Structural Critical Parts
As noted at Section 6.6.2.16, the primary consideration for defining what constitutes a structural critical part should be the certification basis for the aircraft.
Some commonly-used classifications that can apply to structural parts are detailed in Table A.1, along with notes on use of these classifications for identifying structural critical parts under DASR.
In recognition that not all airworthiness codes are equivalent, and some are not explicit on definition of critical parts, DASA provides the following general definition for structural critical parts in DASR AMC 21.A.41:
Any structural part or element where the failure of that part or element could result in a fatality or loss of aircraft. The fatality or loss of aircraft could occur immediately upon failure or subsequently if the failure remained undetected.
A structural part is one that contributes significantly to the carrying of flight, ground, or pressurization loads. For rotorcraft, identification of structural critical parts should include consideration of the rotors, rotor drive systems between the engines and rotor hubs, controls, fuselage, fixed and movable control surfaces, engine and transmission mountings, landing gear, and their related primary attachments.
In the general definition above, DASA considers the consequence of failure (ie fatality or loss of aircraft) to be the minimum acceptable scope for defining structural critical parts. Depending upon the underlying certification basis or design standard(s) more expansive definitions (eg inclusive of US 14CFR §11 / EASA CS 2x.1309 Hazardous consequence or MIL STD 882E Critical consequences) may be used and accepted by DASA.
MTC or MSTC / major change applicants should ensure the definition used to identify structural critical parts is consistent with the aircraft TCB and / or the general definition from DASR AMC 21.A.41.
Table A.1: Common Classifications used for Aircraft Structure
|
Term |
Source |
Classification Definition |
Consequence Definition |
Notes on use under DASR |
|
Principal Structural Element (PSE) |
US 14CFR § / EASA CS 25.571 |
An element that contributes significantly to the carrying of flight, ground, or pressurization loads, and whose integrity is essential in maintaining the overall integrity of the airplane |
Not explicitly defined |
Acceptable basis for identifying structural critical parts under DASR |
|
US 14CFR § / EASA CS 29.571 & 29.573 |
Structural elements that contribute significantly to the carriage of flight or ground loads, the failure of which could result in catastrophic failure of the aircraft |
Catastrophic failure: an event that could prevent continued safe flight and landing [FAA AC 29.571B and FAA AC 29.573] |
||
|
Fatigue Critical Structure |
US 14CFR § / EASA CS 25.571 |
Structure that is susceptible to fatigue cracking, which could contribute to a catastrophic failure |
Not defined |
Subset of PSEs and therefore will be a subset of structural critical parts under DASR |
|
Critical Part |
US 14CFR § / EASA CS 29.602 |
A part, the failure of which could have a catastrophic effect upon the rotorcraft, and for which critical characteristics* have been identified which must be controlled to ensure the required level of integrity |
Catastrophic: the inability to conduct an autorotation to a safe landing, without exceptional piloting skills, assuming a suitable landing surface is available [FAA AC 29.602] |
Applies to all aircraft systems, therefore, may assist with defining structural critical parts under DASR |
|
Fracture Critical (Traceable or Non-Traceable) |
MIL-STD-1530D |
A safety-of-flight structural component |
Failure could cause loss of the aircraft, cause severe injury or death, impair a safety critical function, or cause inadvertent store release. The consequences could occur either immediately upon failure or subsequently if the failure remains undetected. |
Acceptable basis for identifying structural critical parts under DASR |
|
Aviation Critical Safety Item (CSI) |
US DoD Acquisition Regulations: Title 48 Ch. 2 §252.209-7010 |
A part, an assembly, installation equipment, launch equipment, recovery equipment, or support equipment for an aircraft or aviation weapon system if the part, assembly, or equipment contains a characteristic* any failure, malfunction, or absence of which could cause: (i) A catastrophic or critical failure resulting in the loss of, or serious damage to, the aircraft or weapon system; (ii) An unacceptable risk of personal injury or loss of life; or (iii) An uncommanded engine shutdown that jeopardizes safety. |
Inherent in classification definition |
Applies to all components of the air system. Furthermore, in practice only replenishable parts are identified as CSIs, and will typically not include integral structural critical parts. In isolation, the CSI definition may not provide a complete list of structural critical parts under DASR. |
*Critical characteristic (29.602) or characteristic (CSI) refers to any feature throughout the life-cycle, such as any dimension, tolerance, finish, material, assembly, manufacturing or inspection process, or other feature which cannot tolerate variation from the type design and, if non-conforming would cause failure of the part.
ASIP Ongoing Monitoring and Periodic Assessments
Overview. This section provides additional structures-specific guidance in support of Section 6.6.3.
For aircraft structures, the ongoing monitoring and periodic assessment required to ensure continued integrity (DASR 21.A.44(c)) should include:
Ongoing Monitoring: collection and evaluation of relevant service experience data via structural condition monitoring and usage monitoring to identify significant events, exceedances of baseline assumptions, trends or anomalies that require more detailed assessment.
Periodic assessment: analysis of the relevant data to ensure assumptions made during design and certification that could affect structural integrity remain valid for the Defence CRE, and identifying when updates to the MTC (including AwL and operating limitations) or ICA are required.
Structural Condition Monitoring (SCM). The purpose of SCM is to collect and evaluate data that describes the actual structural condition of the fleet and allows comparison of actual degradation / damage with that assumed and predicted during design.
For airworthiness purposes, the SCM system should cover all critical parts and other ‘significant’ structure. Other significant structure should be included as unanticipated degradation or unforeseen failure modes to such structure may warrant re-classification as critical parts. The definition for what constitutes other significant structure will vary by aircraft type. The scope of structure that is monitored by the SCM system should be outlined in the SIMP.
Note that the broader ASIP will ideally be monitoring all aircraft structure in order to maximise cost effectiveness and aircraft availability. The SIMP should therefore articulate, in terms of structural classifications, what structure falls within the remit of the DASR 21.A.44(c) obligation. Separately, and for a broader scope of structure, ASIP data and assessments may also be supporting CAMO RP obligations (refer to Section 6.6.4.17).
As listed at Section 6.6.3.18, relevant data can come from a wide range of sources. The primary sources of structural condition data will be scheduled and unscheduled maintenance and inspections, however, other sources of relevant data such as from health monitoring systems or dedicated inspection of parts with service history should also be captured. The SIMP should outline the sources of data and how these are compiled and evaluated, and specific requirements and responsibilities for data capture. The MTCH may also need to specify certain data capture requirements through ICA.
The data collected should include all structural discrepancies for the scope of structure outlined above, including fatigue degradation, ED, accidental and maintenance-induced damage, and build / assembly non conformances.
The data should include all the information necessary to enable periodic assessment to determine the root-cause of discrepancies and impact on continued structural integrity. At a minimum this information should include:
details of the discrepancy, including damage type, dimensions (length, area, depth, orientation as relevant), and location on part
details of the asset, including tail, part and serial number (as relevant)
when the discrepancy was found, including date, flight hours, other metrics as relevant (eg landings, flight cycles)
how the discrepancy was found, including unscheduled or scheduled maintenance, and inspection technique
disposition (ie use-as-is, repair details or component replacement).
Usage Monitoring (UM). The purpose of UM is to collect and evaluate operational usage and environment data that can be compared to, and used to update, design assumptions. This purpose corresponds to that of the Loads / Environment Spectra Survey (L/ESS) ASIP element described within Section 6.6 of MIL STD 1530D.
The UM system includes the hardware, software, procedures and other tools required to acquire, store and report usage parameters that significantly influence degradation of the aircraft structure. The philosophy and complexity of the system, will vary significantly based on the aircraft type and role, operational variability, certification basis and design philosophy.
For monitoring drivers of fatigue degradation, the UM system should capture sufficient data to allow detailed comparison of actual operational usage to the design or baseline spectrum. This should include all significant sources of repeated loads and other parameters that can impact fatigue of structural critical parts. For most military aircraft types the UM system will need to capture more than just basic usage parameters (eg flight hours, flight cycles, landings, etc). The UM system will typically need to capture parametric time history data22 and / or strain sensor data and have an ability to process this data to allow comparison with the design spectrum and assumptions (often referred to as flight manoeuvre or regime recognition for rotorcraft).
For monitoring drivers of ED, the UM system should capture usage parameters that can be used to describe the operating locations and environment that may influence ED trends for critical parts.
Degradation of structural critical parts, especially fatigue degradation, can be sensitive to small variations in usage or configuration. Therefore, the UM system should be subject to initial validation. Validation should test the entire system (on-board and off-board elements) in an end-to-end fashion (ie from raw / input data through to the processed outputs) and should ensure that the system produces reliable, repeatable and suitably-conservative outputs.
Periodic assessment requires a suitable quantity of valid data, and therefore targets for valid data capture should be established and monitored.
Many UM systems are ‘parametric’-based, which means they capture usage information (ie frequency and duration of manoeuvres and events) but do not directly measure loads. In such systems, prediction of structural loads is accomplished via analytical methods. Other UM systems capture actual loads (ie using strain sensors), however, typically only from a handful of discrete locations. A more comprehensive system for direct loads measurement may be required to support continued airworthiness for certain aircraft types in the following circumstances:
if there is a high degree of uncertainty with regards to the actual in service loads being experienced by structural critical parts
analytical components of the UM system and / or system for tracking life accrual (refer Section 6.6.4) require ongoing validation with actual loads data
to support a major modification or service life extension program.
Paragraph 20 refers to direct loads measurement and when this may be required to support continued airworthiness. The term Operational Loads Monitoring (OLM) is traditionally used to describe a comprehensive suite of strain and other sensor types (eg accelerometers) fitted to a subset of the fleet, along with recording systems and ground-based hardware and software. An OLM system can provide higher-fidelity data and reduce conservatism in ASIP management decisions. The decision to fit an OLM system is usually based on a business case analysis of the capability (cost, availability and safety) benefits of the system.
Note that the L/ESS requirements within MIL-STD-1530D allow for a wide range of potential solutions, from parametric UM systems through to more comprehensive instrumentation suites similar to what has historically been termed Operational Loads Monitoring (OLM) within Defence. The term L/ESS is also sometimes interpreted and applied as an intermittent and sample-based monitoring program, whereas DASA expects that UM for the purposes of DASR 21.A.44(c) is continuous and fleet-wide. Therefore, the MIL STD 1530D L/ESS requirements should be read in conjunction with the guidance in this Annex, and adapted for each aircraft type, when developing detailed UM system requirements.
Periodic ASI Assessments. Periodic ASI assessments should ensure the MTC (including AwL and operating limitations) and ICA remain compliant with the TCB, and therefore that the risk of structural failure remains within that inherent in the TCB. This is primarily achieved by ensuring the assumptions made during design and certification that could affect the integrity of structural critical parts remain valid in light of actual operational usage and condition.
Periodic ASI assessments include both routine assessments and larger semi-routine assessments. For many platforms, routine ASI assessments include fatigue and environmental degradation assessments undertaken every 1-2 years, and semi-routine structural life assessments which are undertaken on an as required basis. However, the type and frequency of assessments will depend upon a number of factors including the platform type, role, age and sustainment support construct, outcomes of previous assessments and stability of operational usage. The type and frequency of periodic ASI assessments should be documented in the SIMP.
The assessment approach will vary between aircraft types, however, some examples that may indicate design assumptions are no longer valid include:
fatigue cracking in an unanticipated location or direction, or at an unanticipated growth rate or time, in critical parts certified under a damage-tolerance philosophy
fatigue cracking in critical parts certified under a safe-life philosophy
environmental, accidental or maintenance-induced damage on critical parts beyond that accounted for in certification, especially when such damage can undermine fatigue crack initiation or growth assumptions
critical parts being routinely rejected / thrown-away during maintenance prior to their mandatory life-limit
evidence of production / build discrepancies in critical parts beyond that accounted for in certification which are symptomatic across the fleet
evidence of widespread fatigue damage (eg multi-element or multi-site damage) in critical parts or other significant structure which has not been accounted for in certification
damage to, or failure of, parts not currently identified as critical in an unanticipated manner that may result in an unsafe condition
fleet or sub-fleet usage severity consistently outside the bounds of the design assumptions (ie baseline spectra) used to develop AwLs
routine exceedance of relevant operational limits, such as gross weight or limit load exceedances (may invalidate strength certification), and airspeed exceedances (may invalidate aeroelasticity / flutter certification).
As defined in DASR AMC 21.A.41, AwLs include not only mandatory life-limit (retirement) and inspection requirements, but any algorithm, equation, factor(s) or other engineering data which must be used to calculate life accrual against the interval (refer Section 6.6.2.24.d.(ii)). Data used to calculate life accrual may be invalidated by in-service changes (eg operational practices or configuration changes). This is especially the case for AwL tracking methods that rely on pre-defined mission severity factors or parametric-based algorithms. Periodic assessment should ensure that these aspects of the AwL, and the system for tracking life accrual more generally, also remain valid.
Dedicated and separate ageing aircraft structural assessments may also be required for ensuring continued airworthiness in certain circumstances. For example, this may include aircraft for which there is:
significant concern around unknown ageing-structure threats
plans to extend the certified service life.
Ageing Aircraft Structural Assessments (AASAs) have proven successful within Defence for capability assurance and successfully managing fleets through to the planned withdrawal date. DASA encourages the conduct of AASAs, but will only require the MTCH to conduct such assessments for continued airworthiness purposes in certain circumstances.
Tracking Structural Life-Limited Components
This section supports Section 6.6.4 and provides additional guidance and examples related to tracking life accrual against structural AwLs. This is often referred to as Individual Aircraft Tracking (IAT) for aircraft structures.
The following guidance and examples are split into the two cases where the authoritative source for the AwL promulgates the interval in either a ‘simple’ or ‘complex’ metric. The processes in the two examples below describe what is often required to ensure that individual aircraft / components do not exceed the underlying basis of AwL intervals, and are therefore considered to be tracking life accrual under DASR M.A.305(d)(4). Note that these are just two common philosophies and there are many other ways life accrual may need to be tracked for structural AwLs.
Simple Metrics. Simple AwL interval metrics such as flight hours, landings or calendar time are easy to measure. However, the system for tracking life accrual must ensure that individual aircraft usage is accounted for and each aircraft / component does not exceed the underlying basis of the AwL.
Consider a critical part AwL that is promulgated in flight hours, but the interval is based on a fatigue life calculation. This is common for rotorcraft critical part AwLs (ie Component Retirement Times (CRTs)). Measurement of flight hour accrual (eg using CAMM2) is straightforward. However, the rate (per flight hour) at which fatigue damage actually accumulates can vary significantly based on the way individual aircraft are operated. Therefore, a process must be established to ensure each aircraft / component does not exceed the underlying baseline fatigue life due to a higher-than-assumed rate of fatigue accrual per flight hour. Ideally, all manoeuvres and parameters that materially influence fatigue damage accrual for the critical part would be recorded. This usage data would then be manipulated into a measure of how fatigue damage is accumulating for each individual aircraft / component compared to the assumptions that underpin the AwL interval33. This process may be performed by the ASIP manager if the requisite design data is available, or ADF usage data could be provided to the OEM to undertake this analysis routinely.
The AwL interval is often calculated based on a ‘worst case’ or ‘severe’ assumed design spectrum, so the severity of actual operations should normally be less than the baseline severity, in which case no action needs to be taken (ie the AwL remains conservative for the aircraft / component). However, if the severity for an aircraft / component is higher than the design assumptions then continuing airworthiness intervention (eg tail / component specific lifing updates) may be required to ensure the aircraft / component does not exceed the underlying basis of the AwL.
Complex Metrics. Complex AwL interval metrics such as Equivalent Flight Hours (EFH), Fatigue Index (FI) or Fatigue Life Expended Index (FLEI) are more complicated to calculate, but will inherently account for variations in individual aircraft usage. However, these metrics often need to be converted back into flight hours for maintenance planning purposes.
Consider a critical part AwL that is promulgated in EFH. The system for tracking life accrual will usually take data (such as parametric data) from the on board system or strain sensor signals and convert this into a cumulative EFH accrual using algorithms, damage equations, etc. The system will therefore contain the current EFH status for each individual aircraft / component based on actual usage. However, for maintenance planning purposes, it may be necessary to project EFH forward and convert into flight hours to establish when the AwL action needs to be performed for each individual aircraft. Whilst this projection and conversion is being performed for maintenance planning purposes, it does impact when the AwL action is scheduled, and therefore is an important step in the overall process.