Table for Failure Condition Categories and Probability Definitions Effect on rotorcraft No effect on Slight Significant Large reduction in Loss of rotorcraft operational reduction in reduction in functional capabilities or capabilities or functional functional safety margins safety capabilities or safety margins safety margin capabilities or (Note 4) Effect on occupants Inconvenience Physical Physical distress, Serious or fatal injury to Multiple Fatalities excluding flight crew discomfort possibly including a passenger or a cabin injuries crew member (Note 2) Effect on flight crew No effect on flight Slight increase Physical Physical distress or Fatalities or incapacitation crew in workload discomfort or a excessive workload that involves significant impairs ability to perform crew actions increase in tasks accurately or well within workload or in completely crew conditions capabilities impairing crew such as routine efficiency flight plan changes DO-178C Software D C B Level (Note 3) Failure Condition Hazardous or Severe- Category No Effect Minor Major Major Catastrophic Qualitative Reasonably Extremely Remote Extremely Probability Frequent Probable Remote Improbable Quantitative No probability 103 (Note 1) 10-5 10-7 10-9 Probability requirement Note 1: A numerical probability range is provided here as reference. The applicant is not required to perform a quantitative analysis, or substantiate by such an analysis, that this numerical criterion has been met for Minor Failure Conditions.
Note 2: This is true if it can be shown that the given failure condition can be contained to a fatal injury of one occupant only.
Note 3: AC 20-174, Development of Civil Aircraft and Systems, provides guidance on using SAE ARP4754A as an acceptable method for establishing a development assurance process.
Note 4: Hazardous Failure Conditions can include events that are manageable by the crew by use of proper procedures which, if not implemented correctly or in a timely manner, may result in a Catastrophic event.
FIGURE AC 27.1309-2 Failure Condition Categories and Probability Definitions Page F - 30.6
(4) Safety Assessment Process Overview.
(i) When showing compliance with § 27.1309, the considerations covered in this guidance should be addressed in a methodical and systematic manner. This section of the AC is provided primarily for the use by applicants who are not familiar with the various methods and procedures generally used in the industry to conduct safety assessments. This section and Figures AC 27.1309-4 and AC 27.1309-5 are not intended to be certification checklists, and they do not include all of the information provided in this guidance. This section contains one method, but not the only method, for showing compliance. The safety assessment process is a structured method of general applicability for showing compliance with the regulation. Other methodologies may be used to show compliance, but the safety assessment process is the only structured process that is defined, and these defined objectives would be the logical criteria applied to any other means of showing compliance. More detailed guidance can be found in SAE ARP4761 on how to perform a safety assessment. SAE ARP4754A, discussed in AC 20-174, includes additional information on how the safety assessment process relates to the system development process.
(ii) The safety assessment process contains several parts that may be necessary as a whole or in part, depending on the criticality or complexity of the system under consideration. The rigor of assessment and analysis performed is also dependent on the system criticality or complexity. At the extremes, some systems may be simple enough such that the entire safety assessment can be performed by observation and compliance shown by a simple statement. Complex and higher criticality systems may require application of all of the safety assessment elements to show compliance. Many states of varying complexity exist between these extremes that may use less than the entire safety assessment process but require more than a simple statement to show compliance. For purposes of the FHA, an autorotation is not considered continued sate flight and landing.
(ill) Typical 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) Page F - 30.7
6/23/23 AC 27-1B, Chg 9 (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 applicantengineering organizations, such as systems, structures, propulsion, and flight test, should be involved in this process. This identification and classification may be done by conducting an 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 establishing the hazard classification of a function, consider for example, the known loss of the function, the erroneous behavior or misleading information provided by the function, and degradation of the function as possible hazards. The hazards are classified using the five failure condition categories found in Figure AC 27.1309-2. The identified Failure Conditions are then 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). SAE ARP4761 provides guidance for determining when and under what conditions multiple failures should be considered.
(C) 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.
(D) Choose the means to be used to determine compliance with § 27.1309. 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 the system is complex (see Figure AC 27.1309-5). 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 Conditions, a thorough safety assessment is necessary. The applicant should obtain early concurrence from the FAA on the choice of an acceptable means of compliance.
(E) Include 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 with the rules; therefore, they must be reviewed by the FAA for approval. Any CMRs must be included in the airworthiness Page F - 30.8
6/23/23 AC 27-1B, Chg 9 limitations section of the instructions for continued airworthiness in the maintenance manual, as required by § 27.1529 and Appendix A to Part 27.
Note: CMRs may be needed to help show compliance with § 27.1309 for significant latent failures. Rational methods, which usually involve quantitative analyses or relevant service experience data, should be used to determine CMR intervals. It is recognized that, for various reasons, component failure rate data is not precise enough to accurately estimate the probabilities of Failure Conditions. This results in some degree of uncertainty, as indicated by the expression "on the order of" in the definition of the quantitative probability terms in paragraph c(1) above. 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 § 27.1309. Any tolerance ability should be built into the interval that is included in the airworthiness limitation section of the maintenance manual.
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, must be approved by the FAA office having cognizance over the type design that relates to the system and its installation, as required by §§ 21.31,
21.95, and 21.97.
(F) 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 that 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 are being used. The probability of failure for those systems should 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 that 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.
Page F - 30.9
-Certifaation Maintennce Repurements (equi ementa Fliniri Menual Requirmments Assessmeni of Rola call Hazard Asseramunil Leven System Safety Assessment Systom Level FMAs Bystem Safely Assesamento Amateu FIGURE AC 27.1309-4: Safety Assessment Process Overview Note: FHA may be based on a design and installation appraisal for these systems.
Page F - 30.10
Conduct Functional Hazard Assessment is there a Safety Effect?
Fattire Condition Is the Venfy by Der ard Installatice Appra al Minori & the System Develop Similanty Anaboe and installation similar to and Obtam Approra freen Previous Design f FAA *Conduct Qualite re ce Quantitative Anale or botl NOTE Castrogise fitre oration will likely requare hotti qualitave anti -o compinar.
FIGURE AC 27.1309-5: Depth of Analysis Flowchart Page F - 30.11
(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 ARP4761.
(i) Design Appraisal. This is a qualitative appraisal of the integrity and safety of the system design.
(ii) Installation Appraisal. 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. SAE ARP4761 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 § 27.1309. 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 or manufacturer component malfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.
Page F - 30.12
AC 27-1B, Chg 9 6/23/23 (Iv) Fault Tree or Dependence Diagram Analysis. This is a structured, deductive, top-down analysis that is used to identity the conditions, failures, and events that would cause each defined Failure Condition. It is a graphical method 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) Markov Analysis. 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) Common Cause Analysis (CCA). 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 tunctions 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 _ Page F - 30.13
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 § 27.1309 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) Software. RTCA DO-178C, Software Considerations in Airborne Systems and Equipment 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 microcode 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 § 27.1309. Although RTCA DO-254 applies specifically to complex custom micro-coded components, Page F - 30.14
applicants are highly encouraged to apply DO-254 to other hardware components up to the LRU and systems level.
(9) Environmental Qualification.
(1) 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 27.863 of this AC provides further guidance on safety from explosion. Another standard that is equivalent to the latest revision of DO-160 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 - 30.15
(A) Temperature Tests.
1. 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 27.1351 of this AC.) Section 27.1043(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 §§ 27.1309 and 27.1581(a)(2). Section 27.1043(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. The § 27.1043(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.
3. 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.
4. The following illustrates an acceptable high temperature evaluation method:
TAT MAx = Maximum outside air temperature at which temperature tests are conducted.
TAx = Maximum temperature to which the installed equipment has been tested in the laboratory.
Page F - 30.16
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 § 27.1043(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 = 275° 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 15,000 274°F 42°F 20,000 270°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., ToR = 100° F - 10,000 ft. × 3.6°/1,000 ft. = 64° F @ 15,000 ft., ToRH = 100° F - 15,000 ft. × 3.6°/1,000 ft. = 46° F @ 20,000 ft., ToR = 100° F - 20,000 ft. x 3.6°/1,000 ft. = 28° F Then at sea level:
Page F - 30.17