ED Decision 2016/010/R
(See AMC 25.301)
(a) Strength requirements are specified in terms of limit loads (the maximum loads to be expected in service) and ultimate loads (limit loads multiplied by prescribed factors of safety). Unless otherwise provided, prescribed loads are limit loads.
(b) Unless otherwise provided the specified air, ground, and water loads must be placed in equilibrium with inertia forces, considering each item of mass in the aeroplane. These loads must be distributed to conservatively approximate or closely represent actual conditions. (See AMC No. 1 to CS 25.301(b).) Methods used to determine load intensities and distribution must be validated by flight load measurement unless the methods used for determining those loading conditions are shown to be reliable. (See AMC No. 2 to CS 25.301(b).)
(c) If deflections under load would significantly change the distribution of external or internal loads, this redistribution must be taken into account.
[Amdt 25/1]
[Amdt 25/18]
ED Decision 2005/006/R
The engine and its mounting structure are to be stressed to the loading cases for the aeroplane as a
whole.
[Amdt 25/1]
AMC No. 2 to 25.301(b) Flight Load Validation
ED Decision 2005/006/R
1. PURPOSE
This AMC sets forth an acceptable means, but not the only means, of demonstrating compliance with the provisions of CS-25 related to the validation, by flight load measurements, of the methods used for determination of flight load intensities and distributions, for large aeroplanes.
2. RELATED CERTIFICATIONS SPECIFICATIONS
CS 25.301(b) “Loads”
CS 25.459 “Special Devices”
3. BACKGROUND
(a) CS-25 stipulates a number of load conditions, such as flight loads, ground loads, pressurisation loads, inertia loads and engine/APU loads. CS 25.301 requires methods used to determine load intensities and distributions to be validated by flight load measurements unless the methods used for determining those loading conditions are shown to be reliable. Although this applies to all load conditions of CS-25, the scope of this AMC is limited to flight loads.
(b) The sizing of the structure of the aircraft generally involves a number of steps and requires detailed knowledge of air loads, mass, stiffness, damping, flight control system characteristics, etc. Each of these steps and items may involve its own validation. The scope of this AMC however is limited to validation of methods used for determination of loads intensities and distributions by flight load measurements.
(c) By reference to validation of “methods”, CS 25.301(b) and this AMC are intended to convey a validation of the complete package of elements involved in the accurate representation of loads, including input data and analytical process. The aim is to demonstrate that the complete package delivers reliable or conservative calculated loads for scenarios relevant to CS-25 flight loads requirements.
(d) Some measurements may complement (or sometimes even replace) the results from theoretical methods and models. Some flight loads development methods such as those used to develop buffeting loads have very little theoretical foundation, or are methods based directly on flight loads measurements extrapolated to represent limit conditions.
4. NEED FOR AND EXTENT OF FLIGHT LOAD MEASUREMENTS
4.1. General
(a) The need for and extent of the flight load measurements has to be discussed and agreed between the Agency and Applicant on a case by case basis. Such an assessment should be based on:
(i) a comparison of the design features of the aeroplane under investigation with previously developed (by the Applicant) and approved aeroplanes. New or significantly different design features should be identified and assessed.
(ii) the Applicant’s previous experience in validating load intensities and distributions derived from analytical methods and/or wind tunnel tests. This experience should have been accumulated on previously developed (by the Applicant) and approved types and models of aeroplanes. The validation should have been by a flight load measurement program that was conducted by the Applicant and found acceptable to the Agency for showing compliance.
(iii) the sensitivity to parametric variation and continued applicability of the analytical methods and/or wind tunnel test data.
(b) Products requiring a new type certificate will in general require flight-test validation of flight loads methods unless the Applicant can demonstrate to the Agency that this is unnecessary.
If the configuration under investigation is a similar configuration and size as a previously developed and approved design, the use of analytical methods, such as computational fluid dynamics validated on wind tunnel test results and supported by previous load validation flight test experience, may be sufficient to determine flight loads without further flight test validation.
(c) Applicants who are making a change to a Type Certificated airplane, but who do not have access to the Type certification flight loads substantiation for that airplane, will be required to develop flight loads analyses, as necessary, to substantiate the change.
In general, the loads analyses will require validation and may require flight test loads measurements, as specified in this AMC.
(d) The Applicant is encouraged to submit supporting data or test plans for demonstrating the reliability of the flight loads methods early in the certification planning process.
4.2. New or significantly different design features.
Examples of new or significantly different design features include, but are not limited to:
— Wing mounted versus fuselage mounted engines;
— Two versus three or more engines;
— Low versus high wing;
— Conventional versus T-tail empennage;
— First use of significant sweep;
— Significant expansion of flight envelope;
— Addition of winglets;
— Significant modification of control surface configuration;
— Significant differences in airfoil shape, size (span, area);
— Significant changes in high lift configurations;
— Significant changes in power plant installation/configuration;
— Large change in the size of the aeroplane.
4.3. Other considerations
(a) Notwithstanding the similarity of the aeroplane or previous load validation flight test experience of the Applicant, the local loads on the following elements are typically unreliably predicted and may require a measurement during flight tests:
— Loads on high lift devices;
— Hinge moments on control surfaces;
— Loads on the empennage due to buffeting;
— Loads on any unusual device.
(b) For non-deterministic loading conditions, such as stall buffet, the applicant should compile a sufficient number of applicable flight loads measurements to develop a reliable method to predict the appropriate design load.
5. FLIGHT LOAD MEASUREMENTS
5.1. Measurements.
Flight load measurements (for example, through application of strain gauges, pressure belts, accelerometers) may include:
— Pressures / air loads /net shear, bending and torque on primary aerodynamic surfaces;
— Flight mechanics parameters necessary to correlate the analytical model with flight test results;
— High lift devices loads and positions;
— Primary control surface hinge moments and positions;
— Unsymmetric loads on the empennage (due to roll/yaw manoeuvres and buffeting);
— Local strains or response measurements in cases where load calculations or measurements are indeterminate or unreliable.
5.2. Variation of parameters.
The test points for the flight loads measurements should consider the variation of the main parameters affecting the loads under validation. Examples of these parameters include: load factor, speeds, altitude, aircraft c.g., weight and inertia, power settings (thrust, for wing mounted engines), fuel loading, speed brake settings, flap settings and gear conditions (up/down) within the design limits of the aeroplane. The range of variation of these parameters must be sufficient to allow the extrapolation to the design loads conditions. In general, the flight test conditions need not exceed approximately 80% of limit load.
5.3. Conditions.
In the conduct of flight load measurements, conditions used to obtain flight loads may include:
— Pitch manoeuvres including wind-up turns, pull-ups and push-downs (e.g. for wing and horizontal stabiliser manoeuvring loads);
— Stall entry or buffet onset boundary conditions (e.g. for horizontal stabiliser buffet loads);
— Yaw manoeuvres including rudder inputs and steady sideslips;
— Roll manoeuvres.
Some flight load conditions are difficult to validate by flight load measurements, simply because the required input (e.g. gust velocity) cannot be accurately controlled or generated. Therefore, these type of conditions need not be flight tested. Also, in general, failures, malfunctions or adverse conditions are not subject to flight tests for the purpose of flight loads validation.
5.4. Load alleviation.
When credit has been taken for an active load alleviation function by a particular control system, the effectiveness of this function should be demonstrated as far as practicable by an appropriate flight test program.
6. RESULTS OF FLIGHT LOAD MEASUREMENTS
6.1. Comparison / Correlation.
Flight loads are not directly measured, but are determined through correlation with measured strains, pressures or accelerations. The load intensities and distributions derived from flight testing should be compared with those obtained from analytical methods. The uncertainties in both the flight testing measurements and subsequent correlation should be carefully considered and compared with the inherent assumptions and capabilities of the process used in analytic derivation of flight loads. Since in most cases the flight test points are not the limit design load conditions, new analytical load cases need to be generated to match the actual flight test data points.
6.2. Quality of measurements.
Factors which can affect the uncertainty of flight loads resulting from calibrated strain gauges include the effects of temperature, structural non-linearities, establishment of flight/ground zero reference, and large local loads, such as those resulting from the propulsion system installation, landing gear, flap tracks or actuators. The static or dynamic nature of the loading can also affect both strain gauge and pressure measurements.
6.3. Quality of correlation.
A given correlation can provide a more or less reliable estimate of the actual loading condition depending on the "static" or "flexible dynamic" character of the loading action, or on the presence and level of large local loads. The quality of the achieved correlation depends also on the skills and experience of the Applicant in the choice of strain gauge locations and conduct of the calibration test programme.
Useful guidance on the calibration and selection of strain gauge installations in aircraft structures for flight loads measurements can be found, but not exclusively, in the following references:
1. Skopinski, T.H., William S. Aiken, Jr., and Wilbur B. Huston,
“Calibration of Strain-Gage Installations in Aircraft Structures for Measurement of Flight Loads”, NACA Report 1178, 1954.
2. Sigurd A. Nelson II, “Strain Gage Selection in Loads Equations Using a Genetic Algorithm”, NASA Contractor Report 4597 (NASA-13445), October 1994.
6.4. Outcome of comparison / correlation.
Whatever the degree of correlation obtained, the Applicant is expected to be able to justify the elements of the correlation process, including the effects of extrapolation of the actual test conditions to the design load conditions.
If the correlation is poor, and especially if the analysis underpredicts the loads, then the Applicant should review and assess all of the components of the analysis, rather than applying blanket correction factors.
For example:
(a) If the level of discrepancy varies with the Mach number of the condition, then the Mach corrections need to be evaluated and amended.
(b) If conditions with speed brakes extended show poorer correlation than clean wing, then the speed brake aerodynamic derivatives and/or spanwise distribution need to be evaluated and amended.
[Amdt 25/1]
CS 25.302 Interaction of systems and structures
ED Decision 2005/006/R
For aeroplanes equipped with systems that affect structural performance, either directly or as a result of a failure or malfunction, the influence of these systems and their failure conditions must be taken into account when showing compliance with the requirements of Subparts C and D. Appendix K of CS-25 must be used to evaluate the structural performance of aeroplanes equipped with these systems.
[Amdt 25/1]
ED Decision 2003/2/RM
Unless otherwise specified, a factor of safety of 1·5 must be applied to the prescribed limit load which are considered external loads on the structure. When loading condition is prescribed in terms of ultimate loads, a factor of safety need not be applied unless otherwise specified.
CS 25.305 Strength and deformation
ED Decision 2005/006/R
(a) The structure must be able to support limit loads without detrimental permanent deformation. At any load up to limit loads, the deformation may not interfere with safe operation.
(b) The structure must be able to support ultimate loads without failure for at least 3 seconds. However, when proof of strength is shown by dynamic tests simulating actual load conditions, the 3-second limit does not apply. Static tests conducted to ultimate load must include the ultimate deflections and ultimate deformation induced by the loading. When analytical methods are used to show compliance with the ultimate load strength requirements, it must be shown that –
(1) The effects of deformation are not significant;
(2) The deformations involved are fully accounted for in the analysis; or
(3) The methods and assumptions used are sufficient to cover the effects of these deformations.
(c) Where structural flexibility is such that any rate of load application likely to occur in the operating conditions might produce transient stresses appreciably higher than those corresponding to static loads, the effects of this rate of application must be considered.
(d) Reserved
(e) The aeroplane must be designed to withstand any vibration and buffeting that might occur in any likely operating condition up to VD/MD, including stall and probable inadvertent excursions beyond the boundaries of the buffet onset envelope. This must be shown by analysis, flight tests, or other tests found necessary by the Agency.
(f) Unless shown to be extremely improbable, the aeroplane must be designed to withstand any forced structural vibration resulting from any failure, malfunction or adverse condition in the flight control system. These loads must be treated in accordance with the requirements of CS 25.302.
[Amdt 25/1]
ED Decision 2005/006/R
(See AMC 25.307)
(a) Compliance with the strength and deformation requirements of this Subpart must be shown for each critical loading condition. Structural analysis may be used only if the structure conforms to that for which experience has shown this method to be reliable. In other cases, substantiating tests must be made to load levels that are sufficient to verify structural behaviour up to loads specified in CS 25.305.
(b) Reserved
(c) Reserved
(d) When static or dynamic tests are used to show compliance with the requirements of CS 25.305(b) for flight structures, appropriate material correction factors must be applied to the test results, unless the structure, or part thereof, being tested has features such that a number of elements contribute to the total strength of the structure and the failure of one element results in the redistribution of the load through alternate load paths.
[Amdt 25/1]
ED Decision 2005/006/R
1. Purpose
This AMC establishes methods of compliance with CS 25.307, which specifies the requirements for Proof of Structure.
2. Related Certification Specifications
CS 25.303 “Factor of safety”
CS 25.305 “Strength and deformation”
CS 25.651 “Proof of strength”
3. Definitions
3.1. Detail. A structural element of a more complex structural member (e.g. joints, splices, stringers, stringer run-outs, or access holes).
3.2. Sub Component. A major three-dimensional structure which can provide complete structural representation of a section of the full structure (e.g., stub-box, section of a spar, wing panel, wing rib, body panel, or frames).
3.3. Component. A major section of the airframe structure (e.g., wing, body, fin, horizontal stabiliser) which can be tested as a complete unit to qualify the structure.
3.4. Full Scale. Dimensions of test article are the same as design; fully representative test specimen (not necessarily complete airframe).
3.5. New Structure. Structure for which behaviour is not adequately predicted by analysis supported by previous test evidence. Structure that utilises significantly different structural design concepts such as details, geometry, structural arrangements, and load paths or materials from previously tested designs.
3.6. Similar New Structure. Structure that utilises similar or comparable structural design concepts such as details, geometry, structural arrangements, and load paths concepts and materials to an existing tested design.
3.7. Derivative/Similar Structure. Structure that uses structural design concepts such as details, geometry, structural arrangements, and load paths, stress levels and materials that are nearly identical to those on which the analytical methods have been validated.
3.8. Previous Test Evidence. Testing of the original structure that is sufficient to verify structural behaviour in accordance with CS 25.305.
4. Introduction
As required by subparagraph (a) of CS 25.307, the structure must be shown to comply with the strength and deformation requirements of Subpart C of CS-25. This means that the structure must:
(a) be able to support limit loads without detrimental permanent deformation, and:
(b) be able to support ultimate loads without failure.
This implies the need of a comprehensive assessment of the external loads (addressed by CS 25.301), the resulting internal strains and stresses, and the structural allowables.
CS 25.307 requires compliance for each critical loading condition. Compliance can be shown by analysis supported by previous test evidence, analysis supported by new test evidence or by test only. As compliance by test only is impractical in most cases, a large portion of the substantiating data will be based on analysis.
There are a number of standard engineering methods and formulas which are known to produce acceptable, often conservative results especially for structures where load paths are well defined. Those standard methods and formulas, applied with a good understanding of their limitations, are considered reliable analyses when showing compliance with CS 25.307. Conservative assumptions may be considered in assessing whether or not an analysis may be accepted without test substantiation.
The application of methods such as Finite Element Method or engineering formulas to complex structures in modern aircraft is considered reliable only when validated by full scale tests (ground and/or flight tests). Experience relevant to the product in the utilisation of such methods should be considered.
5. Classification of structure
(a) The structure of the product should be classified into one of the following three categories:
— New Structure
— Similar New Structure
— Derivative/Similar Structure
(b) Justifications should be provided for classifications other than New Structure. Elements that should be considered are:
(i) The accuracy/conservatism of the analytical methods, and
(ii) Comparison of the structure under investigation with previously tested structure.
Considerations should include, but are not limited to the following:
— external loads (bending moment, shear, torque , etc.);
— internal loads (strains, stresses, etc.);
— structural design concepts such as details, geometry, structural arrangements, load paths;
— materials;
— test experience (load levels achieved, lessons learned);
— deflections;
— deformations;
— extent of extrapolation from test stress levels.
6. Need and Extent of Testing
The following factors should be considered in deciding the need for and the extent of testing including the load levels to be achieved:
(a) The classification of the structure (as above);
(b) The consequence of failure of the structure in terms of the overall integrity of the aeroplane;
(c) The consequence of the failure of interior items of mass and the supporting structure to the safety of the occupants.
Relevant service experience may be included in this evaluation.
7. Certification Approaches
The following certification approaches may be selected:
(a) Analysis, supported by new strength testing of the structure to limit and ultimate load. This is typically the case for New Structure.
Substantiation of the strength and deformation requirements up to limit and ultimate loads normally requires testing of sub-components, full scale components or full scale tests of assembled components (such as a nearly complete airframe). The entire test program should be considered in detail to assure the requirements for strength and deformation can be met up to limit load levels as well as ultimate load levels.
Sufficient limit load test conditions should be performed to verify that the structure meets the deformation requirements of CS 25.305(a) and to provide validation of internal load distribution and analysis predictions for all critical loading conditions.
Because ultimate load tests often result in significant permanent deformation, choices will have to be made with respect to the load conditions applied. This is usually based on the number of test specimens available, the analytical static strength margins of safety of the structure and the range of supporting detail or sub-component tests. An envelope approach may be taken, where a combination of different load cases is applied, each one critical for a different section of the structure.
These limit and ultimate load tests may be supported by detail and sub-component tests that verify the design allowables (tension, shear, compression) of the structure and often provide some degree of validation for ultimate strength.
(b) Analysis validated by previous test evidence and supported with additional limited testing. This is typically the case for Similar New Structure.
The extent of additional limited testing (number of specimens, load levels, etc.) will depend upon the degree of change, relative to the elements of paragraphs 5(b)(i) and (ii).
For example, if the changes to an existing design and analysis necessitate extensive changes to an existing test-validated finite element model (e.g. different rib spacing) additional testing may be needed. Previous test evidence can be relied upon whenever practical.
These additional limited tests may be further supported by detail and sub-component tests that verify the design allowables (tension, shear, compression) of the structure and often provide some degree of validation for ultimate strength.
(c) Analysis, supported by previous test evidence. This is typically the case for Derivative/ Similar Structure.
Justification should be provided for this approach by demonstrating how the previous static test evidence validates the analysis and supports showing compliance for the structure under investigation. Elements that need to be considered are those defined in paragraphs 5(b)(i) and (ii).
For example, if the changes to the existing design and test-validated analysis are evaluated to assure they are relatively minor and the effects of the changes are well understood, the original tests may provide sufficient validation of the analysis and further testing may not be necessary. For example, if a weight increase results in higher loads along with a corresponding increase in some of the element thickness and fastener sizes, and materials and geometry (overall configuration, spacing of structural members, etc.) remain generally the same, the revised analysis could be considered reliable based on the previous validation.
(d) Test only.
Sometimes no reliable analytical method exists, and testing must be used to show compliance with the strength and deformation requirements. In other cases it may be elected to show compliance solely by tests even if there are acceptable analytical methods. In either case, testing by itself can be used to show compliance with the strength and deformation requirements of CS-25 Subpart C. In such cases, the test load conditions should be selected to assure all critical design loads are encompassed.
If tests only are used to show compliance with the strength and deformation requirements for single load path structure which carries flight loads (including pressurisation loads), the test loads must be increased to account for variability in material properties, as required by CS 25.307(d). In lieu of a rational analysis, for metallic materials, a factor of 1.15 applied to the limit and ultimate flight loads may be used. If the structure has multiple load paths, no material correction factor is required.
8. Interpretation of Data
The interpretation of the substantiation analysis and test data requires an extensive review of:
— the representativeness of the loading ;
— the instrumentation data ;
— comparisons with analytical methods ;
— representativeness of the test article(s) ;
— test set-up (fixture, load introductions) ;
— load levels and conditions tested ;
— test results.
Testing is used to validate analytical methods except when showing compliance by test only. If the test results do not correlate with the analysis, the reasons should be identified and appropriate action taken. This should be accomplished whether or not a test article fails below ultimate load.
Should a failure occur below ultimate load, an investigation should be conducted for the product to reveal the cause of this failure. This investigation should include a review of the test specimen and loads, analytical loads, and the structural analysis. This may lead to adjustment in analysis/modelling techniques and/or part redesign and may result in the need for additional testing. The need for additional testing to ensure ultimate load capability, depends on the degree to which the failure is understood and the analysis can be validated by the test.
[Amdt 25/1]