My Airframe Composite Design Capability Studies..pdf

Mr. G. A. Wardle MSc. MSc C.Eng. MRAeS.
MY AIRFRAME COMPOSITE DESIGN CAPABILITY STUDIES.
By Mr. Geoffrey Allen Wardle MSc. MSc. C.Eng. MRAeS. Current Capabilities.
ATDA Project PRSEUS Rib May 2022.
ATDA PRSEUS Upper Wing Cover May 2022.
ATDA PRSEUS Lower Wing Cover May 2021.
ATDA Project Wing Structural Layout May 2021.
ATDA Project PRSEUS Port HT lower skin assembly March 2022.

Figure 1:- My ATDA Port OB Wing section multi material structural assembly model.
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PRSEUS stitched
composite stitched ribs.
Additive Manufacturing
Technology (laser disposition)
Al/Li tip rib.
Additive Manufacturing
Technology (laser disposition)
Al/Li Aileron actuator
attachment ribs.
CFC Thermoplastic
resin spars.

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 ATDA 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 Advanced Technology
Demonstrator 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
ATDA project and encompass my experience in composite design throughout my Cranfield
University MSc in Aircraft Engineering as well as my University of Portsmouth MSc in Advanced
Manufacturing Technology and my working career in aerospace:
 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 ATDA
aircraft wing structure, a ATDA project PRSEUS rib, and the ATDA baseline wing cover skins.
This study will grow over time as more detail structural work is undertaken on the ATDA project and
it is intended to add PATRAN / NASTRAN FEA modeling of ATDA airframe 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.
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OBJECTIVES OF THIS PRIVATE STUDY IN SUPPORT OF FDSA & ATDA.

 Section 1:- Basic Catia V5.R20 CPE capability maintenance exercises:
 Section 2:- Design rules applied to main design exercises from MSc Cranfield studies and texts:
 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 from MSc Cranfield studies and texts:
 Section 7:- Composite structural testing and Qualification:
 Section 8:- Designing component ATDA project parts: (1) Spar design : (2) Skin design :
 Section 9:- Catia V5.R20 Solid part extraction for mock up and assembly evaluation:
 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).
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Contents of this presentation in support of my ATDA & FDSA design studies.

 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), after my Cranfield MSc training modules, Individual Research and
Group Design Projects, and employment experience in composite aerospace design.
 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 I used academic texts and
lecture presentation, and GDP /IRP material from my MSc in Aircraft Engineering at
Cranfield University, and the AIAA Education Series Text Books referenced, and these feed
into my ATDA future commercial aircraft airframe study.
Section 1:- Basic Catia V5.R20 CPE capability maintenance exercises.
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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.
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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.
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 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 2
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).
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Fig 2(a):- 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.
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Fig 2(b):- 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.
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Fig 3:- CT1:-Introduction to Composite Design completed part build and model tree.
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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.
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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.
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Fig 5:- CT1:- Introduction to Composite Design single ply numerical analysis.
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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.
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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.
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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.
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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.
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CT1:- Introduction to Composite Design Balanced Composite Laminate.

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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).

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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.
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Table 2:- CT1:- Composite Design Balanced Laminate Numerical Analysis.
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 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.
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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
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 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).
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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.
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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
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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.
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Fig 14:- Working With Transition Zones Ex 1 completed part build model tree.
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PlyGroup Sequence
Ply/Insert/Cut-
Piece Name
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 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.
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 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).
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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.
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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.
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Fig 17:- Working With Transition Zones Ex 2 completed part build model tree.
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PlyGroup Sequence
Name
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.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.
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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.
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Fig 18:- Limit Contour reference geometry and zones.
10 inch
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 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).
37

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.
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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.
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Fig 21:- Limit Contour Model appearance after ply exploder application.
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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.
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Fig 22:- Creation of the Limit Contour.
Multi-Selection icon
Invert Direction button
Curve C 1
Curve C 2
Limit Contour Icon
42

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.
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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.
44

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.
45

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.
46

Fig 26:- Limit Contour from Cut-Pieces using the Extrapolation Joint Type.
Relimiting Curve Joint Type selection
47

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
48

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.
49

Fig 28:- Limit Contour from Cut-Pieces using the Gap Fill Joint Type.
50

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) ).
51

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.
52

Stagger value input for both curves
Fig 30:- Limit Contour with Staggered Values and Extrapolation Joint Type.
53

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
54

Fig 32:- Limit Contour with Staggered Values completed part and model tree.
55

Table 5:- CT3:- Limit Contour Cut-Pieces Ply Group 1 Numerical Analysis.
PlyGroup Sequence
Name
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.1
0
Ply.10 GLASS 0 152.235 1.07883 0.0779503 0.0649586 10.6049 1.02912 1.25593 0.841512
Plies
Group.1
Sequence.1
1
Ply.11 GLASS 45 152.235 1.07883 0.0779503 0.0649586 10.6049 1.02912 1.25593 0.841512
56

Table 6:- CT3:- Limit Contour Staggered Values Ply Group 2 Numerical Analysis.
PlyGroup Sequence
Name
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
57

This section covers the design rules applied in the detail design of airframe structures based on my
professional experience within aerospace and Cranfield University MSc, and applied by myself in
the design of airframe components covered in my LinkedIn presentations, and further applied to the
ATDA 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. 58
Section 2:- Design rules applied to main design exercises.

59
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(a)/(b) 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(a) to 34(d), 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 ATDA project 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

Figure 33(a):- Fibre Orientation Requirements for CFC Wing Skins / covers.
Tension Bottom Wing Cover Skin.
Compression Top Wing Cover Skin.
0º Plies are to react the wings spanwise bending
(based on references 4 & 5).
The 4 Primary Ply Orientations Used for Wing Skin
Structural Plies (based on references 4 & 5).
60

Figure 33(b):- Fibre Orientation Requirements for CFC Wing Skins / covers.
61
Centre Of Pressure
Engine / Store Loading
Flexural Centre
The 90º plies react the internal fuel tank pressure and aerodynamic suction loads
The 45º and 135º Plies in the Wing Cover Skins react the chordwise shear loads
Pressure Loading
Aerodynamic suction Loading

Fig 34(a):- ATDA Structural Ply Thickness Zones Upper Wing Cover Skin R.6.1
62
PLY LEGEND.
This Legend gives the thickness
of plies in each orientation.
“t”
0º
90º
45º
135º
FWD
IN BD
24.0
6.0
3.0
7.5
7.5
24 mm
20.0
4.0
3.0
6.5
6.5
16.0
4.0
3.0
4.5
4.5
16 mm
12.0
3.0
2.0
3.5
3.5
12 mm
10.0
3.0
2.0
2.5
2.5
10 mm
8.0
3.0
1.0
2.0
2.0
8 mm
6.0
2.0
1.0
1.5
1.5
6 mm
20 mm
PLY DROP OFFS: - 1:20 SPANWISE / 1:10 CHORDWISE.
(For ATDA study un-symmetrical ply drop off e.g. 1:20 in direction
of principal stress and 1:10 in the transverse direction for weight
reduction).
 Outer OML Skin Ply.
 See also figure 28 for lightening strike
protection and figures 29 and 30 for BVID
protection.
6.0
2.0
1.0
1.5
1.5
6 mm

Fig 34(b):- ATDA Structural Ply Thickness Zones Upper Wing Cover Skin PRSUES.
63
PLY LEGEND.
This Legend gives the thickness
of plies in each orientation.
“t”
0º
90º
45º
135º
FWD
IN BD
18.0
4.0
2.0
6.0
6.0
18 mm
16.0
2.0
2.0
6.0
6.0
14.0
3.0
3.0
4.0
4.0
14 mm
12.0
3.0
2.0
3.5
3.5
12 mm
10.0
3.0
2.0
2.5
2.5
10 mm
8.0
3.0
1.0
2.0
2.0
8 mm
6.0
2.0
1.0
1.5
1.5
6 mm
16 mm
(For ATDA study un-symmetrical ply drop off e.g. 1:20 in direction
of principal stress and 1:10 in the transverse direction for weight
reduction).
 See also figure 28 for lightening strike protection and
figures 29 and 30 for BVID protection.
 NB:- These are first pass results and are conservative.
6.0
2.0
1.0
1.5
1.5
6 mm

Fig 34(c):- ATDA Structural Ply Thickness Zones Lower Wing Cover Skin R.6.2
64
(For ATDA study un-symmetrical ply drop off e.g. 1:20 in direction of
principal stress and 1:10 in the transverse direction for weight
reduction).
15 mm
10 mm
10 mm
20 mm
20 mm
15 mm
10 mm
6 mm
6 mm
8 mm
6 mm
6.0
2.0
1.0
1.5
1.5
6.0
2.0
1.0
1.5
1.5
“t”
0º
90º
45º
135º
PLY LEGEND.
8.0
4.0
1.0
1.5
1.5
6.0
2.0
1.0
1.5
1.5
10.0
3.0
2.0
2.5
2.5
10.0
3.0
2.0
2.5
2.5
10.0
3.0
2.0
2.5
2.5
15.0
4.0
2.0
4.5
4.5
15.0
4.0
2.0
4.5
4.5
20.0
4.0
3.0
6.5
6.5
20.0
4.0
3.0
6.5
6.5
This Legend gives the
thickness of plies in each
orientation.
FWD
OUT BD
10 mm
10.0
3.0
2.0
2.5
2.5

Fig 34(d):- ATDA Structural Ply Thickness Zones Lower Wing Cover Skin PRSEUS.
65
(For ATDA study un-symmetrical ply drop off e.g. 1:20 in direction of
principal stress and 1:10 in the transverse direction for weight
reduction).
14 mm
10 mm
10 mm
18 mm
18 mm
14 mm
10 mm
6 mm
6 mm
8 mm
6 mm
6.0
2.0
1.0
1.5
1.5
6.0
2.0
1.0
1.5
1.5
“t”
0º
90º
45º
135º
PLY LEGEND.
8.0
4.0
1.0
1.5
1.5
6.0
2.0
1.0
1.5
1.5
10.0
3.0
2.0
2.5
2.5
10.0
3.0
2.0
2.5
2.5
10.0
3.0
2.0
2.5
2.5
14.0
4.0
2.0
4.0
4.0
14.0
3.0
3.0
4.0
4.0
18.0
3.0
3.0
6.0
6.0
10.0
3.0
3.0
6.0
6.0
This Legend gives the
thickness of plies in each
orientation.
FWD
OUT BD
8 mm
8.0
1.5
1.5
2.5
2.5

<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(a):- Guide to typical effective depths for Sub-structure (reference 4).
66

67
Figure 35(b):- The layout of Sub-structure reduces thickness / weight of the wing skins.
Ti wing boundary and carbon PMR-15 sub-
structure with multi spar layout to resist
buckling of skins with long thin panels.
Concept structural layout for my Advanced Interdiction Aircraft
Cranfield University MSc Individual Research Project.

68
Fig 36(a)/(b):- ATDA Transport aircraft upper cover skin stringer layout to inhibited skin buckling.
Fig 36(b) Upper Cover Skin Stringer Close up of area „A‟.
Fig 36(a) ATDA Upper Cover Skin Stringer layout.
„A‟
As a Rule of Thumb:- The mass of the skins / covers is in the order of
twice that of the sub-structure. Therefore for transports and bombers
with deep wing cross-sections, stiffeners are used bonded to the
internal skin surface as shown in fig 23(a) for the ATDA wing skins.
Where the wing chord thickness is much greater than 11.8 inches.
Figure 23(b) shows a close up of the stringers which are co-bonded „I‟
section and are of constant web depth through thickness zones with
ramped upper flanges. For the PRSEUS Stringer configuration a
variable web depth will be used over the zones.
Constant web height I - section stringers better in
compression (Tear strip peel plies omitted for clarity).
1:20 Skin Zone Transition
Ramps in the direction of
principle stress.

69
Fig 36(c)/(d):- ATDA aircraft upper cover skin stringer layout to inhibited skin buckling.
Fig 11(b) Upper Cover Skin Stringer Close up of area „A‟.
Fig 11(c) ATDA Upper Cover Skin Stringer layout.
„A‟
As a Rule of Thumb:- The mass of the skins / covers is in the order of
twice that of the sub-structure. Therefore for transports and bombers
with deep wing cross-sections. The original RRSEUS Stringer
configuration was to use variable web depth will be used over the zones
to further reduce weight however on simulations the stitching head did
not have sufficient clearance and structural analysis results were
inconclusive, therefore for this study constant height PRSUES stringers
were employed.
Constant web height Pultruded Rod Over Wrap
Chamfered stringers (compression flight loading).
1:20 Skin Zone Transition
Ramp in the direction of
principle stress TYP.

Fig 37(a):- ATDA lower cover skin with co – bonded coaming stringer layout and ports.
Lower cover skin access cut-outs ports require local coaming stringers
on each side to compensate for the reduced stringer number, these have
a higher moment of inertia and smaller cross sectional area to absorb
local axial loads due to the ports.
The stringers next to the local coaming stringers on each
side need to have larger cross sectional areas to absorb a
portion of the coaming stringer load.
Stringers on the lower wing skin cover are of T- section
which are better for panels under tension loading. (Tear –
strip peel plies omitted for clarity).
1:20 Skin Zone
Transition Ramps
in the direction of
principle stress.
70

71
Fig 37(b):- ATDA wing lower cover skin with co-bonded stringer layout and inspection ports.
Note:- lower cover local coaming
stringers run on each side of the
inspection ports for nearly the full
length of the lower cover skin,
however they can be broken or re-
aligned, in this case they re-
aligned as inspection port size is
reduced.
Inspection ports are sized to permit 90 percentile
human to reach all internal structure in each bay with
an endoscope. The port size is reduced outboard as
bay size reduces, and inspection covers are CFC UD
and fabric with kevlar outer plies.
Lower cover skin access cut-outs require local coaming
stringers on each side to compensate for the reduced
stringer number, these have a higher moment of inertia
and smaller cross sectional area to absorb local axial
loads due to the cut out.

Fig 37(c):- ATDA lower cover skin with PRSEUS coaming stringer layout and ports.
72
Constant web height Pultruded Rod Over Wrap
Chamfered stringers (tension flight loading).
Lower cover skin access cut-outs ports require local coaming stringers
on each side to compensate for the reduced stringer number, these have
a higher moment of inertia and smaller cross sectional area to absorb
local axial loads due to the ports.
The stringers next to the local coaming stringers on each
side need to have larger cross sectional areas to absorb a
portion of the coaming stringer load.
1:20 Skin Zone
Transition Ramps
in the direction of
principle stress.
Fig 15(c) ATDA Lower Cover Skin Stringer layout.

73
Fig 37(d):- ATDA wing lower cover skin with PRSEUS stringer layout and inspection ports.
Note:- lower cover local coaming
stringers run on each side of the
inspection ports for nearly the full
length of the lower cover skin.
Inspection ports are sized to permit 90 percentile
human to reach all internal structure in each bay with
an endoscope. The port size is reduced outboard as
bay size reduces, and inspection covers are CFC UD
and fabric with kevlar outer plies.
Lower cover skin access cut-outs require local coaming
stringers on each side to compensate for the reduced
stringer number, these have a higher moment of inertia
and smaller cross sectional area to absorb local axial
loads due to the cut out.

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.
74

Radius fillers are necessary in T - and I – type composite stiffeners and spars. See figure 38(a) 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. NASA experimentation has shown the most effective filler material to be Braided “T”
preform – which gives good to excellent performance. Therefore this filler type will be used in the
ATDA study for both the baseline design, and when necessary in the evolved PRSEUS concept
for example in the base section of the two part PRSEUS rib and in the base of the PRSEUS
stringers.
In figure 38(b) the effects of sloping the feet of the stringer on the Peel stresses in the feet to skin
bond is shown this work conducted by GKN Aerospace and reported as part of the LOCOMACH
research studies indicates a substantial reduction in the peel stress can be achieved by slopping
the feet. However this needs to be traded against the difficulty of any future mechanical (bolted)
repair in service in the case of the baseline ATDA aircraft, and against the limitations / difficulties
such a configuration will pose for PRSEUS stitching when production feasibility studies are
conducted, against the reduction in peel stress and stringer weight.
The capping strips are bonded in place using supported film adhesive to give constant/minimum
glue line thickness of 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.
75

Fig 38(a):- Composite Stringer Types Based on my MSc and reference 5.
“L” Section Stringer (bonded or
mechanically attached panel breaker).
“Z” Section Stringer (mechanically attached to
provide additional stiffness for out of plane loading).
“I” Section Stringer (used as axial load carrying
members on panel under compression loading).
Channel
sections
Capping
strips
Cleavage
fillers
“T” Section Stringer (used as axial load carrying
members on panel under tension loading).
Capping strip
Cleavage filler
Channel
sections
76

77
Figure 38(b):- Composite Stringer design based on MSc, AIAA ES, and reference 5.
Distribution of peel stress in a basic co-bonded stringer subjected
to vertical load validated through „T‟- Pull testing, which can be
modified through redesigning the flange toe as shown.
100%
Square Edge flange toe.
Radius Edge flange toe.
Reduced by ≈ 12%
30º Chamfer flange toe.
Reduced by ≈ 41%
Reduced by ≈ 53%
6º Chamfer flange toe.
Reduced by ≈ 88%
6º Chamfer flange toe
and capping strip.
TRADE STUDY.
 REDUCTION OF PEEL STRESS
AT TOE OF FLANGE.
 REDUCTION IN STRINGER
MASS.
 INCREASED MANUFACTURING
COSTS.
 ISSUES WITH REPAIR /
FASTENERS.

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(a) through (e), 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(a), 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 ATDA 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 40(b), 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 41(a) and the separation of
joggles from changes in laminate thickness are shown in figure 41(b). The support of joggles in
structural assemblies is shown in figure 42.
78
Design of aircraft CFC wing spar structures.

79
Figure 39(a):- 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.

Figure 39(b):- ATDA Outboard Port and Stbd LE CFC Wing Spar and Symmetrical Tool.
Symmetry cut plane.
Port Outboard Leading Edge Spar.
Starboard (Stbd) Outboard Leading Edge Spar.
Two part hollow Outboard Leading
Edge Spar Symmetrical tool with
internal temperature control.
120mm Spar Cut and Trim
Zone to MEP (20mm).
60mm transition zones.
Tool extraction
direction.
Wing
Outboard.
N.B.:-Slat track guide rail cut-outs post lay up activity with
assembly tool hole drilling at extremities rib 35 and splice locations.
(N.B.:- Stbd drill breakout class cloth zones omitted for clarity).
Sacrificial Ply Zone.
Sacrificial Ply Zone.
UP
FWD
OUT BD
Boundary dimensions.
Total spar length = 6.80m :
IB flange to flange height = 0.475m:
OB flange to flange height = 0.407m:
Flange width 224mm 22mm (⅞”) dia bolts in two rows.
80

Figure 39(c):- ATDA Outboard Port CFC Wing Spar as layup and finished part.
10mm Thick Zone.
(46 plies)
7mm Zone
(32 plies)
4mm Zone
(18 Plies)
1:20 Transition zone
(3mm x 60mm)
1:20 Transition zone
(3mm x 60mm)
Slat 7 track guide rail cut-outs.
Fig 30(a) As fibre-placed.
Fig 30(b) As post finishing.
4mm Zone
(18 Plies)
7mm Zone
(32 plies)
10mm Thick Zone.
(46 plies)
Drill breakout Glass Cloth on IML
and OML for spar splice joint.
Drill breakout Glass Cloth on IML for Rib Post
Attachment and tooling holes.
Drill breakout Glass Cloth for track ribs and guide rail
can attachment both IML and OML faces.
Glass Cloth shown in white for clarity.
UP FWD
OUT BD
Tooling Hole
12.7 mm dam
Tooling Hole
12.7 mm dam
Slat track guide rail cut-outs post lay up activity with assembly
tool hole drilling at extremities rib 35 and splice locations.
81

82
Figure 39(d):- ATDA Outboard Port / Stbd CFC Wing Spar assembly.
Port Mid Section
Leading Edge Spar.
Port Outboard Section
Leading Edge Spar.
Ti alloy Rib Post 29
Assembly proposal.
Spar section is to be mounted in jig tool with
pre drilled web fastener holes for rib posts
based on CAD (Catia model). Rib posts with
web pre drilled web fastener holes are then
individually mounted in place with a robot end
effector gripping the rib web, whilst an other
end effector tool insets the bolts IML to OML,
and attaches the collars to complete assembly.
Flange fastener hole would be drilled in
assembly as per the AWBA (see My Robot
Kinematics Presentation LinkedIn).

83
Figure 39(e):- ATDA Outboard Port / Stbd CFC Wing Spar assembly.
OB Leading Edge Ti Rib Post Typical.
Pre-drilled web fastener
holes 22mm (⅞”).
Flange fastener holes
drilled on assembly
22mm (⅞”).
Initial sizing 6mm
web / flange 4mm
rib landing web.
OB Leading Edge section to Mid
Leading Edge section Splice joint.
Port Outboard Section
Leading Edge Spar.
UP
FWD
IN BD

84
Figure 40(a):- Carbon Fibre Composite ply orientations in wing spars MSc ref 3.
-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 ° of monolithic laminate.
 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
CFC Wing Spar

Fig 40(b):- Carbon Fibre Composite ply orientations in wing spars MSc ref 3.
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
85

Fig 41(a):- Separation of Web and Flange Joggles in CFC spars ref 4.
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
86

Fig 41(b):- Separation of Joggles from changes in laminate thickness in CFC spars ref 4.
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)
87

Fig 42(a)/(b):- Support of Joggles in CFC spars in structural assemblies ref 4.
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
88

WING RIBS:- The ribs, an example is shown in figure 43, maintain the determined aerodynamic
shape of the wing cross-section (chord), limit the length of skin stringers or integrally stiffened
panels to an efficient column compressive strength, and to structurally transmit chord-wise loads
across the span-wise torsion box. Hinges and supports for secondary lifting surfaces, flight controls,
are located at the ends of relevant ribs. Ribs also provide attachment points for main landing gear,
powerplants, and act as fuel tank boundaries. Overall the ribs stabilize the spars and skins in span-
wise bending.
The applied loads the ribs distribute are mainly distributed surface air loads and / or fuel loads
which require relatively light internal ribs to carry trough or transfer these loads to the main spar
structures. The loads carried by the ribs are as follows: - (1) The primary loads acting on the rib are
the external air loads which they transfer to the spars: (2) Inertia loads e.g. fuel, structure,
equipment, etc.: (3) Crushing loads due to flexure bending, when the wing box is subjected to
bending loads, the bending of the box as a whole tends to produce inward acting loads on the wing
ribs, and since the inward acting loads are oppositely directed on the tension and compression side
they tend to compress the ribs: (4) Redistributes concentrated loads such as from an engine pylon,
or undercarriage loads to wing spars and cover skins: (5) Supports members such as cover skin –
stringer panels in compression and shear: (6) Diagonal tension loads from the cover skin – when
the wing skin wrinkles in a diagonal tension field the ribs act as compression members: (7) Loads
from changes in cross section e.g. cut outs, dihedral changes, or taper changes.
89
Design of aircraft CFC wing rib structures.

90
Figure 43(a):- Composite Rib 31 from ATDA Prime Baseline typical CFC rib structure.
UP
FWD
OUT BD
Overall Thickness
6mm (28plies)
Rib Integral Cleat for Rib to
Trailing Edge Spar build joint
with single row of 16mm
fasteners (provisional).
Extensive Flange Joggling to accommodate
stringer flanges with 30º chamfer at toe.
Integrated rib web reinforcement to prevent web
buckling under in plane shear and compression
(provisionally additional 6mm 28 plies). Extensive Flange Joggling to accommodate
stringer flanges with 30º chamfer at toe.
Integral Tab for Rib to Leading Edge
Spar rib post attachment two rows of
22mm fasteners (provisional).
Fuel Vent Tank Systems
Penetrations (60mm dia notional).
As design weight in Hercules Inc AS4
Multiaxial fabric CF infused with
Hexflow VRM-34 Epoxy resin = 8.203kg.

91
Figure 43(b):- Composite Rib 31 from ATDA Prime Baseline typical CFC rib assembly.
(N.B.:- As with the metallic ribs the effort is made to use the low level fuel transfer holes
and ventilation holes as assembly tooling holes.)
Aft Low level fuel
transfer hole.
Wing Bottom Cover Skin.
Leading Edge
CFC spar.
Trailing Edge
CFC spar.
Wing Top Cover Skin.
Aft ventilation hole.
Fwd Low level fuel
transfer hole.
Mid Low level fuel
transfer hole.
Aft ventilation.
Leading Edge
Ti Rib Post.
Fwd ventilation.
Aft fuel drain.
Top Cover Skin Co-bonded Stringers.
Fwd Coaming Skin Co- bonded
Stringer.
Aft Coaming Skin Co-bonded
Stringer.
Fwd fuel drain.
Figure 44(b):- Aft Coaming Skin Stringer showing
glass packer zones typical for all stringers.
Glass packers
UP
FWD
Fwd ventilation hole.
Top Cover Skin 20mm fasteners.

92
Figure 43(c):- Composite Rib 31 ATDA Prime Baseline with tapered stringer flange toes.
UP
FWD
OUT BD
Single stage Flange Joggling for
tapered stringer flanges.
Rib Integral Cleat for Rib to Trailing
Edge Spar build joint with single row
of 16mm fasteners (provisional).
(provisionally additional 6mm 28 plies). Single stage Flange Joggling for tapered stringer flanges.
Rib overall Thickness
6mm (28plies)
As design weight in Hercules Inc AS4
Multiaxial fabric CF infused with
Hexflow VRM-34 Epoxy resin = 8.234kg.

93
Figure 43(d):- Composite Rib 31 ATDA Baseline with tapered stringer toe rib assembly.
Aft ventilation.
Aft ventilation hole.
Top Cover Skin Co-bonded Stringers.
Fwd ventilation.
Trailing Edge
CFC spar.
Aft fuel drain.
Aft Low level fuel
transfer hole. Mid Low level fuel
transfer hole.
Fwd Low level
fuel transfer hole.
Aft Bottom Cover Skin Co-
bonded Coaming Stringer.
Fwd Bottom Cover Skin Co-
bonded Coaming Stringer.
Leading Edge
Ti Rib Post.
Leading Edge
CFC spar.
UP
FWD
Figure 46(b):- Tapered Skin Stringer, note
packers required under bonded anchor nuts
Typical.
(N.B.:- As with the metallic ribs the effort is made to use the low level fuel transfer
holes and ventilation holes as assembly tooling holes.)
Fwd fuel drain.

Both the ATDA Prime baseline, and the Developed PRSEUS ATDA wing, employ carbon fibre
composite ribs at 11 locations:-
 In the case of the ATDA Prime baseline wing CFC ribs shown in figures 43(a), and 43(b) they
have top and bottom flanges, with an integral trailing edge spar cleat and a leading edge tab,
the web is stiffened with integral pad-up zones to add buckling resistance under compressive
loading, the webs have standard fuel transfer and vent holes. Both top and bottom flanges of
the rib are bolted to the upper and lower wing cover skins through the stringer flanges with
tolerance compensation, and these flanges are joggled to allow for the interface with stringer
flange toes and fitted with packers these are manufactured on an open male tool and Spring In
will be addressed with mould compression and process control based on statistical analysis. A
variation to this configuration is shown in figures 43(c) and 43(d) where fully tapered co-bonded
stringer flange toes are employed reducing peel stress further and eliminating the joggle
feature.
 In the case of the Developed PRSEUS ATDA wing CFC ribs shown in figures 44(a) to 44(e),
they have a top flange only with a separate stitched bottom integrated flange which is bolted to
the rib web as a proposed method of arresting delamination growth in the lower wing skin in the
same way as the stitched stringers concept, which has been successfully demonstrated through
the joint NASA / Boeing technology demonstration program (reference 10). This structural
assembly concept has the additional advantage of eliminating the need to joggle the rib bottom
flange to accommodate the stringer feet reducing the risk of over dimensioning the tolerance
chain and the effects of laminate thickness variations.
94
Roll and layout of large aircraft wing structural members (CFC wing ribs).

95
Figure 44(a):- Composite Rib 31 ATDA Split Rib, with PRSEUS 30º Chamfer stringers.
Stub Rib to be attached by
fasteners 14mm.
As design weight in Hercules Inc AS4 Multiaxial fabric
CF infused with Hexflow VRM-34 Epoxy resin = 7.22kg.
UP
FWD
OUT BD
Rib Integral Cleat for Rib to Trailing
Edge Spar build joint with single row
of 16mm fasteners (provisional).
Two stage Flange Joggling for
revised stringer flanges.
(provisionally additional 6mm 28 plies).
Rib overall Thickness
6mm (28plies)
Reduced cutout width for PRSEUS
Cover Skin Stringers.

Proposed assembly methodology for Stitched Split Rib 31 subsequent integration into the PRSEUS
tapered stringers / skin assembly is shown below in figures 44(b) to 44(d) follows these procedural
stages:-
1) Production of the Rib Integral Flange / Web unit comprises the bonding of two C-section
preforms, a cleavage filler and a tear strip into one unit using tack adhesive film as shown in
figure 44(b)i. The resulting unit then has the stringer cut-outs and low-level fuel transfer holes
removed, following this the unit is mounted in the stitching tool and the web is stitched with two
rows of 1200 Denier thread infused with Vectran DMS 2479 Type 2 Class 1 VRM epoxy resin,
as shown in figure 44(b)ii. The resulting unit can then be mounted and attached in place on the
Lower Wing Cover Skin, after the PRSEUS lower skin Stringers have been attached figure
44(b)iii all in the dry condition.
2) The Rib Integral Flange / Web unit when mounted over the stringers is stitched into position
using four rows of 1200 Denier thread infused with Vectran DMS 2479 Type 2 Class 1 VRM
epoxy resin, as shown in figure 44(c) the inboard stitching rows are angled at 45º so that
additional interlocking is achieved below the web on the Lower Wing Cover Skin OML this aides
the distribution of loads in the Web area. The complete Lower Wing Cover Skin mounted on the
OML tool and bagged is then infused with DMS 2436 Type 2 Class 72 (grade A) Hexflow epoxy
resin using a Boeing CAPRI type vacuum assisted resin infusion process, and cured.
3) The Upper Rib section swung into place having been inserted between the leading and trailing
edge spars and is bolted to the Leading Edge Rib Post and integral rib cleat is bolted to the
trailing edge spar. The resulting assembly is bolted to the Rib Integral Flange / Web Unit as
shown in figure 44(d). 96
Roll and layout of large aircraft wing structural members (CFC wing ribs).

97
Figure 44(b):- Composite Rib 31 Stitched Stub -Rib Preform assembly.
Tare Strip
(1.5mm)
Figure 44(b)i
J-preform
(4mm)
J-preform
(4mm)
Cleavage filler Tack adhesive film
Two rows of web stitching on three zones.
(Modified lock type)
Aft Coaming Stringer Cut-out
Figure 44(b)ii
Low level fuel transfer holes.
Figure 44(b)iii
Aft Coaming Stringer Section
Section of lower cover skin
(representative)

98
Figure 44(c):- Composite Rib 31 Stitched Stub-Rib PRSEUS Coaming stringers.
Figure 44(c)i Side view on (B)
Figure 44(c)iii Plan view
Figure 44(c)ii Front view on (A)
(Coaming Stringers omitted for clarity.)
(A)
(B)
Flange to Lower Cover Skin Stitching 4 rows 2 per side on all three zones
( Modified Lock type.)
Two rows of web stitching on three zones.
(Modified lock type) Stitching Vectors
OUT BD
FWD

99
Figure 44(d):- Proposed Rib 31/ Flange / Stringer and Spar unit assembly sequence.
(A) :- Post mounting and stitching operations on the PRSEUS Coaming Preform Stringers to
the Lower Wing Cover Skin, the Stub - Rib Flange / Web Preform section is mounted and
stitched in place and the resulting assembly is infused with Hexflow VRM-34 Epoxy Resin
using a similar method to the Boeing CAPRI vacuum assisted resin infusion process.
(B) :- The Rib Post is Bolted on to the Leading Edge Spar, and Split Rib Top
section is inserted between the Leading and Trailing Edge spars and rotated
into position forming with the other ribs the complete build unit.
Lower Wing Cover Skin section.
Stub - Rib Flange / Web Preform Section.
(C) :- The complete Outboard Wing Integral Structure
Build Unit is lowered into the Lower Wing Cover Skin,
and bolted into place, post systems integration with
the Mid Wing Integral Structure Build Unit the Upper
Wing Cover Skin with PRSEUS stringers attached
can be lowered in place on to the assembly and
bolted into place.
Trailing Edge Spar section.
Leading Edge Spar section.
Rib 31 top section. Rib 31 Post.

100
Figure 44(e):- Composite Rib 31 ATDA PRSEUS 30º Chamfer stringer assembly.
Trailing Edge
CFC spar.
UP
FWD
Leading Edge
CFC spar.
Leading Edge
Ti Rib Post.
Aft Bottom Cover Skin PRSEUS
Coaming Stringer.
Fwd Low level
fuel transfer hole.
Mid Low level
fuel transfer hole.
Aft Low level fuel
transfer hole.
Aft fuel drain.
Top Cover Skin PRSEUS Stringers illustration only.
Aft ventilation. Aft ventilation hole.
Fwd ventilation.
Fwd fuel drain.

 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 CA practice CU MSc and academic texts.
10

Figure 45:- Application of ply layup rules in general terms reference 4.
10

Figure 46:- Structural design ply lay-up guidelines reference 4.
The 4 ply separation of coincident ply joints rule.
10

104
Figure 47:- Structural design requirements for Woven cloth reference 4.
General Design Guidelines based on
reference 4 and MSc and AIAA ES.

105
Figure 48:- Structural design requirements for Woven cloth overlap and stagger ref 4.
General Design Guidelines based on
reference 4 and MSc and AIAA ES.

Lay-up Guidelines based on CA practice CU and academic texts (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 my FATA
wing covers figures 34(a)/(b).
 All ply drop-offs must 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 the following limits:-
 „t‟ < 2.5mm, radius = 2t or 3.0mm whichever is greater
 „t‟  2.5mm, radius = 5.0mm
 Plies should not be dropped nor core material run into corner radius, and plies should 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, ply quilting during cure over the core area needs
to be considered, and there is a need for core stabilisation, and reduced cure pressures to be
applied.
 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.
10

Figure 49:- Plie lay-up rosette definition and positioning MSc notes and reference 4.
107
The lay-up rosette definition.
The position of the Ply rosette.
Catia V5.R20 locates the rosette automatically on
the part the Rosette Definition being achieved by
selecting the Absolute Axis System, and the
Rosette Transfer type was set to Cartesian.

108
Figure 50:- Plie stagger rosette definition and positioning MSc notes,& references 4&5.
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.

The staggering Ply Boundaries in Ramps CA /CU and academic texts.
 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
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 12.7mm 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 the 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.
109

 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.
 See reference 4 which gives a minimum fastener spacing for fuel tanks.
More General Design Guidelines from my MSc‟s and academic texts.
110

 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 inspection zone is normally specified on the Engineering
drawing as per reference 2 , however if not the inspection zone shall will be classed as a “Zone
B” for examination purposes.
 Unidirectional Material Limits on Adjacent Plies of Same Orientation:- To avoid matrix micro-
cracking in unidirectional laminates, limit the number of plies of like-orientation be stacked
together for toughened matrix resins: For example a maximum of 0.853mm total thickness
(4 plies of 0.213mm ply material, or 6 plies of 0.135mm ply material).
 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 engineers from different disciplines within a program are involved with the
mapping out of where ply splicing will occur and this requires input from such areas as:-
Manufacturing: Materials: Design: and Stress, to coordinate the required splice locations.
More General Design Guidelines (continued).
111

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 is to be created as
per the process 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
overlap is required, and a overlap of 25.4mm is usual as the guideline shown in figures 47 and
48.
 Splicing Hand lay-Up Carbon / Epoxy Laminates:- Splicing examples for carbon / epoxy
fabric, tape, peel ply, and surface barrier material (scrim) are given in reference 4, for example:-
a minimum stagger distance between splices are for Fabric & Tape >= 300mm width minimum
stagger would be 50.8mm, and for Tape <= 300mm wide the minimum stagger would be
20.4mm. 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.
112

Figure 51:- Control of Ply Joints / splices CA / CU references 2, 4, and 5.
113

 Splicing Hand Lay-Up Carbon / BMI Laminates:- Splicing requirements for carbon / BMI
fabric and tape generally as follows a minimum stagger distance between splices are for Fabric
& Tape >= 300mm wide the minimum stagger is ≈ 50.8mm, and for Tape < 300mm wide the
minimum stagger is ≈ 20.4mm. The splice pattern should not be repeated more often than
every fifth ply of the same orientation for UD 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 generally:- minimum stagger distance between splices are for Fabric & Tape >=
300mm wide is ≈ 50.8mm and for Tape < 300mm wide the minimum stagger is ≈ 20.4mm. The
splice stagger pattern for both tape and fabric should not be repeated more often than every
fifth ply of the same orientation.
 Reducing Splices With Bias Weave Fabric:- Splices can 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 CA/CU/academic text.
114

Figure 51:- Example of Reducing Splice Task by Using Bias Weave Material MSc, AIAA ES.
45º
Warp Fiber Direction. Warp Fiber Direction.
Ply Boundary.
Ply Boundary.
Material Roll Width.
0º/ 90º Weave. 45º/ -45º Weave.
115

 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 is 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, such as flight control surfaces and
spoilers.
 Syntactic Film Core:- Syntactic film is a low-density syntactic core material ordered at either
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 references 4 & 5.
116

117
Figure 52:- 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 5.08mm
(±2.54mm). Used for structures les than
7.366mm thick such as fight control surface
skins see fig 53.

118
Figure 53:- Honeycomb elevator skin structure of a commercial transport aircraft.

 Syntactic Film Core (continued):- Syntactic film requires beveled edges, which are to be
machined or formed at a 5:1 taper with a 0.5mm offset at the edge. The corner radii should be
no less than 25.4mm, with the standard outside radius being 76.2mm. 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 also provides 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.
119
General Design Guidelines for Core Stiffening reference 5.
Syntactic film
Figure 54(a) :- Syntactic film Pinch-off configuration. Figure 54(a) :- Syntactic film Arrowhead configuration.
Symmetry
plane
Syntactic film

My Airframe Composite Design Capability Studies..pdf

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