ED Decision 2003/15/RM
(a) The rotorcraft may have no design features or details that experience has shown to be hazardous or unreliable.
(b) The suitability of each questionable design detail and part must be established by tests.
ED Decision 2003/15/RM
(a) Critical part - A critical part is 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.
(b) If the type design includes critical parts, a critical parts list shall be established. Procedures shall be established to define the critical design characteristics, identify processes that affect those characteristics, and identify the design change and process change controls necessary for showing compliance with the quality assurance requirements of Part-21.
ED Decision 2003/15/RM
The suitability and durability of materials used for parts, the failure of which could adversely affect safety, must:
(a) Be established on the basis of experience or tests;
(b) Meet approved specifications that ensure their having the strength and other properties assumed in the design data; and
(c) Take into account the effects of environmental conditions, such as temperature and humidity, expected in service.
ED Decision 2003/15/RM
(a) The methods of fabrication used must produce consistently sound structures. If a fabrication process (such as gluing, spot welding, or heat-treating) requires close control to reach this objective, the process must be performed according to an approved process specification.
(b) Each new aircraft fabrication method must be substantiated by a test program.
ED Decision 2003/15/RM
(a) Each removable bolt, screw, nut, pin, or other fastener whose loss could jeopardise the safe operation of the rotorcraft must incorporate two separate locking devices. The fastener and its locking devices may not be adversely affected by the environmental conditions associated with the particular installation.
(b) No self-locking nut may be used on any bolt subject to rotation in operation unless a non-friction locking device is used in addition to the self-locking device.
ED Decision 2023/001/R
This AMC supplements FAA AC 27-1B, § AC 27.607 and should be used in conjunction with that AC when demonstrating compliance with CS 27.601, CS 27.602, CS 27.603 and CS 27.607.
(a) Explanation
Designers should consistently take into account the limitations of the standards, including the applicable fastener manufacturing processes and quality controls, to ensure that when a standard part or qualified standard part is selected, its properties and associated level of reliability will meet the applicable certification requirements for the design.
The intent of this AMC is to give further guidance to the design approval holders (DAHs) and applicants for design approvals to help ensure that appropriate measures are considered for initial certification, including associated continued airworthiness aspects, to minimise the risk that the use of standard fasteners might compromise the intended level of safety.
(b) Definitions
(1) Standard fastener: a fastener that is a standard part. Fasteners (nuts and bolts) being produced according to a certain standard which is not directly approved by the Agency. They fall within the category of standard parts as defined in point 21.A.303(c) of Annex I (Part 21) to Commission Regulation (EU) No 748/2012.
(2) Qualified standard fastener: a standard fastener that requires additional verification of compliance with specification and/or control of their source, by methods defined by the DAH.
(3) Critical installation: a structural/mechanical assembly which may include fasteners the failure of which (single or multiple due to common cause) is classified as hazardous or catastrophic.
(c) Procedures
Failures of standard fasteners may have severe consequences at the aircraft level when used in critical installations.
Once demonstrated, conformance to a standard provides a certain level of reliability under known loading and environmental conditions. The reliability of a standard part or any other part specified in the design needs to be assessed and shown to be compatible with the design objectives to be met. Designers should take care to ensure that they select appropriate fasteners to meet the certification objectives for continued function and reliability, taking into account the limitations of the applicable standards including the associated manufacturing processes and applicable quality controls.
This AMC is therefore addressed to DAHs, to provide them with guidance on appropriate actions to ensure appropriate utilisation of standard fasteners in their designs, to help them to instruct production organisations and maintenance organisations as necessary to ensure continued airworthiness, and to provide means by which unsafe conditions related to the use in design of standard fasteners can be prevented.
In order to reduce the risk of critical installations failing, through the inadvertent use of defective standard fasteners or due to the inappropriate selection of standards, the Agency recommends that all applicants for type certificates and design changes perform a design review to ensure that the risk posed by the use of standard parts is mitigated by:
(1) ensuring that fasteners (nuts and bolts) used in the design will meet the certification requirements, taking into account any limitations of the selected standards, the associated fastener manufacturing processes and quality controls, and relevant service experience;
[Note: The degree to which the standard ensures relevant characteristics such as locking functions, static strength and fatigue strength should be evaluated as far as is necessary based on the criticality of the intended use and operating environment of the parts. Consideration should be given to stress levels arising from manufacture, installation requirements, external loading and temperature effects. Particular attention should be paid to standard parts that utilise high-strength alloys in combination with plating or other processes that may increase the risk of hydrogen embrittlement or deformation processes that are not closely specified.]
(2) ensuring that the design standard met and associated procedures followed for the production of the aircraft are maintained throughout the operational life of the aircraft, e.g. through the use of the ICA controlling maintenance of critical installations;
(3) creating, when standard fasteners (nuts and bolts) are selected, a list of critical installations where only qualified standard fasteners (nuts and bolts) may be used. Redundancy of fasteners alone may not negate the need to qualify the fasteners as all the fasteners on a joint could originate from a common defective batch. Similarly, required double locking functions on fasteners may also need consideration of qualified standard fasteners to ensure that the fail-safe design philosophy is maintained when common cause failure of both locking functions is possible;
(4) defining how the standard fastener is qualified wherever necessary;
(5) clearly defining any necessary additional conformity checks as part of the type design standard, specifying requirements for approved suppliers and any other criteria necessary for acceptance, storage and installation of standard fasteners that are appropriate for use in the design;
(6) ensuring through maintenance instructions that qualified standard fasteners are only replaced by other qualified standard fasteners; and
(7) considering introducing a DAH part numbering system for qualified standard fasteners, at which point they would become aviation parts. (Note: If such part numbering is implemented and further part marking is not feasible due to the part’s size or for other reasons, other means such as regular appropriate batch controls should be established, and documentation provided according to point 21.A.804(b) of Part 21.)
In addition, DAHs are reminded that certain existing Certification Specifications and regulations specifically address critical parts. Typically standard parts are not appropriate for use as critical parts. All critical parts are subject to a critical parts plan that controls their critical characteristics during production and service.
[Amdt 27/10]
CS 27.609 Protection of structure
ED Decision 2003/15/RM
Each part of the structure must:
(a) Be suitably protected against deterioration or loss of strength in service due to any cause, including:
(1) Weathering;
(2) Corrosion; and
(3) Abrasion; and
(b) Have provisions for ventilation and drainage where necessary to prevent the accumulation of corrosive, flammable, or noxious fluids.
CS 27.610 Lightning and static electricity protection
ED Decision 2016/024/R
(a) The rotorcraft must be protected against catastrophic effects from lightning.
(b) For metallic components, compliance with sub-paragraph (a) may be shown by:
(1) Electrically bonding the components properly to the airframe; or
(2) Designing the components so that a strike will not endanger the rotorcraft.
(c) For non-metallic components, compliance with sub-paragraph (a) may be shown by:
(1) Designing the components to minimise the effect of a strike; or
(2) Incorporating acceptable means of diverting the resulting electrical current so as not to endanger the rotorcraft.
(d) The electrical bonding and protection against lightning and static electricity must:
(1) Minimise the accumulation of electrostatic charge;
(2) Minimise the risk of electric shock to crew, passengers, and service and maintenance personnel using normal precautions;
(3) Provide an electrical return path, under both normal and fault conditions, on rotorcraft having grounded electrical systems; and
(4) Reduce to an acceptable level the effects of static electricity on the functioning of essential electrical and electronic equipment.
[Amdt 27/4]
AMC1 27.610 Lightning and static electricity protection
ED Decision 2023/001/R
(a) Purpose
This AMC provides an acceptable means of compliance for evaluation of rotorcraft components after lightning strike.
(b) Related Certification Specifications
CS 27.610 ‘Lightning and static electricity protection’
CS 27.571 ‘Fatigue evaluation of flight structure’
CS 27.573 ‘Damage tolerance and fatigue evaluation of composite structures’
CS 27.1529 ‘Instructions for Continued Airworthiness’
(c) Explanation
CS 27.610 requires the protection of rotorcraft structural components, propulsion system, gearboxes, mechanical and hydraulic control systems from lightning effects that could result in catastrophic failures.
However, damage, failure or departure of any rotorcraft component which could endanger the rotorcraft or its occupants should be part of the evaluation.
This AMC provides detailed guidance on damage tolerance evaluation, including residual strength criteria after lightning strike to ensure continuous safe flight and landing.
Each part, the failure of which implies potential catastrophic consequences and that is exposed to damage under lightning conditions, should be subject to further evaluation which includes:
(1) the nature and extent of the lightning damage (threat assessment, damage detectability, etc.);
(2) the demonstration of the functionality of the affected part up to detection;
(3) a static residual strength capability demonstration supported by analysis and/or test;
(4) when found necessary, a fatigue evaluation of a part with lightning damage for the demonstration of the exposure time before detection.
The airworthiness instruction requested after lightning strike (flight manual and maintenance instructions, etc.) should be consistent with the functional, static and fatigue evaluation of the damage consequences (considered to be a partial failure).
A similar approach should be considered for non-metallic components (for composite, see the AMC 20-29 (11c) guidance).
The above approach is also considered to be applicable for parts departure which could preclude continued safe flight and landing.
For non-structural components (e.g. radomes, panels), only static residual strength is requested for part detachment which could preclude continued safe flight and landing.
[Amdt 27/10]
CS 27.611 Inspection provisions
ED Decision 2003/15/RM
There must be means to allow the close examination of each part that requires:
(a) Recurring inspection;
(b) Adjustment for proper alignment and functioning; or
(c) Lubrication.
CS 27.613 Material strength properties and design values
ED Decision 2003/15/RM
(a) Material strength properties must be based on enough tests of material meeting specifications to establish design values on a statistical basis.
(b) Design values must be chosen to minimise the probability of structural failure due to material variability. Except as provided in sub-paragraphs (d) and (e), compliance with this paragraph must be shown by selecting design values that assure material strength with the following probability:
(1) Where applied loads are eventually distributed through a single member within an assembly, the failure of which would result in loss of structural integrity of the component, 99% probability with 95% confidence; and
(2) For redundant structure, those in which the failure of individual elements would result in applied loads being safely distributed to other load carrying members, 90% probability with 95% confidence.
(c) The strength, detail design, and fabrication of the structure must minimise the probability of disastrous fatigue failure, particularly at points of stress concentration.
(d) Material specifications must be those contained in documents accepted by the Agency.
(e) Other design values may be used if a selection of the material is made in which a specimen of each individual item is tested before use and it is determined that the actual strength properties of that particular item will equal or exceed those used in design.
AMC1 27.613 Material strength properties and design values
ED Decision 2023/001/R
COMPOSITE SANDWICH PANEL
(a) Qualification of the manufacturing process
The conditions outlined in the guidance standard AC 21-26, ‘Quality Control for the Manufacture of Composite Materials’ are considered to be relevant to composite sandwich PSE structure.
The qualification is intended to demonstrate that the combination of material, tooling, equipment, procedures, and other controls, making up the process, will produce representative parts having consistent material properties that conform to design requirements.
As part of the process qualification, destructive and non-destructive inspection (NDI) should be conducted to determine conformity to specified design requirements and check the suitability of the resulting product by assessing features such as:
— uniformity of the adhesive fillets between honeycomb core cell wall and skin; in particular, the process should ensure that on both faces of the honeycomb core a regularly shaped fillet (meniscus) be established;
— absence of ‘telegraphing’ effects and waviness on the skins of the sandwich panel;
— distortion of the core cells — this defect could be particularly critical for highly curved panels unless suitable precautions are taken during fabrication (e.g. core thermal performing);
— presence in the adhesive of unacceptable levels of porosity or humidity;
— disbonds between core and cells; and
— weak bonds.
(b) Material strength and determination of design allowables
The strength properties of the sandwich panels should be established in order to ensure that the probability of structural failure due to material and process variability is minimised.
Because of the peculiarity of the sandwich panel construction, the material properties should be established on a specimen that is fully representative of the panel construction in terms of skin, core material and curing cycle.
Design features such as transition zones from solid laminate to core/skin should be also tested with a representative specimen for determination of strength properties.
It is expected that at least the following static allowables be established according to the statistics required in CS 27.613:
— Adhesive shear strength;
— Shear core strength (ribbon and transverse direction);
— Core compression strength;
— Flatwise strength;
— Flexural strength;
— Compressive strength; and
— Bearing strength (for a specimen representative of all the panel areas where fasteners are installed and subject to significant bearing stresses).
In determining the above properties, the effect due to humidity uptake, highest and lowest temperature expected in service, manufacturing defects up to limit of acceptability and allowable in-service damage defined in maintenance documents, if any, should be considered. For PSEs, impact damages should also be assessed in accordance with CS 27.573.
The validity of the engineering formula used to establish analytical design allowables should be always verified by dedicated experimental activity in order to assess the effects of the manufacturing process (e.g. curing pressure which is normally limited to the crush core strength) and environmental conditions on the allowables predicted by these formulas.
(c) Damage tolerance and residual strength
(1) Threat survey and damage modes
Further to good processing, and when meeting the damage tolerance and fatigue evaluation of composite rotorcraft structures requirements of CS 27.573, the applicant should clearly demonstrate that a robust structure has been produced by showing that:
— a thorough damage threat survey has been completed which identifies and defines all threats, including impacts, heat, moisture, etc. and the potential for interaction of these threats is addressed;
— all damage modes have been identified for the configuration when subject to all likely threats, paying particular attention to all likely damage modes which might not be readily detected.
For impact threats, this requires testing throughout the threat impact energy ranges up to a readily detectable damage using a range of appropriate impactor geometries, including blunt impactors up to 4 inches diameter(1), and a range of impactor stiffnesses, e.g. for hail threat damage (if appropriate), such that all competing damage modes can be identified. Representative boundary conditions should be used in the substantiating test campaigns; and
— all potentially undetectable damage modes (not only disbonds and weak bonds) have been simulated in testing (up to appropriate dimensions such that detection becomes possible, and the dimension of such damage has been quantified such that ultimate load (UL) can be maintained up to this level). The possibility of interaction between threats, e.g. impact and heat, should be considered in the simulation and substantiation process.
Note: Witness structures can be used in service, provided that a consistent and conservative correlation can be demonstrated to exist between the witness indications on the witness structure and the damage (all likely modes and extents) considered in the critical structure.
The recommendations for threat assessment and blunt impact evaluation are also addressed in AC 27.573.
(1) An alternative impactor diameter may be proposed by the applicant, based on the results of the damage threat survey.
(2) Residual strength after extensive damage or degradation
The part should be sized to sustain the required residual strength, in accordance with CS 27.573(d)(4)(ii)(B), with extensive damage or degradation of the most critical skin to core bond between available arrestment features. Such damage or degradation should be readily detectable to assure damage tolerance for bond failures which experience has shown not to be extremely improbable.
It is also expected that relevant fatigue testing at specimen level, representative of a design point (e.g. fastened joint) and typical panel configuration, be performed in order to assess the effects of:
— material/manufacturing process variability;
— environmental condition;
— allowables manufacturing defects; and
— impact damages.
(d) Instructions for continued airworthiness (ICA)
The ICA include clear instructions to inspect(2) (and repair), both internally and externally:
— all load paths, e.g. up to load transfer fittings, joints, and other significant changes in stiffness and section, for damage following an overload event, e.g. impact, heavy landing, excessive gust, etc.;
— all structure regularly exposed to extreme temperatures, e.g. local to engine outlets for aircraft used extensively in hot climates, etc. Although inspections intervals should have been justified according to the level of detectability and residual strength capability during certification substantiation based upon a damage threat survey, experience has indicated that the potential for interaction between heat and damage can be problematic.
(2) paying particular attention to:
— repaired structures; and
— any existing, and potentially related, ICA, e.g. existing ADs, etc.
[Amdt 27/10]
ED Decision 2003/15/RM
(a) The special factors prescribed in CS 27.621 to 27.625 apply to each part of the structure whose strength is:
(1) Uncertain;
(2) Likely to deteriorate in service before normal replacement; or
(3) Subject to appreciable variability due to:
(i) Uncertainties in manufacturing processes; or
(ii) Uncertainties in inspection methods.
(b) For each part to which CS 27.621 to 27.625 apply, the factor of safety prescribed in CS 27.303 must be multiplied by a special factor equal to:
(1) The applicable special factors prescribed in CS 27.621 to 27.625; or
(2) Any other factor great enough to ensure that the probability of the part being understrength because of the uncertainties specified in sub-paragraph (a) is extremely remote.
ED Decision 2003/15/RM
(a) General. The factors, tests, and inspections specified in sub-paragraphs (b) and (c) must be applied in addition to those necessary to establish foundry quality control. The inspections must meet approved specifications. Sub-paragraphs (c) and (d) apply to structural castings except castings that are pressure tested as parts of hydraulic or other fluid systems and do not support structural loads.
(b) Bearing stresses and surfaces. The casting factors specified in sub-paragraphs (c) and (d):
(1) Need not exceed 1.25 with respect to bearing stresses regardless of the method of inspection used; and
(2) Need not be used with respect to the bearing surfaces of a part whose bearing factor is larger than the applicable casting factor.
(c) Critical castings. For each casting whose failure would preclude continued safe flight and landing of the rotorcraft or result in serious injury to any occupant, the following apply:
(1) Each critical casting must –
(i) Have a casting factor of not less than 1.25; and
(ii) Receive 100% inspection by visual, radiographic, and magnetic particle (for ferromagnetic materials) or penetrant (for non-ferromagnetic materials) inspection methods or approved equivalent inspection methods.
(2) For each critical casting with a casting factor less than 1.50, three sample castings must be static tested and shown to meet –
(i) The strength requirements of CS 27.305 at an ultimate load corresponding to a casting factor of 1.25; and
(ii) The deformation requirements of CS 27.305 at a load of 1.15 times the limit load.
(d) Non-critical castings. For each casting other than those specified in sub-paragraph (c), the following apply:
(1) Except as provided in sub-paragraphs (d)(2) and (3), the casting factors and corresponding inspections must meet the following table:
|
Casting factor |
Inspection |
|
2.0 or greater......... |
100% visual |
|
Less than 2.0 greater than 1.5 |
100% visual and magnetic particle (ferromagnetic materials), penetrant (non-ferromagnetic materials), or approved equivalent inspection methods. |
|
1.25 through 1.50...... |
100% visual, and magnetic particle (ferromagnetic materials), penetrant non-ferromagnetic materials), and radiographic or approved equivalent inspection methods |
(2) The percentage of castings inspected by nonvisual methods may be reduced below that specified in sub-paragraph (d)(1) when an approved quality control procedure is established.
(3) For castings procured to a specification that guarantees the mechanical properties of the material in the casting and provides for demonstration of these properties by test of coupons cut from the castings on a sampling basis:
(i) A casting factor of 1.0 may be used; and
(ii) The castings must be inspected as provided in sub-paragraph (d)(1) for casting factors of 1.25 to 1.50 and tested under sub-paragraph (c)(2).
ED Decision 2003/15/RM
(a) Except as provided in sub-paragraph (b), each part that has clearance (free fit), and that is subject to pounding or vibration, must have a bearing factor large enough to provide for the effects of normal relative motion.
(b) No bearing factor need be used on a part for which any larger special factor is prescribed.
ED Decision 2003/15/RM
For each fitting (part or terminal used to join one structural member to another) the following apply:
(a) For each fitting whose strength is not proven by limit and ultimate load tests in which actual stress conditions are simulated in the fitting and surrounding structures, a fitting factor of at least 1.15 must be applied to each part of:
(1) The fitting;
(2) The means of attachment; and
(3) The bearing on the joined members.
(b) No fitting factor need be used:
(1) For joints made under approved practices and based on comprehensive test data (such as continuous joints in metal plating, welded joints, and scarf joints in wood); and
(2) With respect to any bearing surface for which a larger special factor is used.
(c) For each integral fitting, the part must be treated as a fitting up to the point at which the paragraph properties become typical of the member.
(d) Each seat, berth, litter, safety belt, and harness attachment to the structure must be shown by analysis, tests, or both, to be able to withstand the inertia forces prescribed in CS 27.561(b)(3) multiplied by a fitting factor of 1.33.
ED Decision 2003/15/RM
Each aerodynamic surface of the rotorcraft must be free from flutter under each appropriate speed and power condition.
ED Decision 2021/016/R
(See AMC1 27.631)
Rotorcraft with six or more passenger seats must be designed to ensure a safe landing after a strike upon the windshield by a 1.0-kg (2.2-lb) bird when the velocity of the rotorcraft relative to the bird along the flight path of the rotorcraft is equal to VNE or VH ‘True Airspeed’ (TAS), whichever is less, at altitudes up to 2 438 m (8 000 ft). The applicant must demonstrate compliance through tests, or analysis based on tests that are carried out on sufficiently representative structures of similar design.
[Amdt 27/9]
ED Decision 2021/016/R
(a) To demonstrate the remaining capability of the rotorcraft after a single bird strike, the applicant should evaluate the parts of the rotorcraft as follows:
(1) the windshield directly in front of the occupants and its supporting frame should be capable of withstanding a bird strike without penetration; and
(2) any systems and equipment (including their controls) that are essential to ensure a safe landing and are installed near the windshield and its supporting frame should remain operative in case of shock loads resulting from a bird strike.
Note: the capability to withstand multiple bird strikes is only evaluated for engines as specified under CS-E 800 ‘Bird Strike and Ingestion’.
(b) For the demonstration under point (a), the altitude range within which the velocity VH is evaluated should be defined and should not exceed 2 438 m (8 000 ft).
[Amdt 27/9]