WP 4 - · PDF filea conclusive chapter S ... of a polymer matrix composites system for a structural ... of internal loads and structural integrity. Compare to analysis

Deliverable D4.1 – Characterization of composite material in railways for structural calculation - 1 -

Towards a REgulatory FRamework for the usE of Structural new materials in railway passenger and freight CarbOdyshells

Grant Agreement no.: 605632

WP 4 Characterization of composite material in railways for structural calculation

Deliverable: D4.1 Due date of deliverable: M12 Submission date: 07/10/2014 Version: final Project co-funded by the European Commission within the Seventh Framework Program


REFRESCO Deliverable D4.1 was produced by AIRBUS DS and received

contributions from the following members of the consortium:

DLR for monolithic modelling procedure and failure criteria, fatigue modelling

CAF and CETEST (laboratory of CAF)

BOMBARDIER TRANSPORT for junctions (test matrix and modelling)

ALSTOM

This document should be referenced as:

“REFRESCO- Characterization of composite material in railways for structural calculation , Deliverable D4.1”

Companies Status Names Dates Document

issue Visas

WP4 Leader Lidia Joana ZUBIA 04/09/14 final

T4.1 leader Sylvain CLAUDEL 18/07/14 final

T4.1 contributor Jean Philippe LEARD 04/09/14 final

T4.1 contributor Frederic HALLONET 04/09/14 final

T4.1 contributor

Janko

KREIKEMEIER 04/09/14 final

cetest T4.1 contributor Mikel MURGA 04/09/14 final

T4.1 contributor David LENGERT 07/10/14 final

T4.1 contributor Patrick RICAUD 04/09/14 final


QUALITY CONTROL INFORMATION

Issue Date Description Revising Authorship

Draft Ed 1.1 11/04/2014 Draft version of REFRESCO D4.1 for TMT COMMENT

S. CLAUDEL / AIRBUS DS

Draft Ed 2 12/06/2014 Draft version of REFRESCO D4.1 for TMT COMMENT


Draft Ed 2.1 07/07/2014 Comments to DLR adds S. CLAUDEL / AIRBUS DS

Final for approval

18/07/2014 Integration of BT’s adds on joints and finalization


Final for approval

04/09/2014 Integration of BT’s adds on joints


Final for approval

07/10/2014 Final comments and add to a conclusive chapter


DOCUMENT HISTORY

Issue Date Pages Comment 1 11/04/2014 All Initial issue

2 12/06/2014 All Draft Ed 2 includes CAF comments, CETEST + AIRBUS DS adds

2.1 07/07/2014 All Comments to DLR adds

final 04/09/2014 All Integration of BT’s adds on joints

Final 07/10/2014 All Final comments and add to a conclusive chapter

DISSEMINATION LEVEL

PU Public [X]

PP Restricted to other program participants (including the Commission

Services)

RE Restricted to a group specified by the consortium (including the

Commission Services)

CO Confidential, only for members of the consortium (including the

Commission Services)


TABLE OF CONTENTS

1. REFERENCE ............................................................................................................ 7

2. EXECUTIVE SUMMARY ........................................................................................... 7

3. GLOBAL CHARACTERISATION LOGIC FOR COMPOSITE MATERIAL ................. 8

4. STATISTICAL-BASED MATERIAL PROPERTIES .................................................. 10

5. PROPOSED TESTS MATRICES ............................................................................. 11

5.1 GENERAL CONSIDERATIONS .................................................................................... 11 5.2 STATIC PROPERTIES AND TESTS – MONOLITHIC COMPOSITE (UD

LAMINA/LAMINATE) ....................................................................................................... 12 5.2.1 Screening test matrix ........................................................................................ 12 5.2.2 Qualification test matrix .................................................................................... 13 5.3 STATIC PROPERTIES AND TESTS - SPECIFICITIES FOR A SANDWICH

PANEL (UD LAMINA/LAMINATE) WITH CORE MATERIAL ........................................... 15 5.3.1 Screening test matrix ........................................................................................ 15 5.3.2 Qualification test matrix .................................................................................... 16

5.4 FATIGUE PROPERTIES AND TESTS ................................................................. 17 5.4.1 Fatigue characterization and test recommendations: ........................................ 18 5.4.2 Specificities for a sandwich panel with core material ......................................... 19

5.5 JOINTS ................................................................................................................ 20

6. MODELLING PROCEDURE .................................................................................... 24

6.1 STATIC MODELLING (*1) ................................................................................... 24 6.1.1 Monolithic composite ........................................................................................ 24 6.1.2 Specificities for a sandwich panel with core material ......................................... 32 6.1.3 Static Modelling of Joints .................................................................................. 37

6.2 FATIGUE MODELLING ........................................................................................ 39

7. CONCLUSION ........................................................................................................ 42

ANNEX 1 - DETAILED TEST PROCEDURES DESCRIPTION ………...………..…….………44

ANNEX 2 - STANDARDS FOR COMPOSITE TEST ………………………………….…….52 ANNEX 3 – COMPLEMENTARY REFERENCES………………………………………………..59


LIST OF TABLES

TABLE 1 – LAMINATE AND STRUCTURAL ELEMENT - MECHANICAL TESTING, SCREENING PROGRAMME ............................................................................................................................................................... 12

TABLE 2 – LAMINATE AND STRUCTURAL ELEMENT - MECHANICAL TESTING, QUALIFICATION PROGRAMME ............................................................................................................................................................... 13

TABLE 2 – LAMINATE AND STRUCTURAL ELEMENT - MECHANICAL TESTING, QUALIFICATION PROGRAMME (CONT) ............................................................................................................................................... 14

TABLE 3 – SANDWICH - MECHANICAL TESTING, SCREENING PROGRAMME ................................... 15

TABLE 4 – SANDWICH - MECHANICAL TESTING, QUALIFICATION PROGRAMME ........................... 16

TABLE 5 – TEST MATRIX FOR STRUCTURAL JOINT ................................................................................... 22

TABLE 5 (CONT) – TEST MATRIX FOR STRUCTURAL JOINT ... ¡ERROR! MARCADOR NO DEFINIDO.23

TABLE 6 - FAILURE CRITERIA AND CORRESPONDING F12 COEFFICIENT ........................................... 28

LIST OF FIGURES

FIG.1 - THE PYRAMID OF TESTS – FROM MIL-HDBK-17-1F .......................................................................... 9

FIG. 2 - MATERIAL/STRUCTURAL QUALIFICATION ....................................................................................... 10

FIG. 3 - LIFETIME DIAGRAM FOR A LAMINATE AT R = 0.1 ......................................................................... 18

FIG.4 - EXAMPLE OF GOODMAN DIAGRAM FOR DIFFERENT R RATIO ................................................ 19

FIG 5. – MECHANICAL JOINTS – FRICTION AND FITTED GRIP JOINTS ILLUSTRATION ................. 20

FIG. 6 - SAMPLES FOR ADHESIVE SHEAR TESTS ACCORDING TO REF [R5], VALUES IN MM.... 22

FIG 7: - PRINCIPAL SKETCH OF HOMOGENIZATION PROCEDURE. ....................................................... 26

FIG.8 - FLOWCHART OF FAILURE ANALYSIS OF A LAMINATE ................................................................ 28

FIG. 9 - INTERACTION BETWEEN TRANSVERSE TENSILE STRESS AND SHEAR STRESS ........... 29

FIG 10 - STRESS DISTRIBUTION ON THE ACTION PLANE AS INTRODUCED BY PUCK .................. 29

FIG. 11 - EXAMPLE OF BEARING STRESS/STRAIN CURVE ........................................................................ 30

FIG. 12 - SANDWICH ELEMENT DEFINITION ................................................................................................... 32

FIG. 13 - LAMINATE FAILURE................................................................................................................................. 33

FIG. 14 - TRANSVERSE SHEAR FAILURE .......................................................................................................... 34

FIG. 15 - LOCAL CORE CRUSHING ...................................................................................................................... 34

FIG. 16 - GLOBAL BUCKLING................................................................................................................................. 34

FIG. 17 - SANDWICH FAILURE MODES ............................................................................................................... 35

FIG. 18 - BONDING BETWEEN TWO PIECES IN HYPERMESH ................................................................... 37


FIG. 19 - STANDARD BOLTED JOINTS WITHOUT INSERT IN HYPERMESH.......................................... 37

FIG 21 - TYPICAL STRENGTH AND STIFFNESS DEGRADATION IN COMPOSITE (SCHEMATIC) .. 39

FIG. 22 – MINER’S SUM............................................................................................................................................. 40

FIG. 23 – MINER’S SUM AND LOAD ORDER ..................................................................................................... 40

FIG. 24: - PERCENT FAILURE RULE TO TAKE INTO ACCOUNT THE INFLUENCE OF THE LOADING ORDER, THE NUMBER OF CYCLES AND THE ULTIMATE LEVEL. ....................................... 41

FIG. 25: - ELASTIC STIFFNESS DEGRADATION AS FUNCTION OF LOAD CYCLES .......................... 41

DEFINITIONS DSC differential scanning calorimetry

DMA dynamic mechanical analysis

CAI compression after impact

CTE coefficient of thermal expansion

RH relative humidity

RT room temperature

TMA thermal mechanical analysis


1. REFERENCE

[R1] : MIL-HDBK-17-1F from June 2002 - Composite Material Handbook

[R2] : DOT/FAA/AR-04/24 from June 2004 - A Comparison of CEN and ASTM Test Methods for Composite Materials

[R3] : A UNIQUE CRITERION FOR DESCRIBING FAILURE OF FOAM CORE SANDWICH MATERIALS - A DESIGN ENGINEERING PERSPECTIVE - J. Feldhusen, S. Krishnamoorthy]

[R4] : ISO 13003:2003 (E) - Fibre-reinforced plastics - Determination of fatigue properties under cyclic loading conditions, 2003

[R5] DIN 6701-3 (2002): Adhesive bonding of railway vehicles and parts – Part 3: Guideline for construction design and verification of bonds on railway vehicles.

[R6] DIN EN 1465 (1994): Adhesives, Determination of tensile lap-shear strength of rigid-to-rigid bonded assemblies.

[R7] ASTM D 3528-96 (2002): Standard Test Method for Strength Properties of Double-Lap Shear Adhesive Joints by Tension Loading”

[R8] Zenkert, D. (1997): “The Handbook of Sandwich Construction”, Engineering Materials Advisory Services Ltd, UK

2. EXECUTIVE SUMMARY

The characterization of composite materials in railways for structural calculations has been studied and the findings have been included in the deliverable 4.1. This document proposes:

- A guideline for determining the properties of a polymer matrix composites system for a structural application, in order to be able to perform analysis with finite element modeling.

- Several test matrices are proposed, based on the Aeronautic/aerospace state of arts.

- This document, and specifically the test matrix hereafter proposed, is mainly based the Composite Material Handbook in reference [R1] written by the US Department of Defense which were modified by Airbus Group internal documents. It covers two kinds of structures using monolithic or sandwich material.

- A guideline for determining the modeling procedure of a composite structure as well

as failure criterion, considering both monolithic and sandwich architecture.

- A specific chapter is dedicated to junctions. A first test matrix is proposed, as well as modeling guideline for composite joints.

It concludes that for the mechanical analysis on a composite structure, some changes on the methodology is needed, primarily due to the heterogeneity and anisotropy of the new material. Chapter 3 of this document gives a general overview on the global methodology that is used for the characterization and testing of composite structures in aeronautic (experimental building-block approach). The important link between the manufacturing process and the material properties is also underlined, as well as the effect of the environment. Chapter 4 underlines specifically the variability of the sources in composite materials, which justifies the consideration of a statistical based method to establish the design values of material properties.


Chapter 5 proposes a selection of relevant tests to characterize the material properties needed in the most standard structure analysis (in plane stress and strain and standard elastic theory). The test matrices proposed have to be considered for new materials and reduced test matrix are proposed for the screening phases of the project. Chapter 6 proposes a number of preferred failure criterions to consider in the analysis of both monolithic and sandwich structures. It gives a brief overview on numerical methods available for calculation of composite structure in static and fatigue and advice for the calculation of joints with a finite element code.

3. GLOBAL CHARACTERISATION LOGIC FOR COMPOSITE MATERIAL

In comparison with common metallic materials, composite materials are characterized by their heterogeneity (fiber + matrix at the ply level) and their anisotropy (depending from the stacking sequence or orientation of each ply at the laminate level). Moreover, it is well know that the final behavior of a composite material is highly dependent on :

- its manufacturing process (prepreg, infusion, hand lay-up,…) and manufacturing parameters (curing cycle, …), and not only from the semi-products (fibber, fabric and matrix) which is made of,

- the unitary thickness of the elementary ply Other important specificities of composite materials are their sensitivity to out-of-plane loads, the multiplicity of failure modes and, finally, the lack of universal failure criteria. As a consequence, the global justification/qualification logic of a composite structure is generally based on an experimental building-block approach. This building-block approach can be summarized in the following steps: 1. Generate material basis values and preliminary design allowables.

2. Based on the design/analysis of the structure, select critical areas for subsequent test verification.

3. Determine the most strength-critical failure mode for each design feature.

4. Select the test environment that will produce the strength-critical failure mode. Special attention should be given to matrix-sensitive failure modes (such as compression, out-of-plane shear, and bondlines) and potential "hot-spots" caused by out-of-plane loads or stiffness tailored designs.

5. Design and test a series of test specimens, each one of which simulates a single selected failure mode and loading condition, compare to analytical predictions (and adjust analysis models or design allowables as necessary).

6. Design and conduct increasingly more complicated tests that evaluate more complicated loading situations with the possibility of failure from several potential failure modes. Compare tests results to analytical predictions and adjust analysis models as necessary.

7. Design (including compensation factors) and conduct, as required, full-scale component static and fatigue testing for final validation of internal loads and structural integrity. Compare to analysis.

The building-block approach is shown schematically here below, usually called the “pyramid of tests”:


Fig.1 - The pyramid of tests – from MIL-HDBK-17-1F

Inside this “pyramid of test”, five structural complexity levels have to be considered: constituent, lamina, laminate, structural element and structural subcomponents. The material form(s) to be tested, and the relative emphasis placed on each level, should be determined early in the material data development planning process, and would likely depend upon many factors, including: manufacturing process, structural application. While a single level may suffice in rare instances, most applications will require at least two levels, and it is common to use all five in a complete implementation of the building-block approach. Regardless of the “structural complexity level” selected, physical and chemical properties characterization of the prepreg (or the matrix, if it is added as part of the process, as with resin transfer molding) is necessary to support physical and mechanical properties test results. The five “structural complexity levels” cover the following areas: Constituent Testing:

This evaluates the individual properties of fibers, fiber forms, matrix materials, and fiber-matrix preforms. Key properties, for example, include fiber and matrix density, and fiber tensile strength and tensile modulus.

Lamina Testing (elementary ply): This evaluates the properties of the fiber and matrix together in the composite material form. For the purpose of this discussion, prepreg properties are included in this level, although they are sometimes broken-out into a separate level. Key properties include fiber area weight, matrix content, void content, cured ply thickness, lamina tensile strengths and moduli, lamina compressive strengths and moduli, and lamina shear strengths and moduli.

Laminate Testing: Laminate testing characterizes the response of the composite material in a given laminate design. Key properties include tensile strengths and moduli, compressive strengths and moduli, shear strengths and moduli, interlaminar fracture toughness, and fatigue resistance.

Structural Element Testing: This evaluates the ability of the material to tolerate common laminate discontinuities. Key properties include open and filled hole tensile strengths, open and filled hole compressive


strengths, compression after impact strength, and joint bearing and bearing bypass strengths.

Structural Subcomponent (or higher) Testing: This testing evaluates the behavior and failure mode of increasingly more complex structural assemblies.

At this point of the document, we propose to consider that the smaller scale for failure modes and analysis/calculations will be performed at the lamina (or ply) level. We will not consider data’s for micromechanics calculations for instance, according to actual practises considered in the aeronautic and space design offices. As a consequence, the test matrix and material data will focus on “lamina testing” (at the elementary ply level), “laminate testing” and “Structural Element Testing” at the material level. NOTE :

- “Constituent testing” gives some material data useful for the process and/or for the material specification acceptance values, but these properties are not directly used for the structure analysis.

- On the other side, the “structural subcomponents (or components) testing” are highly dependent from the application, and will not be considered in the framework of this document.

The boundaries of the material characterization covered by this document is explicated by the graphic below:

Fig. 2 - Material/structural qualification

4. STATISTICAL-BASED MATERIAL PROPERTIES Variability in composite material property data may result from a number of sources including run-to-run variability in fabrication, batch-to-batch variability of raw materials, testing variability, and variability intrinsic to the material. It is important to acknowledge this variability when designing with composites and to incorporate it in design values of material properties. Procedures for calculating statistically-based material properties are not included in this document. Nevertheless, details on these procedures could be found in the document in ref [R1] , Volume 1, Chapter 8 This statistical-based method justifies the realization of a minimum number of samples per material batch “Ns” and to test a minimum number of material batches “Nb”.

Constituent tests Lamina tests Laminate tests

Element

Subcomponent

Component

tests

Full scale test(s)

Physical properties DSC, DMA, …

Basic properties Strength, Stiffness, Environment

Laminate performance Strength, Stiffness

Static/fatigue

Environment

Damage tolerance

Joint

Critical design verification Boundary conditions

Secondary effects

Size effect

Static/fatigue

Damage tolerance Framework of this document


Batch (or Lot) definition: For fibers and resin: a quantity of material formed during the same process and having

identical characteristics throughout. For prepregs/laminate : material made from one batch of fibber and one batch of resin The following test matrices (see §5) give an indication of the number of samples to perform. The notation takes into account: - In the first number : the number of different material batch to test (= Nb), - In the second number : the number of samples to test for each batch (= Ns). As an example, “3x5” mean that it is proposed to perform tests from three different material batches, and with at a minimum five samples for each batch, that mean 15 samples are to be tested.

5. PROPOSED TESTS MATRICES The test matrices hereafter proposed have to be considered for new materials and for the most generic application. The tests to be performed should be simplified taking into account the existing knowledge/available user data, and the specific requirements of the application.

5.1 General considerations

In the constitution of the test matrix, it is implicitly considered that the composite material is mainly loaded by in-plane loads (shell hypothesis). Indeed, this is considered as a “best practice” in the design of composite structures. As a consequence, most of the tests are focused on in-plane mechanical characteristics of the material. In case the structure would be loaded by a high level of out of plane loads (3D state of stress), specific supplementary tests should be considered to address this point.

As it has been said above in this document, the performance properties of composite laminates are directly affected by the specific process used for their manufacturing process. It is critical that the test specimens manufactured through the various levels of the building block approach use the same process, representative of the one that will be used in the manufacturing of the Railway parts.

It is still important to evaluate the resistance of new polymer materials to fluids with which they might come in contact. In case the material is expected to be used in an application where fluid exposure occurs for significant time periods at a different temperature, it is recommended that the test laminates be exposed to the above fluids at room temperature conditions, and tested over the expected range of service temperatures.

Annex2 gives the list of the existing standards related to composite characterization. Concerning the standards proposed in the following matrices, AITM standards have not been selected because they are under the copyright of AIRBUS INDUSTRIE. CEN/ISO standards have been selected as much as possible. In case CEN/ISO standards are not available, ASTM standards have been proposed according to document in reference [R1].

For more detail, the document in reference [R2] performs a complete comparison between CEN and ASTM test methods for composite materials. Moreover, Annexe1 gives, for information purpose, some generic considerations on the test procedure.


5.2 STATIC PROPERTIES AND TESTS – MONOLITHIC COMPOSITE (UD LAMINA/LAMINATE)

5.2.1 Screening test matrix

The objective of the screening process is to reveal key mechanical property attributes and/or inadequacies in new material system candidates, while keeping testing to the minimum number of samples. The screening process identifies, for a particular composite material system, the critical test and environmental conditions as well as any other special considerations. Proper test matrix design enables comparison with current production material systems. The recommended minimum set of cured laminate mechanical properties for general application is defined here after:

glass transition temperature DMA EN 6032 1x5

fibber volume and resin content EN 2564 /

density ISO 1183-1 /

cured ply thickness / /

min T° RT max T° max T°

Tensile strength, tensile modulus and poisson ratio 0° tension EN 2561 B (or A) 1x5 1x5

Flexure strength and modulus (**) 4 points flexure ISO 14125 1x5 1x5

In plane shear strength and modulus tension EN 6031 1x5 1x5

Interlaminar shear strength (ILSS)

short beam shear EN 2563 1x5 1x5 1x5 1x5

bearing strength

Representative of the junction (double lap shear, single lap shear, screw, rivet…)

ASTM D5961 procedure A and B 1x5

plain 1x5notch 1x5 1x5plain 1x5 1x5notch 1x5 1x5

Compression after impact (CAI) EN 6038 1x5

(*) : [% at 0° / % at +45° / % at -45° / % at 90°] or stacking sequence representative from the design

[0°]n 1x5

(**) : not used for mechanical datas. Only to evaluate compressive strength sensitivity to environment (temperature/% moist)

1x5

[0°]n

open hole tensile strength

0° compression

properties layup test type and direction test method (1)

1x5

[0°]n

N batch to be tested x N specimens for each batch

test condition

dry see EN2743

wet see

EN2823

[0°]n or (*) 1x5

[0°]n or (*)

[0°]n

50/20/20/10 (*)

(1) These are recommandations but not to be considered as exclusive test method

ASTM D6484

0° tension ASTM D5766

open hole compression strength

(*)

50/20/20/10 (*)

[±45]ns

[0°]n

25/25/25/25 (*)

Table 1 – Laminate and structural element - Mechanical testing, screening program


5.2.2 Qualification test matrix

The recommended minimum set of cured laminate mechanical properties for general application is defined here after:

glass transition temperature DMA EN 6032 5x3

fibber volume and resin content EN 2564 /

density ISO 1183-1 /

cured ply thickness / /

/

/

moisture diffusivity ASTM D5229 /

thermal diffusivity ISO 1159-2 /

specif ic heat DSC ISO 11357-4 /


Tensile strength and modulus 0° tension EN 2561 B (or A) 1x5 5x5 1x5 1x5

Poisson Ratio 0° tension EN 2561 B (or A) 1x5 1x5

Tensile strength and modulus 90° tension EN 2597 B 1x5 1x5 1x5

Compression strength and modulus 0° compression EN 2850 A (or B) 1x5 5x5 1x5 5x5

Compression strength and modulus 90° compression EN 2850 B 1x5 1x5

In plane shear strength and modulus in plane tension EN 6031 1x5 5x5 1x5 5x5

Interlaminar shear strength (ILSS) short beam shear EN 2563 1x5 5x5 1x5 5x5

1x5 5x5

1x5 5x5 5x5

1x5 5x5 5x5

(*) : [% at 0° / % at +45° / % at -45° / % at 90°] or stacking sequence representative from the design

1x5[0°]n or (*)

1x5

1x5

dry see EN2743

w et see

EN2823

5x3

5x5

1x5

1x5

in plane coefficient of thermal expansion

[90°]nISO 11359-2TMA

[0°]n or (*)

[0°]n or (*)

[0°]n

[0°]n

5x3


25/25/25/25 (*)

10/40/40/10 (*)

50/20/20/10 (*)

bearing strength


properties test method (1)

5x3

test type and direction

test condition

[0°]n

layup

ASTM D5961 procedure A and B

[90°]n

[±45]ns

[0°]n

[0°]n

[0°]n or (*)

[0°]n or (*)

[0°]n

[0°]n

[90°]n

Representative of the junction (double lap shear, single lap shear, screw , rivet…)

Table 2 – Laminate and structural element - Mechanical testing, qualification program



plain 1x5 1x5

notch 1x5 5x5

plain 1x5

notch 1x5 5x5 5x5

plain 1x5 1x5 1x5

notch 1x5 5x5 5x5

plain 1x5 1x5

notch 1x5 5x5

plain 1x5

notch 1x5 5x5 5x5

plain 5x5 1x5

notch 1x5 5x5 5x5

25/25/25/25 (*) notch 1x5 5x5

10/40/40/10 (*) notch

50/20/20/10 (*) notch 5x5 5x5

25/25/25/25 (*) notch 1x5 5x5

10/40/40/10 (*) notch 5x5

50/20/20/10 (*) notch 5x5 5x5

Compression after impact (CAI) EN 6038 5x11 (**) 5x11 (**)

G1c ASTM D5528 3x5 1x5

G2c PREN 6034 3x5 1x5

Resistance to agressive fluid (***)

short beam shear (ILSS) EN 2563

(*) : [% at 0° / % at +45° / % at -45° / % at 90°] or stacking sequence representative from the design (**) : first 1 batch in both condition, other 4 batches at worst case condition

ASTM D6484

0° tension ASTM D5766

ASTM D6742

0° tension

test type and direction

10/40/40/10 (*) 0° compression

layup


test condition

open hole tensile strength

properties dry see EN2743

wet see

EN2823

open hole compression strength

25/25/25/25 (*)

10/40/40/10 (*)

50/20/20/10 (*)

25/25/25/25 (*)

test method (1)


filled hole tensile strength

ASTM D5766 modified according

to MIL-HDBK-17 section 7.4.2.2

0° compression

(***) : optional tests. These include testing of cured laminates after exposure of the laminates to solvents that the part will be subjected to in actual service.

(*)

50/20/20/10 (*)

filled hole compression strength

Table 2 – Laminate and structural element - Mechanical testing, qualification program (continuation)


5.3 STATIC PROPERTIES AND TESTS - SPECIFICITIES FOR A SANDWICH PANEL (UD LAMINA/LAMINATE) WITH CORE MATERIAL

5.3.1 Screening test matrix

The recommended minimum set of cured laminate mechanical properties for general application is defined here after:


tensile strength (out of plane) and core/skin junction under tensile load

tensile ASTM C297 1x5

compresive strength and modulus (out of plane) compression ISO 844 1x5

shear strength and modulus shear ISO 1922 1x5 1x5

core/skin junction under shear load (and shear strength and modulus)

sandwich flexure or

shear

ASTM C393 or

ISO 1922 with skin1x5 1x5

plain 1x5 1x5

notch 1x5 1x5

(*) : [% at 0° / % at +45° / % at -45° / % at 90°] or stacking sequence representative from the design (**) : instead of the open hole compression characterisation test of the composite skin alone (see matrix §4.2.1)

ASTM D 7249

core

core

[(*) / core / (*)]

properties layup (2) test type and direction

[(*) / core / (*)]

test method (1)


test condition

dry see EN2743

wet see

EN2823

skin strength and open hole compression strength

(**)

sandwich long beam flexure

(1) These are recommandations but not to be considered as exclusive test method (2) : If the material is designed to be self-adhesive to the core, then these tests should be conducted on cocured panels fabricated without adhesive. If the material requires an adhesive layer for bonding to the core, then the tests can be conducted on either (or both) cocured panels or precured skins secondarily bonded to the core, depending on the anticipated design and fabrication methods to be used with the material.

[50/20/20/10] / core /

[50/20/20/10] (*)

Table 3 – Sandwich - Mechanical testing, screening program


5.3.2 Qualification test matrix

The recommended minimum set of core and cured sandwich mechanical properties for general application is defined here after:


tensile strength (out of plane) and core/skin junction under tensile load

tensile ASTM C297 1x5 1x5

compresive strength and modulus (out of plane) compression ISO 844 1x5 1x5 1x5

shear strength and modulus shear ISO 1922 1x5 3x5 3x5

core/skin junction under shear load (and shear strength and modulus)

sandw ich f lexure or

shear

ASTM C393 or

ISO 1922 w ith skin

1x5 1x5

plain 1x5 1x5

notch 1x5 5x5

plain 1x5

notch 1x5 5x5 5x5

plain 5x5 1x5

notch 1x5 5x5 5x5

[25/25/25/25] / core /

[25/25/25/25] (*)notch 1x5 5x5

[10/40/40/10] / core /

[10/40/40/10] (*)notch 5x5

[50/20/20/10] / core /

[50/20/20/10] (*)notch 5x5 5x5

(*) : [% at 0° / % at +45° / % at -45° / % at 90°] or stacking sequence representative from the design (**) : instead of the open and f illed hole compression characterisation test of the composite skin alone (see matrix §4.2.2)

(2) : If the material is designed to be self -adhesive to the core, then these tests should be conducted on cocured panels fabricated w ithout adhesive. If the material requires an adhesive layer for bonding to the core, then the tests can be conducted on either (or both) cocured panels or precured skins secondarily bonded to the core, depending on the anticipated design and fabrication methods to be used w ith the material.


core

properties dry see EN2743


w et see

EN2823

test method (1)test type and

directionlayup (2)

test condition

[(*) / core / (*)]

[(*) / core / (*)]

core

f illed hole compression strength (**)

sandw ich long beam f lexure w ith

open holeASTM D 7249

skin strength and open hole compression strength (**)

[25/25/25/25] / core /

[25/25/25/25] (*)

sandw ich long beam flexure ASTM D 7249

[10/40/40/10] / core /

[10/40/40/10] (*)

[50/20/20/10] / core /

[50/20/20/10] (*)

Table 4 – Sandwich - Mechanical testing, qualification program


5.4 FATIGUE PROPERTIES AND TESTS

In the area of fatigue, however, no generalized methodology has yet been devised to predict laminate behavior from unidirectional specimen data.

Hence, the development of fatigue design values becomes a unique problem for each application lay-up.

Many studies have been undertaken, and much has been written concerning life prediction for specific laminates under cyclic loading spectra. Even at this level, empirical methods have been favored due to the inadequacy of results obtained from cumulative damage models, fracture mechanics analyses, and other theoretical approaches.

Fatigue data is generated at the design critical test conditions (room temperature, or hot/wet).

The characteristics for fatigue resistance of materials shall be determined with experimental methods for cyclic loadings associated with static loads under representative service conditions.

The part of the fatigue curve which is characterized shall cover the domain of use (in term of number of cycles, stress or strain amplitude and R ratio) – see below:

σ

With : R = max

min

; σampl = (σmax – σmin ) / 2

For instance, R= 0,1 correspond to a pure traction load, R= -1 correspond to a traction/compression load with σmoy = 0, R = 10 correspond to a pure compression load.

In general, composite structures are assumed to be less sensitive to dynamic loading than metallic structures, but general well accepted hypothesis of damage accumulation are still missing.

The damage itself is therefore strongly influenced by the number of cycles, the ultimate loading and the order of the loading, i.e. if high load levels are applied onto the laminate followed by lower leads to earlier failure than vice versa. This point has to be taken into account when defining the test sequence of a composite structure/substructure under a representative cycling loading.

σ

time

S

σmin

σmoy

σmax

σamp


5.4.1 Fatigue characterization and test recommendations:

In ref [R4] (for instance), the principal procedure of fatigue investigation for composites is described.

In the absence of today’s existing guidelines for fatigue test specimens, the geometries of the specimens therefore is based on the standards for quasi static investigations, (e.g. according to ref [R4]):

for sample in tensile : DIN EN ISO 527-4 or DIN EN ISO 527-5 (for UD)

for samples in flexure samples : ISO 14125, with 4 points flexure preferred in that case

NOTE : The general lack of today’s existing and applied test methods must be seen within the concentration on single mechanisms under investigation only, e.g. constitutive behavior at certain stress ratios and/or single damage phenomena etc. The question of general transferability of results, is not solved yet. Furthermore, the principal phenomenology of fatigue damage on composite materials is not understood. This is due to the very complex structure of composite materials consisting of fiber roving, matrix material and adhesive layers between fibers and matrix. In the near future, strong experimental effort has to be spent for the investigation to answer these fundamental.

The determination of fatigue properties is commonly based on lifetime diagram (“S-N curves”), energy release rates, residual strength and/or residual stiffness properties, (DIN 50100 - Dauerschwingversuch, 1978). With this, the fatigue strength and the fatigue limit are estimated by measuring the number of cycles for given stress ratios.

NOTE : Generally, the low cycle fatigue regime (<105 cycles), the high cycle fatigue regime (<106 cycles) and the very high cycle fatigue regime (>108 cycles) are distinguished.

For each “R ratio” defined, the lifetime diagram (“S-N curve”) shall comprise at least 20 specimens (4 maximum stress or strain levels with 5 samples for each level), for each representative temperature (see example below):

Fig. 3 - Lifetime diagram for a laminate at R = 0.1 For material screening, 8 data points are usually sufficient to establish the preliminary fatigue curve.

It is frequently useful for the justification to perform test for different R ratio. The results of such tests can be draw on diagram called Goodman diagram (see fig below)

static strength R=0.1

Log ( )


Fig.4 - Example of Goodman diagram for different R ratio

It would be very costly, in an experimental point of view, to perform test series for a large number of “R ratio”. As a consequence, it is usual to perform tests for a limited number.

Nevertheless, the number of “R ratio” to test for a material/structure depends on the material database available, on the cyclic load of the application and is finally linked to the margin policy considered. As a consequence, it is difficult to define or propose to consider a unique combination of “R ratio” to test for a generic application. For the most generic application, it is proposed to test at a minimum R=0,1 and R=-1. The frequency of the cycling test should not introduce temperature elevation which could damage the sample. As a consequence it is recommended to not exceed 5Hz. Nevertheless, it is possible to increase this frequency up to 20/30Hz with an adequate cooling device.

The failure can be considered at the complete sample separation or at a given level of stiffness reduction. If the latest failure criterion is considered, a level of stiffness reduction of 20% is a classical value.

5.4.2 Specificities for a sandwich panel with core material

The reduction of stiffness or stability properties of sandwich structures due to cyclic fatigue should be considered. This reduction may be caused by: - a reduction in modulus of elasticity in the facings materials and/or in the core(s)

materials due to various types of damage, e.g. micro cracks - a local debonding between faces and core at the interface. As a consequence, specific tests should be performed at the sandwich level, with 4 points Flexion test according to ASTM D7249 for instance.


5.5 JOINTS

5.5.1 Definition of Joints

In order to limit the number of potential combinations all composite materials to be considered as quasi-isotropic laminates. If it is not the case, tests should be performed for each specific stacking sequence.

Bonded Joints (only pure adhesives considered, without any additives like short-fibers)

Both joints for composite vs. composite pieces and composite vs. metal pieces

Adhesives with high Young’s Modulus, low strains (i.e. known as thin-film), best properties 0.05mm < t < 0.2mm i.e. for epoxy

Adhesives with low Young’s Modulus, high strains (i.e. known as thick-film)

Mechanical Joints

Friction Grip Joints (see illustration fig. 5 here below) For composite vs. composite pieces, composite vs. metal pieces, composite vs. metal pieces covered with composite material o Rivets o Huck-Bolt Type

Standard bolted joints with insert/reinforcement in laminate (with nuts) o Screws Standard bolted joints without insert/reinforcement in laminate (with nuts) Standard bolted joints with insert/reinforcement in laminate (with nuts) Tapped thread joints

Fitted Grip Joints (see illustration fig. 5 here below) For composite vs. composite pieces, composite vs. metal pieces, composite vs. metal pieces covered with composite material o Rivets (i.e. blind rivets) o Huck-Bolt Type

Standard bolted joints with/without insert/reinforcement in laminate (with nuts)

o Screws Standard bolted joints with/without insert/reinforcement in laminate (with

nuts)

Fig 5. – Mechanical joints – Friction and Fitted grip joints illustration


5.5.2 Test matrix The following test matrix is based on experiences and, if possible, based on standards from railway engineering. It does not guarantee to be complete since testing procedures also depend on load cases, on the considered parts (structural relevant or not) and on the test laboratory. Before testing it has to be specifically discussed about testing budget/costs, load cases and agreed with the test laboratory which tests are necessary and possible. Therefore, the following statements are proposals and tests can be omitted/added in specific cases.

Regarding failure criteria, if there are any restrictions or criteria mentioned in the standard for a joint type then these are to use instead of/in addition to given failure criteria in the table.

In general, for static load tests, 5-8 specimen are needed. For fatigue load tests, for every R-ratio, 15 specimen are necessary on different stress amplitudes in order to define a Wöhler-curve.

For bonded joints, some specific points must be taken into account:

- Temperatures (i.e. 3 different levels) as well as other influencing mediums such as moisture have a significate effect on joint behavior. As a consequence, the effect of environment must be considered in the characterization matrix

- Due to peak stress at the edge of the bonded joint, the shear strength is usually not proportional to the bonded length. As a consequence, the sample joint design should be as close as possible from the final design. In the case it is not possible, the local stress distribution along the bonded joint should be evaluated throw a calculation.

For mechanical joint with composite parts, it must be considered that composite material is more sensitive to creeping than metal under compressive load. As a consequence, the loss of pre-load in the screw/rivet have to be considered and characterized by tests. In the case of highly loaded screw/rivet, the admissible surface pressure has to be checked. A metallic insert can also be embedded in the clamping area to improve those points.


Joint Type Sample Dimensions

Standards for test

parameters

Load Type Static (with schematic pictures) (*6) Load Type Fatigue, R=

Criteria for joint verification (*1)

Useful testing? Tensile/C

ompression

Shear Combined Other (i.e. creeping) 0.1 -1 (*3)

Bonded Joints (*2)

Adhesives with high Young’s Modulus (*4)

DIN 6701, (high strength glue) DIN 6701-3 X X X X (Creep tests) X X Shear Strength,

See DIN6701-3 X

Adhesives with low Young’s

Modulus

DIN 6701 (low strength glue) DVS 1618 (app. 3)

see figure 5 DIN 6701-3 X X X X (Creep

tests) X X

Bimodal regulation (force regulation amplitude), see

DIN6701-3

X

*1 In general for composite laminates the following theories are commonly used for evaluation: a) Tsai Wu theory, b) Tsai Hill theory, c) Hoffman’s theory, d) Maximum strain theory. Moreover, critical strains are useful for evaluation *2 Also “ASTM D 3528-96 (2002): Standard Test Method for Strength Properties of Double-Lap Shear Adhesive Joints by Tension Loading” can be considered for composites and metal/composite components *3 In general, testing negative 'R'-ratios depends on the parts which are considered and especially the loads which are applied, i.e. aerodynamical loads. Therefore, in some cases it is necessary to consider R=0, or 0.7 and in some rare cases R=-0.5 to -0.7 as well. For compression fatigue R-values (i.e. R=-1.0), the specimen needs to be stiff enough in order to establish compressive stresses *4 For structural adhesives (thin-film) only a proposal is given because it is not often performed and therefore cannot generally be confirmed by BT. *6 The small schematic pictures are used to show how the load is applied in general but are not explicitly describing the sample geometries. These sample geometries vary depending on the joint.

Table 5 – Test matrix for structural joint

Fig. 6 - Samples for adhesive shear tests according to ref [R5], values in mm


Joint Type Sample Dimensions

Standards for test

parameters

Load Type Static (with schematic pictures) (*6)

Load Type Fatigue, R= Criteria for joint verification (*1) Useful

testing? Tensile/Compression Shear Combined 0.5 -1 (*3)

Mechanical Joints (*5) Friction Grip Joints

Rivets X X X X X For the joint: Loss of preload due to setting or creeping, Measurement of preload either separately

or during static load tests For the laminate: i.e. connection between cylindrical

insert and laminate, plastic deformation, fracture

X

Huck-Bolt Type X X X X X X or DVS/EFB 3435

Screws X X X X X (X) VDI2230

Fitted Grip Joints Rivets X X X X X For the joint: Loss of preload due to setting or

creeping, Measurement of preload either separately or during static load tests

For the laminate: i.e. connection between cylindrical insert and laminate, plastic deformation, fracture,

bearing stress

X

Huck-Bolt Type X X X X X X or DVS/EFB 3435

Screws X X X X X (X) VDI2230

*1 In general for composite laminates the following theories are commonly used for evaluation: a) Tsai Wu theory, b) Tsai Hill theory, c) Hoffman’s theory, d) Maximum strain theory. Moreover, critical strains are useful for evaluation *2 Setting should be investigated no matter if two pure composite parts are joint or if there are inserts (metal) with different surface properties

*3 In general, testing negative 'R'-ratios depends on the parts which are considered and especially the loads which are applied, i.e. aerodynamical loads. Therefore, in some cases it is necessary to consider R=0, 0.1 or 0.7 and in some rare cases R=-0.5 to -0.7 as well. For compression fatigue R-values (i.e. R=-1.0), the specimen needs to be stiff enough in order to establish compressive stresses *4 For structural adhesives (thin-film) only a proposal is given because it is yet not performed and therefore cannot be confirmed by BT. *5 Screws then can be evaluated according to VDI2230 (no tests necessary), Huck-Bolts according to DVS/EFB 3435 or tests. Strength of rivets has to be proven by tests. *6 The small schematic pictures are used to show how the load is applied in general but are not explicitly describing the sample geometries. These sample geometries vary depending on the joint.

Table 5 (continuation) – Test matrix for structural joint


6. MODELLING PROCEDURE 6.1 STATIC MODELLING (*1)

When modeling composites, two ways can be followed: solid elements or shell elements.

Since the thickness of composites is mostly very small compared to its remaining dimensions modeling laminates with a certain amount of plies by using solid elements is very expensive considering model size and solving time.

Therefore in that case, using shell elements for the laminates is probably the best and most efficient choice. Indeed, 3D models are reduced to their middle surface and meshed with shell elements. For laminates, it is defined that the laminate layers are bonded together in order to form a cohesive structure.

In the FE program, a component name for the laminate has to be chosen. After that the property for laminates is defined and a material is allocated. These properties generate shell elements where the following characteristics are defined and may be different for every ply:

- ply number,

- amount of plies

- material ID of each ply

- thickness of each ply

- orientation of each ply (i.e. 0, 90 or +-45 degrees)

With this definition, laminates as well as sandwich structures can be mapped.

*1 Modeling composites is not commonly carried out by BT. Therefore, only a proposal is given here.

6.1.1 Monolithic composite

6.1.1.1 Mixture Rules The consideration of single fibers within the composite cannot be taken into account directly, due to extraordinary high computational effort. Instead, the single plies of the laminated structure have to be modeled as homogeneous continuum.

Homogenization methods or mixture rules are used to define the global behavior on a sufficient level by neglecting local aspects. The mathematical analysis of a laminated structure demands the knowledge about the constitutive relations of the single plies, the definition of a reference placement, the transformation relations for the single plies and the assumptions about through thickness displacements.

In the aircraft design process the Classical Laminate Theory (CLT) is widely used. The CLT assumes homogeneous and orthotropic single plies, plane stress state conditions, the ideal bonding between neighboring plies, linearly shaped membrane displacements and the negligence of transverse shear strains.

The strains itself can directly be deduced from the membrane displacements via differentiation. The corresponding stresses are obtained by multiplying the layer wise constants ply stiffness with the strains. The complete constitutive relation of the laminate is defined by integration of the layer wise material laws over the thickness of the laminate, which results in the well-known matrices of membrane stiffness A, coupling stiffness matrix B and bending stiffness matrix D


xy

y

x

xy

y

x

xy

y

x

xy

y

x

DDDSymDDDBBBABBBAABBBAAA

MMMNNN

33

2322

131211

33321333

2322122322

131211131211

.

. (1)

The in-plane characteristics of the lamina EL, ET, GLT and nuLT need to be define to perform analysis, where E and G are Young’s modulus and shear modulus resp. and the subscripts L, T denote the longitudinal and transverse fiber direction.

When the elementary properties of the fiber and the matrix are the only data’s available, it is possible to evaluate the properties of the lamina (elementary unidirectional ply) by using the simplest assumptions for stiffness homogenization:

mfL EEE )1( (2)

mfLT )1( (3)

mfT EEE /)1(//1 (4)

Where the subscripts f and m denote the fiber and matrix correspondence respectively, and the fiber volume fraction is denotes by .

Nevertheless, the properties in transverse fiber direction commonly need a correction. In case of composites containing isotropic fibers, the following formulae for transverse stiffness moduli were established:

25,1

2

)1(/)85,01(

fm

mT EE

EE (5)

mfT EE

E/)1(5,0/

)1(5,0

(6)

fm

mfT EE

EEE

)1( (7)

)/1(1/ fmmT EEEE (8)

General recommendations when using equations (5) to (8) can not be given. It must be noted, that the range of results can reach more than 10%. As a consequence, theses formulas should be considered only unless material characterization of the lamina is available (see §5.2).

NOTE : concerning GLT, it is assumed to be not so different from a material to another (for epoxy systems). As a consequence, for a early phase of project it is proposed to use previously characterized value as a first order of magnitude, unless characterization test results will be available. IMPORTANT : All the methodologies mentioned above are used to calculate the constitutive properties of composite materials in 2D conditions, only. The constitutive properties in thickness direction are not covered by the CLT. For this, advanced mathematical tools, e.g. analytical and/or numerical homogenization techniques have to be applied.


6.1.1.2 Analytical and Numerical Homogenization Methods for Monolithic Materials

The use of fibers reinforced composite materials in nowadays aerospace applications, i.e. glass and / or carbon fibers are embedded into a matrix material, demands a precise knowledge about the constitutive behavior of the constituents. For this, different analytical and numerical methodologies were developed during the past decades.

A considerable improvement for the estimation of elastic constants was reached by the works of (Hashin & Shtrikman, 1962) and (Hashin & Shtrikman, 1963) by the definition of variational principles for isotropic and anisotropic and heterogeneous multiphase materials. The stress polarization to account for the difference between the true stress and the stress resulting from the true strain acting on a homogeneous trial material was introduced.

It must be noted, that for all the approaches mentioned above, the small strain assumption holds.

In case of very complex microstructures, the use of a representative volume element (RVE) in conjunction with numerical homogenization scheme is recommended, (Böhlke, 2001), (Kouznetsova, 2002), (Lubarda, 2002) or (Nemta-Nasser, 1999).

Therefore the RVE must capture the main features of the microstructure and has to represent a material point on the macrostructure at the same time. This is achieved if the dimensions on the microscale are orders of magnitudes smaller compared to the structural dimensions, which is known as scale separation.

The aim of analytical as well as numerical methods to estimate effective material properties is to obtain macroscopically homogeneous properties from the microscopically very inhomogeneous constituents, which can be used in structural analysis (see fig. 7 from Gross & Seelig, 2007).

It must be noticed that RVE method is considered as an advanced method not frequently used in industry for structure analysis.

Fig 7: - Principal sketch of homogenization procedure.


6.1.1.3 Failure Criteria for Monolithic Composite Different failure modes can be observed at the failure of a composite material (see some

examples here below):

- matrix cracking, matrix/fiber debonding

- delamination,

- fiber failure,

- micro-buckling of fibers (under compressive stress)

Failure criteria are used to evaluate the local or global structural behavior, taking into account the different failure modes as well as the multi axial state of stress of the structure.

6.1.1.4 Failure criteria - First ply failure/ last ply failure In a composite structure nevertheless, the failure of one ply of a laminate doesn’t correspond systematically to the failure of the laminate. As a consequence, failure criteria of laminates can be categorized into first ply failure criteria (FPF) and last ply failure criteria (LPF). FPF are often used due to their very conservative predictions. This is due to the disregard of further load carrying capabilities of the remaining plies.

Practically, this kind of failure prediction approach is favored in industry due to its simplicity and robustness.

In contrast to this, LPF predict the failure of the laminate if the strength limit of the last remaining ply is reached. If the failure criterion of a single ply is reached, a recalculation with reduced stiffness values is carried out until the last ply of the laminate fails as well. A principal sketch of the failure analysis of a laminate is depicted in figure 8. The actual state of stress is assessed with respect to the failure criterion used. If fracture occurs, the user will decide if the degradation is tolerable or if the stacking sequence has to be changed. The rearranged state


of stress is reassessed and evaluated again until the stress analysis is finished (Knops & Böge, 2006).

Fig.8 - Flowchart of failure analysis of a laminate

6.1.1.5 Failure criteria - Differential and non-differential failure criteria Interpolation criteria take into account the general multi axial loading conditions of a composite. In (Gol`denblat & Kopnov, 1965), a tensorial polynomial description was introduced, where in reality the restriction to fourth order strength tensors is commonly made,

1 jiijii FF (14)

where iF and ijF denote strength tensors of second and fourth order, respectively.

It must be noted that a distinction between fiber cracking and matrix failure cannot be made in those criterion : the onset of failure itself is predictable, only.

By different choices of the coupling coefficient F12 numerous well established criteria can be

deduced from the general interpolation criterion (see table 1, where TTLL RandRRR ,,

denote the longitudinal and transverse tensile and compression strength values, respectively).

Criterion F12

Tsai-Wu 0

Hoffmann )/(5,0 LL RR

Norris )/(5,0 TLRR

Tsai-Hahn TTLL RRRR/5,0

Table 6 - Failure criteria and corresponding F12 coefficient

Beside the interpolation criteria mentioned above, some other criterion where introduced in order to perform a distinction between fiber fracture and inter fiber fracture. In that case the inter fiber fracture criterion was introduced, motivated by experimental observations where an interaction of the transverse tensile stress and the shear stress was noticed (see figure 9). The fracture limit obviously is reached before the strength values.


Fig. 9 - Interaction between transverse tensile stress and shear stress

Here below will be detailed two of them: Hashin and Puck

In the Hashin criteria (Hashin, Failure Criteria for Unidirectional Fiber Composites, 1980), failure criteria to distinguish fiber tension (15), fiber compression (16), matrix tension (17) and matrix compression (18) were developed

212

211

LT

tf SX

F (15)

211

C

cf X

F (16)

212

222

LT

tm SY

F (17)

21222

2222 1

22

LCT

C

Tc

m SYSY

SF

(18)

Where the superscripts T and C and L denote tension, compression and longitudinal and X, Y and S are the longitudinal, transverse and shear strength values, respectively.

Puck, in his approach, introduced the concept of fracture process zone (see figure 10), who established the most prominent criterion to take into account the very complex modes of inter fiber fracture.

Fig 10 - Stress distribution on the action plane as introduced by Puck

Thus, three different inter fiber fracture modes can be distinguished:

1. Mode A: the fracture is caused by a tensile stress or by a longitudinal shear stress which leads to a degradation of the Young’s modulus and the shear modulus, respectively.

2. Mode B: the fracture is caused by a longitudinal shear stress. The transverse compression stress acts on the same fracture plane as the shear stress,


simultaneously. Hence there is no further crack opening but the fracture surfaces are pressed on each other.

3. Mode C: if the ratio of compressive normal stress at fracture and the transverse compressive strength exceeds a value of 0,4, the action plane of the external shear stress is no more the fracture plane but the fracture occurs on a plane inclined by an angle 0FP to the action plane of 2 and 21 .

From the physics point of view the Puck criterion should be preferred due to the introduction of the fracture process zone. In comparison, the Hashin theory seems to be more conservative.

6.1.1.6 Failure criteria - Compression Strength for Monolithic Composites The compression strength of composite materials is dominated by two main reasons: fiber buckling and composite delamination, (Martinez & Oller, 2009).

A shear buckling mode as well as the extensional buckling mode depending on the fiber volume fraction was defined by (Rosen, 1965).

Further improvements to describe the compression strength of composite materials can be found in (Jochum & Grandidier, 2004).

6.1.1.7 Failure criteria – Bearing strength The bearing stress correspond to the applied load “P” divided by the projected bearing area onto the area orthogonal to the bearing direction, ie: the product of the nominal bolt diameter “D” and the specimen thickness “t”.

σ_bearing = P / (D x t)

An example of the resulting bearing stress/bearing strain curve is shown in Figure below.

The bearing strain was obtained by normalizing the displacement by the bolt diameter.

The offset bearing strength is the value to consider for calculations. Thus, the 2% offset measurement, which is the default in the proposed standard (see figure below), correspond to a “ovalisazion” of 2% of the hole diameter.

Fig. 11 - Example of bearing stress/strain curve

Offset bearing strength

D

t

P


There is no general consensus as to what the value of the offset bearing strength should be. The usage in the aerospace industry varies from 1%D for stiff double shear joints, to 4%D for single shear joints (the latter being a standard for metal bearing tests).

6.1.1.8 Fracture Mechanical Approaches and Further Extensions for Monolithic Composites

The characterization of cracks within structural components became possible by the definition of stress intensity factors, (Irwin, 1957). The crack modes due to tension, planar shear and non planar shear allow the definition of three different stress intensity factors. One drawback must be seen by the stress singularity at the crack tip where the stress value rise to infinity.

The definition of the J-Integral allows the assessment of linear elastic as well as inelastic material behavior, (Rice, 1968). Due to the path independence of the integration, the stress intensity factor can be calculated as

221

IKE

J (19)

To overcome the problem of stress singularity at the crack tip and to assess especially delamination phenomena, cohesive zone models can be used, (Dugdale, 1960), (Barenblatt, 1962) or (Camanho, Dávila, & Ambur, 2001). The crack tip is extended to a fracture process zone where cohesive forces can cause a critical crack opening. Thereafter, the cohesive forces degrade until the crack surfaces are stress free and the crack further propagates.

Beside the discrete fracture models described above, the continuum damage mechanics which uses a continuous degradation parameter to describe the loss of stiffness in a material, can be used. Therefore, the ratio of actual effective area without any defects to the initial area is assessed. Thus, the actual damage can be quantified in a smeared manner.

The numerical treatment of fracture and damage phenomena was pushed by (Mazars & Pijaudier-Cabot, 1989) by the introduction of a model to describe the damage localization. Due to the smeared character not the crack itself but its action on the continuum is modeled. The use of discrete interface elements based on cohesive zone approaches was pioneered by (Neeleman, 1987).

NOTE : For implementation issues of cohesive zone approaches into finite element program (Xu & Needleman, 1993), (Xu & Needleman, 1995), (Ortiz & Pandolfi, 1999), (Ghosh, Ling, Majumdar, & Kim, 2000), (Camanho, Dávila, & Ambur, 2001), (Camanho & Davila, 2002) or (Camanho, Davila, & de Moura, 2003) should be quoted.


6.1.2 Specificities for a sandwich panel with core material

A sandwich panel is constituted of a light weight core (isotropic materials such as foams or orthotropic such as honeycomb, balsa …) embedded between two composite facesheets.

Fig. 12 - Sandwich element definition

The following sandwich parameters are used in the further analysis:

Ffacesheet properties tf facesheet thickness E1 young’s modulus in longitudinal direction E2 young’s modulus in transversal direction Ef young’s modulus geometrical average value equal to pE1 E2 G12 in-plane shear Modulus 12 in-plane Poisson’s ratio 21 in-plane Poisson’s ratio plasticity factor waviness of facesheet

C core properties tc core height s cell size Ec compressive modulus (in normal direction) G13 core shear modulus in longitudinal direction G23 core shear modulus in circumferential direction c flatwise core compressive strength d total height of sandwich (d=tc + 2 tf)

6.1.2.1 Modeling Depending on the level of through thickness stresses, choice of 2-D or 3-D elements shall be made for sandwich panels analysis. If out of plane stresses may be neglected, in-plane 2-D analysis (shell elements) may be used, otherwise, 3-D elements should be used.

FEA of sandwich structures can be carried out with following element types or combinations:

a single layered shell elements for the entire sandwich material (for in-plane 2-D analysis)


(layered) shell elements for the faces and solid elements for the core (for 3-D analysis)

solid elements for both faces and core (detailed 3-D analysis,).

The solid modeling could be restrained to local areas of complex geometry, load introduction … where out of plane effects are more significant and require locally more complex modelling.

For the analysis of sandwich structures, special considerations shall be taken into account, such as:

elements including core shear deformation shall be selected

for honeycomb cores one shall account for material orthotropy, since honeycomb has different shear moduli in different directions

local load introductions, corners and joints, shall be checked

For many core materials, experimentally measured values of E, G and ν are not in agreement with the isotropic formula. In that case, to assure that the shear response of the core will be described accurately, the measured values for G and ν shall be used, and the E value shall be calculated from the formula : E= 2 G (1+).

Modeling of skin laminates uses same methodologies than described chapter 6.1.1.

6.1.2.2 Failure modes and criteria Failure of a sandwich panel can occur:

In the facesheets

In the core

at the core-facesheets interface

Failure in the facesheets

A laminate failure can occur in facesheets caused by an overstressing (Fig. 13). Determination of strength allowables for the laminates uses same methodologies than described chapter 6.1.1. One can note that allowables material values used in facesheet analysis have to be fully representative of skin materials real strength, process effects on mechanical characteristics should be taken into account (for example, co-curing of composite skins on honeycomb may generate waves on the inner plies of skins and reduce strength) .

Fig. 13 - Laminate failure Both faces shall be checked for failure, since they will be exposed to different stress states if exposed to bending loads.


6.1.2.3 Transverse shear of core This mode is driven by insufficient core shear strength or a low panel thickness (Fig. 14).

Fig. 14 - Transverse shear failure

In many cases the dominant stress in the core material is shear, causing shear yield, ultimate failure, or tensile failure in 45° to the through thickness direction. In that case, it would be checked that the through thickness shear stress does not exceed the shear strength.

When the stress state is more complex, a simplified version of Tsai-Wu criterion could be used for some closed cells foam material: see ref [R3]

Where F1d : compressive strength F1z : tensile strength S : shear strength

6.1.2.4 Local core crushing This mode can be caused at locations with attachments to the panels by a low core compressive strength (see Fig. 15)

Fig. 15 - Local core crushing

It would be checked that the local compression stress does not exceed the compressive strength of the core material.

6.1.2.5 Global buckling of sandwich This mode can be caused by an insufficient membrane or flexural stiffness of the sandwich or an insufficient core shear rigidity (Fig. 16).

Fig. 16 - Global buckling

The global buckling mode of sandwich panels is checked with same methodologies than metallic panels.

This analysis shall evaluate carefully:

boundary conditions effects


influence of geometrical imperfections

influence of high load is that could introduce partial damage in the structure (matrix

cracking or delaminations) , modify stiffness of sandwich panels and finally reduce the

buckling loads

6.1.2.6 “local” buckling modes of sandwich The five sandwich failure modes represented schematically in Fig. 17 have to be analyzes :

wrinkling lower bound (conservative) / wrinkling intermediate value

dimpling

shear crimping

flexural core crushing

Fig. 17 - Sandwich failure modes

Dimpling occurs only with honeycomb materials (intercellular buckling).

The allowable stresses of the single failure modes are given by :

wrinkling lower bound failure (pessimistic) :

wrinkling intermediate value failure :


where

dimpling failure

shear crimping failure

flexural core crushing failure (fcc)

6.1.2.7 skin/core debonding This mode is driven by insufficient strength of the interface between skins and core of the sandwich panel.

It would be checked that :

the out of plane stress at the interface does not exceed the out of plane strength

the resultant shear at the interface 2

232

13 does not exceed the shear strength

If it can be documented that the interface is stronger than the core, core properties can be used to describe the interface. For many sandwich structures made of foam core the interface is stronger than the core.


6.1.3 Static Modelling of Joints

6.1.3.1 Bonded Joints

As mentioned before (see §6.1), laminates (as well as metal pieces) are in general modeled using 2D elements in order to improve solving time.

When bonding two surfaces, no matter if metal or laminate pieces, the adhesive will be preferably modeled with solid elements (usually 2 elements over the thickness for linear elastic evaluation), especially when thick-film bonding is used. This is shown in Fig. 18 here below:

Fig. 18 - Bonding between two pieces in Hypermesh

Between the surfaces, a so called freeze contact is defined. By this function, the bonding between adhesive and each adjacent layer/piece is defined by identical displacements of the nodes in this area.

Nevertheless, in case of large models (analysis on a full carbody for instance), simplified assumption such as no-bonded joint modelization (only perfectly tied) can be assumed in order to safe calculation time. In that case, analysis of adhesive should be performed with a detail model.

6.1.3.2 Standard bolted joints without insert

In the first step modeling, standard bolted joints without insert (such as screws, bolts and rivets) can be done in the following way. 1D elements, with certain properties and cross sections, are connecting the layers where the connection shall be realized (Beam Element with increased thickness in order to prevent artificial bending moments). To map the contact face, which would usually be the bolt head, nut or an underlying washer, so-called 1D rigid element, are used. This procedure is shown in Fig. 19 here after:

Fig. 19 - Standard bolted joints without insert in Hypermesh


After calculating, the bolt forces are obtained. With these forces, an evaluation with respect to the according limiting standard (i.e. VDI2230) and values for screws, bolts or rivets can be undertaken. With respect to composite laminates this procedure is most likely for joints where small loads are applied.

In a second step, if the bolt connection becomes critical or more relevant, preload to the bolts and contact between the surfaces can be modeled in order to evaluate the joint more detailed.

6.1.3.3 Standard bolted joints with Inserts

Joints with inserts in laminates are not commonly modeled by BT. Therefore, the following simple procedure is a suggestion and needs to be verified by tests as well.

As can be seen in the schematic drawing in Fig. 20 the insert is also modeled with 2D shell elements with metal material overlapping the 2D shell elements from the composite laminate. The bolt connection is again modeled as stated in §6.1.3.2 with 1D elements (Rigid 1D elements and 1D element BEAM). If large thicknesses (i.e. for sandwich structures) are existing it is recommended to use 3D solid elements instead of 2D shell elements to map the insert and laminate.

Fig. 20 - Schematic drawing of how to realize standard bolted joints with inserts


6.2 FATIGUE MODELLING

In principal, the distinction between endurance limit and fatigue strength can be made:

In case of endurance limit, the laminate must sustain the maximum load on a continuing basis,

In case of fatigue strength, a damage accumulation occurs during the loading. The damage itself is therefore strongly influenced by the number of cycles, the ultimate loading and the order of the loading.

Fig 21 - Typical strength and stiffness degradation in composite (schematic)

Nowadays, there do not exist general fatigue calculation and assessment methods for composites, in contrary to metallic domain.

Due to this, the general design philosophy for composite structures in industries is based on fact that delamination during cyclic loading has to be prevented.

Some stress design limits (depending on material, stacking sequence, process, etc…) based on previous experience are considered in the early phase of the design, and then has to be consolidated by fatigue tests.

For instance in the aeronautic field, the fatigue load (1 Mcycles) usually correspond to 20 to 30% of the ultimate static load of the composite carbon material, and this level is prone to be acceptable without knockdown factor after residual strength. This kind of assumption can lead to conservative design strategies which act against to the lightweight potentials of modern fiber reinforced composites. Nevertheless, it shall not be considered as a general criterion. because it could be non-conservative in some other cases (other matrix, composite material health, stacking sequences,…).

In (Degrieck & Van Paepegem, 2001), three categories of fatigue models were defined:

1. Fatigue life models : These models use information from the S-N curves or Goodman-type diagrams to propose a fatigue failure criterion. Damage accumulation is not taken into account, whereas the maximum number of cycles for fatigue failure under fixed loading conditions is predicted.

One fatigue life model usually considered is given by the Miner’s sum. In the Miner’s sum method, the results of a counting method and constant amplitude fatigue behavior description are converted into a damage parameter”, D”. Failure criterion of the laminate is considered when D>1 (see. Figure 22)


S

t

N1

N2 Ni

Fig. 22 – Miner’s sum The main limitations of this method are: o the degradation is assumed to be linear, o the potential effect of load order is not taken into account, that is

Fig. 23 – Miner’s sum and load order

As a consequence, this method can be sometimes non conservative. Moreover, the value of the damage parameter only indicates whether or not failure occurred: it does not relate to a physically quantifiable damage. Nevertheless, this method is today in use for the fatigue sizing of current area of wind blades or even helicopter blade for instance, but with integrating specific coefficient based on tests.

2. Phenomenological models to predict residual stiffness/strength : These models predict the degradation of elastic properties during fatigue loading, where a scalar damage variable D=1-E/E0 is commonly used to describe the loss of stiffness. The damage growth rate is then defined as the derivative dD/dN where N denotes the number of cycles.

3. Residual strength models: The residual strength models are distinguished between sudden death models and wearout models. When the composite is subjected to high load levels within the low-cycle fatigue regime, the residual strength is initially constant and decreases drastically when the number of cycles to failure is nearly reached. For this, the sudden death model can be used to describe this phenomenon. If the composite undergoes a state of stress at low load levels, the residual strength degrades more gradually and can be described by wearout models which incorporate the strength-life equal rank assumption, i.e. the strongest specimen has either the longest fatigue life or the highest residual strength at runout, (Degrieck & Van Paepegem, 2001).

n

i i

i

NnD

1

DA = DB 2

2Nn

1

1Nn

3

3

Nn

Loading “A” Loading “B”

1

1Nn

2

2Nn

3

3

Nn


Damage accumulationat high load level

Damage accumulationat low load level

Dam

age

Leve

l

Number of Cycles

High Loads first

Low Loads first

Fig. 24: - Percent failure rule to take into account the influence of the loading order,

the number of cycles and the ultimate level.

Nevertheless, a first step to the mathematical modeling of fatigue behavior by means of finite element method was carried out by the finite element system Abaqus by means of the Direct Cyclic Approach (DCA).

Conventional fatigue analysis determines the fatigue limit by means of well-established S-N curves but does not define the relationship between the number of cycles and the corresponding level of damage.

In DCA, the fatigue life is calculated by using an appropriate damage evolution relation until the structure’s response has stabilized after a number of cycles. Generally, the stabilized constitutive response of the structure subjected to cyclic loading conditions is achieved by applying periodic load cycles repetitively to the unstressed structure until a stabilized state is obtained. The response is calculated at discrete points along the loading history, what makes the DCA a very effective tool for modelling fatigue life, figure 25.

Fig. 25: - Elastic stiffness degradation as function of load cycles

The degradation of the material properties within the upcoming increment, which spans a number of load cycles N , are determined by utilizing the solution at each of these discrete points shown in figure 24. The degraded properties are then further used to evaluate the solution at the next increment in the load history.

The damage initiation is characterized by the accumulated inelastic hysteresis energy per cycle. This level of energy in conjunction with material constants are used to predict the number of cycles needed for the damage in initiation. The damage evolution during further cycling is extrapolated from the current cycle to the next increment of a number of cycles by

43 cNNN wc

LNDD , where c3 and c4 denote material parameters which can be

obtained from characteristic Wöhler curves of the corresponding material and L is a characteristic length associated with an integration point.


7. CONCLUSION

To perform a mechanical analysis with FEM on a composite structure, such as a carbodyshell, it is needed to make some changes on the methodology that is used for metallic structures. Indeed, heterogeneity and anisotropy of the composite material must be taken into account in the intrinsic material properties to consider in calculations and also in the analysis theory itself.

The chapter 3 of this document gives a general overview on the global methodology that is used for the characterization and testing of composite structures in aeronautic (experimental building-block approach). The important link between the manufacturing process and the material properties is also underlined, as well as the effect of the environment (hygrometry ageing, temperature, etc…).

The chapter 4 underlines specifically the variability of the sources in composite materials, which justifies the consideration of a statistical based method to establish the design values of material properties.

Based on these preliminary bases, the chapter 5 proposes a selection of relevant tests to characterize the material properties needed in the most standard structure analysis (in plane stress and strain and standard elastic theory). The test matrices proposed have to be considered for new materials and for the most generic application. They should be simplified taking into account the existing knowledge/available user data’s, and of course the specific requirements of the application. Reduced test matrix are nevertheless proposed for the screening phases of the project, when a selection between several candidates has to be done.

Besides, in chapter 5 one sub-chapter is dedicated to elementary fatigue test in current area, and another to elementary testing of some junctions. It should be noticed that those chapters are more prone to adaptation/modification depending on the application, as there is a lack of currently shared methodologies and standards in these fields over the industry.

The chapter 6 proposes some preferred failure criterions to consider in the analysis of both monolithic and sandwich structures. It also gives a brief overview on numerical methods available for calculation of composite structure in static and fatigue. Chapter 6 gives also some advice for the calculation of joints with a finite element code. It should be noticed that the calculation of fatigue resistance will be more deeply considered in T4.3, which dedicated to this task.


ANNEX 1 – DETAILED TEST PROCEDURES DESCRIPTION 1. Tensile tests

Tensile and tensile shearing tests are performed in a range of different versions.

Tensile Test on Single Filaments (not considered in this document)

The diameter of the individual filaments lies in the micrometer range. The filament is first secured to a small paper frame according to ISO 11566 and then aligned and fixed in the clamping mechanism of the testing machine. After cutting through the frame, the properties can be determined under tensile load.

Tensile Test on Filament Strands (not considered in this document)

Normally, the filament strands are coated in resin first and then cut into lengths. Cap strips made from cardboard or plastic are glued to the ends so that the tensile force can be applied evenly to the specimen. Tools such as extensometers are suitable for measuring elongation.

Fig. 1: Tensile test on filament strands

Tensile Test on Pultruded GFRP Bars (not considered in this document)

Depending on the design and surface structure of the specimen, testing is done with cap strips on the clamping ends, or without cap strips with special jaw inserts for hydraulic or pneumatic specimen grips. This test is described in ASTM D 3916.

Fig 2: Tensile Test on Pultruded GFRP Bars

Tensile Test on Unidirectional Laminates

Unidirectional laminates are normally tested longitudinally for fiber strength and transversely for bond strength. The specimens are reinforced at the ends with cap strips to avoid jaw breaks. This test, which is described in the ISO, ASTM, EN, AITM, BSS, DIN, SACMA and CRAG standards, places high demands on the quality of the extension measurement and on the alignment accuracy.

Fig 3 and 4: Tensile test on laminates

Tensile Test on Multidirectional Laminates


Depending on the size of the textile structure, multidirectional laminates are tested with large specimen widths of 25 mm or even 50 mm. According to ISO 527-4, the thickness of the specimen can also be up to 10 mm. Due to the large specimen cross-sections, very large tensile forces of over 300 kN can occur. To measure strain, strain gages, mechanical extensometers or optical extensometers can be used.

Notch Tensile Test (Open Hole Tensile)

This test characterizes the influence of a hole on the tensile strength of a laminate. The result is usually presented as a notch factor, which gives the ratio of damaged to undamaged specimen

Tensile Test on Bolted Laminates (Filled Hole Tensile)

This test uses the same specimen as the notch tensile test and the hole is closed with a threaded connection.

2. Compression tests

The compression tests are amongst the most difficult tests and are therefore sometimes described as the ultimate class of test. Various procedures have become established in practice End-loading procedure

This procedure is based on ASTM D 695 and has been further developed in various standards. It involves loading the specimen longitudinally between two pressure plates. A buckling support prevents premature failure of the material through bending.

The test consists of two parts:

To measure the compressive modulus, a specimen without cap strips is used. The strain is determined using a strain gage.

To measure the compressive strength, the specimen is reinforced with cap strips to avoid premature failure at the force transmission points.

Fig. 5: End-loading procedure


Shear loading

In this procedure, the compressive force is applied by clamping the specimen, i.e. via a frictional contact. It was standardized in the 1980s in ASTM D 3410 as the Celanese testing tool with conical clamping elements and further developed in various standards.

Under the Celanese arrangement, deviations in specimen thickness lead to unwanted linear support of the clamping elements. DIN 65375 and prEN 2850 offer modified Celanese tools with flat wedges to solve the problem of specimen thickness.

The IITRI developed a similar tool, which nowadays replaces the old Celanese compression tool in ASTM. This tool - like the predecessor model - works on the wedge clamping principle. The wedge jaws are first aligned on the specimen outside the compression tool and then placed into the compression tool.

IMA Dresden developed and patented the hydraulic compression tool HCCF. It has very good accessibility, simple handling and fixed jaws, which remain precisely aligned to each other even during the test procedure. The parallel hydraulic specimen clamping is stick-slip free, unlike the wedge-based principles. The use of a clip-on extensometer is possible.

Fig. 6: Shear loading

Combined loading

This procedure is suitable for the testing of fiber composites under the higher loads that occur with larger specimen cross-sections. Part of the compressive force is loaded via the specimen clamping, the rest on the ends of the specimen. The length of the specimen is matched precisely to the length of the jaws. Very high requirements are placed on the processing of specimen ends, as in the end-loading procedure.

ASTM D 6641 and procedure 2 according to ISO 14126 describe a mechanical testing tool comprising four elements connected to each other by guide columns. The clamping force is generated with 8 screws, which are tightened with a torque key. The test device with integrated specimen is placed between two compression plates in the testing machine for testing.

An enhancement of the procedure was laid down in the Airbus standard AITM 1.0008 edition 2010, which describes both shear loading and combined loading. The application of clamping force is described as a hydraulic parallel clamping principle. Studies showed that more valid break images are achieved when clamping peaks at the transition between free clamping length and clamping area are avoided by the structural design of the jaws. In the Airbus standard, this is described as "soft load introduction".

The HCCF test device that meets the requirements of AITM 1.0008. It was tested and approved in this regard by Airbus in Bremen.

Fig. 6: Combined loading


Notch compression test

Notch compression test (Open Hole Compression)

This test characterizes the influence of a hole on the compression strength of a laminate. The result is usually presented as a notch factor, which gives the ratio of damaged to undamaged specimen.

According to Airbus AITM 1.0008 this test is performed with the fixtures of the compression test.

ASTM and Boeing standards use a 300 mm-long specimen, which is to be clamped in a hydraulic parallel clamping tool with the aid of a supporting device.

Compression test to bolted laminates (filled hole compression)

This test uses the same specimen as the notch tensile test and the hole is closed with a

threaded connection.

3. Flexural test

3-point flexural test

Fig. 7: 3-point flexural test

3-point flexural tests are very common as they are easy to perform. Flexure can be measured with the crosshead travel encoder when the machine deformation is compensated.

The modulus is determined either between 10% and 50% Fmax (EN 2562), or 10% and 25% Fmax (EN 2746), or between two strain limits (ISO and ASTM).

The span-to-thickness ratio is 32:1 in ASTM. ISO uses 20:1 for GRFP and 40:1 for CRFP, EN standards use 16:1 for GFRP and 40:1 for CFRP. This subjects the specimen to low shear forces only.

4-point flexural test

Fig. 8: 4-point flexural test

The advantage of the 4-point flexural compared to 3 points setup is :

- freedom from shear forces in the mid-span area,

- the failure area of the specimen is not located in the area where the load is introduced.

NOTE : In the 3 points flexural tests, the local load introduction under the cylinder can cause local degradation to the material and affect the results.


The flexural modulus is determined between 0.05 and 0.25% strain (ISO 14125) or 0.1 and 0.3% strain (ASTM D 7264).

The central support span can be 1/3 (ISO 14125) or 1/2 (ASTM D 7264) of the lower support span.

The span-to-thickness ratio is 32:1 in ASTM. ISO uses 22.5:1 for GFRP and 40.5:1 for CFRP.

4. Interlaminar Shear Strength (ILSS)

Fig. 9: Interlaminar Shear Strength (ILSS)

ILSS tests are a typical procedure for monitoring quality and are suitable for the comparison of materials.

The procedure supplies only apparent shear properties, as peak stresses occur near to the loading fin.

ILSS tests can be performed by using a sufficiently rigid 3-point bending fixture. The span-to-thickness ratio of 10 mm is very short. This causes high shearing forces and relatively low bending moments in the specimen.

5. Lap shear / In-plane shear

The adhesive strength of a composite is typically characterized by shear tests. Different types are used to determine the shear properties in single directions.

Lap shear (not considered in this document)

Lap shear tests are suitable for comparative test on laminate adhesives, e.g. with film or prepregs, as well as for assessing the bond in a laminate.

Horizontally adjustable grips are needed to test simple single-lap shear specimens. These can be wedge screw grip, screw or pneumatic versions.

Slotted single-lap shear specimens can be tested with symmetrically closing grips. The EN and the DIN standard specify a support to prevent bending.

Double-lap shear specimens can be tested with simple grips. Slotted double-lap shear specimens allow the same grip opening on both sides.

The result is an in-plane shear strength.

Fig. 10: Lap shear


In-plane shear

In-plane shear can be produced by performing a tensile or compression test at ± 45° to the fiber direction.

The specimens are cut from plates at 45° to the fiber direction.

DMS or extensometers are used to measure longitudinal and transverse extension.

The test method is suitable only for extensions of less than 5%.

Fig. 11: In-plane shear

6. V-notch shear

This test arrangement is used both for in-plane and for interlaminar shear. Each of the six possible shear planes can be tested separately. Two characteristics of the procedure are standardized

Losipescu Procedure (not considered in this document)

In this test arrangement, which is described in ASTM D 5379, a specimen notched on both sides is clamped in a special device which is held longitudinally.

When compressed, this creates a zone of torque-free shear load between the notches.

The fibers must lie parallel or perpendicular to the loading axis.

Strain gages are placed at less than 45° in the direction of the shear plane in order to determine the shear extension.

Results are shear response, 0.2% offset stress, max. shear stress and secant shear modulus.

Fig. 12: Losipescu Procedure

V-notch rail shear (not considered in this document)

This procedure is laid down in ASTM D 7078. Compared to the Iosipescu procedure, the shear surface is relatively large.

Results are shear response, 0.2% offset stress, max. shear stress and secant shear modulus.


Fig. 13: V-notch rail shear

7. Compression after impact (CAI)

The test sequence in the Compression After Impact test comprises two parts:

The deliberate pre-damaging of a specimen with the aid of an instrumented drop weight tester.

The static compression test for measuring the residual strength.

Pre-damage

In the instrumented drop weight tester HIT 230F, the specimen is pre-damaged under set conditions.

The specimens are tensioned on a section of:

76.2 x 127 mm (ASTM, Boeing, SACMA, DIN), 75 x 125 mm (EN, Airbus) or 140 mm diameter (CRAG). Only the Airbus AITM requires clamping within the section. To simplify operation, the specimens are tensioned outside the drop weight tester and then inserted into the testing position.

The pre-damaging procedure can be monitored and assessed using the instrumentation of the drop weight tester. The first damage peak on the power-time curve also gives a correlation to the Mode II fracture toughness of the laminate.

CAI compression test

The pre-damaged specimens are tested in a special compression tool to establish the residual compression strength.

The resulting compression forces are usually very large.

To load the test plates without buckling, special compression tools are used that are distinct within the standardization:

ASTM, Boeing, SACMA and DIN: All four sides are guided, but not clamped.

ISO, EN and Airbus standards: The upper and lower ends of the specimen are clamped. The sides are guided with line contact

Fig. 14: Compression after impact (CAI) test


8. Fracture Toughness

Essentially, a distinction is made between three types of mode in the fracture mechanics:

- Mode I: Crack opening

- Mode II: In-plane shear

- Mode III: Out-of-plane shear

Fig. 15: Fracture Modes

Mode I

There are various test arrangements for the individual modes. Mode I and Mode II are usual test procedures, as is Mixed Mode Bending

The test is normally performed with the DCB (Double Cantilever Beam) specimen.

Here, the crack opening is measured as crosshead travel.

The crack's growth is visually tracked on both sides of the specimen.

The test procedures and result evaluations differ depending on the standard applied.

Standards are: ISO 15024, ASTM D 5528, AITM 1-0005, AITM 1-0053, Boeing BSS 7273, CRAG method 600, NASA method RP 1092 ST-5, ESIS TC 4, prEN 6033 (withdrawn)

Fig. 16: Fracture Mode I test.

Mode II

Mode II loads can be generated and measured both in the flexural test, as well as in the tensile and compression test with notched specimens.

The measurement of the Mode II energy release rate is standardized as a flexural test.

The specimens are designated with SENB (Single End Notch Bending), but ENF (End Notch flexure) is another common term.

Deflection is measured by the crosshead travel (with stiffness correction), or by means of a displacement transducer, which is applied centrally.

The crack initiation point is characterized by a force maximum.


At the end of the measurement, the specimen is cooled in liquid nitrogen and then completely broken to measure the fracture surfaces.

Fig. 17: Fracture Mode II test.

Mixed Mode Bending (MMB) (not considered in this document)

"Mixed Mode Bending" can be measured on unidirectional laminates. This involves combining Mode I and Mode II.

9. Fatigue test

Fatigue behavior in fiber composites is measured under a range of loads: as a tensile test, an in-line shear test or at screw or bolt connections.

Different machine types are possible:

Servohydraulically dynamic testing machines, which are able to generate variable frequencies and amplitudes under different control modes.

Vibrophores, which run at constant frequency in sine-loading. This can be a very affordable type of test for thin-walled specimens, for which adequate heat dissipation is guaranteed.


ANNEX 2 – STANDARDS FOR COMPOSITE TEST Type of test Standards Description

Tensile Test ISO 527 - 4 Plastics - Determination of tensile properties - Part 4: Test conditions for isotropic and anisotropic fibre-reinforced plastic composites

Tensile Test ISO 527 - 5 Kunststoffe - Bestimmung der Zugeigenschaften - Teil 5: Prüfbedingungen für unidirektional faserverstärkte Kunststoffverbundwerkstoffe

Tensile Test ISO 4899 Textile-glass-reinforced thermosetting plastics -- Properties and test methods

Tensile Test ISO 11566 Carbon fibre -- Determination of the tensile properties of single-filament specimens

Tensile Test ASTM D 3039 Tensile Properties of Polymer Matrix Composite Materials Tensile Test ASTM D 4018 Continuous Filament Carbon and Graphite Fiber Tows

Tensile Test ASTM D 3916 Properties of Pultruded Glass-Fiber-Reinforced Plastic Rod

Tensile Test ASTM D 5083 Tensile Properties of Reinforced Thermosetting Plastics Using Straight-Sided Specimens

Tensile Test ASTM D 7205 Tensile Properties of Fiber Reinforced Polymer Matrix Composite Bars

Tensile Test DIN 65378 Prüfung von unidirektionalen Laminaten; Zugversuch quer zur Faserrichtung

Tensile Test DIN 65469 Faserverstärkte Kunststoffe; Zugversuch an einlagigen Zugflachprobekörpern Fibre-reinforced plastics; tensile test of monolayer flat tension specimens

Tensile Test EN 2561 Unidirektionale Laminate - Zugprüfung parallel zur Faserrichtung Unidirectional laminates - Tensile test parallel to the fibre direction

Tensile Test EN 2597 Unidirectional laminates - Tensile test perpendicular to the fibre direction

Tensile Test EN 2747 Glass fibre reinforced plastics. Tensile test

Tensile Test prEN 6035 Faserverstärkte Kunststoffe - Prüfverfahren; Bestimmung der Kerbzugfestigkeit an gekerbten und ungekerbten Probekörpern

Tensile Test DIN 29971 Luft- und Raumfahrt; Unidirektionalgelege-Prepreg aus Kohlenstoffasern und Epoxidharz; Technische Lieferbedingungen

Tensile Test Airbus AITM 1.0007

Faserverstärkte Kunststoffe - Bestimmung der Zugfestigkeit an ungekerbten, offen und geschlossen gekerbten Zugproben Fibre reinforced plastics: determination of notched, unnotched and filled hole tensile strength

Tensile Test Boeing BSS 7320 Tensile Testing of Advanced Composites

Tensile Test SACMA SRM 4R-94 Tensile Properties of Oriented Fiber-Resin Composites

Tensile Test SACMA SRM 9-94 Tensile Properties of Oriented Cross-Plied Fiber-Resin Composites

Tensile Test TR 88012 CRAG Methods 300-303 Tensile Testing

Compression test ISO 14126

Faserverstärkte Kunststoffe - Bestimmung der Druckeigenschaften in der Laminatebene Fibre-reinforced plastic composites -- Determination of compressive properties in the in-plane direction

Compression test ISO 604 Plastics -- Determination of compressive properties

Compression test ISO 8515 Textile-glass-reinforced plastics -- Determination of compressive properties in the direction parallel to the plane of lamination

Compression test ISO 3597-3 Textile-glass-reinforced plastics -- Determination of mechanical properties on rods made of roving-reinforced resin -- Part 3: Determination of compressive strength


Compression test ASTM D 3410 Compressive Properties of Polymer Matrix Composite Materials with Unsupported Gage Section by Shear Loading (IITRI - setup)

Compression test ASTM D 695 Compressive Properties of Rigid Plastics

Compression test ASTM D 6641 Compressive Properties of Polymer Matrix Composite Materials Using a Combined Loading Compression (CLC) Test Fixture

Compression test ASTM C 1358

Monotonic Compressive Strength Testing of Continuous Fiber-Reinforced Advanced Ceramics with Solid Rectangular Cross-Section Test Specimens at Ambient Temperatures

Compression test DIN 65375 Prüfung von unidirektionalen Laminaten; Druckversuch quer zur Faserrichtung

Compression test DIN V 65380 Prüfung von unidirektionalen Laminaten; Druckversuch parallel und quer zur Faserrichtung

Compression test prEN 2850

Unidirektionale Laminate aus Kohlenstoffasern und Reaktionsharz - Druckversuch parallel zur Faserrichtung Carbon fibre thermosetting resin unidirectional laminates - Compression test parallel to fibre direction

Compression test JIS K 7076 Testing methods for compressive properties of carbon fibre reinforced plastics

Compression test AITM 1-0008 Fibre Reinforced Plastics - Determination of Plain, Open Hole and Filled Hole Compression Strength.

Compression test Airbus QVA-Z10-46-38

Bestimmung von Druckeigenschaften von CFK-Laminaten aus Gewebe-Prepreg, Prüfung in Kett- oder Schussrichtung

Compression test Boeing BSS 7260 - type III and IV Advanced Composite Compression Tests

Compression test SACMA SRM 1R-94

Compressive Properties of Oriented Fiber-Resin Composites

Compression test SACMA SRM 6-94 Compressive Properties of Oriented Cross-Plied Fiber-Resin Composites

Compression test RAE-TR 88012 CRAG Method 400

Method of test for longitudinal compression strength and modulus of unidirectional fibre reinforced plastics

Compression test RAE-TR 88012 CRAG Method 401 Plain Compression

Compression test ISO 3597-3 Textile-glass-reinforced plastics -- Determination of mechanical properties on rods made of roving-reinforced resin -- Part 3: Determination of compressive strength

Compression after indentation ASTM D 6264

Measuring the Damage Resistance of a Fiber-Reinforced Polymer-Matrix Composite to a Concentrated Quasi-Static Indentation Force

Compression after Impact ISO 18352

Carbon-fibre-reinforced plastics -- Determination of compression-after-impact properties at a specified impact-energy level

Compression after Impact ASTM D 7137 Compressive Residual Strength Properties of Damaged

Polymer Matrix Composite Plates

Compression after Impact prEN 6038 Bestimmung der Restdruckfestigkeit nach

Schlagbeanspruchung Compression after Impact AITM 1.0010 Determination of compression strength after impact-stress

Compression after Impact

Boeing BSS 7260 - type II Compression after impact

Compression after Impact CRAG method 403 Compression after impact


SACMA SRM 2R-94

Compression after Impact Properties of Oriented Fibre-Resin Composites

Compression after Impact DIN 65561 Prüfung von multidirektionalen Laminaten; Bestimmung

der Druckfestigkeit nach Schlagbeanspruchung


NASA RP 1092 ST-1 Compression after Impact


Open hole / Filled hole tensile ASTM D 5766 Open-Hole Tensile Strength of Polymer Matrix Composite

Laminates

Open hole / Filled hole tensile ASTM D 6742 Filled-Hole Tension and Compression Testing of Polymer

Matrix Composite Laminates

Open hole / Filled hole tensile

Airbus AITM 1.0007

Faserverstärkte Kunststoffe - Bestimmung der Zugfestigkeit an ungekerbten, offen und geschlossen gekerbten Zugproben Fibre reinforced plastics: determination of notched, unnotched and filled hole tensile strength

Open hole / Filled hole tensile SACMA SRM 5-94 Open Hole Tensile Properties of Advanced Composites

Materials Open hole / Filled hole tensile

NASA RP 1092 ST-3 Open hole tension

Open hole / Filled hole compression prEN 6036 Bestimmung der Kerbdruckfestigkeit an gekerbten,

ungekerbten und gebolzten Probekörpern

Open hole / Filled hole compression

ASTM D 6484 (Boeing Document 888-10026, Boeing Document 6-83079-71)

Open-Hole Compressive Strength of Polymer Matrix Composite Laminates

Open hole / Filled hole compression ISO/DIS 12817 Fiber-reinforces plastic composites - Determination of

open-hole compression strength Open hole / Filled hole compression

Boeing BSS 7260 - Type 1 Open Hole Compression Test

Open hole / Filled hole compression AITM 1-0008 Fibre Reinforced Plastics - Determination of Plain, Open

Hole and Filled Hole Compression Strength. Open hole / Filled hole compression

SACMA SRM 3R-94 Open Hole Compression Properties


NASA RP 1092 ST-4 Open hole compression


RAE-TR 88012 CRAG Method 402 Open hole compression strength


Northrop NAI-1504C Open Hole Compression Test Method

Flexural tests ISO 14125 Fibre-reinforced plastic composites - Determination of flexural properties Faserverstärkte Kunststoffe - Bestimmung der Biegeeigenschaften

Flexural tests EN 2562 Unidirectional laminates - Flexural tests parallel to the fibre direction

Flexural tests EN 2746 Glass Fibre Reinforces Plastics - Flexural Tests, Three Point Bend Method

Flexural tests EN 13706-2 Spezifikationen für pultrudierte Profile Teil 2: Prüfverfahren und allgemeine Anforderungen

Flexural tests ISO 3597-2 Textile-glass-reinforced plastics -- Determination of mechanical properties on rods made of roving-reinforced resin -- Part 2: Determination of flexural strength

Flexural tests ASTM D 4476 Flexural Properties of Fiber Reinforced Pultruded Plastic Rods

Flexural tests ASTM D 790 Flexural Properties of Unreinforced and Reinforced Plastics and Eletrical Insulating Materials

Flexural tests ASTM D 6272 Flexural Properties of Unreinforced and Reinforced Plastics and Electrical Insulating Materials by Four-Point Bending

Flexural tests ASTM D 7264 Flexural Properties of Polymer Matrix Composite Materials

Flexural tests ASTM D 6415 Measuring the Curved Beam Strength of a Fiber-Reinforced Polymer-Matrix Composite1

Flexural tests TR 88012 CRAG Method 200 Flexural modulus and strength of reinforced plastics

Flexural tests HSR/EPM-D-003-93 Four Point Flexural Testing of Composite Materials

Flexural tests DIN 53390 BIEGEVERSUCH AN UNIDIREKTIONAL


GLASFASERVERSTAERKTEN RUNDSTAB-LAMINATEN

Interlaminar Shear Strength ISO 14130 Determination of apparent interlaminar shear strength by

short beam-method

Interlaminar Shear Strength ISO 3597-4

Textile-glass-reinforced plastics -- Determination of mechanical properties on rods made of roving-reinforced resin -- Part 4: Determination of apparent interlaminar shear strength

Interlaminar Shear Strength ASTM D 2344 Short-Beam Strength of Polymer Matrix Composite

Materials and Their Laminates Interlaminar Shear Strength ASTM D 4475 Apparent Horizontal Shear Strength of Pultruded

Reinforced Plastic Rods By the Short-Beam Method

Interlaminar Shear Strength EN 2377

Glasfaserverstärkte Kunststoffe; Prüfverfahren zur Bestimmung der scheinbaren interlaminaren Scherfestigkeit Glass fibre reinforced plastics; test method; determination of apparent interlaminar shear strength

Interlaminar Shear Strength EN 2563

Kohlenstoffaserverstärkte Kunststoffe - Unidirektionale Laminate; Bestimmung der scheinbaren interlaminaren Scherfestigkeit Carbon fibre reinforced plastics - Unidirectional laminates; determination of apparent interlaminar shear strength

Interlaminar Shear Strength JIS K 7078 Apparent Interlaminar Shear Strength of Carbon Fiber

Reinforced Plastics by Three Point Loading Method Interlaminar Shear Strength SACMA SRM 8-88 Short Beam Shear Strength of Oriented Fiber-Resin

Composites

Interlaminar Shear Strength CRAG method 100 Interlaminar shear strength (ILSS) of reinforced plastics

In-plane shear by Lap-Shear EN 2243-1 Structural Adhesives Test Methods Part 1 - Single Lap

Shear, Aerospace Series

In-plane shear by Lap-Shear EN 2243-6 Structural adhesives. Test methods. Part 6. Determination

of shear stress and shear displacement In-plane shear by Lap-Shear pr EN 6060 Determination of the Tensile Single Lap Shear Strength

In-plane shear by Lap-Shear AITM 1-0019 Determination of lap shear tensile strength of composite

joints In-plane shear by Lap-Shear

Airbus QVA-Z10-46-09

Determination of Interlaminar Tensile Shear Strength of Fiber Composite Structures

In-plane shear by Lap-Shear

Airbus QVA-Z10-46-01 Determination of the Bond Strength of Adhesives

In-plane shear by Lap-Shear DIN 65148 Bestimmung der interlaminaren Scherfestigkeit im

Zugversuch

In-plane shear by Lap-Shear ASTM D 3914 In-Plane Shear Strength of Pultruded Glass-Reinforced

Plastic Rod In-plane shear by Lap-Shear ASTM D 3846 Standard Test Method for In-Plane Shear Strength of

Reinforced Plastics

In-plane shear by Lap-Shear CRAG method 102

In-Plane Shear by ± 45° Laminates test ISO 14129 Bestimmung des Schermoduls und der Scherfestigkeit in

der Lagenebene mit dem ± 45° Zugversuch

In-Plane Shear by ± 45° Laminates test prEN 6031 Determination of in-plane shear properties (±45° tensile

test) In-Plane Shear by ± 45° Laminates test ASTM D 3518 In-Plane Shear Response of Polymer Matrix Composite

Materials by Tensile Test of a ±45° Laminate

In-Plane Shear by ± 45° Laminates test DIN 65466 Prüfung von unidirektionalen Laminaten; Bestimmung der

Schubfestigkeit und des Schubmoduls im Zugversuch In-Plane Shear by ± 45° Laminates test AITM 1-0002 Determination of in-plane shear properties (±45° tensile

test)

In-Plane Shear by ± 45° Laminates test

Airbus QVA-Z10-46-22

Bestimmung der Schubeigenschaften im Zugversuch von ± 45° CFK–Laminaten aus unidirektionalem Prepreg (Tape) oder Gewebe–Prepreg

In-Plane Shear by ± 45° SACMA SRM 7-94 Inplane Shear Properties (±45?)


Laminates test

In-Plane Shear by ± 45° Laminates test

RAE TR 88012 CRAG Method 101 In-plane shear

In-Plane Shear by ± 45° Laminates test JIS K 7079

In-Plane Shear Properties of Carbon Fiber Reinforced Plastics by Plus or Minus 45 Degrees Tension Method and Two Pairs of Rails Method

In-Plane Shear by ± 45° Laminates test ASTM D 3044 Shear Modulus of Wood-Based Structural Panels

In-Plane Shear by plate twist method ISO/CD 15310 Fibre-reinforced plastic composites -- Determination of the

in-plane shear modulus by the plate twist method

Specific Shear Tests ASTM D 4255 In-Plane Shear Properties of Polymer Matrix Composite Materials by the Rail Shear Method

Specific Shear Tests ASTM D 5379 Shear Properties of Composite Materials by the V-Notched Beam Method (Iosipescu)

Specific Shear Tests ASTM D 7078 Shear Properties of Composite Materials by V-Notched Rail Shear Method

Specific Shear Tests ISO 15310 Fibre-reinforced plastic composites -- Determination of the in-plane shear modulus by the plate twist method

Specific Shear Tests DIN 53399-2 Testing of reinforced plastics; shear test on plane specimens Prüfung von faserverstärkten Kunststoffen; Schubversuch an ebenen Probekörpern

Fracture mechanics ISO 13586 Plastics -- Determination of fracture toughness (GIC and KIC) -- Linear elastic fracture mechanics (LEFM) approach

Fracture mechanics ISO 15024 Fibre-reinforced plastic composites -- Determination of mode I interlaminar fracture toughness, GIC, for unidirectionally reinforced materials

Fracture mechanics ISO 17281 Plastics -- Determination of fracture toughness (GIC and KIC) at moderately high loading rates (1 m/s)

Fracture mechanics ISO/CD 15114 Fibre-reinforced plastic composites -- Determination of apparent mode II interlaminar fracture toughness for unidirectionally reinforced materials

Fracture mechanics ESIS TC 4 Protocol for interlaminar fracture testing of composites

Fracture mechanics pr EN 6033 Draft Document - Aerospace series - Carbon fibre reinforced plastics - Test method; determination of interlaminar fracture toughness energy; mode I, GIC

Fracture mechanics prEN 6034

Draft Document - Aerospace series - Carbon fibre reinforced plastics - Test method; determination of interlaminar fracture toughness energy; mode II, G<(Index)IIC>

Fracture mechanics ASTM D 5045 Plane-Strain Fracture Toughness and Strain Energy Release Rate of Plastic Materials

Fracture mechanics ASTM D 5528 Mode I Interlaminar Fracture Toughness of Unidirectional Fiber-Reinforced Polymer Matrix Composites

Fracture mechanics ASTM D 6068 Determining J-R Curves of Plastic Materials

Fracture mechanics ASTM D 6671 Mixed Mode I-Mode II Interlaminar Fracture Toughness of Unidirectional Fiber Reinforced Polymer Matrix Composites

Fracture mechanics ASTM WK22949

New Test Method for Determination of the Mode II Interlaminar Fracture Toughness of Unidirectional Fiber-Reinforced Polymer Matrix Composites Using the End-Notched Flexure (ENF) Test

Fracture mechanics ASTM E 1922 Translaminar Fracture Toughness of Laminated and Pultruded Polymer Matrix Composite Materials

Fracture mechanics CRAG method 600 Method of Test for Interlaminar Fracture Toughness of Fibre Reinforced Composites

Fracture mechanics AITM 1.0005 Determination of interlaminar fracture toughness energy. Mode I.

Fracture mechanics AITM 1.0053 Determination of fracture toughness energy of bonded joints, Mode I, GIC

Fracture mechanics AITM 1.0006 Determination of interlaminar fracture toughness energy. Mode II


Fracture mechanics Boeing BSS 7273 Fracture mechanics Boeing BMS 8-276

Fracture mechanics NASA method RP 1092 ST-5 Hinged double cantilever beam

Fracture mechanics DIN 65563

Thin wall Cylinder ASTM D 5448 In-Plane Shear Properties of Hoop Wound Polymer Matrix Composite Cylinders

Thin wall Cylinder ASTM D 5449 Transverse Compressive Properties of Hoop Wound Polymer Matrix Cylinders

Thin wall Cylinder ASTM D 5450 Transverse Tensile Properties of Hoop Wound Polymer Matrix Cylinders

Lochleibung / Verbindungselemente ASTM D 5961 Bearing Response of Polymer Matrix Composite

Laminates

Lochleibung / Verbindungselemente ASTM D953-02 Bearing strength

Lochleibung / Verbindungselemente ISO/DIS 12815 Fibre-reinforced plastic composites - Determination of

plain-pin bearing strength

Lochleibung / Verbindungselemente DIN 65562 Bestimmung der Lochleibungsfestigkeit

Lochleibung / Verbindungselemente pr EN 6037 Bearing strength

Lochleibung / Verbindungselemente EN 13706-2 Spezifikationen für pultrudierte Profile Teil 2:

Prüfverfahren und allgemeine Anforderungen

Lochleibung / Verbindungselemente AITM 1-0009

Determination of Bearing Strength by either Pin or Bolt Bearing Configuration. Faserverstärkte Kunststoffe; Bestimmung der Lochleibungsfestigkeit mit einer Stift oder Schraubenversuchsanordnung

Lochleibung / Verbindungselemente AITM 1-0065 Fiber reinforced plastics-Determination of joint strength of

mechanically fastened joints Lochleibung / Verbindungselemente

TR 88012 CRAG Method 700 Bearing strength

Lochleibung / Verbindungselemente SACMA SRM 9-89 Bearing Strength Properties of Oriented Fiber-Resin

Composites Lochleibung / Verbindungselemente ASTM D 7248 Bearing/Bypass Interaction Response of Polymer Matrix

Composite Laminates Using 2-Fastener Specimens

Pull tests ASTM D 7332 Fastener Pull-Through Resistance of a Fiber-Reinforced Polymer Matrix Composite

Pull tests ASTM D 7522 Pull-Off Strength for FRP Bonded to Concrete Substrate

Fatigue ISO 13003 Fibre-reinforced plastics -- Determination of fatigue properties under cyclic loading conditions

Fatigue ASTM D 3479 Tension-Tension Fatigue of Advanced Composites

Fatigue ASTM D 6873 Bearing Fatigue Response of Polymer Matrix Composite Laminates

Fatigue ASTM D 671 Flexural Fatigue of Plastics

Fatigue HSR/EPM-D-002-93

Tension-Tension Load Controlled Fatigue Testing of Composite Materials Thermal Mechanical Fatigue (TFM)

Zug ASTM C297 Standard Test Method for Flatwise Tensile Strength of Sandwich Constructions

Zug AITM 1-0025 Flatwise tensile test of composite sandwich panel

Zug prEN 6062 Fibre Reinforced plastics- Test method- Flatwise tensile test of composite sandwich panel.

Druck ASTM C365 Flatwise Compressive Strength of Honeycombs Druck ASTM D351 Edge Strength of Honeycombs

Druck ASTM D 5467 Compressive Properties of Unidirectional Polymer Matrix Composites Using a Sandwich Beam

Scher ASTM C273 Core Shear of Sandwich Honeycombs

Scher Airbus QVA-Z10-46-06

Determination of the Shear Strength of Joint Adhesives in Shear Test of Pipes


Scher DIN 53294 Kernverbunde, Schubversuch

Biege ASTM C393 Flexural Properties of Sandwich Honeycombs Biege DIN 53293 Biegeversuch

Biege AITM 1.0018 Sandwich flexural test - 4 point bending Peel ASTM D1781 Climbing Drum Peel

Peel ASTM D1876 T-Peel Test Peel ASTM D3167 Floating Roller Peel

Peel Airbus QVA-Z10-46-05

Determination of Drum Peeling Force of Adhesives and Adhesive Prepregs During Drum Peeling Test

Peel Airbus QVA-Z10-46-03

Determination of the Peel Strength of Adhesives in Floating Roller Peel Tests (Bell)

Fatigue ASTM C394 Standard Test Method for Shear Fatigue of Sandwich Core Materials


ANNEX 3 – COMPLEMENTARY REFERENCES 6.13, A. (n.d.). Abaqus Anbalysis User's Guide - 6.2.7. Low-cycle fatigue analysis.

(2004). AITM 1-0007 - Determination of Plain, Open Hole and Filled Hole Tensile Strength. Airbus.

(2010). AITM 1-0008 - Determination of Plain, Open Hole and Filled Hole Compression Strength. Airbus.

(2005). AITM 1-0010 - Determination of Compression Strength After Impact. Airbus.

(1997). AITM 1-0019 - Determination of Tensile Lap Shear Strength of Composite Joints. Airbus.

(2009). AITM 1-0065 - Fiber reinforced plastics. Determination of joint strength of mechanically fastend

joints. Airbus.

(2000). ASTM D 3165-00 - Standard Test Method for Strength Properties of Adhesives in Shear by

Tension Loading of Single-Lap-Joint Laminated Assemblies. PA 19428-2959, United States.

(2008). ASTM D 3528-96 - Standard Test Method for Strength Properties of Double Lap Shear Adhesive

Joints by Tension Loading. PA 19428-2959, United States.

(2002). ASTM D 5528-01 - Standard Test Method for Mode I Interlaminar Fracture Toughness of

Unidrectional Fiber-Reinforced Polymer Matrix Composites. PA 19428-2959, United States.

(2006). ASTM D 6671/D 6671M - 06 - Standard Test Method for Mixed Mode I-Mode II Interlaminar

Fracture Toughness of Unidirectional Fiber Reinforced Polymer Matrix Composites. PA 19428-2959,

United States.

Barenblatt, G. (1962). Mathematical theory of equilibrium of cracks in brittle fracture. Advances in

Applied Mechanics, pp. 55 - 129.

Böhlke, T. (2001). Crystallographic Texture Evolution and Elastic Anisotropy - Simulation, Modeling and

Applications. Shaker Verlag.

Camanho, P., & Davila, C. (2002). Mixed-mode decohesion finite elements for the simulation of

delamination in composite materials. NASA/TM, pp. 1-42.

Camanho, P., Dávila, C., & Ambur, D. (2001). Numerical Simulation of Delamination Growth in

Composite Materials. NASA/TP - 2001 - 211041, pp. 1-24.

Camanho, P., Davila, C., & de Moura, M. (2003). Numerical simulation of mixed-mode progressive

delamination in composite materials. J. Comp. Mat., pp. 1415-1438.

Degrieck, J., & Van Paepegem, W. (2001). Fatigue damage modeling of fibre-reinforced composite

materials: Review. Applied Mechanics Reviews, pp. 279-300.

(1978). DIN 50100 - Dauerschwingversuch. FNM im DIN Deutsches Institut für Normung e.V.

(2009). DIN EN 1465. Berlin: DIN Deutsches Institut für Normung e.V.

(1997). DIN EN ISO 527-4. Berlin: DIN Deutsches Institut für Normung e.V.

(1997). DIN EN ISO 527-5. Berlin: DIN Deutsches Institut für Normung e.V.

Dugdale, D. (1960). Yielding of steel sheets containing slits. Journal of the Mechanics and Physics of

Solids, pp. 100 - 104.

Eshelby, J. (1957). The determination of the elastic field of an ellipsoidal inclusion and related problems.

Proceedings of the Royal Society London - A(241), pp. 367-396.

Ghosh, S., Ling, Y., Majumdar, B., & Kim, R. (2000). Interfacial debonding analysis in multiple fiber

reinforced composites. Mech. Mat., pp. 561-591.

Gol`denblat, I., & Kopnov. (1965). Strength of glass-reinforced plastics in the complex stress state.

Mekhanika Polimerov, pp. 54-59.

Hashin, Z. (1980). Failure Criteria for Unidirectional Fiber Composites. Journal of Applied Mechanics,

pp. 329-334.

Hashin, Z., & Rotem, A. (1973). A Fatigue Criterion for Fiber-Reinforced Materials. Journal of Composite

Materials, pp. 448-464.


Hashin, Z., & Shtrikman, S. (1962). A variational approach to the theory of the elastic behaviour of

polycrystals. J. Mech. Phys. Solids., pp. 343 - 352.

Hashin, Z., & Shtrikman, S. (1963). A variational approach to the theory of elastic behaviour of

multiphase materials. J. Mech. Phys. Solids, pp. 127 - 140.

Hill, R. (1952). The elastic behaviour of a crystalline aggregate. Proceedings of the Royal Society

London - A(65), pp. 349-354.

Irwin, G. (1957). Analysis of stresses and strains near the end of a crack traversing a plate. Journal of

Applied Mechanics, pp. 361-364.

(2003). ISO 13003:2003 (E) - Fibre-reinforced plastics - Determination of fatigue properties under cyclic

loading conditions.

Jochum, C., & Grandidier, J. (2004). Microbuckling elastic modelling approach of a single carbon fibre

embedded in an epoxy matrix. Comp. Sci. Tech., pp. 2441-2449.

Knops, M. (2008). Analysis of Failure in Fiber Polymer Laminates. Springer.

Knops, M., & Böge, C. (2006). Gradual failure in fibre/polymere laminates. Composite Science and

Technology, pp. 616-625.

Kouznetsova, V. (2002). Computational homogenization for the multi-scale analysis of multiphase

materials. Technische Universität Eindhoven.

Lubarda, V. (2002). Elastoplasticity Theory. CRC Press.

Martinez, X., & Oller, S. (2009). Numerical simulation of matrix reinforced composite materials subjected

to compression loads. Arch. Comput. Meth. Engng., pp. 357 - 397.

Mazars, J., & Pijaudier-Cabot, G. (1989). Continuum damage theory - application to concrete. J. Engng.

Mech., pp. 345-365.

Needleman, A. (1987). A continuum model for void nucleation by inclusion debonding. Journal of

Applied Mechanics, pp. 525-531.

Nemta-Nasser, S. (1999). Averaging theorems in finite deformation plasticity. Mechanics of Materials,

pp. 493-523.

Ortiz, M., & Pandolfi, A. (1999). Finite-deformation irreversible cohesive elements for three dimensional

crack-propagation analysis. Int. J. Num. Meth. Engng., pp. 1267-1282.

Reuss, A. (1929). Berechnung der Fliessgrenze von Mischkristallen auf Grund der Plastizitätsgrenze.

Journal of Applied Mechanics(35), pp. 379-386.

Rice, J. (1968). A path independent integral and the approximate analysis of strain concentration by

notches and cracks. Journal of Applied Mechanics, pp. 379-386.

Rosen, B. (1965). Fibre composite materials. Am. Soc. Metals, pp. 37-45.

Stadie-Frobös, G. (2000). Mikromechanische Beschreibung von duktiler Schädigung mit der Technik

des repräsentativen Volumenelementes. VDI Verlag.

Voigt, W. (1889). Über die Beziehung zwischen den beiden Elastizitätskonstanten isotroper Körper.

Wiedmanns Annalen der Physik, pp. 573-587.

Xu, X., & Needleman, A. (1993). Void nucleation by inclusion debonding in a crystal matrix. Modelling

Simul. Mater. Sci. Engng., pp. 111-132.

Xu, X., & Needleman, A. (1995). Numerical simulations of dynamic interfacial crack growth allowing for

crack growth away from the bond line. International Journal of Fracture, pp. 253-275.

WP 4 - · PDF filea conclusive chapter S ... of a polymer matrix composites system for a structural ... of internal loads and structural integrity. Compare to analysis

Documents