My Condensed Aircraft Design Career Presentation 9th October 2016

1

Mr. Geoffrey Allen Wardle MSc. MSc. Snr. MAIAA.

Cranfield Terrasoar LMALE

UAS (produced by RTM).

Cranfield AIA F-35D / A-24 Study

(CFC / Al Li / Ti substructure).

Current FATA PRSEUS CFC / Al Li / Ti Airframe

design and automated assembly study.

MY AIRCRAFT DESIGN, STRUCTURES AND MANUFACTURING

RESEARCH AND DEVELOPMENT CAREER OVERVIEW.

This is an overview covering 16 and 1/2 years at BAE SYSTEMS MA&I (Military Air & Information) as a Senior Airframe Design Structures and Manufacturing Research and Development Engineer. The cover illustrates my major Cranfield University and non BAe / BAE Systems research studies.

Also covered is my Cranfield University MSc Aircraft Engineering, my University of Portsmouth Advanced Manufacturing Technology MSc, and British Aerospace (Military Aircraft Ltd) structural test work, as well as my current capability maintenance work, FATA research project, and future career aspirations see also my current work LinkedIn presentations.

Abbreviations and Terms used in this presentation are clarified below:-

STF = Structural Test Facility: ATDF = Advanced Technology Demonstration Facility:

SPF/DB =Super Plastically Formed and Diffusion Bonded (structures formed from Titanium sheets in one process eliminating the need for mechanical fasteners and assembly):

RTM = Resin Transfer Molding non-autoclave method for composite part manufacture:

RAF = Royal Air Force: RN = Royal Navy: CFC = Carbon Fiber Composite: CDA=Concept Demonstration Aircraft: HT = Horizontal Tail: VT = Vertical Tail: SWAT = STOVL Weight Attack Team: UAS = Unmanned Air System: FA-2 = Fighter Attack-2: CTOL = Conventional Take Off and Landing F-35A variant: STOVL = Short Take Off Vertical Landing F-35B variant: CV = Carrier Variant F-35C.

MY CAREER PRESENTATION INTRODUCTION.

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BAe BROUGH STF DEVELOPMENT OF MILITARY AIRWORTHINESS QUALIFICATION TESTS 1990-1993.

SP

F/D

B I

NB

OA

RD

FL

AP

ER

ON

MO

ME

NT

MO

ME

NT

MO

ME

NT

SH

EA

R

SH

EA

R

FIXED L/E

STRUCTURE

SPF/ DB Ti Foreplane structure,

SPF/ DB Ti Engine bay doors

structure,

Figure 1(a) Eurofighter Typhoon

wing showing CFC structure and

SPF/ DB Ti Flaperons,

Figure 1(b) G.A. of Eurofighter

Typhoon SPF/ DB Ti structures,

I developed the structural

qualification test program for

Eurofighter Typhoon SPF/DB Ti major

structural components at RAE

Farnborough and conducted this

program at BAe Brough in 1990-1991,

reporting to the Eurofighter Joint

Structures Committee, and Military

Airworthiness Authority.

This enabled the production of these

components for all subsequent

Typhoon aircraft , and for the process

to be maturely applied to the F-35

engine bay doors.

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BAe BROUGH STF DEVELOPMENT OF MILITARY AIRWORTHINESS QUALIFICATION TESTS 1990 - 1993.

The Eurofighter Typhoon CFC composite wing which are also fuel tanks consist of two wing skins

and an internal structure as shown in the previous slide, the major load bearing structures are the

wing spars and skins. The lower wing skin is co-bonded to the spars eliminating mechanical

fasteners in the highest loaded wing skin reducing not only the overall weight but the thickness of the

wing skin as shown in figures 2(a) and 2(b).

From 1991-1993 my major role was to developed the structural qualification test program for

Eurofighter Typhoon lower wing skin co-bonded “J” spars addressing design configuration issues,

for the Eurofighter Joint Structures Committee and Military Airworthiness Authority, enabling the first

flight target be met and full scale IPA aircraft production to start.

Developing and researching test methodologies i.e. T - pull T – shear rig and environmental chamber,

developing a test proposal with designs based on theses studies in conjunction with stress,

airworthiness (internal BAe and external DRA), and rig design and manufacture. Conducting test

program evaluating the results, report writing and presentation.

I was also responsible for investigating through physical testing Eurofighter Typhoon Co-bonded

Wing Configuration structural issues :-

methods of reduction of bondline peel stress

test „t‟ pull configuration

max stress at flange toe n/mm2

4

5 „FILM‟ ADHESIVE

(BSL.322)

„CLEAVAGE‟ FILLED WITH

UN-CURED CFC WEDGE

RELEASE AGENT

PRE-

CURED

CFC SKINS

UN-CURED „Z‟ & „C‟

SPAR ELEMENTS


SPAR ELEMENTS

CONFORMABLE TOOLING SHOWN THUS:

Figure 2(a):- Co-bonded spar assembled

uncured.

Figure 2(b):- Co-bonded to wing skins.

Figure 2(a) (b):- Composite Spar manufacturing research and test for Typhoon.

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Figure 2(c):- My Composite Stringer based on my STF spar 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.

7

Figure 3(a):- A400M Fatigue test rig mounting.

Figure 3(b):- F-35B Fatigue test rig mounting.

For transports hydraulic jacks apply computer

controlled loading case spectrum through skin

bonded tension pads.

Same methodology applied to fighter

aircraft:- hydraulic jacks apply computer

controlled FALSTAFF or full spectrum

flight by flight loading cases to the

structure through skin bonded tension

pads.


Airframes shown are for illustration only actual

airframe FSAFT were the Hawk TMk1a: Harrier GR5-

GR9: Goshawk: Typhoon Front Fuse (Pressure and

bending Conditioned) Typhoon Wing Substructure:

Typhoon SPF/DB Foreplane: and Tornado F-3 wing.

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Figure 4(a) (b):- Typhoon Fatigue test rig mounting designed and actual airframe in rig.

Figure 4(a)

Figure 4(a)

Figure 4(c) (d) :- Fatigue test loading direct actuators and pads, Figure 4(e):- actuators providing loads via Wiffle tree

and pads.

Figure 4(c) Figure 4(d) Figure 4(e)

My major role running in conjunction with new

airframe structural development and

qualification was the running, fatigue

inspection, and fatigue damage repair

development for the full-scale airframe

structural tests of Harrier GR-5, and Harrier T

Mk4/ Mk2 (which supported the structurally

identical Harrier FA-2 fleet). These MAFT‟s

which were run ahead in fatigue cycles of the

operational aircraft enabled the end users i.e.

RAF and RN Fleet Air Arm to be apprised of

through life structural damage issues and

methods of repair before an aircraft became

unsafe or failed in service. These repair

schemes when approved were certified through

the Military Airworthiness Authority.

One of my major contributions in this field was

the teardown inspection of the Harrier TMk2 /

Mk4 , where major potentially service life

ending damage was discovered in the centre

fuselage. I developed an inspection and repair

methodology for this damage which enabled the

Royal Navies Fleet Air arm FA-2 aircraft to

remain in service for ten years longer than

would have been the case.

Figure 5(a) RN Harrier FA-2,

Figure 5(b) Harrier Centre Fuselage Structure,

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The subject of my Individual Research Project was the

development of the qualification program, and product design

and manufacturing maturation of the Kamar Femoral head

securing implant.

1996 to 1998 MSc Modules undertaken and passed (highest

mark A):- CAD/CAM: Research Methods: Manufacturing

Systems: Composite / Metallics Manufacturing Processes

and Materials: Automation Robotics and Control: Project

Management: Operations. (Full Time) Awarded on 1st May

1998.

MSc Advanced Manufacturing Technology: - University of Portsmouth UK. 1996-1998, graduated 1998.

BAE SYSTEMS Warton ATDC Low Observable Technology Integration IPT.

My first major design role within BAe / BAE

SYSTEMS upon re-joining the company as a

design engineer post University of

Portsmouth MSc in January 1999, was

develop low observable structural concepts

for the wing leading edges and weapons bay

doors for the Anglo French, Future Offensive

Air System project.

Further work on FOAS involved the CFC

structural layout design of the wings of the

non-flying pole signature measurement

airframe shown in figure 6.

Another major work was to investigate new

airframe manufacturing methodologies

required for BAE SYSTEMS to build low

observable aircraft in production quantities.

My final work on FOAS in as part of concept

engineering before moving to JSF, involved

concept design trade studies for engine

intakes for the evolving FOAS aircraft

studies.

Figure 6 (a)The BAE SYSTEMS full-scale FOAS low

observable non flying technology demonstrator ,

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BAE SYSTEMS Warton ATDC FOAS Concept Engineering.

One of my first major design UAS concept design role was to conduct trade studies

for leading edge and intake LO configurations for both the manned and unmanned

elements of the FOAS project from 1999 to 2001 (project cancelled in 2005).

The released concept designs are shown below as figures 6(b) and 6(c).

Figure 6 (b) The BAE SYSTEMS MA&I FOAS

Manned element,

Figure 6 (c) The BAE SYSTEMS MA&I FOAS

Unmanned element,

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BAE SYSTEMS Samlesbury F-35A HT Test Block Structural Design Team.

Figure 7 The BAE SYSTEMS HT Test box design and

structures team Mr G. Wardle Concept Lead fourth in

from left completed HT test box in background.

My first major design role on the JSF/F-35 project

2001, was to design major components of a

structurally representative test article for the CTOL

AV-1 Horizontal Tail (HT) to investigate the

mechanical behaviour of the actual SDD phase HT

when subjected to real flight loads.

Because there was no mature design at this phase of

the program the major components and the

manufacturing methods for this test box would form

the basis for the final production HT, and generically

would form the template for the STOVL production

HT. This would enable both CTOL and STOVL major

control to be produced from cousin parts on the

same production line reducing costs significantly I

took design from concept to detail part design for

manufacture.

This design program was completed to cost and on

time, although there were issues in manufacture with

the new processes, fibre placement of the HT skins

was not continued into the final production program.

The completed HT test box and the team is shown in

for the F-35A shown in figure 7.

The build to responsibility for the production build

articles for HT was given to BAE SYSTEMS Brough

site.

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BAE SYSTEMS Samlesbury F-35C CV Outboard Wing Design Team.

Figure 8a:- The wing fold design incorporated a new multi lug rotary actuator driven

wing fold joint of which neither LMA or BAE Systems had any previous experience.

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Figure 8b:- Preliminary layout of the F-35C Outboard Wing .

Baseline Structural Layout of the F-35C outboard

wing on which the test box design was based.

Outboard wing test box was

representative of this area with

one forward multi-lug Ti alloy

wing fold unit skins and

substructure.

Tier 5 IPT Lead Mark Dugdale / Mike Grant

Build Line Support TBD

Sub-Systems

Integration Russ Brigham

FTI

Integration Joe Cookson

Electrical System

High Cooling Power

Coax

Business

Management TBD

Manuf.

Integration Paul Needham

Assembly Planning

WSTGE

Assembly Tooling

Mechanical Installation

Electrical Installation

Engineering Integration Support

Neil Caruthers

Composite

Skins and

Panels

Wing Fold

Interface &

Fold Rib

L/E, T/E & Tip

Interfaces &

Structure

Internal Sub-

structure

Spars & Ribs

Wing Fold

Building

Block

Wing Structural

Integration Mike Grant

Jo Dewhurst

Jas Sandhu

Geoff Wardle (LD)

Stuart Reid

Paul Metcalfe

Phil Hancock

Ravi Sharma

Mark Dugdale (LD)

C Bridgwood

Paul Metcalfe

Jas Sandhu

FE

Modelling

Alan Church

Bus Mgmt

Structures

Simon Harris

C Bridgwood

Mike Welch

Sub-Systems

Manuf Integ‟n

Design

FTI Integ‟n

KEY

Integration

activities

PAO Mass

Properties Dave Bennett

Empennage

Shared

Resource

Chart 1:- My role as IPT Lead Designer for the CV Outboard wing test box design.

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I was responsible as the F-35C Outboard wing Building Block as IPT Design Leader for

creating a test article to meet the structural validation criteria listed below:-

Validate Structural Analysis,

• Static and Fatigue Load Spectrums.

• Material Design Allowable.

Demonstrate strength and durability of Structure adjacent to Wing Fold Mechanism.

• Multi-Slice Lugs on Fold Rib

• Bolted joint between Skins and Fold Rib flange caps.

• Bolted joint between Forward Spar and Fold Rib.

Reduce Design Risk for SDD test box proposed loading.

I was responsible for a small team consisting of designer / stress / and manufacturing

engineers to develop the test articles to meet the following requirements:-

Manufacture of 2 Outboard Wing Test Articles - (1 Static and 1 Fatigue)

Test Articles will be unconditioned and tested at room temperature.

Testing to be completed by LMA.

The design for these two test boxes was completed approved and signed off by

BAE for manufacture before the outboard wing structure manufacture was handed

over to BAE SYSTEMS Canada as a workload reduction measure.

Building Block IPT Design Leader test box F-35C outboard wing.

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BAE SYSTEMS F-35B STOVL Design Lead VT SWAT design trade studies.

Responsibilities:-

Lead a small team to undertake a series of `near term‟ STOVL Weight Improvement studies including new substructure and structural layouts using my original CTOL designs as the baseline, on STOVL AFT Fuse, Horizontal Tail and Vertical Tail products, to enable selected design solutions to be incorporated into the SDD phase airframe build as soon as possible, examples of these 30 trade studies can be discussed at interview and the overall effort is shown in figure 9.

To deliver results into Empennage team and AFT Fuse team, and ultimately to John Hoffschwelle (LM) - JSF STOVL Weight Improvement Studies – Lead, to complete `near term‟ studies by March 1st 04 however agreed with John Hoffschwelle that this is CTOL personnel availability dependant, I Lead the Vertical Tail SWAT team consisting of two designers (myself and one other, one weights engineer, one stress engineer, and manufacturing engineer, I generated the original concepts and interfaced with the team, and Aft fuse teams and fuel system teams to turn them into viable solutions, reporting weekly to John Hoffschwelle (LM).

The out come of these studies were design solutions enabling the STOVL F-35 SDD aircraft to be completed and reach a weight within 10% of its target weight. I all so produced the detail design of the primary substructure for the STOVL HT-7, and CTOL vertical tail designs which enabled the mass production manufacturing to be handed to BAE SYSTEMS Woodford site of these structural components. I likewise produced the detail design for the STOVL TVT-7 horizontal tail for the mass production of these structural components to be handed over to BAE SYSTEMS Brough site.

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Figure 9:- STOVL General Weight Reduction Studies.

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Figures 10(a)/(b) BAE Systems Chairman's award for Innovation 2005 and the SWAT Team award 2004.

Figure 10(a).

Figure 10(b).

BAE SYSTEMS Samlesbury F-35B B-1 aircraft first VT Pre-shipment photo.

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Figure 11:- Vertical tail F-35B B-1 aircraft manufacturing team BAE SYSTEMS Woodford (Left )

and STOVL SWAT team (Right), manufacturing manager far right, Mr G. Wardle VT Trade

Studies design lead second from right.

Chart 2:- My role in the STOVL BF-1 VT Organisation following my SWAT team studies.

STOVL Vertical Tail Phase 2 Layout Organisational Structure

Root Rib

Rib 1

Rib 2A-E

Rib 3

Rib 4A-C

Rudder Support Rib

Spar Brackets

Jamie Smith

Design

Mark Diamond

Malcolm Downie

Ribs

Fwd Shear Fitting

Aft Spar Fitting

Aft Spar

Rudder Hinge 2

Leading Edge Spar

Tip Rib

VT Aft Fuse ICP

VT Leading Edge ICP

VT Tail Cap ICP

VT Rudder ICP

Geoff Wardle

Stuart Reid

Martin Starkie

Roy Winch

Periphery

Vertical Tail - Metallics

Simon Harris

Claire Bridgwood

Spar 1

Spar 2

Spar 3

Spar 4

Spar 5

Barry Green

Neil Doyle

Spars

Outbd Skin

Inbd Skin

Jo Dewhurst

Richard Coddington

Skins

Outbd Fairing

Inbd Fairing

Design

Craig Hannan

Fairings

Frames

Brackets

Seal Integration

Design

Structures

Substructure

Vertical Tail - Composites

Richard Coddington

Jo Dewhurst

Manufacturing Support

Daniel Parry

Vertical Tail

John Holton

Stuart Huskie

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Figure 12:- My responsibility for Major B1 VT Torsion box substructure component design.

23

Design for manufacture of

the Vertical Tail major

substructure : -

Al ribs / spars:

Ti spars and attachment

fittings:

CFC Intermediate spars.

BAE SYSTEMS Samlesbury / Brough F-35 STRUCTURAL CERTIFICATION TEAM.

Figure 13:- Combined Structural Certification Team in front of CTOL structural mock-up aft

fuse and empennage load pad layout designer Mr G. Wardle third from right on second row

back. 24

Responsibilities in the Combined Structures Certification Team.

I was responsible for developing a structural loading test solution for the rear fuselage and the empennage addressing theses issues, involving extensive liaison with Brough STF and LM:-

What are we trying to simulate?

• Aerodynamic loading

• Inertia loading

• Buffet loading

• Landing and taxiing loads

• Pressurisations (fuel, cockpit, intakes ……)

How sophisticated does the solution need to be?

What standard of test article do we require?

How are we going to support the test article?

How are we going to introduce the loads?

What systems are included in the aircraft for test, bearing in mind this is a flying aircraft subjected to proof loading?

Figure 14:- Proposed structural loading of CTOL

test article from STF Cranfield University MSc

presentation.

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The starting point is a series of „unit‟ load cases for various

elements of the structure

• Aerodynamic and Inertia loads

• Different cases for each of the key performance parameters

• Roll rate, pitch rate, vertical „g‟, etc.

• Different cases for different aircraft configurations

• Fuel state, payload, etc.

Carry out an iterative process to establish a load introduction

methodology which matches the Shear Force, Bending Moment

and Torque at pre-determined „key‟ stations

• Due attention to local strength levels at the point of

introduction

Load introduction points are then combined to provide actuator

positions

Responsibilities in the Combined Structures Certification Team cont.

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Certain „actuator‟ positions will be replaced by fixed reactions to

restrain the test article in all 6 DoF‟s

• Engine thrust loading

• Undercarriages

• Aircraft hoisting points, etc.

Where it is not a full aircraft, means have to be found to replicate the

interface between the test article and the „aircraft‟

With a knowledge of the positions where the aircraft is going to be

loaded, the maximum load likely to be applied and the likely

deflections the test article will experience, the initial concepts of the

rig can be developed.

Responsibilities in the Combined Structures Certification Team cont.

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Chart 3:- BAE SYSTEMS/Cranfield University Terrasoar UAS project organisation chart.

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BAE SYSTEMS Airframe Design Lead for Joint Terrasoar Project MALE UAS.

In addition to my F-35 roles and responsibilities from 2003-2006 I

was responsible as the Airframe Design Lead for a joint Cranfield

University BAE Systems Light MALE UAS Project. The objectives

were to design build and fly a MALE UAV to be built from novel

materials and using new techniques to BAE SYSTEMS, this also

formed the MSc group design project. See Cranfield University MSc

section of my LinkedIn profile for full overview.

Figure 15(a):- The resulting airframe was to have

load CFC bearing fuselage skins with minimal

machined metallic components, for a low cost and

low risk conventional UAS layout with the utility of

preliminary flight trials of new FCS for BAE

SYSTEMS autonomous air vehicles.

Figure 15(b):- Final Terrasoar wing as

built configuration with flight controls

installed. 29

4 plies 90º

4 plies 0°

4 plies +45°

4 plies -45°

4 plies 90º

4 plies 0°

4 plies +45°

4 plies -45°

Outboard wing skin thickness = 3.41mm.

Figure 16:- Initial outboard wing skin ply layup using AIMS IPS 05-01-001-07 Unidirectional

tape/180°C curing class standard modulus fibre.

30

+45º

- 45º

Outboard wing provisional CFC lay up.

N.A.

0º

0º

+45º

+45º

- 45º

- 45º

- 45º

+45º

0º

0º

90º

90º

90º

90º

Symmetrical Balanced Ply Laminate with no consecutive ply orientations

exceeding 60º separation:-

0º Plies reacting bending loads:

+/- 45º Plies reacting chordwise shear loads:

90º Plies reacting aerodynamic suction loads.

Figure 17:- Initial inboard wing skin ply layup using AIMS IPS 05-01-001-07 Unidirectional

tape/180°C curing class standard modulus fibre.

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4 plies 90º

6 plies 0°

4 plies +45°

4 plies -45°

Inboard wing skin thickness = 3.84mm.

Inboard wing provisional CFC lay up.

+45º 0º

0º

+45º

0º

- 45º

- 45º

0º

+45º

0º

0º

+45º

-45º

-45º

N.A.

90º

90º

90º

90º

Symmetrical Balanced Ply Laminate with no consecutive ply orientations

exceeding 60º separation:-

0º Plies reacting bending loads:

+/- 45º Plies reacting chordwise shear loads:

90º Plies reacting aerodynamic suction loads.

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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:

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 / or protective paint / sealant see figure 18 on next slide.

To reduce weight of the airframe structure the airframes primary materials were changed

form glass fibre, to carbon fibre epoxy resin composites, hence the risk of corrosion due to

the galvanic compatibility of materials and coatings became an issue.

Weight reduction, and new technology demo change to RTM Carbon Epoxy March 2006.

Figure 18:- Corrosion prevention methods used for carbon fibre structures in contact with

Aluminium components when change from glass to carbon epoxy was made.

IN ALL AREAS WHERE ALUMINUM AND CARBON

WERE TO BE IN CONTACT THE FOLLOWING

TREATMENTS WERE APPLIED.

33

34

Figure 19:- Resin Transfer Moulding (RTM) Basic process overview.

Port outboard wing in RTM Carbon Epoxy first build components as of March 2006.

Figure 20(a) to (c) :- Terrasoar Outboard wing as

built configuration.

(b)

(c)

(a)

Figure 21:- Terrasoar wing centre tool.

35

Figure 22:- Airframe structure in RTM Carbon Epoxy first build mid 2006.

36

1. MALE UAV built from Carbon Epoxy using resin infusion for

fuselage, RTM, for wing, at BAE SYSTEMS Manufacturing

Technology Centre Samlesbury.

2. Airframe in final assembly tooling, test articles completed and tested.

3. Systems fit due for completed in July 2006.

4. Role out first week in August 2006 with proof test in second week.

5. Engine ground tests and fit checks completed.

6. Completion of all ground tests including high speed taxi testing due

by last week in August 2006.

7. Flight tests due for completion at the end of September 2006, with

handover to Autonomous Air Vehicle Systems in October 2006.

8. Total project cost £100,000.

9. Production rights handed over to BAE SYSTEMS Australia 2008.

BAE SYSTEMS Samlesbury Terrasoar Project MALE UAS project status 2008.

37

38

IRP Background and Mission requirements capture.

During 1995 LMTAS proposed conversion of “mothballed” F-16A fighters into interim

UCAV‟s to meet a USAF fighter aircraft shortfall in 2005-2015 timeframe by replacing the

wing with a 60ft low aspect ratio planform, and removing the cockpit and pilot systems. This

however would not result in an aircraft suitable for today‟s warfighter as this „Defender‟

would be compromised in speed, non-stealthy and cost $3 to $5 million per jet modification.

Whilst Leading the F-35B SWAT trade studies and Leading the design of the joint Cranfield

University / BAE Systems Terrasoar light UAS team as major part of my MSc studies I

developed an Advanced Interdiction Aircraft (AIA) concept design in both manned and

unmanned variants. This proposal study went from requirements capture H of Q (Table 1),

through to preliminary design and produced a modular modifiable manned and unmanned

FB-24 / F-35D / A-24 airframe with an estimated cost of $500,000 to $ 1 million per aircraft,

and was a two year study from concept to preliminary design using USAFA Aerodynamic

MDO toolset for analysis, the final report was submitted to the F-35 project office LM, and

ITAR cleared for Cranfield University and involved Catia V5 surface / solid / and FEA

modelling in V5.R10. An overview slide presentation is in the Cranfield University MSc

section of my LinkedIn profile for the A-24 along with the complete MSc AIA Thesis in pdf

format.

The both the FB-24 / F-35D and A- 24 would employ supercruise and stealth to reach time

critical targets, employing the selected mission profile, and with the F-120 VCE would have

loiter capability for targets of opportunity.

CU / BAE / LMTAS CONCEPT STRUCTURAL AND CONFIGURATION DESIGN FOR AIA.

Table 1:- H of Q requirements capture to evaluate the importance of each AIA requirement.

39

Figure 23:- FB-24 / F35D / A-24 Final down selected configuration side and front views.

18.70

CoG Most Fwd = FS 9.19

CoG Most Aft = FS 10.11

LG = 8.086m

420 53.50

Ground line

16.250 AI View angle

51.60 EOTS Fwd View angle

500 5.945m

13.722m

3.328m

A/C height = 3.79m

Tip back angle

40

41

Tip over angle = 71.90

CoG Most Aft = FS 10.11

CoG Most Fwd = FS 9.19

W = 3.328m

520 15.320

520

19.153m

Figure 24:- FB-24 / F-35D / A-24 Final down selected configuration plan view.

42 Wing structure Ti Carbon CFC with BMI

inner ply skin

Forward fuselage build

module in carbon PMR-15

Weapons bays Ti SPF/DB

Ceramic composite/ Structural

RAM leading edge flap

Center fuselage

build module in Al

and Ti

Ceramic composite / Structural RAM

leading edge flap

Aft fuselage build

module in Ti

Ceramic / Structural

RAM flaperon

Ceramic composite /

Structural RAM flaperon

Wing structure Ti Ti / Structural RAM loaded

core ruddervators

Figure 25:- FB-24 / A-24 AIA Common Structural integration layout within the IML.

Stbd Main u/c bay

Port Main u/c bay

AI module

Fig 26:- To reduce wing skin thickness multi spar pitch was used to inhibit skin buckling.

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

in the order of twice that of the sub-structure.

Therefore where the wing chord thickness is

between 3.9 inches and 11.8 inches, it is more

efficient to increase the number of spars in

order to reduce the skin thickness an hence

reduce weight. Although for highly loaded

combat aircraft spars are used in wings with

root chord thicknesses up to 15 inches in

combination with stiffeners.

N.B. in military combat

aircraft wing ribs are

generally limited to the

weapons carriage and fuel

tank boundary stations.

i.e. long thin panels are more

efficient at resisting buckling

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

CFC intermediate spars and rib trade study. 43

Figure 27:- A-24 Wing metallic sub – structure to Ti boundary joint philosophy.

1.2”

Web to stiffener Outboard Joints.

* Based on 3 x fasteners diameter = 0.1875”

0.34” x 450 Chamfer

r =0.375”

0.4” **

0.2”

0.45” *

1.5”

t = 0.2”

t = 0.1”

Const.

Const.

Al rib Bathtub nested into Ti spar inboard Joints.

* Based on 2 x 0.1875” fasteners diameter + 0.06”

clearance.

** Based on diameter of Eddie bolt installation tool and

footprint of clickbond nutplate.

NB: - Dimensions will vary with web / cap thickness.

0.15”

0.45” *

r =0.16”

0.375” 0.375”

2d

t = 0.12”

0.5

6”

d =0.1875”

44

Figure 28:- A-24 Wing composite sub – structure to Ti boundary joint philosophy.

Web to spar stiffener Outboard Joints.

Tab attachment to integral spar stiffener

considered adequate for outboard joints.

Composite rib nested cap Inboard Joints.

Integral stiffener landing would remove the need

for cleated inboard joints reducing parts count.

d =0.1875” d =0.1875”

Ti boundary spar.

Ti boundary spar.

2.5-d 2.5-d

3-d 3-d

Composite rib secured by two

rib cap bolts and two web bolts

through spar stiffener. Composite rib tab secured by

two web bolts through stiffener.

45

Figure 30:- Group Design Project Terrasoar MALE from concept to flying aircraft

Design Lead for the Terrasoar airframe see chart 1.

46

Figure 29:-Individual Research Project A-24 AIA from Concept to preliminary design.

MSc Aircraft Engineering: - Collage of Aeronautics, Cranfield University UK. 2003-2006, graduated 2007.

Part-Time Student Sponsored by BAE Systems whilst working on F-35.

Figure 31:- F-35 Commonality, the fuel system integration also had to meet this target.

My role was to design and integration of a common fuel system within multi variant

airframe structures of the rear fuselage involving interfaces with the Lockheed Martin

wing and Northrop Grumman centre fuselage fuel systems teams. I conducted

successful detailed designs, and structural integration for the small and large bore fuel

lines and fuel tank gas innerting systems, as well as a common fuel dump system for the

CTOL and STOVL variants incorporating a heat shield.

47

My role in the BAE SYSTEMS Samlesbury F-35 Subsystems Organisation.

48

Subsystems

IPT Lead Brian Cowell

Design Lead

CV Glenn Edmondson

Design Lead

STOVL Ian Lever

Analysis Lead Riz Gulamhussein

4th Site Lead Glenda Dunne

Fokker Elmo WPM Phil Quinn

Business

Mgmnt Lead TBD

Electrical

Group Lead

Steve Brook

Fluid Group

Lead

Colin Ford

Electrical

Group Lead

Nathan Gibbs

Electrical

Group Lead

Nilesh Patel

Fluid Group

Lead

Steve

Reynolds

BM

CV TBD

Horizontal Integration

Electrical

Governance Ian Lever

Fluid Governance Glenn Edmondson

Fluid Group

Lead

Jamie McKay

Analysis Lead Liam Canning

Geoff Wardle

Systems Integration all variants

Design Lead

CTOL Max Kirk

BM

CTOL Rachel Willacy

BM

STOVL Ann Melling

Fluid Group

Lead

Kieran

Bowman

49

BAE SYSTEMS AS&FC MANTIS STRUCTURAL CONFIGURATION DESIGN TEAM

Following the completion of the F-35 design phase and as a result of my design work on the Terrasoar

light UAS I was assigned to the new Autonomous Systems & Future Capability group established within

BAE SYSTEMS to develop the Mantis MALE Multi-role UAS.

At this stage of the only the requirements were known so like Terrasoar the task was Concept design

through to first flight but the time scale was only 18 months.

The basic requirements were as follows:-

Be fully autonomous and all electric flight control system (no hydraulics),

Able to either be transported to a forward operating base or self deploy 66 feet wing span,

Conduct long duration ISR and strike missions with precision guided weapons,

Out-class the US General Atomics Predator A and B aircraft and incorporate advanced cost reducing

manufacturing technologies,

Easily maintained with reduced cost of ownership over manned and competitor unmanned systems

(Global Hawk).

Enabled export productionised examples to markets in Mid and Far east as well as Canada, Europe,

and Australia.

Initial concept and preliminary structural layout design was undertaken by the small Warton team of

which I was a key part, the design of the fuselage was retained by Warton for detailed manufacture, the

wing was subcontracted to BAE SYSTEMS Brough (contracted out to Slingsby for manufacture), the

manufacture of the empennage was also subcontracted to BAE SYSTEMS Brough.

Role – Design and structural layout of Mantis fuselage Spiral 1 and Trade studies for Production aircraft.

BAE SYSTEMS Warton / Preston AS&FC MANTIS MULTI-ROLE UAS.

Conceptual design of the fuselage and structural layout of the forward fuselage:

Manufacturing design of the main load bearing advanced composite fuel tank:

Integration of the forward landing gear and systems:

Detail design and integration of structural components through to manufacture and flight within a concept

demonstration airframe:

Configuration trade studies for the production aircraft for the UK and Export.

On 31st December 2011 I left BAE Systems on VR as part of a mass redundancy program.

Figure 32 :- Mantis Spiral 1 pre flight test at

test site in Australia. November 2009. Figure 33:- Mantis full size model at Farnborough Air Show

50

51

Figure 34:- Honeycomb core transition configurations for composite skins.

To reduce the structural weight of skins honeycomb

cores were used reducing skin thickness whilst

maintaining the same structural loading capabilities.

Used for structures les than 2.9” thick.

Ply/Core Edge Tolerance:- The ply and core Edge

Of Part (EOP) curves shall have a line profile

tolerance of 0.200”(±0.100”) unless otherwise

specified on engineering drawing or other applicable

document.

Side CFC skinned honeycomb structures transition at

frame joint zone. (Pictorial representation only). CFC skinned honeycomb frame structures e

closure at side skin mate, wet cleats used for

frame / skin attachment.

Tapered edges can lead to core

crushing issues requiring either a

reduced processing pressure or

friction grips external to the part to

minimise this 20º is design standard.

52

Figure 35:- Honeycomb core choices for skins based on experience.

Hexagonal Core.

The most common form of core (used for aerospace applications selected).

For soft curvature-(Can be „hot formed‟ to negotiate more severe curvatures.

„NOMEX‟ (Aramid) core is most readily available.

GRP, CFC & Metallic forms are readily available.

Shear load carrying properties are biased towards the „Ribbon‟ or „L‟ direction.

„OX‟ Core (Over Expanded).

„OX‟ core is a hexagonal form which is elongated in the „W‟ direction.

It is used to negotiate pronounced single curvature.

„W‟ shear properties are increased and „L‟ properties are decreased when

compared to Hexagonal core.

(L)

(L)

(L)

(W)

(W)

(W)

„Flex-Core‟.

„Flex-Core‟ exhibits exceptional drape characteristics – making it an ideal

choice for severe compound curvature.

Reduced anticlastic curvature and buckling of cell walls.

Negotiation of tight radii is achieved with minimal loss of load carrying

capability.

It is expensive and therefore should only be selected after a full assessment of

alternatives in the design process.

28th November 2003 Structural Layout of Composite

Components

53

Figure 36:- Woven Cloth Classifications based on experience.

53

Overview of my use of FiberSIM in composite design at BAE Systems.

During my employment as a senior design engineer within BAE Systems MA& I have used

FiberSIM I following VITAGY training for the following:-

Ply Producibility: Creation of design stations and zones: Documents (CATIA drawing objects) and

plybook documents: Flat pattern generation analysis and transfer to manufacturing: Darting:

Splicing: Multi skin core batch producibility.

There is insufficient space in this presentation to detail the procedures however a descriptive

narrative of key points is given below. The following four slides give a generic overview of the

information flow and data required to produce a FiberSIM ply and the catia geometric relationships

for document generation.

Laminate creation:- Chart 4:- Prepare the Catia geometry, create a Catia skin which is the part skin

(tool skin): create Catia boundary curve (net boundary): there are four laminate selections in

FiberSIM:- (1) PART-represents tool skin, (MUST have one PART laminate in every model: (2)

ADD SKIN- represents an over-core surface, if the surface topology changes, you must use a new

skin to represent it and create a new laminate of this type: (3) PLY PACK- an organizational tool

that represents a group of plies that are assembled in a separate process and put into the current

composite part definition, which allows the sub elements of the group of plies to be listed within the

current part: (4) UNI LAYER- an organizational tool used to define uni-directional plies that are laid

on the same layer within a layup. The Laminate Form is presented giving the Non-Geometric

Information and Links to Catia Geometry always lock FiberSIM geometry to prevent modification,

and always save the FORM by choosing ACCEPT or YES END, now create the FiberSIM laminate

using CEE+LAMINATE+CREATE enter new / laminate name / part number / laminate type /

geometry status (locked) / skin (tool skin) / boundary (net) / ACCEPT. 54

55

Chart 4:-FiberSIM design methodologies Laminate Geometry Relationship.

*FAC *SUR

Skin

*CCV

SKIN GEOMETRY.

*LN *CRV

Extended

Boundary Net

Boundary

CURVE GEOMETRY.

Laminate

Rosette creation:- Chart 5:- There are three rosette mapping types in FiberSIM which are as

follows:- (1) Standard-this is the most common, ply origin location is mapped by following the

contour of the surface: (2) Translational-zero direction is parallel to an axis of the part: (3) Radial-

zero direction points out in all directions from the center of the surface of revolution. From the

rosette form select:- Display length this is a magnification factor for the rosette spokes: Rosette

type (as shown in chart 2): and Define the rosette zero direction in one of three ways either:-

Another point / Catia axis / or Line or curve through the origin.

Now the rosette can be created:- CEE+ROSETTE+CREATE entre new / Origin (select point on top

of tool skin / Direction key e.g. x / Adjust Display Length e.g. 100/ ACCEPT, and the rosette is

created.

As can be seen from Chart 6 ply generation for producibility analysis requires material definition,

this is the result of selections made from the Materials Database and inputs on the Ply Form.

The FiberSIM Materials Database contains many common composite materials, the limit angle

being the most important parameter for the FiberSIM producibility simulation. Note not all

information in the materials can be viewed in a single Catia view therefore multiple views are

required to view other material parameters.

The Ply Form is used for entering specific orientations as 0/90, 90/0, +/-45 and -/+45, (note

user must type “+/-”) also the user cannot use CTRL-ALT-U. FiberSIM creates a link between

the non-geometric composite data and the 3D geometry through the ply form.

To create the FiberSIM ply:- CEE+PLY+CREATE / new / Set Step 10 / Select Material (e.g. PPG-

PL-3K) / Lock Geometry / run producibility.

56


57

Chart 5:- FiberSIM design methodologies Rosette types and Geometry Relationship.

90°

0°

45°

-45°

Rosette

*PT

Rosette

Origin

ORIGIN GEOMETRY.

*LN *PT *AXIS

*CRV *CCV

Zero

Direction

DIRECTION DEFINITION.

45°

90°

-45°

0°

Standard

45°

-45°

90°

0°

Y

Z

X

Translational

Radial

58

Chart 6:-FiberSIM design methodologies Requirements for Producibility analysis.

Tool

Surface

Edge of

Part

Laminate

Skin

Net

Boundary

Rosette

REQUIREMENT. DATA COMES FROM. DEFINED BY.

Ply Origin

Fiber

Direction

Rosette

Origin

Zero

Direction

Material

Definition

Materials

Database Ply

From the above ply generation stage ply producibility can now be undertaken:- Click on Flat Net Ply

Boundary / <YES:RUN> (producibility) / <NO:REFUSE> (fiber paths) / <YES:RUN> (flat pattern) /

Change screen to VISTAGY-SPLIT to view flat pattern / Change screen to VISTAGY-SPACE /

<NO:REFUSE> (flat pattern) / <NO:REFUSE> (splice curves) / Save PLY FORM / <ACCEPT> or

<YES:END>.

Sequence and Step in FiberSIM:- The components of a composite part must have an assigned

relationship to each other to define the part‟s layup order. FiberSIM uses SEQUENCE and STEP to

define layup order.

STEP:- is used to define ply order, plies that are laid up at the same time are given the same

step number.

SEQUENCE:- is used to define laminate order , when a new laminate is used to define a new

surface topology it is given a new sequence.

Core sampling conducted in FiberSIM:- Three Core Sample Types are available which are:-

SUMMARY-ply name, orientation, stagger, material, thickness: DETAILED-ply name, orientation,

warp and weft deformation angles: LAMINATE RATING-% symmetry, % laminate balance, %

laminate warpage.

Core sampling is performed via:- CEE+STATION+SAMPLE / Select<none> next to Digitized Points

/ select points / <YES:DONE> / Set Results = SUMMARY / Click on Preform Core Sample / Click

on FWD to toggle through pages of SUMMARY information / <YES:END>.

59


Laminate Rating Core Sample.

Symmetry:- Percent of encountered components pairs equidistant from the laminate centerline

that have identical fiber orientations:

Weighted Symmetry:- Percent of encountered components pairs equidistant from the laminate

centerline that have identical fiber orientations and material thickness:

Mechanical Symmetry:- Percent of encountered components pairs equidistant from the

laminate centerline that have identical fiber orientations and material properties:

Laminate Balance:- Percent of laminate at the core sample location that has the same number

of components with positive and negative fiber orientations:

Laminate Warpage:- Percent warpage of the laminate after undergoing a specified temperature

gradient (default is a Δ250°F), the warpage prediction is based on mechanical symmetry of the

ply layup.

Symmetry:- refers to ply order about the laminate centerline or neutral axis. The ply order must

be mirrored about the centerline to have symmetry.

Balance:- refers to the relative number of +45° and -45° plies in the layup. To have balance

there must be the same number of +45° plies as -45°plies.

This has just been a brief overview of creating a laminate, and core sampling for a laminate layup,

there are many aspects of FiberSIM that I have employed during my time at BAE SYSTEMS MA&I.

60


61

Chart 7:- FiberSIM design methodologies Document Geometry Relationship.

TEXT GEOM

Doc

Template

Skin

Extended

Boundary

Net

Boundary

3D ENTITIES. 2D ENTITIES.

Document

Currently I am conducting a conceptual design study into the application of PRSEUS and other

advanced manufacturing technologies using NASA / Boeing studies as my structural starting point,

and mission adaptive flight control surfaces, to assess their benefits, along with automated

assembly when applied to the airframe of the Future Advanced Technology Aircraft transport, this is

a technical trade study paper for per review and presentation through the AIAA. The baseline

aircraft selected is a CFC twin engine 250-300 seat class point to point aircraft design of

conventional configuration, to determine the structural / weight / and aerodynamic benefits at virtual

trade study level, for commercial aircraft structures to FAR 25 and JAR-25.571, see my LinkedIn

profile.

Charts 8(a) through 8(d) show the FATA airframe and research project studies while, charts 9 to 13

cover the study which consists of three phases:- The first is 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: The second is major structural component layout of

the airframe initial structure with systems integration, using NASTRAN and Catia V5 GSA analysis

for structural sizing. The third is the assembly design study for both versions of the airframe

reference and new build using Catia V5.R20 kinematics, and structural layout and analysis. This will

constitute the core feasibility study to determine the benefits, and constraints of the application of

these new technologies within the limits of the virtual toolsets. Charts 9 through 13 show the design

project research proposal dependencies and time frame. The design philosophy applied is Damage

Tolerance Design using Safe Life and Fail Safe approaches where applicable. Chart 14 shows the

sizing mission and 15 shows basic PRSEUS structural elements which form the basis of the

development elements charts 16 and 17.

My current research and capability maintenance activities in aircraft design.

62

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)

63

Chart 8(a):- Overall configuration and dimensions of the FATA baseline aircraft.

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) 9,321 Range at OWE (km) 15,000

64

Table 2:- Baseline Aircraft Data for the AIAA study (highlighted data used for baseline).

65

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.

Chart 8(b):- My Future Advanced Technology Baseline Aircraft “Tube and Wing” 2030.

66

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.

Wing Carry Trough Box Structure

defined and sized ( section 7).

Wing Torsion Box Structure

defined and sized (section 7).

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

67

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 improvements

will be assessed.

AIRBUS:- A350 XWB

Boeing:- B787

Chart 8(d):- My Future Advanced Technology Aircraft Fuselage Study Baseline .

68

Chart 9:- My current research activity in aircraft design for the AIAA paper.

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).

IN WORK

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.

Major structural loads analysis and

component sizing:- Based on Cranfield

MSc Aircraft Engineering modules

using Catia V5.R20 as tool set.

69

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

70

Chart 11:- Activity dependency for the design trade studies for the FATA paper.

Work book 1:- Composite airframe design

Work book 2:- GSA airframe design

Phase 1:- Baseline composite / metallic

wing box, and wing carry through box

layout design structural component sizing.

Baseline composite / metallic wing

box and wing carry through box

design structural / weight analysis.

Work book 3:- Control surface kinematic

design analysis and sizing.

Phase 2:- Advanced concept composite

PRSEUS wing box, and wing carry through

box layout design structural component

sizing.

Phase 1:- Baseline control surface design,

structural sizing and operational analysis.

Advanced concept composite PRSEUS wing

box and wing carry through box design

structural / weight analysis.

Phase 3:- Future concept full composite

PRSEUS wing box, and wing carry through

box layout 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.

71

Chart 12:- Development Stages of the PRSEUS airframe design for the FATA program.

72

Chart 13:- Design Trade Study Project Milestones for the FATA paper.

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 WING CONCEPT DESIGN STUDY MILESTONES.

Phase 3

Phase 2

Phase 1

Workbook 3

Workbook 2

Workbook 1

73

Chart 14:- Design Trade Study Operational Profile for the FATA paper aircraft sizing.

15,000km (8,099nm) 370km

(200nm)

74

Chart 15:- 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 DMS

2436 Type 1 Class 72 (grade A) and HexFlow VRM 34 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

75

Chart 16:- 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 chart 15.

76

Chart 17:- 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

chart 16.

C/L

Overwrap

Chart 18:- Typical Building Block Methodology used to assess the PRSEUS Structures TRL.

77

Based on this Boeing Technology

Readiness Level Diagram the

PRSEUS structure manufacturing

technology is currently at TRL-6/7 for

primary structures and TRL-9 for

secondary structures.

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

STRUCTURAL PERFORMANCE.

The objective of this work is toolset skills enhancement with the Catia V5.R20 GSA system, below

are the limitations of the Catia V5 R20 FEA toolset which need to be considered when applying this

toolset:-

a)Material Linearity:- In Catia, it is assumed that the stress and strain are linearly related through

Hook‟s law, therefore metals should not be loaded into the plastic deformation region, and rubber

type materials cannot be analyzed by this toolset.

b)Small Strains:- The strains used in Catia are the infinitesimal engineering strains which are

consistent with the limitations above in (a). As an example, problems such as crushing of tubes

cannot be handled by this software.

c)Limited Contact Capabilities:- Although Catia is capable of solving certain contact problems,

they must be within the limitations noted above in (a) and (b). Furthermore, no friction effects can

be modeled by the software.

d)Limited Dynamics:- The transient response in Catia V5 is based on model superposition.

Therefore a sufficient number of modes have to be extracted in order to get good results. The direct

integration of the equations of motion are not available in this version.

e)Beam and Shell Formation:- In these elements shear effects are neglected. Therefore, the

results of thick beams and shells may not be very accurate although not an aerospace issue.

Although these issues seem severe limitations most basic mechanical design problems can be

analyzed using this tool set as such problems are governed by linear elastic analysis.

78

Catia V5.R20, FEA Skills toolset enhancement evaluating system limitations.

There are two types of solid element available in Catia V5.R20 Generative Structural Analysis

which are Linear and Parabolic. Both are referred to as tetrahedron elements shown below.

Limited Hex elements are also available. As are Linear and Parabolic shell elements as well are

limited QUAD elements.

79

Solid Tetrahedron Elements.

Linear. Parabolic.

The Linear tetrahedron elements are faster computationally but less accurate. On the other hand,

the Parabolic elements require more computational power but lead to more accurate results.

Parabolic elements have the very important feature that they can fit curved surfaces better than

Linear elements. In Catia V5 solid machined parts are generally analyzed using solid elements,

where as thin walled and sheet structures are analyzed using shell elements. Linear triangular

shell elements have three nodes each having six degrees of freedom, i.e. three translations and

three rotations, the thickness of the shell has to be provided as a Catia input. As is the case with

the solid tetrahedron elements the Parabolic elements are more accurate.

Linear

18 DoF.

Parabolic

36 DoF.

Sheet Triangular Shell Elements.

Catia V5.R20, FEA Skills toolset enhancement evaluating system components .

The element “size” and “sag” icons appear on each part on entering the Analysis & Simulation >

Generative Structural Analysis toolset. The concept of element size is self explanatory, i.e. the

smaller the element size the more accurate the results at the expense of longer computation time

and processor power. The “sag” is a unique Catia term, in FEA the geometry of a part is

approximated with elements, and the surface of the part and FEA approximation of a part do

exactly coincide. The “sag” parameter controls the deviation between the two, therefore the smaller

the “sag” value generally the better the results.

Catia V5‟s Finite Element Analysis module is geometrically based, therefore the boundary

conditions cannot be applied to nodes and elements. The boundary conditions can only be applied

at the part level. On entering the Generative Structural Analysis workbench, the parts are

automatically hidden. Therefore, before boundary conditions can be applied, the part must be

brought back into the visual working space, and this was carried out by pointing the cursor to the

top of the tree, the Links Manager.1 branch, right-clicking, selecting Show. At this point both the

part is visible and the mesh is superimposed on it, the latter was hidden by pointing the cursor at

Nodes and Elements and right-clicking Hide. This has been the methodology for each worked

example in this presentation, figures 37,39,41,44,45, and 47 show the parts, with constraints and

loading, where figures 38, 40, 43, 46,and 48 show the total displacement magnitude analysis and

Von Mises stress analysis with maximum and minimum values in each case. The three analysis

examples in this presentation form a small part of my Workbook two which is leading into complex

studies of airframe structures.

80

Catia V5.R20, FEA Skills toolset enhancement evaluating system methods.

Four examples of these ongoing studies are given here:-

1)Bearing Shaft Assembly using Analysis Connections:- Problem statement:- The assembly

shown in figure 37 consist of a shaft of 1” diameter and length 6”, and two bearings with dimensions

as shown. All parts are made of aluminum with E=10.15E7 psi and v = 0.346. The bottom faces of

the bearings are clamped and the shaft is subjected to a total downward load of 100lb distributed

on its surface. The objective of this analysis was to predict stresses and deflections in the structure.

Full stress report was produced the results are shown in figures 38(a) and 38(b).

2)Tensile Test Specimen Assembly:- Problem statement:- The assembly consisted of two steel

pins (1”diam x 3” long) and an aluminum block (10”x 4”x1”). The constrained and loaded assembly

is shown in figure 39. The end faces of the bottom pin are clamped, and the end faces of the top

pin are given a displacement of 0.01” (0.254mm) causing the block to stretch. The objective was to

determine the force necessary to cause this deflection and predict the stresses in the structure, for

this analysis Parabolic Tetrahedron elements were used for this analysis. A full stress report was

produced, the results are shown in figures 40(a) and 40(b).

3)Spot Weld Analysis:- Two sheets of made of steel having a thickness of 0.03” are spot welded

together at four dotted points as shown in figure 41. The edge AB of the bottom plate is clamped

and the edge CD of the top L section is loaded with a 10lb force. All the dimensions shown are in

inches. The objective was to use Catia V5.R20 Generative Structural Analysis to predict the

stresses in these parts. Linear Triangular elements were used for this analysis. A full stress report

was produced, the results are shown in figures 42 and 43.

81

Catia V5.R20, FEA Skills toolset enhancement worked examples.

4) Analysis of a fastened assembly:- This assembly consisted of two plates, clamped together

with a preloaded steel bolt. One plate was loaded causing the bending of the entire structure.

The objective of this analysis was to predict the stresses and deflections to which the assembly

was subjected. The top plate was 1” by 1” square with a thickness of 0.125”: the bottom plate

was 1” by 2” with a thickness of 0.125” each had a 0.125” radius hole 0.5” from the trailing edge

as shown in figure 44. The bolt had a shaft radius of 0.125” and length 0.4”, and a head radius

of 0.2” and thickness of 0.1”. The assembly was constructed using Coincidence constraint's and

the material steel was applied. The resultant assembly being meshed, restrained, and contact

connected as shown in figure 45, then a tightening force of 50lbs was applied to the bolt

tightening connection, analysis was then undertaken of displacement, and Von Mises stress in

the assembly, the results are shown in figures 46(a) and (b). Subsequently a distributed load of

100lbf was applied to the leading edge of the lower plate as shown in figure 47 in the Z direction

as a distributed force, and the assembly was re-analysed for displacement and Von Mises

stress values, the results are shown in figures 48(a) and (b).

The final outcome of this workbook will ultimately be the analysis of metallic and composite wing

structures in support of my wing research program, and the method for composite part evaluation

will be based on the procedure overview shown in figure 49, checking against a NASTRAN

component level evaluation.

82

Catia V5.R20, FEA Skills toolset enhancement worked examples.

83

Figure 37:- Example my Catia V5.R20 FEA:- bearing assembly exercise load and constraints.

2 inch 1 inch

Figure 38:- My Catia V5.R20 aluminum bearing beam assembly analysis.

Figure 38(a) :- Total displacement magnitude

analysis of the bearing beam assembly.

Maximum deflection = 0.000881691”

Minimum = 0”

84

Figure 38(b) :- Von Mises Stress (nodal

values) analysis of the same bearing beam

assembly. Maximum stress = 1902.12 psi,

Minimum stress = 17.7862 psi.

85

Figure 39:- Example my Catia V5.R20 FEA:- tensile specimen exercise load and constraints.

86

Figure 40:- My Catia V5.R20 two material tensile test specimen assembly analysis.


analysis of the tensile specimen assembly.

Maximum deflection = 0.01” Minimum = 0”in

the pins and Maximum deflection of 0.00851”

Minimum = 0.00148” in the test block.

Figure 40(b) :- Von Mises Stress (nodal values)

analysis of the same tensile specimen

assembly. Maximum stress = 50732.6 psi, in the

top pin Minimum stress = 51.8327 psi in the

test block.

87

Figure 41:- My Catia V5.R20 FEA Spot welded sheet assembly problem structure.

C

D

A

B

5 in

12 in

3 in

4 in

2 in

2 in

2 in

2 in

2 in

C

D

A

B

5 in

12 in

3 in

4 in

2 in

2 in

2 in

2 in

2 in

1in

10 in

Sheet Material = Steel:

Sheet Thickness = 0.03 inch:

Top L section loaded edge C-D:

Bottom plate clamped edge A-B.

88

Figure 42:- Example my Catia V5.R20 FEA:- Spot welded sheet exercise load and constraints.

89


analysis of the spot welded sheet assembly.

Maximum deflection = 1.38369” Minimum = 0”.


values) analysis of the spot welded sheet

assembly. Maximum stress = 35325.8psi,

Minimum stress = 265.515psi. Maximum

stress was in the weld line as expected.

Figure 43:- My Catia V5.R20 Sheet steel spot welded assembly analysis.

90

Figure 44:- Example my Catia V5.R20 Bolted assembly components for analysis.

91

Figure 45:- Example my Catia V5.R20 Bolted assembly constrained and preload for analysis.

92

Figure 39:- My Catia V5.R20 Bolted assembly preload analysis.


values) analysis of preloaded bolted



stress the bolt as expected.

Figure 46(a):- Total displacement magnitude

analysis of the preloaded bolted plate

assembly. Maximum deflection = 3.35588e-

005” Minimum =1.0” the max value being in

the bolt as expected.

93

Figure 47:- Example my Catia V5.R20 Bolted assembly constrained and preload for analysis.

94

Figure 48:- My Catia V5.R20 Bolted assembly preload with added end load analysis.

Figure 48(a) :- Total displacement

magnitude analysis of the loaded

bolted plate assembly. Maximum

deflection = 0.0448786” Minimum =

1.0” the max value being in the lower

plate edge as expected.


values) analysis of preloaded bolted



stress the bolt region as expected.

95

Figure 49:- Catia V5.R20 composite structural analysis.

CATIA V5 R20 Composite design toolset skills enhancement training.

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

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

these are outlined below see My Composite Capability Maintenance LinkedIn presentation.

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

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

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

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

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

design-to-manufacture period.

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

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

capabilities, CPM is the essential link between engineering design and physical manufacturing,

allowing suppliers to closely collaborate with their OEM‟s in the composite design process. With

CPM, manufacturing engineers can include all manufacturing and producibility constraints in the

composites design process.

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

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

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

The complete CPE / CPM design studies including these exercises, and VT spars and skins as

well as project studies constitutes Workbook 1 and the shortened composite design capability

presentation to be added to my profile.

96

Laminates generated without balanced plies about the Neutral axis will warp during processing.

During the cure cycle a Thermosetting Epoxy resin system hardens (between 120ºC and 140ºC).

When cooling from its maximum processing temperature of 175ºC the resin contracts

approximately 1000 times more than the Fibre, and this mechanism induces warpage of the

Laminate unless the layup is fully balanced about its Neutral axis which can either be a central

plane or an individual ply layer, as shown in figure 50.

97

CT1:- Introduction to Composite Design Balanced Composite Laminate.

Linear Expansitivity (of Fibres) = 0.022

x10^-6 (approximately).

Linear Expansitivity (of Resin) = 28

x10^-6 (approximately).

45º

N A

45º

-45º

-45º

90º

90º

0º

0º

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

angle more than 60º separation angle between

layers.

Figure 50:- Expansitivity difference between fibre and resin matrix illustrating

requirement for balanced ply layups around the Neutral axis.

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

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

selecting the ply group as shown in figure 51. Then reorder the ply sequence so that no adjacent

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

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

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

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

Workbook 1. The resulting laminate for this exercise is shown in figure 52 and the numerical

analysis is shown in table 3.

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

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

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

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

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

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

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

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

model.

98

CT1:- Introduction to Composite Design Balanced Composite Laminate.

99

Figure 51:- Example of my CPE methodology for balanced CFC laminate design.

A balanced ply laminate can be

produced by selecting the ply group

and the balanced ply icon.

Subsequently the ply sequence can be manually reordered so

that adjacent plies are not orientated more than 60º to each

other, manually renumbering the sequence and the ply (use

reorder children).

100

P3 = -45°

P4 = 0°

P5 = 0°

P6 = -45°

P7 = 90°

P8 = 45°

P1 = 45°

P2 = 90°

Detail A

Detail A

Tool face geometry

Laminate Ply Stack

Fig 52 (b):- Composite part laminate lay-up.

Figure 52:- My CPE design of a balanced composite laminate from WB1.

Fig 52 (a):- Final Composite Part Build.

10

0

10

1

Table 3:- Example of my balanced laminate Numerical Analysis.

PlyGroup Sequence Ply/Insert/Cut-Piece

Name Material Direction Area (in2) Volume (in3)

Volumic

Mass(lb) Aerial Mass(lb)

Center Of

Gravity - X(in)

Center Of Gravity

- Y(in)

Center Of Gravity

- Z(in) Cost

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


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



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



Fig 53:- Example of my CPE work e.g. Transition Zones part build model and tree.

102

103

Fig 54(a/b):- Example of Transition Zone completed part and ply stack-up.

(X)

(Y)

(Z)

Figure 54(a) Final Transition Zone Part Geometry.

P10 = 0º

P9 = -45º

P8 = 45º

P7 = 90º

Detail A

Detail A

Reference surface

90º Ply drop 0º Ply drop

0º Ply drop

90º Ply drop

-45º Ply drop

45º Ply drop

Figure 54(b) Ply stagger in transition zone.

104

PlyGroup Sequence Ply/Insert/Cut-Piece

Name Material Direction Area(in2) Volume(in3)

Volumic

Mass(lb) Aerial Mass(lb)

Center Of Gravity -

X(in) Center Of Gravity -

Y(in) Center Of Gravity -

Z(in) Cost

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


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

Plies Group.1 Sequence.4 Ply.4 GLASS -45 105 0.744094 0.0537642 0.0448035 5.25 5 0 0.580412





Plies Group.1 Sequence.9 Ply.9 GLASS -45 150 1.06299 0.0768061 0.0640051 7.5 5 0 0.82916


Table 4:- CT2:- Example of Transition Zones Numerical Analysis.

10

5

Fig 55(a)/(b):- Updated laminate and ply stack Limit Contour with Staggered Values.

Figure 55(a) Updated Laminate Configuration

Figure 55(b) Updated Ply Stack Configuration

New ply stagger

from Curve C 1a

New ply stagger

from Curve C 2a

New ply stack

from Curve C 1a

New ply stack

from Curve C 2a

Figure 56:- Example of my CPE work e.g. Limit Contour with Staggered Values.

106

Below are the ply layup guidelines I used in the design of composite parts at BAE Systems.

Align fibres to principle load direction:

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

to avoid distortion during cure:

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

Overlaps and butting of plies:

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

Woven cloth, no gaps or butt joints, 15mm overlap:

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

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

„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 from BAE Systems MA&I practice detailed in WB1.

107

This is the data required on all 2-D

composite drawings and this is followed

in Design Workbook 1 and the research

project.

Ply Rosette

Stagger Index

Ply Profiles

Lay-up Datum

Honeycomb Core

Profile

Ribbon Direction

Drawing from Cranfield University MSc

presentation.

108

Figure 57:- 2-D drawing annotation based on BAE Systems practice.

Required components of a

composite part 2-D drawing.

Lay-up Table

Assembly Details

Notes

Drawing from Cranfield

University MSc presentation.

Figure 58:- 2-D drawing annotation based on BAE Systems practice.

109

110

Figure 59:- Type 1 ply lay up table for simple detail parts BAE Systems practice.

N.B.:- Drawing and layup table from Cranfield University MSc presentation.

111

Figure 60:- Type 2 ply lay up table for multi-island parts BAE Systems practice.


Figure 61:- 2-D Laminate thickness variation BAE Systems practice.

112

0° 90° 45°

135°

" t "

LEGEND:

1,5 1,5 1,0 3,0

7,0

2,5 1,5 2,0 3,0

9,0

3,0 2,0 2,5 3,0

10,5

2,0 2,0 2,0 4,0

10,0 1,5 1,5 1,0 2,0

6,0

4,5 1,5 2,0 3,0

11,0

5,0 2,0 4,5 4,5

16,0

Detail „A‟

„B - B‟

THICKNESS VARIATION

FROM 4mm TO 22mm.

See Detail „A‟


Fig 62:- FATA Structural Ply Thickness Zones Lower Wing Cover Skin (preliminary).

113

10 mm

10.0

3.0

2.0

2.5

2.5

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

10 mm

6 mm

6 mm

6 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

-45º

PLY LEGEND.

6.0

2.0

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 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

This is an overview of the considerations made in joint design for the Terrasoar and other projects,

it is important to evaluate the advantages and disadvantages of both bolted and bonded

construction methods.

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.

114

Design considerations I used in composite structural assembly joint design.

The disadvantages of bonded assembly are:-

1) Disassembly, in most cases some part of a bonded structural assembly will need to be bolted

instead of bonded to permit access for repair and inspection. An example is the Typhoon

wing structure where the bottom skin is co-bonded to the structural spars, and top skin is

bolted to the same spars, permitting access from one side:

2) Surface preparation, and bond line inspection for porosity even in co-bonded joints using C-

scan ultrasonic inspection, resulting increased costs and time:

3) Need to design for bolted repair access:

4) Environmental degradation due to water absorption leading to degradation in hot / wet

condition, solvent attack:

5) Need for increased qualification testing effort to establish design allowables.

In the case of the vertical tail exercise I created for Workbook 1 based on A-24 studies, I used

bolted construction selected primarily because of the requirement to quickly, inspect, repair, or

replace damaged structural components within a first line servicing environment. For the purpose of

that exercise the external formation light bolted installation was omitted to reduce complexity of the

design and for ITAR. In the vertical tail component and assembly models bolt datum positions were

shown as points and vectors, as was the standard in my BAE Systems MA&I design practice.

115

Design considerations I used in composite structural assembly joint design.

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 63, 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.

116

Design considerations for adhesive bonded joints detailed in WB1.

117

„FILM‟ ADHESIVE

(BSL.322)

„CLEAVAGE‟ FILLED WITH

UN-CURED CFC WEDGE

RELEASE AGENT

PRE-CURED

CFC SKINS


SPAR ELEMENTS


SPAR ELEMENTS

CONFORMABLE TOOLING SHOWN THUS:

Figure 63:- Co-Bonded composite spar manufacture detailed in WB1.

Composite bolted joint design rules:-

1)Design for bolt bearing mode of failure:

2)Counter sink (CSK) depth should not exceed 2/3 of the laminate thinness if required fill

laminate artificially with syntactic core (if design rules permit e.g. not permitted for USN, or

USMC):

3)Minimum bolt pitch is 4D for sealed structures such as fuel tanks, and 6D for non sealed

structures (where D is the bolt diameter):

4)Use only Titanium alloy or stainless steel fasteners to minimise corrosion risk see table 4:

5)Use a single row of fasteners for non sealed structures and a double row for sealed structures

such as fuel tanks:

6)Minimum fastener edge distances are:-

3D in the direction of the principal load path see figure 64:

2.5D transverse to the principal load path see figure 64:

118

Design considerations composite structural bolted joint design detailed in WB1.

Figure 64 fastener edge distances.

2.5xD 3.0xD

4.0 x D

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 Workbook 1.

119

Composite structural mechanically fastened joint design shim guidelines.

120

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 5:- Galvanic compatibility of fastener materials and coatings.

121

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:

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 for aluminium alloys (see also fig 65):-

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 (as required as drill breakout material), and protective sealant (Polysulphide)

„Interfay‟.

Design against metallic corrosion in contact with carbon fibre composites.

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Figure 65:- 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.

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 66). Monolithic laminates are more damage resistant than honeycomb structures,

due to their increased compliance, however if the impact occurs over a hard point such as above a

stiffener or frame, the damage may be more severe, and if the joint is bonded, the formation of a

disbond is possible. The key is to design to the known threat and incorporate surface plies such as

Kevlar or S2 glass cloth. Airworthiness authorities categories impact damage by ease of visibility to

the naked eye, rather than by the energy of the impact: - BVID barely visible impact damage and

VID visible impact damage are the use to define impact damage. Current BVID damage tolerance

criterion employed on the B787 is to design for a BVID damage to a depth of 0.01” to 0.02” which

could be caused by a tool drop on the wing, and missed in a general surface inspection should not

grow significantly to potentially dangerous structural damage, before it is detected at the regular

major inspection interval. This has been demonstrated through a building block test program, and

the wing structures so inflicted have maintained integrity at Design Ultimate Load (DUL). These

design criteria are critical airworthiness clearances ACJ 25.603 and FAA AC20.107A (Composite

Aircraft Structures).

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Composite impact design guidelines detailed in WB1.

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Figure 66:- Structural damage risks to composite structures e.g. the wing.

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CFRP Composite are poor conducting materials and have a significantly lower conductivity than

aluminium alloys, therefore the effects of lightening strikes are an issue in composite airframe

component design and a major issue for airworthiness certification of the airframe. The severity of

the electrical charge profile depends on whether the structure is in a zone of direct initial

attachment, a “swept” zone of repeated attachments or in an area through which the current is

being conducted. The aircraft can be divided into three lightening strike zones and these zones for

the wing with wing mounted engines is shown in figure 67, 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.

Methods of lightening strike protection for military and commercial aircraft wings are shown in figure

68.

Composite lightening strike design guidelines detailed in WB1.

126

Zone 3 Indirect effects.

Zone 2 Swept stroke.

Zone 1 Direct strike.

Lightening Strike

Zones on an

aircraft with wing

mounted engines.

Figure 67:- Lightening strike risks to composite transport wing with podded engines.

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Figure 68:- Lightening strike protection of composite wing structures.

Copper grid

Fig 61(a) Aluminum foil EAP.

Fig 61(b) Copper strip Eurofighter Typhoon. Fig 61(c) Copper mesh grid Boeing 787.

There are metallic structural components employed in the AIAA design project designed by myself

these include the wing ribs which are designed to be produced as double sided machining's from

Aluminium Lithium alloy by 5 axis high speed machining, and figures 69 to 73 illustrate the

machining methods and standards applied in all machined component design. The following are

examples:- figures 74 to 76 are exercises in support of the design activity. Sheet metal design is

shown in charts 15 and 16 and figure 77 to 79 and are sheet metal design worked examples to

maintain capability.

The one of the most effective weight reduction features for the all metallic aircraft wings has been

the adoption of large scale five axis high speed machining of many structural components

previously made by the sheet metal fabrication route. This includes integrally machined wing cover

skin stringers, machined spars (with web crack stoppers), and ribs, thus enabling a reduction in

fastener weight, less scope for fatigue cracking propagating from fastener holes, reduced parts

count and assembly costs. Also joining high speed machined components can be achieved with

bath tub joints or integral end tabs without the need for separate cleats and additional fasteners.

Other weight savings have been gained from the application of titanium alloy in place of steels for

highly loaded or high temperature components produced as near net shape forgings, or even in the

case of Super Plastically Formed titanium alloy structures employed as lower wing access port

panel covers, replacing the formally sheet fabricated covers. Titanium is also more compatible than

aluminum when used with composites in that it is not susceptible to galvanic corrosion and has a

compatible coefficient of thermal expansion. Also the adoption of Aluminium Lithium alloys in such

applications as wing ribs with a density saving of 5% over conventional aluminium alloy structures.

128

Design of Machined and sheet metallic components for design studies.

Figure 69(a) Example of 3 axis machining:-

3 Axis Machining:-

During machining the cutter can move simultaneously

along the X,Y & Z axes. The tool axis orientation is fixed

during machining. Usually used for simple geometries

where missed material is not a major issue.

(This example shows the spiral milling of a shallow

pocket feature on a compound surface).

Figure 69(b) Example of 5 axis machining:

5 Axis Machining:-

During machining the cutter can move along the X, Y &

Z axes and rotate around e.g. the X & Y axes

(designated A & B axes motion) during the machining

cycle. This capability enables the Fanning and Tilting of

the tool during machining for complex deep pockets

where excess material is an issue.

Fig 69 (a/b):- Machining Methods for Metallics applied in the design studies.

X+

Z+

Y+

A

B

Figure 69(b)

X+

Z+

Y+

Figure 69(a)

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Design for Manufacture:-

To machine an External Flange surface produced as a

result of splitting the model with a „complex‟ surface is both

time consuming and costly.

Therefore to aid manufacturing, the „complex‟ surface can

be replaced by a „ruled‟ surface provided the Chord Height

Error (CHE) is within the values specified in Design

Standards. (see Figure 70)

Where the CHE value exceeds the specified maximum, the

flange is produced by splitting the model with a „faceted‟

surface. (see Figure 71).

A bespoke „Flange‟ application will be available in the near

future to automate the creation of the „Faceted Ruled

Surface‟. As this was not available at the time of writing, the

exercise accompanying the course requires manual

generation of this geometry

External Flanges produced by complex surfaces are

permissible, but should only be used in extreme cases and

in agreement with manufacturing due excessive machining

costs

Fig 70/71:- Machined Metallics:- Chord Height Error applied in the design studies.

Figure 70 Figure 71

CHE

Preferred Non-Preferred

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Design for Manufacture:-

In Figure 72 the area shaded in Black indicates the 5

Axis Landing, and is the remaining material following

machining of the internal face of the closed angle

flange, and represents the difference between the „as

designed‟ and „as manufactured‟ part.

In such cases, it is a mandatory requirement for

allowances to be made for the loss of fastener seating

area.

The remaining material can be further reduced by

additional machining.

The area shown in Black in Figure 73 represents the

preferred condition of 5 axis landings following

machining.

Figure 72

Figure 73 Preferred

Fig 72/73:- Machined Metallics :- 5 axis landings applied in the design studies.

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Figure 74(a):- My Catia V5.20 machined Frame X_700 FWD face, from OML surfaces.

132

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Figure 74(b):- My Catia V5.20 machined Frame X_700 AFT face, from OML surfaces.

134

Figure 74(c):- Example of my Catia V5.20 FD&T application to Frame X_700.

135

Figure 75:- My Catia V5.20 preliminary metallic design FATA Al/Li Rib 12.

136

Figure 76:- Example of my Catia V5.20 metallic design of complex components.

137

Generative Sheet Metal is typically used to design parts which are typically manufactured

using „V‟ benders or press tooling. This workbench cannot produce features such as Flanges

which reference surface geometry, or to create „Joggle‟ features.

Aerospace Sheet Metal is typically used to design parts which are typically manufactured

via the „Hydroforming‟ process. This workbench can produce features such as Flanges which

reference surface geometry, and to create „Joggle‟ features.

Functionality Overlap Certain functions are common to both workbenches (sometimes

with limitations), and others are workbench specific. The following table outlines these

functions:

Generative Sheet Metal only icons

Aerospace Sheet Metal only icons

Common Icons

Limited functionality compared to

Generative Sheet Metal workbench

Chart 19:- Design of sheet metallic components for capability maintenance.

Chart 20:-Catia V5.R20 „New Part‟ Sheet metal process overview.

Select Generative Sheet Metal Design from Shareable Products tab in Tools / Options / General

Create New file

Enter Generative Sheet Metal Design workbench

Set Sheet Metal Parameters

Create Wall

Create Features

Check Flattened Component

Create Block and Heel Lines / Curves

Save CATPart

This was a specific BAE Sheet Metal methodology.

138

Figure 77:- Example of my Catia V5.R20 Aerospace sheet metal frame design.

139

140

Figure 78:- Example of my Catia V5.R20 Generative sheet metal design work.

Figure 79:- Example of my Catia V5.R20 Generative sheet metal design work.

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142

Open Model

Analyse Surface

Create Surface

Create Wireframe Geometry

Save Model

Condition Model

Production

Standard?

YES

NO Smooth Surface

Final Checks

Cutting Planes

Distance Analysis

Porcupine Curvature

Connect Checker

Local Smoothing

Smooth Discrepancies

View Modes

Global Smoothing

No Hyperlink

Hyperlink to Task

KEY

Global

deformed

surface ?

Split YES

NO

Chart 21:- Surface process workflow used to create my project surfaces.

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Figure 80(a):- FATA Key datum OML surface assembly model.

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Figure 80(b):- Project Wing torsion box datum surface model.

145

Figure 81:- FATA Wing carry through box datum surface assembly model.

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Figure 82(a):- Main landing gear installation into FATA surface assembly.

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Figure 82(b):- Main landing gear installation into FATA surface assembly.

Create new drawing

Create Project Specific Drawing Border

Filtering Data for Assembly Views

Instantiate Catalogue Details if required

Annotate Views if required

Save CATDrawing

View Creation

View Modification Options

Assembly View Content Modification

Create Drawing Comments

= Hyperlinks

Manual Pre-selection

Scenes

From Scenes

Spatial Query

Lock the Views

Overload Properties

Modify Links

Local Axis System

No Hyperlink

Hyperlink to Task

KEY

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Chart 22:- Catia V5.R20 New Drawing Overview / Process Outline.

Figure 83:- Example of my Catia V5.20 frame X-700 metallic machined part.

149

150

Figure 84:- Example of my Catia V5.20 metallic machined assembly.

151

Figure 85:- Example of my Catia V5.20 metallic sheet metal part.

N

Y

Is a Key

Diagram

available?

Does

Production

Assembly

exist?

Does Data

already

exist?

Is Reference

Geometry modelled

in local axis?

Chart 23:- Catia V5.R20, Adding To / Creating Data in a Production Assembly.

Start

Y

N

Verify Position of Data

N

Open Production Assembly Create Production Assembly

Insert Existing Data Add New Data

Y Snap data to Key Diagram

Position as required

N

Y

No Hyperlink

Hyperlink to Task

KEY

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The BAE Systems methodology of Product Assembly creation.

Is also referred to as the „Vehicle Assembly‟

In CATIA V5 terms, it is the CATProduct holding all

the CATIA data relevant to the design of this

„vehicle‟, in effect, it is the „virtual aircraft‟ - the

DMU

Within this structure, key parts are located with

respect to a Key Datum product which was also

used to position the „reference geometry‟

To ensure engineers working on the project have

access to the correct „reference data‟, the content

of the product structure is organised such that the

data is held within „master models‟ located in the

upper region of the tree structure in a component

node named „REF_REFERENCE_GEOMETRY‟

Designers take the required reference geometry

from the „master model(s)‟ into their own after

inserting and positioning it correctly within the A\C

environment

This „master geometries‟ methodology will be

employed throughout the FATA project.

Production (or Vehicle) Assembly Reference geometry „container‟

„Reference geometry‟

assemblies by „design

discipline‟

„Design assemblies‟

by „design discipline‟

Std. Parts „container‟

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Chart 24:- Catia V5. R20 Assembly Positioning Options.

Various positioning options are available, the majority of which were covered during the Fundamentals course

The functions illustrated are available in the Assembly Design and Digital Mock-Up (DMU) Navigator

workbenches

These functions illustrated have been used by myself at BAE Systems and will be employed in the FATA project.

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Chart 25:- Creating a Production Assembly with Reference Geometry.

Create New Production Assembly

Create a New Reference Component and Fix

Check for latest and Insert Key

Diagram into Reference

Component and Fix

Check for latest and Insert

Reference Geometry into

Reference Component

Snap data to Key Diagram and Fix

Is the Reference Geometry

modelled in local axis?

Y

N

N

Y

Fix Geometry

Is a Key Diagram available?

Insert Reference Geometry into

Reference Component

Position as required

Fix Geometry

No Hyperlink

Hyperlink to Task

KEY

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Symmetry plane.

Symmetrical outboard wing spar section (illustration only) representative

for WING_STBD_LEADING_EDGE_SPAR_SECTION_0001

Figure 86:- Example of my use of assembly in DMU for the symmetrical wing spar tooling.

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Figure 87:- Example of my use of assembly in DMU for the LE rib post assembly.

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Figure 88:- Example of my use of assembly in DMU for the spar assembly.

OB Port Leading Edge

Spar with rib posts and

splice assembled.

Trailing edge spare developed with

datum‟s for Low speed aileron

attachment.

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Figure 89:- Example of my use of assembly in DMU for Rib 12 fit check.

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Figure 90:- Example of my use of assembly AIA F-24 and Human Builders.

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Figure 91:- Example of my Catia V5.20 ABB robots with human builder for kinematics.

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(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 92:- Axis movements / working range of ABB IRB 4400/60 articulated arm Robot.

My future design career aims are within advanced aircraft design.

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Figure 93:- Design and development of aircraft composite and metallic major airframe structures to JAR-25.571

and the development of advanced manufacturing and assembly technology for future aircraft.

My Condensed Aircraft Design Career Presentation 9th October 2016

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Transcript of My Condensed Aircraft Design Career Presentation 9th October 2016