my composite design capability maintenance studies

AIAA Mr. Geoffrey Allen Wardle MSc. MSc Composites Design Capability Maintenance Examples 2012-2015.

MY COMPOSITE DESIGN CAPABILITY MAINTENANCE STUDIES.

By Mr. Geoffrey Allen Wardle MSc. MSc. 2012 to Date.


This is presentation gives examples of composite airframe design work I have undertaken on my

own initiative to maintain my capabilities with the Catia V5.R20 toolset in addition to Workbooks 1

and 2, and my current FATA design study.

The objectives of this capability maintenance work is to preserve my capabilities within the Catia

V5.R20 toolset against future employment and in support of the Future Advanced Technology

Aircraft private research project. As such this work is divided into three areas:-

The first covers baseline capability exercises and lays out the toolset methods:

The second covers the design standards applied in the development of composite parts for the

FATA project and encompass my experience in composite design within BAE Systems MA&I

and my Cranfield University MSc in Aircraft Engineering as well as my University of Portsmouth

MSc in Advanced Manufacturing Technology:

The third covers the application of the Composite Engineering Design (CPE), and Composite

Design for Manufacture (CPM) modules within Catia V5.R20, covering a build up of exercises

and self created examples, such as the outboard leading edge wing spar for baseline FATA

aircraft wing structure, a wing tip fence spar from a Boeing 767 study, and vertical tail skin.

This study will grow over time as more detail structural work is undertaken on the FATA project and

it is intended to add PATRAN / NASTRAN FEA modeling of FATA wing components as they are

evolved to the preliminary design stage. On a month by month basis this will reflect development

progress and is to be taken as an indicator of capabilities and a knowledge base which is applicable

to a range of aerospace industry challenges. The (In Work) designations are sections currently

being completed. 2

OBJECTIVES OF THIS PRIVATE STUDY IN SUPPORT OF FATA.


Section 1:- Basic Catia V5.R20 CPE capability maintenance exercises:

Section 2:- Design rules applied to main design exercises:

Section 3:- Composite component materials and processing overview:

Section 4:- CFRP Post layup conversion processing tooling:

Section 5:- Assembly design and corrosion prevention:

Section 6:- Environmental protection of composite airframe structures:

Section 7:- Composite structural testing and Qualification:

Section 8:- Designing component parts: (1) Spar design : (2) Skin design (In Work):

Section 9:- Catia V5.R20 Solid part extraction for mock up and assembly evaluation (In Work):

Section 10:- Catia V5.R20 Flat pattern and manufacturing data extraction for production (In Work):

Section 11:- Drawing representation by 2-D extraction and annotation (In Work):

Section 12:- FEA structural analysis of the as designed composite components (In Work).

3

Contents of this presentation in support of my FATA study.


The objective of this self study is to develop and enhance the skills set in the application of

the Catia V5 R20 Composite Engineering Design (CPE), and Composite Design for

Manufacture (CPM), post Cranfield MSc and BAE SYSTEMS composite design training

modules.

The required more than 500 hours Catia V5 experience level for these exercises, has been

greatly exceeded by myself with more than 16,800 hours.

The preliminary exercises undertaken used the ABD Matrix tutorials CT1 Basic Composite

Laminate Design: CT2 Working With Transition Zones: and CT3 Creating Limit Contours,

subsequent study used the Wichita State University CATIA Composites text as a guide for

further exercises, as well as the CPDUG Tutorial, the final exercises being the designs for a

military fighter and a commercial airliner vertical tail spar and a multi island vertical tail skin

panel.

At the time of conducting, and creating these study exercises BAE SYSTEMS had no CATIA

V5 composite design methodology choosing to use Catia V5 as a like for like replacement

for legacy Catia V4, therefore this work promotes the application of this powerful toolset to

Cranfield University and the AIAA, and feeds into the FATA future commercial aircraft wing

study.

Section 1:- Basic Catia V5.R20 CPE capability maintenance exercises.

4


CATIA V5.R20 Composite design toolset.

There are two composite design products within Catia V5 Composite Work Bench which are

Composites Engineering Design (CPE) and Composites Design for Manufacturing (CPM) and

these are outlined below.

The Composites Engineering Design (CPE) product provides orientated tools dedicated to

the design of composite parts from preliminary to engineering detailed design. Automatic ply

generation, exact solid generation, analysis tools such as fiber behavior simulation and

inspection capabilities are some essential components of this product. Enabling users to

embed manufacturing constraints earlier in the conceptual design stage, this product shortens

the design-to-manufacture period.

The Composites Design for Manufacturing (CPM) product provides process orientated

tools dedicated to manufacturing preparation of composite parts. With the powerful

synchronization capabilities, CPM is the essential link between engineering design and

physical manufacturing, allowing suppliers to closely collaborate with their OEM‟s in the

composite design process. With CPM, manufacturing engineers can include all manufacturing

and producibility constraints in the composites design process.

5


Learning outcomes:-

From this study I am able to create a simple composite laminate using the Catia V5.R20

Composite Engineering Design module.

From this I am now able to gather important engineering information from the model using

the Numeric Analysis function.

Methodology:-

A reference surface 10 X 10 inches was constructed with four curves and a fill surface in

surface design before entering Mechanical Design – Composite Design.

The composite parameters selected were the default 0:45:-45:90 although the Composite

Parameters screen gives the option of adding, removing, or redefining ply angles. The

material was selected from the materials catalogue as Glass, (Insert – Parameters –

Composite Parameters).

Next the Zone Group Definition menu was accessed using Insert – Preliminary Design –

Zones Group. The default name was used for this example. The reference surface created

earlier was selected to define the Zone group geometry, and the default draping direction

was accepted. The Rosette Definition was achieved by selecting the Absolute Axis System,

and the Rosette Transfer type was set to Cartesian.

CT1:- INTRODUCTION TO COMPOSITE DESIGN.

6


Zone Geometry and Laminate Definition was accomplished the command sequence :-

Insert – Preliminary Design – Create Zone. The Zone Geometry was inputted by selection of

the four boundary curves used to produce the reference surface in ascending sequence 1

through 4. The Laminate Definition was produced using the laminate tab in the Zone

Definition menu, assigning the material (GLASS) from the catalogue and defining the

number of per angle. Figure 1 shows how the maturation of the model incorporates the

Zone Geometry and Laminate Definition.

The next stage was to create the first laminate of 8 plies orientated using the definition

inputted above. To create plies from the zone the following command sequence was used:

Insert – Plies – Plies Creation from Zones. In the Plies Creation window Zone Group 1 was

highlighted and Create plies in new group was selected. Create plies without staggering was

deselected, then OK was selected. This created Plies Group 1 as shown in figure 2

consisting of 8 sequences, one of which is exploded in the tree, also a new geometrical set

was created containing the curves to build each ply in the sequences.

The final stage in creating the build part shown in figure 3 was to apply the Ply Exploder to

show the 3-D stack-up as a 3-D model, enhancing the visual perspective of the Laminate,

allowing the engineer to check the integrity of the virtual component definition. The following

command sequence was used: Insert – Plies – Ply Exploder, and in the Exploder window

the default settings were used checking that Cumulative as per Stacking and Shell Constant

Offset were selected and the scale was set to 20, then OK was selected.

CT1:- INTRODUCTION TO COMPOSITE DESIGN (Cont).

7


Fig 1:- CT1:- Laminate Definition Model Tree Maturation.

This is how the model tree appeared after Zone Geometry

and Laminate Definition see also figure 3 fully matured model

tree.

Laminate definition appears in the

tree when Zone is defined.

These are the results of the laminate

definition data inputs.

8


Fig 2:- CT1:- Plies from Zone Model Tree Maturation.

Using Sequence 1 as an example the way in which Catia

constructs composite parts is revealed.

In this case, Ply 1 is made from glass, has a zero – degree

orientation and is defined geometrically by Contour 7: which is a

derivative of the previously defined Contour 8

The subsequent Sequences shown are built in the same way.

The newly created Geometrical Set 2 holds the 8 curves

needed to build each ply in the sequences. They are

created automatically during the ply creation stage.

9


Fig 3:- CT1:-Introduction to Composite Design completed part build and model tree.

10


CT1:- INTRODUCTION TO COMPOSITE DESIGN (Cont).

The final composite part build is shown in figure 3 with fully matured model tree.

Figure 4 (a) shows the part build with dimensions, and figure 4 (b) shows the ply schematic.

The ply schematic shows the laminate stack in 3-D, and the colors clearly show the varying

angles of each ply in the Laminate as shown in Detail A.

Further engineering design information was obtained from this using the Numerical Analysis

tool, to extract such information as:- ply surface areas: ply or laminate weights: volumetric

mass and much more as an Excel spreadsheet which is shown below as Table 1.

The Numerical Analysis tool is accessed through the Command Sequence:- Insert –

Analysis – Numerical Analysis, and with this tool either a single ply or a complete Composite

Laminate can be investigated.

To determine the Aerial mass of Ply 1 for example entre the Numerical Analysis tool and

select Ply 1 from the model tree as shown in figure 5, the Numerical Analysis dialog box will

update with the analysis parameters for the selected Ply 1, which gave the value as 0.043 lb.

To determine the Aerial mass of the Composite Laminate for example entre the Numerical

Analysis tool and select Plies Group 1 from the model tree as shown in figure 6, the

Numerical Analysis dialog box will again update with the analysis parameters for Plies

Group 1, which gave the value as 0.341 lb, the full data set was exported to Excel using the

Export function shown in figure 6, the results are given in Table 1.

11


Fig 4:- CT1:- Introduction to Composite Design completed part build and detail lay-up.

Plate geometry

Ply Stack

P1 = 0°

Detail A

P2 = 90°

P3 = 90°

P4 = -45°

P5 = -45°

P6 = 45°

P7 = 45°

P8 = 0°

Detail A

Fig 4 (b):- Composite part ply lay-up.

Fig 4 (a):- Final Composite Part Build.

12


Fig 5:- CT1:- Introduction to Composite Design single ply numerical analysis.

13


Fig 6:- CT1:- Introduction to Composite Design composite laminate numerical analysis.

Using the Export function this data was

exported into an Excel spreadsheet and is

presented as Table 1 below.

14


PlyGroup Sequence Ply/Insert/Cut-

Piece Name Material Direction Area(in2) Volume(in3)

Volumic

Mass(lb) Aerial Mass(lb)

Center Of

Gravity - X(in) Center Of

Gravity - Y(in) Center Of

Gravity - Z(in) Cost

Plies Group.1 Sequence.1 Ply.1 GLASS 0 100 0.708661 0.051204 0.04267 -2.80E-16 1.40E-16 0 0.552773



Plies Group.1 Sequence.4 Ply.4 GLASS -45 100 0.708661 0.051204 0.04267 -2.80E-16 1.40E-16 0 0.552773

Plies Group.1 Sequence.5 Ply.5 GLASS -45 100 0.708661 0.051204 0.04267 -2.80E-16 1.40E-16 0 0.552773




Table 1:- CT1:- Introduction to Composite Design Numerical Analysis.

15


The laminate generated in example 1 was not a balanced ply about the Neutral axis therefore

would warp during processing. During the cure cycle a Thermosetting Epoxy resin system hardens

(between 120ºC and 140ºC). When cooling from its maximum processing temperature of 175ºC the

resin contracts approximately 1000 times more than the Fibre, and this mechanism induces

warpage of the Laminate unless the layup is fully balanced about its Neutral axis which can either

be a central plane or an individual ply layer, as shown in figure 7.

16

CT1:- Introduction to Composite Design Balanced Composite Laminate.

Linear Expansitivity (of Fibres) = 0.022 x10^-6

(approximately).

Linear Expansitivity (of Resin) = 28 x10^-6

(approximately).

45º

N A

45º

-45º

-45º

90º

90º

0º

0º

Balanced ply around NA (Neutral Axis) plane. No ply

angle more than 60º separation angle between layers.

Figure 7:- Expansitivity difference between fibre and resin matrix

illustrating requirement for balanced ply layups around the Neutral axis.


The ability to create balanced ply laminates is vital to the construction of real world composite

components and can be achieved for simple laminates using the balanced laminate icon and

selecting the ply group as shown in figure 8. Then reorder the ply sequence so that no adjacent ply

is orientated at angles greater than 60º to the next, in real world situations this requires a more

complex laminate than these simple toolset training examples as we shall see in the tail spar and

cover skin exercises, to react real world loading conditions, this operability is better achieved by

creating a ply layup table in excel and importing it into to Catia V5 model and this is covered later in

Workbook 1. The resulting laminate for this exercise is shown in figure 9 and the numerical analysis

is shown in table 2.

There is also a ply facility in CPE called Plies Symmetry Definition this is used to move a laminate

from one side of a tool surface to the other. In order to use this first crate a symmetry plane about

which the plies will be generated then create a reference surface for the symmetric plies to be

generated from then select the direction about which the symmetric ply is to be generated, select

the ply or ply group to generate the symmetry. This was investigated and will be applied when

appropriate in this study but should not be mistaken as balanced laminate tool.

The rest of the work conducted herein will use balanced ply laminates either using Create

Symmetric Plies method or from balanced ply layup tables generated in excel and imported into the

model.

17

CT1:- Introduction to Composite Design Balanced Composite Laminate.


18

Figure 8:- CT1 Introduction to Composite Design Balanced Composite Laminate.

A balanced ply laminate can be produced

by selecting the ply group and the

balanced ply icon.

Subsequently the ply sequence can be manually reordered so that

adjacent plies are not orientated more than 60º to each other,

manually renumbering the sequence and the ply (use reorder

children).


19

P3 = -45°

P4 = 0°

P5 = 0°

P6 = -45°

P7 = 90°

P8 = 45°

P1 = 45°

P2 = 90°

Detail A

Detail A

Tool face geometry

Laminate Ply Stack

Fig 9 (b):- Composite part laminate lay-up.

Figure 9:- CT1 Introduction to Composite Design balanced composite laminate.

Fig 9 (a):- Final Composite Part Build.

19


20

Table 2:- CT1:- Composite Design Balanced Laminate Numerical Analysis.

PlyGroup Sequence Ply/Insert/Cut-Piece

Name Material Direction Area (in2) Volume (in3)

Volumic


Center Of

Gravity - X(in)

Center Of Gravity

- Y(in)

Center Of Gravity

- Z(in) Cost

Plies Group.1 Sequence.1 Ply.1 U174_T800 45 100 0.511811 0.0277355 0.0412477 -5 -1.26E-15 0 0.299419


Plies Group.1 Sequence.3 Ply.3 U174_T800 -45 100 0.511811 0.0277355 0.0412477 -5 -1.26E-15 0 0.299419



Plies Group.1 Sequence.6 Ply.6 U174_T800 -45 100 0.511811 0.0277355 0.0412477 -5 -1.26E-15 0 0.299419




Learning outcomes:-

From this study I am able to create transition zones within a composite plate that shows the

ply-drops in 3-D; the stagger of each ply, and its respective orientation.

From this study I can now use the module for preliminary design tasks to quickly ascertain

valuable information about the effect a change in ply-drop off will have on weight, location etc.

Methodology:-

In Surface Design a 10in by 15in surface was created on the X-Y plane.

Four edge curves were extracted from the boundaries of this surface, and named curves 1

thru 4 shown in figure 10.

Two mid section curves were created by plane intersection on the surface as shown in figure

10, and named curves 5 and 6.

In the Composite Design module two zones were created as shown in figure 10:

- Zone 1 was created by a contour definition that used curves 1, 2, 6, 4

- Zone 2 was created by a contour definition that used curves 2, 3, 4, 6

The two Zones Laminate Parameters were defined using the same methodology as described

for the CT1 exercise, the parameters being:- Zone 1 - Material = Glass: 1 ply for each of the

orientations 0°/ 45°/ -45°/ 90°: Zone 2 – Material = Glass: 2 plies for each of the orientations

0°/ 45°/ -45°/ 90°.

CT2:- WORKING WITH TRANSITION ZONES.

21


Fig 10:- CT2:- Working with Transition Zones initial geometry.

Left edge

Curve 1

Curve 2

Curve 3

Curve 4

Curve 5

Curve 6

ZONE 1

ZONE 2

22


The next step was to create the Transition Zone between Zone 1 and Zone 2, for this the

Command Sequence – Insert – Preliminary Design – Create Transition Zone was selected.

The Transition Zone Definition dialogue box appeared, Zone 1 was selected as the

Zone/Zone Group input, and the Contours were defined by selecting the following curves:-

5, 2,6,4 (as shown in figure 10), OK was selected to accept the inputs.

Next the Connection Generator was used to check tangency at the edges through the

Command Sequence – Insert – Preliminary Design – Connection Generator, making sure all

dialogue boxes were highlighted Zone Group 1 was selected for analysis, then Apply and

OK were selected.

The resulting Transition Zone is shown in figure 11 with the model and tree maturation that

results from its creation.

The ply stack-up was created using the Plies creation from Zones functionality.

Because the laminate construction consisted of 4 plies in Zone 1, and 8 plies in Zone 2, the

transition zone produced consisted of three staggered plies which were automatically

incremented at a 0.75 inch distance determined by width of the transition zone (i.e. the

distance between curves 5 and 6 being three inches) shown in figure 12.

The 3-D stacking sequence was created using the Ply Exploder with the following settings:-

0.5 Sag: 0.25 step and 20 for the scale. The finished parts stagger transition was examined

as shown in figures 13(a)/(b) and 14, and Numerical Analysis is shown in Table 3.

CT2:- WORKING WITH TRANSITION ZONES (Cont).

23


Fig 11:- CT2:- Interpretation of Connection Generator Output.

The transition zone build sequence

in the model tree.

Zone Connection generation sequenced

in the model tree.

Green line indicates that a connection between a

transition zone and a Top zone exists. (Trans Zone

1 and Zone 2)

Blue line indicates that a edge connection between

two transition zones exists. (Zone 1 and Trans

Zone 1).

Yellow line indicates that a free edge exists at the

conceptual zones boundary (i.e. the boundary of

the reference surface).

Magenta line indicates that a edge connection

between two transition zones exists (i.e. between

Zone 1 and Trans Zone 1)

Numbers Indicate ply count for each zone.

24


Fig 12:- CT2:- Creating plies from zones transition zone schematic.

The first stagger in Zone 2 starts at the white line this is the 0° ply.

The second stagger in Zone 2 starts at the green line this is the -45° ply.

The third stagger in Zone 2 starts at the red line this is the 45° ply.

The fourth stagger in Zone 2 starts at the blue line this is the 90° ply.

0.75 in stagger

25


Detail A

0º Ply drop

-45º Ply drop

45º Ply drop

90º Ply drop

Reference surface

(X)

(Y)

(Z)

Fig 13(a/b):- Working With Transition Zones Ex 1 completed part and ply stack-up.

Figure 13(b) Ply stagger in transition zone.

P8 = 0º

P7 = 45º

P6 = 90º

P5 = -45º

Detail A

Figure 13(a) Final Transition Zone Part Geometry.

26


Fig 14:- Working With Transition Zones Ex 1 completed part build model tree.

27


PlyGroup Sequence Ply/Insert/Cut-

Piece Name Material Direction Area(in2) Volume(in3)

Volumic


Center Of Gravity -

X(in) Center Of Gravity -

Y(in) Center Of Gravity -

Z(in) Cost

Plies Group.1 Sequence.1 Ply.1 GLASS 0 90 0.637795 0.0460836 0.038403 4.5 5 0 0.497496

Plies Group.1 Sequence.2 Ply.2 GLASS -45 97.5 0.690945 0.0499239 0.0416033 4.875 5 0 0.538954


Plies Group.1 Sequence.4 Ply.4 GLASS 90 112.5 0.797244 0.0576046 0.0480038 5.625 5 0 0.62187

Plies Group.1 Sequence.5 Ply.5 GLASS -45 150 1.06299 0.0768061 0.0640051 7.5 5 0 0.82916




Table 3:- CT2:- Working with Transition Zones Exercise 1 Numerical Analysis.

28


On completion of the first working with transition zone exercise, a further exercise was

conducted to determine the effects of changing the numbers of plies in Zone 2 in exercise 2

an extra 0º and 90º ply were added.

The resulting ply build up using the Plies creation from zones function gave the transition

zone schematic shown in figure 15, with 5 stagger lines 0.5 inches apart.

The resulting transition zone ply drop-off started with a single 90º ply followed by two

consecutive 0º ply drops, followed by a -45º, and a 45º, and ending in another 90º ply drop,

as shown in figures 16(a)/(b).

The 3-D ply stack was built using the Ply exploder function and the following settings:- 0.5

Sag: 0.25 step and 20 for the scale and is shown in figure 17.

The addition of these plies resulted in change in the Zone 1 ply stack up as shown in figure

16(b) Detail A, starting with a 90º ply instead of a -45º as in figure 13(b) Detail A, but both

finish with the outer 0º ply as expected.

The Numerical Analysis tool was used to obtain comparative data for this modified

composite configuration and the data is given in Table 4 below.

This exercise concluded the working with transition zones preliminary design tutorial,

applications in the panel and spar designs are given below.

CT2:- WORKING WITH TRANSITION ZONES (Cont).

29


Fig 15:- CT2:- Creating plies from zones transition zone schematic Exercise 2.

0.5 in stagger

The first stagger in Zone 2 starts at the blue line this is the 90° ply.

The second stagger in Zone 2 starts at the grey line this is the 0° ply.

The third stagger in Zone 2 starts at the grey line this is the 0° ply.

The forth stagger in Zone 2 starts at the green line this is the -45° ply.

The fifth stagger in Zone 2 starts at the red line this is the 45° ply.

The sixth stagger in Zone 2 starts at the blue line this is the 90° ply.

Numbers Indicate ply count for each zone.

30


Fig 16(a/b):- Working With Transition Zones Ex 2 completed part and ply stack-up.

(X)

(Y)

(Z)

Figure 16(a) Final Transition Zone Part Geometry.

P10 = 0º

P9 = -45º

P8 = 45º

P7 = 90º

Detail A

Detail A

Reference surface

90º Ply drop 0º Ply drop

0º Ply drop

90º Ply drop

-45º Ply drop

45º Ply drop

Figure 16(b) Ply stagger in transition zone.

31


Fig 17:- Working With Transition Zones Ex 2 completed part build model tree.

32



Name Material Direction Area(in2) Volume(in3)

Volumic


Center Of Gravity -

X(in) Center Of Gravity -

Y(in) Center Of Gravity -

Z(in) Cost



Plies Group.1 Sequence.3 Ply.3 GLASS 0 100 0.708661 0.051204 0.04267 5 5 0 0.552773








Table 4:- CT2:- Working with Transition Zones Exercise 2 Numerical Analysis.

33


Learning outcomes:-

From this study I am able to create limit contour features.

From this study I am able to use limit contouring with Gap Fill and extrapolation techniques.

From this study I am able to use cut-pieces to create a limit contour.

From this study I am able to create a limit contour feature using non - relimited curves.

From this study I have learnt how to manipulate the stagger and step of a limit contour.

From this study I can now use the module for preliminary design tasks to quickly ascertain

valuable information about the effect a change in ply-drop off will have on weight, location

etc.

Methodology:-

The reference surface was created in surface design 10 inches wide by 17.606 inches long

with a 8 inch radius curve section as shown in figure 18.

Two ply zones were created and a transition zone using a transition zone refinement number

of 4, as shown in figure 18.

The Zone Definition consisted of 11 plies in Zone 1 and 5 plies in Zone 2 as detailed below.

CT3:- LIMIT CONTOUR DESIGN.

34


Fig 18:- Limit Contour reference geometry and zones.

10 inch

35


Zone Definition:

Zone 1: 11 plies

4: 0º plies

3: 45º plies

2: -45º plies

2: 90º plies

Zone 2: 5 plies 2: 0º plies

1: 45º plies

1: -45º plies

1: 90º plies

Following creation of the ply zones and the transition zone in Composite Design, the model

was switched back to surface design to create two separate reference curves C 1 and C2

shown in figure 19(a), which were individually projected on to the reference surface as

shown in figure 19(b).

CT3:- LIMIT CONTOUR DESIGN (Cont).

36


Projected Curve:- C

2

Projected Curve:- C 1

Figure 19(b) Projection of reference curves.

Fig 19:- Limit Contour creating reference curves.

Transition Zone Boundary (white line)

Curve:- Ref C 2

Curve:- Ref C 1

Figure 19(a) Creation of reference curves.

37


Figure 20:- Ply Stagger Schematic.

C 1 C 2

Ply stagger lines in transition zone.

Back in Composite Design plies were created using the zones and selecting the default

settings.

The resultant ply stagger schematic is shown in figure 20, the ply orientation of each ply

drop is indicated by the respective colour of lines representing the ply stagger within the

transition zone.

The Ply Exploder was then applied with the tessellated surface option selected with the

following tessellated set:- sag value = 0.25: and step value = 0.20.

The resulting laminate is shown in figure 21.

Figure 20:- Ply stagger lines schematic.

38


Fig 21:- Limit Contour Model appearance after ply exploder application.

39


Exercise 1:- Creating a Limit Contour:

The Create a Limit Contour for a Ply icon was selected for which the alternative Command

Sequence selection was:- Insert – Plies – Limit Contour.

The Limit Contour dialogue screen was presented as shown in figure 22 and Plies Group 1

was selected as the Entity.

The Relimiting Curve multi-selection icon was selected in order to enable the picking of the

two curves previously created (i.e. the blue curves C 1 and C 2) as the Relimiting Curves.

A Blue arrow was generated for each curve indicating the direction that the plies will be

created. The default direction should have pointed outward from the enclosed area bounded

by curves C 1 and C 2, however this was not the case for the arrow on curve C 1, therefore

the Inverse Direction button in the Limit Contour dialogue screen was used to switch its

direction (note changing the arrows direction just by clicking on them will not change

the resultant ply truncation and the Inverse Direction button must be used).

The Multi-selection dialogue screen was then closed and OK was selected in the Limit

Contour creation screen.

The result was a truncation of the transition zone lines at the boundary of the limit curve as

shown in figure 23, then the laminate was rebuilt using the Ply Exploder function to reflect the

new definition as shown in figure 24.


40


Fig 22:- Creation of the Limit Contour.

Multi-Selection icon

Invert Direction button

Curve C 1

Curve C 2

Limit Contour Icon

41


Fig 23:- Updated Transition Zone with Limit Contour.

Limit Contour Boundaries

(Curve C 1 and C 2).

The blue box surrounds the

newly transition zone lines.

42


Fig 24:- Updated Transition Zone with Limit Contour.

A portion of each ply has been removed

based on the boundary conditions set forth by

the limit curve definition (i.e. C 1 and C 2).

Reference Surface. This profile can be modified by simply modifying

the curve sketch and updating accordingly.

43



Exercise 2:- Developing a Limit Contour using Cut-Pieces and the Extrapolation Joint Type:

Using the existing model, the plies and existing geometrical set created for exercise one were

deleted.

Two new curves were then created as shown in figure 25.

These curves were then projected on to the reference surface as in exercise 1, the resulting

curves being designated:- C 1a and C 2a respectively.

The Limit Contour Icon was selected, and Plies Group 1 was selected as the Entity.

The two new curves C 1a and C 2a were selected as the Relimiting Curves, making sure that

the blue directional arrows were pointing outwards as shown in figure 25, and the Multi-

Selection dialogue screen was closed.

In the Limit Contour dialogue screen the Extrapolation Joint Type was selected, and then OK

to implement the input as shown in figure 26.

After selecting OK, the laminate updated to reflect a new transitional zone configuration. Note

the truncation of the step drop off schematic at the boundary curve C 1a, as can be seen in

figure 27(a) which shows the updated Laminate Configuration.

Figure 27(b) shows the updated Ply Stack configuration.

44


Cut-Pieces

Red circle shows gap

between line segments.

Curve C 1a

Curve C 2a

Directional arrow for curve C

1a

Directional arrow for curve C 2a

Fig 25:- Developing a Limit Contour using Cut-Pieces.

45


Fig 26:- Limit Contour from Cut-Pieces using the Extrapolation Joint Type.

Relimiting Curve Joint Type selection

46


Fig 27(a)/(b):- Limit Contour with Extrapolation Joint Type.

Figure 27(a) Updated Laminate Configuration

After selecting OK, the laminate updated to reflect a

new transitional zone configuration. Note the

truncation of the step drop off schematic at the

boundary curve C 1a (extended in red).

The discontinuous blue curves C 1a and C 2a were

joined to form a continuous L-shaped boundary curve

( red ellipse in fig 27(a) ).

The resultant Ply-Stack was as show below in fig

27(b).

Curve C 1a

(extrapolated).

Curve C 2a

(extrapolated).

Figure 27(b) Updated Ply Stack Configuration

47


Exercise 3:- Developing a Limit Contour using Cut-Pieces and the Gap Fill Joint type:

Using the existing model, the plies and geometric set created from the exercise 2 were

deleted, and a new ply group from zones was created, the

The Limit Contour Icon was selected, and Plies Group 1 was selected as the Entity.

The two new curves C 1a and C 2a were selected as the Relimiting Curves, making sure

that the blue directional arrows were pointing outwards as shown in figure 28, and the Multi-

Selection dialogue screen was closed.

In the Limit Contour dialogue screen the Gap Fill Joint Type was selected, and then OK to

implement the input as shown in figure 28.

After selecting OK, the laminate updated to reflect a new transitional zone configuration. Note

the truncation of the step drop off schematic at the boundary curve C 1a, as can be seen in

figure 29(a) which shows the updated Laminate Configuration, and now curves C 1a and

curve C 2a join together by forming an angled segment between the two end points of the

curves.

Figure 29(b) shows the updated Ply Stack configuration.

Therefore this process dose not extrapolate the curves, but simply connects the vertex of

each line segment.


48


Fig 28:- Limit Contour from Cut-Pieces using the Gap Fill Joint Type.


49


Fig 29(a)/(b):- Limit Contour with Gap Fill Joint Type.



Curve C 1a

Curve C 2a.

As in the previous exercises the ply laminate is

updated to truncate at the boundary curve.

The discontinuous blue curves C 1a and C 2a were

joined by an angled segment between the two end

points of the curve to form a continuous boundary

curve ( red ellipse in fig 29(a) ).

50


Exercise 4:- Developing a Limit Contour with Staggered Values and Extrapolation Joint Type:

The Create Plies from Zones Icon was selected, and the Plies Exist dialog box appeared and

No was selected as the answer to “Do you want to delete existing plies”.

A second plies group appeared in the model tree this was Plies Group 2 and this was used to

create the new Limit Contour as shown in figure 30.

Plies Group 2 was selected as the Entity in the Limit Contour dialogue screen, as shown in

figure 30.

The two Relimiting Curves C 1a and C 2a were selected with the Extrapolation Joint Type, as

shown in figure 30.

In the Multi-Section dialogue screen the stagger values were set at 0,1 for curve C 1a and

0.25 for curve C 2a, as shown in figure 30, and OK was selected to accept this input.

The resultant updated laminate configuration is shown in figure 31(a) with the new ply stagger

geometry from both C 1a and C 2a.

The updated ply stack configuration is shown in figure 31(b), and illustrates the power of this

module to emulate a realistic ply build up.

Figure 32 shows the completed limit contour with model tree.

Numerical Analysis was conducted on both Plies Group 1 Limit Contour Cut-Pieces, and Plies

Group 2 Limit Contour Staggered Values and is presented in tables 5 and 6 respectively.


51



Stagger value input for both curves

Fig 30:- Limit Contour with Staggered Values and Extrapolation Joint Type.

52


Fig 31(a)/(b):- Updated laminate and ply stack Limit Contour with Staggered Values.



New ply stagger

from Curve C 1a

New ply stagger

from Curve C 2a

New ply stack from

Curve C 1a

New ply stack from

Curve C 2a

53


Fig 32:- Limit Contour with Staggered Values completed part and model tree.

54


Table 5:- CT3:- Limit Contour Cut-Pieces Ply Group 1 Numerical Analysis.



Volumic Mass(lb)

Aerial Mass(lb)

Center Of Gravity - X(in)

Center Of Gravity - Y(in)

Center Of Gravity - Z(in)

Cost

Plies Group.1

Sequence.1 Ply.1 GLASS 45 45.5 0.322441 0.0232978 0.0194149 3.5 1.75 1.38E-15 0.251512

Plies Group.1

Sequence.2 Ply.2 GLASS -45 52.0814 0.369081 0.0266678 0.0222232 4.00626 1.75 1.38E-15 0.287892

Plies Group.1


Plies Group.1


Plies Group.1

Sequence.5 Ply.5 GLASS 45 71.8258 0.509001 0.0367777 0.0306481 5.52499 1.75 0.00218136 0.397033

Plies Group.1


Plies Group.1

Sequence.7 Ply.7 GLASS -45 152.235 1.07883 0.0779503 0.0649586 10.6049 1.02912 1.25593 0.841512

Plies Group.1


Plies Group.1


Plies Group.1

Sequence.10

Ply.10 GLASS 0 152.235 1.07883 0.0779503 0.0649586 10.6049 1.02912 1.25593 0.841512

Plies Group.1

Sequence.11

Ply.11 GLASS 45 152.235 1.07883 0.0779503 0.0649586 10.6049 1.02912 1.25593 0.841512

55


Table 6:- CT3:- Limit Contour Staggered Values Ply Group 2 Numerical Analysis.



Volumic Mass(lb)

Aerial Mass(lb)

Center Of Gravity - X(in)

Center Of Gravity - Y(in)

Center Of Gravity - Z(in)

Cost

Plies Group.2


Plies Group.2

Sequence.13 Ply.13 GLASS -45 52.8827 0.374759 0.0270781 0.0225651 4.00626 1.7 1.38E-15 0.292321

Plies Group.2


Plies Group.2


Plies Group.2


Plies Group.2


Plies Group.2

Sequence.18 Ply.18 GLASS -45 164.848 1.16822 0.0844091 0.0703409 10.4798 0.76646 1.17903 0.911238

Plies Group.2


Plies Group.2


Plies Group.2


Plies Group.2


56


This section covers the design rules applied in the detail design of airframe structures based on my

professional experience within BAE SYSTEMS and Cranfield University, used by myself in the

design of airframe components covered in my LinkedIn presentations, and further applied to the

FATA design project, primarily this section will deal with wing / empennage design.

Aircraft OML Surfaces:- Peel plies should not be used. Requirements for addition of non-

structural plies on aircraft OML surfaces are listed in the External Surface Features Design

Guide for wing cover skins, fuselage, and empennage.

All Other Aircraft Surfaces:- Internal surfaces of graphite composites in contact with

aluminum or other dissimilar materials shall incorporate a glass ply in the contact area. This

applies to mechanically fastened, co-cured or secondarily bonded joints. For BMI materials, the

glass barrier shall fully cover the laminate surface. For epoxy-based laminates the glass barrier

ply should extend a minimum of 1 inch beyond the contact rejoin of the metallic substructure.

For NDI purposes, the use of a peel ply on the IML surface is encouraged. This peel ply will

enhance the effectiveness of the NDI tools. If sacrificial plies are co-cured to the composite

panel than a peel ply shall not be used. If the outermost structural ply material is fabric, the ply

shall be the least critical ply (generally, but not always a ± 45º fabric ply). If the outermost ply

material is tape, the surface plies shall consist of two tape plies orientated in the least critical

directions (generally one +45º and one -45º ply). However, using a ply of woven fabric on the

exterior surface will reduce “splintering” during trim and drill operations thus requiring less repair

work to be performed on detail parts. Generally, incorporation of carbon fabric or thin glass

scrim ply on part surface is encouraged to prevent shop handling and machining damage to

tape laminates. 57

Section 2:- Design rules applied to main design exercises.


58

COVER SKINS: - The covers form the lifting surface of the wing box and are subjected to span-

wise bending flight loads, the upper wing cover is subjected to primary compression loads, and

lower wing cover is subjected to primary tension loads. The upper wing covers are also subjected to

aerodynamic suction and fuel tank pressures, and both covers are subjected to chord-wise shear

due to the aerodynamic moment on the wing torsion box. Composite wing cover skins shown in

figure 33 can be aeroelastically tailored using: - 0º plies to react span-wise bending: 45º and -45º

plies to react chord-wise shear: and 90º plies to react aerodynamic suction and internal fuel tank

pressures, theses cover skins are monolithic structures and not cored. Combined with co-bonded

stringers, this produces much stronger yet lighter covers which are not susceptible to corrosion and

fatigue like metallic skins. The production method of these cover skins is by Fiber Placement:-

which is a hybrid of filament winding and automated tape laying, the machine configuration is

similar to filament winding and the material form is similar to tape laying, this computer controlled

process uses a prepreg Tow or Slit material form to layup non-geodesic shapes e.g. convex and

concave surfaces, and enables in-place compaction of laminate, however maximum cut angle and

minimum tape width and minimum tape length impact on design process. The wing cover skin

weight in large transports, can be reduced by applying different ply different transition solutions to

the drop off zones as shown in figure 34, maintaining the design standard 1:20 ramps in the

direction of principal stress (span-wise), and using 1:10 ramps in the transverse (chord-wise)

direction, as shown for the Airbus A320 lower wing covers, this requires stress approval based on

analysis. Because the wing chord depth of the transport aircraft considered exceeds 11.8” to reduce

monolithic cover skin weight and inhibit buckling co-bonded CFRP stringers are used as detailed

below and shown in figures 35 to 38.

Design of aircraft wing CFC cover skins structures


Fig 33:- Fibre Orientation Requirements For CFC Wing Skins.

Tension

Compression

Centre Of Pressure

Engine / Store Loading

Flexural Centre

0º MATERIAL TO REACT SPANWISE BENDING

90º MATERIAL TO REACT

INTERNAL FUEL PRESSURES

AND AERODYNAMIC SUCTION 45º AND -45º MATERIAL TO

REACT CHORDWISE SHEAR

59


60

Fig 34:- Weight reduction by of ply drop off design changes to lower wing covers.

PLY DROP OFFS: - 1:20 SPANWISE / 1:20 CHORWISE.

(More usual to have symmetrical ply drop off e.g. all 1:20).

PLY DROP OFFS: - 1:20 SPANWISE / 1:10 CHORDWISE.

(Although in some cases un-symmetrical ply drop off e.g. 1:20 in

direction of principal stress and 1:10 in the transverse direction).

WEIGHT REDUCTION OF COMPOSITE

WING COVER SKINS.


<2.9 inch ~ SQUARE EDGE / TAPERED EDGE

(HONEYCOMB SANDWICH)

2.9 inch - 3.9 inch (WAFFLE STRUCTURE)

3.9 inch - 11.8 inch (RIBS AND SPARS)

> 11.8 inch (STRINGER STIFFENED SKIN PANEL)

Figure 35:- Guide to typical effective depths for Sub-structure.

61


As a Rule of Thumb:- The mass of the skins is in the order of twice

that of the sub-structure. Therefore for transports and bombers with

thick wing sections stringers are used bonded to the internal skin

surface as shown in fig 36(a) for the Airbus A350 wing skins. Where

the wing chord thickness is greater than 11.8 inches,( e.g. the FATA

wing fig 36(b) root chord thickness 42.5 inches) and semi-spans grater

than 20 feet it is more efficient to use integral skin stringers in order to

reduce the skin thickness an hence reduce the skins weight. The semi-

span of the FATA wing is 106 feet., and the weight trade study showed

that that a three spar 13 stringer solution was the best solution to

reduce wing skin (cover) weight.

Fig 36(a) Airbus A350-900

skin stringer layout.

Fig 36(b) Future Advanced Technology

Aircraft baseline aircraft wing box

structural layout.

Fig 36:- For transport and bomber aircraft the skin buckling is inhibited with stringers.

62


Fig 37:- For fighter aircraft the spar pitch is the critical factor in inhibiting skin buckling.

As a Rule of Thumb:- The mass of the skins is in

the order of twice that of the sub-structure.

Therefore where the wing chord thickness is

between 3.9 inches and 11.8 inches, it is more

efficient to increase the number of spars in order to

reduce the skin thickness an hence reduce weight.

Although for highly loaded combat aircraft spars

are used in wings with root chord thicknesses up to

15 inches in combination with stiffeners.

N.B. in military combat aircraft

wing ribs are generally limited

to the weapons carriage and

fuel tank boundary stations. i.e. long thin panels are more

efficient at resisting buckling of

skins. F/A-24 Concept Advanced fighter aircraft wing structural layout.

63


The types of composite stringer which can be used based on my experience.

“L” Section Stiffeners:- are typically used as “panel barkers” and are usually mechanically

attached to skin panels. “L” stiffeners are fabricated on IML tooling with a semi-rigid caul

sheet, often fiberglass, on the OML surface to produce a smooth finish and reduce radius thin

out.

“Z” Section Stiffeners:- are usually mechanically attached to the skin panel and are typically

used to provide additional stiffness for out-of-plane loading. “Z” sections may be fabricated

by the RTM or hand-laid methods.

“I” Section Stiffeners:- are typically used as axial load carrying members on a panel

subjected to compression loading. “I” sections are fabricated by laying up two channel

sections onto mandrels and placing them back-to-back. A minimum of two tooling holes (one

at each end) is typically required to align the mandrels. Two radius fillers (“noodles” or

“cleavage filler”) are placed in the triangular voids between the back-to-back channels. On

one of the two flat sections of the stiffener a “capping strip” is used to tie the two flanges

together. The flanges on the cap side should have a draft (91º ± 1º) to ease mandrel removal

post cure. All “I”- beam flanges should have sufficient width to allow mechanical attached

repair.

“T” Section Stiffeners:- are a simplified version of the “I” section stiffener. “T” sections may

be used as either axial load carrying members or as panel breakers. “T” sections stiffeners

may be used as a lower cost alternative to “I” sections if the panel is designed as a tension

field application and the magnitude of reverse (compression) load is relatively small.

64


Fig 38:- Composite Stiffener Types Based on my design experience.

“L” Section Stiffener (bonded or

mechanically attached panel breaker).

“Z” Section Stiffener (mechanically attached to

provide additional stiffness for out of plane loading).

“I” Section Stiffener (used as axial load carrying

members on panel under compression loading).

Channel

sections Capping

strips

Cleavage

fillers

“T” Section Stiffener (used as axial load carrying

members on panel under tension loading).

Capping strip

Cleavage filler

Channel

sections

65


Radius fillers are necessary in T - and I – type composite stiffeners and spars. See figure 38 for a

2-D depiction of radius / cleavage fillers.

There are several types of filler material that have been used in previous design studies

including:- rolled unidirectional prepreg (of the same fiber / resin as the structure); adhesives; 3-D

woven preforms; groups of individual tows placed in the volume; and cut quasi-isotropic laminate

sections. Experimentation has shown the how effective these have been and a brief summary is

as follows:-

Resin / adhesive noodles – Poor

Tow noodles – Fair

Braided noodle – Good

Braided “T” preform - Good to Excellent.

If rolled prepreg is used, ensure that the volume of the material to be rolled is a close match with

the cavity to be filled and consider using a forming tool to shape the noodle to near final

configuration. Also, it has been found that using a layer of softening adhesive rolled with the

noodle prepreg material will help alleviate cracking due to thermal mismatch between the noodle

and the surrounding material.

The capping strips are bonded in place using BSL322, supported film adhesive to give

constant/minimum glue line thickness of 0.005” per ply, 2 plies max typically, and has applications

in the bonding of primary aircraft structure, bonding honeycomb panels and structural repairs.

Composite Stiffener Radius Fillers (Noodles) based on academics and test experience.

66


WING SPARS: - The spars in conjunction with the covers transmit the bending and torsion loads of

the wing box, and typically consists of a web to react vertical shear, and end flanges or caps to

react the bending moment. In modern transports there are two full span spars, and a third stub

spare in wide chord wings to take engine aft pylon mount loads from the pylon drag strut as in the

case of the A300, A330, A340, and A380, and these spars are currently produced as high speed

machined aluminium structures. However the latest generation of large transport aircraft e.g. the

Airbus A350 and Boeing 787 families use composite spars produced by fiber placement as C -

sections laid on INAVR tooling as shown in figure 39, and are typically 88% 45º / -45º ply

orientation to react the vertical shear loads, in the deflected wing case, the outer ply acts in tension

supporting the inner ply which in compression as shown in figure 40, because the fibers are strong

in tension but comparatively weak in compression. The spars can be C section or I section

consisting of back to back co-bonded C-sections, and for this study the baseline reference wing

spars are C sections, and consists of three sub-sections design, due to the size of component

based on autoclave processing route constraints detailed in the FATA study. Although 0° plies are

generally omitted from the spar design 90° plies are employed in approximately 12% of the spar

lay-up as shown in figure 41, where there are bolted joints, tooling hole sites, to react pressure

differentials at fuel tank boundaries.

The separation of web and flange spar joggles is shown in figure 42 and the separation of joggles

from changes in laminate thickness are shown in figure 43. The support of joggles in structural

assemblies is shown in figure 44.

67

Design of aircraft CFC wing spar structures.


68

Figure 39:- Airbus A350 Composite spar manufacture and assembly.

CFRP Spar C section with apertures for edge control surface attachment.

Wing torsion box section with “C” section spars, ribs, and edge control

surface attachment fixtures.


Fig 40:- Carbon Fibre Composite ply orientations in wing spars.

-45º 45º

Composite Wing Spar Design

Spars are basically shear webs attaching the upper and lower skins together

The lay-up is therefore predominately +45° / -45 °.

Typically 88% of a spar lay-up is made up of +45° and -45° plies.

In the deflected wing loading case (red dashed line) the outer ply is chosen to be acting

in tension which acts to support the weaker compressive ply.

Vertical web stiffeners and rib attachments are bolted or co-bonded to the shear webs.

Wing deflected case

69


Fig 41:- Carbon Fibre Composite ply orientations in wing spars.

90º Plies to react pressure

differentials at fuel tank

boundaries.

90º Plies locally in way of

bolted joints.

Composite Wing Spar Design

0o Plies are generally omitted from spar lay-up however, 90o plies

are added in typically 12% of spar lay-up

70


Fig 42:- Separation of Web and Flange Joggles in CFC spars.

VIEW ON A-A

A A

Joggles in webs are to be offset from flange joggles by

as greater distance as possible, (a minimum distance

of one fastener pitch is standard).

2.5 x d

3 x d 6 x d

2.5 x d

3 x d

71


Fig 43:- Separation of Joggles from changes in laminate thickness in CFC spars.

0.630 in

d = 1.0 in

0.315 in

Internal fillet radius

0.496 in 5.5in

7.5in

(a) Full component spar with web thickness change and web joggle.

30in

d = 1.0 in Web thickness transition

(b) Lower section of spar in (a) showing minimum separation of web thickness change and web joggle.

Origin of ply ramp Sep 5 x d

(minimum)

72


Fig 44(a)/(b):- Support of Joggles in CFC spars in structural assemblies.

Joggle is supported by a GRP tapered packer.

SHIM Packer

(a) TYPICAL BONDED

ASSEMBLY Anti – peel fasteners

Utilize the ability to taper the feet of adjoining members this

simplifies the geometry of the joggle.

(b) TYPICALASSEMBLY

OF PRE-CURED

DETAILS

73


Align fibres to principle load direction.

The lay-up ply orientations must be balanced about the mid-plane (neutral axis) of the

laminate, as so to avoid distortion during cure.

Outer plies shall be mutually perpendicular to improve resistance to barely visible impact

damage.

Overlaps and butting of plies:-

U/D, no overlaps, butt joint or up to 2mm gap.

Woven cloth, no gaps or butt joints, 15mm overlap (see figure 48).

No more than 4 plies (0.125mm per ply) of a single orientation in one stack within a

laminate.

A maximum of 67% of any one orientation shall exist at any position in the laminate.

4 plies separation of coincident ply joints rule (ply stagger rules) shown in figures 45 and 46

below.

Ply separation overlap and stagger requirements for woven cloth laminates are shown in

figures 47 and 48 below.

Lay-up Guidelines based on design best practice.

74


Figure 45:- Application of ply layup rules in general terms.

75


Figure 46:- Structural design ply lay-up guidelines

The 4 ply separation of coincident ply joints rule.

76


77

Figure 47:- Structural design requirements for Woven cloth


78

Figure 48:- Structural design requirements for Woven cloth


Lay-up Guidelines based on best practice (continued).

Changes in the laminate thickness should occur evenly with a taper rate of 1 in 20 in the

principal load direction. This can be reduced to 1 in 10 in the traverse direction as with the

A320 lower wing skin figure 34.

All ply drop-offs to be internal and interleaved with full plies

Internal corner radii of channels are important because, sharp corners result in bridging and /or

wrinkling of the prepreg, thus weakening the part, and sharp also result in high internal stress

under bending loads which can lead to premature failure therefore the designer shall make the

internal radii as large as practical within these limits:-

„t‟ < 2.5mm, radius = 2t or 3.0mm whichever is greater

„t‟ 2.5mm, radius = 5.0mm

For F-35 a internal radius of 0.25” is preferred, with the minimum corner radius being the

larger of the following two values 0.12” or the laminate thickness.

The designer shall not drop plies or core material in the corner radius, and plies can only

dropped at a distance equal to or greater than whole laminate thickness from the tangent of the

corners outer radius.

While co-curing honeycomb sandwich panels, beware of ply quilting during cure over the core

area, need for core stabilisation and reduced cure pressures.

The minimum skin thickness over honeycomb sandwich panels to prevent moisture ingress to

be respected (typically 1mm for UD and 1.5 for cloth). Use of surface films on thin skin panels

such as Tedlar can be considered. 79


Figure 49:- Plie lay-up rosette definition and positioning.

80

The lay-up rosette definition.

The position of the Ply rosette.


81

Figure 50:- Plie stagger rosette definition and positioning.

START POINT

Lay-up Guidelines

A ply stagger rosette is displayed on the drawing face:

This defines the position of joints in successive ply

courses, ensuring that they are controlled to within the

project requirements. Generally the four ply separation

rule applies.


Guidelines for Staggering Ply Boundaries in Ramps.

Changes in laminate thickness are usually accomplished by dropping two plies at one (one

on each side of the neutral axis N.A. plane of symmetry).

Only one ply should be dropped at any location if the ply is equal to or grater than 0.3302mm

(0.013 inch) thick.

Sequence the ply terminations to produce a smooth transition in stiffness through the

transition region (do not drop all the 0º plies, then all 45º plies, etc.).

No more than 4 adjacent plies shall be terminated between continuous plies, good design

practice is a maximum of two – ply terminations.

Sequence the ply terminations the total thickness in order to maximize the distance between

ply terminations in adjacent plies, maximum strength is achieved if ply terminations in

adjacent plies are a minimum of 0.5 inches apart.

Ply drop-offs shall be avoided near concentrations such as cutouts, corners, and joggles.

Ply drop-offs shall be balanced with respect to the neutral axis (N.A.) of the laminate to

maintain symmetry and avoid warpage.

Balance and symmetry may be relaxed over very short distances.

For uni-directional material avoid tape buildups shorter than 12.7mm (0.5 inches) (tape might

migrate during the cure cycle).

Avoid dropping a 0º ply that is adjacent to a 90º ply. A 90º ply has little load carrying

capability relative to the 0º ply as there are no reinforcing fibers in the 0º direction.

82


Adhesives are best when used in shear – dominated applications. Avoid bonded structures in

areas that have high delta pressure loads.

Avoid as much as possible out – of – plane loading of laminates. The thru-thickness (z-

direction) properties of the laminate are significantly lower than the in-plane properties of the

laminate, (e.g. composite angles used as tension clips).

Use a rub strip (or Teflon paint) on moving surfaces to prevent abrasion of the load carrying

composite structure.

Bonding adhesive, when used in composite structures shall be non-hydroscopic (i.e. non-

moisture absorbing.).

The designer should take advantage of composite material capabilities to reduce part counts,

fastener counts and assembly complexity by combining parts, even if they are separated later

during trim operations. The inclusion of co-cured stiffeners or longerons with the skin are

examples of this practice.

To avoid delamination at a “rabbet” step (sharp step change in laminate thickness) details

during un-bagging, wrap a continuous ply over the step feature. This ply can be non-structural

such as fiberglass.

General Fastener Spacing And Edge Guidelines, contains the direction on fastener spacing

and minimum edge distance as used in this study.

Design Guide For Fuel Tanks, contains the minimum fastener spacing for fuel tanks.

More General Design Guidelines.

83


Inspection Zones are defined as separate levels or classes into which composite laminates and

bonded assemblies shall be divided for evaluation using ultrasonic and / or radiographic

techniques. In addition, each part or assembly may have different zones specified for different

regions of the part or assembly. The zones are labeled “A” through “E” and are defined in

reference (2) “Inspection, Acceptance Criteria, for Composite Parts and Assembles”. The

inspection zone is normally specified on the Engineering drawing, however if not the inspection

zone shall default to “Zone B”.

Unidirectional Material Limits on Adjacent Plies of Same Orientation:- To avoid matrix micro-

cracking in unidirectional laminates, it is recommended that a limited number of plies of like-

orientation be stacked together for toughened matrix resins: For skins, webs and web caps, a

maximum of 0.0336” total thickness (4 plies of 0.0084”ply material, or 6 plies of

0.0053”ply material). If this guideline is violated, analysis is required to show that the

interlaminar shear allowable is not exceeded.

Ply Splicing Overview:- Due to material width constraints, one piece of material is not always

large enough to make the entire ply. Splices are the interfaces within the ply between two or

more pieces of material in order to create a ply of the necessary size. Splices can be made in

two ways:- butt splice and overlap splice. Plies with dissimilar ply orientation shall not be

spliced. A group of IPT members must be involved with the mapping of where the splicing of

plies will occur the input of:- Manufacturing: Materials: Design: and Stress, are all needed to

coordinate the required splice locations.

More General Design Guidelines (continued).

84


General Design Guidelines for Ply splicing.

Butt Splices:- A butt splice (also known as a course splice when referring to unidirectional tape

materials) is created by placing the two pieces of material side by side with no overlap and

within accepted gap limits. This type of splice is typical for unidirectional materials and is always

parallel to the fiber direction as shown in figures 45 and 46. Butt splicing of fabric plies can only

be done in circumstances where a detailed stress analysis has found that this splice type is

acceptable. In cases where analysis determines a part does not meet design requirements with

a butt splice, then an overlap splice must be used. If a butt splice is used it shall be created as

per the appropriate process specifications guidelines outlined in the following slides.

Overlap Splices:- An overlap splice is formed by one piece of material laying over the adjacent

piece of material by a specified distance. Overlap splices are not used with unidirectional

material. This splice type is only used with woven fabric material. A minimum of 12.7mm (0.5”)

overlap is required, and a overlap of 25.4mm (1.0”) is recommended as the guideline shown in

figures 47 and 48.

Splicing Hand lay-Up Carbon / Epoxy Laminates:- Splicing requirements for carbon / epoxy

fabric, tape, peel ply, and surface barrier material (scrim) are given in company specifications

LMA-PL005 for example:- Minimum stagger distance between splices are for Fabric & Tape >=

300mm (12”) wide minimum stagger is 50.8mm (2”), and for Tape <= 300mm (12”) wide the

minimum stagger is 20.4mm (1”). The splice stagger pattern shall not be repeated more than

every fifth like-orientated ply for tape. The splice stagger pattern shall nor be repeated for fabric.

85


Figure 51:- Control of Ply Joints / splices.

86


Splicing Hand Lay-Up Carbon / BMI Laminates:- Splicing requirements for carbon / BMI

fabric and tape are given in company specifications:- Minimum stagger distance between

splices are for Fabric & Tape >= 12” (300mm) wide the minimum stagger is 2” (50.8mm) and for

Tape < 12” (300mm) wide the minimum stagger is 1” (20.4mm). The splice pattern shall not be

repeated more often than every fifth like-orientated ply for tape, and the splice stagger pattern

shall not be repeated for fabric.

Splicing Resin Transfer Molding (RTM) Laminates:- Splicing requirements for RTM fabric

and tape are given in company specifications:- Minimum stagger distance between splices are

for Fabric & Tape >= 12” (300mm) wide 2” (50.8mm) and for Tape < 12” (300mm) wide the

minimum stagger is 1” (20.4mm). The splice stagger pattern for both tape and fabric shall not

be repeated more often than every fifth like-oriented ply.

Reducing Splices With Bias Weave Fabric:- Splices may be minimized by substituting 45º

bias weave fabric for traditional, non-bias weave fabric, see figure 51 for an example of how

bias weave fabric can reduce the amount of splicing for some plies. However 45º bias weave

fabric is more costly than non-bias weave fabric and should only be used in special cases

where the added cost has been justified. These cases are typically where the minimum ply

dimension is less than the material roll width.

General Design Guidelines for Ply splicing.

87


Figure 51:- Example of Reducing Splice Task by Using Bias Weave Material.

45º

Warp Fiber Direction. Warp Fiber Direction.

Ply Boundary.

Ply Boundary.

Material Roll Width.

0º/ 90º Weave. 45º/ -45º Weave.

88


Honeycomb Core:- All composite / honeycomb sandwich structures shall utilize positive means

to prevent water intrusion into core areas. Core panels (metallic and non-metallic) shall seal

against water intrusion, and each panel will be checked for leaks before delivery for installation.

The designer shall include a fabric glass scrim ply between honeycomb core and structural plies

as shown in figures 52 and 53. The structural facesheets should be fabric. If tape is used in the

facesheet then the outermost structural plies and the plies adjacent to the core should be 45º

fabric. Each facesheet on a honeycomb panel shall be symmetric and balanced about the

facesheet mid-plane. The susceptibility of thin sandwich structures to FOD should be

considered in the design and appropriate actions should be taken to insure that such parts are

easy to repair and / or replace, especially when located in damage prone areas.

Syntactic Film Core:- Syntactic film is a low-density syntactic core material ordered at either

0.06 or 0.12-inch ( 1.5mm or 3.0mm) thickness as a core for sandwich construction. It is

moisture resistant, and co-curable with a wide variety of thermoset curing epoxy prepreg

systems. This type of core is a pliable film that can be cut or formed to the desired shape using

standard shop practices. Due to its tack, a small amount of pressure is all that is needed to

secure the edge of the film to the prepreg stack. The syntactic film is placed in the center of the

laminate ply stack-up as shown in figures 54(a) and (b). Fastener hole machining is prohibited

in portions of the laminate where this type of core is present, and the syntactic film shall not be

exposed at a trimmed edge.

General Design Guidelines for Core Stiffening.

89


90

Figure 52:- Honeycomb core insert build up structure.


91

Figure 53:- Honeycomb core transition configurations.

Tapered edges can lead to core

crushing issues requiring either a

reduced processing pressure or

friction grips external to the part to

minimise this 20º is design standard.

Ply/Core Edge Tolerance:- The ply and

core Edge Of Part (EOP) curves shall have

a line profile tolerance of 0.200”(±0.100”)

unless otherwise specified on engineering

drawing or other applicable document. Used

for structures les than 2.9” thick.


Syntactic Film Core (continued):- Syntactic film requires beveled edges, which are to be

machined or formed at a 5:1 taper with a 0.020” (0.5mm) offset at the edge. The corner radii

shall be no less than 1” (25.4mm) unless approved by stress analyst, the standard outside

radius is 3” (76.2mm) and should be used whenever possible. For improved damage tolerance,

a 45º fabric ply may be placed on either side of the syntactic film. The 45º fabric ply adjacent to

the syntactic film will provide a smoother stiffness transition between the film and the composite

laminate. Each facesheet on a syntactic film panel shall be symmetric and balanced about the

facesheet mid-plane.

92

General Design Guidelines for Core Stiffening.

Syntactic film

Figure 54(a) :- Syntactic film Pinch-off configuration. Figure 54(a) :- Syntactic film Arrowhead configuration.

Symmetry

plane

Syntactic film


In order to achieve lower cost production and hence wider aerospace application of composite

materials to commercial aircraft, and large military bomber and transports, lead to the

development of automated composite processing. Initially these were developed for the

Northrop B2 flying wing bomber in the 1980‟s, but since then these processes are now the most

widely used methods for large aircraft primary structures, as well as fighter aircraft examples

are:- fuselage components: empennage structures: fighter aircraft intake ducts: engine nacelle

components: wing structures and space launch vehicle components. The major manufacture of

the machines and developer of the processes is Cincinnati Lamb of the US. The two main types

of automated composite process machines covered here are Fibre Placement machines and

Tape laying machines and are shown in figure 55.

Fibre Placement:- This is a hybrid of filament winding and automated tape laying, the machine

configuration is similar to filament winding and the material form is similar to tape laying, this

computer controlled process uses a prepreg Tow or Slit material form to layup non-geodesic

shapes e.g. convex and concave surfaces, and enables in-place compaction of laminate,

however maximum cut angle and minimum tape width and minimum tape length impact on

design process .

Tape Laying:- Allows high deposition rates 10-15kg/hour, but has limited curvature +/- 15º and

maximum cut angle and minimum tape width and minimum tape length impact on design

process.

93

General Design Guidelines for automated composite processing.


94

Figure 55:- The two main automated composites processing methods available.

For complex curvature parts. For simple curvature or flat panel parts.

FIBRE PLACEMENT TAPE LAYING


95

Figure 56:- Limitations of Tape Laying compared to Fibre Placement.

56(a) Automated Tape Laying

56(b) Automated Fibre Placement

300mm tape with.

Manufacturing Edge Of Part.

6mm tape with.

Quasi Isotropic laminates require the plies to be laid up in the 0º: 45º: 90º: and 135º orientations,

and as the majority of ply orientations are in the 45º and 135º directions for automated tape laying a

large excess of waste material is generated as a triangle which over hangs the manufacturing edge

of part, as can be seen in figure 56(a), however this is significantly reduced when automated fibre

placement is used for the same laminate as shown in figure 56(b), so for such laminates fibre

placement is recommended to reduce material waste.


Minimum Course Length:- When generating ply shapes, the designer must consider the

shortest material length that the machine can lay down. This criterion is driven by the distance

between the compaction roller and the cutter. Different machines have different limits, so the

designer must design for the particular machine capability that will be used to manufacture the

part.

96

Fibre Placement Specific Design Guidelines based on design experience.

MINIMUM COURSE LENGTH

57(b) ALTERED DESIGN FOR FIBRE PLACEMENT

(NO HAND LAYUP REQUIRED.

45º COURSES

57(a) UNALTERED DESIGN.

Figure 55:- Minimum Tow Length.


Figure 57 on the previous slide indicates how ply design can be optimised for this minimum

course length. The corner regions in this example cannot be created by fibre placement

machine because the lengths are below limits so one design step could be to trim the ply

boundary as shown. Another option is to add material to the corners so that the minimum

course length is maintained, but this only works on exterior corners.

There are three techniques used to eliminate areas of missing tows:-

Exterior ply boundary extension past the required part shape, for example creating tabs on

45º plies which are subsequently trimmed back to the required part shape:

The reshaping of curved interior plies to match the fibre angles:

Re-distribution of holes to the full coverage plies having the same fibre angles.

Ply Edge Definition:- Each tow is cut perpendicular to its direction since individual tows can

be added or deleted, the edge can be any general shape. For edges not perpendicular to the

fibre direction the actual edge is a stair-step or “pinking shear” appearance. Design definition of

fibre place ply edges shows the “smooth” theoretical ply edges. The recommended practice is

to cut the material when the centerline of the intersects the theoretical ply boundary, this is

referred to as 50% overlap.

97

Fibre Placement Specific Design Guidelines (continued).


Gap / Overlap Allowance:- Individual tows are spaced based on nominal tow width to

achieve a set fibre areal weight and per-ply thickness. This tow spacing is not adjustable. The

finite width variation of the material and fixed spacing leads to the occurrence of small gaps

and overlaps between adjacent tows and bands. In addition, band convergence / divergence

due to part contour and fibre orientation definition leads to gaps / overlaps internal to plies.

The allowed gap / overlap values shall be included in the process specification for fibre

placement.

Surface Contour Capability:- Surface Geometry Limits are driven by several aspects of the

overall machine geometry, head and roller geometry, and the conformability of the roller

across the width of the material being placed. Generally convex tool geometry is more

producible than concave geometry, see figure 62. If male and female radii exist on a given

part the tighter radii should be made on the male features of the lay-up tool.

Fibre Placement Programming Methods:- The fibre placement process can use a variety of

programming approaches when building a laminate. However, the methods available for use

are dependent upon the type of fibre placement machine figures 58-60. The decision on

which fibre placement method to use is important and should therefore be made by a

representative IPT group at the beginning of the part sizing phase. The various methods

which can be employed and their advantages and disadvantages are discussed below.

98



Figure 58:- Large and Small VIPER fibre placement systems.

99

Cincinnati has supplied over 25 Large and Small VIPER Fibre Placement

Systems for r the USA, Europe, and the UK, over the past 14 years.

Figure 58(a) VIPER FPS-3000 Figure 58(b) VIPER FPS-1200


1) Natural Path Method:- The fibre is allowed to follow its natural course and is not steered

along a specific orientation. This method is only used on flat panels because contour would

generate gaps and overlaps between bands. This describes the path followed by tape and

fabric that are hand-laid over contoured surfaces.

2) Controlled Angle Method (Fixed Angle):- The paths follow a set fibre angle without

deviation. Tows will be either added or dropped to control angle deviation and create a

uniform ply thickness. Can be less efficient than parallel paths throughout a full ply

stack.

3) Band Off-Set Method (Parallel Paths):- A single guide path (guide band) is established for

each ply, which aligns with the true fibre orientation relative to a designated reference axis.

Allowable deviation over full path same as prepreg broadgoods. All remaining bands align

edge-to-edge with the guide band, which produces grater angle deviation the further it is away

due to change in surface curvature / area. Individual tow cuts and adds are not programmed

which produces all constant width bands. Can be more efficient than fixed paths if

resulting angular deviation is acceptable.

4) Controlled Offset Method:- This method is a hybrid of the controlled angle method and the

band offset method. It is used when inter-band tow dropping and adding is not desired, and

some degree of fibre angle compliance must be maintained. The ply filling process begins

with parallel paths as in (3), and moves to fixed paths if the pre-determined fibre angle

tolerance is exceeded.


100


Figure 59:- Typical fibre placement system components.

101

Fibre Placement

Head, Mounted to a

Roll-Bend-Roll

Wrist


Figure 60:- Details of the fibre placement system.

102

Fibre Placement Head

Fibre Placement System – VIPER 6000


5) Hoop Method:- This method is used when a continuous, tightly would helical path on a

closed surface is desired. The hoop method has the advantage of continuous fibres and

efficient machine lay-up for 90º plies, but it sacrifices fibre angle compliance. Also, if the hoop

path converges or diverges upon / from itself, tows are dropped.

Orientation Axis for Complex Contours:- For flat parts, a rosette is adequate to define a

reference axis system, but for complex parts as shown in figure 61 the fibre reference system

becomes more complex. Typical selection of the fibre reference system is based on primary

load paths, the ability to analyze changing fibre direction within a ply, and the design

allowable database for specified laminates. The fibre axis reference system greatly

influences the producibility of fibre placed parts. Establish reasonable fibre orientation

reference axis relative to the fibre placement surface, otherwise, the steering required may

exceed the physical limits of the tow / machine and cause degradation in part quality.

Radius Guidelines:- The recommended minimum corner outside (or male) radius for faceted

shapes is 0.5 inches (12.7mm), and for complex contours the minimum outside (or male)

recommended radius is 3.0 inches ( 76.2mm). The minimum fibre place-able inside (or

female) radius is 1.5 inches (38.1mm) depending on the roller size and conformability: ply

angle, bandwidth, fiber placement head envelope, surrounding geometry and radius location.

For shallow drops, without immediate reverse curvature, a 6 inch (152.4mm) radius is

recommended.


103


Figure 61:- Fibre placement applied to commercial aircraft.

Radius Guidelines (continued):- The recommended in plane radius (fibre steering) is 24

inches (609.6mm). Manufacturing engineering familiar with fiber placement or representatives of

manufactures of fiber placed parts should be consulted for additional information on radii.

104

Examples of VIPER Fibre Placement Systems applied to the Airbus A380 Aft Fuselage.


Figure 62:- Fibre placement tooling.

105


Figure 63:- Tape laying machine and tape head components.

106


Figure 64:- Contour Tape Laying machines.

107

Cincinnati has supplied over 36 Contour

Tape Layers for the USA, Europe, Japan,

UK, and Indonesia over the past 21 years.


Figure 65:- Flat Tape Laying machines.

108


Figure 66:- A330/A340 Contour Tape Laying applications.

109


Figure 67:- A330/A340 Flat Tape Laying applications.

110


In the proceeding slides have focused on hand and automated ply laminates, but to ensure build

tolerances are met in composite skin and spar joint assemblies sacrificial plies are used, for large

structures as an alternative to match tooling. The following is a brief outline of them and their design

requirements which will be applied in the design of composite structure in the FATA project.

As discussed above carbon fibre composites are fabricated using individual plies in orientations

defined by engineering to specific thicknesses in order to carry the design loads. Due to parent

material thickness variation for the raw material as well as those introduced as part of the post

layup cure process, the resulting laminate product will have varying thickness. Therefore in order

attain a specific thickness to aid assembly and meet aerodynamic OML mismatch requirements a

procedure has been adopted to predict the amount of variation expected in the structural laminate.

A sufficient amount of sacrificial plies are added to the laminate at the interface location to the

substructure to compensate for the expected variation. Finally, the thickness is machined to the

specific desired thickness without infringing into the structural plies.

In the fabrication of a laminate, a “buffer or waviness layer” is used to isolate the structural plies

from the sacrificial machining as shown in figure 68(a). This buffer or witness ply is designed to

provide a visual indicator to manufacturing of machining through the sacrificial plies and into the

structural plies. The specific buffer layer on the laminate is dependent on the laminate material and

will be issues in project guidelines. Considerations must also be given to laminate thickness

changes i.e. ramped ply-drop areas, and the locating accuracy of ply-drops must be compensated

with sacrificial plies in the footprint of the substructure. The assembly process of mating the skin to

the substructure adds the positioning accuracies of the locating holes to require a designed in gap

at these ply-drop ramps. 111

The design of composite sacrificial plies for assembly tolerance control.


112

Figure 68:- Sacrificial ply design to meet assembly requirements.

Figure 68(a):- CFC sacrificial incorporation in ply lay up to meet assembly tolerance.

Figure 68(b):- CFC laminate thickness constituents to meet assembly tolerance.


The following design details need to be considered prior to computation of the sacrificial ply

thickness (see figure 68(b)): -

1. Determine the buffer layer material thickness:- (a) Fiberglass ply scrim: (b) Adhesive, use cured

thickness or carrier thickness if any.

2. Determine the corrosion barrier thickness and type:- e.g. Fiberglass: Polysulfide with glass

carrier: or Polysulfide alone: Substrate and laminate Surface finish with faying sealant.

3. Determine which finish to apply:- Determine primer / paint to be applied to skin / door / cover

IML if the land is in a fuel bay: Apply secondarily bonded corrosion barrier if applicable and then

Paint / Primer after IML machining: Paint / Primer is added after IML machining or the corrosion

protection layer.

4. Determine other details in the laminate:- Determine land width to allow for ply drops in sacrificial

plies: Plan where there may and may not be overlaps in sacrificial or structural ply layers

(overlap splices will count as additional thickness in the laminate in local areas): Determine

Slopes for Ramps (recommended 10:1 minimum ramp for ply drop and 5:1 minimum ramp for

joggles): Determine land width to allow for ply drops in sacrificial plies.

The composite laminate and the MSP (machined sacrificial plies) have a Nominal thickness which

is used to calculate the laminate IML and the substructure OML surface (figure 69). Both the

laminate and the MSP also need a minimum “before-machined” thickness which compensates for

thickness and machining variation. The following two steps must be taken to determine the laminate

IML.

113

The design of composite sacrificial plies for assembly tolerance control (continued).


Step 1:- Determine the Total Laminate Thickness at the lands where the Composite skin is

attached to the substructure. Laminates within the substructure footprint must include an additional

layer of sacrificial plies to account for manufacturing and assembly tolerances. For constant

thickness laminates, the Total Laminate thickness = Structural Ply thickness + OML fibermat plies +

lightening strike ply + a Buffer ply and / or film thickness + Sacrificial Ply thickness + a corrosion

barrier (as applicable) + finish primer / paint.

Step 2:- Determine Ply Ramps. To avoid machining into the structural plies, design the ramp to be

machined in the maximum material condition (MMC) (+0.150” to +0.200”) location. However, if the

ramp exists in the least material condition (LMC) (-0.150” to -0.200”) location, there must be

sufficient sacrificial plies on the ramp to produce a machined ramp slope.

114

Ramp Offset Distance = (Ply Location Accuracy) /2+ Ply Drop

Depth x Tan (Slope). Example:- Ply Location Accuracy = 0.300”:

Ply Drop Depth = 13 x 0.0083 + 0.002 = 0.1099. Hence Ramp

Offset Distance = 0.300” / 2 + 0.1099” x 1/10 = 0.161”

N.B.:- If the plies are placed by hand with a ply

projector, location, ply projector and ply pack trim

tolerances must be accounted for.

Also note the thickness of sacrificial plies on a

constant laminate section will be less than the

thickness at the top of a ramp which has to account

for ply drop location accuracies.

Figure 69:- CFC laminate thickness constituents in a Taper Region.

Sacrificial Ply Thickness

Top of Ramp Thickness

Bottom of Ramp Thickness

Corrosion Protection

Sacrificial Ply Thickness

Machined IML

Buffer

(Witness)

Layer

Fibermat OML Layer

OML Surface

The design of composite sacrificial plies for assembly tolerance control (continued).


This section gives an overview of the materials and processing knowledge base required for the

design composite structural component. The properties data for design analysis is drawn from the

AIAA Aerospace Design Engineers Guide (4th edition), and RAeS Aeronautical Engineers Data

Book. The detail structural composite component design will follow the Catia V5.R20 CPD

methodologies of CU/CoA/AAO/1.

Fibre types see figures 70 and 71:- (1) Carbon Fibre used for primary aircraft structures,

combining low density with high specific strength and stiffness: (2) Glass Fibre of which three types

are in major use:- E-Glass low cost systems applications, R-Glass electromagnetic properties used

in radomes, S-Glass ballistic properties used as surface plies (see cover skin impact damage

protection) with low scrap rate and is applied where required: (3) Aramid Fibre of which there are

two types of interest:- Kevlar 49 used for ballistic protection and in fairings and panels, and Nomex

paper used for honeycomb cores.

Resin types see figure 72:- (1) Epoxy Resin easily processed, low cost and good performance

and can be used in 80% to 90% of airframe primary structural applications with an operational

temperature ceiling of 120°C, but suffers from environmental degradation with moisture and

temperature: (2) Bismaleimide (BMI) resin has higher operational temperature than epoxy resin i.e.

170°C to 230°C and provides high temperature operability for medium cost, however requires more

complex processing than epoxy resins: (3) Polymide Resin high temperature resin for engine case

applications, complex processing with long high temperature cure cycles, and health and safety

issues.

115

Section 3:- Composite component materials and processing overview.


Figure 70:- Composite Component Fibre Material Types.

116


117

Figure 71:- Composite Component Fibre Material Properties.


Figure 72:- Composite Component Resin Material Types.

118


Material forms available for section:- There are three main material forms of composite

material available which are as follows:-

Prepregs:- These consist of resin and fibre combined to form a ready to mould material form

and can be in the form with uni-directional fibres or woven cloths. A paper or polythene backing

material is applied in order to protect the material prior to moulding. The resin is in the “B” stage

cure state in order to hold it in place on the fibre, and as such prepregs have a limited shop life

at room temperature typically 5 to 31 days for an epoxy depending on resin type, and a fridge

life of 6 to 12 months depending on resin type at -18°C.

Preforms:- Dry fibre (fabric or NCF) held by a binder approximately 4% to 6% resin by weight

prior to conversion by Resin Transfer Moulding (RTM) / Vacuum Assisted Resin Transfer

Moulding (VARTM) or Resin Film Infusion (RFI).

Dryfibres:- Used for wet layups, and applications in non-structural repairs.

Unidirectional prepreg material comes in two classifications as shown in figure 73, which are:-

Broadgoods with a width greater than 300mm and Tape with a width less than 300mm.

119

Composite component materials and processing overview (continued).


Figure 73:- Composite Uni-Directional material classifications.

120


CFRP Post layup conversion processing methods studied.

The majority of aerospace composite parts with thermosetting matrices are cured at elevated

temperatures and pressures (conversion), to ensure that the service temperatures of the composite

is sufficiently high. As a typical example, a carbon / epoxy composite cured at 180°C for 2 hours

might have a glass transition temperature (Tg) of 200°C when dry, but only 160°C when saturated

with moisture. This would allow the composite to be used at a maximum service temperature of

around 135°C.

There are four major conversion processes used in industry which are as follows:- (1) Vacuum

assisted oven:- Used in mainly for repairs and adhesive bonding: (2) Autoclave processing:- The

most common method used for curing prepregs for primary aircraft structures (covered below): (3)

Resin Transfer Moulding:- And the related process Resin Infusion Moulding where dry preforms

are injected with resin in a heated matched tool in the former or half tool with caul plate in the latter.

These processes give good tolerance parts although with high tooling costs (covered below): (4)

Press Moulding:- Commonly used for the production of aircraft floor panels and other flat thin

skinned honeycomb panels.

This study will consider designing for Autoclave processing for the FATA baseline reference

aircraft and Resin Transfer Moulding / Resin Infusion Moulding for the FATA developed aircraft

and future concepts. For the purpose of this document an overview of these processes is given to

highlight the design for processing issues.

Autoclave processing:- The autoclave is basically a very large, internally heated pressure vessel,

with internal connections for vacuum hoses and sensors such as thermocouples, figure 74

illustrates the autoclave general internal arrangement, as well as layup bagging requirements. 121


122

Figure 74:- Basic autoclave design and component preparation for processing.

Figure 74(a):- A modern autoclave general layout

arrangement from ASC systems.

Figure 74(b):- Example of the size of modern autoclaves and tooling

i.e. A350 XWB fuselage skins.

Figure 74(c):- Example laminate / tooling bagging and support.


Autoclaves are computer controlled and pressurized with an inert gas either nitrogen or carbon

dioxide to reduce the risk of an internal fire. A standard autoclave for epoxy composites would be

capable of temperatures of more than 200°C and pressures of 700 kPa, where as an autoclave for

processing thermoplastic composites or high temperature thermosets may be capable of 400°C

and 1200 kPa or more. The part is normally heated by convection of heat from the fan-forced air

circulation, however electrically heated mould tools can be used. Although the latter tools are more

expensive there are several advantages in heated tooling which include more rapid and uniform

heating, and the ability to use higher temperatures without heating the walls of the autoclave.

Normally the layup or part is under vacuum from the time it leaves the layup room and while it is

loaded into the autoclave, to keep the layup in position and help remove air and volatiles. The

vacuum and sensor connections are checked before the tool / bagged layup are sealed in the

autoclave and the cycle commences, figure 75 shows the steps of a part through autoclave

processing from layup profile to inspection. Figure 75 also shows a basic autoclave temperature /

pressure / time profile, and this is a generalized example of the process, as soon as the autoclave

is sealed the pressurization and heating cycle begins, and the target pressure is reached in 30

minutes, whereas in thick layups the target temperature may not be reached for several hours.

After more than 100kPa (gauge) pressure is reached in the autoclave, the space under the vacuum

bag is vented (connected to the atmosphere) to eliminate bubbles from entrapped gases and

volatiles, in the resin as the part is heated. Heat-up and cool-down rates are controlled to ensure

even curing throughout the part and to reduce the possibility of residual stresses causing structural

deficiencies or distortions.

123

CFRP Post layup conversion processing methods studied (continued).


Figure 75:- Basic outline of the autoclave processing from laying to inspection.

124 TIME

TE

MP

ER

AT

UR

E (

°C) P

RE

SS

UR

E (k

Pa) TEMERATURE

PRESSURE

TYPICAL AUTOCLAVE CYCLE WITH

TEMPERATURE DWELL.


The viscosity of the resin falls with increasing temperature until the resin begins to chemically

cross-link (gel). It is important to maintain full pressure up to and throughout gelation process to

allow the expulsion of entrapped gases and the removal of excess resin from the layup. Under

certain circumstances a dwell is incorporated (isothermal hold), as shown in autoclave cycle chart

in figure 75, to prolong the time for consolidation and volatile removal. The hold also pre-reacts the

resin and reduces the danger of large damaging exothermic reactions that can occur in thick

laminates, e.g. over 50 plies thick. A hold will also allow the temperature to become more uniform

which is very important in components with large variations in thickness.

The requirement for complex heating / pressure cycles is important when using less viscous epoxy

resins and high-temperature resins because they are required to accommodate the requirements of

the chemical reactions and to ensure that resin viscosity is at its optimum state when the pressure

is increased. Most modern non-bleed epoxy prepregs, however can be processed with a simple

“straight-up” cure cycle, provided that the component is not too thick or complex.

When co-curing or co-bonding complex components internal conformal tooling is required, and in

some cases internal pressurization if silicon rubber pressure bagging is used this is covered in

workbook 1 and reference 1.

Processing problems:- The main processing problems encountered in autoclave processing which I

will be designing against include:- overheating (caused by excessive exothermic reactions):

porosity: resin–rich areas: resin-dry areas, poor surface finish, insufficient consolidation, uneven

cure, and distortion. Many of these problems can be resolved by correct timing of the application of

temperature and pressure, and the use of prepregs with low exothermic cures. 125



The formation of voids is generally caused by the entrapment of volatiles, water, and air that have

remained after debulking. At the high processing temperatures in the autoclave, more solvents are

liberated, and the volume of the solvents and other entrapped gases increases. To avoid the

formation of severe porosity, it is necessary that the hydrostatic pressure in the resin before

gelation exceeds the partial pressure of the gases, allowing them to be expelled. Once the resin

cures (gels) no further void removal or consolidation is possible. Water is considered to be the main

cause of void formation so that the applied pressure needs to exceed the partial pressure of water.

While a low temperature hold is often used to increase the time at low resin viscosity for the

reasons stated above, excessive pressure or over-efficient resin bleed when the resin is in a low

viscosity state can lead to dry zones. Resin rich areas on the other hand result from areas of the

layup have lower resistance to resin flow and insufficient pressure is applied before gelation.

To reduce porosity, a surfacing resin film or fine class / epoxy scrim ply is usually placed on the

mould surface before the prepreg is placed. The part should be smooth on the tool side, but unless

matched moulds are used there will be some texture or roughness on the bag side of the part,

however this can be minimized if a stiff caul plate is used. Due to the variations in prepreg fibre

areal weight and resin content, and resin bleed during curing, it is difficult to specify the thickness of

a prepreg part to less than +/- 5%, which is a serious concern in thicker parts such as wing cover

skins, where the choice would be between having a smooth outside surface with the correct

aerodynamic contour (Outer Mould Line tooling), and controlling the inner surface dimensions

(Inner Mould Line tooling) to allow easy assembly to the substructure. This has to be resolved by

OML tooling, with sacrificial inner mould line plies, and shims, as detailed in WB1, as also are part

spring in and honeycomb core crushing issues. 126



Resin Transfer Moulding processing:- The resin transfer moulding process shown below in

figure 76 involves first placing the dry fabric preform into the cavity of a matched mould and then

filling the mould and thereby the preform with liquid resin. The mould and the resin being preheated

before injection. After injection, the mould temperature is increased to cure the part. In some cases

the resin can be injected into a mould that has been preheated to the cure temperature. The resin

preheat, injection time, and mould temperatures being determined by the characteristics of the

resin system selected. If the temperature is too high, the resin will gel before the mould is filled,

conversely if the temperature is too low, the viscosity may be too high to permit flow through the

preform. A vacuum is typically applied at the exit port to evacuate air and any moisture from the

mould / preform before resin injection, and injection pressures of around 700 kPa are usual. The

application of a vacuum during injection is useful in order to prevent void entrapment, and as a

supplement to the injection pressure, however care must be taken to ensure that the resin injection

temperature is not above the resins vacuum boiling point as this would result in unacceptable

porosity. When high injection pressures are used, there is a possibility of fibre – wash (i.e.

reinforcement distortion) exists. Loose weaves and unidirectional plies will have a greater tendency

to fibre-wash than tightly woven preforms, such as plane weaves. Additionally, high injection

pressures will cause an increase in resin flow speed between tows, without complete fibre wetting,

resulting in voids within tow bundles, alternatively if the pressure is too low it can also result in voids

between tows.

A large range of resins can be used for RTM, including polyesters, vinyl esters, epoxies,

bismaleimides (BMI‟s), phenolics, and cyanate esters.

127



128

Figure 76:- Basic outline of the Resin Transfer Moulding (RTM) process.


Vacuum Assisted Resin Transfer Moulding:- The Vacuum Assisted RTM process is a single-

sided tooling process, and involves laying a dry fibre preform onto a mould, then placing a

permeable membrane on top of the preform, and finally vacuum bagging the assembly. Inlet and

exit feed tubes are positioned through the bag, and a vacuum is pulled at the exit to infuse the

preform. The resin will quickly flow trough the permeable material across the surface, resulting in a

combination of in-plane and through thickness flow and allowing rapid infusion times. The

permeable material is usually a large open area woven cloth or plastic grid. Commercial “shade-

cloth” is often used for this process. In foam cored sandwich structures, the resin can be

transported through grooves and holes machined in the core, eliminating the need for other

distribution media. The VARTM process results in lower fibre / volume fractions than RTM because

the preform is subjected to vacuum compaction only. However for the PRSEUS process this is

addressed by stitching the preform before layup as shown in figure 77(a), and in additional soft

tooling (bagging aides) are also used figure 77(b) and in the Boeing Controlled Atmospheric

Pressure Resin Infusion process figure 77(c), resin infusion takes place in a walk in oven at 60°C,

and following injection the assembly is then cured at 93°C for five hours, and then finally with the

vacuum bag removed post cured for two hours at 176°C with a final CNC machining to remove

excess material. The full process is documented in NASA/CR-2011-216880. The main advantages

of the CAPRI process over conventional VARTM is increased performance for airframe standard

parts, and over RTM reduced tooling costs and production of larger components, and over

conventional processing the elimination of a specialist autoclave. The full process and

manufacturability of large airframe components by this process will be a major focus of the FATA

project. 129



130

Figure 77:- Boeing Controlled Atmospheric Pressure Resin Infusion (CAPRI) process.

Fig 77(b):- Soft tooling (bagging aids) installation over stiffeners.

Fig 77(a):- Robotic stitching of dry preform assembly.

Fig 77(c):- Vacuum bag installation over dry preform assembly.


This section gives an overview of conversion process component tooling which is essential for the

design of producible CFRP laminate components.

In the selection of tooling materials the designer must attempt to match the thermal expansion in

the tool to the coefficient of thermal expansion (CTE) of the laminate component to be produced on

the tool. When both the tool and the laminate are subjected to elevated temperatures for forming

and consolidation as in autoclave processing, invar, steel, carbon (graphite), or ceramic tooling

materials must be used. In forming operations where only the laminate is heated and then pressed

into cold tooling, a range of materials can be used, such as aluminium, MDF, wood, rubber and

silicone. Table 7 gives a guide to the selection of tooling materials and is taken from reference 2.

The requirements for composite laminate tooling differs significantly those of metallic (sheet) tooling

in the following aspects:-

Tolerance build-up is much more critical:

The final machined dimensions of the tool are not necessarily the final dimensions of the

composite part and the degree of disparity is dependant on :- the type of tooling and the CTE

characteristics:

Final part dimensions are those present at the ultimate gelation temperature of the matrix

system.

There are currently no definitive rules to specify tooling selection figure 78 offers some guidelines to

the most cost effective choice of tooling, but this is still an area of much research and development

activity. Table 8 gives a rating of tooling priorities (factor 1 being the lowest and factor 5 being the

highest). 131

Section 4:- CFRP Post layup conversion processing tooling.


Tooling

Material.

Coefficient of

Thermal

Expansion.

Heat

Conductivity. Material Cost.

Fabrication

Cost. Durability.

Aluminium. Poor. Good. Good Fair. Fair.

Steel. Good. Good. Good. Poor. Respectable.

Graphite. Excellent. Good. Good. Good. Poor.

Ceramics. Excellent. Poor. Good. Fair. Fair.

Fibreglass Resin

Composites. Poor to Good. Fair. Good. Good Poor.

Graphite Epoxy

Composites. Excellent. Fair. High. Fair. Poor.

132

Table 7:- CFRP tooling material guidelines.


133

Fig 78 (a):- Male Mould Tooling characteristics.

Most commonly used for aircraft parts because of its low cost:

Lowest layup cost:

Small radius producibility > 0.05 inch (1.27mm):

Baseline (non-aerodynamic surfaces):

Surface control one side only:

Localised control from vacuum bag surface. (a) Male mould tool

(b) Female mould tool

(c) Matched die mould

Fig 78 (b):- Female Mould Tooling characteristics.

Limited use in contour applications because of bend radius:

High layup cost:

Radius producibility > 0.25 inch (6.35mm):

Surface control one side only:

Localised control from vacuum bag surface.

Fig 78 (c):- Matched Die Mould Tooling characteristics.

Used male / female tooling to control laminate thickness and is very

expensive:

Best thickness control:

Highest tooling cost:

Moderate layup costs:

OML / IML control (smooth surface both sides).

Figure 78:- Mould Tooling Types and Characteristics.


Tooling Properties. Importance Factor. Tooling Properties. Importance Factor.

Dimensional accuracy. 5 Ease of tool fabrication. 3

Dimensional stability. 5 Ease of repair. 3

Durability. 5 Tool weight. 3

Thermal mass. 4 Ease of inspection. 2

Surface finish. 4 Resistance to handling

damage. 2

Ease of reproducibility. 4 Ease of thermocouple

implantation. 2

Temperature uniformity. 4 Release agent compatibility. 1

Material cost. 3 Sealant compatibility. 1

134

Table 8:- CFRP tooling properties rating factors guide.


Tooling solutions for composite parts cover a very wide field and are based on requirements,

schedules, costs, incentives, and other such issues, table 9 gives the pros and cons of the most

commonly used tooling types and is intended as a guideline when considering tooling and

component design.

The general requirements for composite tooling are very similar to those for sheet metal fabrication

dies or compression moulding dies:- Tool contact with the deforming material should occur in such

a way that the sheet surface pressure is uniform at all times: In geometries where this is not

possible e.g. where the loading direction is perpendicular to the surface, flexible tool halves should

be used to provide a type of hydrostatic pressure: Normally tools should be designed for a draft

angle of 1º to 2º to counteract the effects of “closure” or “spring-in” after cure to facilitate ease of

part removal from the tooling.

The ideal tooling requirements are listed below:-

CTE characteristics compatible with the composite part material:

Ability to withstand sever temperature and pressure conditions without deterioration:

Dimensional stability:

Relatively low cost:

Reproduce pattern with high dimensional accuracy:

Retain mechanical properties at high temperatures.

135

CFRP Post layup conversion processing tooling (continued).


Type. Potential application. Pro. Con.

Autoclave.

Large components:

Low volume production:

Honeycomb sandwich

assemblies:

Co-cured parts:

Parts having vertical walls:

Bonding.

Low cost:

Internal heating possible:

Undercut feasible:

Vertical walls attainable:

Versatile:

Complex co-cured parts:

Thermal expansion can be

made to match each part.

Low production rates:

High labour costs due to ancillary material layup

see fig 73:

Loose dimensional control of bag surface:

Low moulding pressure, relative to matched

dies, requiring more generous radii:

Curing temperatures limited by ancillary

materials unless internally headed tools are

used with insulation installed between bag and

layup:

More process variables involved than with

matched dies:

Bag failure usually causes the part to be

scrapped.

Matched metal dies.

Relatively small parts:

Both surfaces dimensionally

controlled.

High productivity:

Good dimensional control:

High moulding temperatures:

Good quality surfaces on all

faces:

High fabrication pressures:

Durable:

Internal heating feasible:

Good thermal response and

control:

Compression moulding tool

technology available:

Minimising ancillary material

use.

High cost due to machining, stops guides etc.:

Tool thermal expansion different from

composite:

Limited ability to selectively reinforce the tool:

Undercuts require multi part tooling:

Draft angles required where vertical wall

preclude part removal from tool:

Large components present tool flexibility,

heating uniformity, and air / volatile removal

from tool difficulties:

Difficult to repair or modify:

136

Table 9:- Summary of CFRP tooling types pros and cons.


Type. Potential application. Pro. Con.

Elastomeric.

Allows complex geometries:

Large components feasible.

Considerable part design flexibility:

More complex parts feasible than with

matched metal dies due to casting of

elastomeric elements:

Ability to layup on numerous

elastomeric mandrels and install these

in the metal tools allows complex, parts

to me produced.

Limited life:

Volatile and air removal less

than ideal:

500ºF / 260ºC processing limit:

Low conductivity of elastomeric

elements can cause undesirable

thermal gradients in the part.

Monolithic Graphite.

Tight dimensional control of

complex components:

Rapid cure cycles (high heat up

rates):

Prototype parts.

Low CTE matched to graphite fibre

composites:

Temperature capability 600ºF / 316ºC:

Lower cost than metals:

Easily machined in specialised facility:

High thermal conductivity:

Easy part release:

Easy to repair or modify.

Susceptibility to impact damage:

Special precautions needed

when machining:

Not suited for matched die

moulding.

Ceramics

Tight dimensional control of high

temperature components.

Can be cast into complex shapes:

Low CTE which can be controlled :

Electric and fluid heating systems can

easily be cast into the tool:

Temperature capability 600º / 316ºC.

Susceptible to impact damage:

Difficult to repair.

137

Table 9:- Summary of CFRP tooling types pros and cons (continued).


When the complexity of a composite component indicates serious tooling issues, design

modification should be sought early in the development phase to allow mutually acceptable

compromises to be reached between design and manufacturing teams, this is best implemented by

the use of integrated project teams for product development in which all parties are represented

and work collectively. However, after production tools have been made it is almost always too late

(due to both schedule and cost) to make major changes to the design. There is significant cross

over in the basic information on the design of metallic fabrication tools and the design of tools for

fabricating composite components. A highly detailed discussion of tooling design is given in both

references 1 and 2 and will not be reproduced here.

The following is a set of manufacturing guidelines and practices applied to the FATA study

composite structural component design and to the main exercise components in this presentation.

1) Size limitations:- The available facilities impose constraints on the size of composite

components and assemblies for example:- autoclave size (diameter and / or length), or die

size: the ability to provide the required time – temperature – pressure throughout the cure cycle

over the whole part: or the consolidation cycle requirements for the particular matrix of the

composite system selected. For automated systems the planform size may be limited by the lay

up area capability or the limits of the tape laying or fibre placement machine (see section 2),

and filament wound components are limited by the size of the winder and mandrel.

Hand layup process size is unlimited but may require a special facility:

Out of autoclave methods should be considered for processing large components:

Automated equipment should be used for large components against high cost hand layup. 138



2) Complexity of the composite component shape:- The shape limitations during construction

result from the drape quality of the tape and fabric selected. Tight radii and abrupt changes in

surface features should be avoided as they result in bridging between plies. Attention is paid to

the following in the examples and the FATA component design:-

Fabrics are generally easier to form than unidirectional plies:

Shape limitations are the result of the materials drape capability:

Changes in surface contours should be avoided where possible:

Shape restrictions are also a function of access requirements for layup machine heads and

tooling:

Graphite / epoxy thermoset prepreg material has a drape and tack which allows the

fabrication of components with grater changes in contour than is possible with thermoplastics:

Thermoplastic prepregs require different methods such as thermoforming to fabricate

contoured parts.

3) Tolerance:- Within practical limits tolerances should always be as large as applicable to the

function of the component will allow. Length and width tolerances for composite parts should be

kept to the same standards as metal production component counterparts, and where very high

tolerances are required sacrificial plies can be employed (see section 2). However each

material and method of fabrication will produce parts with some thickness variation.

Normal cured or consolidated ply thickness:- 0.00xx ± 0.0003 inch (0.0076mm) for tape and

0.0xxx ± 0.0012 inch (0.0305mm) for fabrics:

139



Nominal cured thickness should be taken as reference:

Factors affecting component thickness are as follows:- amount of resin bleed during the cure

cycle: cure pressure: tooling type (matched die or bagged curing): resin content:

Controlled thickness without sacrificial plies is very difficult to achieve except by expensive

matched die tooling:

Male or female tool selection affects tolerances as shown in figure 79:

Fastener grips should be specified for maximum laminate thickness:

Structural joint pull - up or nesting parts should be analysed for min / max tolerance

conditions and shims provided as required:

Figure 79 shows tolerance requirements for general composite airframe structure

applications.

4) Surface smoothness and flatness (see figure 80):- A smooth surface is required on surfaces of

composite aircraft skins and components either: - for aerodynamic considerations: to ensure an

adequate roughness for keying in adhesive bonds or painting. The required surface condition

(smoothness and / or flatness) can be achieved through:-

Specifying the tool surface side of the composite laminate:

Specifying requirements for a defined area:

Consideration of laminate variation with: - material: assembly method: and selection of tool

side for layup and cure:

140



141

Figure 78:- Effects of Mould Tool Selection on Tolerance.

Tool surface

± 0.030

(a) Male Tool.

(b) Female Tool.

Tool surface

± 0.030


Selection of outer plies, where aerodynamically smooth surfaces are more readily produced

by using tape outer plies rather than fabric:

The fabrication process does not insure mating surface flatness on the bag face:

Specifying requirements for a defined area (common manufacturing option is to employ caul

plates).

5) Laminate thickness, in most cases the control of laminate thickness via the drawing to

tolerances greater than the manufacturing specifications is not required.

6) Drilling and countersink of mechanical fastener holes in carbon or graphite thermoset laminates

requires the use of carbide tools and the application of glass ply outer layer to control drill

breakout, and Kevlar and thermoplastic composite laminates generally require special drilling

procedures as well as tools.

7) Engineering drawing face data requirements for components:- Ply Rosette: Stagger Index: Ply

Profiles: Lay-up Datum: Honeycomb Core: Profile: Ribbon Direction.

8) Additional Engineering drawing face requirements for assemblies:- Lay-up Table: Assembly

Details: Notes.

9) Material selection:- The most desirable material form is one which meets the required strength,

and allows net shape forming in a matter of minutes.

142



143

Figure 79:- General Tolerance Requirements for Airframe Laminates.

XXX

REF

XXX

Requires matched die tool to achieve this dimension.

PLY

drop-off


144

Figure 80:- Methods of achieving surface smoothness and flatness.

Vacuum bag.

Laminate part.

Rough surface.

Smooth surface.

Tool.

Figure 80(a) Smooth surface on tool side only.

Figure 80(b) Smooth surface on both sides using Caul plate.

Laminate part. Vacuum bag.

Smooth surfaces.

Tool.

Caul plate.


145

Figure 80:- Methods of achieving surface smoothness and flatness (continued).

„FILM‟ ADHESIVE

(BSL.322)

„CLEAVAGE‟ FILLED WITH

UN-CURED CFC WEDGE

RELEASE AGENT

PRE-CURED

CFC SKINS

UN-CURED „Z‟ & „C‟

SPAR ELEMENTS


SPAR ELEMENTS

CONFORMABLE TOOLING SHOWN AS:-

Assembly Tooling.

Figure 80(b) Smooth surface on both sides using Pre-cured skins.


10) Symmetrical balanced laminate about the Neutral Axis as described in Section 1 is essential in

order to eliminate warping during cure and consolidation, and extension – bending coupling.

The symmetrically balanced laminate has these characteristics:-

Symmetrical about the neutral axis (NA) which can be either a datum plane or a ply layer:

Ply drop-offs are also symmetrical about the neutral axis (NA):

Each 45º ply is balanced with a 135º ply (-45º):

On laminate outer surfaces either 45º or 135º (-45º) are used as the final plies:

No more than six plies of the same orientation are grouped together:

Groups of 90º plies are avoided:

Where the laminate thickness exceeds 0.762mm (0.030 inch) with a ply thickness of

0.127mm (0.005 inches), adjacent ply angles will not differ more than 60º therefore there are

no combinations of 0º and 90º or 45º and 135º plies:

There is a minimum of 10% of the plies in each direction:

The minimum number of plies for a basic laminate is seven plies.

11) The minimum bend radius for internal corner radii of channels are important because, sharp

corners result in bridging and /or wrinkling of the prepreg and resin rich internal pools, thus

weakening the part, and sharp also result in high internal stress under bending loads which can

lead to premature failure therefore the designer shall make the internal radii as large as

practical within the guidelines in section 2.

146



12) Bend angle Spring Back see figure 81. The degree of spring back and the resulting variation

between the part and the tool depends on the following factors:- resin type: degree of

debulking: fibre orientation: part thickness. This spring back occurs because of the differential

expansion between the inner and outer surfaces of the part. Consequently this is taken into

account in tool configuration design, and female tooling is designed with slightly opened angle

as shown in figure 81. Currently there is much research on improving predictive methods for

calculating the degree of spring back by GKN Aerospace and others in order to improve tooling

and part design.

13) Laminate thickness dimensions are referenced (unless specifically required for fit up or mating

on assembly) to avoid excessive clamp-up force damaging the parts, because there is no

plastic deformation in composites as there is in metals hence applied clamping loads are not

redistributed. The final thickness of a laminate includes a cumulative tolerance based on the

number of structural plies used in the laminate.

14) Material placement is one of the key aspects of composite manufacturing and the following

should be considered in tool selection and design:-

Plies are laid on a convex tool surface and as the tool expands during the cure cycle, the

plies which do not expand, must slide relative to the tool, especially those plies adjacent to

the tool. This presents no problem for cloth over a one dimensional curve but for sharp two

dimensional curves cloth may tend to change orientation to take up a relatively strain free

state. During compaction the outer plies are being forced toward the hard tool face and

wrinkling may occur. The thicker the laminate, greater care in layup design is required to

prevent wrinkling. 147



148

Figure 81:- Female tooling to compensate for spring back.

2º 2º

90º 90º

Tool

Part

Figure 81(a):-Angled female tool to accommodate spring back.

Figure 81(d):-Post-cure.

Tool

Part

90º 92º

Figure 81(c):-Pre-cure.

Tool

Part

90º

2º

Figure 81(b):- Cured part in

unmodified tooling.

Part

Tool


For plies laid up on a concave tool surface, the layup must allow for the tool moving away

from the plies during cure or consolidation. The compaction of plies furthest from the tool

face requires that all plies move in the same direction to compensate for tool thermal

expansion. Failure to allow for this is a primary cause of bridging.

When plies are laid up over a sharply curved tool surface, in the case of unidirectional tape

care must be taken due to its relative weakness normal to fibres. Movement of plies laterally

in that weak direction can easily result in tape failure during cure or consolidation resulting

in strength degradation.

15) Interlaminate slip issues, when using laminates on a complex contour allowance should be

made for slippage between plies when forming radiused parts overlapping adjacent curvature

zones to prevent wrinkling as shown in figure 82.

16) Additional or drop off plies (ADP) should be symmetrical and balanced about the NA with a

distance between ADP steps of 3.18 – 6.35mm minimum (0.15 – 0.25inch), and a slope angle

no greater than 10º. ADP should not involve more 6 to 8 plies based on ply thickness of

0.127mm (0.005inch) or 2 to 3 plies for thicker plies, and should not be positioned on the outer

surface of a laminate to avoid this risk of peeling. All ADP‟s should be covered with at least one

continuous outer ply in order to aid load redistribution, and prevent edge delamination. All pre-

cured or consolidated inserts should be compatible with the ADP guidelines.

149



150

Figure 82:- Slippage allowance design.

Wrinkle here.

Wrinkle here. Figure 82(a):- Without slippage design.

Figure 82(b):- With slippage design.

Lap joint allows

for slippage.


The methodology of assembly of the complete structure is as important as the manufacturing

process selection in the design of structural components and their analysis, it is important to

consider the advantages and disadvantages of both bolted and bonded construction methods and

the impact of corrosion on composite assemblies.

The advantages of bolted assembly are:-

1) Reduced surface preparation:

2) Ability to disassemble the structure for repair:

3) Ease of inspection.

The disadvantages of bolted assembly are:-

1) High stress concentrations:

2) Weight penalties incurred by ply build ups, and fasteners:

3) Cost and time in producing the bolt holes, and inspection for delamination's:

4) Assembly time.

Corresponding issues for bonded assembly are set out below.

The advantages of bonded assembly are:-

1) Low stress concentrations:

2) Small weight penalty:

3) Aerodynamically smooth.

151

Section 5:- Composite structural assembly joint design and corrosion.


Composite structural assembly joint design and corrosion (continued).

The disadvantages of bonded assembly are:-

1) Disassembly, in most cases some part of a bonded structural assembly will need to be bolted

instead of bonded to permit access for repair and inspection. An example is the Typhoon

wing structure where the bottom skin is co-bonded to the structural spars, and top skin is

bolted to the same spars, permitting access from one side:

2) Surface preparation, and bond line inspection for porosity even in co-bonded joints using C-

scan ultrasonic inspection, resulting increased costs and time:

3) Need to design for bolted repair access:

4) Environmental degradation due to water absorption leading to degradation in hot / wet

condition, solvent attack:

5) Need for increased qualification testing effort to establish design allowables.

In the case of the vertical tail subject of this exercise bolted construction was selected primarily

because of the requirement to quickly, inspect, repair, or replace damaged structural components

within a first line servicing environment. For the purpose of this exercise the external skin

formation light bolted installation is omitted to reduce complexity of the design and for ITAR. In

the vertical tail component and assembly models bolt datum positions are shown as points and

vectors, as was the case for the real article.

152


Co-Curing:- This is generally considered to be the primary joining method for joining composite

components the joint is achieved by the fusion of the resin system where two (or more) uncured

parts are joined together during an autoclave cure cycle. This method minimises the risk of

bondline contamination generally attributed to post curing operations and poor surface

preparation. But can require complex internal conformal tooling for component support.

Co-Bonding:- The joint is achieved by curing an adhesive layer added between a co-cured

laminate and one or more un-cured details. This also requires conformal tooling as shown in

figures 83 and 84, and as with co-curing the bond is formed during the autoclave cycle, this

method was used on Eurofighter Typhoon wing spars which were co-bonded to the lower wing

cover skins, and proposed for the F-35B VT lower skin stringers in SWAT trade studies. Care

must taken to ensure the cleanliness of the pre-cured laminate during assembly prior to the

bonding process.

Secondary Bonding:- This process involves the joining of two or more pre-cured detail parts to

form an assembly. The process is dependent upon the cleaning of the mating faces (which will

have undergone NDT inspection and machining operations). The variability of a secondary

bonded joint is further compounded where „two part mix paste adhesives‟ are employed.

Generally speaking, this is not a recommended process for use primary structural applications.

Design considerations for adhesive bonded joints.

153


154

Figure 83:- Fighter aircraft Co-Bonded composite spar assembly.


„FILM‟ ADHESIVE

(BSL.322)

„CLEAVAGE‟ FILLED WITH

UN-CURED CFC WEDGE

RELEASE AGENT

PRE-CURED

CFC SKINS


SPAR ELEMENTS


SPAR ELEMENTS

CONFORMABLE TOOLING SHOWN THUS:

Figure 84:- Co-Bonded composite spar bonding and installation bonding to wing skin.

155


Composite bolted joint design rules:-

1) Design for bolt bearing mode of failure:

2) Counter sink (CSK) depth should not exceed 2/3 of the laminate thinness if required fill

laminate artificially with syntactic core (if design rules permit e.g. not permitted for USN, or

USMC):

3) Minimum bolt pitch is 4D for sealed structures such as fuel tanks, and 6D for non sealed

structures (where D is the bolt diameter):

4) Use only Titanium alloy or stainless steel fasteners to minimise corrosion risk:

5) Use a single row of fasteners for non sealed structures and a double row for sealed

structures such as fuel tanks see figure 86 next slide:

6) Minimum fastener edge distances are:-

3D in the direction of the principal load path see figure 85:

2.5D transverse to the principal load path see figure 85:

156

Composite structural assembly joint design and corrosion (continued).

Figure 85:- Fastener edge distances.

2.5xD 3.0xD

4.0 x D


157

Figure 86:- Corrosion / leek prevention methods for carbon fibre structures.


158

FASTENER

MATERIAL / COATING COMPATABILITY

• Monel. Marginally acceptable.

• Alloy Steel.

• Silver Plating.

• Nickel Plating.

• Chromium Plating.

Excellent compatibility and are

recommended for use in CFC structures

• Cadmium Plating.

• Zinc Plating.

• Aluminium Coating.

Not compatible, and will deteriorate rapidly

when in intimate contact with CFC.

• Titanium Alloy.

• Corrosion Resistant Steel.

Excellent compatibility and are

recommended for use in CFC structures

• Al. Alloys.

• Magnesium Alloys.

Not compatible

Not compatible

Table 10:- Galvanic compatibility of fastener materials and coatings.


159

The use of carbon composites in conjunction with metallic materials is a critical design

factor :-

Improper interfacing can cause serious corrosion :

Problem for metals e.g. Fasteners see table 6 above:

This corrosion problem is due to the difference in electrical potential between some of the

materials widely employed in the aircraft industry, and carbon:

When in contact with carbon and in the presence of moisture (electrolyte), anodic materials

will corrode sacrificially (galvanic corrosion).

Corrosion prevention methods:-

1) Prevent moisture ingress:

2) Prevent electrical contact carbon / metal:

3) Anodise aluminium parts:

4) Seal in accordance with project specifications:

5) Protective ply of inert cloth (glass) between contact surfaces extending 1” beyond edge on

metal part, and protective sealant (Polysulphide) „Interfay‟ see figure 87 on next slide.

Corrosion due to the galvanic compatibility of materials and coatings.


160

Figure 87:- Corrosion prevention methods for carbon fibre structures.

EPOXIDE PRIMER (15 to 25 Microns THICK)*

ANODIC TREATMENT*

Pu. VARNISH or EPOXIDE PAINT FINISH (ONE COAT)*

Al ALLOY COMPONENT

POLYSULPHIDE „INTERFAY‟ SELANT

EPOXIDE PRIMER**

GRP (As required as a „Drill

Breakout‟ material.)**

CARBON FIBRE COMPOSITE

* = Applied over the entire Al component.

** = Applied over the entire CFC

component – or a minimum of 5mm

beyond the contact area.


161

1) Stress concentrations exert a dominant influence on the magnitude of the allowable design

tensile stresses. Generally, only 20-50% of the basic laminate ultimate tensile strength is

developed in a mechanical joint:

2) Mechanically fastened joints should be designed so that the critical failure mode is in

bearing, rather than shear out or tension, so that catastrophic failure is prevented. To

achieve this an edge distance to fastener diameter ratio (e/D), and a side distance to

fastener diameter ratio (s/D) relatively greater than those for metallic materials is required,

(see figure 85 above). At relatively low e/D and s/D ratios, failure of the joint occurs in shear

out at the ends, or in tension at the net section. Considerable concentration of stress

develops at the hole, and the average stresses at the net section at failure are but a fraction

of the basic tensile strength of the laminate:

3) Multiple rows of fasteners are recommended for unsymmetrical joints, such as shear lap

joints, to minimize bending induced by eccentric loading:

4) Local reinforcement of unsymmetrical joints by arbitrarily increasing laminate thickness is to

be avoided because the resulting eccentricity can give rise to greater bending stress which

negates the increase in material thickness:

5) Since stress concentrations and eccentricity effects cannot be calculated with a consistent

degree of accuracy, it is advisable to verify all critical joint designs by testing of a

representative sample joint.

Composite structural mechanically fastened joint design guidelines.


162

6) If a laminate is dominated by 0° fibers with few 90° fibers it is most likely to fail by shear out,

unlike metals, in which shear out resistance can be increased by placing the hole further

from the edge, laminates are weakened by fastener holes regardless of distance from the

edge. Reinforcing plies at 90° to the load direction helps prevent both shear out and

cleavage failures: Use larger fastener edge distances than with aluminum design, e.g. e/D

>3: Use a minimum of 40% of ± 45° plies (for their influence on bearing stress at failure: Use

a minimum of 10% of 90° plies.

7) Net tension failure is influenced by the tensile strength of the fibers at fastened joints, which

is maximized when the fastener spacing is approximately four times the fastener diameter

(see figure 85 above). Smaller spacing's result in the cutting of too many fibers, while larger

spacing‟s result in bearing failures in which the material is compressed by excessive

pressure caused by a small bearing area: Use minimum fastener spacing as shown in figure

66 with 5D spacing between parallel rows of fasteners: Pad up to reduce net section

stresses.

8) To avoid fastener pull-through from progressive crushing / bearing failure:- Design joint as

critical in bearing: Use pad up: Use a minimum of 40% of ± 45° plies: Use washer under

collar or wide bearing head fasteners: Use tension protruding heads when possible.

9) To avoid shear failure:- Use large diameter fasteners: Use higher shear strength fasteners:

Never use a design in which failure will occur in shear.

Composite structural mechanically fastened joint design guidelines (continued).


10) Use two row joints when possible, as the low ductility of advanced composite material confines

most of the load transfer to the outer rows of fasteners.

11) The choice of optimum layup pattern for maximized fastener strength is simplified by the

experimentally established fact that quasi – isotropic patterns (0°/±45°/90°), or (0°/45°/90°/-45°)

are close to optimum, in practice this reduces experimental costs and simplifies analysis and

design of most fastened joints.

12) The effects of eccentricities on joints:- if eccentricities exist in a joint, the moment produced

must be resisted by the adjacent structures: eccentrically loaded fasteners patterns may

produce excessive stresses if eccentricity is not considered.

13) Mixed fastener types should not be used, i.e. it is not allowed to use both permanent fasteners

and removable fasteners in combination on the same joint, this is due to the better fit of the

permanent fasteners, which would result in the removable fasteners not picking up their

proportionate share of the load until the permanent fasteners have deflected enough to take up

clearance of the removable fasteners in their holes.

14) Do not use a long string of fasteners in a splice joint, because the end fasteners will load up first

and hence yield early. Therefore use three or four fasteners per side as the upper limit unless a

carefully tapered, thoroughly analyzed splice is used (wherever possible use a double shear

splice).

163



15) Use tension head fasteners for all applications (because potentially high bearing stress under

the fastener head cause failure). Shear head fasteners may be used in special applications.

16) Where local buildup is required for fastener bearing strength, total layup should be at least 40%

± 45° plies.

17) Installation of fasteners wet with corrosion inhibitor may be required in some cases.

18) Use of large diameter fasteners in thicker composite assemblies (for example to transfer critical

joint loads, fastener diameters should be about equal to the laminate thickness) to avoid peak

bearing stress due to fastener bending. Fastener bending is much more significant for

composites than for metals, because composite are thicker for a given load, and more sensitive

to non-uniform bearing stresses due to brittle failure modes.

19) N.B. the best fastened joints can barely exceed half the strength of unnotched laminate.

20) Peak hoop tension stress around fastener holes is roughly equal to average bearing stress.

21) Fastener bearing strength is sensitive to through - the – thickness clamping force of laminates it

is highest for a 30% / 60% /10% (0º/± 45°/90°) ply lay up stack, and much lower for

50%/40%/10% (0º/± 45°/90°) ply lay up stack.

22) Production tolerance build ups:- proper tolerances should be determined with manufacturing to

minimize the need for shimming: shim allowance should be called out on engineering drawings:

N.B. since production tolerances can easily be exceeded in the thickness tolerance, fastener

grip length can be adversely affected.

164



Shims are used in airframe production to control structural assembly and to maintain aerodynamic

contour and / or structural alignment. With composite joints the allowable unshimmed gaps are only

¼ as large as those for an similar metallic structural joints. Therefore, the assembly of composites

generally require more extensive use of shims than comparable metal components.

Engineering can reduce both cost and waste by controlling shim usage through design and

specifications. Design can control where to shim: what the shim taper and thickness should be:

what gap to allow: and whether the gap should be shimmed or pulled up with fasteners.

Shim materials currently available are:-

1) Solid shims:- titanium: stainless steel: precured composite laminates: etc.

2) Laminated (or peelable) shims {with a laminate thickness of about 0.003” (0.0762mm) ±0.0003”

(0.00762mm)}

Laminated titanium shims:

Laminated stainless steel shims:

Laminated Kapton shims.

3) Moldable shim, which is a cast – in – place plastic designed for use in filling mismatches

between metal or composite parts. It can be used at any location to produce custom mating

molded surfaces examples are given in the reference works given in the end of this

presentation.

165

Composite structural mechanically fastened joint design shim guidelines.


Section 6:- Environmental protection of composite airframes.

Impact damage:- Impact damage in composite airframe components is a major concern of

designers and airworthiness regulators. This is due to the sensitivity of theses materials to quite

modest levels of impact, even when the damage is almost visually undetectable. Detailed

descriptions of impact damage mechanisms and the influence of mechanical damage on residual

strength can be found in reference 3. Horizontal, upwardly facing surfaces are the most prone to

hail damage and should be designed to be at least resistant to impacts in the order of 1.7J (This is

a worst case energy level with a 1% probability of being exceeded by hail conditions). Surfaces

exposed to maintenance work are generally designed to be tolerant to impacts resulting from tool

drops (see figure 88). Monolithic laminates are more damage resistant than honeycomb structures,

due to their increased compliance, however if the impact occurs over a hard point such as above a

stiffener or frame, the damage may be more severe, and if the joint is bonded, the formation of a

disbond is possible. The key is to design to the known threat and incorporate surface plies such as

Kevlar or S2 glass cloth. Airworthiness authorities categories impact damage by ease of visibility to

the naked eye, rather than by the energy of the impact: - BVID barely visible impact damage and

VID visible impact damage are the use to define impact damage. Current BVID damage tolerance

criterion employed on the B787 is to design for a BVID damage to a depth of 0.01” to 0.02” which

could be caused by a tool drop on the wing, and missed in a general surface inspection should not

grow significantly to potentially dangerous structural damage, before it is detected at the regular

major inspection interval. This has been demonstrated through a building block test program, and

the wing structures so inflicted have maintained integrity at DUL. These design criteria are critical

airworthiness clearances ACJ 25.603 and FAA AC20.107A (Composite Aircraft Structures). Also

around fuselage cutouts and doors CFRP aircraft have thicker skins to resist ramp rash (figure 89).

166


Figure 88:- Structural damage risks to composite wing structures.

167


168

Figure 89:- Structural damage risks to composite fuselage structures.


Lightening strike protection is a major issue for composite airframes because, CFRP composite are

poor conducting materials and have a significantly lower conductivity than aluminium alloys,

therefore the effects of lightening strikes are an issue in composite airframe component design and

a major issue for airworthiness certification of the airframe. The severity of the electrical charge

profile depends on whether the structure is in a zone of direct initial attachment, a “swept” zone of

repeated attachments or in an area through which the current is being conducted. The aircraft can

be divided into three lightening strike zones and these zones for the wing with wing mounted

engines is shown in figure 90, and can be defined as follows:-

Zone 1:- Surface of the aircraft for which there is a high probability of direct lightening flash

attachment or exit: Zone 1A- Initial attachment point with low probability of flash hang-on, such

as the nose: Zone 1B- Initial attachment point with high probability of flash hang on, such as a

tail cone.

Zone 2:- Surface of the aircraft across which there is a high probability of a lightening flash

being swept by airflow from a Zone 1 point of direct flash attachment: Zone 2A- A swept-stroke

zone with low probability of flash hang-on, e.g. a wing mid-span: Zone 2B- A swept-stroke zone

with high probability of flash hang-on, such as the wing trailing edge.

Zone 3:- Zone 3 includes all of the aircraft areas other than those covered by Zone 1 and Zone

2 regions. In Zone 3 there is a low probability of any direct attachment of the lightening flash arc,

but these areas may carry substantial current by direct conduction between some Zone1or Zone

2 pairs.

169

Environmental protection of composite airframes (continued).


Zone 3 Indirect effects.

Zone 2 Swept stroke.

Zone 1 Direct strike.

Lightening Strike

Zones on an

aircraft with wing

mounted engines.

Figure 90(a):- Lightening strike risks to composite wing structures with podded engines.

170


171

Figure 90(b):- Lightening strike risks to composite podded engine aircraft structures.

Zone 1 direct strike. Zone 1 direct strike.

Zone 1 direct strike.










Zone Key.



172


Lightening effects can be divided into direct effects and indirect effects:-

(1) Direct Effects: - Any physical damage to the aircraft and / or electrical / electronic systems due

to the direct attachment of the lightening channel. This includes tearing, bending, burning,

vaporization or blasting of aircraft surfaces / structures and damage to electrical / electronic

systems: (2) Indirect Effects: - Voltage and / or current transients induced by lightening in aircraft

electrical wiring which can produce upset and or damage to components within electrical /

electronic systems.

The areas requiring protection in this study are:-

1) Non-conductive composites (e.g. Kevlar, Quartz, fiberglass etc.):

Do not conduct electricity:

Puncture danger when not protected.

2) Advanced composites skins and structures:

Generally non-conductive except for carbon reinforced composites:

Carbon fibre laminates have some electrical conductivity, but still have puncture danger for skin

thickness less than 3.81mm.

3) Adhesively bonded joints:

Usually do not conduct electricity:

Arcing of lightening in or around adhesive and resultant pressure can cause disbonding.


173


4) Anti-corrosion finishes:

Most of them are non-conductive:

Alodine finishes, while less durable, do conduct electricity.

5) Fastened joints:

External fastener heads attract lightening:

Usually the main path of lightening transmission between components:

Even the use of primers and wet sealants will not prevent the transfer of electric current from

hardware to structure.

6) Painted Skins:

The slight insulating effect of paint confines the lightening strike to a localized area so the that

the resulting damage is intensified:

Lightening strikes unpainted composite surfaces in a scattered fashion causing little damage to

thicker laminates.

7) Integral fuel tanks:

Dangers are melt-trough of fasteners or arc plasma blow between fasteners and the resulting

combustion of fuel vapors in the tanks.

Methods of lightening strike protection for composite airframe structures have been developed and

are illustrated in figures 91 and 92.


174

Figure 91:- Lightening strike protection of composite wing structures.

Copper grid

Fig 91(a) EAP Aluminium Foil.

Fig 91(b) Typhoon Copper Strip. Fig 91(c) B787 Copper grid.


175

Figure 92:- Lightening strike protection of composite fuselages ( A350 and B-787).

Electrical network following frames and floorgrid.

Grounding

Bonding

Voltage

HIRF Protection

CFRP

Lightening Direct Protection:

CFRP + Metallic Mesh.

Figure 92(b) Airbus A350 system.

The Boeing 787 employs

Inter-Woven Wire Fabric

(IWWF) Lightening strike

protection.

Figure 92(a) Boeing B787 system.


Figure 93(a):- Design philosophies applied to composite fuselages ( A350 XWB).

176

13m 20m 14m

Side panels.

Top panel.

Keel panel.

A350 Design philosophy:- In order to reduce the operating costs and environmental impact through

reduced fuel burn the airbus A350 adopted the use of a four composite panel layout for the

fuselage skins in the areas shown above.

The key attributes of this layout:-

The skin panels are as long as possible to reduce the number of circumferential joints:

The longitudinal joints participate in the fuselage resistance to bending hence increasing

bending strength:

Each panel can be optimised for its design case:

Significant weight reductions can be achieved by this design philosophy.


Figure 93(a):- Design philosophies applied to composite fuselages ( B-787).

Contoured section. Constant section. Nose section.

Section 48. Section 47. Section 46. Section 44 / 45. Section 43. Section 41.

Fwd body joint. Aft body joint. Centre section

joints.

Aft section joint.

Boeing 787 Design philosophy:- Multiple filament wound barrel sections with major circumferential

splice joints between sections 41 to 43, and 46 to 47. These barrel sections allow a single

manufacturing process to be applied to constant, contoured, and nose sections of the fuselage.

Resitting hoop stresses better than metallics, this allows higher cabin pressures, and larger

windows. 177


Airframe structural design requires a continuing assessment of structural function to determine

whether or not the requirements have been satisfied. The expected service performance must be

satisfied before the structure enters the service environment. This assessment is the structural

testing which will ensure and substantiate structural integrity per certification for either civil or

military requirements. The basic “building block” and TRL approach, shown in figures 94 through

97, for testing of anisotropic laminate structures should be established at the early stages of

development because the validation process for composite structures is very dependant on testing

of all levels of the manufacturing process to meet AMC No1 to CS25.603 reference 10.

Composite structural testing is similar to most metallic structural testing (the majority of metallic

testing procedures are applicable to composite structures) in that it requires knowledge of design

and analysis. The difference is that composites behave anisotropically and need thorough

experimental testing, not only of the structure as a whole, but also of test specimens at the coupon,

element, and component levels.

Design with composite materials requires a knowledge of lamination theory and appropriate failure

criteria, as well as related analysis. These analyses must deal with the new set of material

properties that result from the making of the laminate. Laminate properties test results are not

useful to the engineer until the data is reduced, and translated into design allowables, and then

reported in a standard format that can be clearly understood with no ambiguity.

The purpose of a structural test program is to establish failure modes, demonstrate compliance with

criteria, and correlate test results with theoretical predictions and thus assure confidence in the part

or overall airframe structure that it will perform satisfactorily throughout its service life.

178

Section 7:- Composite testing and Qualification.


179

Extensive risk management is required certify a new structural material and / or manufacturing

process, and processing and process variability can significantly impact structural performance.

Figure 94:- Certification route for new composite structural materials and processes.


180

Figure 95:- Building block Technology Readiness Level risk reduction maturation.


Figure 96:- Building block Certification route test article examples for a CFC wing.

181


182

Figure 97:- Building block Certification route augmented by analysis for a CFC wing.


The Building Block assisted by analysis approach which I have been involved in my career,

consists of three stages as illustrated above in figure 94 namely:- Materials property evaluation:

Design-value development: and Analysis verification, and these are covered below.

1) Materials Property Evaluation:- This consists of primary coupon testing, and is fundamentally

important in that a structures constituent components and materials are studied under an

encompassing range of service conditions before a program is locked into a production design.

For example, expensive redesigns may be avoided by an early screening of matrix materials to

assess moisture degradation effects. A broad range of material and component characterisation

tests should be completed to establish lamina material properties and establish lamina design

allowables (design criterion varies for particular applications. A large number of tests are

required to satisfy these requirements. It is vital that emphasis is placed on accurate material

property characterisation, as modern computer design techniques e.g. FiberSIM TM and Catia

based CPM, FEA e.g. figure 98, used in analysis of composite anisotropic materials are

extremely dependant on and sensitive to the quality of the material property data parameters

which are furnished from coupon testing results, directed to establish lamina material properties

and establish lamina design allowables (design criterion varies for particular applications).

Single ply (lamina: tape or fabric) properties are obtained experimentally from multi-ply

unidirectional laminate specimens where all plies have the same orientation. For tape laminates

with all fibres aligned in the same direction (also tests on cross – plied laminates can be

considered to determine unidirectional properties), the ply properties needed for design are: -

Ultimate strength values: Elastic constraints: and Poisson‟s ratio values.

183

The Building Block approach for composite testing and qualification.


184

Figure 98:- Catia V5.R20 composite structural analysis.


Test coupons that are designed to be weighed during the conditioning process should be

weighed immediately after fabrication. All of the coupons are then stored in a dry desiccated

chamber prior to conditioning. It is vitally important that the fibre volume and void content of

each coupon is known. Moisture is absorbed by the matrix, so the percentage of matrix in a

given coupon will affect the amount of moisture absorbed. The size and concentration of voids

present in the coupon must also be known. The relative humidity in the conditioning

environmental chamber will determine the maximum moisture content of the conditioned test

coupons in this conditioning. Table 11 illustrates the effect of Fibre Volume Percentage (FVP)

on the mechanical properties of laminate test coupons.

There are several basic coupon tests which would form the basis of a building block test

program aimed at validating a composite wing box structure, and these would deliver an

adequate design database, for establishing the design properties of the material system and

identify the most critical environmental exposures including humidity and temperature (AMC

No1 to CS25.603 section 4: - Material and Fabrication Development). These tests are listed

here and detailed in reference 4: - (1) Tensile tests: (2) Compression tests: (3) Shear tests: (4)

Flexural tests: (5) Short Beam tests: (6) Moisture and temperature (hot-wet) tests: (7) Notch

tests: (8) Impact tests: (9) Fastener Bearing and Pull-trough tests: (10) Process control tests.

Some elevated temperature moisture coupon testing data would be used in support of element

testing to meet certification requirement 5.3(a) of AMC No 1 to CS 25.603.

185

The Building Block approach for composite testing and qualification (continued).


Properties

Effect of FVP on laminate mechanical properties.

0ºnt 90ºnt ±45ºns [(±45º)5/0º16/90º4]c

Ultimate strength Varies directly

with FVP

Not sensitive to

FVP

Varies directly

with FVP

Varies directly

with FVP

Ultimate strain Not sensitive to

FVP

Not sensitive to

FVP

Not sensitive to

FVP

Not sensitive to

FVP

Proportionality

limit stress

Varies directly

with FVP

Not sensitive to

FVP

Varies directly

with FVP

Varies directly

with FVP

Proportionality

limit strain

Varies directly

with FVP

Not sensitive to

FVP

Varies directly

with FVP

Varies directly

with FVP

Poisson‟s ratio Not sensitive to

FVP

Not sensitive to

FVP

Varies directly

with FVP

Not sensitive to

FVP

Modulus of

elasticity

Varies directly

with FVP

Varies directly

with FVP

Varies directly

with FVP

Varies directly

with FVP

186

Table 11:- Effect of Fibre Volume Percentage (FVP) on laminate mechanical properties.


2) Design Value Development:- This consists of structural element and component testing

represents the second block in the building block composite certification test program shown in

figures 94 through 96, although there is a growing opinion that these tests can be reduced and

substituted by computer analysis tools based on FEA and structural optimisation software, as

shown in figure 96. However currently and for the foreseeable future representative structural

element testing will play a key role in composite airframe certification programs. Figures 96 and

97 show typical structural elements and components used for allowables verification, and

fulfilment of both static and fatigue / damage tolerance structural integrity requirements. Such

elements contain detail features such as holes, notches, stringer run–outs, joggles, and the

objective of element and component testing is to determine what effect these features have on

the total structure, for example:-

• An access hole through a skin structure may drastically alter the stress concentration and

redistribution in the surrounding area:

• A fastened bonded and / or fastened joint may also produce significant stress perturbations in the

joints immediate vicinity:

• These sections of components may induce large stress perturbations in the constitutive material

and induce failure modes very different from those predicted by laminate theory:

• In addition to inplane axial and shear loads, concentrated normal tension load on a composite

integrally stiffened panel, can be used to determine the flatwise tension and peel strength between

the skin and stiffener which are much lower than inplane laminate strengths, hence stiffener pull –

off strength tests would be conducted as part of the wing structure qualification program.

187



Element and component testing will require much more instrumentation and have more

complicated load introduction and test fixtures than coupon testing therefore this form of testing

is more expensive, but yields a much more accurate picture of structural behaviour. The

element and component testing would be used to cover both the Proof of Structure under static

loading and Fatigue / Damage Tolerance requirements of AMC No1 to CS 25.603. The types of

test covered include the following:-

a) Joint Design evaluation:- One of the most difficult aspects of joint testing is inducing the loads

into the joint in a fashion which is representative of the boundary conditions of a test article. For

example, it may be difficult or virtually impossible to determine, much less duplicate in a test,

the stiffness boundary conditions which are present at the joint in actual service. The choice of

boundary conditions which are readily reproducible in most tests consist of either free or fixed

supports, which usually have a very high reserve factor on them for BAE Systems STF in the

order of 4. Based on previous testing on legacy aircraft information may be available as to the

procedure and gripping hardware which would be most appropriate for approximating in situ

conditions, such as historical tests on Airbus A320‟s empennage which could be applied to the

A400M wing testing. The service stress distribution in the components which border the joint

would then have to be predicted by analytical methods probably FEA modelling. Then it is

possible to approximate the same stress proportions by using boundary control techniques

which are related to an active feedback signal from the component under test. Such a test

would be expensive, but the application may be critical enough to warrant resorting to such a

technique, for example: - the adhesively bonded spar to bottom wing skin joints.

188



b) Cut–outs:- Obviously, small coupon test specimens are inappropriate for evaluating the effects

of cut – outs, or large flaws unless the imperfection being assessed is small compared to the

coupon width; coupons can also be adversely affected by free edge stress effects. Therefore

panel tests with major and minor dimensions close to that of the actual structure would be used

for notch, cut – out or imperfection tests. In composite structures, a large cut – outs such as

access holes, or fuel transfer holes, will present significantly different stress redistribution

around the edge of the cut – out and therefore an array of strain gauges would be used to

quantifying the strain distribution. But the tips of cracks cause steeper strain gradients which

would be measured by photo – elastic coatings or Moiré fringe analysis.

c) Free Edge Effects:- The delamination problem which is associated with free edges in cross –

ply laminates detailed in answer to question 2(a) above will be more severe in laminates with

cut – outs because large stress concentrations exist in the vicinity of cut - outs. Therefore

measurements of through – the – thickness deformation should be made at the cut - out edge

since this may be the most relevant measurement to support analytical characterisation studies.

Also strain gauges, displacement sensors, and optical methods could be used for delamination

strain characterisation.

d) Damage Tolerance testing:- Damage tolerance testing is significantly different for composites

than for metal. Damage tolerance in metals is related to the rate of propagation of a crack of a

given size and location, where as damage tolerance in composites is primarily dependent on

resistance to impact. Composite material structures must be designed to support design loads

after an impact that has a reasonable probability of occurring during fabrication or during the

service life of the structure.


189


d) Damage Tolerance (continued):- To define a strain allowable to account for impact damage

compression stress is similar to defining a fatigue allowable for metal structure (tensile stress is

critical). The fatigue allowables are selected based on limited tests and previous design

experience. However, final fatigue substitution is based on durability (fatigue) tests conducted

on full – scale components or the complete airframe. Compression tests are conducted on

impact damaged coupons to select preliminary compression design stress allowables, and then

compression tests of impact damaged structural panels and subcomponents are conducted to

substantiate the design allowable. To define design allowables for impact damage, tests would

be conducted on flat laminates loaded in compression. These may have varying amounts of

impact damage, dependent on panel thickness and damage tolerance requirements for damage

visibility and maximum impact energy. The panels must be large enough to nullify size effects,

e.g., 25.4cm x 30.48cm. The results being representative of impact damage to areas of the

structure between reinforcements (e.g. stiffeners). The effect of impact damage where

reinforcements are attached to the skin or the effect on the reinforcements themselves would

be determined by tests on reinforced structurally representative panels. Because strength and

damage sustained can vary as a function of lay - up configuration, several variations of each

laminate would be tested. The effects of environmental degradation would also be evaluated

with tests at given moisture content and temperature, with pre – conditioned structural panels,

tested in environmentally controlled test chambers. Some tests would also be conducted with

higher impact energies to determine the trend of data for wider damage widths. It would also be

necessary to conduct sufficient cyclic tests to ensure that no detrimental damage growth will

occur during the structures expected service life.


190


d) Damage Tolerance (continued):- The damage requirements vary considerably, depending on

mission and life – time requirements. For example the requirements for a typical military

composite structure (as this aircraft may be used as a military transport) are as follow: -

(1) Low level impact damage: -

An impact of 8.4J from an impactor with a 12.7mm diameter hemispherical head:

The damage laminate should have the capability of carrying static ultimate loads.

(2) High level impact damage: -

An impact of 140J from an impactor with a 25.4mm diameter hemispherical head:

Or an impactor which would not cause a dent deeper than 2.54mm:

The damaged laminate should have the capability of carrying static limit load.

e) Durability (Fatigue) testing:- Durability testing in composites must consider the effects of

environmental exposure on static and dynamic behaviour. Therefore the durability testing of the

composite wing components becomes a function of load cycling and environmental exposure.

Airframe durability testing would be would be accomplished using a flight by flight real – time

loading spectrum based on the aircrafts life – time and, concurrently, environmental exposure

based on flight temperatures and ground based moisture environments. In addition, accelerated

flight spectrum loading and accelerated moisture / temperature environments could be used to

simulate real – time testing but care would need to be taken in correlation of these accelerated

tests with real – time loading and environmental conditioning.

•


191


e) Durability (Fatigue) testing (continued):- Fibre – dominated laminates are considerably more

efficient in load – carrying ability than are matrix – dominated laminates: however the latter are

sometimes needed, for multidirectional loadings and damage tolerance requirements. It is

generally assumed that matrix – dominated laminate design is governed by durability strength,

where as fibre – dominated laminate design is governed by static strength. Therefore, durability

testing for structural integrity verification of matrix – dominated laminates such as those for the

bottom wing skins would have to include bonded joints to the spars. Element and component

tests would address requirements for Proof of Structure by static: fatigue: damage tolerant: and

fail safe evaluations meeting sections 5 especially 5.8 and 6 of CS 25.603 and CS 25.571.

3) Analysis Verification:- Full – scale testing (FST) of the complete airframe, or the testing of a

major structural component, is the major test in an airframe structural test program, and is the

final building block in figure 93. FST is one of the primary methods of demonstrating that the

airframe or major structural component e.g. the FATA wing, can meet the structural

performance requirements and is extremely important because it tests all of the related

structures in the most realistic manner. Typical FST include: - static: durability (fatigue): and

damage tolerance. The use of FST must take into account the unique characteristics of

composite structures and their response to the expected service conditions as simulated by the

test. FST is necessary check in the process of developing satisfactory structural systems,

although analytical techniques have significantly improved in recent years with more capable

computer analysis techniques and the wide – spread use of finite element analysis, the

complexity of composite structural systems still requires FST verification programs.


192


Test requirements such as limit and ultimate loads are often established on the basis of

material test scatter derived from coupon testing, and composites usually exhibit higher scatter

than metallics therefore this raises difficulties in establishing values for the test. Also, composite

laminates exhibit relative brittleness, low interlaminar strength and differences in coefficient of

thermal expansion (CTE) in contact with metal parts, and all of these factors would present

serious problems for the FST program. There are three considerations which would need to be

addressed when choosing the size of the FST article for the wing test but are equally applicable

to all FST programs: -

• The test article must be large enough to allow for proper complex loading and for the load

interactions at interfaces that would otherwise would be difficult to simulate:

• If the component is small enough it is less expensive to use a FST environmental test to certify the

structure, (an example of scale is the environmentally controlled cabin pressure and bending test

conducted by myself at BAE Systems Brough STF on the Eurofighter Typhoon single seat test

article in 1992 which was a complete forward fuselage contained within a purpose built chamber

under load):

• Structural configuration also has an important role in the environmental condition test: - Primary or

secondary structure: Type and complexity of loading.

For example a wing test would be a production representative half span article with a dummy

counter balance as in the case of the Nimrod MR4 test at BAE Systems STF. This would suffice

as the wing / fuselage would be included in the test article and the port wing design features

would be mirrored in the starboard wing, if there were any non – symmetrical details these

would be tested as component test articles.


193


The major FST objectives are as follows: -

• To verify analysis with actual internal load distribution were a test article is used which may

not representative of the final production configuration of the structure, (as was the case with

the JSF Horizontal Tail Test box for which I was responsible for designing).

• To observe any unexpected discrepancies occur:

• To evaluate whether durability and damage tolerance have been adequately assessed:

• To evaluate the durability of combinations of composite and metal parts, particularly in

interface areas where glass cloth packers are required due to galvanic corrosion and to

investigate differential thermal expansion problems.

Instrumentation (all data would be electronically recorded and controlled by computer data

logging and control system) used on the FST structures would include: - (1) Strain gauges: (2)

Deflection indicators: (3) Accelerometers: (4) Stress coatings: (5) Acoustic emission detectors:

(6) Evener systems.

Pre – test prediction of the test article FST structural failure loads, locations and mechanisms

are important as they will profoundly influence the test loadings, rig design and load application.

These would be based on minimum margin of safety calculations and the known statistical

variation of the material allowable developed from coupon tests and used in analysis.

Appropriate “knock – down” factors are applied to test margins after completion of the

mechanical property and environmental testing program.


194


These results would be verified by long – term aging tests on critical structural components,

which are subjected to real – life environments and tested at various intervals throughout the

duration of the test program. This would clear the FST article of the requirement to be

environmentally conditioned because perfect duplication of moisture / temperature / and time

histories for such a large and complex structure would impossible and even attempting it would

be unacceptably costly, and component testing is considered validation under section 6 of AMC

No1 to CS 25.603. Careful consideration of the method of inducing loads into the FST of the

wing would be required, generally: -

a) In tension testes:– The mating structures must be sufficiently strong that they must not fail

before the structure under test.

b) In compression tests:– The mating structure must be simulated and the loads applied to it

in such that the rotational characteristics are approximated. This subjects components

which are in buckling critical to appropriate end – fixity conditions and ensures adequate

load diffusion into the test structure.

Example Static FATA Wing Test:- The static FST a most important test in the qualification of

composite airframe structures because of their brittleness and sensitivity to stress concentrations

compared with the same structures in metal therefore to meet AMC No1 to CS 25.603 section 5 the

following methodology would be applied to the test article described above:-


195


Example Static FATA Wing Test (continued):-

1) The parameters considered for the static test would be:

• Type of test structure

• Type and number of load conditions

• Usage environment to be simulated

• Type and quantity of data to be obtained.

As stated above the environmental effects would be addressed at the analysis, coupon,

structural element, and component level “building block” stages (see figure 94). The sums of

these tests would be consolidated to validate and satisfy the consideration of environmental

testing.

2) The method of loading the FST article requires careful consideration due to the composites

weak through the thickness strength (tension) and sensitivity to stress concentrations, possible

methods for the wing test are outlined below:

a. Tension – patches method (see figure 99):

• Offers uniform load distribution with a closer representation of the real structure load but is

expensive:

• Involves a more complex test set – up (higher cost and longer set – up time):

• Introduction of load directly into a composite bonded surface must be done more carefully

than with metal surfaces because of their inherent through – the thickness weakness.


196


197

Figure 99(a):- A400M Fatigue test rig mounting.

Figure 99(b):- F-35B Fatigue test rig mounting.

For transports hydraulic jacks apply computer

controlled loading case spectrum through skin

bonded tension pads.

Same methodology applied to fighter

aircraft:- hydraulic jacks apply computer

controlled FALSTAFF or full spectrum

flight by flight loading cases to the

structure through skin bonded tension

pads.

Figure 99:- Full scale airframe structural testing loading pad method.



b. Loading frame method (see figure 100):

• Less complex loading set – up and less costly method:

• All loads are converted into numerous compressive concentrated loads (this is not as

effective as the tension – patches method but it is acceptable:

• The attachment of substructures such as spars, ribs etc. at locations of concentrated

loads needs careful investigation to make sure there is sufficient strength reserves are

present in the substructure.

3) The following FST sequence would be followed in accordance with reference 4:

a. Checking of the test set – up, which would involve functional testing of:

• Loading jacks and evener system:

• Instrumentation:

• Data recording:

• Real – time data displacement (this check would be accomplished by applying a simple

load case at low levels to ensure that the loads are induced as expected.

b. A strain and deflection survey would be run to determine whether the strain distribution and

deflections are as predicted.

c. The lowest of the loads to be certified are applied first i.e. the conditions for which there is the

highest confidence are run first and the conditions with the highest risk of premature failure

are run last.


198


199

Figure 100:- Full scale wing structural testing loading frame method.



d. The early test results could be extrapolated to the predicted design ultimate load level for

analysis validation.

e. If a risk of failure before design load is determined then the test would be stopped and a

careful review and investigation would need to be conducted.

4) Ultimate load requirements – i.e. the type of load required by the qualifying or certification

agencies to meet their validation requirements includes:

a. U.S. FAA requires the structure to the limit load (same as that governing metallic

structures):

b. U.S. Military requires testing to the ultimate load.

c. AMC No 1 to CS 25.603 requirements call for testing to ultimate load for the article like the

U.S. military requirement.

5) The final step is a review of data obtained the test and supporting evidence from element and

sub – component testing and evaluation of its correlation with the analytical stress

analysis. The structure should be able to withstand static loads to be expected during

completion of a flight on which damage resulting from obvious discrete sources occurs.

Durability FST of the FATA wing:- Cyclic Full Scale Testing of airframe structures used to

evaluate metal structures is also applied to composite structures. In general, FST cycle testing is

limited to 2 to 4 lifetimes of spectrum loading (2 for civil aircraft) in the presence of BVID, including

a spectrum load enhancement factor such as environmental effects.


200


Periodic inspections must occur during FST durability testing at specific intervals between the limits

of detection and the time when limits of residual strength capability have been reached. These

inspections are conducted to determine whether any damage is progressing due to cyclic loading in

order to: -

• Obtain the durability performance of the structural details:

• Detect any critical damage whose growth would result in failure of the test article during the

durability test.

For example stiffness changes in a composite structure has been found to be an indication of

fatigue damage, hence crack and delamination (very difficult to detect) inspections are conducted at

intervals throughout the test, after a given number of cycles which would be based on coupon,

element and sub - component level testing. The inspection plan would use the minimum detectable

damage / defect size established in the materials qualification and manufacturing development

coupon, element, and sub - component test level of the building block test program and would

determine: - the frequency, and extent of the inspections, the methods employed, intervals,

inspection for zero growth, and the residual strength associated with assumed damage. Non –

Destructive Inspection techniques likely to be employed are ultrasonic C – scan, x-ray, acoustic

detection by microphones in the structure to listen for delaminations. Finally a post – test inspection

of the test article after the FST durability test would be conducted to ensure that no damage had

occurred that would threaten the structural integrity of the composite wing box.


201


Damage Tolerance FST of the FATA wing:- Testing composite FST structures for damage

tolerance is especially important because it addresses the concerns associated with both the static

and durability FST‟s. The damage tolerance test, like the static test, is a qualification requirement to

meet the Proof of Structure requirements of AMC CS 25.603, and is also required by the U.S. FAA,

and military regulatory authorities. The load specified by civil and military requirements varies )both

specify a residual strength requirement which in this case is equal to or greater than the strength

required for the specified design loads considered as the ultimate load) and requirements also vary

depending on: -

• Ability to inspect damage:

• Type of service inspection used:

• Type of aircraft.

As in durability tests the critical flaw or damage may be associated with either its initial state or its

growth after cyclic loading. The environmental effect during the cyclic test is not easily defined but

the load enhancement of the spectrum as recommended for the durability test would be the best

option. Because the FST damage tolerance test has many similarities to the static, and durability

tests, all the testing considerations which apply to them are also applicable to this test. If the

residual strength test is successfully passed the structure can then be loaded to failure to further

evaluate its damage tolerance capability. The flutter proof of structure requirement section 7 of

AMC No 1 CS 25.603 would be met by sub – component testing. The test program outlined above

would meet the damage tolerance / environmental degradation / impact evaluation requirements of

AMC No 1 CS 25.603 for a large civil aircraft composite wing box certification criteria.


202


1) AIAA Aerospace Design Engineers Guide 4th edition: by ADEG Subcommittee, AIAA Design Engineering Technical

Committee 1801 Alexander Bell Drive, Reston VA 20191-4344 USA, Published by the American Institute of Aeronautics

and Astronautics, 1998.

2) CU/CoA/AAO/1 Issue 3 Cranfield College of Aeronautics Design Manual: Published by Cranfield College of

Aeronautics September 1999.

3) Aircraft Loading and Structural Layout: Professional Engineering Publishing: by Prof Denis Howe: 2004: ISBN

186058432 2.

4) Composite Airframe Structures: Conmilit Press Ltd Hong Kong: by Michael Chun-Yung Niu: 1992: ISBN 962-7128-06-6.

5) Composite Materials for Aircraft Structures second edition: AIAA Education Series: by Alan Baker et al: 2004: ISBN 1-

56347-540-5.

6) Airframe Structural Design: Conmilit Press Ltd Hong Kong: by Michael Chun-Yung Nui: 1992: ISBN 962-7128-04X.

7) Catia V5.R20 Composite Design Engineering Workbook 1: Private Study 2013: Mr. Geoffrey Wardle (not a published

document).

8) Catia V5.R20 FEA in Airframe Design Workbook 2: Private Study 2014: Mr. Geoffrey Wardle (not a published

document).

9) Technology and Innovation for the Future of Composite Manufacturing GKN Aerospace Presentation: by Ben Davis

and Sophie Wendes.

10) FATA Airframe Design Study private study.

11) Work book 1 Catia V5. R20 Composite Design private study.

12) NASA Perspectives on Airframe Structural Substantiation: Past Support and Future Developments : Richard Young:

NASA Langley Research Centre Hampton Virginia: Presented at the FAA / EASA / Industry Composite Damage

Workshop Tokyo on June 4-5 , 2009.

13) Aerospace Structural Material Certification BOE021711-120: Dave Furdek, Manager Next Generation Composite

Materials Boeing Research and Technology: 28th February 2011.

14) Damage Tolerance in Aircraft:- by Prof P.E. Irving Damage Tolerance Group School of Engineering Cranfield

University: Published by Cranfield University 2003 / 2004.

Reference material in use for this presentation.

203

my composite design capability maintenance studies

Documents

spar design

assembly design

skin design

design rules

design standards

capability maintenance

current fata design

main design exercises