ATDA Commercial Transport Airframe Part 2.pdf

Mr. Geoffrey Allen Wardle. MSc. MSc. ATDA Airframe Design Study 2012-2020
ATDA STUDY PRESENTATION PART: - 2 ATDA WING DESIGN, STRUCTURAL
LAYOUT, MANUFACTURE, AND MAJOR COMPONENT INTEGRATION.
By Mr. GEOFFREY ALLEN WARDLE. MSc. MSc. MRAeS. CEng. Snr MAIAA.
ATDA PRSEUS Lower Wing Cover May 2019.
ATDA Project Wing Structural Layout May 2019.
ATDA Project Wing Carry Through Box May 2019.
ATDA PRSEUS Upper Wing Cover May 2019.

This presentation has been created, for the sole purpose of private study and is not the work of a
company or government organisation it entirely the work of the author using resources in the public
domain. The final paper will be submitted for peer - review to the American Institute of Aeronautics
and Astronautics, Design Engineering Technical Committee, and the RAeS Structures and
Materials Group, for pre submission assessment. Readers must be aware that the work contained
may not be necessarily 100% correct, and caution should be exercised if this project or the data it
contains is being used for future work. If in doubt, please refer to the AIAA, Design Engineering
Technical Committee and the author.
All of the views and material contained within this document are the sole research of the author and
are not meant to directly imply the intentions of the Boeing Company, Airbus Group, GKN
Aerospace, or any contractor thereof, or any third party at this date. Although the USAF and NASA
have awarded contracts for studies into stitched composite transport aircraft structures, this work is
not the product of their results or any part of their body of research, and should not be considered
as such.
This document contains no material what so ever generated or conceived by myself or others
during my employment with BAE SYSTEMS (PLC), or that is governed by ITAR restrictions. This
work is solely my own creation and is based on my own academic studies and literature research
and the distribution of all information contained within this document is unlimited public release and
has been approved through the AIAA. This document and any part thereof cannot be reproduced
by any means in any format or used for any other research project without consultation with AIAA
Design Engineering Technical Committee or the author.
2
Presentation “Health” Warning.
.

3
This is an overview covering my current private design trade studies into the incorporation of new
structural technologies and manufacturing processes into a future transport airframe design, and
the incorporation of mission adaptive wing (MAW) technology for per review through the AIAA
This study has been undertaken after my 13 years at BAE SYSTEMS MA&I, in airframe design
development as a Senior Design Engineer, and my Cranfield University MSc in Aircraft Engineering
completed in 2007(part-time), and was commenced in 2012 and I aim to complete it at the end of
2020. This utilises knowledge and skills bases developed throughout my career in aerospace,
academic studies and new research material I have studied, to produce a report and paper
exploring the limits to which an airframe research project can be perused using a virtual tool set,
and how the results can be presented for future research and manufacturing. The toolsets used are
Catia V5.R20 for design / analysis / kinematics / manufacturing simulation: PATRAN / NASTRAN for
analysis of composite structures: AeroDYNAMIC™ for analysis of aircraft OML / Structural Loads /
performance. This work will also form the basis for a PhD study, it is the product of my own
research, and has not in any part been produced or conceptualised during my employment with
BAE SYSTEMS or any company which is any part thereof.
About this presentation:-
This presentation is Part 2 of a series of 5 presentation Parts which cover the airframe major
structural component development and engine and landing gear integration, and assembly
manufacturing technologies. The contents of this presentation are given in the following slide.
Overview of my current research activities in aircraft design for the ATDA paper.

 Section 1:- Wing planform and aerofoil selection aerodynamics of the ATDA wing.
 Section 2:- Roll, manufacturing methods, and layout of large aircraft wing structural members:
 Section 3:- Flight control surfaces sizing and design:
 Section 4:- The design and structural layout of the ATDA wing box:
 Section 5:- Wing fuel tank and engine / pylon integration into the ATDA wing:
 Section 6:- Main Landing Gear Integration in wing torsion box and wing carry through box (in
work):
 Section 7:- The design, structural layout and sizing of the ATDA wing torsion and carry through
box (in work):
 Section 8:- Wing flight control system and high lift device mechanical integration (in work):
 Section 9:- Wing assembly automation of the wing torsion and carry trough box (in work).
THIS WORK MAY NOT BE REPRODUCED WITHOUT EXPRESS PERMISSION OF MYSELF, RAeS, AND
AIAA.
4
Table of contents of this ATDA Study Presentation Part 2.

As stated in Part 1 the fundamental reasons for this research is to reduce the structural weight of
the airframe and make it easer to produce through PRSEUS technology. The former is intended to
reduce the amount of CO2 emitted from kerosene – burning aircraft engines which is solely
dependent on the amount of fuel consumed (discussed below), and the latter will reduce costs both
of acquisition and ownership through life maintenance (discussed in Part 1). Figure 1 gives the
overall dimensions of the ATDA, and Table 1 gives the configuration desired performance and
baseline dimensions for the ATDA.
The variables influencing fuel consumption can easily be examined using the Breguet range
equation. One form of the range equation for the special case of constant lift coefficient – i.e. at
constant cruise / climb – reads:-
WF = WTO * 1 – exp R equation (1.0)
X
With WTO representing aircraft take - off weight, R the mission range;
X = L / D * V = L/D *ᶇ * H equation (1.1)
TSFC *g g
V = the cruise speed;
TSFC = the thrust specific fuel consumption;
ᶇ = the overall engine efficiency;
H = the cabrific value of the fuel.
5
Section 1:- Wing planform and aerofoil selection aerodynamics of the ATDA wing.

Now equation 1.0 can be rewritten to give fuel consumption in kg per kg payload as:-
WF 1 + WE * exp R*g πeb² -1 equation (1.3)
Wp Wp CL ᶇth ᶇprop H
Where:- CDo = the zero-lift drag;
S = the wing area;
e = the Oswald factor;
B = the wing span;
CL = the aircrafts lift coefficient;
ᶇth = the engines thermal efficiency;
ᶇprop = the engines propulsive efficiency.
Minimising fuel weight, with respect to CO2 emission for a given payload and range can be obtained
by:-
 Aerodynamics:- Maximise CL, e, and b: Minimise CDo and S;
 Structure:- Minimise WE / Wp ( Weight empty / Weight payload);
 Engine:- Maximise ᶇth and ᶇprop ;
 Fuel:- Maximise H
This work is to modify the structural weight parameters, improve aerodynamics an efficiency.
6
My requirements research breakdown for the ATDA aircraft design project.
=
S
C²L
CDo +

Figure 1:- Overall configuration and dimensions of the ATDA baseline aircraft.
7
70.52m (231ft 3.3in) Code F
18.34m (60ft 7in)
11.51m (37ft 1.6in)
30.58m (100ft 3.8in)
O/A 75.87m (248ft 1.3in) Code E
74.47m (244ft 3.8in)
34.45m (113ft 2.4in)
O/A 75.27m (246ft 10.7in)
Fuselage sized for
twin aisle 9 abreast
2 LD-3 containers
5.99m (235.85in)
Section on „A‟
„A‟
„A‟
17.85m
(58ft 4.6in)
11.92m (39.136ft)
7.771m
14.154m
17.248m

IMPERIAL DATA. METRIC DATA.
Wing Span (ft / in) 231 / 3.3 Wing Span (m) 70.52
Length (ft / in) 240/88 Length (m) 75.88
Wing Area (sq ft) 4,375.49 Wing Area (sq m) 406.481
Fuselage diameter (in) 235.83 Fuselage diameter (m) 5.99
Wing sweep angle 35° Wing sweep angle 35°
Fuselage Length (ft /in) 244 / 3.8 Fuselage Length 74.47
Engine number / type 2 X RR Trent XWB Engine number / type 2 X RR Trent XWB
T-O thrust (lb) 83,000 T-O thrust (kN) 369.0
Max weight (lb) 590,829 Max weight (tonnes) 268.9
Max Landing (lb) 451,940 Max Landing (tonnes) 205.0
Max speed (mph) 391 Max speed (km/h) 630
Mach No 0.89 Mach No 0.89
Range at OWE (miles) 9,631 Range at OWE (km) 15,500
Cruise Altitude (ft) 45,000 Cruise Altitude (m) 13,716
8
Table 1: - Initial Configuration Aircraft Data for the baseline ATDA study.

Starting with the wing, the major drivers in the baseline wing structural design considered in this
study are: - Sweep angle: Front and rear spar locations: Main undercarriage location to be aft of the
Centre of Gravity (C of G) and its sizing, weight, and actuation system: Engine pylon installation
and mounting: Flying control surface actuator and mounting positions: Fuel tank boundaries and
system couplings employed and systems installation to ensure there is no trapped fuel within the
wing structure: The rib layout to support load transfer and structural stability of the wing box:
Materials selection and manufacturing and assembly methods stitching and bolting for CFC wing
structures, and the reference ATDA is shown in figure 2.
The major parameters of wing definition as follows: - Size: Aspect Ratio: Sweep angle: Taper Ratio:
Wing Loading and Thickness, which are derived from: - (1) LE = wing leading edge sweep angle:
(2) A = wing planform area: (3) Ĉ = Mean Aerodynamic Chord: (4) Cr = Root Chord: (5) Ct = Tip
Chord: (6) t / c = Thickness chord ratio: (7) b = Span = 2 x s (where s = semi-span): (8) S = wing
area: (9) yMAC = the y station of the Mean Aerodynamic Chord (10) Xac = aerodynamic centre of
pressure in the x axis mapped on the MAC.
For the baseline wing: - the Aspect Ratio from b² / S = 10.15: the MAC Ĉ length = 5.89m (259”) and
yMAC = 15.14m (596”) (from graphical evaluation number 1 in figure 2): LE = 35º: A = 406.481m²
(4,375ft²): Cr = 13.97m (550”): Ct = 3.81m (150”): t / c = 0.27: b = 64.76m (2,549.5”): and S =
413.02m² (640,199 inch²): the Centre of Gravity (number 2 in figure 57) was determined as 35%
root chord this allows for fuselage length growth (as per reference 4) = 4.89m (192.5”): taper ratio λ
= Ct / Cr = 0.27. The initial estimated wing loading is 10,309kN/m² (124.6lbs/ft²) within 82.7kN/m²
(1lb/ft²) of published figures for the Airbus A350: Xac = 12.07m (475”). See figure 57 for MAC,
aerodynamic centre of pressure, and C of G mapping on the reference wing. 9
The ATDA wing planform selection and aerofoil geometry.

10
Figure 2:- ATDA baseline reference wing graphical determination of MAC.
1
Croot
13.97m
(550”)
Croot
13.97m
(550”)
Ctip 3.81m
(150”)
Ctip 3.81m (150”)
b/2 32.37m (1274.5”)
MAC (Ĉ) length 5.89m (232”)
50% Chord reference wing.
100% Chord reference wing 7.69m (303”).
2
Diagonal Construction Line.
Aircraft Centre Line
CL.
yMAC (Ĉ) 15.14m (596”)
Aerodynamic centre of a subsonic swept wing is
approximately located at Xac = yMAC tan LE+ 0.25MAC
the value = 12.07m (475”) in X from reference wing tip.
3
3
Engine Pylon Centre Line.
35º

The important parameters in long range transport aircraft wing design are:-
 The Aspect Ratio (b²/S): - Increased Aspect Ratio gives improved Lift and Drag and a greater
Lift curve slope, and for subsonic transports AR values between 8-10 are considered typical.
For initial design purposes an Aspect Ratio from historical data can be used, but trade studies
using MDO toolsets are needed for definitive values. Selecting a higher value AR has beneficial
effects at high altitude cruise to give greater range and endurance, and when usable take-off
incidence is restricted by ground clearance, however this is not the case for tactical military
aircraft in low altitude high-speed flight where profile drag is the dominant factor. Historically the
Aspect Ratio has been used as a primary indicator of wing efficiency based on the square of
the wing span divided by the wing reference area. In fact the AR could be used to estimate
subsonic Lift / Drag where Lift and Drag are most directly affected by the wing span and wetted
area but for one major problem i.e. drag at subsonic speeds is composed of two parts:-
“Induced“ drag caused by the generation of lift and therefore primarily a function of the wing
span: and “Zero-lift” or “Parasitic” drag which is not related to lift but is primarily skin-friction
drag, and as such is directly proportional to the total surface area of the aircraft exposed
(“wetted”) to the air. Therefore the ratio of the wetted area of the full aircraft to the reference
wing area ( Swet / Sref ) can be used along with the aspect ratio as a more reliable early estimate
of L/D, as the wetted-area ratio is clearly dependent on the actual configuration layout. This
suggests a new parameter “Wetted Aspect Ratio” which is defined as the wingspan squared
divided by the total aircraft wetted area. This is very similar to the aspect ratio except that it
considers total wetted area instead of the wing reference area. AeroDYNAMIC™ MDO toolset
enables this to be done within its design module and compared against the Catia V5 model.
11
The ATDA wing planform selection and aerofoil geometry (continued).

 The leading edge sweep angle LE: - The greater the sweep angle the higher the lift dependent
drag and requires increased roll control for cross wind take-offs. However, it delays drag rise „M‟
and reduces the lift curve slope. For commercial transports the leading edge sweep angle
ranges between 28º to 35º with the A350 being at the top of this range and this was adopted for
the ATDA study wing as a result of AeroDYNAMIC analysis for high altitude cruise at Mach 0.89
at 39,000ft (11,887.2m).
 Taper ratio Ct / Cr: - Taper transfers load from the tip towards the root, thus increasing the
likelihood of tip stall (which gives wing droop and pitch up on a swept wing). For swept wing
increased taper gives lower trailing edge sweep, which enhances the effectiveness of trailing
edge flaps and controls (giving reduced take-off and landing speeds and improving
controllability in cross winds), the taper ratio selected for the baseline wing was 0.27 based on
AeroDYNAMIC analysis.
 Thickness: - Thick section wings incur a Profile Drag Penalty. Increasing thickness dose
however, give increased maximum lift, eases mechanisation of flaps and slats, generates a
lighter structure and presents a greater internal volume for fuel carriage.
 Camber: - Camber is added to enhance lift. It is however detrimental at low speeds.
 High Lift Devices: - There are of primary benefit on thin swept wings at supersonic speeds,
although high lift leading edge slats are used by most subsonic transports, and are incorporated
into the baseline wing design as described below.
 Winglets:- Described below which reduce induced drag.
 Aerofoil: - Section selection see figures 3 through 6 this has a major effect on drag. 12
The ATDA wing planform selection and aerofoil geometry (continued).

13
Figure 3:- Typical Breakdown of aircraft drag by form and component.
Total Drag
Parasitic
Wave / Interference
Lift
Dependant
Drag
Friction Drag
Friction Drag
Pylons and Fairings
Nacelles
Horizontal Tail
Vertical Tail
Wing
Fuselage
From this it can be seen that the largest
contributions to Friction drag are the wings and
fuselage. In this study the ATDA attempts to
reduce both:-for the wing by selection of
supercritical aerofoil selection below: and for the
fuselage by applying aerodynamic tailored
shaping for the nose and the rear sections see
the Part 3 presentation.

14
Figure 4:- Aerofoil profile selection based on Friction Drag Reduction.
Figure 4a/b:- Flow fields around 4(a) conventional aerofoil 4(b) supercritical aerofoil.
Figure 5(a) Figure 5(b)
Figure 4(c):- Sketches of root NASA SC(2) 0414 and tip NASA SC(2) 0410 aerofoil profiles.

15
Figure 5:- Aerofoil supercritical profile selection to reduce wing friction drag.
LAMINA
TURBULENT
Moment reference centre.
0.25 0.50 0.75 1.00
Reference line.
NASA SC (2) 0410 Aerofoil.
0.1
- 0.1
V= freestream
Laminar
Boundary
Layer
V= freestream Turbulent
Boundary
Layer
NB: - The Laminar boundary layer
has much lower friction drag.

16
Figure 6:- Drag Coefficient (Skin Friction Drag).
0.005
0.0 10.0 Mio 20.0 Mio
0.001
0.002
0.003
0.004
C
l
Reynolds Number
Turbulent
50% Laminar
Laminar
NB: -A substantial drag reduction is possible, even if only part of surface is Laminar.

17
The structural layout of the reference wing, and evolved wing based on the following fundamentals,
the wing has structurally to be both a span-wise and chord-wise beam and posses adequate
torsional stiffness and therefore be able to react the loads outlined in figure 7. Figure 8 illustrates
the control surfaces on the wing of the ATDA subsonic composite concept airliner, and shows how
the numerous leading and trailing edge devices occupy a significant portion of the chord. The
consequence of this is that only approximately half of the chord is available for the span-wise beam
of the torsion box, however it is the deepest portion and this is preferable for both bending and
torsion.
The primary load direction is well defined and is span-wise and therefore wings are good
candidates for the application of carbon – fibre composites providing the overall size is such that it
can be built with the minimum number of joints.
The primary wing box components of the baseline wing as is common with large transport aircraft
are:- the wing skin covers which form the lifting surface and transmit wing bending and torsion
loads, and these are stabilized with span-wise stringers to inhibit cover skin buckling, the stringers
reduce cover skin thickness requirements and hence cover weight as outlined below, (either CFC or
metallics are used for cover skins e.g. A380 uses 7449 and 7055 Al upper skins and 2024 and 2026
Al lower skins): the front and rear spars which in conjunction with the stringer stiffened skin transmit
bending and torsion loads, and consist of a web to react vertical shear loads, and edge flanges to
react the wing bending loads (and can be CFC or metallic e.g. A380 uses 7085 and 7040 Al for
spars: and ribs which maintain the aerodynamic shape of the wing cross-section, and structurally
transmit local loads chord-wise across to the span-wise torsion box, the ribs stabilize the spars and
skins in span-wise bending. In this study CFC cover skins / spars / and some ribs is the baseline.
Section 2:- Roll and layout of large aircraft wing structural members.

Figure 7:- ATDA Wing complexity as a complete structural component.
Aircraft Sizing Determined by Wing
Architecture (e.g. Tail sizing: Landing
gear geometry: Belly Fairing: etc.
Complex systems installation
(Fuel: Pneumatics: Electrical and
FTI provision)
Aircraft Configuration influenced by
wing definition (e.g. C of G: Ground
Line: Cargo hold position: PAX
evacuation: etc.).
Determines Aircraft High and Low
speed performance.
Complex assembly and equipping.
Critical Structures and Systems
Integration (Root Joint: Landing
Gear: Pylon: and Moveable's).
Optimising aerodynamics / structural
geometry (Twist: Taper: Camber:
Sweep: and Gulling: etc.).
Particular Risk Mitigation (Bird
Strike: WTF: Lightening Strike:
UERF).
Managing High Load Inputs
(Pylon: Landing Gear: Flaps
Integration).
Optimise design and manufacture of
thick and complex structures (e.g. FTE:
Bottom Cover Skin: etc.).
Combined Loading effects
(Ground: Manoeuvre: Gust:
and Flight loads).
18

19
Figure 8:- High lift devices and control surface layout of the ATDA concept airliner.
Six Outboard Leading edge slats.
Droop nose Leading edge slat.
Two Inboard
Spoilers with
droop function.
Five Inboard
Spoilers with
droop function.
Outboard Flap
single pivot.
Inboard Flap
single pivot.
All Speed Aileron.
Low Speed Aileron.
Rudder.
(Planform area 15m²)
Port Elevator
(Planform area 10 .18m²)
Stbd Elevator.
(Planform area 10.18m²)

20
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 9(a)(b)(c) 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 transition
solutions to the drop off zones as shown in figure 10(a) through (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 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 stiffeners are used as
detailed below and shown in figures 11, 12, and 13 for the baseline ATDA reference structure.
Roll and layout of large aircraft wing structural members ( CFC cover skins).

Figure 9(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.
The 4 Primary Ply Orientations Used for Wing Skin Structural Plies.
21

Figure 9(b):- Fibre Orientation Requirements for CFC Wing Skins / covers.
22
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

23
Figure 9(c):- ATDA Design Load Cases for PRSEUS Lower Wing Skins / Covers.
VERTICAL GUST
(CLEAN WING)
VERTICAL GUST
JACKING
VERTICAL GUST*
*ENVELOPE OR QUISI-FLEXIBLE
WING GUST
ALTITUDE WEIGHTED
TUNED GUST
STR1
STR2
STR3 STR4
STR5
STR6
STR7
STR8
STR9
Note:-
Rib 0 C = Closure Rib:
STR1 = Stringer One (etc.)
This is the PRSEUS ATDA Skin layout.

Fig 10(a):- ATDA Structural Ply Thickness Zones Upper Wing Cover Skin Baseline
24
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 10(b):- ATDA Structural Ply Thickness Zones Upper Wing Cover Skin PRSUES.
25
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 10(c):- ATDA Structural Ply Thickness Zones Lower Wing Cover Skin Baseline
26
(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 10(d):- ATDA Structural Ply Thickness Zones Lower Wing Cover Skin PRSEUS.
27
(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

28
Fig 11(a)/(b):- 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(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 11(a) for the ATDA wing skins.
Where the wing chord thickness is much greater than 11.8 inches.
Figure 11(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.
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.

29
Fig 11(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.

Composite cover skin stringer types: -
 “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. 30
Roll and layout of large aircraft wing structural members (CFC cover skins).

31
Figure 12:- Baseline composite stringer selection based on design experience.
“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
“Z” Section Stringer (mechanically attached to
provide additional stiffness for out of plane
loading).
“L” Section Stringer (bonded or
mechanically attached panel breaker).

Composite wing cover skin stringer radius fillers (noodles):-
Radius fillers are necessary in T - and I – type composite stiffeners and spars. See figure 12
(previous slide) 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. Research 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. Figure 13 and 14
show how peel stresses and manufacturing weight can be reduced in stringer design. Figures 15(a)
through (d) shows the ATDA lower cover skin stringer arrangement and special considerations for
the inspection cut outs, either side of which coaming stringers are installed.
32

33
Figure 13:- Composite Stringer design based on design / research experience.
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.
8.5 N/mm²
Square Edge flange toe.
Radius Edge flange toe.
7.5 N/mm²
30º Chamfer flange toe
(selected for PRSUES
Flange ATDA).
5 N/mm²
4 N/mm²
6º Chamfer flange toe strip
(desired for developed
PRSEUS ATDA but could
give rise to stitching
induced delamination ).
1 N/mm²
6º Chamfer flange toe and capping.
TRADE STUDY.
 REDUCTION OF PEEL STRESS
AT TOE OF FLANGE.
 REDUCTION IN STRINGER
MASS.
 INCREASED MANUFACTURING
COSTS.
 ISSUES WITH REPAIR /
FASTENERS.

Fig 14(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 example CFC stringers and CFC ribs.
b) TYPICALASSEMBLY OF
PRE-CURED DETAILS
34

 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 and as with co-
curing the bond is formed during the autoclave cycle, this method has been used on some CFC
fighter wing spars which were co-bonded to the one wing cover skin, and is proposed for the
ATDA baseline, as this technology has used to bond the wing cover skin stringers for current
large CFC transport aircraft wings, see section 7. 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.
35

Fig 15(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.
36

37
Fig 15(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 15(c):- ATDA lower cover skin with PRSEUS coaming stringer layout and ports.
38
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.

39
Fig 15(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.

Conventional Co-bonding of laminated two-dimensional composites are not suitable for applications
where trough thickness stresses may exceed the (low) tensile strength of the matrix (or matrix /
fibre bond) and in addition, to provide residual strength after anticipated impact events, two–
dimensional laminates must therefore be made thicker than required for meeting strength
requirements. The resulting penalties of increased structural weight and cost provide impetus for
the development of more damage-resistant and tolerant composite materials and structures.
Considerable improvements in damage resistance can be made using tougher thermoset or
thermoplastic matrices together with optimized fibre / matrix bond strength. However, this approach
can involve significant costs, and the improvement that can be realized are limited. There are also
limits to the acceptable fibre / matrix bond strength because high bond strength can lead to
increased notch-sensitivity.
An alternative and potentially more efficient means of attaching the stringer to the cover skins and
increasing damage resistance and through-thickness strength is to develop a fibre architecture in
which a proportion of fibers in the composite are orientated in the z-direction. This fibre architecture
can be obtained, for example, by three-dimensional weaving or three-dimensional breading.
However a much simpler approach is to apply reinforcement to a conventional two-dimensional
fibre configuration by stitching: although, this dose not provide all of the benefits of a full three-
dimensional architecture. In all of these approaches, a three dimensional preform produced first
and converted into a composite by either RTM / VARTM, or CAPRI (see later in this presentation).
Even without the benefits of three-dimensional reinforcement, the preform approach has the
important advantage that it is a comparatively low-cost method of manufacturing composite
components compared with conventional laminating procedures based on pre-preg. 40

41
The structural benefits of 3-D stitched for stringer over conventional laminates.
(a) Lock stitch (b) Modified Lock stich
(c) Chain stitch
Needle
Thread
Bobbin
Thread
Needle
Thread
Bobbin
Thread
Figure 16:-Schematic diagram of three commonly used stitches for 3-D reinforcement.
Indeed, preforms for resin transfer molding (RTM) and other liquid molding techniques are often
produced from a two dimensional fibre configuration by stitching or knitting Stitching was selected
for the ATDA wing and fuselage.
Stitching:- This is best applied using an industrial-grade sewing machine where two separate
yarns are used. For stitching composites, the yarns are generally aramid (Kevlar), although other
yarns such as glass, carbon, and nylon have also been used. A needle is used to perforate a pre-
preg layup or fabric preform, enabling the insertion of a high–tensile-strength yarn in the thickness
direction. In the case of the PRSEUS process a Vectran thread impregnated with epoxy resin is
used. The yarn, normally referred to as the needle yarn, is inserted from the top of the layup /
preform, which is held in place using a presser foot. When the yarn reaches the bottom of the
layup / preform it is caught by another yarn, called the bobbin yarn, before it re-enters the layup /
preform as the needle is withdrawn from the layup / preform, thus forming a full stich. The layup /
preform, is then advanced a set distance between the presser foot and a roller mechanism before
the needle is used to apply the next stitch. This process is repeated to form a row of stitches.
Figure 16 shows the various types of stitches commonly used to create z-direction reinforcement.

Among the three stitches shown in figure 16, the modified lock stitch in which the crossover knot
between the bobbin and needle threads is positioned at either laminate surface, to minimize in-
plane fibre distortion is considered the best, and is the preferred method. Apart from improving z-
direction properties, stitching serves as an effective means of assembling preforms of dry two-
dimensional tape or cloth, for example, attaching stringers to skin preforms, that can then be
consolidated using liquid molding.
Mechanical Properties Improvements: - (1) Out-of-Plane properties are significantly improved by
stitching, increasing the interlaminar delamination resistance for fibre reinforced plastic laminates
under mode I (tensile loading KIC) and to a lesser extent mode II (shear loading KIIC) loadings. In
order achieve this, the stiches need to remain intact for a short distance behind the crack front and
restrict any effort to extend the delamination crack. With such enhanced fracture toughness stitched
laminates have better resistance to delamination cracking under low energy, high energy and
ballistic impacts as well as under dynamic loading by explosive blast effects. Stitched laminates
also possess higher post-impact residual mechanical properties than non-stitched laminates.
Studies (ref 6) have shown that the effectiveness of stitching for improving residual strength is
dependent on factors such as the stitch density, stitch type, and stitch thread. Although the best
improvement in compression post impact strength has been found in relatively thick laminates, and
though similar improvements in residual strength have been observed in toughened matrix
laminates the latter is two to three times more expensive than stitching. Stitching also improves
shear lap joint strength under both static and cyclic loading, largely due to reducing the peel
stresses. Stitching can delay the initiation of disbonds and provide load transfer even after bond line
failure. Stitching is also effective in suppressing delamination due to free edge effects.
42
The structural benefits of 3-D stitched and pinned composites over conventional laminates.

(2) In-Plane properties of a two dimensional composite laminate can also be affected by stitching,
due the introduction of defects in the final laminate during needle insertion or as a result of
presence of the stitch yarn in the laminate. These defects may occur in various forms including
broken fibres, resin-rich regions, and fine scale resin cracking. Fibre misalignment however
appears to have the greatest detrimental effect on mechanical properties, particularly under in
plane tensile and compressive loading.
In order to keep defects resulting from stitching to a minimum, careful selection and control of the
stitching parameters (including:- yarn diameter: yarn tension: yarn material: stitch density: etc.), are
essential. Analysis of the effects of stitching on in-plane material properties of two dimensional
composite laminates in general have been somewhat inconclusive (ref 6), with studies showing that
stiffness and strength of the composites under tensile and compressive loadings can be either
degraded, unchanged, or improved with stitching, depending on the type of composite, the stitching
parameter, and the loading condition. The improvements in tensile and compressive stiffness have
been attributed to the increase in fibre / volume fraction that results from a compaction of the in-
plane fibres by stitching. The enhancement in compressive strength is attributed to the suppression
of delamination's. The stiffness in tension and compression is mainly degraded when in-plane fibres
are misaligned by the presence of the stitching yarn in their path. Premature compressive failure
can result from the stitching being too taut, which in turn can cause excessive crimping of the in-
plane fibres. Conversely, insufficient tension on the stitching yarn can cause the stitches to buckle
under consolidation pressures and render them ineffective as a reinforcement in the thickness
direction, which was the original intention. Tensile strength however is normally degraded due to
fibre fractures arising from damage inflicted by the stitching needle. 43

Enhancements of tensile strength, which has been observed, is attributed to an increase in fibre /
volume fraction resulting from compaction of the in-plane fibres by the stitching. The in-plane
fatigue performance is also considered to be degraded due to the same failure mechanisms
responsible for degradation of their corresponding static properties.
Finally, it appears that the flexural and interlaminar shear strengths of two-dimensional laminates
may also be degraded, unchanged, or improved with stitching. In general, the conflicting effects of
stitching, in increasing fibre content and suppressing delamination, on one hand, and introducing
misalignment and damage to in-plane fibres on the other, are possibly responsible for the reported
behaviors.
Z-Pinning:- Was also considered, this is a simple method of applying three-dimensional
reinforcement with several benefits over stitching. However, unlike stitching, z-pinning cannot be
used to make preforms and therefore is included here for completeness. In the z-pinning process,
thin rods are inserted at right angles into a two-dimensional carbon / epoxy composite laminate,
either before or during consolidation. The z-rods can be metallic, usually titanium, or composite,
usually carbon / epoxy, and these are typically between 0.25mm (0.0098 inch) and 0.5mm (0.0197
inch) in diameter. These rods are held with the required pattern and density in a collapsible foam
block that provides lateral support, this prevents the rods from buckling during insertion and allows
a large number of rods to be inserted in one operation. The z-rods are typically driven into the two-
dimensional composite by one of two methods as shown in figure 17. The first method (figure
17(a)) involves placing the z-rod laden foam on top of an uncured pre-preg and autoclave curing.
During the cure, the combination of heat and pressure compacts the collapsible foam layer, driving
the rods orthogonally into the composite. 44

45
Figure 17 (a)/(b):- Z-Pinning process an alternative to stitching.
TOOL
Vacuum Bag
Prepreg Composite
Z-Fibre Preform
TOOL
PRESSURE
TOOL
Remove & Discard Foam
Cure Z-Pinned Composite
Stage 1:- Place Z-Fibre Preform on top of Prepreg and then enclose in vacuum
bag.
Stage 2:- Standard cycle or debulk cycle, heat and pressure compact preform
foam, forcing the Z-pins into the Prepreg composite.
Stage 3:- Remove compacted preform foam and discard Finish with cured Z-
pinned composite.
Figure 17(a). Figure 17(b).
Remove Used
Preform
Uncured Composite
Z-Fiber Preform
Ultrasonic Insertion Transducer
(a) Primary insertion stage and residual preform removal.
(b) Secondary insertion stage.

When curing is completed, the residual foam preform is then removed and discarded, and the z-
rods sitting proud of the surface of the cured laminate are sheared away using a sharp knife.
The second method uses a purpose built ultrasonic insertion transducer to drive the z-rods into the
two-dimensional composite and is shown schematically in figure 17(b). This is a two stage process,
and during the first stage the preform is only partially compacted using the ultrasonic insertion
transducer, and thus the z-rods are not fully inserted. The residual foam is then removed, and a
second insertion stage is carried out with the ultrasonic insertion transducer making a second pass
to complete the insertion of the z-rods. If the z-rods are not flush with the part surface, the excess is
sheared away. In principle, the part to be z-pinned could take on any shape provided there is an
appropriate ultrasonic insertion transducer. Research indicates that the ultrasonic insertion
technique can be used to insert metallic pins into cured composites for the repair of delamination's,
although a considerable amount of additional damage to the parent material results and further
trade studies are required to determine its true viability.
Of the two z-pinning insertion methods the vacuum bag method is more suitable when a large or
relatively flat and unobstructed area is to be z-pinned. The ultrasonic method is more suitable for z-
pinning localized or difficult to access areas by configuring and shaping an appropriate ultrasonic
insertion transducer.
Mechanical Properties Improvements: - (1) Out-of-Plane properties indicate a significant
improvement in both mode I (tensile loading KIC) and mode II (shear loading KIIC) fracture
toughness, achieved through z-pinning based on published data, which would translate into
superior damage resistance and tolerance, as well as improved skin stiffener pull out properties.
46

(2) In-Plane properties current research (ref 6) indicates that the improvements in out-of-plane
properties are achievable without much if any, sacrifice of in-plane properties, although other work
indicates that the z-pins can introduce excessive waviness to the in-plane fibres, resulting in
compressive properties being severely degraded. As with the stitched 3-d reinforcement, the
degree to which the in-plane properties are detrimentally affected, and the out-of-plane properties
are improved, depends on the pinning parameters, such as pinning density and pattern
configuration.
Z-direction reinforcement:- Research into z-direction reinforcement of traditional 2-D laminate
mechanical properties has been particularly extensive, and the impetus is derived from the potential
of both stitching and z-pinning to address the poor out-of-plane properties of conventional 2-D fibre
reinforced composites, in a cost-effective method. The amount of z-direction reinforcement needed
to provide a substantial amount of out-of-plane property improvement is small and values of 5% are
typical. The improvements in fracture toughness resulting from these processes mean that higher
design allowables could be used in the design of composite structures. Stitched and z-pinned
components could reduce the layup complexity, and weight for structures subjected to: - the risk of
impact damage (e.g. due to dropped tools), high peel stresses (e.g. in joints and at hard points),
and cut-outs (e.g. edges and holes) that are difficult to avoid in aircraft design. Stitching and z-
pinning also provide the opportunity for parts integration to be incorporated into the production of
composite components, thus improving the ease of handling in automated assembly processes,
and the overall cost-effectiveness of the manufacturing process. When used in conjunction RIM /
RTM stitching provides pre-compaction of the preform that enables reduces the mold clamping
pressures while ensuring a high fibre / volume fraction in the finished product. 47

48
The PRSEUS structural concept was developed for the HWB fuselage pressure and bending load
issues that have held back the development of this aircraft type. This ATDA study examines the
feasibility of using the same structural concept to attach stringers, and frames, as well as lower
cover rib feet to reduce composite skin thickness / weight in a large conventional configuration
transport aircraft.
As conceived in NASA/CR-2011-216880, the PRSEUS panels were designed as a bi-directionally
stiffened panel design, to resist loading where the span wise wing bending are carried by the frame
members (like skin / stiffeners on a conventional transport wing), and the longitudinal (fuselage
bending loads in a HWB aircraft), and pressure loads being carried by the stringers. In the ATDA a
similar concept be used to take the bending, torque, and fuel pressure loads in a conventional wing,
and fuselage pressure and bending loads. Based on the NASA sponsored Boeing stitched / RFI
wing demonstrator program of 1997, which produced 28m (92ft) structure 25% lighter and 20%
cheaper than an equivalent aluminium structure the answer would appear to address the structural
weight reduction desired.
The highly integrated nature of PRSEUS is evidenced by figure 18 (a)(b) which shows the stringer
structural assembly of dry warp-knit fabric core, pultruded rods, materials, which are then stitched
together to create the optimum structural geometry. Load path continuity at the stringer – frame
intersection is maintained in both directions. The 0º fiber dominated pultruded rod increases local
strength / stability of the stringer section while simultaneously shifting the neutral axis away from
the skin to enhance overall panel bending capability. Stringer elements are placed directly on the
IML (Inner Mold Line), skin surface and are designed to take advantage of carbon fiber tailoring by
placing bending and shear – conductive layups where they are most effective.

49
 All detailed parts were constructed from AS4 standard modulus 227,526,981kPa (33,000,000
lb/in²) carbon fibers and DMS 2436 Type 1 Class 72 (grade A) Hexflow VRM 34 epoxy resin.
Rods were Toray unidirectional T800 fibres with a matrix of 3900-2B resin. The preforms were
stitched together using a 1200 denier Vectran thread, and infused with a DMS2479 Type 2 Class
1 (VRM-34) epoxy resin (dimensions in mm). PRSEUS Upper wing cover skin stringer is shown
as a typical example, each stack is of 18 ply layup (0.21336mm ply) giving a ply stack thickness
of 4.0mm in the following configuration: -
Pultruded rod 0º
Each stack: - (-45º/+45º/-45º/+45º/-45º/0º/90º/0º/90º/90º/0º/90º/0º/-45º/+45º/-45º/+45º/-45º).
The stringer stack is overwrapped around the pultruded rod and the web is formed by stitching
the overwrapped stack together with two stitching runs 14.8mm from the radius ends to allow
needle clearance and any defects that the stitching. The flanges are formed from continuations
of the same stack and are stitched to the tear strip (same as a capping strip) with a braided
noodle cleavage filler. Two stitching runs secure each flange to the tear strip and skin, again the
inboard stitching runs are offset 8mm from the radius ends, and the outboard runs are 15mm
inboard of the edge. For standard wing stringers the flange with is 77mm and the stringer height
is 77mm overall.
 The PRSEUS Coaming Stringers have an 18 ply stack layup of 0.21336mm ply giving a
thickness of 4.0mm, in the following configuration:-
Each stack: - (-45º/+45º/-45º/+45º/-45º/0º/90º/0º/90º/90º/0º/90º/0º/-45º/+45º/-45º/+45º/-45º).
Flange Stitching runs are angled at 45º inboard, and normal to the flange surface outboard. The
height is 126mm and the flange with is 120mm.
My construction of the ATDA study PRSEUS wing skin stringers.

50
Figure 18(a):- Section layout of a typical ATDA study PRSEUS wing skin stringers.
Flange Stitching runs
and vectors
30º Chamfer of the Stringer
flange to reduce peel stress
Web Stitching runs
and vectors
Stringer Ply stack
Overwrap
Pultruded Rod (10mm Dia)
Lower Wing Cover
Skin Section
Tear Strip
C/L

51
Figure 18(b):- Section layout of the ATDA Study PRSEUS Coaming Stringers.
Web Stitching runs
and vectors
30º Chamfer of the Stringer
flange to reduce peel stress
Flange Stitching runs
and vectors
Stringer Ply stack
Overwrap
Pultruded Rod (10mm Dia)
Lower Wing Cover
Skin Section
Tear Strip
C/L

The stitching is used to suppress out-of-plane failure modes, which enables a higher degree of
tailoring than would be possible using conventional laminated materials.
In addition to the enhanced structural performance, the PRSEUS fabrication approach is ideally
suited to compound curvatures as may be found in advanced transport concepts. The self
supporting stitched preform assembly feature that can be fabricated without exacting tolerances
and then accurately net molded in a single oven-cure operation using high precision OML (Outer
Mold Line) tooling is a major enabler in low cost fabrication. Since all of the materials in the stitched
assembly are dry, there is no out-time or autoclave limitations as in a prepreg system, which can
restrict the size of an assembly as it must be cured within a limited processing envelope.
Resin infusion is accomplished using a soft-tooled fabrication method where bagging film conforms
to the IML, surface of the preform geometry and seals against a rigid OML tool, this eliminating the
costly internal tooling that would be required to form net-molded details. The manufacture of
multiple PRSEUS panels for the NASA/CR-2011-216880 program validated this feature of the
concept, and demonstrated that the self supporting preform that eliminates interior mold tooling is
feasible for application to the geometry of the airframe. An example of my stitched wing rib integral
flange assembly using PRSUES technology is shown in figure 19(a)(b), and the integration of the
rib / spar assembly is shown in figure 20 and my developed PRSEUS wing stringers for this ATDA
airframe project are shown in figures 18 (NB analysis under baseline loading has enabled a
reduction in flange size over previous iteration from 172mm to 120mm), the lock stitch stitching
machine, and assembly is shown in figures 21 and 22, respectively this will also be used for frame
and rib stitching.
52
PRSEUS stringer and rib cleat design and stitching to respective cover skins.

53
Figure 19(a):- Composite Rib 31 Stitched Stub -Rib Preform assembly.
Tare Strip
(1.5mm)
Figure 19(a)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 19(a)ii
Low level fuel transfer holes.
Figure 19(a)iii
Aft Coaming Stringer Section
Section of lower cover skin
(representative)

54
Figure 19(b):- Composite Rib 31 Stitched Stub-Rib PRSEUS Coaming stringers.
Figure 19(b)i Side view on (B)
Figure 19(b)ii Plan view
Figure 19(b)iii 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

55
Figure 20:- 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.

56
Figure 21:- RS 545 and RS 543 Lock stitching machines proposed for the ATDA stringers.
Figure 21(a):- The RS 545 Lock stitching machine mounted on a KUKA
robot used in a KL 500 robot sewing workstation by Eurocopter to
stitch I – beam webs. Reference KSL Composites Europe 2014 VDMA
forum.
Figure 21(b):- Detailed view of the stitching head proposed
for the two rows of stitching on PRSEUS stringer webs.
Figure 21(d):- Detailed view of the stitching head proposed
for the two rows of stitching on PRSEUS stringer flanges.
Figure 21(c):- The RS 545 Lock stitching machine mounted on a KUKA
robot used in a KL 500 robot sewing workstation by Eurocopter to
stitch I – beam flanges.

57
Figure 22:- Schematic factory of the future proposal for stitching wing structures.
Stitching
Cutting
Tooling
Assembly
Trim and Drill
*Note Horizontal PRSEUS wing assembly this study covers not only stitched stringers but also
stitched rib cleats and fuselage frames.

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 23(a), and in additional soft
tooling (bagging aides) are also used figure 23(b) and in the Boeing Controlled Atmospheric
Pressure Resin Infusion process figure 23(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 using this process will be a major focus of this project, and are covered in the
companion Composite Design and Capability Research presentation.
PRSEUS stringer /cover skin and rib cleat post assembly processing overview.
58

59
Figure 23:- Boeing Controlled Atmospheric Pressure Resin Infusion (CAPRI) process.
Fig 23(b):- Soft tooling (bagging aids) installation over stiffeners.
Fig 23(a):- Robotic stitching of dry preform assembly.
Fig 23(c):- Vacuum bag installation over dry preform assembly.
NASA Public released report concept.

To maintain the aerodynamic smoothness of the external surface Outer Mold Line, of the composite
wing cover skins, the surface is always laid on the tooling face and non-structural surface ply is
added at the tool interface, to ensure smooth OML surface.
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 aircraft with wing mounted engines is shown in figure 24(a)/(b), 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. 60

Zone 3 Indirect effects.
Zone 2 Swept stroke.
Zone 1 Direct strike.
Lightening Strike
Zones on an
aircraft with wing
mounted engines.
Figure 24(a):- Lightening strike risks to composite wing structures with podded engines.
61

62
Figure 24(b):- Lightening strike risks to composite podded engine aircraft structures.
Zone Key.

63
Lightening effects can be divided into direct effects and indirect effects:-
 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.
 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.
Reference wing box layout key structural members (CFC cover skins).

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.
The main method of lightening strike protection for composite aircraft wing structures is illustrated in
figure 25, this commercial aircraft system will be employed in this study (see also ref 5).
64
Roll and layout of large aircraft wing structural members (cover skins).

65
Figure 25:- Lightening strike protection of composite commercial aircraft wing.
Reference Cranfield MSc lecture notes AIAA ES, and ref 4&5.
Lightening Strike on CFC airframe wings, as described above
requires the following protection:-
 Wing (with exception to wing tips):
 Copper strip embedded in the ply lay up:
 Fastener heads exposed.
Copper grid
Dielectric
Cap
seal
Stringer
CFC Skin

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 ref 6. 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 26(a)/(b)/(c)). 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 see figure 27. 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 Design
Ultimate Load (DUL). These design criteria are critical airworthiness clearances ACJ 25.603 and
FAA AC20.107A (Composite Aircraft Structures) a full treatment is given below.
66

From practical experience damage to composite structures due to accidental damage on the flight
line or weather damage cannot be eliminated, therefore composite airframe structures must be
designed with adequate reserves to function safely after damage i.e. be damage tolerant.
Designing for damage tolerance includes selecting damage resistant materials (in particular matrix
resin systems), identifying sources and types of damage, knowledge of damage propagation
mechanisms, and criticality of damage. Damage tolerance in composite airframes depends on
details such as ply layup, frame / rib and stringer pitch attachment details, crack arrest features,
structural redundancy etc. By understanding damage and being able to predict the growth rate, as
well as being able to detect critical damage enables the designer to design a structure that can
withstand given levels of damage that can be detected within regular inspection intervals.
Chart 1 (ref 21) categorises the types of damage which can occur to a composite airframe into five
categories of damage severity as detailed below:-
 Category 1:- is allowable damage that may go undetected by scheduled inspections which
includes;- classical low energy BVID; allowable manufacturing defects; and in service damage
which dose not result in degradation of the ultimate load carrying capacity over a reliable
service life of the airframe.
 Category 2:- is defined as damage that can be reliably detected by scheduled or directed
inspections. Typical examples of this type being;- visible impact damage; deep scratches;
detectable delamination or disbonding; the resulting residual strength of the composite structure
resulting from this damage must be significantly above the limit load level for the chosen
inspection interval.
67
Classification of impact damage by severity for composite aircraft structures.

68
Chart 1:- Design load levels vs damage severity for composite aircraft structures.
Design
Load
Level
1.5 Factor
of Safety.
Ultimate
Limit
~ Maximum load
per lifetime.
Continued
safe flight.
Allowable
Damage Limit
(ADL)
Critical Damage
Threshold
(CDT)
Increasing Damage Severity.
Category 1 Damage:- BVID:
Designed for Mfg damage.
Category 2 Damage:- VID: requiring
repair per normal inspection process.
Category 3 Damage:- Obvious damage
found first few flights after occurring:
requiring immediate repair.
Category 4 Damage:- Discrete
damage obvious to flight crew :
requiring repair post flight.
Category 5 Damage:-
Anomalous damage not
covered in design but known
to operations: requiring
immediate repair.

 Category 3:- is damage detectable within a few operational flights by ramp servicing personnel
this would include;- large visual impact damage; damage easily detected by a pre-flight walk
around or drone visual inspection. The design of the airframe to meet Category 3 damage
requires features that provide a sufficient damage tolerance capability that it retains limit load
levels for a short time detection interval.
 Category 4:- is discrete damage known to the pilot that limits flight manoeuvres;- this includes
damage due to bird strike; tyre-burst; or sever in-flight hail. This requires sufficient damage
tolerance in the airframe to complete the flight.
 Category 5:- is severe damage of the airframe caused by ground or flight conditions not
covered by design criteria this my include;- severe impact with a ground vehicle with an aircraft
fuselage; flight overload condition; in-flight loss of a component e.g. control surface; hard
landings; or blunt impacts. The criticality of this category is highlighted by the fact that there are
no clear visual prior indicators of damage.
Often impacts with ground vehicles can generate Category 2 or 3 damage, which must be
managed with a Certification process i.e. using substantiated scheduled inspections for detection,
and immediate repair action when detected. Alternatively such an impact may result in Category 5,
damage which must be reported and repaired immediately, although this category is outside the
immediate aircraft design Certification process the need to report such damage is identified in
documents such as AMC 20-29. Therefore the boundaries between Category 2/3 and Category 5
damage should be clearly understood.
69
Classification of impact damage by severity for composite aircraft structures.

70
Figure 26(a):- Structural damage risks to composite wing structures.
Dropped hand
tool - 8J
All internal structure - 8J
Gravity refuelling point - 30J
Fig 26(a) i:- ATDA Upper wing cover skin Fig 26(a) ii:- ATDA Lower wing cover skin
Engine debris
- 160J zone
Runway stones - 17J
(6mm 140 Knts) zone
Dropped hand
tool - 8J zone
Low Energy Impact Damage Threats:-
 Barely Visible Impact Damage (BVID) threat from:- dropped hand tools: runway stones etc.
Solution:- Design for known threat level: Incorporate surface plies such as Kevlar or S2
glass cloth: Use hybrid ply lay-ups combining UD and woven surface plies.

Figure 27:- Woven Cloth Classifications and surface ply BVID protection options trades.
71

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 airliners 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 28, 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 29, 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. 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 30, where there are
bolted joints, tooling hole sites, to react pressure differentials at fuel tank boundaries, and spar
section splicing, figures 31 to 33 show preliminary outboard wing spar design, and figure 34 shows
a spar splice joint concept and 35 shows the outboard spar assembly. The chord-wise location of
the spars is restricted by the numerous leading and trailing edge devices that occupy a significant
portion of the wing chord as shown in figure 8. Generally the front spar should be as far forward as
possible, subject to: - (a) The local wing depth being adequate to enable vertical shear loads to be
reacted efficiently: (b) Adequate nose chord space for leading edge devices and their operating
mechanisms, and de-icing systems. 72
Roll and layout of large aircraft wing structural members (CFC wing spars).

Therefore the front spar of a two-spar wing torsion box is usually located in the region of 12-18% of
local wing chord.
In two spar modern transport wings the rear spar should be as far aft as possible being limited to
being in front of the trailing edge flaps, control surfaces, and spoilers, and their operating
mechanisms. Thus the rear spar is typically at 55-70% of the chord.
Any intermediate spars are usually spaced uniformly across the chord-wise section except where a
particular pick-up point is required for a powerplant as in the case of the A300, A330/A340/A380,
and the B-747, and auxiliary spars are used to support main landing gear attachment and some
trailing edge surfaces.
Although there have been cases where the width of the structural torsion box has been limited to
give rise to high working stresses in the distributed flanges, and consequent good structural
efficiency, this is achieved at the expense of potential fuel volume. This approach therefore has not
been adopted in these trade studies as the wing is to be employed as a primary integral fuel tank,
and in general for a transport aircraft the opportunity should always be taken to maximize the
potential fuel volume for future growth development.
Spar location should not be stepped in plan layout as this gives rise to offset load paths, but a
change of sweep angle at a major rib position is acceptable.
Returning briefly to metallic ribs, current practice is to integrally machine them from aluminium alloy
rolled or forged plate, this method of construction gives weight savings at reasonable cost over
fabricated construction. Each section of spar has a continuous horizontal stringer crack stopper
introduced approximately 1/3 of the way up the shear web from the predominantly tension flange.
73
Roll and layout of large aircraft wing structural members (CFC wing spars).

74
Figure 28:- Composite spar manufacture and assembly example.
CFRP Spar C section with apertures for control surface guide rails.
Wing torsion box section with “C” section spars, ribs, and edge control
surface attachment fixtures.

75
Figure 29:- 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 ° 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

76
Figure 30:- Carbon Fibre Composite ply orientations in wing spars continued.
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

Figure 31:- 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.
77

Figure 32:- FATA Outboard Port CFC Wing Spar as layup and finished part (preliminary).
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 Thick Zone
(18 Plies)
7mm Thick 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.
78

Figure 33:- ATDA Outboard Port / Stbd CFC Wing Spar preliminary part layup.
Zone (1):- 4mm THK 18 plies see Table 6(a)
Zone (2):- 7mmTHK 32 plies see Table 6(b)
Zone (3):- 10mmTHK 46 plies see Table 6(c) (parts 1 and 2)
14ply symmetrical drop
14ply symmetrical drop
79
Based on Carbon / Epoxy 3501/6 QI unidirectional composite tape
material with a ply thickness of 0.21336mm (see table 2(a),2(b),and 2(c)).

Structural Ply No Only. Material
Nominal ply thickness
(mm)
Ply orientation
1 Fabric 0.25000 45º/135º
2 UD 0.21336 135º
3 UD 0.21336 45º
4 UD 0.21336 90º
5 UD 0.21336 45º
6 UD 0.21336 135º
7 UD 0.21336 45º
8 UD 0.21336 135º
9 UD 0.21336 45º
10 UD 0.21336 45º
11 UD 0.21336 135º
12 UD 0.21336 45º
13 UD 0.21336 135º
14 UD 0.21336 45º
15 UD 0.21336 90º
16 UD 0.21336 45º
17 UD 0.21336 135º
18 Fabric 0.25000 45º/135º
80
Table 2(a):- Outboard Leading Edge Spar Zone (1) 18 ply stacking sequence.

Structural Ply
No Only.
Material
Nominal ply
thickness (mm)
Ply
orientation
Structural
Ply No Only.
Material
Nominal ply
thickness (mm)
Ply
orientation
1 Fabric 0.25000 45º/135º 17 UD 0.21336 45º
2 UD 0.21336 45º 18 UD 0.21336 135º
3 UD 0.21336 135º 19 UD 0.21336 45º
4 UD 0.21336 45º 20 UD 0.21336 135º
5 UD 0.21336 135º 21 UD 0.21336 45º
6 UD 0.21336 45º 22 UD 0.21336 90º
7 UD 0.21336 90º 23 UD 0.21336 45º
8 UD 0.21336 45º 24 UD 0.21336 135º
9 UD 0.21336 135º 25 UD 0.21336 45º
10 UD 0.21336 45º 26 UD 0.21336 90º
11 UD 0.21336 90º 27 UD 0.21336 45º
12 UD 0.21336 45º 28 UD 0.21336 135º
13 UD 0.21336 135º 29 UD 0.21336 45º
14 UD 0.21336 45º 30 UD 0.21336 135º
15 UD 0.21336 135º 31 UD 0.21336 45º
16 UD 0.21336 45º 32 Fabric 0.25000 45º/135º
81
Table 2(b):- Outboard Leading Edge Spar Zone (2) 32 ply stacking sequence.

Structural Ply No Only. Material Nominal ply thickness (mm) Ply orientation
1 Fabric 0.25000 45º/135º
2 UD 0.21336 135º
3 UD 0.21336 45º
4 UD 0.21336 135º
5 UD 0.21336 45º
6 UD 0.21336 135º
7 UD 0.21336 45º
8 UD 0.21336 135º
9 UD 0.21336 45º
10 UD 0.21336 135º
11 UD 0.21336 45º
12 UD 0.21336 135º
13 UD 0.21336 45º
14 UD 0.21336 90º
15 UD 0.21336 45º
16 UD 0.21336 135º
17 UD 0.21336 45º
18 UD 0.21336 90º
19 UD 0.21336 45º
20 UD 0.21336 135º
21 UD 0.21336 45º
22 UD 0.21336 135º
23 UD 0.21336 45º
82
Table 2(c):- Outboard Leading Edge Spar Zone (3) 46 ply stacking sequence (part 1).

Structural Ply No Only. Material Nominal ply thickness (mm) Ply orientation
24 UD 0.21336 45º
25 UD 0.21336 135º
26 UD 0.21336 45º
27 UD 0.21336 135º
28 UD 0.21336 45º
29 UD 0.21336 90º
30 UD 0.21336 45º
31 UD 0.21336 135º
32 UD 0.21336 45º
33 UD 0.21336 90º
34 UD 0.21336 45º
35 UD 0.21336 135º
36 UD 0.21336 45º
37 UD 0.21336 135º
38 UD 0.21336 40º
39 UD 0.21336 135º
40 UD 0.21336 45º
41 UD 0.21336 135º
42 UD 0.21336 45º
43 UD 0.21336 135º
44 UD 0.21336 45º
45 UD 0.21336 135º
46 Fabric 0.25000 45º/135º
83
Table 2(c):- Outboard Leading Edge Spar Zone (3) 46 ply stacking sequence (part 2).

84
Proposed C section wing spar section splice joint design methodology.
Due to the ± 5% thickness control limitations on composite parts the spar splice joints will have to
be multi component adjustable assemblies. Using a mirrored internal female tool on which port and
starboard spar sets are formed by fibre placement and then split on the long axis. Sacrificial plies
will be used on the external mating surfaces and machined back using the methods. Although this
adds a further manufacturing stage it would reduce joint complexity and weight. The material of
choice is Titanium alloy Ti 6Al 4V. Full joint design is shown in figure 34 (a) through (d) and
proposed installation shown in figures 34 (e) and (f) (notional sizing 6mm thk on initial analysis).
Figures 35(a) and 35(b) show the outboard to mid leading edge spar assembly.
The concept is for a two part assembly the insert section mounted on the IML spar web and flange
faces and the doubler mounted on the spar web OML, the web attachment being made with 30 Hi-
Lok Ti alloy PAN head bolts for a high shear strength joint, with head washers, mounted OML to
IML through pre-drilled holes in both the insert section and the doubler plate, three vertical rows are
used each side of the splice, because the end fasteners will load up first and hence yield early. The
spars currently would be fully drilled from the Master fastener model data prior to assembly, post
machining of their sacrificial ply zones, and loaded with assembly pins for determinant assembly.
Interface sealant would for the whole assembly will be Polysulphide (PRC) as per fuel tank sealing.
The flange to spar and cover skin joint is made using two rows of NAS 1221 Ti alloy Countersunk
bolts, and domed (flange IML) bonded anchor nuts with dielectric seals beneath the nut plate as per
figure 25 for lightening strike protection. The wing cover skins would also be tailored to carry the
balance of the flange shear loads from the splice joint. Currently the flange holes would be pilot
drilled for drill on assembly as per spar flange drilling in tooling, the rib post would be pilot drilled for
drill on assembly.

85
Figure 34(a) (b) (c) (d):- Proposed C section wing spar section splice joint.
A
2 x rows of NAS 1221, 22mm (⅞”) Countersunk Ti Flange bolts.
6 x rows of Hi-Lok, 22mm (⅞”) PAN head Ti Web bolts.
Fig 34 (a) Inboard Front (View on B)
Integral rib post
Fig 34(b) Top (View on A)
B
Fig 34 (d) Doubler (View on C)
C
3d to edge of spar TYP.
2d to edge of part TYP.
3 x vertical rows of Hi-
Lok, 22mm (⅞”) PAN
head Ti Web bolts
each side of splice
(pre-drilled).
3d to edge of spar TYP.
2d to edge of part TYP.
Fig 34 (c) ISO Splice plate.
2.5d to edge of part TYP.

Figure 34 (e) (f):- Proposed C section wing spar section splice joint methodology.
Fig 34(e):- Outboard Leading Edge Splice
plate assembly looking on IML.
Fig 34(f):- Outboard Leading Edge Splice
plate assembly looking on OML.
Splice plate pre drilled installed with integral rib
post (flange pilot holes drilled on assembly).
Leading Edge Spar Mind Section
Joint (sacrificial ply zone).
Leading Edge Spar
Outboard Section Joint
(sacrificial ply zone).
Top cover skin tailored to react
OML flange shear loads.
Bottom cover skin tailored to react
OML flange shear loads.
Leading Edge Spar
Outboard Section Joint
(sacrificial ply zone).
Leading Edge Spar Mind Section
Joint (sacrificial ply zone).
Splice doubler pre drilled installed.
FWD
UP
OUT BD
OUT BD
UP
AFT
86

87
Figure 35(a):- 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).

88
Figure 35(b):- ATDA Outboard Port / Stbd CFC Wing Spar assembly.
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 Ti Rib Post Typical.
OB Leading Edge section to Mid
Leading Edge section Splice joint.
Port Outboard Section
Leading Edge Spar.
UP
FWD
IN BD

ATDA Commercial Transport Airframe Part 2.pdf

ATDA Commercial Transport Airframe Part 2.pdf

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ATDA Commercial Transport Airframe Part 2.pdf