my latest fata airframe design research project current status overview 17th february 2017

Mr. Geoffrey Allen Wardle. MSc. MSc. FATA Wing Design Trade Study 2012-2019

FUTURE ADVANCED TECHNOLOGY AIRCRAFT (FATA) AIRFRAME

DEVELOPMENT STUDY PROGRESS OVERVIEW PRESENTATION.

By Mr. GEOFFREY ALLEN WARDLE. MSc. MSc. C.Eng. Snr MAIAA.

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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 wing 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

2019. 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 proposal, 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 company which is any part thereof.

Sections which are defined as in work Sections 14 through 17 will be presented on completion as

the overview is updated and in depth studies of some supporting sections will be moved to the

capability maintenance supporting presentations, and referenced as such.

This structured overview should be read in conjunction with the following LinkedIn presentations: -

(1) My Composite Design Capability Maintenance Examples: (2) New Metallic Design and FEA

Capability Maintenance Examples: (3) New Kinematics and Aircraft Assembly Robotics Study.

Overview of my current research activities in aircraft design for the FATA paper.


Section 1:- Overview of the FATA airframe design development study.

Section 2:- Benefits of Z- direction reinforcement in composite laminates:

Section 3:- PRSEUS Structural element design and processing:

Section 4:- Overall loading on transport aircraft primary structures:

Section 5:- Structural design philosophies employed in the design of wing components:

Section 6:- Roll and layout of large aircraft wing structural members:

Section 7:- The design and structural layout of FATA wing:

Section 8:- The design and structural layout of the FATA fuselage (in work):

Section 9:- The design and structural layout of the FATA empennage (in work):

Section 10:- Assembly of baseline aircraft wing torsion box structural members:

Section 11:- Robotic assembly in the development of the Baseline wing (see also Robotic Kinematic for

FATA wing Study LinkedIn presentation):

Section 12:- Integration of baseline and developed aircraft main landing gear:

Section 13:- Integration of baseline and future concept engines:

Section 14:- FATA baseline airframe structural analysis and component sizing (in work):

Section 15:- FATA baseline airframe systems integration (in work):

Section 15:- FATA PRSEUS developed airframe structural layout and sizing analysis (in work):

Section 16:- FATA PRSEUS developed airframe systems integration (in work):

Section 17:- FATA MAW control surface integration (in work).

ONLY WORK FROM REFERENCED STUDIES MAY BE REPRODUCED WITHOUT EXPRESS PERMISSION

OF MYSELF AND THE AIAA.

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Contents of this FATA study overview presentation.


Currently I am conducting a conceptual design research into the application the Future Integrated

Structure (FIS) technology PRSEUS (using NASA/TM-2009-215955 (ref 1) and NASA/CR-2011-

216880 (ref 2), as my starting point) and mission adaptive flight control surfaces, to future large

transport aircraft, as detailed in charts 1 to 5, chart 6 shows the projected baseline operational

profile used in loads and fuel tank sizing calculations. This is a technical report for per review

through the AIAA, future PRSEUS studies and the work breakdown are shown in charts 7,8,9.

The reference baseline aircraft selected is for a CFC twin engine 250-300 seat class aircraft design

of conventional configuration. Table 1 presents design data and figure 1(a) illustrates the

configuration of the Baseline FATA aircraft, and figure 1(b) shows the supercritical airfoil selected

the FATA aircraft. This conventional design using the current materials technology shown in figure

2, and will be compared with an improved baseline design incorporating PRSEUS (FIS) technology

figures 5, 6, 7 and 8, and Mission Adaptive Wing MAW Control surfaces, figures 9 and 10, to be

designed using Catia V5.R20, to determine the structural / weight / and aerodynamic benefits at the

trade study level and finally more advanced aircraft design configurations will be used to determine

future potential applications. The study consists of three phases:- (1) The overall airframe

configuration design and parametric analysis using both classical analysis and the Jet306 /

AeroDYNAMIC V2.08 analysis tool set based on my Cranfield MSc: (2) The second is major

structural wing component layout of the airframe initial structure with preliminary systems

integration, and using Cranfield University methods and Catia V5.R20 GSA for structural sizing. (3)

The final design study for both versions of the wing reference and new build will consist of

parametric analysis, initial optimisation and structural layout and analysis and constitutes a

feasibility study proposal to determine the benefits, and constraints on such an application.

Section 1:- Overview of the FATA airframe design development study.

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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

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Table 1:- Baseline Aircraft Data for the AIAA study (highlighted data used for baseline).


Figure 1(a):- Overall configuration and dimensions of the FATA baseline aircraft.

70.52m (231ft 3.3in) Code F

18.34m (60ft 7in)

11.51m (37ft 1.6in)

30.58m (100ft 3.8in)

75.87m (248ft 1.3in) Code E

74.47m (244ft 3.8in)

34.45m (113ft 2.4in)

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)

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Figure 1(b):- Aerofoil profile selection based on C-17 transport.

Figure 2a/b:- Flow fields around 1(a) conventional aerofoil 1(b) supercritical aerofoil.

Figure 2(a) Figure 2(b)

Figure 2(c):- Sketches of root NASA SC(2) 0414 and tip NASA SC(2) 0410 aerofoil profiles.


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AL/Li Alloy

CFRP MONOLITIC

CFRP SANDWICH

TITANIUM

QUARTZ GLASS

By weight percentage.

Composites 50%

Titanium 15%

Steel 10%

Other 5%

AIRBUS A350-900 XWB Airframe.

BOEING 787-8 Airframe.

Figure 2:- Materials utilization on current generation commercial airframes.


My current research activities in aircraft design for the FATA project.

Aircraft design studies are a detailed and iterative procedure involving a variety of theoretical and

empirical equations and complex parametric studies. Although aircraft specifications are built

around the basic requirements of payload, range and performance, the design process also

involves meeting overall criteria in terms of, for example, take-off weights, airport constraints,

maintainability and operating cost.

The main issues come from the interdependency of all of the design variables involved, in

particular the dependency relationship between wing area, engine thrust, and take-off weight which

are complex and often require an initial study of existing aircraft designs to get a first impression of

the practicality of the proposed design, this is the process adopted by myself in designing the

reference wing based upon the most recent fielded technology. An aircraft design trade study can

be considered to two phases:- the initial „first approximation‟ methodology: followed by „parametric

analysis‟ stages, although in practice the process is more iterative than purely sequential. Table 2

shows the basic steps to generate configuration data for AeroDYNAMIC MDO toolset, with some

general rules of thumb, based on concept design experience.

Chart 4 illustrates the development stages, for evaluation the Baseline, PRSEUS airframes and

future concepts employing this technology. The AeroDYNAMIC™ toolset was used to produce

parametric study plots showing the bounds of the design which fitted the chosen design criteria and

are incorporated in the full study paper.

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Requirements Cascade.

Starting with the customer needs, Top Level Aircraft Requirements (TLAR) are formulated

e.g.:- Number of passengers / seats: Weight target: Cargo / baggage payload: Range: etc.

These requirements were broken down into requirements for the Major Airframe Components

of the aircraft Top Level Structural Requirements Studies (TLSRS) e.g. for:-

Fuselage:

Wing:

Empennage:

Systems.

These were further broken down to Section Level Requirements (SLR) for each structural

component e.g. for:-

Skins:

Stringers:

Floor Beams.

All of these are governed by Design Principles and Standards for which for commercial aircraft I

have researched the AIAA ARC :- (Reference Structural Design Principles and Systems

Installation Design Principles). For Airbus there are the RSDP which are Design Principles

for airframe structural design, and SIDP which are Design Principles for designing and

integrating aircraft systems.

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My requirements research breakdown for the FATA aircraft design project.


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FATA Project Top Level Aircraft Requirements for this design project.

Key Requirements Current Best Standard Target

Low Cost Current Unit Cost Gap 20% reduction

High Rate Manufacture Production rate 32 pa 50% increase

Rapid ramp up and Cut Over 30 pa over 2 years 10 fold improvement

Significant Performance Improvements A350 Standard 268.9 tonne 242.01 tonne (10%>)

Aerodynamics A350-1000 Standard 3% drag reduction

Cruise altitude A350-1000 Standard 45,000ft / 13,716m

Range A350-1000 Standard 9,631m / 15,500km

Capacity two class seating A350 -1000 Standard 350 in 9 abreast

Capacity three class seating A350 -1000 Standard 315in 9 abreast

Cargo Capacity A350 -1000 Standard 8pallets +18LD3‟s

Cabin Altitude Pressure A350-1000 Standard 6000ft / 1,829m (20% humidity)

Minimise NRC A350-1000 Standard 50% reduction

DMC improvement 64$/FH 30% reduction


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Table 2:- Example of the „first approximation‟ methodology used in the FATA study.

Estimated parameter. Basic relationship. Rule of thumb.

(1) Estimate wing loading

W/S.

W/S = 0.5 pV² C˪ (in the

„approach‟ condition).

Approach speed lies between 1.45 and 1.62 Vstall.

Approach C˪ lies between C˪max /2.04 and C˪max /2.72

(2) Check C˪ in cruise. C˪ = 0.98(W/S) /q

Where q = 0.5 pV² .

C˪ generally lies between 0.44 and 0.5

(3) Check gust response

at cruise speed.

Gust response parameter

α1wb .AR / (W/S)

α1wb is the wing body lift curve slope obtained from

data sheets.

(4) Estimate size. Must comply with take-off

and climb performance.

The aircraft type considered i.e. long range transport

have engines sized for top of the climb requirements.

(5) Estimate take-off wing

loading and T/W ratio as

a function of C˪V2

s =kM²g²/(SwT. C˪V2 )

1.7< C˪max < 2.2 and 1.18< C˪V2 <1.53

(6) Check the capability

to climb (gust control) at

initial cruise altitude.

17< L/D < 21 in the cruise for most civil airliners.

(7) Estimate take-off

mass.

MTO = ME + MPAL + Mf 0.46< OEM / MTOM <0.57


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Chart 1:- My current research activity for aircraft design trade studies FATA project.

The development and application of

advanced structural concepts, and

mission adaptive control surfaces to

commercial aircraft. Estimated at:-

6,240hrs (15 hour weeks over 8 years)

Work book 1:- Composite airframe design

and manufacture incorporating Catia

V5.R20. (exercises vertical tail fighter a/c

design / commercial aircraft vertical tail

design) COMPLETED.

Work book 2:- FEA using Catia V5.R20.

(exercises airframe structural component

design and analysis) COMPLETED.

Work book 3:- Control surface kinematics

Catia V5.R20. (exercises airframe flap

deployment analysis) COMPLETED.

Major structural layout:- Based on

Cranfield MSc Aircraft Engineering

modules using Catia V5.R20 as tool

set.

Defining airframe study concept:- MSc

Aircraft Engineering modules using

Catia V5.R20 as tool set and

AeroDYNAMIC V3 MSc / BAE skills

sets.

Major structural loads analysis and

component sizing:- Based on Cranfield

MSc Aircraft Engineering modules using

Catia V5.R20 as tool set.


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DETERMINE AIRFRAME CONFIGURATION.

DEVELOP BASELINE STRUCTURAL LAYOUT

Wing size, sub structure layout, control surface

layout, interfaces and LG / fuel tankage integration.

Fuselage diameter, internal structural layout plus

cutouts, and structural interfaces with the wing,

empennage and LG.

Empennage size, structural internal layout, control surface layout and

sizing, interfaces with surfaces and fuselage.

DETERMINE STRUCTURAL LOADING AND LOAD

PATHS

Structural sizing of all major airframe components.

Detailed structural analysis of selected

airframe components.

Chart 2:- Activity dependency for the design trade studies of the FATA airframe.


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Chart 3:- Phases of the FATA airframe PRSEUS design trade study program.

Work book 1:- Composite airframe design

Work book 2:- GSA airframe design

Phase 1:- Baseline composite / metallic wing

box, and wing / fuselage and empennage

layout design structural component sizing.

Baseline composite and metallic wing /

fuselage / empennage design structural /

weight analysis.

Work book 3:- Control surface kinematic

design analysis and sizing.

Phase 2:- Advanced concept composite

PRSEUS wing / fuselage and empennage

layout design structural component sizing.

Phase 1:- Baseline control surface design,

structural sizing and operational analysis.

Advanced FATA concept composite PRSEUS

Airframe design Wing: Fuselage: Empennage

conduct structural / weight analysis.

Phase 3:- FATA concept full composite

PRSEUS Airframe layout, Landing gear,

and MAW control surface integration,

design structural component sizing and

weight analysis.

Phase 2:- MAW control surface design

trades, structural sizing, weight and

operational analysis.


STAGE 1:-DEVELOPMET OF BASELINE AIRFRAME.

Generate concept iterations for parametric analysis using AeroDYNAMIC™ to give sizing of major airframe components against mission requirements, first pass airframe structural loads drop.

Use initial loadings for preliminary sizing of airframe sub-structure, integrating between major airframe component interfaces and installations (power plants, landing gear, fuel tankage) as a Composite / metallic airframe build to Airbus / Boeing design standards meeting FAA / CAA design regulations.

Produce a preliminary airframe design using Catia V5.R20 and Patran / Nastran toolset, to be using current manufacturing technology which forms the baseline for the PRSEUS trade study.

STAGE 2:- EVOLUTION OF BASELINE TO PRSEUS STRUCTURE.

Using the baseline airframe for a twin engined twin aisle long range transport develop a PRSEUS stitched airframe alternative retaining the same sub structure layout and OML, to be produced using RTM and RIM techniques. Analyse the resulting airframe structure and compare with the conventional baseline airframe in terms of weight, complexity, ease of imparting design intent to manufacturing.

Conduct airframe assembly studies, to determine possible automated assembly of major airframe components.

Conduct integration studies of proposed mission adaptive flight control systems for the wing and empennage, factoring these into complexity and performance trades.

STAGE 3:-FUTURE CONCEPTS.

Apply the results and experience gained in stages 1 and 2 to the design and development of advanced configuration airframes to maximise the benefits of PRSEUS stitched composite structural technology, advanced manufacturing and automated assembly technology, and mission adaptive control surfaces.

These airframe concepts are to be in both single aisle medium range, and twin aisle long range transports.

Also to be explored is the application of thermoplastic resin matrix composites and processing technologies.

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Chart 4:- Development Stages of the PRSEUS airframe design for the FATA program.


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Chart 5:- Design Trade Study Project Milestones for the FATA Project.

0% 10% 20% 30% 40% 50% 60% 70% 80% 90% 100%

2011

2012

2013

2014

2015

2016

2017

2018

2019

MILESTONE % COMPLETED.

PR

OJ

EC

T Y

EA

R.

ADVANCED AIRFRAME CONCEPT DESIGN STUDY MILESTONES.

Phase 3

Phase 2

Phase 1

Workbook 3

Workbook 2

Workbook 1


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Chart 6:- Design Trade Study Operational Profile for the FATA paper aircraft sizing.

15,500km (9,631m) 370km

(230m)

45,0

00 f

t 13,7

16m


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Chart 7:- My Future Advanced Technology Baseline Aircraft “Tube and Wing” 2030.

Composite Wings and

Empennage applied PRSEUS

stitched composite

technology.

All electric control system with

MAW technology and advanced

EHA actuation system.

Hybrid Laminar Flow

Control on wing

upper surface.

Composite Fuselage

applied PRSEUS stitched

composite stringers.

Natural Laminar

Flow on nacelles.

Advanced

Engines.

Variable Trailing

Edge Camber.

Wing aspect ratio >10.

Riblets on fuselage.

Hybrid Laminar Flow Control

on Vertical and Horizontal tails .

SOFC/GT Hybrid APU.

Positive control winglets.

HT Thermoplastic

composite engine pylons.

Thermoplastic composite

fuselage frames.


Belly Fairing.


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PRSEUS stitched

composite technology

empennage 2016-2018.

PRSEUS stitched composite

technology wing in work

2013-2017.

Automated Assembly of wing

structure fall 2016-2017.


fuselage frames 2017-2019.

Positive control winglets

2016-2017.

Composite Fuselage applied


stringers 2017-2019.


Belly Fairing 2017-2019. HT Thermoplastic

composite engine pylons

proposed fall 2016-2018.

Chart 8:- My Future Advanced Technology Aircraft Study Project Work Breakdown.

Wing Carry Trough Box Structure

defined and sized ( section 7).

Wing Torsion Box Structure

defined and sized (section 7).


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Chart 9:- My Future Advanced Technology Aircraft Fuselage Study Baseline .

Composite Fuselage applied


stringers 2017-2019.


fuselage frames 2017-2019.

Stringer Co-Bonded to Skin.

Clip PPS Thermoplastic Matrix composite with quasi-isotopic layup

Frame CFRP prepreg.

80mm

120mm

Frame lay up [30º/90º/-30º]

with 0º reinforcement.

The proposed fuselage PRSEUS and thermoplastic application design and structural

development will use either Airbus or Boeing composite fuselage structural design

philosophies as the baseline against which PRSEUS improvements will be assessed.

AIRBUS:- A350 XWB

Boeing:- B787


Conventional 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 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.

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Section 2:- The structural benefits of 3-D stitched and pinned composites.


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The structural benefits of 3-D stitched and pinned composites over conventional laminates.

(a) Lock stitch (b) Modified Lock stich (c) Chain stitch

Needle

Thread

Bobbin

Thread

Needle

Thread

Bobbin

Thread

Figure 3:-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:- 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 3 shows the various types of stitches commonly used to create z-direction reinforcement.


Among the three stitches shown in figure 3, 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. 24



(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. 25



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:- 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 4. The first method (figure 4(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. 26



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 4(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. 27



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Figure 4 (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 4(a). Figure 4(b).

Remove Used

Preform

Uncured Composite

Z-Fiber Preform

Ultrasonic Insertion Transducer

(a) Primary insertion stage and residual preform removal.

(b) Secondary insertion stage.


(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.

29



30

The PRSEUS structural concept illustrated in figures 5-7 is seen as the answer to the HWB

fuselage pressure and bending load issues that have held back the development of this aircraft

type. This study proposes to examine the feasibility of using the same structural concept to reduce

the wing rib structure and composite skin thickness / weight in a large transport aircraft wing.

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 figure 5. Could

a similar concept be used to take the bending, torque, and fuel pressure loads in a conventional

wing? 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 be yes.

The highly integrated nature of PRSEUS is evidenced by figure 6 which shows the structural

assembly of dry warp-knit fabric, precured rods, foam core 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. 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.

Section 3:- PRSEUS Structural element design and processing.


31

Figure 5:- Examples of the PRSEUS airframe technology explored.


The currently PRSEUS for HWB airframe design with its continuous load paths higher notch design

properties, and larger allowable damage levels represents a substantially improved level of

performance beyond that which would be possible using unstitched materials and designs.

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 figure 6, 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 HWB airframe. Boeing and NASA have used this type

of technology in a stitched wing in the 1990‟s figure 6 insert, and in all 8 C-17 landing gear doors.

The dimensions of the NASA test articles and the ply layups are shown in figure 7 and my

developed PRSEUS wing stringers for this FATA wing project are shown in figures 8/9 (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 10/11.

32

PRSEUS Structural element design derived from NASA/CR-2011-216880.


33

Figure 6:- The PRSEUS Structural concept used based on NASA/CR-2011-216880.


34

Figure 7:- PRSEUS Structural element dimensions in mm based NASA/TM-2009-215955.

Rohacell

foam core

(b) NASA Test Frame stiffener

(a) NASA Rod stiffener

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).

Ply orientations:- Pultruded rod 0º :

Each stack : - 7 Plies in +45º / -45º / 0º / 90º / 0º / -45º /+45º pattern

knitted together. Percent by fiber area weight (44/44/12) using

(0º/45º/90º) nomenclature.

The NASA test box layout was 152.4mm stringer pitch and 508mm

frame pitch, analysis conducted using PS SHELL / MAT2 smeared

properties locally sized using HyperSizer as true skin-stringer

geometries this will be used for comparison with Catia V5 baseline

FATA stringer assembly / NASTRAN 2000 modeling.

31.75mm 37.85mm

86.36mm

152.4mm

Test Skin.

101.6mm

12.7mm

152.4mm


35

Figure 8:- Section of the FATA study PRSEUS Upper wing skin Stringer 1.

Pultruded Rod (10mm Dia)

Web Stitching runs

and vectors

Overwrap

C/L

77mm

120mm

Tear Strip

Flange Stitching runs

and vectors Stringer

Ply stack

Lower Wing Cover Skin

Section

PRSEUS Lower wing cover skin stringer 5 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:-

(-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. The same

materials are used stated above in figure 7.


36

Figure 9:- Section of the FATA Study PRSEUS Coaming Stringer.

Pultruded Rod (10mm Dia)

Lower Wing Cover Skin Section

126mm

Web Stitching runs

and vectors

Tear Strip

Flange Stitching runs

and vectors

120mm

Stringer

Ply stack

The PRSEUS Coaming Stringers have an 18 ply stack layup of 0.21336mm

ply giving a thickness of 4.0mm, in the following configuration:-

(-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. All other features and materials as other main stringers see

figure 8.

C/L

Overwrap


37

Figure 10:- RS 545 and RS 543 Lock stitching machines proposed for the FATA stringers.

Figure 10(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 10(b):- Detailed view of the stitching head proposed

for the two rows of stitching on PRSEUS stringer webs.

Figure 10(d):- Detailed view of the stitching head proposed

for the two rows of stitching on PRSEUS stringer flanges. Figure 10(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.


38

Figure 11:- Schematic factory of the future proposal for stitching wing structures.

Stitching

Cutting

Tooling

Assembly

Trim and Drill


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 12(a), and in additional soft

tooling (bagging aides) are also used figure 12(b) and in the Boeing Controlled Atmospheric

Pressure Resin Infusion process figure 12(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 figures 13 and 14

show current PRSEUS structures and NASA‟s road map for development.

PRSEUS component post assembly processing overview.

39


40

Figure 12:- Boeing Controlled Atmospheric Pressure Resin Infusion (CAPRI) process.

Fig 12(b):- Soft tooling (bagging aids) installation over stiffeners.

Fig 12(a):- Robotic stitching of dry preform assembly.

Fig 12(c):- Vacuum bag installation over dry preform assembly.


41

Figure 13:- NASA‟s PRSEUS (CAPRI process) Development Roadmap.


42

Figure 14:- NASA / Boeing Block development of PRSEUS Structures to achieve TRL.


43

1970‟s

Extended Arm Single Needle.

1) Limited preform size.

2) Limited thickness.

3) Low speed manual operation (<100 stitches / minute).

1980‟s

X-Y Computer Controlled Single Head Quilting Technology.

1) 2.4m x 4.6m planform.

2) 38.1mm thick preforms.

3) Medium speed (200 stitches / minute).

1990‟s

Multi-Head Multi-Needle Computer Controlled Gantry.

1) 3.0 x 15.2m planform.

2) Vertical stitching.

3) 38.1mm thick preforms.

4) High speed walking needle concept.

5) 800 stitched / minute.

2000‟s

Robotic Multi-Needle End Effectors.

1) Off axis stitching.

2) Complex shapes.

3) High speed (>800 stitches / minute).

Chart 10:- Evolution of stitching technology for complex airframe applications.


Chart 11:- NASA / Boeing Building Block Methodology for PRSEUS Structures TRL.

44

Based on this Boeing Technology Readiness

Level (TRL) Diagram the PRSEUS structure

manufacturing technology is currently at TRL-

5 for primary structures and TRL-6/7 for

secondary structures see also figure 14.

NOTE:- PROSESSING AND PROCESS VARIABILITY CAN HAVE A SIGNIFICANT IMPACT ON

STRUCTURAL PERFORMANCE.


Overall loading on lifting surfaces:- Figure 15 illustrates the symmetrical flight case forces and

moments to be considered in wing structural design. The structural role of the wing includes the

following (ref 4):-

The transmission of lift the force, which is balanced at the root by the air loads on the fuselage

and the stabilizer and by the inertial loads:

The collection of the chord-wise air loads and the loads from control surfaces and high-lift device

hinges and the transfer of them to the main span-wise beam structure, which has to be achieved

by a series of chord-wise beams and gives rise to a torque on the span-wise structure as well as

contributing to the span-wise bending of the wing:

The transfer to the main beam of the local inertia loads from the wing mounted powerplants, and

retracted main landing gear units:

The reaction of landing loads from the main landing gear units:

The pressure and inertia loads from integral fuel tanks and fuel:

The provision of adequate torsional stiffness of the wing in order to satisfy the aeroelastic

requirements:

The reaction of wing and landing gear drag loads and possibly, thrust loads in the plane of the

wing.

Figures 16(a) through (c) illustrate Symmetric:- span-wise, chord-wise, and fuselage loading.

Figures 17(a) through (d) illustrate Asymmetric (roll):- span-wise, fuselage torque, and fuselage

sideslip and yaw loading, and figure 14(a) and (b) illustrates overall ground loading. Figure 15

illustrates overall fuselage loading 45

Section 4:- Overall loading on the aircraft primary structures.


46

Figure 15:- Overall loading on the aircraft wing surfaces.

Lw

Dw Lc

T

R

S

D

Wing inertias (structural / fuel) – relieve

all vertical and in-plane effects. Main landing gear.

R= Vertical – wing vertical shear, moment, torque.

D= Drag – wing in-plane shear, moment, torque.

S= Side – wing vertical moment.

Lw= Wing lift – wing vertical shear, moment, torque.

Lc= Control /high-lift devices – wing vertical shear, moment, torque.

Dw= Wing drag – wing in –plane shear, moment.

T = Thrust – wing in – plane shear, moment, torque.


Symmetric flight cases:- Figure 16(a) illustrates the loading and corresponding form of the shear

force diagram across the wing of a twin engined low wing commercial airliner configuration similar

to the baseline study aircraft. Symmetric wing lift is relieved by the inertia of the structure, engines,

systems and fuel (see section 6). The overall loading on the wing is reacted at the side of the

fuselage at the wing root joint, and the bending moment is constant across the fuselage.

The loads on a typical chord-wise wing section are illustrated in figure 16(b), the sum of the

moments of the forces about a given chord-wise reference point yields the torque at that section,

and the integration of the local values of the torque across the span of the wing yields the overall

torque diagram.

Finally figure 16(c) illustrates the loading and the basic form of the shear force diagram along the

length of the fuselage of a twin engined low wing commercial airliner similar to the baseline study

aircraft. The shear force and bending moment due to the horizontal air-load are relived along the

fuselage by the transitional and rotational inertia effects. The net fuselage bending moment at the

fore and aft centre of gravity (c.g.) position is balanced by the sum of the wing torques at the sides

of the fuselage.

Asymmetric flight case:- The asymmetric flight cases are more complex than the symmetric

cases. A simplified example is the instantaneous application of aileron control on a wing having no

initial lift results in an asymmetric loading case, although in practice there is no true symmetry

between the up-rising and down-lowering ailerons. A more usual case is when the ailerons are

applied as the aircraft is in steady level flight as shown in figure 13(a).

47

Overall loading on the aircraft primary structures (continued).


Figure 16(a):- Symmetric span – wise loading steady level flight condition.

48

Horizontal stabilizer load.

Span-wise airload. Net distributed span-wise load.

Fuselage reactions.

Powerplant inertia. Powerplant inertia.

Span-wise inertia load. Span-wise inertia load.

SHEAR FORCE DIAGRAM.


49

Powerplant weight.

Thrust -T

Aerodynamic moment - M

Control / Flap moment.

Aerodynamic Lift - L

Aerodynamic Drag - D

Control Force.

Control surface drag.

Wing structural systems

and fuel weight.

Figure 16(b):- Symmetric loading chord – wise torques on the aircraft wing.


50

Figure 16(c):- Symmetric flight case fuselage loading.

Thrust.

Drag.

Horizontal stabilizer airload.

Aerodynamic moment from wing.

Wing lift. Fuselage lift.

Centre of gravity.

Fuselage reaction.

Aircraft inertia.

Fuselage reaction

Stabilizer load

SHEAR FORCE DIAGRAM.


Asymmetric flight case (continued):- The initial steady level flight condition will have a symmetric

loading as shown in figure 16(a). The aileron and the consequent roll effects are approximately anti-

symmetric in form figure 17(a). Figure 17(b) shows the shear force distribution due to this anti-

symmetric condition as well as the overall result of combining it with the symmetric diagram. In a

general rolling motion the couple resulting from the application of the aileron is balanced both by

the acceleration effect of the roll inertia and the aerodynamic effect due to the rate of roll (ref:-4).

The torque loading on the rear fuselage as a consequence of the application of the rudder control to

cause a sideslip motion is shown in figure 17(c). The torque due to the fin side load is increased by

the effect of asymmetric distribution of the trimming load on the horizontal stabilizer.

Figure 17(d) shows the plan view of the fuselage, illustrating how the fin side load is reacted by side

forces along the fuselage. The lateral bending along the fuselage is relived by sideslip and yaw

inertial effects and the net value at the wing root is balanced by wing aerodynamic forces and yaw

inertia. The torque on the fuselage is mainly reacted by the rolling inertia of the wing group.

Fuselage cyclic pressure loading is also very important and is considered in fuselage loadings later

in this study.

Ground loading cases:- The ground loading cases unlike the flight cases occur from local ground

forces. The take - off case is effectively a static balance of the aircraft weight by the vertical loads

on the nose – and main – wheels. However, the landing cases are not static in that even after the

wheels have made contact with the ground there is a translational motion of the centre of gravity of

the aircraft, as well as a rotation in pitch and, possibly, roll. It is also usual for the wing to be

providing lift at the time of wheel contact with the runway. Figures 18(a) and (b) illustrate the nature

of the landing gear span-wise loading, and the longitudinal loading.


51


52

Figure 17(a):- Asymmetric (roll) span – wise loading flight condition.

Force due to aileron

application.

Net wing load in steady level flight.

Load due to rate of rotation in roll (roll damping).

Fuselage reactions – balance net

vertical force and rolling moment.

Resultant force and moment at fuselage Net moment is the difference of aileron, roll rate, and inertia effects.

Force due to aileron

application.


53

Figure 17(b):- Asymmetric (roll) span – wise loading flight condition shear force diagrams.

Aircraft C/L

Powerplant inertia. Anti-symmetric load.

Aircraft C/L

Fuselage reaction.

Overall.


54

Reacting fuselage side

load (balanced by inertia

and wing-body air-load.

Fin side load.

Asymmetrical trim load on horizontal tail.

Reacting fuselage torque (balanced

mainly by wing rolling inertia.

Aircraft C/L

Figure 17(c):- Asymmetric loading flight condition fuselage torque.


55

Figure 17(d):- Loading on the fuselage (sideslip, yaw and pressure).

Resultant side force – balanced by lateral (horizontal) inertia.

Fuselage side air-load (distributed along fuselage length.

Fin side load.

Moment at centre of gravity due to side loads –

balanced by yawing (rotational) inertia.

Cabin Pressurisation creates cyclic hoop tensile

stresses in the fuselage skin.

P

A

A

Section view on A


Ground loading cases (continued):- The various forces and moments are balanced in the same

way as those arising in the flight cases, that is primarily by inertial effects. For this reason here the

ground contact forces are regarded as applied loads rather than as reacting forces.

Overall loading on the fuselage:- The loading determining the design of the fuselage is shown in

figure 19. The roles of the fuselage includes the following:-

Provision of a pressurized (in commercial aircraft) envelope and structural support for the

payload (passengers and freight) and crew. The skin thickness required to limit hoop tensile

stresses to acceptable values is given by:- tp = ∆ρR / σρ Where:- ∆ρ is the maximum working differential

pressure: R is the local radius of the shell : and σρ is the allowable tensile working stress.

To react landing gear, pressurization (in commercial aircraft), and powerplant loads when these

items are located on, or within the fuselage, the nose gear being always present.

To transmit the control and trimming loads from the stabilizing / control surfaces to the centre of

the aircraft, and to provide support and volume for equipment and systems.

These requirements imply that to perform its structural role the fuselage has to be a longitudinal

beam loaded both vertically and laterally, it also has to react torsion and local concentrated loads,

the provision of a pressurized envelope implies a cylindrical encapsulated construction, with

pressure bulkheads. Therefore a conventional commercial airliner fuselage of circular cross section,

cabin floor, and cargo bay floor, with pressurized cabin, and external powerplants is the baseline

FATA airframe.

56



57

Figure 18(a):- Ground loading span – wise.

S

Ground vertical loads = R R R

Ground side loads = S

Resultant force and moment at fuselage.

Net wing load.

Fuselage reaction to balance vertical and side loads and rolling moment

due to side load – balanced by roll, vertical and horizontal inertias.


58

Ground vertical loads = R R

D Ground drag loads = D

Fuselage vertical force – reacted

by vertical (translational) inertia.

Fuselage bending moment – reacted

by pitch (rotational) inertia.

Overall lift and weight in balance.

Figure 18(b):- Ground loading longitudinal.


59

Figure 19:- Overall loading on the fuselage.

LF

LT

D

R

S

D

R

S

Main landing gear.

Nose landing gear.

LF = Fin load – fuselage horizontal shear, moment, torque:

LT = Tail load – fuselage vertical shear, moment, torque.

R = Vertical - fuselage vertical shear moment:

D = Drag – fuselage vertical shear moment:

S = Side – fuselage horizontal shear, moment, torque.

Cabin Pressurisation creates cyclic hoop tensile stresses in the fuselage skin.

P


Aircraft structures fall into 3 categories which are as follows:-

Class 1:- structural component the failure of which will result in structural collapse; loss of control;

failure of motive power, injury or fatality (to any occupant); loss of safe operation of the aircraft.

Class 2:- Stresses components but not Class1.

Class 3:- Unstressed or lightly stresses component which is neither Class 1 or 2.

Structural integrity is defined as the capability of the structure to exceed applied design loading

throughout its operational life, and the selection of a design philosophy to achieve this from the start

of the design process is extremely important as this selection impacts on:- airframe weight;

maintainability; service life; and any future role change of the airframe. The approaches available to

the designer are:- Static Strength; Safe Life; Failsafe; Damage Tolerance; and Fatigue Life, the last

four of which, are expanded below (ref:-4). See tables 3 through 5 for FATA candidate materials

selection.

(a) Safe Life:- The important criterion in this approach is the time before a „crack or flaw‟ is initiated

and the subsequent time before it grows to critical length. It can be seen from a typical S-N

curve that low levels of stress at high frequency of application theoretically do not cause any

fatigue damage. However it is necessary to allow for them, possibly by introducing a stress

factor such that effectively damage dose not occur.

(b) Fail-safe:- In this approach the dominate factors are the crack growth rate and the provision of

structural redundancy in conjunction with appropriate structural inspection provision.

60

Section 5:- Structural design philosophy of airframe structural components.


There are several ways of ensuring that fail safety is achieved:-

i. By introducing secondary, stand-by components which only function is in the event of a

failure of the primary load path, to carry the load. This may consist of a tongue or a stop

which is normally just clear of the mating component. A mass penalty may be implied but in

same circumstances it is possible to use the secondary items in another role, for example

the need for a double pane assembly on cabin windows for thermal insulation purposes.

ii. By dividing a given load path into a number of separate members so that in the event of the

failure of one of them the rest can react the applied load. An example of this is the use of

several span wise planks in the tension surface of metallic wing boxes. When the load path

is designed to take advantage of the material strength the use of three separate items

enables any two remaining after one has failed to carry the full limit load under ultimate

stress. In some instances the „get home‟ consideration may enable a less severe approach

to be adopted.

iii. By design for slow crack growth such that in the event of crack initiation there is no danger

of a catastrophic failure before it is detected and repaired.

c) Damage tolerant:- With this philosophy it becomes necessary to distinguish between

components that can be inspected and those that cannot. Effectively either the fail-safe or

safe-life approaches are then applied, respectively, in conjunction with design for slow crack

growth and crack stopping (e.g. panel braking web stiffeners).

61

Structural design philosophy of airframe structural components.


A. Safe-life and Fail–safe design processes (see Chart 11):- There is a commonality in the design

process for the safe –life and fail-safe concepts. The material to be used for the structure must

be selected with consideration of the critical requirements for crack initiation or crack growth

rate, as most relevant, together with the operating environment. A vital consideration for fail-

safe design is the provision of the alternative load paths, possibly together with crack

containment or crack arresting features. When these decisions have been made it is possible to

complete the design of the individual components of the structure and to define the

environmental protection necessary.

In the case of the safe-life concept the life inclusive of appropriate life factor follows directly

from the time taken initiation of the first crack to failure. Inspection is needed to monitor crack

growth. In the fail-safe concept the life is determined by the structure possessing adequate

residual strength subsequent to the development and growth of cracks.

In both cases it essential to demonstrate by testing, where possible on a complete specimen of

the airframe, that the design assumptions and calculations are justified. Further, in fail-safe

design it is necessary to inspect the structure at regular, appropriate intervals to ensure that any

developing cracks do not reach the critical length and are repaired before they do so.

As the design process is critically dependent upon assumed fatigue loading it is desirable, if not

essential, to carry out load monitoring throughout the operational life of the airframe. This is

used either to confirm the predicted life, or where necessary, to modify the allowable

operational life.

62

Structural design philosophy application processes.


63

Safe-Life.

Crack Initiation time.

Fail-Safe.

Crack growth rate.

Provision of redundancies.

Crack containment.

Environment.

Material: Component Design:

Corrosion protection: Testing.

Life. Residual strength.

In service load monitoring.

Chart 11:- Application of Safe-life and Fail-safe structural design philosophies.


B. Damage Tolerant Design process (see Chart 12):- The damage tolerant approach commences

with the assumption that cracks or faults are present in the airframe as manufactured.

Experience suggests that these vary in length from 0.1mm to as much as 1.5mm.

Those items of the structure which may be readily inspected can be designed by selecting an

appropriate material and then applying essentially a fail-safe approach. The working stress

level must be selected and used in conjunction with crack stopping features to ensure that any

developing cracks grow slowly. Inspection periods must be established to give several

opportunities for a crack to be discovered before it attains a critical length.

When it is not possible to inspect a particular component it is essential to design for slow–crack

growth and ensure that the time for the initial length to reach its critical failure value is greater

than the required life of the whole structure. Since this approach is less satisfactory than that

applied to parts that can be inspected it is desirable to develop the design of the airframe such

that inspection is possible, wherever this can be arranged. As with safe-life and fail-safe

philosophies testing is needed to give confidence in the design calculations. Likewise, in-service

load monitoring is highly desirable for the same reason. This design philosophy is employed on

this project using techniques from ref:-4, JAR 25, and data sheets, MSc F&DT module notes.

C. Fatigue-life Design process (see Chart 13):- The first stage in the fatigue-life approach is the

definition of the relevant fatigue loads and the determination of the response of the aircraft

structure to these loads. The analysis for this follows that for limit load conditions, which

enables the loading on individual components of the airframe to be determined, and the

airframe structural response to be assed and the best design philosophy to be applied. 64

Structural design philosophy application processes.


Chart 12:- Application of the Damage Tolerance structural design philosophy.

Damage Tolerant.

Crack in structure as manufactured.

Is the component inspectable?

Yes. No.

Fail-safe approach.

Slow crack growth.

Crack arrest features.

Inspection periods.

Crack growth to initiate

failure to be more than

service life.

Testing.

In service load monitoring (FTI / G monitors / SHM). 65


66

Chart 13:- Application of the Fatigue-life structural design philosophy.

Fatigue-life.

Aircraft structural response.

Fatigue load spectra.

Design philosophy selection.

Damage Tolerant. Safe-Life. Fail-Safe.


Fatigue Design Requirements:- The emphasis of the requirements specified to ensure the integrity

of the airframe design under fatigue loading is on the methods of analysis and the means of

determination of a satisfactory fatigue life. Only in the United States military code is there a

specification of a magnitude and frequency of repeated loading and this is outlined below. Loading

conditions for all categories of aircraft are discussed below.

1) Civil transport aircraft JAR 25.571:- This standard outlines the basic requirements for fatigue

evaluation and damage tolerance design of transport aircraft. The paragraph outlines the

general requirements for the analysis and the extent of the calculations. Amplification of the

details is given in the associated „acceptable methods of compliance‟ given in JAR 25.ACJ

25.571.

2) UK Military Aircraft:- The basic requirements for fatigue analysis and life evaluation are

specified in Def Stan 00970 Chapter 201. This covers techniques for allowing for variances in

the data as well as overall requirements and the philosophy to be adopted. Detail requirements

of the frequency and magnitude of the repeated loading are given in the particular specification

for the aircraft.

3) US Military Aircraft:- The United States military aircraft stipulations are to be found in three

separate documents:- In MIL-A-8866A the emphasis is on the detail of the required magnitude

and frequency of the repeated loading rather than on analysis the data covers;- maneuver;

gust; ground and pressurization conditions for fighter, attack, trainer, bomber, patrol, utility and

transport aircraft. 67

Structural design fatigue requirements for design philosophy application.


MIL-A-8867 prescribes the ground testing to be undertaken as part of the demonstration of the

life of the airframe. The final document the is MIL-8868 paragraph 3.4 and 3.5 which stipulate

the information to be provided in the form of reports outlining the analysis and testing

undertaken to substantiate the life of the airframe.

The types of repeated airframe load data required for design against fatigue and to apply in the

selected component design philosophy are outlined below.

1) Symmetric manoeuvre case:- Extensive information is available in relation to symmetric

manoeuvres of both military and civil aircraft, e.g. Van Dijk and Jonge‟s work which outlines a

fatigue spectrum obtained from flying experience of fighter / attack aircraft which is known as

the FALSTAFF spectrum, based on the maximum value of peak stress (s) and loading

frequency (n) the peak stress selected being the Input Parameter.

2) Asymmetric manoeuvre case:- Fatigue loading data for asymmetric manoeuvre loading is

sparse, and these originate from the roll and yaw controls, the texts of Taylor derives data from

early jet fighter experience. As for civil aircraft it has been determined that atmospheric

turbulence is of much greater significance.

3) Atmospheric turbulence:- Fatigue loading due to encounters with discrete gusting or the effect

of continuous turbulence is of importance for all classes of aircraft, but especially for those

where operational role does not demand substantial manoeuvring in flight. ESDU data sheets

69023 (Average gust frequency for Subsonic Transport Aircraft (ESDU International plc. May

1989) is used in this study.

68



There are two main types of turbulence which are:- (a)Symmetric Vertical Turbulence, and

(b)Lateral Turbulence.

a) Symmetric Vertical Turbulence:- where gust magnitude is a function of both flight altitude and

terrain over which the aircraft is flying, e.g. low level penetration bombing missions B-1B,

Tornado, and B-52H, where there are more up gusts than down, these are allowed for by

using correction factors.

b) Lateral Turbulence:- there is less information on the frequency and magnitude of lateral

turbulence for aircraft but it has been suggested that at altitudes below about 3km the

frequency of a given magnitude is some 10-15% greater than those of the corresponding

vertical condition.

4) Landing gear loads:- these fall into three categories;- (a) loads due to ground manoeuvring e.g.

taxiing; (b) the effects of the unevenness of the ground surface e.g. unpaved runways, rough

field poor condition runways, major consideration in troop / cargo military transports, and

forward based CAS aircraft; (c) landing impact conditions. The texts of Howe (ref:-4): Niu: and

MIL-A-8866A are employed in this project.

5) Buffeting Turbulence:- Flow over the aircraft may break down at local points and give rise to

buffeting. This induces a relatively high – frequency variation in the aerodynamic loads,

possibly resulting in the fatigue of local airframe components such as metallic skin panels.

6) Acoustic Noise Turbulence:- local high frequency vibration or flow field loading, and ESDU Data

sheets 75021 and 89041 were used in this project.

69



Materials Code ρ

Kg/m

E

GPa

σe

MPa kht khc kdt kdc kθ

Carbon /

Epoxy. 3501/6 QI 1605.44 67 736 0.61 0.65 0.55 0.38 0.83

Carbon /

Epoxy. 3501/6 O 1605.44 80 880 0.55 0.62 0.55 0.38 0.83

Ti Alloy Ti6Al4V 4428.8 110 902 0.94 0.94 0.20 0.94 1.00

Al/Li Alloy 8090 T3X 2530 80 329 0.94 0.94 0.39 0.94 0.90

Al Alloy 7075 T76 2768 72 483 0.94 0.94 0.29 0.94 0.90

Al Alloy 2024 T351 2768 72 325 0.94 0.94 0.31 0.94 0.90

70

3

Table 3(a):- Materials Properties of FATA wing / empennage materials (Ref.6).


Table 3(b):- Materials Properties for FATA fuselage materials (Ref.22).

Property. Unidirectional Tape /Slit Tape. Plain Weave Fabric.

Thickness per ply 0.15mm 0.25mm

Density (ρ) 1790kg/m 1570kg/m

Longitudinal modulus E1 (GPa) 137.3 62.6

Transverse modulus E2 (GPa) 7.8 59.3

In-plain Shear modulus G12 (GPa) 5.23 4.6

Poisson's ratio V12 0.36 0.062

Longitudinal tensile strength F1t (MPa) 2057 621

Transverse tensile strength F2t (MPa) 46.9 594

Longitudinal compressive strength F1c (MPa) 1610 760

Transference compressive strength F2c (MPa) 207 707

In-plain shear strength F6 (MPa) 135 125

Standard Width ATL* 460mm / AFP* 12.5mm 1600 mm

71 *ATL= Automated Tape Laying : AFP = Automated Fibre Placement.

3 3


Category. Failure Mode. Weight Ratio (W2 / W1)

1 Tensile strength. ρ2 / ρ1 σe1/σe2 [kth1/ kth2 kθ1/kθ2]

2 Compressive strength. ρ2 /ρ1 σe1/σe2 [kch1/kch2 kθ1/kθ2]

3 Crippling ρ2 / ρ1 [Es1 σe1 / Es2 σe2]

4 Compression surface column and crippling ρ2/ρ1 [Es1 Et1 σe1/Es2 Et2 σe2]

5 Buckling compression and shear ρ2 /ρ1 [E1 / E2]

6 Aeroelastic stiffness ρ2/ρ1 E1/E2

7 Durability and damage tolerance ρ2/ρ1 [kd1kθ1/kd2kθ2]

72

Table 4:- Weight Ratio Equations for Various Failure Categories (based on Ref.6).

0.25

0.2

1/3


Material Code

Weight Ratio (S1/S2) (ρ2/ρ1)

Cat 1 Cat 2 Cat 3 Cat 5 Cat 6 Cat 7(a) Cat 7(b)

Carbon /

epoxy 3501/6QI 0.4 0.4 0.5 0.4 0.6 0.2 0.7

Carbon /

epoxy 3501/6O 0.4 0.3 0.4 0.4 0.5 0.1 0.6

Titanium Ti6Al4V 0.5 0.5 1.1 1.0 1.0 0.8 0.5

Aluminium /

Lithium 8090T3X 0.9 0.9 0.9 0.9 0.8 0.7 0.9

Aluminium

alloy 7075 T76 0.7 0.7 0.9 0.9 1.0 0.7 0.7

Aluminium

alloy 2024 T3 1.0 1.0 1.0 1.0 1.0 1.0 1.0

73

Table 5:- Weight Ratios for Airframe Materials for Various Failure Categories (Ref.6).

n


74

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 the previous slides. Figure 20

illustrates the control surfaces on the wing of the FATA subsonic composite transport aircraft, 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 6:- Roll and layout of large aircraft wing structural members.


75

Figure 20:- Control surface layout of the FATA composite transport aircraft.

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)


76

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 21 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 22(a) and (b), 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 FATA 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 23, 24, and 25.

Roll and layout of large aircraft wing structural members ( CFC cover skins).


Figure 21(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.

77


Figure 21(b):- Fibre Orientation Requirements for CFC Wing Skins / covers.

78

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 22(a):- FATA Structural Ply Thickness Zones Upper Wing Cover Skin R.6.1

79

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 FATA 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 22(b):- FATA Structural Ply Thickness Zones Lower Wing Cover Skin R.6.1

80

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

(For FATA 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

Outer OML Skin Ply.

10 mm

10.0

3.0

2.0

2.5

2.5


81

Fig 23:- FATA Transport aircraft upper cover skin stringer layout to inhibited skin buckling.

Fig 23(b) Upper Cover Skin Stringer Close up of area „A‟.

Fig 23(a) FATA 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 FATA 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 RRSEUS 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.


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.

82

Roll and layout of large aircraft wing structural members (CFC cover skins).


83

Figure 24:- 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 24

(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. Experimentation has shown the

how effective these have been and a brief summary is as follows:-

Resin / adhesive noodles – Poor

Tow noodles – Fair

Braided noodle – Good

Braided “T” preform - Good to Excellent.

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

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

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

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

surrounding material.

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

constant/minimum glue line thickness of 0.005” per ply, 2 plies max typically. Figure 25 and 26

show how peel stresses and manufacturing weight can be reduced in stringer design. Figures 27(a)

and (b) shows the FATA lower cover skin stringer arrangement and special considerations for the

inspection cut outs, either side of which coaming stringers are installed.


84


85

Figure 25:- Composite Stringer design based on design / test 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 Prime

baseline FATA).

5 N/mm²

4 N/mm²

6º Chamfer flange toe strip

(selected for Developed

PRSEUS FATA). 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 26(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

86


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 was used on Eurofighter

Typhoon wing spars which were co-bonded to the lower wing cover skins, and proposed for the

F-35B VT lower skin stringers in SWAT trade studies, and is used to bond the wing cover skin

stringers for large transport aircraft 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.

Design options for stringer adhesive bonded joints detailed in WB1.

87


Fig 27(a):- FATA 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 panels under tension loading. (Tear –

strip peel plies omitted for clarity).

1:20 Skin Zone

Transition Ramps

in the direction of

principle stress.

88


89

Fig 27(b):- FATA 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.


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 28(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.

Reference wing box layout key structural members (CFC cover skins).

90


Zone 3 Indirect effects.

Zone 2 Swept stroke.

Zone 1 Direct strike.

Lightening Strike

Zones on an

aircraft with wing

mounted engines.

Figure 28(a):- Lightening strike risks to composite wing structures with podded engines.

91


92

Figure 28(b):- Lightening strike risks to composite podded engine aircraft structures.

Zone 1 Direct strike. Zone 1 Direct strike.











Zone Key.



93

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.

Methods of lightening strike protection for composite aircraft wing structures have been developed

and are illustrated in figure 29(a)/(b), these range from layers of aluminium foil on EAP wing, to the

sophisticated copper mesh and fastener insulations used on Eurofighter Typhoon, and the Boeing

787 transport, and the latter will be employed in this study (see also ref 5). 94

Roll and layout of large aircraft wing structural members (cover skins).


95

Figure 29:- Lightening strike protection of composite wing cover structures (ref 5).

Copper grid recessed into skin.

Fig 29(a) Aluminum foil EAP.

Fig 29(b) Copper strip Eurofighter Typhoon. Fig 29(c) Copper mesh grid Boeing 787.

COPPER STRIP RECESSED INTO SKIN.

TUFTHANE INSULATED RIVETS.

INDIVIDUAL STRIP.

SKIN.

SPAR.

(See My Composite Design Capability Maintenance

Studies LinkedIn presentation for fuselage lightening

strike protection methods).


96

Figure 29:- Lightening strike protection of composite fuselages ( A350 and B-787) cont.

Electrical network following frames and floorgrid.

Grounding

Bonding

Voltage

HIRF Protection

CFRP

Lightening Direct Protection:

CFRP + Metallic Mesh.

Figure 29(e) Airbus A350 system.

The Boeing 787 employs

Inter-Woven Wire Fabric

(IWWF) Lightening strike

protection.

Figure 29(d) Boeing B787 system.

* FATA uses an Electro Mesh™ IWWF lightening strike protection.


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 30(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 31. 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.

97



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 14 (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.

98

Classification of impact damage by severity for composite aircraft structures.


99

Chart 14:- 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. 100

Classification of impact damage by severity for composite aircraft structures.


101

Figure 30(a):- Structural damage risks to composite wing structures.


102

Figure 30(b):- Structural damage risks to composite fuselage structures.

Cut-out skin reinforcement:-

20:1 ply dropoff ramps

Heavy skin around door special

criteria to resist ramp rash.


103

Figure 30(c):- Structural damage risks to composite airframe structures.

CFC Fuselage.

CFC Empennage.


Figure 31:- Woven Cloth Classifications and surface ply BVID protection options trades.

104


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 32, 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 33, 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 34, where there are

bolted joints, tooling hole sites, to react pressure differentials at fuel tank boundaries, and spar

section splicing, figures 35 to 37 show preliminary outboard wing spar design, and figure 38 shows

a spar splice joint concept and 39 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 20. 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. 105

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. 106

Roll and layout of large aircraft wing structural members (CFC wing spars).


107

Figure 32:- Airbus A350 Composite spar manufacture and assembly.

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.


108

Figure 33:- 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


109

Figure 34:- 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 35:- FATA 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.

110


Figure 36:- 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. 111


Figure 37:- FATA 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 112

Based on Carbon / Epoxy 3501/6 QI unidirectional composite tape

material with a ply thickness of 0.21336mm (see table 6(a),6(b),and 6(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

113

Table 6(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

114

Table 6(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

115

Table 6(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

116

Table 6(c):- Outboard Leading Edge Spar Zone (3) 46 ply stacking sequence (part 2).


117

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 shown in

figures 64 and 65. Although this adds a further manufacturing stage it would reduce joint complexity

and weight. The material will be choice will be Titanium alloy Ti 6Al 4V. Full joint design is shown in

figure 38 (a) through (d) and proposed installation shown in figures 38 (e) and (f) (notional sizing

6mm thk on initial analysis). Figures 39(a) and 39(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 pilot drilled with final holes drilled on assembly, post machining of their

sacrificial ply zones. Interface sealant would for the whole assembly will be Polysulphide (PRC) as

per figure 70 for 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 29(c) 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.


118

Figure 38(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 38 (a) Inboard Front (View on B)

Integral rib post

Fig 38(b) Top (View on A)

B

Fig 38 (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 38 (c) ISO Splice plate.

2.5d to edge of part TYP.


Figure 38 (e) (f):- Proposed C section wing spar section splice joint methodology.

Fig 38(e):- Outboard Leading Edge Splice

plate assembly looking on IML.

Fig 38(f):- Outboard Leading Edge Splice

plate assembly looking on OML.

Splice plate pre drilled installed with integral rib

post (flange 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

119


120

Figure 39(a):- FATA Outboard Port / Stbd CFC Wing Spar assembly.

Port Mid Section

Leading Edge Spar.

Port Outboard Section

Leading Edge Spar.

Ti alloy Rib Post 29



Ti alloy Rib Post 32 Ti alloy Rib Post 33 Ti alloy Rib Post 34

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).


121

Figure 39(b):- FATA 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


WING RIBS:- The ribs, an example is shown in figure 40, 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 figure 40(d), 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.

122

Roll and layout of large aircraft wing structural members (wing ribs).


Rib construction for large transports fall into three types and this off course influences the way in

which they distribute the external loads and reaction forces categorized above, the three types of

metallic construction are:- (a) Truss type: (b) Shear web type: (c) Webs stiffened ribs with fuel

transfer holes (shown in figures 40(a) and 40(c) is the FATA baseline Al/Li Rib 12).

The way in which the rib structure resists the external loads and reaction forces the rib is subjected

to is dependent on the construction methods employed as outlined below:-

In the truss rib construction distributed external loads and reaction forces are applied as

concentrated loads at the joints and the structure can be analysed as a simple truss. The outer

members on which the distributed loads are relied upon to transfer these loads, in shear, to the

points where they can then be considered as concentrated loads. These outer members are

therefore subjected to combined bending and compression or bending and tension, structural

analysis of one such rib is given in Workbook 2.

Shear web rib construction is usually employed in to either distribute the concentrated loads,

such as the engine pylon or main landing gear, or to distribute fuel tank bulkhead boundary

pressure loads to the shear beams.

Web with lightening hole and stiffener construction are used to resist bending moments by the

rib cap members and shear by the web.

Simple beam structural analysis can be applied to ribs design checking the following:- Shear in the

web, or axial loads in the truss members: Rib cap bending loads: Shear attachment to the spars

and wing cover skins: Tension attachment of the wing cover skins: Crushing loads: Shear load

effects from local cut outs: Fuel pressure loads which are normal to the rib plane. 123

Roll and layout of large aircraft wing structural members (Al Li wing ribs).


Figure 40(a) / (b):- Inner Metallic rib 12 design for FATA aircraft baseline study in Al/Li.

124

Fig 36(a):- Advanced metallic aircraft rib 12, for the FATA baseline study

using the methodology employed in the B787 and for composite wing

skins with CFRP „I‟ stringers using the contour of the rib flange for

attaching both skin and bonded stringer to the rib (stressing for FATA

baseline ribs sizing is in work this model uses nominal sizing).

FWD

UP

IN BD

Fig 40(b):- Boeing 787 metallic rib with „I‟ stiffeners.

Leading edge spar bath tub attachment.

Ventilation holes.

Fwd Mass Flow Fuel Transfer

Hole with web reinforcement.

Aft Mass Flow Fuel Transfer Hole

with web reinforcement.

Low Level Fuel Transfer Hole


Low Level Fuel Transfer Holes

with web reinforcement typical.

Trailing edge spar bath tub attachment .

Shear load web stiffeners

typical.

Fuel Transfer System

Penetration Holes with

web reinforcement.

Web panel breakers

typical.

Low Level Fuel Transfer Hole


Initial design weight in Al/Li = 78.581kg.


125

Figure 40(c) / (d):- Inner Metallic rib design for FATA baseline study in Al/Li.

Fig 40(d):- Wing crushing loads due to flexure bending.

Leading edge spar. Wing top cover skin.

Wing bottom cover skin.

Wing bottom skin stringers.

Fwd coaming skin stringer.

Low level fuel

Transfer Hole.

Aft fuel drain.

Fwd ventilation. Aft ventilation.

Fig 40(c):- Metallic Al/Li Rib 12 installed in FATA wing view looking outboard.

(N.B. Low level fuel transfer holes and ventilation holes double as tooling holes

sizing and location of all ventilation and fuel transfer holes is based on TSM-08).

Trailing edge spar. Wing top skin stringers.

Fwd fuel drain.

Low level fuel

Transfer Holes.


Both the FATA Prime baseline, and the Developed PRSEUS FATA wing, employ carbon fibre

composite ribs at 11 locations, figures 41and 42 illustrate the basic rules for interface strategy.

In the case of the FATA Prime baseline wing CFC ribs shown in figures 43, and 44 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 (see figure 25) 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 45 and 46 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 FATA wing CFC ribs shown in figures 47 to 51, 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 2). 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, and is aligned with LOCOMACH research

(reference 19).

126

Roll and layout of large aircraft wing structural members (CFC wing ribs).


127

Figure 41:- Basic rules of interface design strategy FATA composite structure.

1) Design critical joining areas to

permit elastic gap closure during

assembly.

2) Avoid over dimension in the flange

design.

3) Ensure that laminate thickness

variations are not critical for the

tolerance chain.

(1)

(2)

(3)


128

Figure 42:- Geometric accuracy strategy FATA composite rib structures.

SPRING IN

SCATTER

SHRINKAGE

As covered in reference 10, the major issues influencing the ability to

control and maintain the geometrical accuracy of production composite

components are:- Spring In component deformation: Spring In Scatter:

and Part thickness Shrinkage. Tooling design to reduce these effects is

detailed in reference 10 section 4.

In order to determine the extent of these problems for the individual ribs

the Pheno-numerical strategy to predict Process Induced Deformation

developed in reference 18 is proposed using:- Simulation of systematic

“Spring In” deformation for the planned processing operation, and for the

“Spring In scatter” using statistical identification of scatter sensitive

production parameters. The shrinkage phenomenon will be dealt with by

experimental identification of modified Coefficient of Thermal Expansion

CTE and process sensitivity.

The desired outcomes of the application of the PRSEUS integrated rib lower flange concept are:-

an improvement in composite rib interface strategy for wing box assembly: comparable component

geometrical accuracy to Al Li wing ribs: structural weight reduction: and reduced manufacturing

costs.


129

Figure 43:- Composite Rib 31 from FATA 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.


130

Figure 44:- Composite Rib 31 from FATA 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.


131

Figure 45:- Composite Rib 31 FATA 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.

Fuel Vent Tank Systems


Rib overall Thickness

6mm (28plies)




As design weight in Hercules Inc AS4

Multiaxial fabric CF infused with

Hexflow VRM-34 Epoxy resin = 8.234kg.


132

Figure 46:- Composite Rib 31 FATA 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.

Wing Top Cover Skin.


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.

Top Cover Skin 20mm fasteners.


Proposed assembly methodology for Stitched Split Rib 31 subsequent integration into the PRSEUS

tapered stringers / skin assembly is shown below in figures 48 to 50 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 48(a). 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 48(b). 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

48(c) 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 49 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 the Boeing CAPRI 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 50. 133

Roll and layout of large aircraft wing structural members (CFC wing ribs).


134

Figure 47:- Composite Rib 31 FATA Split Rib, with PRSEUS tapered stringers.

As design weight in Hercules Inc AS4 Multiaxial fabric

CF infused with Hexflow VRM-34 Epoxy resin = 7.22kg.

UP

FWD

OUT BD Fuel Vent Tank Systems


Rib Integral Cleat for Rib to Trailing

Edge Spar build joint with single row

of 16mm fasteners (provisional).

Single stage Flange Joggling for

tapered stringer flanges.






(provisionally additional 6mm 28 plies).

Rib overall Thickness

6mm (28plies)

Reduced cutout width for PRSEUS

Cover Skin Stringers.

Flange attachment fasteners 14mm (provisional).


135

Figure 48:- Composite Rib 31 Stitched Integral Flange / Web Preform assembly.

Tare Strip

(1.5mm)

Figure 48(a)

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 48(b)

Low level fuel transfer holes.

Figure 48(c)

Aft Coaming Stringer Section

Fwd Coaming Stringer Section

Section of lower cover skin

(representative)

Fwd Coaming Stringer Cut-out


136

Figure 49:- Composite Rib 31 Stitched Integral flange and PRSEUS Coaming stringers.

Figure 49(a) Side view on (B)

Figure 49(b) Plan view

Figure 49(c) Front view on (A)

(Coaming Stringers omitted for clarity.)

(A)

(B) Aft Coaming Stringer Section Fwd Coaming Stringer Section

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


137

Figure 50:- 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 Integral Rib Flange / Web Preform section is mounted and

stitched in place and the resulting assembly is infused with Hexflow VRM-34 Epoxy Resin

using 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.

Aft Coaming Stringer Section

Fwd Coaming Stringer Section

Integral 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.


138

Figure 51:- Composite Rib 31 FATA Prime with PRSEUS tapered stringer assembly.

Trailing Edge

CFC spar.

UP

FWD

Leading Edge

CFC spar.

Wing Top Cover Skin.


Leading Edge

Ti Rib Post.

Fwd Bottom Cover Skin PRSEUS

Coaming Stringer. 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. Top Cover Skin 20mm fasteners.

Aft ventilation. Aft ventilation hole. Fwd

ventilation.


Fwd fuel

drain.


The rib alignment and rib spacing has to be established at an early stage in the preliminary design

phase, since the weight of the ribs contributes significantly to the total wing box structural weight,

therefore rib layout configurations were run through the AeroDYNAMIC™ MDO toolkit at the start of

the wing design process. It is advantageous to select a lager rib spacing; equal structural weight it

leads to cost savings and less fatigue risks. The rib spacing will increase with the depth of the wing

box, hence considering the typical wing which is tapered in planform and depth, the optimum wing

structure would have a variable rib spacing with the maximum spacing inboard and minimum

spacing outboard.

The wing rib arrangement outside the root interface is critical for designing the compression

structural stability of the wing box members especially the upper cover skin, and the rib spacing is

as important as the root joint design, ideally the rib spacing should be determined to ensure

adequate overall buckling support to the distributed flanges, and this requirement gives the

maximum theoretical pitch of the ribs. However other practical considerations are likely to

determine the actual rib locations such as:- (a) Hinge positions for control surfaces and attachment

/ operating points for flaps, slats, and spoilers: (b) Attachment locations of powerplants and landing

gear structure (and stores for military derivative airframes P-8 etc.): (c) The need to prevent or

postpone skin local shear or compression buckling, as opposed to overall buckling: (d) Ends of

integral fuel tanks where a closing rib is required.

For the swept wing configuration there are two main options for rib alignment which are:- (1) In the

direction of flight shown in figure 52(a) and: (2) Orthogonal to the rear spar direction shown in figure

52(b). 139

Roll and layout of large aircraft wing structural members (wing ribs).


140

Figure 52:- Rib layout options for large swept wing aircraft.

Fig 52(a) Ribs laid out in direction of flight. Fig 52(b) Ribs laid perpendicular to the rear spar.

Front spar.

Rear spar.

Auxiliary spar.

Ribs.

Front spar.

Rear spar.

Auxiliary spar.

Ribs.

Front spar.

Rear spar. Auxiliary spar.

Transition

Rib.

Fig 52(c) Ribs laid in hybrid fan from line of flight to perpendicular to rear spar.

Perpendicular

Ribs.

Fight line

Rib.


While the direction of flight alignment for the ribs, option 1 (figure 52(a)) gives greater torsional

stiffness, but the ribs are heavier, connections are more complex, and in general the disadvantages

outweigh the stiffness gains. The orthogonal direction alignment of the ribs, option 2 (figure 52(b))

with the ribs at right angles to the rear spar is more satisfactory in facilitating hinge pick-ups, but

they cause layout issues in the root regions. It is possible to overcome these issues by fanning the

ribs so that the alignment changes from perpendicular to the spars outboard portion of the wing to

stream-wise over the inboard portion of the wing, (with the special exceptions for powerplant

mounting ribs which are best located in the fight direction), as shown in figure 52(c), and it was this

hybrid configuration which gave the best MDO analysis results and was selected for the baseline

wing configuration.

FIXED SECONDARY STRUCTURE:- A fixed leading edge is usually stiffened by a large number of

closely pitched ribs, span-wise members being absent. Providing care is taken in the detail design

of the skin attachments it is possible to arrange for little span-wise end loading to be diffused into

the leading edge and hence avoid buckling of the relatively light structure. Therefore these are

usually in short span-wise sections. The incorporation of thermal de-icing system, this is

traditionally performed using hot bleed air from the engines ducted along the wings leading edge

via a “piccolo” tube, with the spent air being exhausted through holes in the lower surface of the

wing or slat. However new systems like that developed for the Boeing 787 use an electro-thermal

system made up of several electrically heated elements contained within a sprayed metal matrix

bonded to the inside of the leading edges by a polymer composite material and can be energised

simultaneously or sequentially fig 53, and would be more compatible with NAW leading edges. 141

Rib alignment and fixed secondary wing structures.


In addition to the anti-icing system major influences on the detail design of the leading edge

structure are the installation of high lift slats and other devices driven by EHA‟s as shown in figure

53(a) and 53(b), as well as bird strike protection. The A350 Droop nose leading edge figure 53(a)

installed inboard of the engine, reduces low speed drag thus reducing engine thrust requirements,

and also reduces control surface noise.

Installation also affects the trailing edge structure where much depends on the type of flaps, flap

gear, controls and systems. It is best aerodynamically to keep the upper surface as complete and

smooth as possible, therefore where possible spoilers should be incorporated in the region above

flaps or hinged doors provided for ease of access. There are many types of trailing edge flaps used

to increase the maximum lift coefficient of the wing to shorten aircraft take-off and landing

distances. The design flap systems is more complex than leading edge systems and poses very

challenging design issues to be covered in this design study. The flap applied to the trailing edge of

a wing cross section usually takes up 25-35% of the chord length, and for some special mission

requirements this can rise to as high as nearly 40%. The determination of the flap chord length is

also a function of wing box structural stiffness and strength requirements as well as the volume

required for the wing fuel tank requirements to achieve the aircrafts performance requirements.

Therefore trade studies to investigate trailing edge requirements for the reference and advanced

wing were conducted before freezing the final configuration. Figure 53(c) illustrates the typical

trailing edge arrangement control arrangement. New innovations in flap design are being

incorporated on the Airbus A350 XWB an example being the Drooped Hinge Flap as an alternative

to the Fowler Flap, which has the benefits of being able to be used as both a high lift device and in

flight adaption of the cruise wing shape figure 55. 142



143

Figure 53:- Control surface arrangement on large swept wing aircraft.

Fig 53(c) B787 trailing edge control surfaces.

Fig 53(a) A350 Droop nose leading edge,

driven by Electro-Hydrostatic Actuators

(EHA‟s) with EBHA‟s.

Fig 53(b) A350 Control surface general arrangement. Fig 53(d) B787 leading edge ice protection.


As the baseline aircraft is conceived as an all electric aircraft using only electric power for all of the

systems for control surface actuation, as is the case on the Boeing 787, and in the Airbus A350

family, although the A350 also has an emergency backup hydraulic actuation system as per figure

53(a). This replaces the former massive hydraulic systems of other aircraft, which required several

miles of high-pressure piping, the costly non-inflammable hydraulic fluids, large and heavy linear

and rotary actuators, and the very large number of spool valves, seals, accumulators and other

auxiliary devices which comprise a conventional hydraulic control actuation system. This reduces

the weight and complexity of the flight control actuation system, with the additional impact on the

engine of the removal of the need for engine driven pumps and engine mounted accessory

gearboxes. To generate aircraft electrical power a Rolls Royce proposal to use shaft-mounted

starter generators, starting the engine and interchanging energy between shafts of the proposed

triple-shaft engine in flight, also driving PM machines off the Low Pressure fan is a possibility where

emergency power could be extracted when the engine is windmilling.

In place of the hydraulic system the FATA baseline wing incorporates an extensive direct-current

electrical system reflecting the current state of the art employing fault tolerant 270DCV electrical

power generation systems. The actual actuation system was a choice between the Electro-

Mechanical Actuation System (EMAS), or the Electro-Hydrostatic Actuators (EHA‟s), the EMAS

usually requires a speed reducing gearbox and although these can be quite light it is viewed as

adding to the overall complexity of the system, therefore EHA‟s have been selected for this baseline

technology study, where electric power drives a self-contained hydraulic actuator. The baseline

sized EHA‟s and their applications are shown in figures 54(a) and (b) and the Low Speed Aileron

EHA actuator integration to the outboard trailing edge spar is shown in figures 54(c). 144



Height = 0.28m.

Weight = 8.01kg (est.)

145

Height = 0.86m.

Weight = 29.0kg (est.)

Figure 54(a)(b):- Control surface actuators proposed for baseline FATA wing study.

Figure 54(a): - EHA actuator for flap actuation

(used in structural sizing). Source authors

private collection.

Figure 54(b): - EHA actuator for leading edge

slat and aileron actuation (used in structural

sizing). Source authors private collection.


14mm Thick Zone

(66 Plies)

10mm Thick Zone

(46 Plies)

6mm Thick Zone

(28 Plies)

1:20 Transition Zone

(4 x 80mm )

1:20 Transition Zone

(4 x 80mm )

UP

FWD

OUTBD

EHA Actuator Supports Hinge Ribs

Figure 38(g): Outboard Trailing Edge Spar.

EHA Support Brace 1

Datum‟s

Low Speed Aileron

Hinge Rib 1 Datum

Aileron Attachment Pin CL

Low Speed Aileron

Hinge Rib 2 Datum

EHA Support Brace 2

Datum‟s

Aileron

Attachment Pin CL Aileron

Attachment Pin CL

EHA Drive Lines

FWD

INBD

UP

146

Figure 54(c):- Trailing edge outboard spar actuator supports and hinge ribs datum's.


147

Figure 55:- Current advanced control surface on the A350 large swept wing aircraft.

Figure 55 shows the Advanced Drooped Hinge Flap of the Airbus A350 XWB which replaces the

current flap tracks or linkages. This is a multifunction trailing edge flap system which can be

integrated for use as a high lift device and for in flight adaption of cruise wing shape.

The ADHF significantly improves

High-lift efficiency without increasing

weight and complexity, yielding load

alleviation and cruise efficiency

enhancements.


148

Figure 56:- Examples of the types of mission adaptive wing technology explored.

For the FATA evolved wing design the application of the

MAW technologies shown in figure 41 will be explored , the

possible benefits of mission adaptive wing technology are:-

1) Enhanced performance:

2) Fuel savings:

3) Drag reduction:

4) Noise reduction:

5) Weight reduction:

6) Reliability:

7) Gust load alleviation:

8) Ease of integration:

9) Reduced wing bending moment :

10) Cost effectiveness.

(ref 3)


Starting with the wing, the major drivers in the baseline wing structural design considered in this

study are: - 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 e.g. single point bonding for CFC wing

structures.

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 57): 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. 149

Section 7:- The design and structural layout of the FATA wing.


150

Figure 57:- My baseline aircraft 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. 151

The design and structural layout of the FATA wing.


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 baseline 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 see figure 58, which reduce induced drag.

152



153

Figure 58:- Examples of Winglet devices for modern single and wide body aircraft

Figure 58(a):- Boeing 737MAX wingtip

device increases efficiency by:-

Combining rake tip technology with a dual

feather winglet concept:

Reduces fuel burn up to an additional

1.5%:

Fits within current airport single-aisle gate

constraints:

Validated by wind tunnel testing.

Figure 58(b):- Airbus A350-900 wingtip device

increases efficiency by:-

Raked saber winglet of advanced composite

manufacture:

Reduces fuel burn by reducing induced drag:

Fits within current airport wide body gate

constraints:

Validated by wind tunnel testing and flight

testing.


Winglets.

A variety of devices have been used on aircraft to reduce induced drag figure 58 shows two of the

latest such devices for the Boeing 737 Max single-aisle transport figure 58(a), and for the Airbus

A350 XWB wide-body transport figure 58(b). These devises inhibit the formation of wing tip vortices

and therefore reduce downwash and induced drag.

A similar effect could be achieved by extending the wing to increase its span and aspect ratio ,

however, the increased lift far out at the end of the wing will increase the bending moment at the

wing root and create greater loads on the wing root structure, requiring larger and heavier wing root

fittings and skins.

The winglet only increases the wing span slightly and therefore achieves the increase in aspect

ratio without significantly increasing the wing root structural loading. The winglet configuration

selected for the baseline wing study is based on the saber design for the A350 XWB made from

epoxy carbon fibre composite, with an internally co-bonded Waffle structure preforms (see figure

58(c) below), in the blade where the depth is less than 4” (100mm to 75mm), the root section being

CFC spars, based on GKN Aerospace technology shown in figure 58(d) on the next slide.

154

Bondline.

Figure 58(c) Proposed internal structure of baseline winglets.



155

Figure 58(d):- Examples of Winglet devices for modern single and wide body aircraft.

One possible option for FATA winglet construction based on GKN Aerospace STeM

research see reference


Higher wing spans improve aerodynamic efficiency and reduce fuel burn as demonstrated in the

Boeing / NASA Subsonic Ultra Green Aircraft Research (SUGAR) Super Refined SUGAR aircraft

study (ref 12), amongst others, however high span wings crate airport compatibility issues in terms

of gate space / footprint and minimum spacing. Although to some extent this can be alleviated by

the adoption of a Code E wing with a winglet as discussed above a solution for very high span

wings is to incorporate a wing fold to a Code F wing (ref 13) figure 59(a), maintaining performance

of the high span wing whilst being able to meet gate requirements, and this has been proposed by

Boeing to meet current airport requirements for the 777-9. This solution was previously proposed

for the 777-200 but was found to be an over complex solution, and was not adopted, however as

can be seen from figure 59(b) the amount of wing to be folded is considerably less and does not

include any control surfaces so should be lighter and much less complex. Figure 59(c) gives an

early indication of the proposed hinge line and drive mechanism.

Such a folding wing tip has the following system safety and functional requirements:- (1) The wing

tip is designed and tested to the same requirements as flight critical control surfaces that is it will

have redundant load paths; can be isolated in flight; in failure mode is latched and locked : (2) The

tip will take 20 seconds to reach a pilot commanded position, from the position input from the pilot:

(3) With the tips folded (up) the aircraft can withstand Cat 1 hurricane winds (74 - 95mph) without

ground support equipment (GSE) installed or hydraulic power: (4) It must be possible to conduct

pre-flight control checks and de-icing with the wing tips in any position: (5) The pilot must be able to

determine the folding wing tip status without external input from either ground crew or the control

tower staff. These requirements will be met by the Boeing 777-9 wing tips. 156



157

Figure 59(a): - Advantages of the application of a wing fold for high span wings.

*The trade between weight and complexity against fuel burn

savings will be an interesting future trade study.


158

Figure 59(b): - Early proposals for the wing tip fold on the 777-9 code F wing.


159

Figure 59(c): - Early proposals for the wing tip hinge and actuator layout.


Leading and trailing edge device integration:- The integration of leading and trailing edge

devices requires that the following criteria must be considered:- Leading Edge devices are subject

to bird strike and the actual Leading Edge must be replaceable: Erosion protection of the Leading

Edge must be considered: All devices must be bonded for EMC and lightening strike protection see

figures 28 and 29: Selection and retention of bearings is critical: Actuation must allow for wing

deflection: Clearance checks are required between inboard and outboard flaps during deployment

especially if the hinge line is kinked: Trade studies will be required to determine the optimum

method of actuation, and for sealed versus non-sealed gaps at the interface with the wing torsion

box.

For trailing edge flaps on swept wings a real difficulty arises when the effective hinge-line is swept.

It is possible to arrange the geometry so that the flap is deployed at right angles to the hinge line,

that is, along circular arcs on the conical surface. This often implies that any external hinge

brackets or tracks are positioned across the airflow with a consequent drag penalty. Alternatively a

swept flap may be moved along the line of on elliptical paths described on the surface of a circular

cone, which leads to complex geometry. (The deployment of the outboard single pivot flap is to be

validated using the Catia V5 Kinematic Simulation following the principles of Kevin Beyer and Lee

Krueger presentation „Design Validation Through Kinematic Simulation: Airplane Flap Design‟

presented at the PLM Conference 2010 Las Vegas Nevada USA).

Leading edge slats move out on circular arc tracks, which are usually attached to the slat, with the

support rollers being mounted in the fixed leading edge structure. Most designs use a short length

of slat located on two attachments, with actuation also usually at the track position, often by means

of leavers, or rack and pinion gears driven by EHA‟s. 160



The base line aircraft wing control surfaces as shown in figure 60 on the port exposed wing surface,

consisting of the following:- One Inboard Slat: Six Outboard Slats: One Inboard Flap: One Outboard

Flap: Three Inboard Spoilers: Four Outboard Spoilers: One Flaperon: One All Speed Aileron: and

One Low Speed Aileron and these were duplicated on the starboard wing. These control surfaces

were sized using classical methodology from reference 4 and outputs from AeroDYNAMIC™ MDO

toolset, these are initial evaluations and are subject to revision as the project progresses the first

pass sizings in surface area are given below.

Trailing Edge Surfaces:-

Inboard Flap = 6.118m² (9482in²): Spoiler Inboard (1) = 2.02m² (3135in²): Spoiler Inboard (2) = 2.02m²

(3135in²) :

Outboard Flap = 8.597m² (13324in²): Spoiler Outboard (1) = 1.71m² (2644in²): Spoiler Outboard (2) = 1.71m²

(2643in²): Spoiler Outboard (3) = 1.71m² (2642in²): Spoiler Outboard (4) = 1.70m² (2641in²): Spoiler Outboard

(5) = 1.70m² (2640in²).

All Speed Aileron = 4.07m² (6310in²): Low Speed Aileron = 4.07m² (6305in²).

Leading Edge Surfaces:-

Inboard Slat = 6.282m² (9737in²).

Outboard Slat (1) = 3.361m² (5209in²): Outboard Slat (2) = 3.336m² (5170in²): Outboard Slat (3) = 3.310m²

(5130in²): Outboard Slat (4) = 3.284m² (5089in²): Outboard Slat (5) = 3.258m² (5049in²): Outboard Slat (6) =

3.232m² (5008in²).

The final structural sizing was conducted after freezing of the control surface sizing:- wing semi

span = 32.37m (106ft 2in), root chord = 13.97m (45ft 10in), tip chord = 3.81m (12ft 6in), semi span

area = 226.291m² (2,435.78ft²).

161

Layout of FATA aircraft wing flight control surfaces.


162

Figure 60:- My baseline aircraft wing flight control surface layout model.

Six Outboard Leading edge slats.

Engine Center Thrust line.

Wing Carry

Trough Box

Attachment

Joint line.

Low Speed Aileron.

All Speed Aileron.

1 2

Outboard Flap

single pivot.

Inboard Flap

single pivot.

Two Inboard

Spoilers with

droop function.

Five Inboard

Spoilers with

droop function.

Droop nose Leading edge slat.

Note: - Three flap track fairings, one on the inboard flap,

and two on the outboard flap.


Wing torsion box layout is shown in figures 61(a),(b),(c) and constitutes a datum structural layout of

the primary structure. This has the best performance over the AeroDYNAMIC simulation mission

and is the Prime Baseline Wing and will be carried forward to structural detailed layout, and detailed

part sizing, with conventional materials. The Prime Baseline Wing will be reconfigured for PRSEUS

based stitched structure technology, as the Advanced Baseline Wing, for comparison with the

Prime Baseline, in teams of weight, structural integrity, manufacture, and assembly. Figure 62

illustrates what the datum surfaces represent for metallic and composite structures.

Baseline wing structural components:- Leading edge spar:- 35.826m (117.54ft) divided into 3

sections:- inboard spar 12.11m (39.73ft): mid spar 17.04m (55.92ft): outboard spar 6.68m (21.92ft):

C-section carbon fibre epoxy resin fibre placed monolithic construction with sacrificial plies for

interface control of the titanium splice joints and fittings of bolted assembly.

Trailing edge spar:- 33.31m (109.28ft) divided into 3 sections:- inboard spar 9.49m (31.14ft): mid

spar 17.11m (56.14ft): outboard spar 6.70m (21.98ft): C-section carbon fibre epoxy resin fibre

placed monolithic construction with sacrificial plies titanium splice joints and fittings of bolted

assembly.

Centre Spar:- 9.07m (29.77ft) single unit C-section carbon fibre epoxy resin fibre placed monolithic

construction with sacrificial plies and titanium fittings.

Ribs:- 37 in total:- 1 stub rib to support engine pylon fwd attachment, 25 Al li ribs, plus 11 CFRP ribs

with integral leading edge cleats.

Auxiliary Gear Spar:- Ti double sided 5 axis machining „I’- section integral stiffeners 7.64m

(25.07ft).

163

Datum layout of baseline aircraft wing torsion box structural members.


Figure 61(a):- Baseline wing torsion box key datum layout structure model R.4.

Wing Torsion Box Structure with transparency applied to top cover skin

to show spar, revised rib, and revised stringer layout.

All stringers I – section

co-bonded to the skin.

Ti I-section Gear

auxiliary spar.

Three monolithic Carbon Fibre

Epoxy Resin C-section Spars.

Upper cover skin monolithic

Carbon Fibre Epoxy Resin.

(Transparent for structure view)

Slat track ribs currently machined

but possible candidate for AM.

Engine Center Thrust Line with wing box main rib

on thrust line for pylon fwd attachment plate and

additional firewall L/E ribs Ti and Ti engine fire

wall on spar and upper cover.

Slat track ribs currently machined

but possible candidate for AM.

Carbon Fibre Epoxy Resin

ribs with integral Leading

edge cleat (green)

Al Li monolithic ribs

(dark blue).

Ti I-section MLG

Kick spar.

164


165

Figure 61(b):- Baseline wing torsion box key datum layout structure model R.4.

Lower Wing Torsion Box Structure with top cover skin and stringers

removed for clarity to show spar, revised rib, revised inspection cut

outs and revised stringer layout.

Inner Spar section

(Leading Edge).

Lower cover skin monolithic

Carbon Fibre Epoxy Resin.

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.

Outer Spar section

(Leading Edge).

Outer Spar section

(Trailing Edge).

Rib attached by countersunk

bolts through skin and to

anchor nuts bonded to the rib

internal flange surface.

30º Chamfered edges to

reduce toe peel stresses.

All stringers I – section

co-bonded to the skin.

Mid Spar section

(Leading Edge).

Inner Spar section

(Trailing Edge).

Ti I-section Gear

auxiliary spar. Mid Spar section

(Trailing Edge).

Coaming stringers.

Inspection cut outs.

Carbon Fibre Epoxy Resin

ribs with integral Leading

edge cleat (green)

Al Li alloy ribs (dark blue).

Co-bonded Wing cover skin stringer design chamfered edges

to reduce peel stresses (see also figs 25, 40, and 47).

Intermediate Spar


166

Figure 61(c):- Baseline wing torsion box key datum layout structure model R.4.

Spar splice joints.

Low speed aileron

hinge ribs and

actuator supports.

Spar splice joints.

All speed aileron

hinge ribs and

actuator supports.

Lower Wing Torsion Box Structure with top cover skin

and stringers removed for clarity to show spar, revised

rib, revised inspection cut outs and revised stringer

layout.

Flap track ribs.

Spar splice joint.

Spar splice joints.

Actuator supports.

Flap track ribs.

Ti I-section Gear

auxiliary spar.

Gear bay skin support

structure CFC.

Flap track ribs.

Actuator support. Ti I-section MLG

Kick spar.


(1):- Metallic ‘I’- beam also applies to

CFRP „I’- section (back to back „C‟

sections).

(2):- Metallic „C‟- section. (3):- CFRP „C‟- section.

Datum plane / surface

In middle of web.


On tool face of web.


On tool face of web.

167

Figure 62:- Key datum's in the layout structure models.

Key datum models show datum positions upon which actual detailed structure will be located when

sized this slide is intended for non / new designers and shows what the model datum‟s represent.


The wing carry through box layout is shown in figures 63(a), (b), and (c). The following structural

layout has been used:-

Spars:- CFRP monolithic laminate C section with co-bonded web stiffeners (shown in dark

green):

Skins:- CFRP monolithic laminate (shown in green), with 11 top and 11 bottom spanwise

tapered flange „I’ section CFRP solid laminate stiffeners with tapered flanges (shown in dark

green) on developed airframe these are PRSEUS stringers:

The seven internal upper and seven lower chordwise Al/Li load beams (dark blue) to which are

attached 56 angled CFC tube struts (shown in light blue) and 21 vertical CFC tube struts

(shown in orange), The Tube Struts are produced from unidirectional tape automated cross-ply

wound around a foam core, and subsequently is autoclave cured, with Ti alloy embedded

fittings:

The root ribs are currently Al/Li alloy (shown in dark blue), but there is the option to change this

to CFC for the evolved FATA wing depending on further structural analysis:

The seven over wing floor beams are „I’ section CFRP solid laminate which are co-bonded to

the top WCTB and MLGB cover skins, and have splice attachments at each end, also shown in

figure 63 is the port / starboard Al/Li trap panel.

The Wing Carry Trough Box to Fuselage attachments which are Ti alloy machining's attached

to the Root Ribs and leading /trailing edge Spars and to the CFC Fuselage frames:

Keel Beam is a CFC box beam made by RTM, a secondary support keel runs through the

MLGB and supports the APU fuel line and bracketing.

168

Structural layout of baseline aircraft wing carry through box initial sizing.


169

Figure 63(a):- My FATA baseline wing carry through box full structural model R.6.

Titanium alloy Root Rib to Fuselage Interface

beam attachments (7-off Port /7-off Stbd).

I section Carbon Fibre Epoxy Composite Over-wing

Floor beams (7-off WCTB and 7-off MLGB)

WCTB top CFC Cover Skin. MLGB CFC Cover Skin.

Main Landing Gear Bay


pressure bulkhead Al/Li.

Cargo Bay closure bulkhead Al/Li.

Trap panels Port and Starboard Al/Li. Port / Stbd Al Li Root ribs.

WCTB tank closure

seal caps Ti alloy.

UP

FWD


170

Figure 63(b):- My baseline wing carry through box internal structural layout R.6.

Al/Li Load Beams (14-off) 7 top and 7 bottom.

WCTB CFC Leading Edge Spar.

WCTB CFC Intermediate Spar.

WCTB CFC Trailing Edge Spar.

Top Horizontal Triform.

Vertical CFC Tube struts (21-off)

three locations (brown).

Angled CFC Tube struts

(56-off) (light blue).

MLG Bay Keel I Beam CFC

RIM for APU fuel line support.

Fuselage Keel Bottom Box Beam

CFC RIM Extrusion.




171

Figure 63(c):- My baseline wing carry through box structural layout R.6.

WCTB CFC Cover co-bonded Skin stiffeners

(22off) 11 on top skin and 11on bottom skin.

MLGB CFC Cover Skin.

WCTB top CFC Cover Skin.

WCTB Bottom CFC Cover Skin.

Bottom Horizontal Triforms.

Cargo Bay trailing edge spar

closure bulkhead Al/Li.

MLG Bay Keel I Beam CFC RIM

for APU fuel line support.

APU fuel line support brackets.



Fuselage Keel Bottom Box Beam

CFC RIM Extrusion.

APU Fuel line.


The careful arrangement of the wing fuel tank layout (see figure 64(a) and (b) for the initial FATA

baseline wing), from the initial design stages of a commercial aircraft can result in a lighter

structural weight through bending moment relief. The fuel management system is an important

consideration in the structural design of an aircraft, and in addition to the wing tankage the wing

carry through box is also usually a fuel tank.

The way in which the tank fuel tank layout and fuel management in commercial aircraft wings

influences wing bending moment relief is shown by the three cases considered in figure 64(c)

below, i.e. the weight of fuel in the tanks acts down at its centre of gravity (c.g.), thus creating a

downward bending moment which is counter to the lifting upwards bending moment at the root, and

these downwards bending moments are subtracted from the root lift bending moment to obtain the

final root bending moment.

Case A (figure 64(c)):- In this case there are two wing fuel tanks, and by feeding first from the

inboard tank and subsequently from the outboard tank, a fuel weight wing bending moment

relief corresponding to track A is obtained:

Case B (figure 64(c)):- In this case there are also two wing fuel tanks however the inboard

tank is much longer than the inboard tank in case A. Therefore its c.g. remains further

outboard and the fuel weight wing bending moment relief corresponding to track B is obtained:

Case C (figure 64(c)):- In this case there are three wing fuel tanks and by feeding first from the

root tank, next from the mid wing tank, and finally the outboard tank, a wing bending moment

relief corresponding to track C is obtained, which is of the highest magnitude. This latter case

has been selected for the FATA baseline wing box. 172

Wing fuel tank layout effect on bending moment relief.


173

Main inboard fuel tank

Main mid wing fuel tank.

Outboard reserve fuel tank, and surge and tip vent tanks.

Main fuel tanks are shown with nominal off set for skin

thickness (light blue, and pink) the initial estimated

total maximum capacity of all tanks is 95,500lts

(21,007 Imperial gallons) estimated from volume

envelope.

Figure 64(a):- My baseline FATA wing torsion box initial fuel tank layout.


174

Figure 64(b):- My baseline FATA initial wing fuel tank layout overview.

Wing Carry Through

Box fuel tank

Port Mid Wing fuel tank

Stbd Mid Wing fuel tank

Port Inboard Wing fuel tank Port Inboard Wing fuel tank

Port Outboard reserve fuel tank,

and surge and tip vent tanks.

Port Outboard reserve fuel tank,

and surge and tip vent tanks.


175

Root Tip

CASE (A)

Inboard fuel tank Outboard fuel tank

Root Tip

CASE (B)

Inboard fuel tank Outboard fuel tank

Root Tip

CASE (C)

Inboard fuel tank Mid wing fuel tank

Outboard fuel tank

Tip Root

WIN

G B

EN

DIN

G M

OM

EN

T R

EL

IEF

.

CASE (A)

CASE (B)

CASE (C)

Figure 64(c):- Fuel tank layout for maximum bending moment relief.


The wing root joint design is one of the most critical areas of the aircraft structure, especially for

fatigue considerations of a long life structure. The joint for the Prime Baseline Wing will be carried

over for the Advanced Baseline Wing and then will be re-evaluated for the Future Concept Wings.

The types of joint available for fixed swept wing large transports are outlined table 7 below, and the

option which has been selected as the design solution has been skin splice joints across the root rib

flanges and splice plates attachments for the spars as shown in figure 66(a) through (c).

Table 7:- Wing Root fixed joints.

176

The wing torsion box to wing carry through box root fitting.

JOINT TYPE. ADVANTAGES. DISADVANTAGES.

Spliced plates. Widely used due to its light weight and

more reliable and inherently fail-safe

nature.

Higher cost, and manufacturing and

fitting issues, the latter of which could

be reduced with cover skin sacrificial

plies.

Tension bolts. Less manufacturing, easy to assemble

and remove and inspect, common on

fighter aircraft

Heavy weight penalty.

Discrete lug fittings with shear

bolts.

As for tension bolts and I have greater

experience with designing this type,

common on fighter aircraft.


Combinations of tension bolts / or

lug fittings, and spliced plates

Reliable and inherently fail-safe feature,

and less manufacturing and fitting

issues.



The torsion box root loads are described below and the distributed loads on discrete fittings are

illustrated in figure 65(a). Figure 65(b) illustrates a splice plate arrangement for a metallic integrally

stiffened lower wing skin joint. The solution to be used here is splice joints for the spars and upper

Triform and lower Triform splice joints for the skins with the skins landing on the Root Rib flanges

and capped by the Triforms taking loads into the WCTB and Fuselage.

177

Figure 65:- The wing torsion box to wing carry through box root fitting.

Shear Shear

Shear Shear

Moment Moment End Load

End Load

Minimal intrusion into the fuselage.

Drag Drag

Uneven load

distribution across

fittings.

Fittings carry end load + shear +drag.

Figure 65(a): - Distributed Fitting Loads.

Figure 65(b): - Splice shown for a

metallic integrally stiffened wing.


178

Figure 66(a):- The wing carry through box root rib integration.

Wing Carry Through Box Top Cover

Skin (floor beams omitted for clarity).

Wing Torsion Box Cover

Skin Sections Spliced to

Root Rib by Triforms.


Bottom Horizontal Triform. IB wing Leading Edge

Spar Spliced to Root Rib.

IB wing Intermediate Spar

Spliced to Root Rib.

IB wing Trailing Edge Spar

Spliced to Root Rib.

Port Root Rib.

Wing Carry Through Box Root Rib to Fuselage Frame attachments

bolted to Root Rib inboard face and to frames.

IN BD

FWD

UP


Figure 66(b):- Component assembly of the wing root splice (Port and Stbd).


Bottom Horizontal Triform

Port Root Rib.

Port Rib Crown

Sealing Strips.

Wing Carry Through Box

Leading Edge Spar.


Intermediate Spar.


Trailing Edge Spar.

IB wing Leading Edge

Spar end section.

IB wing Intermediate Spar

end section.

IB wing Trailing Edge Spar

end section.

LE Spar OB

Splice Plate.

LE Spar IB Splice

Bathtub.

IM Spar IB Splice

Bathtub.

IM Spar OB

Splice Bathtub.

TE Spar IB Splice

Bathtub.

TE Spar OB

Splice Plate.

UP

IN BD

179


180

Figure 66(c):- Component assembly of the wing root splice (Port and Stbd).

Wing Carry Through Box Root Rib

to Fuselage Frame attachments.

Port Rib Crown

Sealing Strips.

Wing Carry Through Box Root Rib

to Fuselage Frame attachments.

IB wing Trailing Edge Spar end section.

TE Spar IB

Splice Plate.

TE Spar OB

Splice Plate.

Top Horizontal Triform lands on both cover skins and bolts through rib flange and crown.

Bottom Horizontal Triform lands on

both skins and bolts through flange.

UP

OUT BD

Port Root Rib.


The FATA Fuselage Baseline will use the skin panel structural layout of the Airbus A350 as the

initial structural starting point as the longitudinal joints participate in the fuselages bending

resistance and therefore imparts superior bending strength, and each panel can be optimised for

its specific design case as shown in figure 67(a), these panels being as long as practical to

minimise the number of circumferential joints. Although this is a more complex manufacturing

option than the Boeing 787 barrel option which allows a single process to be used throughout, as

shown in figure 67(b). The Airbus system offers greater compatibility with FATA Prime fuselage

PRSEUS stitched stringer manufacturing technology, although the noes and tail are barrel sections

as shown in figure 67(c), figure 67(d) shows initial cabin window sizing.

The FATA Fuselage Baseline frame stations and build joints are shown in figure 68(a):- Section

11/12 to Section 13/14 joint is at Frame 25: Section 13/14 to Section 15 joint is at Frame 52:

Section 15 to Section 16/18 is at Frame 91: Section 16/18 to Section 19 join is at Frame 118: and

the Section 19 to Section 19.1 joint is at Frame 135. The build joint philosophy selected for the

FATA Fuselage Baseline was the Airbus style lap joint with a splice strip located on frames, as

shown in figure 68(c). This was adopted in preference to the Boeing butt joint with splice strap

located between frames as shown in figure 68(b) which although easier to inspect and maintain but

was not see to enable the direct coupling for high loads that Airbus style lap joint ensures. The

FATA Fuselage Baseline frame design concept adopted the Boeing full depth bolted Z – section

frames mounted over co-bonded hat stringers as shown in figure 68(d), this offered lower

complexity and weight over the Airbus clip based philosophy, more closely resembling the FATA

Fuselage Prime build for analysis, with full depth frames. Table 8 gives a comparison of frame and

stringer types and pitches from which the FATA baseline values were derived.

181

Section 8:- The design and structural layout of the FATA fuselage.


Figure 67(a):- Design philosophies for composite fuselages ( A350 XWB).

182

A350 Design philosophy:- In order to reduce the operating costs and

environmental impact through reduced fuel burn the airbus A350 adopted the use

of a four composite panel layout for the fuselage skins in the areas shown above.

The key attributes of this layout:-

The skin panels are as long as possible to reduce the number of

circumferential joints:

The longitudinal joints participate in the fuselage resistance to bending hence

increasing bending strength:

Each panel can be optimised for its design case:

Significant weight reductions can be achieved by this design philosophy.

13m 20m 14m

Side panels.

Top panel.

Keel panel.

Port Side panel.


Figure 67(b):- Design philosophies for composite fuselages ( B-787).

Contoured section. Constant section. Nose section.

Section 48. Section 47. Section 46. Section 44 / 45. Section 43. Section 41.

Fwd body joint. Aft body joint. Centre section

joints.

Aft section joint.

Boeing 787 Design philosophy:- Multiple filament wound barrel sections with major circumferential

splice joints between sections 41 to 43, and 46 to 47. These barrel sections allow a single

manufacturing process to be applied to constant, contoured, and nose sections of the fuselage.

Resitting hoop stresses better than metallics, this allows higher cabin pressures, and larger

windows. 183


Figure 67(c):- FATA Fuselage design showing frame stations and build joints.

C25 C52

Section 11/12

C91

Section 13/14

C118 C135

Section 15 Section 16/18 Section 19

Section 19.1

FWD Fuselage Centre Fuselage AFT Fuselage

Ground Line

Radome.

Nose Landing Gear.

Frame 20 FWD

Pressure Bulkhead

(*Flat design).

Frame 20 Fwd Pressure Bulkhead being of Flat

design, reacts the pressure loads in bending

and will be machined from with a grid of vertical

and horizontal support beams.

FWD Body Mate Joint.

FWD Centre Body Mate Joint.

Wing Carry

Through Box.

AFT Centre Body Mate Joint.

Main Landing Gear.

Aft Body Mate Joint.

Frame 124 AFT Pressure Bulkhead

(*Curved Membrane design).

Frame 124 Aft Pressure Bulkhead being of Curved Membrane

design, reacts the pressure loads in tension and will have radial

and annular crack stoppers. These are lighter than the flat design,

and provisionally a stitched CFC membrane similar to the A380 aft

bulkhead is proposed.

CFC Barrel

section

CFC Barrel

section

4 CFC Panel

section

4 CFC Panel

section 4 CFC Panel

section

184


Figure 67(d):- FATA Fuselage cabin window showing sizing and spacing.

540mm

508mm

305mm

Proposed cabin window size is 508mm high

by 305mm wide viewing area, with a frame of

51mm. The proposed separation is 540mm

and the fuselage skin thickness in the window

band will increase to 4.25mm (based on initial

calculations).

51mm

51mm

185


186

Figure 68(a):- FATA Fuselage design showing fuselage build and panel joints.

Fwd CFC Barrel Section

(B787 example).

Aft CFC Barrel Section

(A350 example).

Section 15 CFC Keel Panel

(A350 example).

Section 13/14 Crown panel:-

Area = 68.051m²: Length = 12.792m

Section 15 Crown panel:-

Area = 110.074m²: Length = 20.645m

Section 16/18 Crown panel:-

Area = 74.569m²: Length = 14.391m

Port Section 13/14 Side panel:-

Area = 39.176m²: Length = 12.792m

Port Section 15 Side panel:-

Area = 63.635m²: Length = 20.645m

Port Section 16/18 Side panel:-

Area = 47.865m²: Length = 14.391m

Section 13/14 Keel panel:-

Area = 87.734m²: Length = 12.792m

Section 15 Keel panel:-

Area = 92.423m²: Length = 20.645m

Section 16/18 Keel panel:-

Area = 94.923m²: Length = 14.391m

Analysis will be of Section 15 Panels initially. (*Note Port and Stbd side panel dimensions are identical).


187

Figure 68(b)(c):- FATA Fuselage baseline build joint concept options (reference:-20).

MID SECTION SKIN

FWD SECTION SKIN

SPLICE STRAP

STRINGER

SPLICE ANGLE

SPLICE FILLER

SPLICE C-CHANNEL

Figure 68(b):- Example Butt-Joint with splice

strap located between frames, therefore this

type of build joint can be easily inspected and

maintained.

PANEL

STRINGER

BUTT-STRAP

DOUBLER

STRINGER COUPLING (U-SECTION)

FRAME CLIP WITH STABILIZER

ON STRINGER COUPLING

FRAME CLIP ON BUTT-STRAP

Z- CROSS SECTION FRAME

Figure 68(c):- Example Lap-Joint with splice strap

located on frames, enables direct coupling for high

loads but is more of a challenge to inspect.

Fwd Fuse Mid Fuse


188

Ω-Stringer Co-Bonded to Skin.

Clip PPS Thermoplastic Matrix composite with quasi-isotopic layup

Frame CFRP prepreg.

120mm

Z- Frame lay up [30º/90º/-30º]

with 0º reinforcement.

80mm

Airbus A350 Co-Bonded CFC Ω-Stringers. Boeing 787 Co-cured CFC Hat-Stringers.

Figure 68(d):- FATA Fuselage baseline frame / stringer concept options (reference:- 20).

Airbus A350 Bolted Z- Frame assembly. Boeing 787 Bolted Z-Frame assembly.


Considering the actual lay out of the FATA fuselage baseline structure it is a combination of both

Boeing and Airbus philosophies as described above.

In the case of the Airbus A350 all of the fuselage panels (except the cockpit cabin see figures 2 and

66(a)) are CFC as are stringers, window frames, clips, and doors, with hybrid door frames

structures consisting of CFC and titanium alloy. Depending on the location on the fuselage the

stringers are either Ω (shown in figure 68(d) note extended flanges for frame clip attachment) or T

profile and these are produced separately to the skin panels, and subsequently co-bonded on to the

panels. As shown in figure 68(d) the frames are attached by a mixture of thermoplastic clips and

special fasteners and are not bonded to the stringers or skin panels. Based on the difference in

fuselage length between the A350-900 XWB and the A350-1000 XWB fuselages which is 7m and

the reported difference in frame number of 11 frames the frame pitch can be estimated as 635mm.

Also the CFC window frames are of L – section profile but have the same strength as the heavier

traditional metal T - section frames (reference 20). From released photographs and presentations

from Airbus (reference 8) the Ω-section stringer pitch is estimated at ~ 230mm .

The Boeing 787 series are 50% CFC by airframe weight figure 2 shows the structural breakdown,

as with the A350 the 787 fuselage has Titanium alloy / CFC hybrid door frames as shown in figure

30(b). The construction modules of the fuselage are complete barrel sections as shown in figures

67(b) and 68(d), and reportedly (reference 20) each barrel section contains 80 hat /Ω-section

stringers, where the constant section is 5.46m (Airbus data) so the stringer pitch would be

~214mm. The average frame pitch on the B787 based on public data is between 574mm and

610mm and the latter figure is taken as the most representative.

189

The design and structural layout of the FATA fuselage (continued).


190

Table 8:- Comparison of FATA Fuselage frame / stringer layout with current aircraft.

Aircraft

type.

Fuselage

diameter (m)

Skin

material

Frame

material

Frame

profile

Frame

pitch (mm)

Stringer

material

Stringer

profile

Stringer

pitch (mm)

A320 3.96 Aluminium

alloy Aluminium Z 533 Aluminium Z ~150

A330 /

A340 5.64

Aluminium


A350 5.96 Composite Composite Z, L 635 Composite Ω, T ~230

A380 Width 7.14

Height 8.47 Aluminium

Glare Aluminium J 635 Aluminium Z ~210

B737 3.76 Aluminium

alloy Aluminium Z 508 Aluminium Ω 152-178

B747 6.40 Aluminium

alloy Aluminium Z, C 508 Aluminium Ω, Z 203-254

B777 6.20 Aluminium


B787 5.46 Composite Composite Z 610 Composite Ω ~214

FATA

Baseline 5.99 Composite Composite Z 533 Composite Ω 250


Based on the historical data in table 7 the FATA fuselage structural layout, for analysis using

AeroDYNAMIC™ and following data was used:-

Materials:- By weight, 50% of the latest large commercial transports are CFC namely the A350 and

B787 series, and the fuselage structure is predominantly CFC skin, stringers, and frames, with

Titanium / CFC hybrid door frames. Previous Airbus and Boeing CS-25 fuselage structures have

been metallic comprising of aluminium skins, frames, and stringers. The materials selected for the

FATA baseline fuselage was Cytec industries unidirectional tape X840 Z60 12K, and plain weave

fabric X840 Z60 PW, for skin, frames, and stringers and the provisional thickness and ply layups

are shown in Tables 8(a) through 8(c) the initial skin thickness was 2.9mm.

Geometry:- Frame cross-sections have generally been Z-section or C-section with a height of 85 -

100mm, with a thickness of 2 - 3mm, and a flange width of 25mm:

Stringers have been Z-section with a height of 30mm, with a thickness of 2mm, and a flange width

of approximately 15mm:

CFC Omega - profile stringers (current designs) have a height of 25 – 35mm, a thickness of 1.5 –

2.0mm, and a head width of 25mm and a flange width of 100 – 130mm:

Skin thicknesses dependant on location is reported to be between 1.0mm and 2.6mm thick.

Frame pitch from table 7 ranges 475.2 mm to 533.4mm with some extending this to 610mm to

635mm. A frame pitch of 533mm was selected FATA baseline fuselage although conservative this

was for direct comparison with the FATA PRSEUS Prime fuselage.

Stringer pitch ranges from 150mm to 254mm. A stringer pitch of 250mm was used for FATA. The

structural ply lay up tables for these components are given in table 9(a) through (c).

191

The design and structural layout of the FATA fuselage (continued).


Structural Ply No Only. Material Nominal ply thickness


1 Fabric 0.25 0º/90º

2 UD 0.15 0º

3 UD 0.15 45º

4 UD 0.15 90º

5 UD 0.15 135º

6 UD 0.15 0º

7 UD 0.15 45º

8 UD 0.15 90º

9 UD 0.15 135º

10 UD 0.15 135º

11 UD 0.15 90º

12 UD 0.15 45º

13 UD 0.15 0º

14 UD 0.15 135º

15 UD 0.15 90º

16 UD 0.15 45º

17 UD 0.15 0º

18 Fabric 0.25 0º/90º

192

Table 9(a):- Initial proposed FATA Fuselage 2.9mm thick skin ply stacking sequence.


193

Table 9(b):- Proposed FATA Fuselage 2.6mm thick Ω-stringer ply stacking sequence.

Structural Ply No Material Nominal ply thickness



2 UD 0.15 0º

3 UD 0.15 45º

4 UD 0.15 135º

5 UD 0.15 90º

6 UD 0.15 45º

7 UD 0.15 135º

8 UD 0.15 0º

9 UD 0.15 0º

10 UD 0.15 135º

11 UD 0.15 45º

12 UD 0.15 90º

13 UD 0.15 135º

14 UD 0.15 45º

15 UD 0.15 0º

16 Fabric 0.25 0º/90º


194

Table 9(c):- Proposed FATA Fuselage 3mm thick Z-frame ply stacking sequence.

Structural Ply No Material Nominal ply thickness


1 Fabric 0.25 ±45º


3 Fabric 0.25 ±45º


5 Fabric 0.25 ±45º



8 Fabric 0.25 ±45º


10 Fabric 0.25 ±45º

11 Fabric 0.25 0º/90º

12 Fabric 0.25 ±45º


FATA Baseline Vertical tail design.

The vertical tail presents a set of design issues which are different from those of the wing, or

horizontal tail and theses are be itemised below:-

a) It is not unusual for the vertical tail of a large transport to be integrally attached to (but still

removable from) the rear fuselage, the leading and trailing edge spars of the vertical tail being

attached to dedicated fuselage frames. A root integration plate is built into the vertical tail to

coincide with the upper surface of the fuselage and is used to transmit the vertical tail root skin

shear loads directly into the fuselage skin, this is the case with the Boeing 787 and 777 CFC

vertical tails which use a tension fitting plate to interface with the fuselage with attachment to

this plate at the VT torsion box leading and trailing edge spars. Vertical tail span-wise bending

results in a fuselage torsion. In some cases it is logical to incline the rear spar bulkhead to

continue the line of the rear spar of the vertical tail torsion box, as this is usually at the end of

the fuselage well aft of the rear pressure bulkhead, although no current airliner produced by

either Airbus or Boeing has adopted this layout. All of the current large Airbus and Boeing

passenger aircraft, based on published data from literature surveys and examination of aircraft

cutaways attaching the rear spar to perpendicular frames. The front spar and any intermediate

attachments to frames are also made to perpendicular frame stations within the aft fuselage

Section 19, with the transition being made at the Vertical Tail root rib or integration plate in the

case of the B-777 , and B-787, shown in figure 69(a)ii The structural layout is generally the

same format as the wing with front and rear spars and ribs forming the vertical tail torsion box,

with additional rudder hinge ribs and auxiliary front spar to support de-icing equipment and

other systems in the vertical tail leading edge fairing.

195

Section 9:- The design and structural layout of FATA empennage.


196

Figure 69(a):- FATA Vertical tail showing fuselage attachment design philosophy.

Figure 69(a)i:- Airbus A350 XWB Vertical tail to

fuselage attachment philosophy i.e. leading edge and

trailing edge star attachment lugs (Flight International

and Airbus gallery). Rudder is CFC and Nomex

honeycomb skins with aluminium ribs. Figure 69(a)ii:- Boeing 777 and 787 Vertical tail to fuselage

attachment philosophy i.e. Tension fixtures and integration

plate (Flight International and Boeing gallery). Rubber is CFC

and Nomex honeycomb skins with aluminium ribs.


b) Alternatively the vertical tail is designed to be readily detached as in the case of fighter aircraft

and modern large transports, in this case attachment is through a system of lugs attached to

the leading edge and trailing edge spars as shown for the Airbus A350 in figure 69(a)i. The

vertical attachment lugs are arranged in both lateral and fore and aft directions so that in

addition to vertical loads they react side and drag loads. The normal layout being that the lugs

attached to the leading edge spar arranged laterally and react the vertical and drag loads, and

the lugs attached to the trailing edge spar are arranged in the fore and aft direction and react

the vertical and side loads. This lug attachment philosophy was selected for the FATA vertical

tail which is attached at the leading edge and trailing edge spars with lateral and fore and aft

lugs to perpendicular fuselage frames.

c) The rudder attachment to the vertical tail is invariably supported by a number of discrete hinges

and number and location of these hinges depends on the length and weight of the rudder, and

the other major points to consider in rudder attachment design are as follows:-

i. The bending distortion of the control surface relative to the fixed vertical tail must be limited

so that the nose of the control does not foul the fixed shroud:

ii. The control hinge loads and the resulting shear forces and bending moments should be

equalized as far as possible:

iii. Structural failure of a single hinge should be tolerated unless each hinge is of fail-safe

design and can tolerate cracking in one load path.

197

The design and structural layout of FATA empennage (continued).


198

Figure 69(b):- FATA Vertical tail showing internal structural layout key datum positions.

Fwd Attachment Frame interface:- two

lateral lugs the Leading Edge spar root.

Aft Fuselage

barrel section.

Mid Attachment Frame interface:- two Fore

and Aft lugs on the Mid spar root.

Aft Attachment Frame interface:- two Fore

and Aft lugs on the Trailing Edge spar root.

Triplex rudder

EHA actuators.

CFC Stringers.

CFC Leading Edge Spar.

CFC Vertical Tail Leading

Edge box with Al/Li Ribs.

CFC Mid (Intermediate) Spar.

CFC Trailing Edge Spar.

Al/Li alloy all Ribs.

Rudder CFC Honeycomb

skins with Al/Li ribs .

Figure 69(b)i:- VT / Frame interface.

FWD

Figure 69(b)ii:- VT internal layout

Port skin and stringers removed.

UP

CFC Skins. Vertical Tail Pf Area = 35.36m²

Rudder Pf Area = 15.00m²


These points suggest the use of a relatively large number of discrete hinges but there are

issues associated with this solution. There is the obvious issues of high assembly complexity

and maintenance, and hinge alignment difficulties. Additionally the loads likely to be induced in

the rudder by the distortion under load of the vertical tail to which it is attached may be

significant. These problems do not arise if only two hinge points are used as any span-wise

distortion or misalignment can be accommodated by designing one of the hinges so that it can

rotate about a vertical axis a so called „floating‟ hinge. When more than two hinges are used

this „floating‟ hinge concept cannot fully overcome the problems. However it is possible to

design the control surface so that it is flexible in bending and indeed the more hinges there are

the easier this is to accomplish. One hinge must always be capable of reacting side loads in the

plane of the control surface, the hinges being supported near to the aft extremities of the

vertical tail ribs. For the initial internal structural layout concept the FATA Baseline Vertical Tail

the rudder attachment layout of the Airbus A330 was used as a starting point for analysis using

AeroDYNAMIC™ of loads and detailed structural analysis.

FATA Baseline Horizontal tail design.

When the horizontal tail is constructed as a single component across the centreline of the aircraft

the basic structural requirements are the very similar to the wing see above. Therefore to address

this the concept structure was designed as two spar multi rib torsion box, with two actuator

positions for the elevator on the Port and Stbd Horizontal Tail Planes figure 70(b), this is similar to

the Airbus A350 WXB, and A330. The Boeing 787 takes a different approach with the horizontal tail

torsion box being multi spar construction. The all moving FATA horizontal tail is attached to the

fuselage by the fwd Screw Jack actuator fitting and aft pivot lugs figure 70(a).

199

The design and structural layout of FATA empennage (continued).


200

Figure 70(a):- FATA Horizontal tail showing internal and interface design philosophy.

Figure 70(a):- Airbus A350 XWB Horizontal tail to fuselage attachment philosophy i.e. stiffened centre box is attached to

the screw jack actuator at the front, and at trailing edge is attachment with two pivot outer lugs. The same basic layout is

used by Boeing (Flight International and Airbus gallery). Elevators are constructed of CFC skinned Nomex honeycomb

skin panels with aluminium ribs, and mesh instead of electrical bonding straps.

Port Lug Stbd Lug

HT Composite leading edge spars.


201

Figure 70(b):- FATA Horizontal tail showing internal structure key datum positions.

CFC Leading Edge Spar.

CFC Trailing Edge Spar.

View on arrow „B‟ of Jack Screw

actuator attachment lug.

A

CFC Torsion Box Skins.

View on arrow „A‟ of Port and

Stbd HT Pivot lugs.

B

CFC Horizontal Tail Leading

Edge box with Al/Li Ribs.

CFC Horizontal Tail

Leading Edge box

with Al/Li Ribs.

Al/Li alloy all Ribs.

Out-BD Elevator

EHA actuator.

In-BD Elevator

EHA actuator.

Elevators CFC Honeycomb

skins with Al/Li ribs.

CFC Stringers.

Four-part Hard Back.

Stiffened CFC Centre Box.

FWD

UP

PORT

Port or Stbd Horizontal Tail Pf Area = 23.49m²

Port or Stbd Elevator Pf Area = 10.18m²


Type-2 assembly:- Current metallic aircraft assembly is a Type-2 process in that it requires precise

fixtures and jigs to support the metallic components in build, and the majority of these jigs and

fixtures are specific to the airframe model and configuration, therefore the manufacture has to rely

on a relatively long production run to for the tooling to be economic. The reason for this, is that it

would be prohibitively expensive to attempt to make large and temperature sensitive – sensitive

structures to tolerances as small as 3x10 on a relative basis, and such fine tolerances are

required to reduce locked in stresses. The resultant structures are relatively stiff compared to the

component parts so small deformations can usually be eliminated by bending the structure,

however this induces local stresses which detract from the flight load carrying capabilities of the

assembly and therefore should be avoided where ever possible.

Boeing seeks to avoid such stresses by building structural components with multiple slip joints by

ensuring that there is empty space at maximum material condition in many joint conditions. These

spaces are filled by peel-apart metallic shims until the gap is small enough to be safely pulled

together with fasteners.

Airbus seeks to avoid such stresses by making their structural components through high precision

5 axis NC machining (see career presentation), which produces a very accurate part that only

requires occasional application of liquid shim (this methodology is also used in UK military aircraft I

have worked on).

Type-1 assembly:- Bridges and skyscrapers are classed as Type-1 assembles as their materials

are thick section and rugged, and they are assembled from hole pattern features, although hole

filling requirements are not as critical as for aircraft and the materials are less temperature

sensitive, and there is not the same need to conserve weight. 202

-4

Section 10:- Airframe structural assembly philosophies.


Therefore bridges and sky scrapers constitute Type-1 assemblies in that they are assembled by

joining of their component parts rapidly at their features without dedicated single use tooling.

Given the costs and the number of non-added value operations involved in airframe assembly the

direction of assembly research in this design trade study will be focused on reducing the amount of

manual labour and the specificity of the fixtures required to assemble the wing torsion box and wing

carry through box. The weight of the components in theses structures will still need require support

to avoid collapse in assembly, so fixture-like structures will still be necessary but they might not

need to be as accurate or as specific as they are now, lessons learnt from the Mantis UAS field

assembly will be used to modularize these structures into a kit form facilitating autonomous

assembly, of major build units. Three major research activates will be perused in common with

other current research these are:- (1) move towards new composite manufacturing and assembly

methods using preforms and RIM and sacrificial plies: (2) the broad attempt to move aircraft

structures from Type-2 to Type-1 assembly: (3) autonomous assembly.

Activities (1) and (3) are covered in some detail in the latter sections of this research status update,

so here I will briefly cover activity (2).

Develop aircraft structures for Type-1 assembly:- If aircraft parts can be made to net size and

shape with assembly fixtures incorporated in them then they could be tacked together to achieve

the desired final assembly dimensions and relationships just by joining these features (as was

achieved with the Terrasoar light UAS wing / boom assembly). Then they could be given their final

assembly fasteners as before. The savings would arise from the elimination accurate and specific

fixtures.

203

Assembly of baseline aircraft wing torsion box structural members (cont.).


While progress has been made in this area it is not felt possible to pre-drill every fastener hole as is

done in building and bridge construction. The holes in aircraft construction must be essentially

exactly opposite each other or the fastener cannot fill the hole causing fretting which leads to hole

elongation, corrosion and fatigue, because the fastener will wobble when exposed to oscillating

shear loads normal to its axis rapidly enlarging the hole until to can carry no load at all. The only

current method to achieve many holes that are exactly opposite each other is to match drill on

assembly when the parts are clamped together in their correct relative position. The focus of

attention of current research in this area is therefore on tack fastening to create mates that pass the

dimensional location constraints between the parts, and achieving this would create Type-1 aircraft

assembly. The work I intend to undertake in this area is to identify which critical fastener locations

could become tack fasteners and to look at additional features which could be designed in for a

Lego type build solution.

Figure 71 on the next two slides illustrates proposed join concepts for the rib to leading and trailing

edge spars here there are two possible innovations:- One is the integral cleat shown in figure 71(c)

which would remove the need for additional spar / rib cleating reducing parts count and assembly

time, although the possibility of a resin rich area at the bend must be considered, I have designed

an actual rib based on my key datum model and the current loads drop figures 43 to 47. The other

innovation is the composite post which would be produced from back to back RTM moldings I am in

the process of conducting drape trials and calculations for the flow required to realistically mold

such articles, woven cloth would be used in preference to UD ply to reduce the risk of fibre-wash,

this would then be co-bonded into the Leading edge spar.

204

Assembly of baseline aircraft wing torsion box structural members (cont.).


205

Figure 71:- Metallic rib build joints selected for assembly of the baseline wing.

Ti Rib post.

Fig 71(a):- Dry Bay Al rib Bathtub nested into CFC Trailing Edge Spar joints.

* Based on 2 countersunk x 1.25” diameter fasteners + 0.06” clearance.

** Based on diameter of Eddie bolt installation tool and footprint of clickbond

nutplate.

Top wing cover skin.

Bottom wing cover skin.

Rear spar.

Bonded anchor nuts.

* 2.5 d

**

Wing rib to spar bathtub.

Fig 71(b):- Rib to Leading Edge Spar post joints.

*Based on 3 x fasteners.

This joint employs a rib attachment post mounted in the spar

for the rib tab to land on which could be bonded or bolted in

place although shown here as a Ti fitting a CFC co-bonded

post is to be studied.

Front spar.

Wing rib to spar tab.

*

2d+1.5mm

4d


206

Figure 71:- Composite rib build joints selected for assembly of the baseline wing.

Top wing cover skin.

Fig 71(c) CFC Rib to CFC Trailing Edge integral cleat joints.

Integral cleat removes the need for cleated joint reducing parts count

and easing assembly this is a concept for illustrative purposes an

actual rib design will be included in the next update.

Bottom wing cover skin.

Front spar.

Wing rib to spar tab.

CFC Rib post. 3-d 2.5-d

6-d

Fig 71(d) CFC Rib to CFC Leading Edge Spar post joints.

Rib tab attachment bolted to co-bonded RTM integral spar post

joints composed of two back to back filled C sections.

Wing rib to spar integral cleat.

Rear spar.

Bolted through Rear spar web.

4d

3d edge of cleat


Before proceeding with the conventional baseline design, it is important to consider the

advantages and disadvantages of both bolted and bonded construction methods and the impact

of corrosion on composite assemblies.

The advantages of bolted assembly are:-

1)Reduced surface preparation:

2)Ability to disassemble the structure for repair:

3)Ease of inspection.

The disadvantages of bolted assembly are:-

1)High stress concentrations:

2)Weight penalties incurred by ply build ups, and fasteners:

3)Cost and time in producing the bolt holes, and inspection for delamination's:

4)Assembly time.

Corresponding issues for bonded assembly are set out below.

The advantages of bonded assembly are:-

1)Low stress concentrations:

2)Small weight penalty:

3)Aerodynamically smooth.

207

Composite structural assembly joint design and corrosion.


Composite structural assembly joint design and corrosion (continued).

The disadvantages of bonded assembly are:-

1) Disassembly, in most cases some part of a bonded structural assembly will need to be bolted

instead of bonded to permit access for repair and inspection. An example is the Typhoon

wing structure where the bottom skin is co-bonded to the structural spars, and top skin is

bolted to the same spars, permitting access from one side:

2) Surface preparation, and bond line inspection for porosity even in co-bonded joints using C-

scan ultrasonic inspection, resulting increased costs and time:

3) Need to design for bolted repair access:

4) Environmental degradation due to water absorption leading to degradation in hot / wet

condition, solvent attack:

5) Need for increased qualification testing effort to establish design allowables.

In the case of the baseline wing configuration both bolted and co bonded construction will be

selected primarily because of the requirement to quickly, inspect, repair, or replace damaged

structural components within a first line servicing environment. In the assembly models bolt

datum positions are shown as points and vectors, as was the practice within BAE Systems MA&I,

and for this level of study only selected detail fastener models will be created.

208


Co-Curing:- This is generally considered to be the primary joining method for joining

composite components the joint is achieved by the fusion of the resin system where two (or

more) uncured parts are joined together during an autoclave cure cycle. This method minimises

the risk of bondline contamination generally attributed to post curing operations and poor

surface preparation. But can require complex internal conformal tooling for component support.

Co-Bonding:- The joint is achieved by curing an adhesive layer added between a co-cured

laminate and one or more un-cured details. This also requires conformal tooling as shown in

figure 72, and as with co-curing the bond is formed during the autoclave cycle, this method was

used on Eurofighter Typhoon wing spars which were co-bonded to the lower wing cover skins,

and proposed for the F-35B VT lower skin stringers in SWAT trade studies. Care must taken to

ensure the cleanliness of the pre-cured laminate during assembly prior to the bonding process.

Secondary Bonding:- This process involves the joining of two or more pre-cured detail

parts to form an assembly. The process is dependent upon the cleaning of the mating faces

(which will have undergone NDT inspection and machining operations). The variability of a

secondary bonded joint is further compounded where „two part mix paste adhesives‟ are

employed. Generally speaking, this is not a recommended process for use primary structural

applications.

Design considerations for adhesive bonded joints.

209


UN-CURED „Z‟ & „C‟

SPAR ELEMENTS

„FILM‟ ADHESIVE

(BSL.322)

„CLEAVAGE‟ FILLED WITH

UN-CURED CFC WEDGE

RELEASE AGENT

PRE-CURED

CFC SKINS

UN-CURED „Z‟ & „C‟

SPAR ELEMENTS

CONFORMABLE TOOLING SHOWN AS:-

Figure 72:- Co-Bonded composite spar manufacture.

210


Composite bolted joint design rules:-

1) Design for bolt bearing mode of failure:

2) Counter sink (CSK) depth should not exceed 2/3 of the laminate thinness if required fill

laminate artificially with syntactic core (if design rules permit e.g. not permitted for USN, or

USMC):

3) Minimum bolt pitch is 4D for sealed structures such as fuel tanks, and 6D for non sealed

structures (where D is the bolt diameter) figure 73:

4) Use only Titanium alloy or stainless steel fasteners to minimise corrosion risk:

5) Use a single row of fasteners for non sealed structures and a double row for sealed

structures such as fuel tanks see figure 74 next slide:

6) Minimum fastener edge distances are:-

3-D in the direction of the principal load path see figure 73:

2.5-D transverse to the principal load path see figure 73:

211

Composite structural assembly joint design and corrosion (continued).

Figure 73:- Fastener edge distances.

2.5xD 3.0xD

4.0 x D


212

Figure 74:- Corrosion / leek prevention methods for carbon fibre structures.

Prevent moisture ingress:

Prevent electrical contact between CFC and Al alloy:

Anodise, Prime and Paint all Al alloy parts:

Seal in accordance with Project specifications.

Corrosion / leak prevention (fuel tank) example.

Titanium Alloy Fasteners

(NOT cadmium plated)

Dipped prior to assembly.

Al alloy Rib

CFC Spar

Polysulphide

Sealant (PRC)

- FUEL TANK -

- CFC SKIN -


213

FASTENER

MATERIAL / COATING COMPATABILITY

• Monel. Marginally acceptable.

• Alloy Steel.

• Silver Plating.

• Nickel Plating.

• Chromium Plating.

Excellent compatibility and are

recommended for use in CFC structures

• Cadmium Plating.

• Zinc Plating.

• Aluminium Coating.

Not compatible, and will deteriorate rapidly

when in intimate contact with CFC.

• Titanium Alloy.

• Corrosion Resistant Steel.

Excellent compatibility and are

recommended for use in CFC structures

• Al. Alloys.

• Magnesium Alloys.

Not compatible

Not compatible

Table 10:- Galvanic compatibility of fastener materials and coatings.


214

The use of carbon composites in conjunction with metallic materials is a critical design

factor :-

Improper interfacing can cause serious corrosion :

Problem for metals e.g. Fasteners see table 10 above:

This corrosion problem is due to the difference in electrical potential between some of the

materials widely employed in the aircraft industry, and carbon:

When in contact with carbon and in the presence of moisture (electrolyte), anodic materials

will corrode sacrificially (galvanic corrosion).

Corrosion prevention methods:-

1) Prevent moisture ingress:

2) Prevent electrical contact carbon / metal:

3) Anodise aluminium parts:

4) Seal in accordance with project specifications:

5) Protective ply of inert cloth (glass) between contact surfaces extending 1” beyond edge on

metal part, and protective sealant (Polysulphide) „Interfay‟ see figure 75 on next slide.

Corrosion due to the galvanic compatibility of materials and coatings.


215

Figure 75:- Corrosion prevention methods for carbon fibre structures.

EPOXIDE PRIMER (15 to 25 Microns THICK)*

ANODIC TREATMENT*

Pu. VARNISH or EPOXIDE PAINT FINISH (ONE COAT)*

Al ALLOY COMPONENT

POLYSULPHIDE „INTERFAY‟ SELANT

EPOXIDE PRIMER**

GRP (As required as a „Drill

Breakout‟ material.)**

CARBON FIBRE COMPOSITE

* = Applied over the entire Al component.

** = Applied over the entire CFC

component – or a minimum of 5mm

beyond the contact area.


216

1) Stress concentrations exert a dominant influence on the magnitude of the allowable design

tensile stresses. Generally, only 20-50% of the basic laminate ultimate tensile strength is

developed in a mechanical joint:

2) Mechanically fastened joints should be designed so that the critical failure mode is in bearing,

rather than shear out or tension, so that catastrophic failure is prevented. To achieve this an

edge distance to fastener diameter ratio (e/D), and a side distance to fastener diameter ratio

(s/D) relatively greater than those for metallic materials is required, (see figure 73 above). At

relatively low e/D and s/D ratios, failure of the joint occurs in shear out at the ends, or in tension

at the net section. Considerable concentration of stress develops at the hole, and the average

stresses at the net section at failure are but a fraction of the basic tensile strength of the

laminate:

3) Multiple rows of fasteners are recommended for unsymmetrical joints, such as shear lap joints,

to minimize bending induced by eccentric loading:

4) Local reinforcement of unsymmetrical joints by arbitrarily increasing laminate thickness is to be

avoided because the resulting eccentricity can give rise to greater bending stress which

negates the increase in material thickness:

5) Since stress concentrations and eccentricity effects cannot be calculated with a consistent

degree of accuracy, it is advisable to verify all critical joint designs by testing of a representative

sample joint.

Composite structural mechanically fastened joint design guidelines.


217

6) If a laminate is dominated by 0° fibers with few 90° fibers it is most likely to fail by shear out,

unlike metals, in which shear out resistance can be increased by placing the hole further from

the edge, laminates are weakened by fastener holes regardless of distance from the edge.

Reinforcing plies at 90° to the load direction helps prevent both shear out and cleavage failures:

Use larger fastener edge distances than with aluminum design, e.g. e/D >3: Use a minimum of

40% of ± 45° plies (for their influence on bearing stress at failure: Use a minimum of 10% of 90°

plies.

7) Net tension failure is influenced by the tensile strength of the fibers at fastened joints, which is

maximized when the fastener spacing is approximately four times the fastener diameter (see

figure 68 above). Smaller spacing's result in the cutting of too many fibers, while larger

spacing‟s result in bearing failures in which the material is compressed by excessive pressure

caused by a small bearing area: Use minimum fastener spacing as shown in figure 73 with 5D

spacing between parallel rows of fasteners: Pad up to reduce net section stresses.

8) To avoid fastener pull-through from progressive crushing / bearing failure:- Design joint as

critical in bearing: Use pad up: Use a minimum of 40% of ± 45° plies: Use washer under collar

or wide bearing head fasteners: Use tension protruding heads when possible.

9) To avoid shear failure:- Use large diameter fasteners: Use higher shear strength fasteners:

Never use a design in which failure will occur in shear.

Composite structural mechanically fastened joint design guidelines (continued).


10) Use two row joints when possible, as the low ductility of advanced composite material confines

most of the load transfer to the outer rows of fasteners.

11) The choice of optimum layup pattern for maximized fastener strength is simplified by the

experimentally established fact that quasi – isotropic patterns (0°/±45°/90°), or (0°/45°/90°/-45°)

are close to optimum, in practice this reduces experimental costs and simplifies analysis and

design of most fastened joints.

12) The effects of eccentricities on joints:- if eccentricities exist in a joint, the moment produced

must be resisted by the adjacent structures: eccentrically loaded fasteners patterns may

produce excessive stresses if eccentricity is not considered.

13) Mixed fastener types should not be used, i.e. it is not allowed to use both permanent fasteners

and removable fasteners in combination on the same joint, this is due to the better fit of the

permanent fasteners, which would result in the removable fasteners not picking up their

proportionate share of the load until the permanent fasteners have deflected enough to take up

clearance of the removable fasteners in their holes.

14) Do not use a long string of fasteners in a splice joint, because the end fasteners will load up first

and hence yield early. Therefore use three or four fasteners per side as the upper limit unless a

carefully tapered, thoroughly analyzed splice is used (wherever possible use a double shear

splice).

218



15) Use tension head fasteners for all applications (because potentially high bearing stress under

the fastener head cause failure). Shear head fasteners may be used in special applications.

16) Where local buildup is required for fastener bearing strength, total layup should be at least 40%

± 45° plies.

17) Installation of fasteners wet with corrosion inhibitor may be required in some cases.

18) Use of large diameter fasteners in thicker composite assemblies (for example to transfer critical

joint loads, fastener diameters should be about equal to the laminate thickness) to avoid peak

bearing stress due to fastener bending. Fastener bending is much more significant for

composites than for metals, because composite are thicker for a given load, and more sensitive

to non-uniform bearing stresses due to brittle failure modes.

19) N.B. the best fastened joints can barely exceed half the strength of unnotched laminate.

20) Peak hoop tension stress around fastener holes is roughly equal to average bearing stress.

21) Fastener bearing strength is sensitive to through - the – thickness clamping force of laminates it

is highest for a 30% / 60% /10% (0º/± 45°/90°) ply lay up stack, and much lower for

50%/40%/10% (0º/± 45°/90°) ply lay up stack.

22) Production tolerance build ups:- proper tolerances should be determined with manufacturing to

minimize the need for shimming: shim allowance should be called out on engineering drawings:

N.B. since production tolerances can easily be exceeded in the thickness tolerance, fastener

grip length can be adversely affected.

219



Shims are used in airframe production to control structural assembly and to maintain aerodynamic

contour and / or structural alignment. With composite joints the allowable unshimmed gaps are only

¼ as large as those for an similar metallic structural joints. Therefore, the assembly of composites

generally require more extensive use of shims than comparable metal components.

Engineering can reduce both cost and waste by controlling shim usage through design and

specifications. Design can control where to shim: what the shim taper and thickness should be:

what gap to allow: and whether the gap should be shimmed or pulled up with fasteners.

Shim materials currently available are:-

1)Solid shims:- titanium: stainless steel: precured composite laminates: etc.

2)Laminated (or peelable) shims with a laminate thickness of about 0.003” (0.0762mm) ±0.0003”

(0.00762mm)

Laminated titanium shims:

Laminated stainless steel shims:

Laminated Kapton shims.

3)Moldable shim, which is a cast – in – place plastic designed for use in filling mismatches between

metal or composite parts. It can be used at any location to produce custom mating molded surfaces

examples are given in the reference works given in the end of this report.

220

Composite structural mechanically fastened joint design shim guidelines.


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:- (a) U/D, no overlaps, butt joint or up to 2mm gap: (b) Woven

cloth, no gaps or butt joints, 15mm overlap:

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):

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:

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

Internal corner radii of channels:- (a) „t‟ < 2.5mm, radius = 2t or 3.0mm whichever is greater: (b)

„t‟ 2.5mm, radius = 5.0mm

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

area, need for core stabilisation and reduced cure pressures.

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.

Composite ply layup guidelines applied to FATA wing based BAE Systems MA&I practice.

221


In the proceeding slides I have referred to the use of sacrificial plies to ensure build tolerances are

met in composite skin and spar joint assembly. In this section I will give a brief outline of them and

their design requirements which will be applied in the design of composite structure in this project

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

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

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

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

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

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

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

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

specific desired thickness without infringing into the structural plies.

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

from the sacrificial machining as shown in figure 76(a). This buffer or witness ply is designed to

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

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

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

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

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

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

at these ply-drop ramps.

222

Composite sacrificial plies for assembly tolerance control.


223

Figure 76:- Sacrificial ply design to meet assembly requirements.

Skin OML

Figure 76(a):- CFC sacrificial incorporation in ply lay up to meet assembly tolerance.

Substrate finish coating

Machined Shim

Faying Sealant

Machined Sacrificial Plies

Adhesive /Fabric Buffer (Witness) Layer

Laminate finish coating

Structural Plies

Fibermat

OML

Figure 76(b):- CFC laminate thickness constituents to meet assembly tolerance.

Machined Shim

Faying Sealant

Machined Sacrificial Plies

Laminate finish coating

Adhesive /Fabric Buffer (Witness) Layer Structural Plies OML Plies

Substructure OML

Machined Skin IML

Nominal

Laminate

Thickness


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

thickness (see figure 76(b)): -

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

thickness or carrier thickness if any.

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

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

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

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

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

protection layer.

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

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

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

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

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

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

is used to calculate the laminate IML and the substructure OML surface (figure 77). Both the

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

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

IML.

224

Composite sacrificial plies for assembly tolerance control.


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

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

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

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

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

barrier (as applicable) + finish primer / paint.

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

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

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

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

N.B.:- If the plies are placed by hand with a ply projector, location, ply projector and ply pack

trim tolerances must be accounted for.

Also note the thickness of sacrificial plies on a constant laminate section will be less than the

thickness at the top of a ramp which has to account for ply drop location accuracies.

225

Composite sacrificial plies for assembly tolerance control (Workbook 1).


Figure 77:- CFC laminate thickness constituents in a Taper Region.

226

Fibermat OML Layer OML Surface

Machined IML

Structural Plies

Buffer

(Witness)

Layer

Sacrificial Ply Thickness

Sacrificial Ply Thickness

Top of Ramp Thickness

Bottom of Ramp Thickness

Corrosion Protection

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

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

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

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


Faced with ballooning order backlogs for airframes, and an ever increasing demand for a shorter

development and build cycle times for future aircraft as well as current models, aerospace

manufactures and automation system suppliers are exploring new ways to automate a broader

range of aircraft manufacturing processes beyond drilling and filling. The objective of this section is

to incorporate the work of the Kinematics and Aircraft Assembly Robotics studies by applying the

modeling techniques developed in the Kinematics studies to the design of assembly robots,

assembly fixtures, and the human builder to study the automated assembly of airframe wings in

support of my private Future Advanced Technology Aircraft study. These robot designs will be used

to asses: - assembly clearances, space envelopes with human interactions, fixtures, and individual

part features required for structural assembly and systems installation using the FATA outboard

wing section as an example structure.

Within the limits of Catia V5.R20 Kinematic modeling (which include the inability to combine the

simulation of sub-mechanisms in to a single larger mechanism, which means that the entire robot

mechanism will have to be modeled as a single large mechanism), to model full sized ABB

IRB4400, and ABB IRB6650S robots figure 78(a), from datasheet and surface model dimensions,

simulating their functional space envelopes in combination with the human builder in assembly

activities required for the baseline and developed wings of the FATA aircraft project. Path

simulations and tracker simulations will be undertaken to ensure that line of sight is preserved

between the robots optical sensors and a simulated Leica laser tracker (modeled from catalogue

data, see section 3 of this presentation ). Studies will also address the behaviour of the robots

under load conditions to determine deflections using GSA, leading to placement errors. This work is

focused on establishing a Type-1 feature based assembly methodology for the FATA wing project.

227

Section 11:- Robotic assembly in the development of the Baseline wing.


228

Fig 78(a):- ABB IRB 4400/60 and ABB IRB6650S articulated arm Robots.

ABB IRB6650S_90 Robot.

ABB IRB4400/60 Robot. Male Human Builder

(Jerry).

Female Human

Builder (Betty).


229

(B) Arm

axis

+96º /-70º

(A) Arm

axis

65º/-60º

(C) ± 165º

axis Rotation

(B) Arm

axis

+96º /-70º

(D) ± 200º

axis Wrist

(E) ± 120º

axis Bend

(P) ± 400º

axis Turn

Fig 78(b):- Axis movements / working range of ABB IRB 4400/60 articulated arm Robot.


230

Table 10:- ABB Data sheets for the IRB 4400/60 robot from reference 18.


Robots by functions, fall into four basic categories:-

1)Pick and Place (PNP) this is the simplest of robots and its function is to pick up a part and move it

to another location. Typical applications include machine loading and unloading and general

materials handling tasks:

2)Point to Point (PTP) some which are similar to PNP robots, in that they move material from one

location to another, hence point to point, however it can move to literally hundreds of points in

sequence. At each point sophisticated PTP robots can stop and perform an action such as spot

welding, gluing, drilling, deburring, or a similar task:

3)Continuous path (CP) robot also moves from point to point but the path it takes is critical. This is

because it performs its task while it is moving. Paint spraying, seam welding, cutting and inspection

are typical applications of this type:

4)Robotic assembly (RA) articulated arm robot shown figure 78(b) is the most sophisticated robot

type of all and combines the path control of CP robots with the precision of machine tools. RA often

work faster than PNP and perform smaller, smoother and more intricate motions than CP robots.

A full description and definition of the proposed automated assembly study will be released in as

my full Kinematics and Aircraft Assembly Robotics Study by the middle of 2017. However currently I

have produced overview of industry development projects and a SCARA (selective compliance

assembly robot arm) kinematic model Human manikin and started interface studies (which can be

demonstrated at interview), which are covered in the accompanying presentation Kinematics and

Airframe Assembly Robotics Study (posted on my LinkedIn profile).

231

Robotic assembly in the development of the Baseline wing (continued).


The landing gear loads and reactions are the largest local on the aircraft structure, and therefore

transmitting such large local loads into the semi-monocoque structure of the wing box requires

extensive local reinforcement. Since the landing gear loads are large, there can be severe weight

penalties in the use of indeterminate structural load paths. An indeterminate structure is one in

which a given load may be reacted by more than one load path with the distribution being subject to

the relative total stiffness of these paths. In practice the manner in which the members share the

load can be determined but only when the design is finalized, and often overlapping assumptions

are made of the load paths which results in an over deigned heavy structure.

Often the gear loads can be spread out so as to keep the local reinforcement to a minimum, in the

case of the FATA as with the A350 family of aircraft the use of carbon fibre reinforced plastic

(CFRP) required a reduced point loading to reduce the amount of structural reinforcement required

in the aft spar. So as shown in figures 79(a), 79(b) for FATA a double side-stay landing gear was

developed similar to the Messier-Dowty A350 configuration where the aft side-stay is attached to

the auxiliary spar (or gear kick beam), thus reducing the reinforcement weight for the aft CFRP

spar. The support structure in the wing is designed to higher loads than the gear itself to ensure

that in the event of impact the gear will break off cleanly with the wing and not precipitate a fuel

tank rupture. The installation of the landing gear aft of the wing carry through box is shown in figure

79(c) and the requirement is for a 4.1m fuselage bay. For this study the landing gear loads are

developed using the methods in references 4 and 7.

232

Section 12:- Integration of baseline and developed aircraft nose and main landing gear.


233

Figure 79(a)/(b):- FATA Nose and Main landing gear used in for the design study.

Aft Stay

Figure 79(a) Sized Nose landing gear general

arrangement for integration in the FATA Design

Study.

Figure 79(b) Sized Port Main landing

gear general arrangement for integration

in the FATA Design Study.

FWD

FWD

Fwd Stay

Sliding piston

Main Fitting

Upper Torque Link

Lower Torque Link

Bogie Unit

Attachment fore

and aft pintle

Upper Drag Stay

Lower Drag Stay

Main Fitting

Steering

Assembly

Upper Torque Link

Lower Torque Link

Retraction Actuator Fwd Stay

Lock

Note:- Although Landing Gear Wheels and Struts have been sized these are not detail designs.

Lateral pintle


234

Figure 79(c):- FATA Main Landing Gear Installation checks in 4.43m bay.

4 wheel bogie MLG Installation check

(view from below).

FWD

OUT BD

UP UP OUT BD

FWD

4 wheel bogie MLG installation Section 15 Keel

panel kinematics clash check (view from below).


Section 13:- Integration of baseline and future concept engines.

The engine installation used on the baseline and developed aircraft in this study is in the standard

form of an under-wing nacelle pod, which for current designs has least effect on the aerodynamic

characteristics of the wing. For jet engines the wing nacelle pod mounting is the preferred option,

freeing more space in the wing to be used for integral fuel tanks, and imposing a torsional moment

on the wing which is desirable to offset wing wash-out at high angles of attack, and under

accelerating flight conditions. The thrust and inertia loading on the engine and the air loading on its

attached structure are carried back to the aircraft structure via the engine mounts. The engine and

support structure will react loads in any direction as Px (thrust), Py (side loads), Pz (vertical loads)

and the three corresponding moments Mx, My, and Mz as shown in figure 80(a). The nacelle,

nacelle strut, and engine mounts are designed to the ultimate load factors given in reference 7 for

this preliminary design study.

The pylon options for mounting the under-wing nacelle pod are shown in figures 80(b),(c),(d), where

the engines are supported by box beams of aluminium, titanium, or steel construction. The pylon is

attached to the wing front spar and lower skin panel with pylon loads distributed to the wing

structure in such a manner that wing box secondary deformation is minimized. In figure 80(b) the

pylon bulkheads take the engine loads onto the wing box and the pylon is attached to the front spar

by the pylon upper longeron, utilizing a rear drag strut to transfer the pylon lower longeron loads to

a point between the front and rear spar requiring skin reinforcement and not favored for this study.

In figure 80(c) the pylon is a box beam design and although this design puts more weight into the

pylon it saves weight in the wing box and reduces fatigue issues, and is the basis for the Alliance

pylon used on the A380 and is favored for this study.

235


236

Figure 80:- Possible wing pylon arrangements for the baseline aircraft.

Pz

(Side)

Mx

Mz

My Px

Py

(Thrust)

(Vertical)

Figure 80(a) Engine Loads.

Figure 80(b) Drag strut pylon installation.

Figure 80(c) Box beam pylon installation.

Figure 80(d) Drag strut pylon installation with

upper support arm (redundant support).


The pylon is attached to the wing through a fitting on the front spar for vertical and side loads, to a

fitting beneath the front spar on the wing lower surface for thrust loads, and to a fitting attached to

the wing box structure on the wing lower surface at the end of the pylon for vertical and side loads.

Spherical bearings are used at the pylon-to-wing attachments to avoid over constraint to the wing

lower front spar. Side fairing panels, with attached bulb seals cover the gap between the pylon

structure and the lower skin, and the pylon structure is identical left and right and is therefore

interchangeable. However the front spar fitting is complicated. In figure 80(d) the pylon has a

complex redundant support structure as detailed in reference 7 this is shown here for completeness

of options considered, although it is an inherently structurally fail safe design due to its redundant

load paths it is heavy and complex and was not considered for this study.

Figures 81, and 82 shows the proposed engine to be used, and the study engine nacelle and pylon

dimensions and OML layout which has an impact on the pylon and wing box structural design.

Table 11 gives approximate data for the Rolls Royce Trent 1700 for the A350-1000, in comparison

with the Rolls Royce Trent 772 for the A330 to illustrate the requirements growth.

Figures 83 shows the engine configurations for long and short / medium haul aircraft, and 84 shows

the additional loading introduced by the application of turbofan engine thrust reverses. Figure 85

shows current materials, and figures 86 through 89 show possible future engine concepts

considered for the future concept airframes to be studies in the third phase of this project, and

figure 90 shows the basis for the new airframe configurations to be studied in the third phase.

237

Baseline and future concept engines used in this study.


238

Figure 81:- RR-Trent 1700 used for wing and pylon loading, 87,000lbs thrust, 3 shaft.

The Forward engine

mount takes vertical

and side loads . The Aft engine mount takes engine

thrust loads, vertical side loads,

and torque moment Mx .

The Fan 118” diameter

SPF/DB Ti or monolithic

CFC blades with kevlar or

R2 glass faces and Ti

blade edges.

Low pressure Fan stage

compressor SPF/DB Ti alloy

or monolithic CFC with Ti

leading / trailing edge blades.

Intermediate 8 stage

pressure compressor

machined solid Ti blades.

High 6 stage pressure compressor

machined solid Ti blades BLISK.

High 1 stage pressure

turbine with directionally

solidified hollow Nickel alloy

air cooled blades.

Low 5 stage

pressure turbine

with directionally

solidified hollow

Nickel alloy air

cooled blades.

Intermediate 1 stage

pressure turbine

Nickel alloy blades.


239

Table 11:- RR-Trent 1700 & 772 data used in wing and pylon loading design calculations.

TRENT 772 Data. TRENT 1700 (Approximations).

Fan diameter 97.40” (2.474m) Fan diameter 118” (2.997m)

Basic engine Length 154” (3.912m) Basic Engine Length 191.7” (4.868m)

Basic engine weight 10,550lbs (4,785kg) Basic engine weight 13,700lbs (6,214kg)

Max thrust 71,100lbs Max thrust 87,000lbs

Number of shafts 3 Number of shafts 3

Compressor stages 1LP+8IP+6HP Compressor stages 1LP+8IP+6HP

Turbine stages 1HP+1IP+4LP Turbine stages 1HP+1IP+5LP

On wing podded length 236” (6.00m) On wing podded length 330” (8.40m)

On wing max podded

diameter 105” (2.67m)

On wing max podded

diameter 126” (3.20m)


Figure 82:- FATA Engine Nacelle and Pylon from AeroDYNAMIC™ sizings.

Dimensions used to model baseline

aircraft wing pylon and engine nacelle

derived from AeroDYNAMIC™ based on

the engine size and performance data for

the RR Trent 1700 HBT (public domain).

*Engine Nacelle Ground Clearance:- FWD

C of G position 0.76m and AFT C of G

position 0.78m.

A

6.40m

2.09m

11.10m

1.14m

2.15m

A

3.94m

3.17m

1.18m

240


241

Requires high:-

Overall pressure ratio:

Turbine entry temperature:

Bypass ratio.

Range

Fuel consumption.

Long / Medium-Haul (40,000-100,000lbs thrust):

Three-Shaft Configuration.

Short / Medium-Haul (8,000 - 40,000lbs thrust):

Two-Shaft Configuration.

Acquisition Cost

Maintenance

Simpler engine, hence moderate:-

Overall pressure ratio

Turbine entry temperature

Bypass ratio

Figure 83: - Engine type selection long and medium / short haul (RR), pylon implications.


242

Figure 84: - Engine thrust reversal conditions need to be considered for pylon loads.

Net 25% to 30% of engine thrust

acting in reverse thrust condition

through exit apertures.


243

Figure 85: - Current engine materials are considered for engine weights and pylon loads.

Low pressure Fan stage compressor

either SPF/DB Ti or monolithic CFC

with Ti leading / trailing edge blades.

Titanium.

Nickel.

Steel.

Aluminium.

Composites.


244 DATA SOURCE :- JEFFREY M. STRIKER Chief Engineer Turbine Engine Division Propulsion

Directorate Airforce Research Laboratory USAF Presentation POWERING UP 8th March 2007

Figure 86:- Highly Efficient Embedded Turbine Engine used in my future project studies.


245

Figure 87:- Highly Efficient Embedded Turbine Engine project focus.

DATA SOURCE :- JEFFREY M. STRIKER Chief Engineer Turbine Engine Division Propulsion

Directorate Airforce Research Laboratory USAF Presentation POWERING UP 8th March 2007


246

New Engine Architecture with reduced

parts count, weight, advanced cooling,

aerodynamics and lifting.

All engine

accessories

are electrically

driven.

Pylon/aircraft mounted engine

systems controller connected to

engine via digital highway.

Internal active magnetic bearings and

motor/generators replace conventional

bearings, oil system and gearboxes

(typical all shafts)

Generator on fan shaft

provides power to airframe

under both normal and

emergency conditions

Air for pressurisation / cabin

conditioning supplied by

dedicated system

Figure 88:- Example of Rolls Royce Electric Engine concept pylon mounted.


247

Gas generator

Large diameter

duct

Contra-rotating

fan

Contra-rotating

turbine

Blended wing aircraft may offer up

to 30% reduction in fuel

consumption - 40% if combined with

electric engine concepts

Figure 89:- Example of Rolls Royce advanced engine concept pylon mounted.


248

2030 - 2040

2040 - 2050

2050 – Beyond.

Figure 90:- Advanced propulsion concept and disruptive technology projected timeline.

New Engines: - Advanced turbofan: CROR: Incorporated Engines.

Greater Aerodynamic efficiency: - Sharklet: Laminar Flow: Future Concepts.

Innovative Structures: - Thermoplastic Composites: PRSEUS: AMT: Bionic Structures.

Design and Structural Analysis Capabilities: - Virtual Design: Improved simulation.

Advanced Assembly Technologies: - Virtual Simulation of assembly methodologies:

Robotic / human interface with collaborative working in shared work space safely:

Advanced assembly processes reducing build time.


249

Figure 91:- Concept basis for application of Future Integrated Structures and MAW.

Figure 91(b):- NASA BWB Aircraft Concept

Design. Figure 91(a):- Airbus Advanced Concept Aircraft

Design.


The structural layout and initial sizing of the major airframe structural components is an iterative

process, and implies that a synthesis phase is required to establish overall details before structural

analysis can be undertaken and the design refined, involving capture of the loads identified in

sections 4 and 5, and their integration into the major structural components, applying the

methodologies from references 4, 5, 6,7 and JAR 25.

Traditionally this synthesis phase has been based on the experience of the concept designer in

conjunction with the application of simple standard equations. However “expert” programs are

becoming more readily available which encapsulate previous experience and enable the synthesis /

analysis / refinement process to be undertaken in one seamless operation (e.g. AeroDYNAMIC™

see also my Cranfield University MSc thesis on Advanced Interdiction Aircraft on LinkedIn).

However, in order to use such systems effectively it is essential to have an understanding of the

means by which a structure reacts and transmits loads. All expert programs require an initial input

of some type for example to generate the structural layout of a wing the program may only require

the external geometry of the wing, and consequently the structural configuration produced will be

determined by the historical data built into the program. The ability to input a basic internal

configuration for the structure results in more versatility and more rapid convergence to a

satisfactory solution.

The approach applied in this project to accomplish the initial sizing of the main structural members

is a combination of both the „classical‟ approach where use is made of loading data obtained from

initial loading capture and analysis outlined in reference 4, to derive shear force, bending moment,

and torque diagrams, to evaluate the initial sizes of the main structural members of the airframe. 250

Section 14:- FATA baseline wing structural analysis and component sizing.


These initial sizings based on elementary theory will then be refined as defined 3-d solid structural

assemblies using Catia V5 GSA, and NASTRAN, for detailed analysis and sizing refinement, and

systems installation. This approach provides a good basis for understanding the way in which the

structure will function, and provides an early validation of the concept and serves as a datum

against which to check the output of a more advanced analysis.

Analysis of requirements-structural design data capture:- With the exception of specific ground

loading conditions an aircraft can effectively be considered as a free body in space. Therefore in

general the airframe will be in a state of acceleration in all six degrees of freedom. Therefore it is

necessary to include all of the inertial forces and moments in the analysis used to derive the basic

structural design data which is defined as:- shear force, bending moment, and torque diagrams.

This procedure consists of the following stages:-

1) Interpreting the loading requirements as defined in the design requirements:

2) Evaluating the consequent aerodynamic loads, wing lift:

3) Calculating the implied translation and rotational accelerations, using overall moments of inertia

consistent with local load distribution (masses and centre of gravity):

4) Distributing the aerodynamic loads and local inertia effects appropriately across the airframe.

When finite element modelling is applied these distributions will be allocated as local loads at

the structural nodes:

5) Employing the „classical‟ approach the loads are initially integrated across the airframe with

respect to length to obtain shear forces, and integrated a second time to get the bending

moments or torques.

FATA baseline wing structural analysis and component sizing (continued).

251


5) (continued) In this analysis integration starts from the extremities of the airframe and proceeds

towards the centre of gravity, because working from the outside in results in any accrued errors

being relatively small in comparison with the magnitude of the local data. Additionally any errors

due to inconsistent assumptions are more likely to occur in the wing – body interface region.

N.B. when the direction of integration is changed so is the sign of the result.

The process is applicable to all overall aircraft components for example, wing, fuselage, flaps,

engine nacelles, however in all cases the moments must be in total equilibrium.

Following load capture the synthesis procedure for initial sizing of the structural members will

require the following data to be determined and researched:-

a) Reasonably comprehensive load distributions, which may be used to derive the shear force,

bending moment, and torque diagrams, together with any particular concentrated load inputs:

b) Any relevant airframe life requirements and if appropriate, stiffness criteria, (see section 5 also):

c) An initial definition of the location of the main structural members, although there is always the

possibility of revision as the design progresses and the layout is refined (see section 6):

d) An initial choice of the airframe construction materials and assembly methods (see also

sections 7,8,9, and 10).


252


Reference and datum lines:- It is important to define reference points and lines at the outset of

the structural design. Ideally a set of orthogonal axes passing through the centre of gravity of the

aircraft would be used. However this is not the most convenient since the centre of gravity moves

both longitudinally and vertically with differing fuel and payload conditions and therefore a

compromise has to be made to yield a consistent reference. A fore and aft reference located at the

nose of the aircraft is sometimes used but it is not helpful in terms of indicating the magnitude of the

forces and moments actually applied, and becomes inconvenient if the fuselage is stretched. A fore

and aft datum in the region of the centre of gravity range is better as shown in figure 92. Overall the

most suitable reference axes are considered to be:-

a) Aircraft centreline:

b) Fuselage horizontal datum in the side elevation unless the mean vertical position of the centre

of gravity is significantly removed from it :

c) Fore and aft axis located at a point 35% to 40% of the root chord, which has the advantage of

being in the region of the location of the aft centre of gravity and is close to the local mid-point

of the main span-wise structure, especially when the wing is unswept.

Swept lifting surfaces:- A particular difficulty arises when the layout of the aircraft uses swept

wings as in the case of the FATA configuration as shown in figure 93. It is logical to treat the outer

parts of the surface as an isolated structural member and to fix the span-wise reference axis along

the locus of say the 40% chord point. The problem arises in the root region where it is necessary to

resolve the bending and torsion couples into those appropriate to the overall axis system of the

aircraft. Thus what is a convenient definition for the analysis of local structural conditions becomes

inconvenient overall.

253



254

x

z

y

Span-wise.

Locate at 35% root chord.

Centreline and fuselage datum.

Figure 92:- Structural design reference axes – (datum lines).


Figure 93:- Swept lifting surface datum lines (wing skin stringers omitted for clarity).

255

Orthogonal axes

Bending moment

Centreline

Oblique aircraft axes

Bending moment

Torque

Resolve at root station


The alternative is the use of an orthogonal axis across the whole span of a swept wing implies that

in the outer region the actual torsion couples are derived as a difference between two relatively

large numerical values and it implies the local resolution of couples at each span-wise station.

Often the most satisfactory approach is the former with careful thought given to obtain the correct

components of couples at the root junction. This problem is dealt with automatically when finite

element analysis tools are applied (Catia V5,R20 GSA, or NASTRAN), although care must be taken

in the selection of the element geometries.

In either of the approaches discussed above, when defining the bending moments, and torques it is

necessary to identify the load distribution across chord-wise strips. This is straightforward when

overall orthogonal axes are used since the chord-wise strips are in the flight direction used

conventionally to define the aerodynamic loading. When the wing is treated as an isolated structural

member the structural chord-wise strips lie across the stream direction and hence it is necessary to

resolve the aerodynamic information appropriately. For this study the former approach is applied to

the wing analysis using oblique aircraft axis at 40% wing chord.

Span – wise loading of swept lifting surfaces.

A simple evaluation of the additional load distribution on uncranked swept wings may be made

using the method of Stanton-Jones (ref 20). This method uses the Weissenger approach to

interpret experimental results. The shape of the distribution is completely defined by the position of

the span-wise centre of pressure of one half of the surface ŷ, which must be betwwen the limits

0.4< ŷ < 0.5 for the method to give acceptable results. 256



257

(C)0

(C)b/2

b/2

Sweep of 0.25 chord line Λ = 35º

ŷ

y

Figure 94:- Swept wing span-wise loading notation for the Stanton-Jones method.


The notation for the application of the Stanton-Jones method of span wise wing loading estimation

for swept wings is shown in figure 94, and the equation for ŷ is as follows:-

ŷ = 0.42 + Am’ (4.45 + 5λ) tanΛ/m’ + 10.4𝜆1/2-6.7 x 10−3………(eq.1)

Where:-

A is the aspect ratio (b²/ S) = 10.15:

λ is the ratio of the tip to root chord = 0.27:

Λ is the sweep of the 0.25 chord line = 35º:

m’ = (1-MN²) where MN is the subsonic Mach number m’ = 0.2079.

Therefore:-

ŷ = 0.42 + 2.110185 (5.8) 3.368 + 1404 – 6.7 x 10−3 = 0.41

Hence the value of ŷ falls within the operational limits for acceptable results in the following

analysis.

*Let the value of ƞ = 2y/b, the for ƞ < 0.7:

c(y) CL(y) / (ċCL) = 1.28(1-ƞ²)1/2 + (14.13ƞ – 6.35)(ŷ – 0.425) .......(eq.2)

*and for ƞ > 0.7

c(y) CL(y) / (ċCL) = 1.28(1-ƞ²)1/2 + 4.25-53.8 (ƞ – 0.815)² (ŷ – 0.425) .....(eq.3)

According to ref 4 this method is likely to be as accurate as the votex-latice approach for this range

of lifting surfaces, a more general method is given in ESDU TD Memo 6403.


258


AeroDYNAMIC Jet Designer 3.0 is a Whole Aircraft Analysis tool capable of determining and

reporting the following aerodynamic behaviour of a given configuration: - Lift: Parasite Drag:

Induced Drag: Supersonic Wave Drag: Mcrit: supporting the following analysis: - constraint analysis:

manoeuvre analysis: performance and specific excess power analysis: stability and control

analysis: sizing: weight prediction: optimisation: and cost analysis. The above are based on

geometry, mission and engine data as detailed below. The initial data required in order to construct

the AeroDYNAMIC Jet Designer 3,0 analysis model in the Main spreadsheet consist of data

parameters defining the fuselage, and parameters defining the wing, and empennage geometry for

determining the initial wetted area the configuration, and the fuselage is then further defined in the

geometry spreadsheet with the 20 specific geometry data points at 20 frame stations, and this

much more accurate data is used for the Swet.

The ultimate outputs of Jet Designer 3.0 are:-Customer Focus – Needs, House of quality: Design

Synthesis – Aircraft configuration modeling: Geometric Modeling – Areas and Volumes: Aero

Analysis – Parametric aerodynamic analysis: Propulsion Modeling – Parametric Trent XWB

published data: Constraint Analysis – Design Point: Mission Analysis – Better Mission Fuel

Fraction: Structural / Weight Prediction – Weights analysis: Sizing – Sized Wing Area, WTO, and

TSL: Performance Analysis – Ps: Cost Analysis – Acquisition and operating costs, life cycle:

Sensitivity / Optimization – “Best” Design, Cost trade: evaluated against the design mission figure

95, built in the mission builder module.

In order to generate the initial trade space the baseline FATA airframe employed the same fuselage

horizontal and vertical tail dimensions, as the Airbus A350, and the FATA baseline wing the

parameters required are shown in figures 96(a) and (b). 259



260

Figure 95:- FATA baseline analysis using AeroDYNAMIC Jet Designer 3.

Twelve Disciplines: Definition

1 Customer Focus Determining the customer's needs and the priority which the customer gives to each need

2 Design Synthesis Creating design concepts that can reasonably be expected to meet the customer's needs

3 Geometry Modeling Representing the the shape of a design concept in sufficient detail that it can be analyzed

4 Aerodynamic Analysis Calculating the non-dimensional aerodynamic characteristics of the geometry model

5 Propulsion modeling Choosing an engine cycle and representing the variation of its characteristics with speed and altitude

6 Constraint Analysis Determining what combinations of wing loading and thrust loading will allow the concept to meet the ciustomer's needs

7 Mission Analysis Determining what fuel fraction the concept requires to fly customer-specified design missions

8 Weight Prediction Predicting the weight per unit area of the design concept's various components

9 Sizing Determining how large the concept will need to be in order to meet all the customer's requirements

10 Cost Analysis Determining how expensive the concept will need to be in order to meet all the customer's requirements

11 Optimization Finding the variation of the design concept that meets the customer's needs for the least cost

12 Performance Analysis Calculating the expected performance of the final optimal design, and comparing it to the customer's requirements

15,500km (9,631m) 370km

(230m)

45,0

00 f

t 13,7

16m


Figure 96(a):- FATA baseline geometry inputs for AeroDYNAMIC Jet Designer 3.0

Root Chord

Croot

Y

X

XWing

Tip Chord

Ctip

Wing Sweep ΛLE

Xengine

Yengine

Fuselage Length

Fuselage

Dia

Span b

Airframe Plan View.

Spreadsheet Nomenclature and Coordinate

System used in Jet Designer 3.0 to define

the FATA aircraft for evaluation. Note

Vertical Tail and Horizontal tail surfaces are

defined in the same way as the wing.

Root Chord

Croot

Tip Chord

Ctip

Span b

VT Side View.

261


262

Figure 96(b):- FATA baseline geometry inputs for AeroDYNAMIC Jet Designer 3.0

Y

X

Root Chord

Croot

XWing

X Exposed Wing

Exposed Root Chord

C exposed root

Note Vertical Tail and Horizontal

tail surfaces are defined in the

same way as the wing.

Airframe Plan View.

Span b

Exposed Semi Span b exposed


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Specimens: by Dawn C. Jegley : NASA Langley Research Center: Dec 2009.

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Report: by Alex Velicki et al: NASA Langley Research Center: Jan 2011.

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March 2008.

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Howe: 2004: ISBN 186058432 2.

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10) My Composite Design Capability Maintenance Studies: Private Study 2016: Mr. Geoffrey

Wardle published on LinkedIn.

11) My Metallic Design and FEA Capability Maintenance Studies: Private Study 2016: Mr. Geoffrey

Wardle published on LinkedIn. 263

Current reference material in use for the FATA paper for the AIAA list will be extended.


264


12) NASA N+3 Subsonic Ultra Green Aircraft Research: by Marty Bradley (Principal Investigator

Boeing Research and Technology) et al: Boeing Research & Technology: Published April 20th

2010.

13) Boeing 777x Airport Compatibility ECCN:9E991: by Boeing Airport Compatibility Engineers:

Boeing Commercial Airplanes: Published July 2013.

14) Automated Assembly of Aircraft Structures: by Vorobyov. Yu. A. et al : Published by the

National Aerospace University “KhAl”: Kh-Al – ERA Consortium 2013.

15) Technology and Innovation for Future Composite Manufacturing GKN Aerospace Presentation:

by Ben Davies and Sophie Wendes.

16) ABB Robotics at www.abb.com/robotics for all product datasheets and surface models of the

IRB4400/60 robot and the 6650S_90 robot.

17) Damage Tolerance in Aircraft:- by Prof P.E. Irving Damage Tolerance Group School of

Engineering Cranfield University: Published by Cranfield University 2003 / 2004.

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No 32, 1951: by Stanton-Jones.

19) Fully controlled production environment for autoclave injection processes: LOCOMACHS

Consortium: by M. Kleineberg et.al: Published 11/03/2015.

20) CODAMEIN Research Project EASA.2010.C13 Final Report 20120312: European Aviation

Safety Agency: by Zoltan Mikulik and Peter Haase: Published 12/03/2012.

21) AMC 20-29, Composite Aircraft Structure, Annex II to ED Decision 2010/003/R of 19/07/2010.


265

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Flight Control September 14-16, 2009: S.J. Campos. SP, Brazil: by Marco Aurelio Rossi and

Sergio Frascion Muller de Almdeida ITA Mechanical Engineering Department.
