysis is used to help show compliance with the regulations for equipment or systems, which are installed and only required for a specific operating condition for which the rotorcraft is thereby approved, credit is usually not allowed due to the fact that the operating condition does not always exist. If an applicant does request that such credit be allowed, they must obtain the approval of the FAA/AUTHORITY.
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AC 29-2C, Chg 4 5/1/2014 Certification Requirements Maintenance Requirements Flight Manual Requirements Rotorcraft Level Functional Hazard Assessment Assessment of Rotorcraft Level System Safety Assessment System Safety System Level FHA’s Assessments Analyses FIGURE AC 29.1309-4: Safety Assessment Process Overview Page F - 24
Note: FHA may be based on a design and installation appraisal for these systems.
Conduct Functional Hazard Assessment Is There a Safety Effect? No Yes Yes No Yes No Verify by Design and Installation Appraisal.
Is the Failure Condition Minor?
Is the System and Installation Similar to a Previous Design?
Develop Similarity Analysis and Obtain Approval from FAA/Authority * Conduct Qualitative or Quantitative Analysis, or both.
* Note: Catastrophic and Hazardous/ Severe-major failure conditions will likely require both qualitative and quantitative analysis, depending on system complexity.
FIGURE AC 29.1309-5: Depth of Analysis Flowchart Page F - 25
AC 29-2C, Chg 4 5/1/2014
(5) Assessment Methods Design Appraisal Installation Appraisal . 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 analys is are provided below and in Document ARP 4761.
(i) . 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 installation. 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. Document 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 is not practically feasible with the present state of the art. In that context, a 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) careful choice of 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 are (1) military aircraft component failure or fault reports or handbooks, such as MIL-HDBK-217C; (2) operator or manufacturer component malfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.
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(iv) Fault Tree or Dependence Diagram Analysis Markov Analysis Common Cause Analysis (CCA) . Structured, deductive, topdown analyses which are used to identify the conditions, failures, and events that would cause each defined Failure Condition. They are 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 nonoperational. 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 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 which 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 faults or malfunctions. 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 Common Cause Analysis is sub-divided into three areas of study:
(A) Zonal Safety Analysis (ZSA). This analysis has the objective of ensuring that the 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 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.
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AC 29-2C, Chg 4 5/1/2014 (B) Particular Risk Analysis. Particular risks are defined as those events or influences that are outside the systems concerned. Examples are fire, leaking fluids, bird strike, tire burst, high intensity radiated fields exposure, lightning, uncontained failure of high energy rotating machines, etc. 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 malfunctions could occur in the same flight for all redundant subparts of a system with a Catastrophic or Hazardous or Severe-Major Failure 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 or Severe-Major. 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 is used after the design is complete to determine if the FTA goals have been met.
(6) Documentation Software Airborne Electronic Hardware (AEH). For airborne complex custom micro- . All laboratory, ground and flight tests, and failure analyses, must be documented in sufficient detail to show compliance with § 29.1309 and 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 Document DO-178C, “Software Considerations in Airborne Systems and Equipment Certification,” dated December 13, 2011, is the latest standard and is recommended to be used for qualification and subsequent approval of airborne software. See AC 20-115C for guidance on using DO-178C and earlier standards.
(8) coded electronic components assessed as level A, B, or C in the safety assessment, RTCA Document 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 the guidance to show compliance to 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 through to the LRU and systems level.
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(9) Environmental Qualification.
(i) Laboratory Tests.
(A) Environmental Standards. In order to assure 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 it 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 the Radio Technical Commission for Aeronautics (RTCA) document DO-160 is an acceptable environmental standard for laboratory qualification of aircraft equipment.
(B) Adverse environmental variables for all types of required and critical equipment include, but are not necessarily limited to temperature, humidity, vibration, shock, altitude, overpressure, and power source transients.
(C) For electrical and electronic equipment, adverse environmental variables include all of (b) above plus overvoltage and undervoltage. Electronic equipment should also be tested for electromagnetic interference (EMI). These tests should include both emission and susceptibility evaluations of both conducted and radiated EMI.
(D) 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. If another standard is used that is at least as good as 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 assured 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 assured 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, also 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;
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AC 29-2C, Chg 4 5/1/2014 and (c) Any other cases 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.
(A) Temperature Tests.
( ) Temperature tests may be accomplished by instrumenting the installed equipmen 1 t 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 assure 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 to be at least 100° F for powerplant-mounted accessories such as starter generators, vacuum pumps, etc. 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. 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.
(2 3 4 ) 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 assure 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 will illustrate 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.
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TORH = The high reference outside air temperature. It varies w ith 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:
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:
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