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HELSINKI UNIVERSITY OF TECHNOLOGY Department of Mechanical Engineering

Juha H a l m e

Development Testing of a Composite Wing Rib

Thesis submitted in partial fulfillment of the requirements for the degree of Master of Science in Technology

Espoo, 6 May, 2002

Supervisor Professor Olli Saarela Instructor Jyrki Laitila Master of Science (Technology)

HELSINKI UNIVERSITY OF TECHNOLOGY

ABSTRACT OF THE MASTER'S THESIS

Author: Juha H a l m e Title of the thesis: Development Testing of a Composite Wing Rib Date: 6.5.2002 Department: Mechanical engineering Professorship: Kul-34 Aeronautical Engineering Supervisor: Professor Olli Saarela Number of pages: 127

Instructor: Jyrki Laitila Master of Science (Technology)

The objective of this thesis was to define, design and execute a testing program for 5 composite wing ribs in the development design phase of the TANGO project. TANGO is a multinational European aerospace program involved by the key European airframe manufacturers and their major suppliers. The wing rib development test program needed to be consistent with the testing philosophy that was defined for the TANGO components. As a part of the work the loading of the composite wing ribs was studied. The wing ribs are finally manufactured by the Resin Transfer Moulding (RTM) manufacturing method. Basic steps of the RTM method are described. In the material screening part manufacturing and mechanical performance of 7 different thermoset resin and carbon fibre reinforcement material combinations were studied. As a result of the material screening program a single material combination was selected as the wing rib material. The required design data properties were defined and measured for the selected material. Fibre compaction tests were conducted in order to assess the mould closing forces in the RTM process. Limitations of the Non-Destructive Inspection of complex composite web geometry were studied. During the conceptual design of the ribs a testing fixture for producing multiaxial loading for rib panels was designed. Also significant upcoming element tests defined in the test program are outlined. As a result of the material screening program the G0926INJ reinforcement in the RTM6 matrix was selected for the wing rib materials.

Development testing of a composite wing rib -

Preface

Preface

This thesis work was completed at the Helsinki University of Technology in the Laboratory of Lightweight Structures for Patria Finavicomp Oy in the TANGO project. A great deal of new contacts as well as knowledge of this field of technology was gained. For this, I would like to like thank my Supervisor Professor Olli Saarela for directing this opportunity into my hands and also for his patience during the work. I also would like to thank my instructor Jyrki Laitila and the whole TANGO project team for creating a comfortable and challenging work environment. I express my gratitude to the personnel of the laboratory that has often provided advises and answers to the many questions I had. Last, but not least, I would like to thank my parents, who have at the time influenced me to write this thesis.

Otaniemi 6.5.2002

Juha Halme

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Contents

Contents

PREFACE ..........................................................................................................................I CONTENTS ..................................................................................................................... II ABBREVIATIONS.........................................................................................................IV LIST OF SYMBOLS ....................................................................................................... V 1. INTRODUCTION ..................................................................................................... 1 1.1. 2. DESCRIPTION OF THE TANGO PROJECT ............................................................... 2

COMPOSITE WING RIB ........................................................................................ 5 2.1. 2.2. LOADING OF A COMPOSITE WING RIB ................................................................... 5 MANUFACTURE OF THE COMPOSITE WING RIB ................................................... 17

3.

BASIS OF THE WORK.......................................................................................... 23 3.1. 3.2. 3.3. 3.4. 3.5. TESTING PHILOSOPHY IN GENERAL AVIATION .................................................... 23 TESTING PHILOSOPHY IN TANGO-PROJECT....................................................... 23 STRUCTURE OF DEVELOPMENT TESTING IN TANGO-PROJECT .......................... 26 INVESTIGATED RIB CONCEPTS ............................................................................ 27 TEST METHODS FOR COMPOSITE WING RIBS ....................................................... 28

4.

FIBRE COMPACTION TESTS ............................................................................ 30 4.1. 4.2. 4.3. 4.4. 4.5. 4.6. 4.7. 4.8. INTRODUCTION ................................................................................................... 30 TEST METHOD A FOR DRY REINFORCEMENTS .................................................... 30 TEST METHOD B FOR RESIN APPLIED REINFORCEMENTS .................................... 31 CALCULATION OF THE PARAMETERS.................................................................. 32 MATERIALS ........................................................................................................ 34 TEST SETTING ..................................................................................................... 34 TEST RESULTS..................................................................................................... 34 SUMMARY........................................................................................................... 36

5.

MATERIAL SCREENING .................................................................................... 37 5.1. 5.2. 5.3. 5.4. 5.5. 5.6. 5.7. MATERIALS SELECTED FOR EVALUATION .......................................................... 37 MANUFACTURE OF THE TEST LAMINATES .......................................................... 41 TENSILE TESTS .................................................................................................... 51 COMPRESSION TESTS .......................................................................................... 55 INTERLAMINAR SHEAR STRENGTH TESTS .......................................................... 59 LAYER LEVEL ANALYSIS .................................................................................... 62 SUMMARY OF MATERIAL SCREENING TESTS ...................................................... 64

6.

DESIGN DATA GENERATION ........................................................................... 69 6.1. INTRODUCTION ................................................................................................... 69

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Development testing of a composite wing rib 6.2. 6.3. 6.4. 6.5. 6.6. 6.7. 6.8. 6.9. 6.10. 6.11. 6.12. 6.13. 7.

Contents

PREPARING OF THE TEST LAMINATES ................................................................. 70 TENSILE TESTS .................................................................................................... 71 COMPRESSION TESTS .......................................................................................... 73 IN-PLANE SHEAR TESTS ...................................................................................... 75 OPEN HOLE TENSILE TESTS ................................................................................. 82 OPEN HOLE COMPRESSIVE TESTS........................................................................ 82 INTERLAMINAR SHEAR TESTS ............................................................................. 83 BEARING TESTS .................................................................................................. 83 THROUGH THICKNESS TESTS ........................................................................... 86 SUPPORT TESTS ............................................................................................... 91 MOISTURE BEHAVIOUR OF THE SELECTED MATERIAL .................................... 92 SUMMARY ....................................................................................................... 97

LOADING RIG FOR MULTIAXIAL TESTING ................................................ 99 7.1. 7.2. 7.3. 7.4. 7.5. 7.6. INTRODUCTION ................................................................................................... 99 REQUIREMENTS FOR THE TEST FIXTURE ............................................................. 99

HIGHLIGHTING THE PROBLEMS OF THE MULTIAXIAL LOADING ................................ 99

7.2.1.

TEST FIXTURE ................................................................................................... 102 ANALYSES ........................................................................................................ 103 AMBIVALENT FACTORS .................................................................................... 107 CONCLUSION .................................................................................................... 108

8.

NON-DESTRUCTIVE INSPECTION OF SINE-WAVE WEB ....................... 109 8.1. 8.2. 8.3. 8.4. INTRODUCTION ................................................................................................. 109 PROBLEM AREAS .............................................................................................. 109 SCANNING RESULTS.......................................................................................... 111 CONCLUSIONS .................................................................................................. 112

9.

OTHER UPCOMING TESTS.............................................................................. 114 9.1. 9.2. 9.3. L-PULL TEST ..................................................................................................... 114 FULL SCALE RIB SECTION TEST......................................................................... 115 FULL SCALE WINGBOX PRESSURE TEST ............................................................ 117 SUMMARY ........................................................................................................ 118 REFERENCES................................................................................................... 119 APPENDIXES .................................................................................................... 122

10. 11. 12.

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Abbreviations

Abbreviations

RTM SMT NDI PFC HUT LLS c.g. CFRP LRI GRP VARTM Tg CTE CNC FE(M) RT AH ILSS NCF SO SE UT AITM ABT MEK UD ERFI OHT OHC TMA DSC DOF RF Resin Transfer Moulding Shear, Moment and Torque (loads) Non-Destructive Inspection Patria Finavicomp Oy Helsinki University of Technology Laboratory of Lightweight Structures centre of gravity Carbon Fibre Reinforced Plastics Liquid Resin Infusion Glass (Fibre) Reinforced Plastics Vacuum Assisted Resin Transfer Moulding Glass Transition temperature Coefficient of Thermal Expansion Computer Numerical Controlled Finite Element (Method) Room Temperature Ambient Humidity InterLaminar Shear Strength Non-Crimp Fabric Symmetric Odd Symmetric Even Ultrasonic Testing Airbus Industries Test Method British Aerospace Airbus Test Specification Methyl Ethyl Ketone Uni-Directional Epoxy Resin Film Infusion Open Hole Tension Open Hole Compression Thermal Mechanic Analysis Differential Scanning Calorimetry Degree Of Freedom Reserve Factor

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List of symbols

List of symbols

Cp Vf

xy

F A n m h l t b Xnor Xmea E G

P HL w

Centre of pressure Fibre volume fraction Poisson's Ratio in longitudinal direction Force Area number of layers mass density thickness length thickness stress width Normalised test results Measured test result Young's Modulus Shear Modulus shear stress shear strain normal strain Force Absorbed Moisture mass fraction scanning track angle nominal angle

Subscripts ph A f t u n max nor c m r press head Area of the fabric fibre, fabric tensile ultimate, unnotched notched maximum normalised (value) compressive matrix resin

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Introduction

1. Introduction

The continued growth in air transport has placed an increasing demand on the aerospace industry to manufacture aircraft at lower cost, whilst ensuring the products are efficient to operate, friendly to the environment and fulfil the safety requirements. This, combined with the necessity to develop larger aircraft to cope with increased numbers of passengers on already crowded routes, has established a number of key business drivers for the future products from the industry. This thesis involves issues related to testing of a composite wing rib in development phase. The purpose of the work committed for this thesis was to establish an understanding what kind of testing is needed in terms of different fields of testing; mechanical, chemical and non-destructive testing. When this knowledge is established tests are to be conducted. Mechanical testing of a structure involves strongly analysis of the structure. In order to analyse a structure, in this case a wing rib, it is essential to know how the structure is loaded. Loading of a composite wing rib is discussed in Chapter 2. Furthermore, manufacture of a composite part by Resin Transfer Moulding (RTM) is described. New carbon fibre fabrics and resin systems suitable for structural RTM applications are available from many companies in increasing numbers. This increased supply leads to difficulties when selecting materials. Processability and structural performance are regularly unknown factors. Even if manufacturers present injection parameters, cure cycles and mechanical properties, the suitability of resin system and mechanical behaviour, at least in terms of design data, still remain unknown. Processability of materials and structural performance was studied in material screening. One component aircraft industry approved resins and also cheaper two component resin systems are screened. Required design data values are measured for the selected material. Operative purpose is to establish mechanical properties for the materials rather than to study the processability in viewpoint of this thesis. A lateral wingbox of an aircraft is mainly loaded by shear, moment and torque (SMT) in normal in-flight situation. These loads combine multiaxial loading on a wing rib inside the lateral wingbox. Testing a full scale wing rib or a wing rib section under SMT loads requires test fixture that is capable of producing multiaxial loading. A test fixture for multiaxial testing of wing rib sections is studied in Chapter 7. Problems with multiaxial testing are considered and solutions for these problems are presented. Possibility and restrictions of mechanical Non-Destructive Inspection (NDI), in this case ultrasonic scan, of composite parts incorporating variable geometry is studied. In many applications inspection is forced to be conducted by hand. Mechanical inspection over hand inspection offers many advantages, from which the most significant are time saving (cost) and traceability (scanning record).

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Introduction

1.1.

Description of the TANGO project

The word TANGO stands for: Technology Application to the Near term business Goals and Objectives of the aerospace industry. TANGO is a four-year multi-partner, multi-nation European aerospace project running from April 1st 2000, with the objectives of integrating emerging and near-to-market technologies to provide airframes with lower acquisition and operating costs. Total value of the work proposed is some 87.9 Million Euros. This will be 50% funded by the European Commission 5th Framework programme [1]. TANGO aims to improve structural efficiency and reduce manufacturing costs by achieving further reductions in airframe weight thereby lowering fuel consumption and decreasing the impact on the environment lowering the acquisition cost of the aircraft by using cheaper or more efficient materials, designs with improved producibility and more efficient manufacturing processes To achieve these goals the following specific industrial objectives and challenging targets have been set 20% reduction in weight in comparison to current structures 20% cost reduction in comparison to both current manufacturing processes and state of the art design These levels of product improvement can only be achieved if new technologies are validated and integrated into the design and manufacturing process by the key European airframe manufacturers and their major suppliers. TANGO represents an integrated approach to the validation of these technologies through the construction of the following main airframe components, wherein significant weight and cost reductions can most likely to be generated Composite wing box and metal to composite joint Composite centre wing box Composite fuselage section Advanced metallic fuselage section A holistic view of these airframe elements is presented in Figure 1.

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Introduction

M - METAL C - COMPOSITE

Figure 1. Primary airframe structures to be built within TANGO.

The TANGO consortium has 7 principal contractors: Airbus Industrie (AI), BAe Systems Airbus UK (AUK), DaimlerChrysler Aerospace AG (DA), Aerospatiale Matra, Construcciones Aeronáuticas (Casa), Alenia and SAAB with and a further 26 associate contractors. Figure 2 shows the involvement of nationalities by the number of participating organisations. Figure 3 shows participating organisations and their involvement in TANGO.

BELGIUM UK FINLAND FRANCE

GERMANY SWEDEN SPAIN GREECE HOLLAND IRELAND ITALY ISRAEL

Figure 2. Breakdown of nationalities involved by number of participating organisations.

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Introduction

Figure 3. Participating organisations and their involvement in TANGO.

Patria Finavicomp Oy's (PFC) involvement in co-operation with Helsinki University of Technology / Laboratory of Lightweight Structures (HUT/LLS) is to design and manufacture 5 wing ribs located in the left side outer wingbox. The lateral wingbox includes a total of 15 wing ribs. PFC is designing and manufacturing in co-operation with HUT/LLS the wing ribs 9, 10, 11, 12 and 14.

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Composite wing rib

2. Composite wing rib

The genesis of strong, tough metals, especially iron (steel) and aluminium alloys has meant that in recent times sophisticated structures have been constructed with these materials and there are overriding differences between them and composites; they are isotropic whereas composites can have significant anisotropic behaviour. The use of anisotropic fibre composite materials in structural design requires more complex analysis than isotropic materials. If a composite structure is designed for maximum efficiency, it is necessary to consider the advantages and disadvantages of composite materials.

2.1.

Loading of a composite wing rib

Present-day stress manuals for wing ribs are generated typically for metallic (usually aluminium) parts. However, even if the final structure between metallic and composite within the wingbox structure will be significantly different and the behaviour under loading is different, this does not affect the design approach. The same global loading philosophy can be used for metallic and composite parts. The major influence when incorporating composites is the manufacturing, joining methods and tolerance issues that will set another challenging aims for technology down selection and assembly procedure. To fully understand how the rib is loaded, it is essential to be familiar with the whole wing structure and how the loads acting on wingbox elements are generated. Hence the examination of the overall wing loading is necessary. Also for analyse purposes it is essential to examine the rib in terms of loading. Ribs are part of the wing structure and are influenced by other structural members in the wing. This dependence means that a view of the ribs in isolation will give only partial understanding of wing loading. However, by omitting some of the details do allow some simplification to help clarification. Figure 4 shows the wingbox structure of a typical civil airliner. Skins, spars and the stringers define the geometrical boundary for the ribs. Figure 5 shows the wing structure of a typical two-spar wing configuration. The wing loads can be categorised as shown in Table 1 [2]:

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Composite wing rib

Figure 4. Typical wing structure of a civil airliner [ 3].

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Composite wing rib

Figure 5. A typical two spar wing construction [3].

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Composite wing rib

Table 1. Aircraft loads by category.

Airloads Manoeuvre Gust Control deflection Component integration Buffet

Inertia loads Acceleration Rotation Dynamic Vibration Flutter

Landing - Vertical load factor - Spin-up - Spring-back - Crabbed - One wheel - Arrested - Braking Other loads Towing Jacking Pressurisation Bird strike Actuation Crash

Takeoff Catapult Aborted

Power plant Thrust Torque Gyroscopic Vibration Duct pressure

Taxi Bumps Turning

For analysis it is practical to view some of rib functions in terms of loading. Below are listed the major loads and factors affecting the wing ribs:

Local air loads on the wing collected by the skins Fuel pressure Fuel and structural Inertia loads Slat devices and hold-down loads at the front spar Spoiler/Airbrake loads at the rear spar Aileron loads at the rear spar Flap loads at the rear spar Pylon loads Jacking loads

Figure 6 illustrates a typical airliner configuration and shows the principal forces acting on it during normal flight condition. Normal flight condition is considered as 1g level flight. Aircraft weight includes the weights coming from the aircraft structure, engines and passengers with payload. This group of weights does not vary during the flight. Fuel weight is a quantity that decreases as the progression of the aircraft travel continues. Figure 7 shows the torque generated by other forces. Engine thrust is transferred to the wingbox via engine pylon. The location and design of the engine installation has a significant impact on the loads transmitted through the wing. In a modern airliner the engine is usually located below and in front of the wing thus creating clockwise torque (nose up) by thrust and anti-clockwise torque (nose down) by engine weight. Another torque load is generated by the location of Centre of pressure. Typically for civil airliner the Centre of pressure (Cp) is located on 25-35% of the chord giving an additional clockwise torque to the wingbox [4], assuming that the wingbox torque centre is located after Cp. The influence of these loads and forces is shown in the bending diagram in Figure 9. Figure 9 also illustrates a way how the fuel and engine weight reduce the

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Development testing of a composite wing rib -

Composite wing rib

bending moment across the wing. Fuel is fed to the engines from inner wing to outer wing giving also an additional reduction to the bending moment.

Figure 6. Principal forces during normal flight.

Figure 7. Wing torque during normal flight.

Figure 8. Lift pattern around typical airfoil.

Figure 9. Wing bending diagram.

All forces and loads can be further simplified into 4 basic components especially for structural design purposes. These basic components are shown in Figure 10. Wing loading is represented by these components that are based on aircraft lift and refer to a certain point along the wing span. The shear load is increasing towards the wing root, because every section transfers the load generated by that section but everything outboard of it as well. Shear force creates also an increasing bending moment. Torque along the wing also increases towards the wing root with airload, structural, fuel and engine weight. Engine mounting has a torque reducing effect when located ahead and below the wing.

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Composite wing rib

Figure 10. Basic loading components acting upon the wing in normal flight.

Above was presented wing loading in normal (1g) flight. Other important conditions are landing and ground handling. Figure 11 illustrates the effect that prior to flight mission structural, fuel and engine weights no longer have reducing loading impact on the wing, because lift is not compensating the wing weight. In this situation the whole weight of the aircraft is concentrated at the landing gear attachment points. Usually in swept wing configuration the landing gear is located towards the rear of the wing. Landing gear location is strongly limited by the centre of gravity (c.g.) position and reference [3] gives a rule of 50-55% of the mean aerodynamic chord for the main landing gear location. Figure 12 illustrates the wing torque and loading on the ground at the landing gear position. This generates an anti-clockwise torque that is opposite to the normal flight condition. The engine thrust decreases the anti-clockwise torque. Figure 13 shows the wing behaviour under the effect of undercarriage on the ground.

Figure 11. Principal forces on the ground.

Figure 12. Wing torque and forces on the ground.

Figure 13. Wing behaviour on the ground.

The final factor that sizes the wing and wing elements is often an unusual flight condition. For example in the case of the rib the sizing condition can be the crash case. JAR 25.561 General [5] defines the following ultimate inertia forces acting separately relative to the surrounding structure:

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Development testing of a composite wing rib i. ii. iii. iv. v.

Composite wing rib

Upward, 3,0g 1 Forward, 9,0g Sideward, 3,0g on the airframe and 4,0g on the seats and their attachments Downward, 6,0g Rearward, 1,5g

These inertia forces, the swept wing configuration and the total fuel tank volume decrease in crash case generate internal fuel tank pressure which sets the design goals for many components within the wing. Such components in the case of the ribs are e.g. the rib foot corners (thickness and radius) and bolting/bonding characteristics. The main function of the wing ribs is to maintain the shape and integrity of the wingbox. Below are presented principal functions of the wing ribs from a structural design point of view with illustrating pictures. Maintain the aerodynamic profile of the wing,

React the fuel pressure loading collected from the skins

Transfer local air pressure collected from the skins to the spars

Diffuse local inputs

1

g is defined in JAR 25 as acceleration due to gravity.

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Composite wing rib

Provide end support for the skin/stringers in each rib bay

Resist brazier loading

End boundaries for the fuel tank

Mounting for systems instrumentation inside the wingbox

The rib resists the loads and force inputs presented above by functioning in the following way:

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Development testing of a composite wing rib a) b) c) d) e)

Composite wing rib

In-plane shear loading Rib crushing (brazier load) due to wingbox bending Bending loading due to the Poisson' ratio effect in wing covers s Fuel/Air pressure loading Lateral fuel loading

In-plane shear loading A rib provides a stiff shear connection between spars and skins. A rib collects the loads from the skins and the vertical shear load differences between the spars. The torsion loading about the wingbox structural axis is reacted by skins and spars as a constant shear flow. The shear flow from the skins is transferred to the rib main body (shear panels) via rib feet and castellation. Castellation can be referred to as the area above the shear panels and between mouseholes. Mouseholes are cut outs in the rib for enabling the skin stringers run through the ribs. There are differences in the castellation design between metallic and CFRP ribs mostly due to manufacturing related issues. Yield property of metals, especially aluminium, also leads to different design approach when dealing with castellations and mouseholes. A typical metallic rib castellation is presented in Figure 14, which shows the shear force load path, strengthened castellation edges and horizontal stiffener. Horizontal stiffeners carry the end loads from spars, react to the overall rib bending, divide the shear panels into smaller sections and distribute the shear loads from separate castellations into a more continuos shear along the panel edges. The main body of the rib, the shear panel, carries the shear loads through the rib (Figure 15). At high levels of shear loads the panel may buckle to form diagonal tension fields. In this condition the panel with vertical stiffeners produces a truss type structure that is still capable of transferring the shear load (Figure 16). However, this buckling allowing is used only with metallic ribs. One design driver for CFRP ribs in TANGO is not to allow buckling. Again, the yield properties of metals are favourable in buckling conditions, while composite structures tend to be damaged already in the state of initial buckling.

Wingbox skin

Large radius ­ small stress concentration

Castellation

Edge members

Figure 14. Typical rib castellation in metallic construction.

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Composite wing rib

Shear panel

Figure 15. Rib shear panels.

Figure 16. Rib shear panels in higher levels of shear, post-buckling.

Rib crushing (Brazier load) Rib crushing is generated by flexural wing bending - when a wing box is subjected to bending loads, the compression and tension in curved wing skins generates an inward acting loads on the wing ribs. This effect is shown in Figure 17. The compression in the ribs is reacted by vertical stiffeners or shear panel in case the rib does not have any vertical stiffeners or when geometrical stiffening is used (e.g. sine-wave web shape). Figure 18 illustrates the compression in the rib. Compression is mainly transferred via skin stringers and in an ideal design situation in the rib viewpoint the vertical stiffeners are vertically aligned with a pair of stringers on upper and lower skins. However, it was indicated during the conceptual design in TANGO that locating the stringers against each other was difficult. Structural efficiency was not the primary

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obstructing factor, but the fuel system requirements and lack of co-operation in upper and lower cover design1.

Figure 17. Wing rib crushing load (Brazier load).

Figure 18. Brazier load impact on rib.

Bending loading Chordwise bending loads are generated by stresses in the wing skins. In a normal flight wing bends upward creating a spanwise compression in the upper skin and tension in the lower skin. Due to the materials Poisson' ratio wing skins under stresses generate a s bending load to the ribs. Upper cover generates outward tension in the rib upper flange and lower cover generates inward compression in the rib lower flange, this combined load condition tends to bend the rib that is shown in Figure 19. The induced bending load is critical for mouseholes due to stress concentrations at the mousehole edges.

1

Upper cover is designed by Alenia and lower cover by DASA.

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Development testing of a composite wing rib Upper cover Poisson' effect s

Composite wing rib

Lower cover Poisson' effect s

Figure 19. Rib bending due to spar end loads and Poisson' ratio effect in skins. s

With CFRP ribs the Poisson' ratio effect is highly dependent upon the lay-up of the s 1 skins, since longitudinal and transverse Poisson' ratios for CFRP laminates are s different. Figure 20 shows a typical (AS1/3501, Vf=60%) CFRP laminate longitudinal Poisson' ratio xy values with different ply orientations. s

Figure 20. A Poisson' ratio in longitudinal loading as a function of ply orientation [ 6]. s

The fact that wing skins are primarily designed to resist the bending load in the wing [3], leads to significant alignment of the fibres in the direction of the loading. In the case of wing skins this means that fibres are heavily aligned to the spanwise direction, which is

1

Longitudinal direction refers to laminates 0 -direction, which further refers to spanwise direction.

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Composite wing rib

an advantageous situation in the rib point of view, since the longitudinal Poisson' ratio is s small in this case. Fuel/air pressure loading The wing bending generates vertical compression at the rib, while fuel and air pressure work in the opposite direction generating tension loads across the rib depth. The rib vertical stiffeners work in this situation as connection rods between the top and bottom stringers and skins. This tension loading pulling the skins apart from the ribs drives to the use of separate cleats for some areas as shown in Figure 21. There are lots of different cleating possibilities between ribs/skins and ribs/stringers. Fuel / air pressure

Angle cleat

Rib foot

Fuel/air pressure load reaction

Figure 21. Fuel and air pressure loads on the rib.

Lateral fuel loading Moving fuel inside the wing generates lateral fuel loading. In normal in-flight condition gusts and manoeuvres cause fuel movement from inboard to outboard. This creates an increasing pressure towards the tip of the wing. The skins pick up most of the pressure, and mainly ribs and spars carry loads. Highly loaded parts due to the fuel pressure are the sealed fuel tank end rib and the ribs at the end of the fuel tank.

2.2.

Manufacture of the composite wing rib

This chapter describes few methods considered for manufacturing TANGO wing ribs and highlights some of the benefits and disadvantages of the chosen manufacturing method. Also a few examples of advanced composite parts developed of aerospace industry are presented. The full potential of the composite wing rib is achieved in cost and effectiveness viewpoint when the manufacturing issues are also considered. First of all, from the manufacturing point of view it is not practical to copy the structure of a modern aluminium wing rib. The structures based on that kind of design philosophy are

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commonly referred to as " black aluminium" parts. Aluminium wing ribs are machined from solid billet while the manufacturing approach for the composite parts is totally different. Within PFC' involvement in TANGO two manufacturing methods were s investigated for manufacturing the wing ribs: Liquid Resin Infusion (LRI) and Resin Transfer Moulding (RTM). RTM was chosen for the manufacturing method for all five wing ribs. The choice of RTM was based on existing RTM technology knowledge at HUT/LLS and PFC. Also the driving factors of TANGO project to lower the manufacturing costs as well as lowering the structural weight was considered to be met by choosing RTM as a method of manufacture. Table 2 illustrates the major benefits and disadvantages of the RTM process. RTM is not a single manufacturing process that can be dealt with in a monolithic manner. RTM is better thought of as a philosophy of manufacturing in which the resin and fibres are held apart until the last possible moment. Most of the RTM processes comprise out of the illustrated steps in the Figure 22.

Table 2. Some of the benefits and disadvantages of the RTM process [ 7, 8].

The benefits of RTM: The disadvantages of RTM: Low-cost materials High tooling cost Non-autoclave cure Difficulty of using certified prepreg resins Reduced lay-up time Reduced cost benefits for low-volume Improved reproducibility applications High fibre volume fraction Excellent surface finish Net or near-net shape manufacturing Improved laminate quality Improved tolerance control Integration of parts Advanced textile preforming: stitching, 2D braiding, 3D weaving is possible A. Preforming

Pressure / Heat

B. Injection

Pressure / Heat

C. Curing

Pressure / Heat

Preform Resin In Tool Pressure / Heat

Air Out

In

Plug

Pressure / Heat

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Development testing of a composite wing rib D. De-moulding

Open mould

Composite wing rib F. Trimming

E. Post cure

Heat

Cured part

Heat

Figure 22. Schematic process sequence of RTM.

Preforming (Figure 22A) A preform of reinforcement is prepared by hand or by automated equipment . A separate tool for preforming is often used. Wide variety of materials (wood, tooling foam, casting resin, GRP and metals) can be applied for the preform-tool [9]. Preform tooling can be an open or closed type tool. Usually in preforming the reinforcements contain a small amount of binder resin 1, which is applicable with the injection resin. After the reinforcements are placed in mould or moulded on tool the bindered reinforcements are heated to cure the binder and further to keep the reinforcements in the performed state. Cured preform can be easily or at least easier than dry reinforcements, shaped and worked before it is placed in the mould. Injection (Figure 22B) Once the dry preform is placed in the mould and the mould is closed the resin is injected with pressure in the mould. A release agent is naturally applied to the mould surface before injection. Injection pressure varies from rather small pressures of 0.5 bar up to higher pressures of 6­ bar [7]. Injection pressure is highly dependent upon application, 7 RTM equipment, available shop air pressure, mould design and desired fibre volume fraction (Vf). Vacuum can be also applied to the air outlet port to help the resin flow through the preform. In this case the process is often referred to as Vacuum Assisted Resin Transfer Moulding (VARTM). However, in some relations VARTM is described as only vacuum injection process possibly with open mould. The mould is usually at elevated temperature (injection temperature) for lowering the viscosity of injected resin and to further accelerate and enable the resin to flow through the whole preform. In almost every RTM aerospace application one-part thermoset resin 2 is used. The primary reason for the use of thermosets in RTM applications is their processability. Thermosets have optimal low viscosity and sufficient pot-life at injection temperatures. Thermosets also have high Glass Transition Temperature 3 (Tg) when cured that is often a requirement for aerospace structures. Depending of the part size multiple inlet and outlet ports can be

Also referred as binder or tackifier. One-part thermoset resins are in liquids or semi-solids at room temperature. Thermosets can be cured with catalyst or heat and once cured they cannot change form again. 3 Glass Transition temperature is a temperature where at the physical properties of the resin rapidly drops and resin transforms from glassy state to rubbery state.

2

1

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Composite wing rib

used to accelerate the injection process. Table 3 presents some one-part commercial thermosets reported being used in aerospace applications.

Table 3. One-part epoxy systems for aerospace resin transfer moulding Applications [ 10. 11, 12].

Viscosity at injection temp. (mPa.s, cPs) *) Hexcel: RTM 6 3M: PR 500 Cytec: Cycom 875RTM 60-70 (120 C) <45 (160 C) 150-200 (60 C)

Pot-life at injection temp. (min) >240 30-40 24 hours

Tg

( C)

183 (DMA1) 205 (DSC2) 193 (DMA)

*) Applicable injection temperature in brackets.

Curing (Figure 22C) Once resin is bleeded through the air outlet the outlet hole is plugged and pressure in the inlet line is maintained or possibly risen. The amount of bleeded resin depends of the applications and injection parameters. Good results in TANGO manufacturing trials were achieved by bleeding the resin to the point were air bubbles in resin were not visually detectable. In practise this meant rather small (20 ­ 100 g) amount of bleeded resin. This is further discussed in chapter 5.2. De-moulding (Figure 22D) After the initial cure is finished the mould is opened and the cured part is removed from the mould. Part separation from the mould is not always obvious and should be taken into account in mould design. A separate tool can be used for extracting the part or mould can incorporate integrated de-mould features or tool, blowing air into resin inlet or outlet port is also used in de-moulding when ports are positioned in applicable manner. Mould material has a significant effect on whether the mould should be let cooled to room temperature or the part should be removed when the mould and moulding are still warm/hot. When i.e. an aluminium mould with a high Coefficient of Thermal Expansion (CTE, 23 10-6 / C [13]) is used, it is disadvantageous to let the part to cool down before removal, especially with Carbon/Epoxy parts, which has considerably smaller CTE (CTE for a Carbon/Epoxy laminate with quasi-isotropic lay-up is 2.9 10-6 / C [14]) than for aluminium. During the cooling the geometry of the mould and moulding change at a different rate when resulting in different geometry between mould and mouldings when

1

Dynamic Mechanical (Thermal) Analysis (DMA) is a method for determining Tg value: a sample is subjected to a small sinusoidal load and the response of the sample is analysed. 2 Differential Scanning Calorimetry (DSC) is a method for determining Tg value: the amount of energy (heat) absorbed or released by a sample is measured as it is heated, cooled, or held at a constant temperature.

- 20 -

Development testing of a composite wing rib -

Composite wing rib

cooled down. From this a following rule can be implemented; The higher the difference is in CTE between mould and moulding the harder the part removal is when the part is let to cool down. Postcure (Figure 22E) After curing and removal from mould the part is postcured at elevated temperature to improve final properties and/or degree of cure. For most of the resins used in RTM applications, especially thermosets, the final degree of cure is often achieved at higher temperature than curing temperature. In postcuring the part is freestanding. Trimming (Figure 22F) Trimming can be conducted before or after postcuring. However, in some cases postcuring produce changes to the part geometry and in this situation trimming is sensible to conduct after postcuring. Mostly trimming is conducted with pneumatic hand drills or when higher tolerances are required with a CNC machine. Figure 23 shows a wing spar for a modern fighter. The spar is manufactured by RTM and, according to the manufacturer, cost saving is 20 % compared to the corresponding prepreg part. The number of reinforcement parts has also cut in half, which is significant in this case since the fighter contains 46 composite spars [15]. Table 4 lists parts of a modern fighter manufactured by RTM. This list points out the high amount of RTM parts also in load carrying primary structures; wing spars, highly loaded ribs and loaded fuselage parts.

Figure 23. F22 RTM sine-wave wing spar [ 15].

- 21 -

Development testing of a composite wing rib -

Composite wing rib

Table 4. Examples of modern fighter Resin Transfer Moulded parts [7].

- 22 -

Development testing of a composite wing rib -

Basis of the work

3. Basis of the work

3.1. Testing philosophy in general aviation

When composites are to be used in structural components, a design development program is generally initiated during which the performance of the structure is assessed prior to use. This process of substantiating the structural performance and durability of composite components generally consists of a complex mix of testing and analysis. Testing alone can be prohibitively expensive because of the number of specimens needed to verify every geometry, loading, environment, and failure mode. Analysis techniques alone are usually not sophisticated enough to adequately predict results under every set of conditions. By combining testing and analysis, analytical predictions are verified by test, test plans are guided by analysis, and the cost of the overall effort is reduced while reliability is increased. An extension of this synergistic analysis/test approach is to conduct analysis and associated tests at various levels of structural complexity, often beginning with small specimens and progressing through structural elements and details, sub-components, components, and finally the complete full scale product. Each level builds on knowledge gained at previous, less complex levels. This substantiation process, using both testing and analysis in a program of increasingly complex levels, has become known as the "Building Block" approach. Building block approach is illustrated in Figure 24 below.

Figure 24. Building block approach of testing.

3.2.

Testing philosophy in TANGO-project

This chapter explains the approach for testing the lateral wingbox structures within TANGO. This approach was created during the first year of TANGO and is based on the knowledge of the participating organisations. However, heavy input was fed from the principal contractors, especially AUK (BAe), since the level of experience in these companies was higher compared to smaller organisations.

- 23 -

Development testing of a composite wing rib -

Basis of the work

To ensure exportability of TANGO data to other aircraft programmes, all data generated within TANGO was to be of sufficient quality such that it could be used as part of the certification of that technology for aircraft use if required. Therefore all testing was to be carried out to the traceability and quality standards required for a full certification programme. At the coupon and element levels, test articles were to be manufactured using the same technology and processing route (e.g. stitched fabric preform) as is proposed and preferred for the corresponding part of the lateral wingbox structure. Due to cost and time constraints, some element and detail level tests were possibly replaced by analysis if sufficient confidence in the calculation methodology and supporting data could be demonstrated. The TANGO lateral wingbox is not designed as a flying aircraft structure and was not to be subjected to certification by the Airworthiness Authorities. There was no requirement to formally certify or qualify the novel materials and processes being evaluated for the lateral wingbox, and testing within TANGO is, therefore, confined to that required to validate the design methodology. For generating design data B-values or A-values, defined in MIL-HDBK-17 [16], a different method was agreed. This different statistical treatment of data was agreed to reduce the number of test specimens for generating design data B-values. For an aircraft structure, design allowables are generated statistically using, for example, the procedure defined in MIL-HDBK-17 to give mean, A- and B-basis values as required. To be statistically acceptable this approach requires large numbers of specimens, typically 5 batches of 6 specimens per material property. In order to achieve the statistical confidence levels required by a formal certification approach, the limited set of data generated within TANGO would result in large knockdowns being applied. As this would lead to an overly conservative design in TANGO the following approach was created where batch sizes prevent the sensible application of the methods outlined above: - A minimum of eight specimens was to be tested at the coupon level for each mechanical property. This will provide a credible mean and standard deviation for each property. - For design allowables based on means (e.g. tensile modulus), the average measured value for that property is used - For design allowables based on statistically derived B-basis values, an approximate B-basis value for that property is derived as follows: Standard deviation 0.05 Mean then B-basis value = (0.85 x measured mean value). A. If

- 24 -

Development testing of a composite wing rib -

Basis of the work

Standard deviation 0.05 Mean then B-basis value = (0.80 x measured mean value). B. If For coupon level testing the minimum level of material data required for design of the lateral wingbox structure within TANGO is defined in Table 5. The element and detail level tests required for designing the lateral wingbox components are presented in Table 6 and Table 7

Table 5. Level of data required for design of different lateral wingbox components.

Property In plane tensile modulus and strength In plane compressive modulus and strength In plane shear modulus and strength Through-thickness tensile modulus and strength Through-thickness compressive modulus and strength Interlaminar shear strength Tensile notch factor (OHT) Compressive notch factors (OHC and FHC) Bolt bearing bypass diagram Ramp angles and ply drops Bolt bearing strength Bolt pull through Tension after impact strength Compression after impact strength Adhesive material data Shim material data Fastener fit and hole quality Bolt torque effect Creep relaxation Spar Ribs Top Skin Bottom Skin Joints

Table 6. Element level tests required for lateral wingbox structures.

Structure Spar Element Tests Spar cap corner radius L-pull Rib post corner radius T-pull Spar web panel in-plane shear (with and without hole) Rib Rib foot corner radius L-pull (with and without bolts) Rib web and stiffener in compression Rib web panel in-plane shear (with and without hole) Top Skin Bottom Skin Stringer pull-off (if cleated) Stringer pull-off (if cleated)

- 25 -

Development testing of a composite wing rib Joint

Basis of the work

Bonded joint elements (with and without shim) Bolted joint elements (with and without shim)

Table 7. Detail level tests required for lateral wingbox structures.

Structure Spar Detail Tests Spar 3 or 4 point bend Fuel pressure test of spar section Rib Top Skin Bottom Skin Joint None Stiffened compression panel Stiffened tension panel (including access hole and impact damage) Mid span interface section

The testing philosophy in TANGO is based on building block approach. Figure 25 outlines the structure of testing, where development testing is concentrated in laminate, detail and element level testing.

Component Complete wing Sub-Components, e.g. Outer wing box Landing gear support Pylon attachment Details, e.g. Skin/stringer panels Mid-span joint Spar or rib box test Elements, e.g. Bolted joint elements L-pulls and T-pulls Fatigue and damage tolerance Coupons, e.g. Materials & processes Design data generation Environment

Component Tests Certification / Qualification

Development Tests Proof of concept Design data generation

Figure 25. Pyramid of testing.

3.3.

Structure of development testing in TANGO-project

In the TANGO program from the PFC and HUT/LLS point of view the structure of development testing is presented in Figure 26.

- 26 -

Development testing of a composite wing rib -

Basis of the work

In the material screening program different fabrics and resin systems were tested by measuring few basic mechanical properties of the given material combination. Material screening program gives mechanical performance input for the selection of the final fabric and resin system. In this point it is good to notice that mechanical performance was only one part affecting the final ranking of all tested materials. Feasibility for manufacture and drapeability of the fabric were actually the selection drivers in the material screening program. Design data generating for the selected material includes measuring the needed properties (Table 5) and moisture absorption behaviour of the selected material. Design data is then input to the Finite Element (FE) Model of the ribs. FE analyses reveal the critical areas in the ribs and further refine the need of element tests. Sub-component design is finally verified in the full-scale component test.

Mechanical performance input

Material screening program ­ranking of materials

Selection of material

Design data input

Design data generation for selected material

FEM - Sizing and stressing calculations

Critical areas and detail analyses

Detail tests

and

element

Element Tests: - L-Pull tests - Compression/shear panel tests

Detail tests: - Integrated concept

cleat

Rib design definition and Manufacture

Full Scale Wingbox Tests

Wingbox Assembly

Figure 26. Structure of testing within TANGO.

3.4.

Investigated rib concepts

A number of rib web stiffening options were identified during the concept design phase as shown in Figure 27. These web stiffening options were developed to understand how a

- 27 -

Development testing of a composite wing rib -

Basis of the work

complete rib design would result from these various options. The rib concepts resulting from the different web stiffening concepts are presented in Figure 28.

Flat plate

Blade stiffened panel

Corrugated web

Beaded, symmetrical web

Beaded, closed web

Figure 27. Investigated web stiffening options.

Figure 28. Investigated rib concepts.

3.5.

Test methods for composite wing ribs

A literature study to find test and verification methods for wing ribs was part of the work of this thesis. However, the study practically did not reveal anything useful.

- 28 -

Development testing of a composite wing rib -

Basis of the work

Composite wing programs are all still strongly research and development type programs in civil aviation. Due to the nature of the programs very little information is shared by the responsible organisations. Composite parts in military airframes, also in primary load carrying structures, are incorporated very early compared to the airframes in civil aircrafs. But naturally data from these structures and programs is also very limited. Discussions with representatives from Airbus UK revealed also that the testing of metallic wing ribs is quite exiguous. In fact, no tests for metallic wing ribs have been conducted. Since detail and element level testing of metallic wing ribs is scarce, the resource material is also very limited. Metallic wing ribs are usually tested by subcomponent and component level tests as a part of a bigger structure such as a whole wingbox structure. In the TANGO project component and sub-component level testing are carried out when the final full-scale lateral wingbox structure is tested.

- 29 -

Development testing of a composite wing rib -

Fibre compaction tests

4. Fibre compaction tests

4.1. Introduction

The fibre compaction tests were conducted in order to assess the mould-closing forces and required pressures for an intended fibre volume fraction of the final product. The objective of the test was to produce a graph of the achieved fibre volume fraction as a function of the applied pressure on the reinforcements. Two types of tests were used; one for dry reinforcements and the other for resin wetted reinforcements. A 100 kN servohydraulic testing machine was used in the test.

4.2.

Test method A for dry reinforcements

In the test method A the reinforcements were compressed with a flat press head, which was pivot jointed to the loading plate. The reinforcements were placed on a flat rigid steel base plate. The applied load and thickness of the reinforcements under the press head were recorded during the test. A schematic picture of the test method is presented in Figure 29. A photograph of the test setting is presented in Figure 30.

F

Loading plate Pivot joint Press head Reinforcements Base plate

Figure 29. Reinforcement compression test arrangement for dry reinforcements.

- 30 -

Development testing of a composite wing rib -

Fibre compaction tests

Figure 30. Test arrangement of the test method A.

4.3.

Test method B for resin applied reinforcements

The test method B is similar to the method A except that resin is applied to the fibres. Also a resin sump is used for keeping the reinforcements filled with resin. A schematic picture of the test method is in Figure 31. A photograph of the test setting is presented in Figure 32.

F

Loading plate Pivot joint Press head Resin Reinforcements Resin sump Base plate

Figure 31. Reinforcement compression test arrangement for resin applied reinforcements.

- 31 -

Development testing of a composite wing rib -

Fibre compaction tests

Figure 32. Test arrangement of the test method B.

4.4.

Calculation of the parameters

Recorded parameters during the test were applied load and stroke of the load plate. Maximum applied load was 80 kN. At 80 kN the test arrangement is under certain amount of yield. The yield results from the loading machine frame and elements such as the pivot joint. The total yield of the test setting can be measured by running the test without the reinforcements. The recorded stroke in this case is directly the total yield of the test setting. When load and stroke are recorded it is possible to fit a function for load ­stroke relation. This function gives the amount of yield (mm) for given a load (N). The amount of the yield needs to be added to the measured thickness of the reinforcements so that the right fibre volume fraction can be calculated. Yield of the test setting is demonstrated in Figure 33. The fibre volume fraction can be calculated when the thickness of reinforcements, density of the fibres, areal weight of the reinforcements and the number of the layers are known. Theoretical percentual fibre volume fraction is calculated with equation n mA 10 f h is the number of layers is areal mass of the fabric, ( g/m 2 ) is the density of the fibre, ( g/cm 3 )

f

(1)

where

n mA

f

- 32 -

Development testing of a composite wing rib h

Fibre compaction tests

is the thickness of the cured laminate, (mm)

It is assumed that all the applied force to the press head is transferred through the reinforcements only under the press head. In other words, only the reinforcements under the press head are assumed to be compressed and no fibres are compressed outside the press head area. It is possible that small amount of fibres is compressed near the press head, but the calculation of the area under compression outside the press head is somewhat undetermined. Applied pressure (bar) is F Aph is applied force on the reinforcements (N) is the area of the press head ( mm 2 )

c

10

(2)

where

F Aph

Yield of the test setting

0,40

0,35

0,30

0,25

0,20

Stroke - yield (mm) Poly. (Stroke - yield (mm))

0,15

0,10

0,05 y = -2,9383E-11x6 + 8,0079E-09x5 - 8,4799E-07x4 + 4,4082E-05x3 - 1,1757E-03x2 + 1,8599E-02x + 2,3230E-02 0,00 0 10 20 30 40 50 Load / kN 60 70 80 90

Figure 33. Yield of the test setting as a function of applied load. Curve fit is shown.

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Development testing of a composite wing rib -

Fibre compaction tests

4.5.

Materials

The materials and conditions studied are presented in Table 8.

Table 8. Materials and conditions. n = Number of layers, D = Dry, R = Resin applied.

Material Hexcel G0926 Hexcel G0926 INJ Hexcel G0926 Hexcel G0926 INJ (Devold) TYPE 2B NCF Devold DB550 Sigmatex 43364 8HS Interglas 92140 * Fibre, density (g/cm3) HTA, 1.78 HTA, 1.78 HTA, 1.78 HTA, 1.78 HTS, 1.78 HTS, 1.78 HTA, 1.78 E-glass, 2.54 Binder RTM6 * RTM6 * EB6 n 7, 14, 21 7, 14, 21 14 14 8 9 16 19 Condition D D R R D D D D Stacking order [0/90]n [0/90]n [0/90]n [0/90]n [-45/90/45]n [45/-45]n [0/90]n [0/90]n Weight (gsm) 370 391 370 391 660 560 364 390

Binder is Hexcel' RTM6 compatible. s

The used resin in the test was Ciba LY 5052. Sigmatex 43364 and Interglas 92140 materials were tested for comparison.

4.6.

Test setting

Fabrics were cut in 120 x 120 mm2 pieces for dry test and 110 x 120 mm2 pieces for resin applied test. The press head was round with a diameter of 108.0 mm. The press head edge was chamfered with 1 x 1 mm chamfer. The press head speed was 0.25 mm/min for measuring the test setting yield and 1.0 mm/min when the reinforcements were compressed. The material for load plates, base plate, press head and pivot joint elements was high strength steel. The resin sump base plate material was Al 2024-T6.

4.7.

Test results

Figure 34 shows the binder effect and the effect of the number of layers on the pressure needed to achieve a specified fibre volume fraction. Figure 35 shows the differences in the required pressure for selected materials. Figure 36 shows the effect of the resin applied to the reinforcements. The results from already once pressed reinforcements are also presented.

- 34 -

Development testing of a composite wing rib -

Fibre compaction tests

Fiber volume fraction - Pressure

30

G0926 - 7 layers G0926 - 14 layers G0926 - 21 layers

20

G0926 INJ - 7 layers G0926 INJ - 14 layers G0926 INJ - 21 layers

10

0 55 57 59 61 63 65 Fiber volume fraction / %

Figure 34. Effect of number of layers and binder effect.

Fiber volume fraction - Pressure

60

G0926 - 14 layers

50

G0926 INJ - 14 layers TYPE 2B NCF - 8 layers

40

DB550 - 9 layers SC548 8HS - 16 layers 92140 - 19 layers

20

30

10

0 55 57 59 61 63 65 Fiber volume fraction / %

Figure 35. Differences between materials.

- 35 -

Development testing of a composite wing rib -

Fibre compaction tests

Fiber volume fraction - Pressure

40

G0926 - 14 layers G0926 INJ - 14 layers

30

G0926 +Resin - 14 layers G0926 INJ +Resin - 14 layers G0926 INJ +Resin 2. press - 14 layers

20

10

0 55 57 59 61 63 65 Fiber volume fraction / %

Figure 36. Effect of resin applying and results for second press.

4.8.

Summary

Different materials and conditions were tested in order to establish an estimation of the required pressure for intended fibre volume fraction for a given material and condition. In this case G0926 INJ was the most interesting material since it was chosen for the final wing rib material. The results show that an increasing number of layers leading to a thicker laminate does not have a significant effect on the required pressure. Adding binder to the fibres has a moderate effect on the required pressure. By applying binder to the fibres leads to increase of the required pressure for intended fibre volume fraction. Also by applying resin to the fibres leads to increase of the required pressure. However, it has to be noted that speed of testing has a significant effect to the results when testing resin applied fibres. Lower speed of testing leads to decrease of the required pressure, since it enables better bleeding of the resin through the fibres. In practise this result has somewhat little relevance, since in closed mould methods of manufacturing compressed reinforcements are usually dry and the resin is injected after the reinforcements are already compressed to the final thickness or fibre volume fraction. Differences in the materials were as anticipated. Glass fibre reinforcements require more pressure to achieve the same fibre volume fraction than carbon fibre reinforcements. When the reinforcements are compressed the second time the required pressure is much lower.

- 36 -

Development testing of a composite wing rib -

Material screening

5. Material screening

The purpose of the material screening phase was to give a ranking for the tested materials in terms of: Structural performance Feasibility of manufacture General applicability of the material for wing ribs Structural tests also qualified the manufacturing process designed for the particular test laminate. Finally, fabric and resin system for all five ribs were selected. In order to rank the tested materials in terms of mechanical performance three different basic properties of the test laminates were measured in room temperature (RT) and ambient humidity (AH). These properties were: Tensile strength and modulus Compressive strength and modulus Interlaminar shear strength (ILSS)

5.1.

Materials selected for evaluation

The materials that were selected for evaluation are presented in Table 9. The selection of the materials was based on discussions with representatives from AIRBUS UK and on the existing knowledge at PFC and HUT/LLS of the RTM applicable resin systems and fabrics.

Table 9. Preliminary selection of materials for screening. FABRIC Hexcel G0986 INJ Hexcel G0926 Hexcel G0926 INJ Hexcel NC2 / 45 Devold DB 550-C06 Devold Type 2 NCF1 Devold Type 6 UD RESIN Hexcel RTM6 Bryte EX1545-1 Cytec 875 RTM Shell Epikote 862 with Epicure DX6509 ACG EF7199 = XLVR15-2

Since the delivery of Hexcel' NC2 and Advanced Composite Group' (ACG) EF7199 s s was uncertain while material screening was ongoing, it was decided to leave these two materials out of the material screening program. Also at the early stage of the concept

1

Type 2 and type 6 fabrics refer to Airbus UK notation used for categorising fabrics.

- 37 -

Development testing of a composite wing rib -

Material screening

design work it was decided to leave Hexcel' G0986 out of the mechanical performance s screening program. This decision was based on the fact that the 5H Satin (G0926) weave has better draping properties over the 2x2 Twill (G0986) weave. In theory the 5H Satin has also better mechanical performance than 2x2 Twill, since the level of crimp is smaller with the 5H Satin. G0986 was used only for demonstrating and manufacturing trial purposes. In Twill fabrics one or more warp fibres alternately weave over and under two or more weft fibres in a regular repeated manner. This produces the visual effect of a straight or broken diagonal 'rib' to the fabric. Satin weaves are fundamentally twill weaves modified to produce fewer intersections of warp and weft. The harness number used in the designation (typically 4, 5 and 8) is the total number of fibres crossed and passed under, before the fibre repeats the pattern. Satin weaves are very flat, have good wet out and a high degree of drape. The low crimp gives good mechanical properties. Satin weaves allow fibres to be woven in the closest proximity and can produce fabrics with a close tight weave. Examples of these two fabrics are presented in Figure 37.

a

B

Figure 37. 2x2 Twill (a) and 4H Satin (b) fabrics.

Non-crimp fabric (NCF) is a multiaxial fabric, whose manufacturing technique combines aspects of weaving and knitting. The benefits of NCF are straight fibres and lower manufacturing costs compared to woven fabrics. Straight (uncrimped) fibres naturally perform better under loading than crimped fibres. NCF is manufactured using warp knitting or stitch bonding and can contain relatively uncrimped fibres oriented at angles that can vary from 0 to 90 . One type of NCF is presented in Figure 38.

- 38 -

Development testing of a composite wing rib -

Material screening

Figure 38. Non-crimped fabric (NCF).

More detailed information on the evaluated fabrics is presented in Table 11. The fibre orientations of the fabrics are presented in Table 12. Detailed information on the evaluated resins is presented in Table 13. Summary of the material combinations for the material screening program is presented in Table 14. The selection of the material combinations was based on the facts and presumptions presented in Table 10. The G0926INJ was selected as the basic material that was used with all resin systems. The four resin systems could be compared to each other. Cycom 875 RTM and EX 1545-1 resin systems were selected as the matrix material to be used with all fabrics. These two material were considered as the two best choices. Also availability supported the selection.

Table 10. Selection criteria of the material combinations. MATERIAL G0926INJ ADVANTAGES + Base material + Good drapeability + Familiar reinforcement form + Inexpensive + Good mechanical performance + 1-component + Already qualified aerospace material + High Tg + Good mechanical performance + High Tg + 1-component + Already qualified aerospace material + Inexpensive + High Tg DISADVANTAGES - More expensive than NCF

NCF in general RTM6

- Poor drapeability - Limited directionality - Expensive

EX 1545-1

Cycom 875 RTM Epikote 862/Amicure

- Expensive - Hard to process - 2-component - Low Tg - Cheaper than RTM6 - 2-component - Poor injection viscosity

- 39 -

Development testing of a composite wing rib -

Material screening

Table 11. Reinforcements selected for evaluation. FABRIC FIBRE TYPE AREAL WEIGHT ( g/m ) Hexcel G0986 INJ Hexcel G0926 Hexcel G0926 INJ Devold DB 550-C06 Devold Type 2 NCF Devold Type 6 UD HTA, 6k HTA, 6k HTA, 6k HTS, 12k HTS, 12k HTS, 12k 2x2 Twill, 5H Satin 5H Satin NCF NCF NCF 280 370 392 550 652 652 RTM6 1 RTM6 EB6 EB6

2

BINDER

Table 12. Fibre orientations of the fabrics FABRIC Hexcel G0986 INJ Hexcel G0926 Hexcel G0926 INJ Devold DB 550-C06 Devold Type 2 NCF Devold Type 6 UD FIBRE ORIENTATION (0 / 45 / -45 / 90 ) 2 (52% / 0% / 0% / 48%) (50% / 0% / 0% / 50%) (50% / 0% / 0% / 50%) (0% / 50% / 50% / 0%) (0% / 41.7% / 41.7% / 16.6%) (100% / 0% / 0% / 0%)

Table 13. Resin systems selected for evaluation. RESIN TYPE Cured Tg ( C) RTM6 EX1545-1 Cycom 875 RTM Epikote 862/Amicure 1-component epoxy 2-component syanate ester 1-component epoxy 2-component epoxy 180-190 180 130 160

3

RTM6 binder is RTM6 applicable. 0 denotes warp direction of the fabric. 3 Dry value from manufacturer.

2

1

- 40 -

Development testing of a composite wing rib -

Material screening

Table 14. Selected material combinations for material evaluation.

CODE

1-a 2-a 2-b 2-c 2-d 3-a 4-a 4-b 5-a 5-b 6-a 6-b

FABRIC / FIBRE

Hexcel G0986 / HTA Hexcel G0926 / HTA

RESIN

Hexcel RTM6 Bryte EX1545-1 Cytec Cycom 875 RTM Shell Epikote 862/Amicure ACG XLVR15-2 Hexcel RTM6 Bryte EX1545-1 Cytec Cycom 875 RTM Bryte EX1545-1 Cytec Cycom 875 RTM Bryte EX1545-1 Cytec Cycom 875 RTM

Hexcel G0926 / HTA / RTM6 binder Hexcel NC2 / 45 / 600 g/m^2 / HTS Devold DB 550-C06/ HTS / 45 Devold Type 2 NCF + Type 6 UD / HTS

5.2.

Manufacture of the test laminates

The test laminates, from which the test specimens were extracted, were manufactured at the HUT/LLS. The manufacturing method for the test laminates was RTM, since RTM was chosen to be the final manufacturing method for the wing ribs. A special flat RTM mould was designed for producing 3 mm thick test laminates. The mould consisted of 3 aluminium plates presented in Figure 39. The manufacturing process is illustrated in Figure 22. The reinforcements were cut slightly smaller than the plate 1 frame for avoiding the fibres from getting between plate edges and for creating enough space for resin to flow. The reinforcements were then placed on the plate 1 and plates 2 and 3 are then laid on the plate 1 to close the mould. The stacked plates are then placed in the 300 Ton-press (Figure 40). The HUT/LLS RTM injection system was then connected to the injection inlet of the mould with 6 mm inner radius copper pipe. The pipe was connected with olive connectors. The RTM system consisted of the pressure vessel, controllable pressure valve and temperature control. The mould was electrically heated using the two heating plates of the press. The press heating plates are located above and under the mould. The mould and the resin inside the RTM system were heated to appropriate temperature. After the mould and the resin achieved the correct temperature the injection valve was opened and the resin injected in the mould. The injection line (copper pipe) was also heated during the injection. At the beginning of the injection the pressure was kept low at 0.5 to 1.0 bar in order to let the resin first to fill the space between the reinforcements and the plate 1 frame edge. After a minute or two the injection pressure was risen to the correct injection value. Once the space between the reinforcements and the plate 1 edge was filled the resin starts to flow towards the centre of the reinforcements. In the plate 2 there are two bleeding holes. After the resin fronts collide at the centre the resin is bleeded through the bleeding holes. The resin sump in the plate 3 then collects the bleeded resin. Once the resin sump is filled, the resin is bleeded through the outlet hole. By monitoring the amount of air (bubbles) in the bleeded resin the amount of air could be assessed in the mould. When the bleeded resin was air free the

- 41 -

Development testing of a composite wing rib -

Material screening

outlet hole was closed with a conical steel pin. The mould temperature and the injection pressure were then adjusted to the correct values. After curing the press was opened, and the laminate demoulded when the mould was still hot. After demoulding the laminate was post-cured freestanding. After the postcuring the laminates were ready for cutting and testing. Double silicone ring sealing was used between the plates 1 and 2. Single silicone ring sealing was used between the plates 2 and 3. The manufactured laminate dimensions were 540 x 340 mm 2 . The individual injection parameters are presented in Figure 41 through Figure 44.

Resin sump

Double sealing

3.

Injection outlet

Bleed holes

1.

2.

Injection inlet

Figure 39. RTM flat mould used for manufacturing test laminates.

- 42 -

Development testing of a composite wing rib -

Material screening

Figure 40. 300-Ton press used for manufacturing test laminates.

The 3 mm nominal thickness of the mould was agreed to be used in the material screening. For comparability nominal fibre volume fraction for all test laminates was intended to be 64%. This with the nominal thickness defined the number of layers in the laminate. However, the intended fibre volume fraction could not be exactly achieved with fabrics of different thickness. It was decided to have the fibre volume fraction in the range from 60 to 65%. The theoretical fibre volume fraction can be calculated with Equation 1. Laminates were designed to be symmetrical and balanced when applicable, and for reasonable comparability laminate lay-ups were designed to have approximately the same percentual amount of fibres in all directions (0 , 45 and 90 ). At the time of launching the material screening phase the actual stacking orders for the wing rib laminates were not decided. However, at that stage it was agreed that significant directioning was not going to be used and the final lay-up will be approximately quasiisotropic. The material screening test laminates were designed to incorporate the same lay-up and stacking order, when applicable, as the final parts at level of knowledge at the time. The test laminate lay-ups and stacking orders are presented in Table 15, where layup shows the percentual amount of the fibres in the directions 0 / 45 / 90 in the same order. This kind of presentation is often used especially when laminates have many layers. Stacking order shows the actual layer stacking order. In the stacking order field all the directions of the layer are presented.

Table 15. Test laminate lay-ups in material screening.

Fabric

Lay-up

G0926

DB 550

Type 2 NCF

10% / 83% / 7%

16.7% / 66.7% / 16.7% 16.7% / 66.7% / 16.7%

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Development testing of a composite wing rib Stacking order Layers Nominal vf [0/90 45/-45 45/-45 45/-45 0/90]SO 9 62.4% [0/90 45/-45 45/-45]SE 6 61.8%

Material screening

[45/0/-45 45/90/-45 45/0/-45 45/90/45 -45/0/45] 5 61.0%

The cure cycles were defined by the used resin system. The same initial and post cure cycles were applied to the laminates with the same resin system to achieve the best possible comparability. Manufacturers normally recommend a number of possible cure cycles and the user can choose the most applicable for the given application. In this case, the cure cycles were chosen by making few initial curing tests and by applying the recommendations from the manufacturers. The cure cycles for the resin systems used in the material screening are presented in Figure 41 through Figure 44. The rate of temperature rise was 3-5 C/min that was dependent on the press heating system. The heating controller of the press did not allow to control the ramp. The rate of cooling was slower and highly dependent upon the size of the cooled block.

T, ( C)

Curing 75 min

p, (bar)

160 Bleeding time

120 70

5 4 Injection start

Resin temperature during the cure cycle Mould temperature at the beginning of the injection Injection pressure, p

t, (h)

Figure 41. Cure cycle for RTM6.

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T, ( C)

p, (bar)

Curing 60 min 120 Bleeding time 90 60 5 4 Injection start

Resin temperature during the cure cycle, T Mould temperature at the beginning of the injection, T Injection pressure, p

t, (h)

Figure 42. Cure cycle for 875RTM. T, ( C)

2. Curing 90 min 175 150 1. Curing 30 min

p, (bar)

60 RT

Bleeding time

4 3

Injection start

Resin temperature during the cure cycle Mould temperature at the beginning of the injection Injection pressure, p

t, (h)

Figure 43. Cure cycle for EX-1545-1.

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T, ( C)

2. Curing 120 min 170

p, (bar)

1. Curing 30 min 80 Bleeding time 30 RT Injection start 5 6

Resin temperature during the cure cycle Mould temperature at the beginning of the injection Injection pressure, p

t, (h)

Figure 44. Cure cycle for Epikote 862/Amicure.

Post curing was conducted freestanding after demoulding. Used post cure cycles are presented in Table 16.

Table 16. Post cure cycles for used resin systems.

Resin

RTM6 875RTM EX-1545-1 Epikote 862/Amicure

Post cure cycle

120 min at 180 C (1-2 C/min) 60 min at 175 C (1-2 C/min) 120 min at 200 C (1-2 C/min) 120 min at 170 C (1-2 C/min)

After a laminate was manufactured and post cured it went through a non-destructive inspection (NDI). The purpose of the NDI was to verify the integrity of the test laminates and further usability for the mechanical testing. Also several test standards for composite coupons include recommendation for NDI, usually an ultrasonic inspection, prior to the testing. The test laminates were inspected at PFC facilities by an ultrasonic inspection. The Ultrasonic-Testing System (UT-System) manufactured by Nukem GmbH consisted of two main parts: a processor controlled Transmitter-Receiver-Unit TRU 20.x a logarithmic Analogue/Digital Converter ADU 20.x

The data representation and manipulation is done by a DEC/Micro VAX-computer system including data storage. The data can be represented as a C-Scan or a D-Scan as well as combined presentation. The UT-system is presented in Figure 45.

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The UT-system is capable of two inspection methods: immersion and squirter operation, from which the first method was used to inspect the laminates. In this method both the part to be inspected and the transducers are submerged in water. Different measuring techniques can be used with immersion. In this case the Through Transmission technique was used. In the through transmission technique the ultrasonic signal passes through the part to the receiving transducer on the other side of the part. This technique is specially used when the part to be inspected has high attenuation values as composites typically have (as high as 1000 dB/m even at 0.5 MHz) [17]. Transmitter-Receiver set-up in water immersion is presented in Figure 46.

Figure 45. Nukem UT-system.

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Water Transmitter Test specimen

Receiver

Figure 46. Through transmission technique in immersion operation [ 18].

All the manufactured laminates passed the NDI. Significant attenuation at the ultrasonic scan revealing high porosity, voids or delaminations was not found. However, few observations were made: All the Bryte resin laminates had somewhat higher attenuation change in the laminate than the other materials. The level of porosity in the laminate can be assessed by the level of the ultrasonic scan attenuation according to reference [19]. However, using this method requires a calibration of the ultrasonic scan system for the given material. The calibration can be conducted by destructive testing e.g. acid burn test (Appendix 3). Figure 47 shows calibration curves for a typical CFRP laminate. The level of attenuation change in the laminate reveals the amount of porosity. The level of the attenuation change in the Bryte laminates was around 10-12 dB when attenuation change was around 6-8 dB for the other laminates. By comparing the measured 10-12 dB attenuation change to the curves in Figure 47c this could indicate 1.5-2.0 % void content for the 2.5-3.0 mm thick laminate. Also the higher deviation at the results of the mechanical tests indicated a higher void content in the laminate. Less than 2.0 % void content in the laminates was considered to be acceptable for the used RTM process. C-Scans of the Bryte laminate G0926-B-T-2 are presented in Appendix 1. Laminates including Devold's NCF fabrics (DB-) had large attenuation at the laminate edges. Normally a narrow strip right next to the laminate edge shows different attenuation values, since all the reinforcement layers do not reach to the edge. This attenuation occurrence was further investigated, but no reason for was found. Destructive and microscopic examination did not reveal anything specific. Also representatives from Devold had no explanations for this. It was decided to extract the specimens only from the area with a normal attenuation value. An example of this occurrence is presented in Appendix 2, where a G0926 laminate is compered with a DB laminate. Both laminates are manufactured with the Cytec's resin matrix.

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Figure 47. Typical production calibration curves for CFRP laminate [ 19].

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After the test laminates were manufactured and no faults inhibiting the mechanical testing were detected by the NDI, the test coupons were cut from the laminates. The cutting was conducted by using a small Computer Numerical Controlled (CNC) machine shown in Figure 48. However, after a number of cutting operations the CNC-machine lost its tolerances and after that the remaining coupons were cut at PFC facilities with another CNC-machine.

Figure 48. CNC-machine used for cutting the test laminates.

In some places the quality of the milling on the specimen edges was poor, hence the edges were trimmed and polished using a high-grade abrasive paper. The manufactured laminates were coded at the following manner (Figure 49): the first part describes the fabric, the second part describes the resin system, the third part describes the mould type and the last part is the running order number of that specific laminate type.

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G0926-R-T-1

Running order number Mould type Resin System Fabric G0926

DB NCF R = RTM6 C = 875RTM B = BRYTE E = EPIKOTE T = Flat mould (fin. Tasomuotti)

Figure 49. Coding of the laminates in material screening.

5.3.

Tensile tests

Tensile tests were conducted according to the standard Airbus Industrie Test Method (AITM) 1.0007 issue 2 [20], Determination of notched and unnotched tensile strength. This standard is defined and used by Airbus Industries. The nominal dimensions of the tensile test specimens are presented in Figure 50. Tab length l is a function of the testing machine; in this case l was 40 mm.

Figure 50. Nominal dimensions of a tensile test specimen.

Tensile tests were conducted with tabs to reduce local stress concentrations and to prevent the failure of the test specimen next to the hydraulic jaw grip. This has been the most usual reason for rejecting a single test result at the recent tensile tests conducted at the Laboratory of Lightweight Structures.

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Tab material was Integlas 92140 (390 gsm, 50/50%, E-Glass) in SP Systems Ampreg Epoxy matrix. Tab laminate (1000mm x 1000mm) was manufactured by VARTM with [0/90/0/90]SO stacking order giving the nominal thickness of 2.15 mm. After post curing 16h at 50 C, 40 mm wide strips were cut with diamond edge grinding machine from the tab laminate at 45-degree angle. The other longitudinal edge of the tab strip was chamfered to 45 . Tabs were then bonded onto the test specimen with FM 300-2 film adhesive. FM 300-2 offers service range from -55 C to 150 C and has good moisture resistance [21]. Some of the FM 300-2 properties are presented in Table 17.

Table 17. FM 300-2 Film adhesive properties [ 21].

Resin

FM 300-2M

Nominal weight (gsm)

145

Nominal thickness (mm)

0,14

Before bonding the tabs the test specimen end surfaces (the bond area for tabs) were finished with 180 grade abrasive paper and cleaned/degreased with Methyl Ethyl Ketone (MEK) solvent. A special aluminium bonding jig was used for bonding the tabs onto the tensile test specimens. The jig was used in order to bond the tabs precisely to correct positions and to prevent the tabs from sliding during curing. The jig is shown in Figure 51. The red foil is film adhesive. Upper tabs are not yet placed on the film adhesive. The left side end support is movable. After the film adhesives and tab laminates are placed in the jig the left side end support is moved to close the test specimen tightly between the two end supports. Then external weight is applied onto the tabs and the jig is placed into the oven. The used cure cycle for curing the film adhesive was 90 min at 120 C (30 min to 120 C) at approximately 0.10 MPa of pressure.

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Material screening

Figure 51. Bond jig for tensile test specimens.

The tensile tests were conducted at the Laboratory of Lightweight Structures with Dartec MK 100 + 9500 servo-hydraulic testing machine, which has maximum force of 100 kN. Tensile test set-up is presented in Figure 52. There were 3 specimens in each batch, which was agreed with other TANGO partners to be adequate number of specimens per batch for material screening/comparison purposes. The speed of testing was 2.0 mm/min. Unnotched tensile strength was calculated with equation Fmax b t is ultimate tensile strength, (MPa)

tu

(3)

where

tu

Fmax is the tensile failure load, (N) b is the width of the specimen, (mm) t is the thickness of the specimen, (mm) Thickness and width of the specimens were measured at three points and the smallest measured cross sectional area ( b t ) was used in equation 2 for determining tensile strength. Young' Modulus was measured with 5/50mm extensionmeter from the strain range of s 0.05 ­0.25%.

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Tensile strength and modulus were normalised to fibre volume fraction value 65% with Equation 4. Normalisation was executed in order to achieve better comparability of the materials. v f nor vf

X nor

X mea

(4)

where

X nor is the normalised value, (MPa, GPa) X mea is the measured value, (MPa, GPa) v f nor is the normalised value for fibre volume fraction, (%) v f is the measured fibre volume fraction, (%)

Table 18 shows the tensile test results. Results are normalised to 65% fibre volume fraction.

Figure 52. Tensile test set-up.

Table 18. Material screening tensile test results.

Laminate

G0926-R-T-1 G0926-C-T-2 G0926-B-T-2 G0926-E-T-1 DB-B-T-1 DB-C-T-1 NCF-C-T-2

Tensile Strength

Tensile Modulus

Mean, MPa Max., MPa Min., MPa cv, % Mean, GPa Max., GPa Min., GPa cv, % MP MP 497.6 519.9 464.1 5.81 45.2 47.3 44.1 3.97 450,4 453,6 446,7 0,77 47,4 475.3 477.6 473.1 0.47 45.5 48.6 43.5 5.99 515.1 533.6 504.7 3.12 43.5 45.9 41.4 5.27 470.7 482.8 460.2 2.41 42.5 44.9 39.8 6.11 433.0 457.1 391.1 8.42 47.5 49.1 46.1 3.16 304.6 306.6 302.5 0.68 33.5 34.4 32.8 2.48

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5.4.

Compression tests

Compressive tests were conducted according to the standard British Aerospace Airbus Test Specification (ABT) 1.0008 issue 2 [22], Determination of compressive strength and modulus of carbon, glass and aramid fibre composites. This standard is defined and used by AIRBUS UK. The nominal dimensions of the compressive test specimens are presented in Figure 53.

Figure 53. Compression test specimen nominal dimensions (in mm).

Compression tests were conducted utilising the " Celanese"type test fixture. The fixture was originally built according to standard ASTM D3410, but was modified for the specimen defined by ABT 1.0004. The compression test fixture is shown in Figure 54 and Figure 55. The test fixture has split collect-type grips at both ends: A, B and A' B' The grip cavities have filed face , . linings and alignment pins for proper closure. The grip nominal width and thickness were after modification 10.1 mm a 4.0 mm, respectively. The grips have an outer 10 conical taper and fit into sleeve (C) with a matching inner taper. These sleeves are received into a

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compactly fitting cylindrical shell (D) for ease of assembly and alignment but which is not load bearing during the test. A 10.0 mm thick preload spacer (E) is applied to separate the grips and allow them to be closed with a preload up to 75 kg without preloading the specimen. The assembled fixture with specimen inside is loaded between two steel platens. The other steel platen was also pivot jointed allowing swivelling around the centre line of the fixture.

Figure 54. Compression fixture [ 23]

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Development testing of a composite wing rib -

Material screening

Figure 55. Compression fixture dimensions [23]

New collect grips were designed and machined. The new collect grip cones incorporated removable jaws in order to provide wider thickness range of the test specimens by changing the removable jaws to favourable thickness. The thickness of the removable jaws was defined by the total specimen thickness with tabs. The collect grips fit to the sleeves only when the gap between the opposite collect grips is correct, in other words the thickness of the specimen is correct. It was tested that the correct gap between the collect grips was 1-2 mm. The nominal thickness of the post cured specimens was about 2.90 mm, tab thickness 1.25 mm, and cured FM300-2 adhesive film 0.14 mm giving total theoretical thickness t ctotal tc 2 t film 2 t tab 2,90mm 2 0,14mm 2 1,25mm 5,68mm

(5)

which defined the dimensions for removable jaws and jaw slots. Grip thickness was machined to 8 mm and grip slot to 10 mm. With this configuration the fixture would function with 5-6 mm thick specimens (including tabs).

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Development testing of a composite wing rib -

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1,0 ­2,0mm 5,0 ­6,0mm

Tab Collect grip Jaw

Adhesive film Test specimen

Figure 56. Celanese type test fixture grip thickness design.

Tab laminates were manufactured with the flat RTM mould to achieve good thickness tolerance. Since the nominal thickness of the flat mould was 3 mm, two tab laminates were manufactured in single injection. The desired thickness of 2.5 mm was achieved using four 0.13 mm peel plies in the mould. The stacking order is presented in Figure 57. Tab material was Interglas 92140 in Ciba Araldite 5025 LY matrix. The resin was injected at 30 C and cured at 60 C. Post curing was 60 min at 120 C The cured tab laminate thickness was 1.15-1.25 mm.

Peel ply 0.13 mm Plastic foil Glass woven fabric Flat RTM mould Figure 57. Stacking order for compression specimen tab laminates.

Tabs were bonded with FM300-2 film adhesive as referred above. Special bonding jig was used for achieving best possible tab location and orientation control. The bonding procedure was similar to the tensile test specimens described above. After bonding the test specimens were cut and finished by hand to nominal dimensions. Five specimens were tested per batch. Compressive modulus was measured from one specimen per batch using strain gages on both sides of the specimen. The active gage length was 2 mm (KYOWA KFG-2-120-C1-23). Strain gages were also used for confirming that no bending occurred during compression by monitoring strain differences of the two gages. The speed of testing was 1.5 mm/min. Compressive modulus was measured from the range of 0.05 ­0.25%.

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Development testing of a composite wing rib The compression strength was calculated with equation Fmax b t

Material screening

cu

(6)

where

cu

is ultimate compression strength of the unnotched specimen, (MPa)

Fmax is the maximum recorded tensile force, (N) b is the width of the specimen, (mm) t is the thickness of the specimen, (mm) Preload spacers were used in order to create a pre-compression onto the tabs from the grips. The preloading proved to be essential, since without it there was not enough compression to the tabs from the grips, and grips hardly managed to create enough compression to the tabs. The result of poor compression to the tabs was that tabs were sliding against the grips and no compression was applied to the specimen. Table 19 shows the compression test results. The results are normalised to 65% fibre volume fraction.

Table 19. Material screening compression test results.

Laminate G0926-R-T-1 G0926-C-T-2 G0926-B-T-2 G0926-E-T-1 DB-B-T-1 DB-C-T-1 NCF-C-T-2 Compressive Strength Compressive Modulus Mean, MPa Max., MPa Min., MPa cv, % Mean, GPa Max., GPa Min., GPa cv, % 340.3 357.5 320.4 4.38 41.8 42.3 41.2 1.89 286.1 312.0 270.9 5.75 46.8 47.4 46.2 1.71 301.8 324.0 291.1 4.42 42.4 44.3 40.6 6.13 318.9 325.3 306.2 2.34 37.6 41.4 33.8 14.32 291.9 300.0 279.0 3.20 41.2 43.1 39.3 2.69 293.1 302.8 269.9 4.57 39.7 41.0 38.5 4.47 354.4 375.3 326.1 5.28 33.5 36.6 30.5 12.80

5.5.

Interlaminar Shear Strength tests

Interlaminar Shear Strength tests (ILSS) tests were conducted according to the standard BS ISO EN 14130 [24], Fibre-reinforced plastic composites ­Determination of apparent interlaminar shear strength by short-beam method. European Committee approves this standard as a European standard. This method is a 3-point bending of a beam, where the distance between outer supports is short enough to create a significant interlaminar shear stress compared to the stresses in the laminate layers. The nominal dimensions of the test specimen are relative to the thickness of the specimen. The specimen has the same length/width to thickness ratio as the standard specimen (10 x 20 mm2), accordingly: the length of the specimen is l 10t the width of the specimen is b 5t

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The nominal thickness of the specimen is 3 mm. According to above rules the nominal length is 30 mm, and nominal width is 15 mm. Nominal diameter of the roller struts is 4 mm, and the press nominal diameter is 10 mm. The roller strut nominal distance is 3t. The test set-up is presented in Figure 58 and Figure 59. The edges of the test specimens were polished with high-grade abrasive paper for exposing the layers and verifying that interlaminar failure occurred at the test. Standard BS ISO EN 14130 defines acceptable and unacceptable failure modes in the similar manner as in Figure 60 [25], where failure modes a through c describe the acceptable failure modes and the rest are more or less unacceptable failure modes. Unacceptable failure modes are non-shear (e and f), mixed (d and e), compression (e), tension (f) and plastic shear (d and e) modes. 10 4 t

3t

Figure 58. ILSS test set-up.

Figure 59. ILSS test set-up.

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Material screening

Figure 60. Types of failure in ILSS tests.

BS ISO EN 14130 recommends for the speed of testing 1 mm/min in the absence of any better information. However, 1 mm/min seemed to be slight too fast, because failure occurred in 5-10 seconds, hence 0.5 mm/min was chosen for the speed of testing. Five specimens per batch were tested. Interlaminar shear strength was calculated according to BS ISO EN 14130 3Fmax 4bt

ILSS

(7)

where

Fmax is the maximum recorded force before first failure (N)

ILSS material screening results are presented in Table 20. The results are normalised to 65% fibre volume fraction.

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Table 20. ILSS material screening results. Laminate G0926-R-T-1 G0926-C-T-2 G0926-B-T-2 G0926-E-T-1 DB-B-T-1 DB-C-T-1 NCF-C-T-2 Interlaminar Shear Strength Mean, MPa Max., MPa Min., MPa cv, % 60.3 62.7 58.2 3.09 55.0 56.6 53.0 2.64 54.4 56.7 50.8 4.33 58.4 60.8 56.0 3.42 50.7 57.9 45.5 9.33 47.0 50.1 45.2 4.38 37.2 38.9 36.3 2.74

5.6.

Layer level analysis

All the tested laminates did not incorporate the same stacking order or lay-up. In order to compare the tested materials a layer level detail analysis was conducted. The analysis was conducted using ESAComp 2.0 composite analysis software. For comparison purposes the tested laminates were constructed with ESAComp. These laminates were constructed out of single uni-directional (UD) plies in order to assess the layer stresses at each fibre direction. The tested laminates were constructed by the following manner: The thickness of a single layer was calculated by dividing the thickness of a test laminate by the number of the layers in the laminate. A multiaxial layer was further divided into UD layers. The thickness of a single UD layer was derived according to the percentual areal mass of the UD layer in the fabric, e.g. 0.335 mm thick 0 /90 5H Satin fabric was constructed of 0.1675 mm 0 -layer and 0.1675 mm 90 -layer. The fact that layers in the woven fabrics cross each other can not be simulated with ESAComp; thus layers in the ESAComp models locate at different depths. However, since the layers are considerably thin the error is negligible. Values provided by a European aircraft manufacturer were used for Young's modulus E21, shear modulus G12 and Poisson's ratio 12. These were considered as the best data available. Material values were measured for G0986 (2x2 Twill, HTA) and RTM6 material. Material properties for HTA and HTS fibres are similar. HTS fibre has only improved strength properties over HTA fibre. These UD laminate ply properties are presented in Table 21.

Table 21. Material properties for layer level analysis.

Property

E21 G12

12

HTA

4.65 GPa 4.65 GPa 0,35

HTS

4.65 GPa 4.65 GPa 0,35

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The properties are not exactly the same for all tested materials, but better UD ply data was not available. Furthermore, the properties in Table 21 do not have great significance for layer level stress analysis. Acquired precision is well acceptable for comparison purposes. Young's modulus E12 was iterated by setting values for E12 for a single UD ply and by comparing constructed laminate values to the measured tensile and compression mean E12 values. The constructed laminates were loaded with the mean ultimate loads derived by the actual mechanical testing. As the results the layer stresses at the ultimate loads are derived. An example of the layer stress analyses is shown in Figure 61, where the small deviation in the tensile stress inside the laminate derives of the unbalanced stacking (the laminate is constructed with UD-layers). The ultimate layer stresses from tensile and compressive tests are collected in Table 22. The results shows that the G0926 in RTM6 matrix has the best overall mechanical performance. No significant difference in the mechanical performance between the materials was detected. However, the compression performance of the NCF-C-T-2 laminate was noticeable. Few reasons for the exceptionally good compression performance was considered: smaller amount of fibres at 0 -direction the surface fibres at 45 -direction. This subject is further discussed in Chapter 5.7. the level of crimp of the fibres is smaller with the NCF.

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Layer stresses: Actual - sig_1

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Laminate : G0926-R-T-1 Modified : Tue Jan 09 11:54:13 2001 Lay-up : (0a/90a/+45a/-45a/+45a/-45a/+45a/-45a/0a/90a/-45a/+45a/-45a/+45a/-45a/+45a/90a/0a)

Ply a HTA - G0926-R-T-1

Load : Veto Modified : Tue Jan 09 10:57:07 2001 Type : Forces and moments (Var.;E)

N_x = 1.44133e+006 N/m N_y = 0 N/m N_xy = 0 N/m M_x M_y M_xy = 0 Nm/m = 0 Nm/m = 0 Nm/m

Figure 61. G0926-R-T-1 laminate layer stresses at tensile failure.

Table 22. Test laminate layer stresses at the ultimate loads Laminate G0926-R-T-1 G0926-R-T-1 G0926-C-T-2 G0926-C-T-2 G0926-B-T-2 G0926-B-T-2 G0926-E-T-1 G0926-E-T-1 DB-B-T-1 DB-B-T-1 DB-C-T-1 DB-C-T-1 NCF-C-T-2 NCF-C-T-2 Test type Tensile Compression Tensile Compression Tensile Compression Tensile Compression Tensile Compression Tensile Compression Tensile Compression Ultimate load (MN/m) 1.44 -0.98 1.28 -0.80 1.35 -0.84 1.48 -0.90 1.33 -0.81 1.24 -0.83 0.92 -0.97 Ultimate layer stress (MPa) 1730 -1180 1530 -951 1660 -1030 1660 -1010 1580 -970 1470 -990 1390 -1480

5.7.

Summary of material screening tests

The purpose of the material screening program was to create clear understanding of the mechanical performance of the selected materials compared to each other. This information was successfully gained.

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Manufacturing the test laminates also gave valuable information on the injection parameters and on general suitability of these materials for the future RTM process parts. Following notifications significantly affecting to the material selection and results were made during the manufacturing and testing the test laminates: During the compression test it was detected that the surface layer (0-degree layer) with G0926 and DB laminates had the tendency to buckle before the actual compressive strength of the coupon was reached. Figure 62 shows an example of this state. In some cases the buckling was also detectable in the stress-strain curve. With NCF-C-T laminate this was not the case apparently due to the 45-degree surface layer. Also the maximum strain with NCF-C-T was significantly higher.

Figure 62. Buckling of the surface layer during the compression test.

Apparently the EB6 binder applied to the AIRBUS NC fabrics was not compatible with Cycom 875RTM resin. This was noticed after the NCF-C-T laminates were cured. The binder did not dissolve into the resin matrix; instead it was left in the matrix as detectable spots. Apparently this happened due to the fact that EB6 is RTM6 compatible and the curing temperature for RTM6 is 160 C when the curing temperature for Cycom 875RTM is 120 C. Clearly the curing temperature of 120 C is too low for the EB6 binder. This incompatibility was shown also in the mechanical performance of the NCF-C-T laminates, especially in the ILSS result, which was considerably lower than for the other materials. The material screening results are presented in more comparable form in Figure 63 through Figure 67. The NCF-laminate lay-up was different from the other laminates. NCF laminate included only 10% of 0-degree fibres. The difference is accounted by a

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Development testing of a composite wing rib -

Material screening

following manner: the tensile modulus of the NCF laminate was 34.4 GPa. The Young's modulus of the NCF laminate is 45.2 GPa with 16.7% / 66.6% / 16.7% lay-up. With the same ultimate stain level the ultimate tensile stress in the laminate is 45.2 / 34.4 304.6 MPa = 400.2 MPa. Same kind of compensation is used for the compression result. This compensation is marked in Figure 63 through Figure 66 with diagonal lines. This kind of calculation presumes that the same ultimate strain level is reached with both lay-ups. Based partly on the mechanical performance and primarily on the processability, the RTM6 resin was selected as a matrix material for all five wing ribs. Drapeability of the bindered 5H Satin over thick bindered NC fabrics is significantly better and lead to the selection of the G0926 INJ fabric. NC fabrics have been considered to have better mechanical strength over 5H Satin fabrics. This assumption is usually based on the fact that fibres are straighter or the level of crimp is smaller in NC fabrics than in 5H Satin fabrics. For example in reference [26] NC Fabrics have been reported to have over 30% higher mechanical and stiffness properties than woven fabrics. This data was based on the same modulus reinforcement fibre, the same fibre pattern but using different resin matrix and manufacturing method. The woven fabric used a " stage epoxy prepreg product and autoclave cure where the B" NC fabric used an Epoxy Resin Film Infusion (ERFI) process and autoclave cure. The results of the study are presented in Table 23. However, this kind of behaviour was not detected in this material screening program as the mechanical performance of NCF laminates was not better than that of 5H Satin laminates but vice versa.

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Tensile Strength (MPa)

NCF-C-T

DB-B-T

DB-C-T

G0926-B-T

G0926-C-T

G0926-E-T

G0926-R-T 250 300 350 400 450 500 550

Figure 63. Tensile strength comparison.

Tensile Modulus (GPa)

NCF-C-T

DB-B-T

DB-C-T

G0926-B-T

G0926-C-T

G0926-E-T

G0926-R-T 30 32 34 36 38 40 42 44 46 48 50

Figure 64. Tensile modulus comparison.

Compressive Strength (MPa)

NCF-C-T

DB-B-T

DB-C-T

G0926-B-T

G0926-C-T

G0926-E-T

G0926-R-T

250

300

350

400

450

500

Figure 65. Compressive strength comparison.

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Compressive Modulus (GPa)

NCF-C-T

DB-B-T

DB-C-T

G0926-B-T

G0926-C-T

G0926-E-T

G0926-R-T 30 32 34 36 38 40 42 44 46 48

Figure 66. Compressive modulus comparison.

ILSS (MPa)

NCF-C-T

DB-B-T

DB-C-T

G0926-B-T

G0926-C-T

G0926-E-T

G0926-R-T 30 35 40 45 50 55 60 65

Figure 67. ILSS result comparison.

Table 23. Mechanical properties of NCF and woven fabric from reference [26].

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Design data generation

6. Design data generation

6.1. Introduction

As a result of the material screening program and manufacturing trials RTM6 resin and G0926INJ reinforcement were selected for the final material of all five wing ribs. The selection of the material properties to be tested for G0926INJ in RTM6 matrix was based on: the general test requirements (Table 5) discussions with AUK the agreement with stress department. In order to perform stressing analyses of the wing ribs the design data values (material properties) had to be generated. This design data includes material allowables and Young' modulus in different directions. To reduce the number of the specimens and s effort put into the design data generating the method described in Chapter 3.3 was used to define B-values. All the test laminates were prepared at HUT/LSS and some of the tests were conducted at PFC in Halli. The test matrix of the design data generation tests is presented in Table 24.

Table 24. Test matrix of the design data generation tests.

TEST Tensile

RT / AH Hot/Wet

STANDARD SPECIMENS TABS WHERE

AITM 1.0007 AITM 1.0007 8 8 Yes Yes PFC HUT/LLS

Compression

RT / AH Hot/Wet ABT 1.0004 ABT 1.0004 8 8 Yes Yes PFC HUT/LLS

OHT 1

RT / AH Hot/Wet AITM 1.0002 AITM 1.0002 8 8 No No PFC HUT/LLS

OHC 2

RT / AH BS ISO EN 14130 Hot/Wet BS ISO EN 14130 8 8 No No PFC HUT/LLS

Shear

RT / AH Hot/Wet ASTM 5379M ASTM 5379M 8 8 Yes Yes HUT/LLS HUT/LLS

ILSS

RT / AH BS ISO EN 14130 Hot/Wet BS ISO EN 14130 8 8 No No PFC HUT/LLS

1 2

Open Hole Tensile Open Hole Compression

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RT / AH Hot/Wet 6 4

Design data generation

No No

HUT/LLS HUT/LLS

Bearing Strength

RT / AH Hot/Wet AITM 1.0009 AITM 1.0009 8 8 No No HUT/LLS HUT/LLS

6.2.

Preparing of the test laminates

The test laminates were manufactured in the similar manner as in the material screening program. The only difference was that the design data was to be generated for the quasiisotropic laminate with Vf of 62%. Such laminates could not be manufactured with 3 mm flat mould available without modifications. The thickness was adjusted using extra peel plies in the mould. Theoretical nominal thickness of the laminate with Vf of 62% is n mA 10

f

h

vf

8 370 mm 2,68 mm 10 1,78 62

(8)

A test laminate was manufactured in order to determine the amount of peel plies needed to achieve nominal 2.68 mm laminate thickness. One 0.18 mm thick peel ply was added in the mould on both sides of the laminate covering half of the laminate area. The effect of adding peel plies on the thickness of the laminate was determined by measuring the thickness of the cured laminate from the area with applied peel plies and without peel plies. Mean thickness of the test laminate without the peel plies was 2.94 mm and with the peel plies 2.70 mm. According to the results laminates were to be manufactured by using two extra 0,18 mm peel plies in the mould. After manufacturing two laminates the thickness of the laminates from the edge area was too high resulting in too low Vf. Measured mean thickness was 2.82 mm. Two 0.13 mm peel plies were added to the mould for the rest of the test laminates. All manufactured test laminates are presented in Table 25. Layer directions presented are warp directions of the layers. G0926 without binder was used in the three last laminates -T7 to -T9, because the bindered G0926 ran out.

Table 25. Manufactured test laminates for design data generation. A is 0.18 mm Peel ply B is 0.13 mm Peel ply.

MATERIAL

G0926 I + RTM6 G0926 I + RTM6 G0926 I + RTM6 G0926 I + RTM6 G0926 + RTM6 G0926 + RTM6

LAMINATE CODE

G0926-RT-DD-T2 G0926-RT-DD-T3 G0926-RT-DD-T4 G0926-RT-DD-T5 G0926-RT-DD-T7 G0926-RT-DD-T8

STACKING ORDER

[0,5A/45/0/90/-45/-45/90/0/45/0,5A] [B/45/0/90/-45/-45/90/0/45/B] [B/45/0/90/-45/-45/90/0/45/B] [B/45/0/90/-45/-45/90/0/45/B] [A/B/45/0/90/-45/-45/90/0/45/B/A] [A/B/45/0/90/-45/-45/90/0/45/B/A]

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Design data generation

[A/B/45/0/90/-45/-45/90/0/45/B/A]

All the test laminates were inspected by ultrasonic C-Scan (Chapter 5.2) before further processing. No detectable defects were in any of the manufactured laminates. Test coupons were extracted from the test laminates by CNC machining. IGES files were first created with I-DEAS-software for establishing the specimen geometry. The IGES files were imported into the CNC machine' controller. s

6.3.

Tensile tests

Tensile tests were conducted according to the standard Airbus Industrie Test Method (AITM) 1.0007 issue 2. In this standard two test specimen shapes are defined; unnotched and notched. AUK uses the geometry in Figure 68 for unnotched tensile tests without the hole. For creating a consistent design data record with the other partners, especially with AUK, notched shape test specimens without the hole were used as the tensile test specimens. The notched shape in this case refers also to the Open Hole Tensile (OHT) specimen. The notched shape tensile specimen nominal dimensions according to AITM 1.0007 are presented in Figure 68. Tab length l was 40 mm.

Figure 68. Notched tensile test specimen.

The tensile tests were conducted as described in Chapter 5.3. The room temperature tensile tests were conducted at PFC, Halli and hot/wet tensile tests at HUT/LLS. Hot/wet testing was conducted at the environmental chamber at 80 C temperature. The test specimens were exposed at least 1000h at 70% relative humidity in the Weiss Technik SB22 climate chamber. More detailed information about hot/wet behaviour is presented in Chapter 6.12.

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Tabs were used for ensuring a correct failure position. The tab material was the same as used in the material screening tests (Chapter 5.3). Room temperature compression modulus was measured using a 1/70 mm extensionmeter. The hot/wet specimens were instrumented with KYOWA KFG­ 5-120-C1-11L1M2R strain gages that were suitable for temperatures up to 80 C. The temperature compensation of the test was carried out with a " dummy-gage"set-up. " Dummy-gage"is an identical configuration with the active gage and it is subjected to the same thermal environment as the active gage but remains unaffected by the mechanical forces applied to the structure on which the active gage is mounted. The identical dummy-gage coupon was extracted from the same laminate as the active coupon. The leadwires of the two gages were of equal length and were subjected to identical environment. The active and the dummy gages were installed in adjacent arms of a Wheatstone bridge circuit resulting in automated thermal compensation. The dummygage compensation system is illustrated in Figure 69 [27]. An instrumented test specimen in the environmental chamber is presented in Figure 70.

Figure 69. " Dummy-gage"Wheatstone bridge compensation system.

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Test specimen

Dummy specimen

Figure 70. Instrumented tensile test specimen in the environmental chamber. Dummy specimen is on the left.

6.4.

Compression tests

The compressive tests were conducted according to the AITM 1.0008 issue 1 [28], Determination of notched, unnotched and filled hole compression strength. The length of the specimen is a function of the testing machine. The 40 mm tab length was preferred, therefore the nominal length of the specimen was 110 mm. The nominal dimensions of the compressive test specimens are presented in Figure 53.

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Figure 71. Compression test specimen according to AITM 1.0008.

Test specimen was attached directly to the testing machine jaws. No specific test fixture was used. In the DARTEC testing machines of HUT/LLS the stroke actuator cannot be restrained from rotating around its axis. However, actuator seemed to be stiff enough and rotation during the testing was not detected. The recommended thickness of the specimen is 4.0 mm for a quasi-isotropic lay-up. The nominal thickness of the specimens was 2.68 mm, hence stability during the testing was a concern. Monitoring of the stability was implemented by instrumenting one strain gage on both sides of the laminate. A large difference in the strain readings prior to failure would indicate instability of the specimen. The speed of testing was 0.5 mm/min. The compressive modulus was measured from the strain range of 0.05 ­ 0.25%. Mean value from the two strain readings was used for elongation. An identical system was used for testing hot/wet specimens as in tensile tests (Chapter 6.3). Special extension posts were designed for hot/wet compression test in the environmental chamber. Due to the dimensions of the chamber the minimum distance of the jaws is limited. According to AITM 1.0008 the required distance is 30 mm. The detachable jaws are mounted on the four extension posts (Figure 72). The extension posts were manufactured from 50 mm diameter steel bar. The length of a post is 140 mm and M10 thread is machined on both ends for attaching the post to the testing machine and the lower jaw to the post. The upper jaw is attached normally to the load head. Instability was not detected with the RT/AH specimens, but the hot/wet specimens did indicate instability during the test. A stress-strain curve of a single test (G0926-RT-DDT7_7 HW) is presented in Figure 73. The two curves clearly diverge prior to the failure. This means that the specimen starts to bend. In this particular case the buckling starts approximately at 320 MPa. However, the deviation of the compression strength was quite

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Development testing of a composite wing rib -

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small (5.09%) and the results were even higher than expected, therefore the effect of instability is not significant.

Test specimen

Extension posts

Dummy specimen

Figure 72. Hot/Wet compression test.

-400,00 -350,00 -300,00 -250,00 -200,00 -150,00 -100,00 -50,00 0,00 0,00

-0,20

-0,40 -0,60 Strain (%)

-0,80

-1,00

Figure 73. Stress-strain curve of a Hot/Wet compression test indicating instability.

6.5.

In-plane shear tests

In-plane shear tests were conducted according to the standard ASTM 5379M-98 [29], Standard Test Method for Shear Properties of Composite Materials by the V-Notched Beam Method. This test method is also referred to as the Iosipescu shear test.

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In this test method a material coupon is a rectangular flat strip with symmetrically centrally located v-notches (Figure 74). The coupon is loaded with a mechanical testing machine in a specially designed testing fixture. A schematic picture of the fixture is presented in Figure 75. The specimen is inserted into the test fixture with the notch located along the axis of loading. The two halves of the testing fixture are compressed by the testing machine while monitoring the load. The relative displacement of the two halves loads the notched specimen. By placing two strain gage elements, oriented at 45 to the loading axis, in the middle of the specimen and along the loading axis, as illustrated in Figure 76, the shear response of the material can be measured. The loading can be idealised as asymmetric flexure, as shown by the shear and bending moment diagrams in Figure 77. The notches influence the shear along the loading direction, making the distribution more uniform than would be seen without the notches. The value of the dimension b in Figure 77 is not critical to the concept.

Figure 74. Iosipescu shear test specimen.

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Development testing of a composite wing rib -

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Figure 75. V-notched test fixture schematic.

Figure 76. Strain gage locations.

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Design data generation

Figure 77. Idealised Force, Bending and Shear diagrams.

While the degree of uniformity of the shear loading is a function of material orthotropy, the best overall results, when testing in 1-2 plane, have been obtained on [0/90]ns -type laminates [29]. The test fixture used was designed in the early 90' at HUT/LLS before the standard, s which was introduced in this context, was released. However, the operational principle of the two fixtures is exactly the same. The only difference is that the alignment pin and tightening rods are missing from the HUT/LLS fixture. Also the structures of the two fixtures differ slightly. However, the test specimen is supported in HUT/LLS fixture by two rectangular beams that can rotate in a way that beams are able to adapt to different specimen dimensions and produce evenly distributed loading, therefore tightening is not needed. Test specimen was placed in the test fixture exploiting the lines marked that

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Development testing of a composite wing rib -

Design data generation

represent the axis of loading, on the test fixture. The HUT/LLS test fixture is presented in Figure 78.

Figure 78. HUT/LLS test fixture.

Preliminary tests demonstrated that tabs should be applied to the specimens, since an undesirable failure (Figure 79) occurred. Typical acceptable failure modes are presented in Figure 80. The used tab configuration is presented in Figure 81. The width of a tab is the same as the width of the specimen. The used tab material was the same described in chapter 5.3.

Failure

Failure

Figure 79. Undesirable failure mode

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Design data generation

Figure 80. Typical acceptable failure modes.

Figure 81. Tab locations.

The ultimate shear strength of the material is Fmax A where Fmax A is the ultimate shear strength (MPa) is the maximum recorded force (N) is the cross-sectional area of the specimen at notch (mm2)

(9)

The shear strain is determined from indicated normal strains, measured with the strain gages. The shear strain is

i

45

45

(10)

where

i 45 45

is shear strain at each data point ( ) is +45 normal strain at each data point ( ) is -45 normal strain at each data point ( ) 200

The shear modulus is calculated using Equation 10 nominally from 4000 (0.4% 0.02%) strain range starting from the nominal strain of 1500 (0.15%).

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Design data generation

G

(11)

where G

is shear modulus (GPa) is difference in shear stress between two applied shear points (MPa) is difference between two shear points (nominally 0,4%)

ASTM 5379M recommends 1.5 mm for active gage length. However, that gage length was not available, therefore 1.0 mm rosetta gages were agreed to be used. It would have been favourable to use longer gages especially when testing woven fabrics. Now the available space is limited and previous experience on the same material and gage types supported the use of shorter gages as 1.5 mm. The speed of testing was 1.0 mm/min. Figure 82 shows typical stress-strain curves of the two strain gages for RT/AH specimen. +45 and ­ refer now to the directions at which they are aligned at the specimen co45 ordinate system. In this case, comparing to the Figure 76, the left side of the specimen moves up and right side moves down resulting in a stress-strain curve as below. Summing the strains at single stress points gives the stress-strain curve from which the shear modulus can be determined. This curve is presented in Figure 83.

300,00 250,00 200,00 150,00 100,00 50,00 0,00 -0,75 Strain +45 Strain -45

-0,50

-0,25

0,00

0,25

0,50

0,75

1,00

Strain (%)

Figure 82. Typical stress-strain curves of the two strain gages for RT/AH specimen.

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Design data generation

300,00 250,00 200,00 150,00 100,00 50,00 0,00 0,00 Shear strain

0,50

1,00 Shear strain (%)

1,50

2,00

Figure 83. Stress-strain curve of typical RT/AH specimen.

6.6.

Open hole tensile tests

Open hole tensile tests (OHT) were conducted according to the AITM 1.0007 issue 2. OHT tests were conducted in the similar manner as the unnotched tensile tests. The only difference was that the specimens were notched. The nominal dimensions of the specimen are presented in Figure 68. Tabs were not used in the OHT tests, since failure occurred without exception at the hole area due to the stress concentrations around the hole. The ultimate tensile strength of the OHT specimens is Fmax b t is ultimate tensile strength of the notched specimen (MPa) is maximum recorded tensile load (N) is the width of the specimen (mm) is the thickness of the specimen (mm)

tn

(12)

where

tn

Fmax b t

6.7.

Open hole compressive tests

Open hole compression (OHC) tests were conducted according to the AITM 1.0008 issue 1. The nominal dimensions of the specimen are presented in Figure 84. The ultimate compression strength of the OHC specimen is

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Design data generation

cn

Fmax b t

(13)

where

tn

is ultimate compression strength of the notched specimen, (MPa) is maximum recorded tensile load (N) is the width of the specimen (mm) is the thickness of the specimen (mm)

Fmax b t

Figure 84. The nominal dimensions of the notched compression specimen.

6.8.

Interlaminar shear tests

The ILSS tests were conducted according to the standard BS ISO EN 14130. The testing was similar as conducted in the material screening program (Chapter 5.5).

6.9.

Bearing tests

The bearing tests were conducted according to the standard AITM 1.0009 Issue 2 [30], Determination of bearing strength. The principle of the test is simple. A tensile test is carried out on a flat laminate containing a hole; a bolt passing through the hole applies the load. A large specimen width is chosen, so that bearing failure occurs. The yield and ultimate bearing strengths can be determined. This test required a special loading fixture to be used. AITM 1.0009 defines two possible means to attach the specimen to the testing machine: attaching directly to machine jaws or attaching by bolt to the testing machine. For convenience it was decided to use the first choice. Tabs were not used, since large area of the specimen was available for hydraulic jaw grips. The recommended nominal thickness of the test specimen is 4 mm, but this

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Development testing of a composite wing rib -

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thickness was not available, instead, 2.68 mm thick specimens were used. The nominal dimensions of the bearing test specimen are presented in Figure 85. A schematic picture of the bearing test fixture is presented in Figure 86. The test fixture was designed based on the features in Figure 86 and noticing that later demand for bearing tests could include different thicknesses and shapes. The bearing test fixture is presented in Figure 87. The clamping strips of the test fixture are adjustable allowing different thicknesses to be tested. A spring on the clamping bolt between the clamping strips holds the strips against the adjustment screws. Plastic foils were used between the test specimen and strips to ensure that the required 0.1mm gap existed during the test. The foils were placed above the pin high enough not to provide any support for the test specimen at the hole area. An external clamp and rubber band was used to ensure the alignment of the strips. The external clamp and rubber band kept the strips lightly compressed against the plastic foils (clamp not showing in Figure 87).

Figure 85. Nominal dimensions of a bearing test specimen according to AITM 1.0009.

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Design data generation

Figure 86. Bearing test fixture according to AITM 1.0009.

Hydraulic jaws

Test specimen

Pin

Plastic foils

Adjustement screws

Clamping bolt

Figure 87. The bearing test fixture.

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AITM 1.0009 defines the procedure to measure yielding stress of the specimen. Yield stress was not measured, since yielding was not detected during the test. Yield stress is a bearing stress, which results in 0.02 times the diameter permanent set of the joint. Yield stress calculation is illustrated in Figure 88.

Figure 88. Calculation of bearing test yield stress.

6.10. Through thickness tests

No standard has yet been established for measuring through thickness properties of a composite laminate. For measuring through thickness properties it was agreed to use a draft proposal from National Physical Laboratory (NPL), UK. The NPL draft proposes three different types of test specimens. The through thickness properties, which can be determined by these three specimen types, are: Strength, Modulus, Strength and Modulus. The last one was used in this program. Test specimen geometry and dimensions are presented in Figure 89.

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R12 12 40

5 16 25

Figure 89. Through thickness test specimen.

Through thickness test specimens were extracted from 39 mm thick laminate manufactured by RTM. Originally this RTM laminate was designed for other purposes, but it was suitable for the use in TANGO, since the materials were the same. The RTM block material was Hexcel` G0986I (2x2 Twill weave, 285 gsm, HTA, 6K) in RTM6 s matrix. The laminate consists of 117 layers and the actual stacking order is: [lam1 / lam2(5) / lam2 / lam2(5) / lam1] where lam1 is [0/45/45/45/0]SO lam2 is [0/45/0/45/0]SO The mean height (or thickness of the laminate) of the test specimens was 38.7 mm. Using the equation 1 we get the fibre volume fraction of 117 285 % 48.4% . 10 1.78 38.7

vf

A steel block with a hole was bonded with Hysol EA934 structural adhesive on both ends of the specimen (Figure 90). Holes were aligned perpendicularly to each other. A thin scrimp fabric was applied for the bond to prevent a too thin bond line.

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Design data generation

Figure 90. Through thickness tensile test specimen.

Figure 91 shows the test set-up for through thickness tensile test. Two rods going through the metal blocks loaded the specimen. The rods were aligned perpendicular to each other in order to give free rotation in two directions and further to ease setting up the test and decreasing undesired bending moment during the set-up and the test. Two strain gages (Kyowa KFG-5-120-C1) were used on perpendicular sides (Figure 91) to determine mean Young' modulus in 31-direction. s

Figure 91. Through thickness tensile test set-up.

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The problem in the test is that the radius along the measuring range creates a stress concentration on the radius area. An FE-model of the specimen was built with PATRAN/NASTRAN software in order to assess the stress concentration factor and the stress concentration areas. Figure 92 shows the FE-model and boundary conditions. On the upper flat surface all nodes are restricted and on the lower surface all but z-direction is restricted. There is also a restriction that keeps the lower surface planar.

Figure 92. FE-model of the through thickness test specimen.

Table 26 shows the material properties used for analyses. Elastic modulus 33 is input from the through thickness tensile tests.

Table 26. Material properties used for through thickness analyses.

Property Elastic Modulus 11 (GPa) Elastic Modulus 22 (GPa) Elastic Modulus 33 (GPa) Poisson Ratio 12 Poisson Ratio 23 Poisson Ratio 31 Shear Modulus 12 (GPa) Shear Modulus 23 (GPa) Shear Modulus 31 (GPa)

Value 41.5 41.5 8.90 0.253 0.100 0.021 14.1 5.0 5.0

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Figure 93 illustrates the stress concentrations at the radius area. Analysis results were adjusted in the way that the stress tensor in z-direction at the measuring range in the middle was 1.0. Results show that the stress concentration factor is 1.44 for this orthotrophic configuration. The fact that stress concentrations exist was detected in the actual test also. All specimens failed at same point/area where the stress concentrations located according to analyses (Figure 94).

Figure 93. Result of the FE-analysis.

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Design data generation

Figure 94. Failure mode of the through thickness tensile test.

Through thickness compression tests could not be continued up to the specimen failure, since the maximum compressive force of the testing machine (100 kN) was inadequate to achieve failure. When failure in tensile test occurred at 25-30 MPa the compression test specimens sustained 390 MPa (at 100 kN) compression stress. Even if the compressive modulus can be obtained from the test, the compression strength result is not useful in terms of design data. The critical load is tension in terms of through thickness stress. Therefore, it is rational to use the tensile properties as the design values.

6.11. Support tests

In order to ensure the quality of the laminates, their fibre volume fractions and glass transition temperatures were measured chemically. Fibre volume fraction was measured by acid burn test according to a military standard. Reference is not available for this standard. The method is simple and a similar method is presented in Appendix 1. Glass transition temperature (Tg) was measured by Thermal Mechanical Analysis (TMA) at Pirelli Cables and Systems Oy. The used TMA system was Mettler TMA-40 Expansion TMA system. Expansion TMA measures the Coefficient of Thermal Expansion (CTE) as a function of temperature. Tg is determined at the intersection of the CTE line tangents. An example of Tg measurement is presented in Appendix 2. Another purpose of Vf measurement was to create a curve for Vf as a function of laminate thickness. However, data gathered by acid burn test was not very reliable. Thicker

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specimens (all with the same number of layers) seem to have higher Vf according to test results. Fibre volume fraction can be calculated from fibre mass volume fraction with equation wf / wf /

f

Vf

f

wm /

(14)

m

where wf

f

wm

m

is the fibre mass fraction (%) is the density of the fibre (g/cm3) is the matrix mass fraction (%) is the density of the matrix (g/cm3)

The fibre volume fraction results are presented in Table 27. The glass transition results are presented in Table 28

Table 27. Fibre volume fraction results by acid burn method. 1

Laminate DD-T8 DD-T9 DD-T6

Number of specimens 3 3 3

Thickness (mm) 2.50 2.56 2.70

wf (%) 74.0 76.0 76.0

Stdv of wf (%) 5.45 1.32 1.32

Vf Stdv of Vf (%) (%) 64.6 67.0 67.0 5.45 1.32 1.32

Table 28. Tg results by TMA.

Laminate Number of specimens DD-T6 DD-T3 3 3

Tg (C) 203 203

Stdv of Tg (%) 5.42 0.05

6.12. Moisture behaviour of the selected material

Most of polymer composites absorb moisture. In CFRP laminates the moisture is absorbed by the resin matrix. Carbon fibres have not been stated to absorb moisture. As a result of moisture absorption the resin matrix is softened and strength of the laminate is decreased. The strength of polymer composite is also decreased when the temperature is raised. Materials exposed to moist conditions are usually tested at raised temperature in order to define the lowest design value imaginable.

1

1.14 g/cm3 used for matrix density, 1.78 g/cm3 used for fibre density.

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When aircraft structures are designed the moisture exposure and surrounding test temperature of the test elements is a function of conditions of the part life cycle inside the aircraft. These environmental parameters are defined. The purpose of hot/wet testing is to simulate these conditions. Naturally test elements cannot be exposed to these conditions as long as they are inside the aircraft. Instead the moisture absorption is accelerated. Basically two different methods are used: fluid immersion or environmental test chamber. Fluid immersion is much faster method than using a test chamber where relative humidity and temperature can be controlled. Figure 95 [10] shows a moisture absorption of RTM6 resin in fluid immersion. The moisture equilibrium is reached in 14 days.

Figure 95. RTM6 moisture absorption in fluid immersion.

Materials that have not been exposed to moist conditions are referred to as RT/AH 1 and exposed materials are referred to as Hot/Wet. In TANGO the structural items are designed with RT/AH design values and the final " weight penalty" is added based on the Hot/Wet material properties. The exposure to humid conditions in TANGO was agreed to be 1000 hours at 70 C / 85% or saturated. Saturated means in this context that moisture uptake that is monitored by weighting is lower for a given interval in the given standard. All Hot/Wet test specimens were exposed to humid conditions at least 1000h at 70 C / 85%. For the most specimens exposure was even longer. The exposure to humid conditions was monitored and conducted according to the standard EN 2823, Test method for the determination of the effect of exposure to humid atmosphere on physical and mechanical characteristics. This standard is referred in all of the applied test standards (Table 24) when test specimens are to be exposed to humid conditions except in the

1

Standard conditions are stated in most standards as 23 C / 50% of relative humidity.

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Development testing of a composite wing rib -

Design data generation

standard ASTM D5379M, which refers to the standard ASTM D5229. However, the standard ASTM D5229 is similar to EN 2823. The exposure of the test specimens to humid conditions was conducted at HUT/LLS in the Weiss Technik SB22/160 climate chamber (Figure 97). The moisture absorption of the test specimen was monitored using traveller specimens that accompany mechanical and physical test specimens. The dimensions of the traveller specimens were 25 x 25 mm2. Since those were extracted from the same laminates used for testing, their the nominal thickness was 2,68 mm. The accuracy requirement for weighting the travellers was 0.1 mg. The specimens were weighted at the HUT Laboratory of Machine Design with the Mettler AE200 precision balance (Figure 96). The surface finish can have a significant effect on the moisture absorption. The surface finish of the traveller specimens was the same as for the test specimens: peel ply finish, edges not sealed. The parts were free-standing in the chamber. Prior to exposure the traveller specimens were dried in order to establish the moisture absorption from zero point. The sequence for drying was as recommended in EN 2823: 72h at 50 C + 72h at 70 C + 5h at 90 C. However, change in the weight during the drying could not be detected.

Figure 96. Mettler AE200 precision balance at HUT/LMD.

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Development testing of a composite wing rib -

Design data generation

Figure 97. Weiss Technik SB22/160 climate test chamber.

The percentage of absorbed moisture is determined according to the following formula mi m0 100% m0 is the absorbed moisture (%) is the mass of the traveller specimen after drying is the mass of the traveller specimen after exposure

HL

(15)

where

HL m0 mi

The required time to achieve moisture equilibrium or constant mass is considered to be achieved when the difference between three successive weightings carried out at 168h interval is less than 0,5 mg for each gram of the traveller specimen. Resin mass fraction of the laminate is [6] (1

f f

wr

r r

r

r

f

)

r f f

(1

f

)

(16)

r

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Development testing of a composite wing rib -

Design data generation

If the resin only absorbs moisture, then the resin moisture absorption is mi m0 m r0 m0 (1 m0 (1 mi m0 100% m0 wr0

f r f

HL r mi

100%

f

)

f

(17)

r

)

100%

where mr0 wr0

is the initial weight of the resin (g) is the initial mass fraction of the resin

2,5

2,0

1,5 Laminate Uptake [%] Resin Uptake [%] 1,0

0,5

0,0 0 1000 2000 3000 Time [h] 4000 5000 6000

Figure 98. Moisture uptake of G0926 INJ and RTM6 (5000h period).

2,5

2,0

1,5 Laminate Uptake [%] Resin Uptake [%] 1,0

0,5

0,0 0 200 400 600 Time [h] 800 1000 1200

Figure 99. Moisture uptake of G0296 INJ and RTM6 (1000h period).

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Development testing of a composite wing rib -

Design data generation

6.13. Summary

The results are collected in Table 29 and normalised to 62% Vf (2.68 mm nominal thickness) in Table 30. B-basis values are calculated according to the method presented in Chapter 3.3. Only the strength values are downgraded, since downgrading elastic modulus properties would easily lead to unrealistic behaviour of the structure. The stress concentration factor of 1.44 was applied for the through thickness tensile test strength.

Table 29. Measured properties of G0926INJ and RTM6 (Non-normalised).

Property

Tensile strength Tensile modulus Tensile strength Tensile modulus Compressive strength Compressive modulus* Compressive strength Compressive modulus In-plane shear strength In-plane shear strength In-plane shear modulus In-plane shear modulus Through thickness tensile strength Through thickness tensile modulus Through thickness tensile strength Through thickness tensile modulus Bolt bearing strength Bolt bearing strength ILSS ILSS Open Hole Tension Open Hole Tension Open Hole Compression Open Hole Compression

Standard

AITM 1.0007 AITM 1.0007 AITM 1.0007 AITM 1.0007 AITM 1.0008 AITM 1.0008 AITM 1.0008 AITM 1.0008 ASTM D5379 ASTM D5379 ASTM D5379 ASTM D5379 NPL Draft v.02 NPL Draft v.02 NPL Draft v.02 NPL Draft v.02 AITM 1.0009 AITM 1.0009 BS ISO EN 14130 AITM 1.0007 AITM 1.0007 AITM 1.0008 AITM 1.0008

Cond.

RT/AH RT/AH Hot/Wet Hot/Wet RT/AH RT/AH Hot/Wet Hot/Wet RT/AH Hot/Wet RT/AH Hot/Wet RT/AH RT/AH Hot/Wet Hot/Wet RT/AH Hot/Wet RT/AH RT/AH Hot/Wet RT/AH Hot/Wet

Mean Std.dev. B-basis (%) value

664.0 55.0 552.3 53.7 394.0 55.5 359.7 51.9 228.2 170.0 22.3 19.0 28.7 8.5 18.8 353.9 328.4 56.0 42.1 423.0 397.6 292.0 202.2 1.71 3.68 1.90 2.78 8.04 4.38 5.06 4.91 7.81 7.88 6.14 7.07 11.47 9.00 32.91 8.78 10.44 3.96 3.05 3.62 2.54 6.54 5.67 564.4 469.5 334.9 287.8 182.6 136.0 22.9 15.0 283.1 279.1 44.8 33.6 359.6 338.0 248.2 171.9

BS ISO EN 14130 Hot/Wet

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Design data generation

Table 30. Measured properties of G0926INJ and RTM6 (Normalised to 62% Vf).

Property

Tensile strength Tensile modulus Tensile strength Tensile modulus Compressive strength Compressive modulus* Compressive strength Compressive modulus In-plane shear strength In-plane shear strength In-plane shear modulus In-plane shear modulus Through thickness tensile strength Through thickness tensile modulus Through thickness tensile strength Through thickness tensile modulus Bolt bearing strength Bolt bearing strength ILSS ILSS Open Hole Tension Open Hole Tension Open Hole Compression Open Hole Compression

Standard

AITM 1.0007 AITM 1.0007 AITM 1.0007 AITM 1.0007 AITM 1.0008 AITM 1.0008 AITM 1.0008 AITM 1.0008 ASTM D5379 ASTM D5379 ASTM D5379 ASTM D5379 NPL Draft v.02 NPL Draft v.02 NPL Draft v.02 NPL Draft v.02 AITM 1.0009 AITM 1.0009 BS ISO EN 14130 AITM 1.0007 AITM 1.0007 AITM 1.0008 AITM 1.0008

Cond.

RT/AH RT/AH Hot/Wet Hot/Wet RT/AH RT/AH Hot/Wet Hot/Wet RT/AH Hot/Wet RT/AH Hot/Wet RT/AH RT/AH Hot/Wet Hot/Wet RT/AH Hot/Wet RT/AH RT/AH Hot/Wet RT/AH Hot/Wet

Mean Std.dev. B-basis (%) value

611.5 50.9 533.6 51.9 368.65 353.9 48.9 219.7 21.5 41.3 8.5 35.3 348.9 337.8 55.6 42.3 399.09 407.0 275.86 197.2 1.71 3.68 1.90 2.78 8.04 5.06 4.91 7.81 6.14 11.47 9.00 32.91 8.78 10.44 3.96 3.05 3.62 2.54 6.54 5.67 519.8 453.6 294.9 283.1 175.8 33.0 28.2 279.1 270.2 47.3 36.0 339.2 346.0 220.7 157.8

BS ISO EN 14130 Hot/Wet

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Development testing of a composite wing rib - Loading rig for multiaxial testing

7. Loading rig for multiaxial testing

7.1. Introduction

Multiaxial testing of laminates as well as aerospace structures has been under investigation for a longer period of time in the aerospace industry. At the beginning of the test program planning phase quite a lot of consideration was given for combined loading test set-up and how to arrange combined loading case for the wing ribs. In this context combined load case is considered as simultaneous compression and shear or tension and shear load. A wing rib is loaded primarily by shear force (torque) and brazier/crushing load (wing bending) in normal flight as described in Chapter 2.1.

7.2.

Requirements for the test fixture

The primary function of the test fixture was to generate simultaneous compression and shear loading on a wing rib panel. Maximum wing rib test panel dimensions, based on TANGO rib9 dimensions, were: 600 mm length (chordwise) 400 mm height The ratio of the compression/tension and shear load was to be controllable. Hence, two independent actuators were planned to be used for loading. Maximum actuator loads for applying compression and shear loads were 300kN due to the limitations at the Laboratory for Mechanics of Materials where the rib panels were planned to be tested. The stability of the specimen was to be ensured in a way that the natural behaviour of the specimen was to be simulated as closely as possible.

7.2.1. Highlighting the problems of the multiaxial loading

The problems of multiaxial loading of the representative wing rib panel focuses on the free edges of the test panels. Loading a test panel with compression and shear loads creates unsymmetrical stresses in the test panel due to the free edges. This behaviour is demonstrated by the following FEM analyses. An FE model of a test panel was created using I-DEAS and PATRAN/NASTRAN software. The test panel in this case also incorporates a vertical stiffener. The test panel is presented in Figure 100. Material properties used in analyses are presented for the CFRP laminate in Table 31. The test panel was loaded using representative loads from A310 aircraft: compressive load 23.4 kN applied to vertical stiffener, shear load 26.5 kN evenly distributed on the upper flange. The test panel is attached by restraining all z-direction

- 99 -

Development testing of a composite wing rib - Loading rig for multiaxial testing displacements at the upper flange (not shown in Figure 101) and all DOFs at lower flange. Test panel loading and boundary conditions are presented in Figure 101. Deformed test panel is presented in Figure 102. Unsupported test panel behaviour under stresses is highly unsymmetrical due to the free edges. Vertical y-stresses of the loaded test panel are presented in Figure 103. The results show that one free edge is under compression and the other free edge is under tension. Also shear stresses in the panel were highly unsymmetrical. Free edges also lead to stability problems. The first buckling mode is confronted much before the maximum stresses for the CFRP material are reached. The problem concentrates on the fact that wing ribs are designed and sized by examining wing rib panels separated by vertical and horizontal stiffeners inside the whole wing rib. Testing these separated panels is difficult due to the free edges of the test panel that do not exist inside the real structure. Thus, the actual mission is then to design a fixture that is able to transfer the representative loads to the test panel, while compensating the free edge effect.

Figure 100. Free edge test panel. Height 310 mm, width 300, red area 3mm thick, blue area 5 mm thick.

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Figure 101. Free edge test panel loads and boundary conditions.

Figure 102. Deformations of the free edge test panel.

Figure 103. Vertical Y-stresses in the free edge test panel.

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Figure 104. First buckling mode of the free edge test panel.

7.3.

Test fixture

A 3D model of the loading rig is presented in Figure 105. The loading rig was modelled using IDEAS software. The test rig consists of two support plates, alignment frame and two load input brackets. The test panel is also modelled in the picture. The functional concept of the rig is to achieve symmetrical shear flow and compression around the test panel. The support plates create an evenly distributed compression to the test panel when load is applied to the vertical input bracket and shear flow is created by applying load to the horizontal input bracket. During the loading the alignment frame holds the two support plates aligned horizontally to each other. By keeping the support plates aligned an evenly distributed shear flow and compression is generated. The stability of the horizontal strut is ensured by arranging the alignment frame and horizontal load bracket as in Figure 105. The horizontal struts are under tension stress instead of compression. All components except the alignment frame are designed of steel (Formax material properties). The alignment frame is made of aluminium (2024-T6). The test panel is modelled as quasi-isotropic CFRP laminate.

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Compressive load input

Support plate

Shear load input

Support plate Test panel

Alignment frame / Horizontal struts

Figure 105. The test fixture for multiaxial loading of wing rib panels. The support plate thicknesses are not to scale.

7.4.

Analyses

The analyses of the test fixture were conducted at the PFC stress department. An IGES file from the 3D model was transferred to PATRAN/NASTRAN software in order to execute FEM analyses for the test fixture. Both the fixture and the test panel in the fixture were under investigation. The purpose of the FEM analysis was to demonstrate the structural integrity of the test fixture and to expose the behaviour of the loaded test panel. The alignment frame of the test fixture was meshed using 3D solid elements, horizontal struts using beam elements and the rest of the components using shell elements. The meshed 3D model is presented in Figure 106. Different components are shown in different colours. Loads were applied through the hole edges of the load input brackets. The lower support plate was to be bolted to the massive base set at the Laboratory for Mechanics of Materials. The FEM model is attached to the ground by similar manner by setting the boundary conditions as all DOF restrained at location of the bolt hole node. The boundary conditions and loads are presented in Figure 107. Material properties used in analyses are presented in Table 31. A deformed model under loads in Figure 108 highlights the behaviour of the loaded test fixture. First analyses showed quite high deformations in the upper support plate; hence adding flange plates and increasing the thickness of the plate further stiffened it. These

- 103 -

Development testing of a composite wing rib - Loading rig for multiaxial testing modifications however are not shown in the figures. Deformations of the modified test fixture were stated to be acceptable. The vertical and shear stresses of the test panel inside the test fixture are presented in Figure 109 and Figure 110.

Figure 106. Meshed 3D model of the test fixture.

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Figure 107. Boundary conditions and loads.

Table 31. Material properties used in test fixture FEM analyses.

MATERIAL 2024-T6 [31] Formax [32] CFRP laminate

E11 (GPa) 72.4 210.0 39.56

E 22 (GPa) 72.4 210.0 39.56

G12 (GPa) 28.0 79.0 15.35

12

0.33 0.30 0.29

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Figure 108. Deformed model of the test fixture. Deformations are exaggerated to reveal behaviour.

Figure 109. Vertical Y-stresses in the test panel.

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Development testing of a composite wing rib - Loading rig for multiaxial testing

Figure 110. XY-stresses in the test panel.

7.5.

Ambivalent factors

Even if the stress distribution is satisfactory, the free edges still cause problems: the stability issue, the uneven load distribution at flanges and applying the compression load. Stability of the panel with free edges must be ensured. Stability can be accomplished by material addition, modifying geometry or external support at free edges. All these alternatives lead to changes in the results and compensations in post-processing. Uneven load distribution at the flanges can not be avoided as long as the test panel edges are free and flanges transfer shear loads. This can be demonstrated simply through equilibrium equations in Figure 111a. If we calculate moment around point A, to achieve an equilibrium we end up having load distribution on the flanges as in Figure 111b. Similar vertical stress distribution is shown also in Figure 109.

a.

b.

A

Figure 111. Equilibrium of the test panel.

- 107 -

Development testing of a composite wing rib - Loading rig for multiaxial testing Compression loads on ribs in the true wing are generated through wing bending and occur as brazier loads. These loads are transferred mainly through stringers since the spanwise stiffness of the wing skin is much higher at stringer locations. Further, to fully utilise the benefit of vertical stiffeners in the ribs should those be vertically located at stringer locations. Hence, it was agreed that it should be possible to input compressive loads to the test panels through vertical stiffeners as it is done in Figure 101. This leads to another dilemma. How to pass loads only locally through support plate when it is at the same time stiff enough to keep the flanges horizontally aligned.

7.6.

Conclusion

Due to the following facts it was decided not to start assembling the multiaxial testing fixture: Financial issues and lack of time The critical load cases showed to be an in-flight fuel pressure case, not the in-flight shear, moment and torque (SMT) case when compression and shear acts on the wing rib section simultaneously. The ambivalent factors described in the previous chapter Multiple different solutions to remove the uncertainty around the test set-up were laid out. Consulting with colleagues from CRC-ACS within TANGO revealed that this similar situation had been under investigation for years, but ran up against similar problems. CRC-ACS solution to the problem was to test the full-scale rib. The test set-up is presented in Figure 112 and was made during a DaimlerChrysler Aerospace Airbus GmbH (DA) program related to similar area as TANGO.

Figure 112. CRC-ACS rib test during a DASA program [ 33].

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Development testing of a composite wing rib - Non-destructive inspection of sinewave web

8. Non-destructive inspection of sine-wave web

8.1. Introduction

Non-destructive inspection is one of the most time consuming and costly phase when manufacturing aircraft parts. Especially, when NDI is conducted manually. Besides the literature references this was noticed during the first cost analysis for PFC TANGO wing ribs. Since cost effectiveness is one of the primary goals in TANGO, mechanical NDI is preferred over manual NDI. Additionally, manual inspection does not file any records of the inspection while mechanical inspection outputs record in useful form e.g. C-Scan or attenuation record, thus traceability is better when conducting NDI mechanically. One of the first goals during the evaluation and concept design phase was to investigate which kind of problems are related to NDI of the whole wing rib. The purpose of the NDI study concentrated on the problems with web area of the rib: How to inspect the web mechanically if the web incorporates various out of plane features such as sinusoidal, dimpled or corrugated web.

8.2.

Problem areas

NDI was naturally restricted to the available resources at PFC. It was agreed that ultrasonic water immersion was the best way to inspect the ribs. The equipment used for inspection is described in Chapter 5.2. It consists of a through transmission system that sets restrictions for scanning. Transmitter/receiver unit cannot follow the outlines of the inspected part, instead the scanning track is linear. Problems exist when the angle between the inspected part and the scanning track (Figure 113) is too high. At wide enough angles the attenuation value is not readable. The maximum inspectable angle was established by using 3 mm thick aluminium plates, which were bent to incorporate angles as in Figure 113. Only one angle was used in single plate. The nominal angle definition is presented in Figure 114. The nominal angle values of the aluminium plates are presented in Table 32. Four plates attached to the scanning rig (Figure 115) were scanned at a same time.

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Development testing of a composite wing rib - Non-destructive inspection of sinewave web

Transmitter Rib web

Receiver Scanning track

Figure 113. Definition of the angle between scanning track and rib web.

Figure 114. Nominal angle definition for the aluminium plates.

Figure 115. Test rig set 2 attached to the scanning rig.

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Development testing of a composite wing rib - Non-destructive inspection of sinewave web

Table 32. Nominal angles of the aluminium plate. Values are in degrees.

Test rig set 1. Test rig set 2.

60 65 70 75 75 77.5 82.5 85

8.3.

Scanning results

The panels were scanned in two sets. Results are presented in Figure 116 and Figure 117. The results of set 1 show unacceptable attenuation values that are shown dark blue / black in Figure 116. Acceptable results are attained from the first two panels from the left in the set 2 (length 0 - 600 mm).

Figure 116. C-scan results of the test rig set 1.

Figure 117. C-scan results of the test rig set 2.

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Development testing of a composite wing rib - Non-destructive inspection of sinewave web

8.4.

Conclusions

According to the scanning results the readable scan data can be extracted from the panel 2 (Figure 117, at length 0 ­ 300 mm) in test rig set 2 and possibly from panel 1 from left (Figure 117, at length 300 ­600 mm). The nominal angles of these panels are 77.5 and 80 degrees. Relative angles between the scanning track and part can be calculated with the following equation 180 2

(18)

The actual angles from the panels near the readable transition area were measured using a manual bevel. The results of the measurements are presented in Table 33. By taking the mean values of the angles from Table 33 of panels 77,5 and 80 and substituting into equation above, we get a boundary value for the maximum inspectable angle 180 2 180 2 77.2 73.7

Panel77.5

53.2

Panel80

51.4

Table 33. The actual angles in the test panels, values are in degrees.

Test panel Angle # 1 2 3 4 5 6 7 8 9 10 11 12 Mean Dev. Max. Min. Beta 71.5 71.5 71.0 71.5 71.0 71.5 69.5 70.0 70.5 70.5 70.5 69.5 70.7 0.75 71.5 69.5 54.6 73.5 73.0 74.0 73.0 74.0 74.0 73.5 73.5 74.0 74.0 74.0 76.0 79.0 76.0 77.0 75.0 78.0 77.0 77.5 78.0 77.5 78.0 78.0 79.0 79.5 79.5 80.0 79.0 78.5 79.0 79.0 79.5 75 77.5 80 82.5

73.7 0.40 74.0 73.0 53.2

77.2 1.15 79.0 75.0 51.4

79.1 0.57 80.0 78.0 50.5

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Development testing of a composite wing rib - Non-destructive inspection of sinewave web According to the results the maximum inspectable angle by NDI between the scanning track and the rib web is 51 degrees. However, it must be noted that this value applies only for 3 mm thick and thinner specimens. If the inspected part is thicker than 3 mm the inspectable angle is most likely smaller and more tests should be conducted to verify the inspectable thickness.

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Development testing of a composite wing rib -

Other upcoming tests

9. Other upcoming tests

The tests presented in this thesis are only a part of the whole wing rib development testing program. More tests will be conducted as it appears in Chapter 3.3. The upcoming tests include L-pull tests, full-scale panel tests and a full-scale wingbox pressure test. A short description of these tests is presented ahead below.

9.1.

L-Pull test

The pressure inside the lateral wingbox () creates pulling forces to the wing rib feet. This pulling force has an opening effect to the rib foot corner. The purpose of this test is to determine the strength of the rib foot corner. The term L-pull test comes out from the shape of the test specimen. L-Pull test is a test of a C-shaped rib flange/web corner. An L-shaped part of the wing rib is bolted from the flange to a wing skin panel and pulled from the web. The wing skin panel is supported by two rollers allowing free the rotation of the skin panel under the rollers. The L-pull specimen will be damaged from the corner area.

Figure 118. L-pull test set-up. Nominal 4.0 mm skin and 2.68 mm specimen.

During the L-pull test program at least four thicknesses (2.68, 5.36, 8.05 and 10,73 mm nominal thickness) are tested with two different corner inner radius (5.0 and 10.0 mm nominal inner radius). These specimens with different dimensions are tested respectively with a correct wing skin thickness. The corresponding wing skin panel thickness is attained from the lateral wingbox wing skin panel manufacturer, which is ALENIA in this case. A special RTM steel mould is milled to manufacture the L-pull specimens. The

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Development testing of a composite wing rib -

Other upcoming tests

specimens will incorporate the same Vf and lay-up as the current design (62%, 25%/50%/25%).

9.2.

Full scale rib section test

The design of the TANGO wing ribs 9,11 and 12 incorporates a so-called Sigma design concept. In the Sigma concept the cross-section of the wing rib (drafted in Figure 119) has a shape of the Creek letter Sigma. This kind of design significantly reduces the bending of the rib web under compressive loads.

Figure 119. A cross-section of the Sigma design rib.

The Brazier effect creates compressive loads for the rib web. However, the web buckling or the compressive strains are not critical according to the current stress analyses. The purpose of this test is to verify the stability of the Sigma concept and assess the accuracy of the FE buckling analysis for the Sigma concept. The Sigma concept rib panel is manufactured by VARTM with an aluminium mould. The flange of the test panel is not curved as the actual rib shape instead it is straight for facilitating the test set-up. The injected and untrimmed test panel is shown in Figure 120. The test panel is bolted to the wing skin panel. One stringer is co-bonded to the skin panel. The wing skin panel and stringers are manufactured of AS4/3506-1 prepreg material in the RTM press. The wing skin panel incorporates the same properties as the current TANGO wing skin design. The wing skin panel is supported by two rollers allowing the panel to rotate freely on the rollers. The test panel has also small flanges at the free edge. An aluminium strut is constructed between the flanges to hold the alignment of the two flanges parallel. This will reduce the influence of the free edges to the buckling mode and load. The compressive load to the test panel is applied through the whole width of the test panel. The movement of the skin and the test panel is monitored with displacement transducers during the test. The test set-up is presented in Figure 121.

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Development testing of a composite wing rib -

Other upcoming tests

Figure 120. Injected Sigma panel.

Figure 121. The Sigma concept test set-up.

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Development testing of a composite wing rib -

Other upcoming tests

9.3.

Full scale wingbox pressure test

A high pressure inside the lateral wingbox due to the in-flight tuned up-gust case combined with in-flight SMT loads generates the critical loading for the TANGO ribs 11, 12 and 14. To verify the whole integrity of the wing rib under high pressure loads a full scale pressure test is conducted. Also the load distribution for the rib feet (bolt loads) is examined. In the full-scale pressure test a part of the lateral wingbox is pressurised with water. The test wingbox is composed of 3 wing ribs, front/rear spars, 3 fuel system pipes and upper/lower cover. The wingbox is presented in Figure 122. The skins and spars are aluminium. The ribs are manufactured by RTM with TANGO rib 14 trial mould. The Rib14 trial mould was manufactured in order to verify the manufacturing process and to study the manufacturing process of the actual full-scale wing rib. These trial ribs were used in the pressure test. The Rib 14 trial ribs incorporate the same properties as the current design. The ribs located at the wingbox ends are also strengthened, since only the rib in the middle of the wingbox is under investigation at the pressure test. Two types of tests are to be conducted: Unsymmetrical case and symmetrical case. At the unsymmetrical case only the inner rib bay is pressurised. This test simulates the conditions for the rib14 pressure case, since the rib 14 is a sealed fuel tank end rib. At the symmetrical test the both rib bays are pressurised. This test simulates the conditions for the rib 11 or 12. In the both tests the lower skin displacements are monitored. Also, the middle rib is instrumented with strain gages. Strains, displacements and the pressure inside the wingbox are recorded by one data logging system. Thus, strains and displacements can be plotted as a function of the pressure.

Figure 122. Assembled full-scale pressure test wingbox.

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Development testing of a composite wing rib -

Summary

10.

Summary

The thesis describes how the testing of a composite wing rib in the development phase of the rib is approached. Loading of a wing rib is studied. The study clarifies how a composite wing rib is loaded. The study also establish a basis for the upcoming coupon and element tests in terms of what kind of material properties and more detailed information are needed for the design of composite wing ribs. Mechanical properties for different materials were measured in the material screening. The most suitable material combination was selected based on the mechanical performance and processability. More extensive measurements were conducted for the selected wing in order to establish design data values. These properties were measured both in RT/AH and hot/wet conditions. Different chemical tests were conducted in order to verify the quality of the test laminates. Compressive tests for reinforcement stacks were conducted in order to assess the mould closing forces of the press system in the RTM process. Mechanical NDI of complex shapes was also investigated. In this case an ultrasonic C-Scan of a sinusoidal wing rib web. Multiaxial testing of wing rib sections was studied and a test fixture was designed to produce multiaxial loading for the wing rib sections. However, this test fixture was not manufactured due to the fact that critical load cases showed to be an inflight fuel pressure case, not the in-flight shear, moment and torque (SMT) case when compression and shear acts on the wing rib section simultaneously. This thesis was finished nowhere near within the original planned timeframe. The definition of the thesis at the beginning was somewhat open, since no clear estimation could be made what was to be confronted in the field of testing in the whole TANGO project. The first document of lateral wingbox testing philosophy or approach was issued only after 8 months of the beginning of the thesis work. Originally coupon testing was intended to be minor, but coupon testing could not be avoided. The emphasis was intended to be on element and detail level testing. However, during first year of the programs there simply was not enough data to conduct element level tests. Materials, loading and master geometry was still under design and consideration. Part of the element and detail tests, e.g. L-Pull and panel compression tests, conducted was also intended to be included in this thesis, but due to the limited resources those needed to be bounded off. A short descriptions of the upcoming tests are presented. More detailed analysis of the test results could also be presented, but in terms of required design data for TANGO all significant factors are considered to be presented. This applies also for the wing rib loading examination. Analytical hand calculation stressing methods could be presented, but would not match with the stressing methods used at PFC during TANGO.

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Development testing of a composite wing rib -

References

11.

[1]

References

A PROPOSAL FOR FUNDING FROM THE CEC UNDER THE FIFTH FRAMEWORK PROGRAMME, PART B, June 1999. Niu, Michael C. Y. Airframe Structural Design, second printing. Conmilit Press LTD. 1989. Raymer, Daniel P. Aircraft Design: A Conceptual Approach. American Institute of Aeronautics and Astronautics, Inc. 1989. Roskam, Jan. Aircraft Aerodynamics & Performance. DAR Corporation. 1997. http://www.jaa.nl/jar/jar/jar/jar.25.561.htm. Airasmaa, I. & Kokko, J. & Komppa, V. & Saarela, O. Muovikomposiitit. Muoviyhdistys ry. Helsinki. 1991. Kruckenberg T. & Paton R. Resin Transfer Moulding for Aerospace Structures. 1998. Chapman & Hall. London. ISBN 0412731509. ADVANCED COMPOSITE MATERIALS: New developments Applications. 1991. ASM International. USA. ISBN 0-87170-431-5. and

[2]

[3]

[4] [5] [6]

[7]

[8]

[9]

Potter, K. Resin Transfer Moulding. Chapman & Hall. London. 1997. ISBN 0412725703. RTM 6 Product Data. Hexcel Composites. Publication ITA 065a. 1998. 6 p. PR 500 Product Data. 3M Advanced Technologies for Aerospace Structures. 3M Adhesives Division. USA. 1999. 48 p. Cycom 875RTM Product Data. Cytec Fiberite, Issue ­August 1999. 7 p. ENGINEERED MATERIALS HANDBOOK; Volume 1. Composites. ASM International. USA. 1987. ISBN 0-87170-279-7 (v. 1). 983 p. Holmberg, Anders J. Resin Transfer Moulded Composite Materials, Doctoral Thesis. Department of Materials and Manufacturing Engineering. Sweden. April 1997. ISNN: 1402-1544. Mechanical Engineering. The American Society of Mechanical Engineers. 1997. USA.

[10] [11]

[12] [13]

[14]

[15]

- 119 -

Development testing of a composite wing rib [16] [17] MIL-HDBK-17-1E, Volume 1. Chapter 8. 1997.

References

Krautkrämer, J. Ultrasonic Testing of Materials, 4th fully revised Edition. Springer-Verlag, 1990. ISBN 3-540-51231-4. 677 p. Non-destructive system for advanced materials. Nukem GmbH. 1991. 3176-AM0900.4. 22 p. Teagle, P. R. The quality control and non-destructive evaluation of composite aerospace components. Composites, Volume. 14, No. 2. 1983. AITM 1.0007. Determination of notched and unnotched tensile strength. Airbus Industrie. France. 1994. 9 p. FM 300-2 Film Adhesive product description. Cytec Fiberite Ltd. 1990. 6 p. ABT 1.0004, Issue 1. Determination of compressive strength and modulus of carbon, glass and aramid fibre composites. British Aerospace Airbus Test specification. UK. 1993. 5 p. STM D3410-75, Compressive properties of unidirectional or crossply fiber-resin composites. The American Society for Testing of Materials. USA. 1982. 9 p. BS ISO EN 14130:1998. Fibre-reinforced plastic composites ­ Determination of apparent interlaminar shear strength by short-beam method. British Standard. 1998. 17 p. BS 2782:Part 3:Method 341A:1977. Determination of apparent interlaminar shear strength of reinforced plastics. British Standards Institution. 3 p. Griffiths, G. R. & McGarthy, R. F. Making the difference with innovative design and processes. Sampe Europe. 2001. ISBN 3-9520477-8-3. 691 p. Pendleton, Richard L. Manual on Experimental Methods for Mechanical Testing of Composites. Elsevier Applied Science Publishers. U.K. 1989. ISBN 1-851663754. 169 p. AITM 1.0008. Determination of notched, unnotched and filled hole compression strength. Airbus Industries. France. 1994. 11 p. STM 5379M-98. Standard Test Method for Shear Properties of Composite Materials by the V-Notched Beam Method. The American Society for Testing of Materials. USA. 1999. 13 p.

[18]

[19]

[20]

[21] [22]

[23]

[24]

[25]

[26]

[27]

[28]

[29]

- 120 -

Development testing of a composite wing rib [30]

References

AITM 1.0009. Determination of bearing strength. Airbus Industries. France. 1994. 12 p. Davis, J. R. Aluminium and Aluminium Alloys. ASM International. USA. 1993. ISBN: 0-87170-496-X. 781 p. Uddeholm product catalogue. Oy Uddelholm Ab. 2000. Discussion with Rob Ness from CRC-ACS, Australia, http://www.crcacs.com.au.

[31]

[32] [33]

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Development testing of a composite wing rib -

Appendixes

12.

Appendixes

APPENDIX 1: C-SCAN OF THE G0926-B-T-2 LAMINATE .................................................123 APPENDIX 2: C-SCANS OF THE G0926-C-T-3 AND DB-C-T-1 LAMINATES ... ... ...........124 APPENDIX 3: DETERMINATION OF FIBRE VOLUME FRACTION BY ACID DIGESTION .......125 APPENDIX 4: TG TMA MEASUREMENT RECORD ... ... ... ... ... ... ... ... .............. ... ... ...127

- 122 -

Appendix 1: C-Scan of the G0926-B-T-2 laminate.

- 123 -

Appendix 2: C-Scans of the G0926-C-T-3 and DB-C-T-1 laminates.

- 124 -

Appendix 3: Determination of Fibre volume fraction by acid digestion.

- 125 -

Appendix 3: Determination of Fibre volume fraction by acid digestion.

- 126 -

Appendix 4: Tg TMA measurement record

- 127 -

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