Terrasoar 9th intake thesis Geoff Wardle

CRANFIELD UNIVERSITY

SCHOOL OF ENGINEERING

MSc THESIS

Academic years 2003 – 2006

GEOFFREY A WARDLE MSc CEng Airframe Design Lead.

WING DESIGN AND SYSTEMS INTEGRATION FOR THE

TERRASOAR HALE UAV GROUP DESIGN PROJECT

Incorporating: - Conceptual design of the complete aircraft and the

wing: Major Component layout and detailed design of the wing:

Manufacture: and Structural qualification assessment proposal.

Supervisor: Mr Phillip Stocking

November 2005

This thesis is submitted in partial (40%) fulfilment of the requirements for the degree of Master of Science

in Aircraft Engineering.

Cranfield University 2005. All rights reserved. No part of this publication may be reproduced without the

written permission of the copyright holder.

TERRASOAR BAE Systems / Cranfield University G. A. Wardle MSc CEng MSc Group Design Project Thesis.

i

Theses “Health” Warning

This thesis has been assessed as of satisfactory standard for the award of a Master of

Science degree in Aircraft Engineering. This thesis covers part of the assessment

concerned with the Individual Research Project. Readers must be aware that the work

contained is not necessarily 100% correct, and caution should be exercised if this thesis

or the data it contains is being used for future work. If in doubt, please refer to the

supervisor named in the thesis, or the Aerospace Engineering Group.


ii

Abstract.

This thesis covers the group design project of the 2003 9th

Intake of the Cranfield

University part-time Master of Science Degree in Aircraft Engineering, focusing on the

work of the author as part of the project team.

The objective of the Terrasoar project was to design, manufacture, ground test, and fly a

High Altitude Long Endurance small low cost Unmanned Air Vehicle with CAA

certification for flight at 10,000ft (3,048m) in uncontrolled airspace, with an endurance of

5hours, and payload of 5kg. However although CAA certification was not possible at this

time all of the structures and methodology is in place, and this aircraft will be certificated

to fly within controlled airspace.

The final solution has been to design and manufacture an aircraft which will meet a set of

less demanding missions criteria up to an altitude of 400ft (121.92m), and an endurance

of 4.5hours, which was within the scope of the project. This aircraft has the capability of

being modified with additional systems, new outboard wing, and a new engine, to fly at

the original target altitude. These modifications could reasonably be considered for a

future intake to the part – time Cranfield University / BAE Systems Aircraft Engineering

MSc course, and pursuit of eventual certification is a realistic goal as the route to full

certification has been established by the 9th

intake.

This thesis covers the author‟s contribution to the conceptual design phase and

configuration investigation, and the author‟s role as the Airframe Lead and wing designer

of the final Terrasoar configuration, which has lead to the manufacturable design solution

which meets the current missions identified for this aircraft.

At the time of writing Terrasoar tooling OML (Outer Mould Line) has been

manufactured for all major airframe units, the FCS has been test flown and the engine has

been ground tested and the final assembly jig and tool has been designed. The materials

and other long lead time items are due to be ordered and manufacture is due to begin in

February 2006, with ground testing in June and first flight targeted for July or September

2006.

This thesis covers the wing design maturation up to the 30th

November 2006 when

authority for further detail design changes, such as any minor manufacturing easement

changes of the wing design was handed over to BAE Systems Warton / Samlesbury New

Business unit, and effectively the author passed over responsibility for the design to BAE

Systems.


iii

Acknowledgements

Firstly I would like to thank BAE Systems for allowing me to participate in this MSc

course, in particular my thanks are offered to Andy Bruce who supported my application

for the course and maintained an interest throughout even during my movements around

the Joint Strike Fighter / F-35 Integrated Product Team.

From Cranfield University, School of Engineering, Aerospace Engineering Group I wish

to thank Mr Philip Stocking whose strong chairmanship and guidance in the face of

external forces has ensured that this project did not veer off course.

I would like to thank BAE Systems Manufacturing Technology team in particular Mr

Robert Cross and Mr Craig Carr (who was originally a member of the 9th

intake) for their

time and energy in helping the team in developing a manufacturing methodology which

will ensure this aircraft is built.

Particular thanks go to my colleagues from the 9th

intake of the Aircraft Engineering MSc

course:-

Paul Gilligan James Pennington Bob Currie

Damian Adams Alan Barnes Dave Baird

Rob Sneddon Vernon Hind Craig Carr

Jon Baggaley Rachael Cunliffe Rob Cunliffe

Emma Bradley.

All of who participated towards the successful completion of the Terrasoar project with

varying degrees of theoretical and practical work. Additionally I would like to thank Mr

Ian Isenburge of JSF IT support for transferring the CATIA models and documents I

crated at BAE SYSTEMS on to disc enabling me write this thesis at home.

Finally I would like to thank my partner for encouragement understanding, coffee and

food throughout the duration of this particularly demanding course.


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Contents Page

Theses “Health” Warning i

Abstract ii

Acknowledgements iii

Contents iv

Figures vii

Tables‟ xiv

Glossary xv

1.0 Introduction 1

1.1. Project Statement of Work 3

1.2. 9th

Intake Roles and responsibilities 11

1.3. Design methodology (Design Manual) 14

2.0 Requirements capture 22

2.1. Aircraft design specification document divergence 23

2.2. Revision of mission requirement specification 29

3.0 Airframe Conceptual Design Phase 36

3.1 Airframe configuration submissions and design decisions 39

3.2 Airframe configuration down selection 76

4.0 Terrasoar wing design 84

4.1 Structural definition for wing design 85

4.2 Review of 5th

intake wing design 89

4.3 (A) Status wing layout and structural definition for PDR 93

4.4 (A) Status wing layout and structural definition post PDR 96


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4.5 (B) Status wing layout and structural definition for CDR 101

4.6 (C) Status revised CDR wing layout incorporating stressing 109

4.7 Detailed design for BAE Systems manufacture 112

5.0 Wing interface joint concept design 117

5.1 (A) Status PDR Interface Joint Trades 117

5.1.1 Post PDR (A) Status design study 117

5.1.2 (A) Status wing to empennage boom joints 118

5.1.3 (A) Status wing to fuselage interface joint 119

5.1.4 (A) Status inboard to outboard wing joints 120

5.2 (B) Status interface joint design studies 123

5.2.1 (B) Status wing to fuselage joint for CDR 124

5.2.2 (B) Status common wing and boom interface joint for CDR 124

5.3 (C) Status interface joint maturation for manufacture 127

6.0 Flight control surface and systems integration 129

6.1 (A) Status aileron attachment studies 129

6.2 (B) Status aileron attachment for CDR 131

6.3 (C) Status aileron attachment for manufacture 132

6.4 (A) Status aileron design studies 133

6.5 (B) Status aileron design 135

6.6 (C) Status aileron design for CDR 136

6.7 (C) Status FCS systems integration COTS 137

7.0 Materials and manufacturing methodology 141

7.1 Materials selection and aircraft weight 141


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7.2 Manufacturing methods and tool design 143

7.3 Structural test and qualification proposal 145

8.0 Conclusions 163

9.0 Further work required 163

10.0 References 164

Appendices:-

A-1: - Wing weight tables 166

A-2: - Major wing component isometric view sketch book

after signing off by the customer: structure engineering:

and manufacturing 170

A-3: - Post CDR Wing sign – off document 180

A-4: - Terrasoar wing major component dimensioned drawings 187


vii

Figures: - Page

1: - Intake 5 original Terrasoar design concept 2

2: - Intake 9 CDR Terrasoar design 3

3: - 9th

Intake GDP team mock-up inspection 4

4: - 9th

Intake GDP project framework 5

5: - 5th

Intake PDR aircraft launch proposal 8

6: - 5th

Intake PDR aircraft recovery proposal 8

7: - 9th

Intake Terrasoar Project Team Organisation Chart 13

8: - Terrasoar Starboard Outboard Wing structure Key datum model 20

9: - Key datum planes for composite and metallic details 20

10: - Terrasoar nosecone / payload bay cover 24

11: - BAE Systems Standard Bill of Materials 21

12: - Cambeltown Airport illustrating its layout and proximity to the sea 29

13: - Micropilot representation of the S1 mission flight plan 31



16: - Damian Adams proposed configuration for Terrasoar 40

17: - Chart of supporting notes for Damian Adams submission 40

18: - Mr Alan Barnes submission for an alternative Terrasoar 41

19: - Mr James Pennington‟s submission for an alternative Terrasoar 41

20: - Supporting data for Mr James Pennington‟s submission 42

21: - Mr Robert Currie‟s submission for an alternative Terrasoar 42

22: - Dimensioned drawing of Mr Robert Currie‟s submission 43


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23: - Supporting configuration notes for Mr Robert Currie‟s Submission 43

24: - The authors first alternative Terrasoar submission Configuration 1 44

25: - Aft isometric view of Configuration 1 highlighting key details 44

26: - Supporting key data and design notes for Configuration 1 45

27: - The authors second alternative Terrasoar submission Configuration 2 45


29: - Supporting key data and design notes for configuration 2 46

30: - The authors third alternative Terrasoar submission Configuration 3 47



33: - The authors fourth alternative Terrasoar submission Configuration 4 48



36: - The authors fifth alternative Terrasoar submission Configuration 5 50



39: - Overview design notes for all of the authors concepts 52

40: - Mr Craig Carr‟s submission a picture of the Pioneer UAV 51

41: - Mr Robert Sneddon‟s submission the Boeing Scan Eagle UAV 52

42: - Chart of configuration scoring based M1 module (Red) teams selection 53

43: - Raptor tractor propeller UAV used by NASA 57

44: - Predator pusher propeller UAV used by the USAF 57

45: - Major parameters of wing definitions 59


ix

46: - Induced drag factor as a function of taper ratio for different wing AR 60

47: - Example of geometric twist in a tapered wing 63

48: - Comparison of high-wing, mid-wing, and low wing configurations 64

49: - Characteristics of an airfoil section 66

50: - Airfoil forces and moments 69

51: - Airfoil centre of pressure 69

52: - Aerodynamic centre 1 70

53: - Aerodynamic centre 2 70

54: - NACA 4412 chart analysis 1 71



57: - Wartmann FX 63 – 137 airfoil 75

58: - Aircraft reference axes and corresponding aerodynamic moments 76

59: - Alternative Terrasoar configuration 6 (A) fwd quarter 77

60: - Alternative Terrasoar configuration 6 (A) aft quarter 77

61: - Author‟s supporting data for alternative Terrasoar configuration 6(A) 78

62: - Alternative Terrasoar configuration 7 isometric view 79

63: - Alternative Terrasoar configuration 7 external features 1 79

64: - Alternative Terrasoar configuration 7 external features 2 80

65: - Alternative Terrasoar configuration 7 fuselage internal features 81

66: - Alternative Terrasoar configuration 7 wing internal features 82

67: - Author‟s supporting data for alternative Terrasoar configuration 7 83


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68: - Sketch of the final down selected modified Terrasoar configuration 84

69: - Typical I spar constriction 86

70: - Typical Rib construction 87

71: - Typical Stringer stiffened and stressed skin wing layouts 87

72: - Major fixed aerodynamic surface airfoil selection (5th

Intake) 89

73: - Wing dihedral on outboard wing panels (5th

Intake) 89

74: - Aerodynamic twist in wing panels (5th

Intake) 90

75: - Major fixed aerodynamic surfaces (5th

Intake) 90

76: - Proposed wing manufacturing methodology (5th

Intake) 91

77: - Vehicle transportation breaks (5th

Intake) 91

78: - (A) Status PDR wing general arrangement model 93

79: - (A) Status PDR internal structure model showing key features 94

80: - (A) Status PDR internal structure of the outboard wing 95

81: - (A) Status PDR wing to fuselage location model 95

82: - (A) Status PDR wing / fuselage attachment rib 96

83: - (A) Status Post PDR Study wing G.A. model external configuration 99

84: - (A) Status Post PDR Study wing G.A. model internal configuration 99

85: - (A) Status Post PDR Study Outboard wing with skin construction detail 100

86: - (A) Status Post PDR Study Outboard wing with spar construction detail 100

87: - Interim Maturation wing external features 102

88: - Interim Maturation wing internal features 103

89: - Actuator integration concept 104

90: - (B) Status CDR Revision wing external configuration 105


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91: - (B) Status CDR Revision wing internal configuration 105

92: - (B) Status CDR Revision outboard wing external detail 106

93: - (B) Status CDR Revision outboard wing internal detail 106

94: - (B) Status CDR Revision aileron integration 107

95: - (C) Status Post CDR wing design with stress sizing at sign off 110

96: - (C) Status Post CDR centre wing skin at sign off 110

97: - (C) Status Post CDR centre wing structure at sign off 111

98: - (C) Status Post CDR outboard wing exterior at sign off 111

99: - (C) Status Post CDR outboard wing interior at sign off 112

100: - Manufacturing maturation of signed off centre wing structure 113

101: - Manufacturing maturation of signed off centre top wing skin 113

102: - Manufacturing maturation of signed off centre lower wing skin 114

103: - Manufacturing maturation of signed off outboard lower wing skin 114

104: - Manufacturing maturation of signed off outboard top wing skin 115

105: - Manufacturing maturation of signed off outboard wing structure 115

106: - (C) Status aileron structure 116

107: - (A) Status wing to empennage joint concept 2 118

108: - Post PDR (A) Status design study wing / fuselage joint concept 1 120

109: - (A) Status Inboard / Outboard wing joint (authors) concept 1 121

110: - (A) Status Inboard / Outboard wing joint (authors) concept 2 122

111: - (A) Status Inboard / Outboard wing joint (Peter Hamilton) concept 3 123

112: - (B) Status wing / fuselage joint presented at CDR (James Pennington) 124

113: - (B) Status common outboard wing / empennage joint (James Pennington) 125


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114: - (C) Status common outboard wing / empennage joint (author) 126

115: - The basic forces acting on an aircraft in flight 126

116: - (C) Status common wing joint weight reduction proposal (Pete Hamilton) 127

117: - Manufacturing maturation of the common interface joint 127

118: - Manufacturing maturation of the wing / fuselage joint 128

119: - Post PDR (A) Status design study wing / aileron interface 130

120: - General arrangement of a 2 hinge flap illustrating the floating hinge 130

121: - (B) Status wing / aileron interface presented at CDR 131

122: - Post CDR wing / aileron interface for manufacture 132

123: - Spigot support lug integration 132

124: - Pre PDR (A) Status David Baird first aileron concept 133

125: - Pre PDR (A) Status David Bird second aileron concept 133

126: - PDR (A) Status David Bird third aileron concept 134

127: - (B) Status Aileron design for Post PDR wing design studies 135

128: - (C) Status Aileron design for CDR 136

129: - Schematic of Terrasoar FSC 137

130: - FCS aileron wiring and connections in the centre wing interface 138

131: - FCS aileron wiring in the outboard wing 138

132: - Actuator wiring channels in the outboard wing 139

133: - Aileron actuator installation 139

134: - Aileron actuator drive mechanism 140

135: - Aileron horns and actuator drive rods 140

136: - First outboard wing skin re-stressing 141


xiii

137: - First inboard wing skin re-stressing 142

138: - Second outboard wing skin re-stressing 142

139: - Second inboard wing skin re-stressing 143

140: - Terrasoar wing centre section tool design 144

141: - Manufacturing spar changes 144

142: - Building block test plan 146

143: - Materials qualification testing 151

144: - Component testing 153

145: - Tension patch full scale wing testing 160

146: - Loading frame full scale wing testing 161


xiv

Tables: - Page

1: - The disparity between the Projected and Actual Project Timescales 10

2: - Documentation requirements for the S1 Mission 31



5: - Configuration decision making summary table 54

6: - Stress data for the Terrasoar post CDR wing using raped skin 109

7: - Effect of FVP on mechanical properties of test laminates 148

8: - ASTM Shear coupon composite shear tests 148

9: - CDR wing submission wing weight table based on Phenolic fibreglass and

R63.80 foam core and Styrofoam and Al 2024-T351 densities 166

10: - Post CDR sign - off wing redesign weight table based on revised fibreglass

and R63.80 foam core and Tricast 6 and Al 2024-T351 densities 168


xv

Glossary: -

AAC Air Combat Command

ADR Air Data Relay

AFB Air Force Base

AIAA American Institute of Aeronautics and Astronautics

AT Aerial Target

AV Air Vehicle

ATDC Advanced Technology Demonstration Centre

BDA Battle Damage Assessment

BM Bending Moment

CAD Computer Aided Design

CATA Control, Automation and Task Allocation

CCD Charged Couple Device

CDL Common Data Link

C of G Centre of Gravity

CFC Carbon Fibre Composite

CFD Computational Fluid Dynamics

DARO Defence Airborne Reconnaissance Office

DARPA Defence Advance Research Projects Agency

DERA Defence Evaluation and Research Agency

D o D Department of Defence

EO Electro - Optical

ERAST Environmental Research Aircraft and Sensor Technology


xvi

EW Electronic Warfare

ESDU Engineering Science Data Unit

FLIR Forward Looking Infrared

FEA Finite Element Analysis

G.A. General Arrangement (Drawing or Model)

GA General Atomics Aeronautical Systems

GCS Ground Control Station

GDP Group Design Project

GDT Ground Data Terminal

GPS Global Positioning System

GRP Glass Fibre Reinforced Plastics

GSE Ground Support Equipment

HALE High Altitude Long Endurance (= or > 30,000ft)

IFF Information Friend or Foe

INS Inertial Navigation System

IML Inner Mould Lines (internal surface limits)

LAD Landing Assist Device

LASS Low Altitude Surveillance System

LOS Line Of Sight

MAE Medium Altitude Endurance

MALE Medium Altitude Long Endurance (< 30,000ft)

NACA National Advisory Committee on Aviation

NASA National Aeronautics and Space Administration


xvii

NOLO No Live Operator

OML Outer Mould Lines (external surface limits)

OTH Over The Horizon

RAF Royal Air Force

RPV Remote Piloted Vehicle

RN Royal Navy

SEAD Suppression of Enemy Air Defences

SFC Specific Fuel Consumption

UAV Unmanned Air Vehicle

UCAV Unmanned Combat Air Vehicle

UHF Ultra High Frequency

URAV Unmanned Reconnaissance Aerial Vehicle

USAF United States Air Force

USCG United States Coast Guard


1

1: - Introduction.

The UAV or Unmanned Aerial Vehicle is an aircraft that is specifically designed to fly an

entire mission profile in the same way as a manned operational aircraft would i.e. take-off

fly to somewhere to perform a specific task, return to base, and land, with the exception

of the pilot being aboard. These aircraft were known as remotely piloted vehicles

(RPV‟s) until the 1980‟s, because they were primarily directed by an external source of

control either on the ground or in an accompanying aircraft. However many modern

vehicles are no longer remotely controlled and are pre-programmed to operate

autonomously and the development of an operational UAV is the subject of this thesis.

While UAV‟s are not designed to be expendable meaning that the end user services do

not intend to lose them every time they are sent on a mission, they are, in the terminology

of the United States of America‟s Department of Defence “attritable”. This means that a

commander can afford to lose one through attrition, especially when the alternative is the

loss of a manned aircraft or an aircrew. Although UAV‟s do not put pilots lives at risk in

uncontrolled airspace they have the potential to put manned aircraft from light aircraft to

airliners and populated area at risk from air to air collision or crashing, therefore very few

are FAA or CAA certified to fly outside closed government ranges and testing

establishments. Also because the UAV is unmanned there is a misconception that they

are cheap (mostly founded on the belief that they are like the cruise missile and

expendable), although generally they are less expensive than manned aircraft and the cost

of fully developed UAV‟s capable of performing useful civil and military missions in the

first decade of the twenty first century ranges from around a half – million dollars for a

medium utility aircraft to nearly forty million for something with the capabilities of

Global Hawk (a fully FAA certified long range HALE UAV). The payloads, sensors,

airframes, and control and communication networks that are combined to provide high

quality capabilities are of themselves not inexpensive.

The types of mission for which UAV‟s and UCAV‟s (Unmanned Combat Air Vehicles)

are best suited are: - (1) Dull repetitive reconnaissance missions such as coastal patrol, or

boarder security: military and civilian long range reconnaissance such as target location

or disaster area searches, or oil spill monitoring, and air quality and resource surveys: (2)

Dirty investigations of areas contaminated with biological or chemical weapons which in

this time global terrorism could be a city as easily as a battlefield, as well as nuclear

contamination for example nuclear power stations and reprocessing facilities: (3)

Dangerous missions which are mostly military such as the suppression of enemy air

defences, and deep penetration tactical conventional and nuclear strike missions,

equivalent high risk civil missions could be hurricane penetration flights and volcano

science missions. To these current and near term missions could be close support of

combat troops, and police surveillance (replacing police helicopters and the Optica light

aircraft.


2

The role of the UAV selected for the 9th

Intake GDP (Group Design Project) initially fell

into the first category i.e. general reconnaissance and survey, but to ensure funded

manufacture of the aircraft by BAE Systems, and an actual flight worthy product the

Terrasoar UAV became a vehicle to demonstrate new manufacturing processes, and small

aircraft design philosophies, as well as validation of the Micropilot flight control system.

As such this aircraft should be viewed as a low risk concept demonstration aircraft, and

not as a new high risk airframe configuration, and the OML (Outer Mould Line)

configuration selection is covered in depth in section three of this thesis.

The original intention stated at the 9th

Intake Group Design Project launch meeting of

Monday 9th

June 2003, was to review the Intake 5 original Terrasoar design to determine

this concepts feasibility from the data set they presented at PDR (Preliminary Design

Review) which was the end point in their Group Design Project. Introducing

modifications to this design for manufacture, and proceed with detail design manufacture,

and assembly leading to a first flight by the end of the 9th

Intake MSc course.

However on examination of the 5th

Intakes Terrasoar configuration, structural layout,

manufacturing methodology, and launch and recovery procedure as well as their choice

of flight control system and systems integration, little was found to be acceptable, and the

decision was made by the 9th

Intake team to start over with a fresh design incorporating

the engine which the 5th

Intake had already purchased.

Although the 5th

Intake was working on the philosophy of designing a real aircraft for

flight test and an ideal aircraft to full-fill the actual mission requirements based on the

real aircraft, the 9th

intake had grave doubts if either aircraft could undertake realistic

flight operations, in the as designed configuration. For comparison the 5th

intake aircraft

and the 9th

intake aircraft are shown in figures 1 and 2 respectively. From these figures a

superficial resemblance can be inferred i.e. both were rear engine aircraft with twin

booms, however in terms of structural layout, sizing, materials, and manufacturing

methods they were completely different.

The initial intention of both Cranfield University and BAE Systems was to seek full CAA

Certification to fly the aircraft in uncontrolled airspace, in addition to designing and

building the Terrasoar UAV to fly on Category B certification within sanitised airspace,

and this aspect of the Group Design Project is covered within reference 1:- Mr Alan

Robert Barnes thesis: - UAV Configuration Management, Certification / Qualification &

Control of Mass / C of G / Inertia. To this end the Terrasoar UAV was designed to meet

as closely as practical the requirements of JAR-VLA, the compliance of the final aircraft

in respect to these requirements id covered within reference 2:- Miss Emma Bradley

thesis: - Terrasoar Unmanned Aerial Vehicle Group Design Project Thesis. From Mr

Barnes thesis it will be seen that the requirements for full CAA certification could not be

met within the financial constraints of the Terrasoar budget and the current aircraft seeks

Category B certification only at the time of writing.


3

Figure 1:- The 5th

Intake Terrasoar PDR design. (Reference 3:- Preliminary Design

Review (PDR) for “Terrasoar” High Altitude Long Endurance Unmanned Air

Vehicle: Cranfield University College of Aeronautics: Wednesday 7th

February 2001)

Figure 2:- The 9th

Intake Terrasoar CDR Design submission. (Reference 4:- Authors

private collection).


4

1.1: - Project Statement of Work.

The 9th

Intake Group Design Project Launch meeting was held on the 9th

June 2003 at

BAE Systems Samlesbury in building S210 (Man Tech) and attended by representatives

from BAE Systems Air Systems and Aerostructures: Cranfield University : and all of the

9th

Intake with the exception of Miss Emma Bradley.

The aims of the launch meeting were to: - introducing the 5th

Intake configuration work

on the Terrasoar aircraft and the level of maturity which they had achieved (preliminary

design phase being their exit point from the GDP): identifying the customers for the

aircraft: possible applications: the design requirements: and the projects framework:

concluding with a mock up review, shown in figure 3.

Although initially the Cranfield University as the final joint customer with BAE Systems

of the Terrasoar UAV expressed great enthusiasm for retaining the 5th

Intakes

configuration as the basis for detail design and manufacture, the 9th

Intake design team

had some grave reservations about the preliminary design work undertaken to date which

was exasperated when the mock up was examined outside the meeting building.

Figure 3:- The 9th

Intake Terrasoar team with the customer representatives inspect

the mock up of the 5th

Intakes preliminary design concept at BAE Systems

Samlesbury during the GDP launch meeting of the 9th

June 2003 note no engine or

propeller was fitted to the mock up and no dummy systems were installed.


5

Building on Cranfield University‟s enthusiasm for the previous intakes configuration the

GDP Framework was formally set out in document AVT-0215 in detail and followed the

established process phases of concept demonstration aircraft design and manufacture

followed within the aerospace industrial sector. This is shown as a flow chart in figure 4

and detailed below.

Figure 4: - The 9th

Intake Group Design Project original framework for the design

and manufacture of the Terrasoar scaled technology demonstrator for a High

Altitude Long Endurance (HALE) Unmanned Air Vehicle as defined in reference 5:-

AVT-0215.

Phase 1:- Concept Confirmation: - The objective of this phase was to review the 5th

Intakes preliminary design concept in order to determine the feasibility of carrying this

concept in the as designed condition to the detail design phases of the project frame work,

and to justify any reason why this could not be accomplished by the 9th

Intake team.

Some major review areas were identified in the Launch meeting document AVT-0215

and are summarised below, also as stated above the 9th

Intake had several key issues of

concern in addition to the areas identified in AVT-0215, which are also presented below.

CONCEPT CONFORMATION

(PHASE 1)

PRELIMINARY DESIGN

(PHASE 2)

DETAIL DESIGN

(PHASE 3)

MANUFACTURING AND SYSTEMS INTEGRATION

(PHASE 4)

GROUND TEST

(PHASE 5) FLIGHT TEST

(PHASE 6)

MSc COMPLETION


6

1. Methods of take off and recovery (landing) the methods proposed by the 5th

Intake were not deemed practical on terms of safety for the ground handling

personnel in either: - deploying: launching: or recovering the aircraft. Also these

proposals were seen as an over complication of the aircrafts support infrastructure

as well as the aircraft systems, and furthermore the 9th

Intake team had serious

doubts that if the aircraft could get airborne using these proposals it could be

recovered without sever structural damage, probably resulting in the loss of the

aircraft.

2. The airworthiness of the 5th

Intakes design was reviewed and found to be

unsuitable for the mission requirements which are captured in section 2 and

although the aerodynamics of original configuration may well have been suitable

for the 40,000ft altitude requirement for the 5th

Intakes GDP it represented a point

design and exactly how the original design was intended to achieve this region of

operation was unclear. This aspect of the project is detailed within reference 1:-

Mr Alan Robert Barnes thesis: - UAV Configuration Management, Certification /

Qualification & Control of Mass / C of G / Inertia.

3. Rationalisation of the 5th

Intakes airframe to ease manufacture, and reduce cost of

the vehicle. Because the lack of maturity and the unsuitability of the 5th

Intakes

design to meet the requirements set-out in section 2 this work package became a

complete redesign of the aircraft with the pre-purchased engine being and the

name being the only common features. The redesign of the wing is subject of this

thesis, with the fuselage redesign is covered in reference 6:- Mr Paul Francis

Gilligan thesis: - Fuselage Design and Integration for a High Altitude Long

Endurance Aireal Vehicle, and the empennage redesign and aircraft aerodynamics

are covered within reference 7:- the thesis of Mr James Pennington Terrasoar

aerodynamic configuration, performance and design.

4. Removal of the 5th

Intakes proposed gearbox as well as the propeller indexing

system and replacement with a toothed belt drive. All of the powerplant and

propeller issues are covered in reference 8:- Mr Damian Adams thesis: -

Terrasoar Chief Engineer, Powerplant and Manufacturing-Group Design Project.

5. Simplification of the flight control system which involved substitution of a

modern and more appropriate avionics package than the legacy „XREA‟ avionics

crate, which was the original choice for the 5th

Intakes design because of its

availability within Cranfield University being owned by the Flight Dynamics

Group. After extensive research the FCS selected for the 9th

Intakes aircraft was

the Micropilot system which was a low cost but very capable FCS and is detailed

with representative flight test data in reference 9:- Mr Robert Currie thesis: -

Development of a Cost Effective UAV Avionics and Flight Control System.


7

6. Examination of possible payloads and the design requirements for these payloads.

The payload provision for the aircraft was set as 5kg and to be inclusive of power

and / or cooling requirements, the payload required have a separate data channel

and not to generate any electrical interference with the flight control system, and

therefore a self contained and RF / electrically isolated payload was required.

Several investigations into possible payloads were conducted by both Mr David

Baird and Mrs Rachel Cunliffe which are fully documented in reference 10:- Miss

Rachel Cunliffe thesis: - Terrasoar UAV Payload, Stability and Flutter Group

Project.

7. Possible use of new technology in areas defined by the BAE SYSTEMS –

Cranfield University namely: - Low cost high altitude payload vehicles: Novel

secondary power systems: Novel flight control systems and control surfaces: High

reliability, zero maintenance: and Novel manufacturing methods. Of these the

design of a low cost UAV and the use of novel manufacturing methods were at

the forefront of the 9th

Intakes Terrasoar aircraft design.

8. Aircraft major component sizing for the revised mission requirements in terms of

wing, empennage, and propeller sizing. As stated above the 5th

Intakes design was

sized for a mission to 40,000ft (with a different engine), and not for the 10,000ft

mission the 9th

Intake was charged with, therefore factors such as the aspect ratio

were in appropriate for the mission as was the propeller sizing, the control

surfaces were considered to be ineffective both in terms of size and location with

no redundancy, and the empennage was considered to be equally in effective.

9. Employment of either a retractable or fixed undercarriage was considered

essential after reviewing the 5th

Intakes proposed launch and recovery systems

shown in figures 5 and 6 respectively and this area is covered in reference 11:-

Mr Robert Sneddon‟s thesis (yet to be titled), as well as references 6 and 7.

10. Airworthiness clearance to JAR-VLA and the appropriateness of these

requirements, as well as the cost of certification of the aircraft to fly within

unrestricted airspace which is covered within reference 1.

11. The interface / transportation joints between the major components namely: -

centre wing to fuselage: centre wing to outboard wing: wing to empennage

booms: all of which are covered within this thesis.

12. Systems integration for the aileron actuators: rudders: elevator actuators: and

navigation lights which are covered within this thesis, and references 6 and 7.

This resulted in the Concept Confirmation Phase being widened to include completely

new aircraft configurations which are captured in this thesis and this phase became

Aircraft Conceptual Design.


8

Figure 5: - The 5th

Intake Terrasoar PDR design Vehicle launch proposal.

(Reference 3)

Figure 6:- The 5th

Intake Terrasoar PDR design Vehicle recovery proposal.

(Reference 3)


9

Phase 2:- Preliminary design: - The objective of this phase was to produce an outline

concept design which could meet the Terrasoar UAV specification at the lowest cost

while meeting all of the functionality: flight and ground safety: reliability:

maintainability: and manufacturability targets set by the customer. To achieve this level

of maturity all major risks had to be identified and reduced to an acceptable level, but as

will be seen in later sections maturation weight gain and manufacturing costs were to

play key roles design decisions making even after the Critical Design Review this is a

situation all too common in both civil and military aircraft projects and even small

aircraft such as ours was not exempt.

From the start of the project the key decision was made to use 3D CATIA solid and

surface models rather than sets of paper drawings and lofts to communicate design intent

because these could be readily interrogated by the prime airframe contractors the 9th

Intake considered capable of producing this aircraft, and modified much easer than paper

drawings. Also 2D drawings could be readily extracted from the 3D model as and when

required. This decision has enabled the design to progress to the level of maturity

presented within this thesis which would not have been possible with manual drawings,

although because this aircraft required flight certification only licensed CATIA V4 could

be used within BAE Systems which has lead to a heavy work load on the only two

qualified design engineers within this team namely myself and Mr James Pennington, and

the one manufacturing engineer we trained to use CATIA V4 namely Mr Paul Gilligan.

The exit criteria for the preliminary design phase was the Preliminary Design Review

(PDR) at which the final Outer Mould Lines (OML) were frozen for the aircraft and the

basic airframe assembly philosophy was defined, and the majority of the major problem

areas had been resolved to a point where detail design could be undertaken and this phase

was completed on the 3rd

March 2004.

Phase 3:- Detail design: - The objective of this phase was to produce the matured detail

design, assembly methodology, systems installation models, and manufacturing models,

of a standard for release to manufacture. However although the design layout was mature

the detail stressing of the airframe was delivered late and consequently no models could

be released except for OML tooling could be released to manufacture. Although in the

unstressed condition the aircraft was considered capable of complying with the

specification and design requirements. At the Critical Design Review of the 21st April

2005 the aircraft was judged by Cranfield and BAE Systems sufficiently matured to

proceed to an additional final design for manufacturing phase when final stress data

became available on the 23rd

May 2005, although this design phase proved to be more

protracted than originally foreseen due to weight growth, materials and manufacturing

methodology changes, and stress data revisions and is only now coming to a close in

November 2005.


10

Phase 4:- Manufacture and systems integration: - This phase is scheduled to start in

February 2006 although initial tool design is currently taking place at BAE Systems Man

Tech facility, and the manufacture of all metallic details have been approved for

manufacture with BAE Systems.

Phase 5:- Aircraft Ground Testing: - Planning for this phase has been completed and

can be reviewed in references 1, 2, 8, and 9, also the engine has been ground tested to

determine the specific fuel consumption figures for the engine to enable accurate fuel

tank sizing prior to CDR and this is covered in references 2 and 8.

Phase 6:- Aircraft Flight Testing: - Planning for this phase has been completed and is

covered in reference 1, also the Micropilot FCS has been successfully flown on several

simulated missions in a model aircraft at BAE Systems Samlesbury airfield and these

qualification flights are covered in reference 9.

Table 1:- The disparity between the Projected and Actual Project Timescales.

Task / Phase Projected timescale. Actual timescale.

Project Launch Meeting. 9th

June 2003 9th

June 2003

1st GDP Meeting (project

handover from 5th

Intake). 25

th June 2003 25

th June 2003

Phase 1:- Concept

Conformation. 17

th September 2003 4

th September 2003

Phase 2:- Preliminary

Design Review. 11

th February 2004 3

rd March 2004

Student Design Week 31st May to 4

th June 2004 31

st May to 4

th June 2004

Phase 3:- Critical Design

Review. 13

th October 2004 21

st April 2005

Manufacturing and Systems

Integration Completed. End of April 2005

Tool design started in

November 2005

Ground Test Phase

Completed. End of July 2005

TBD:- Provisional May

2006

Flight Readiness Review. September 2005 TBD:- Provisional June

2006

First Flight. October 2005 TBD:- Provisional July 2006

GDP Thesis Hand in date 30th

November 2005 30th

November 2005


11

From Table 1 above it can be seen that there was a marked disparity between the

projected timescales and the actual (achieved) timescales of the Terrasoar project and

they strongly diverged after the Preliminary Design review. This was due to four primary

reasons which were: -

(1) The projected time scales were based on Cranfield University‟s initial assumption that

the 5th

Intakes concept would require relatively little modification and had reached a level

of maturity close to that required for detail design to commence, this however was not a

view shared by the 9th

Intake or BAE Systems.

(2) The volume of work required to produce a design suitable for manufacture was to

great for a team of three designers working in their own time without the commitment for

hours or support of their line management, and the compelling commitments to the BAE

Systems flagship project JSF / F-35 which itself is under resourced in design and stress

disciplines with the design team on Terrasoar working a 7:00am to 7:00pm day on JSF

not much “free time” could be devoted to the detail design phase.

(3) The need to use only licensed CAD software i.e. CATIA V4 which was only available

within BAE Systems, this was paramount for aircraft certification.

(4) The late issuing of stress data due to the full-time work commitments of the 9th

Intake

stessman, and the use of novel airframe materials.

Therefore the projected phase completion dates could not be met and should be

considered as over ambitious. This concludes discussion of the statement of work as

issued in the 9th

Intake GDP Launch Presentation document AVT-0215.

1.2: - 9th

Intake Roles and Responsibilities.

After completion of Phase 1 Conceptual Design (detailed in section 3 of this thesis) as it

was re-designated after the scope of this phase was widened as described above, the 9th

Intake selected roles and responsibilities within the Terrasoar Project Team based on their

normal full – time working functional discipline e.g. design, stress, manufacturing, etc or

an area in which they would had specialist expertise outside the daily work environment

such as large scale model building and flight operations, or even an area they wished to

explore e.g. aerodynamics, or payload integration. The 9th

Intake was unusual in that for

the first time two streams of the course were run concurrently of which one was the

engineering stream and other was the manufacturing stream. This lead to a greater degree

of detailed manufacturing planning, and early process definition than in the 5th

Intake,

however the number of designers remained small i.e. two compared with previous

intakes, although one manufacturing engineer Mr Paul Gilligan undertook the major

design role as the fuselage designer and fuselage systems integrator much effort was

required on his part to learn the skills set required to use CATIA V4 and myself and Mr

James Pennington took time form our design maturations to train him fully in the design

toolset.


12

The other manufacturing engineers who helped the design effort Mr Damian Adams who

provided detail design support in CATIA V5 to Mr James Pennington and Mr Paul

Gilligan in respect of the engine mounting frame, and Mr David Baird who supported my

wing design with conceptual aileron designs based on sizing generated by James

Pennington and Miss Rachel Cunliffe. Additional detail design support was given by Mr

Alan Barnes for empennage / boom interface and Mr Robert Currie for the avionics tray

and from Mr Vernon Hind for the full fuel tank design and the fuel system for which he

was responsible. No other members of the team offered any design support although Mr

Robert Sneddon was specifically charged with undercarriage design in addition to his

primary stress role no work in this area was forthcoming and the design of this system

was undertaken by James Pennington and Paul Gilligan.

Figure 7 shows the 9th

Intake Organisational structure as of November 2005, although the

original intake was as follows:-

Paul Gilligan James Pennington Bob Currie

Damian Adams Alan Barnes Dave Baird

Rob Sneddon Vernon Hind Craig Carr

Jon Baggaley Rachael Cunliffe Rob Cunliffe

Emma Bradley.

Where Craig Carr had the role of airworthiness certification and materials selection but

left the course in March 2004: and John Baggaley had the role of control surface and

structures design support but left the course in January 2005 and the loss of both of these

talented individuals was a blow to the team. Although Craig Carr has been a great help to

the team in his new as a senior engineering specialist within the Man Tech organisation

as the design of Terrasoar has neared manufacturing design readiness. The developments

of organisational structures and inter - team dynamics is covered within reference 8 and

will not be covered within this thesis, and in my view the team have done the best they

can within the timeframe available.

The design challenges of the Terrasoar were undertaken by the Airframe Design Team,

which I was selected to lead as the Airframe Lead shown in figure 7, with responsibilities

for: - co - ordination of all of the airframe design and systems integration activities,

project planning and scheduling, and sole design responsibility for the wing, and wing

flight control surface and systems integration, as well as the wing mounted navigation

lights.


13

Fig

ure

7:-

9th

IN

TA

KE

TE

RR

AS

OA

R P

RO

JE

CT

TE

AM

OR

GA

NIS

AT

ION

CH

AR

T.


14

As Airframe Lead my priorities were to provide oversight of the design activities and out

brief these to the Chief Engineer, and the rest of the Terrasoar project team for their bye

in and discussion, identification of design information requirements from structures,

aerodynamics, systems, propulsion and manufacturing. Provide project planning, and

identification of design milestones, design training, and design procedures

documentation, for configuration control.

As Wing Chief Designer my priorities were to develop a wing concept which satisfies the

following requirements: - transportation breakdown into sections, load transfer through a

multi section structure, systems integration and access, mate joint philosophy, and

manufacturing philosophy. To this end I have focused my seven years of aircraft airframe

design experience to produce an engineering solution which meets the Terrasoar design

requirements. Additionally I also developed a wing test programme based on my four

years as a Senior Structural Test Engineer at the BAE Systems Structural Test Facility

Brough.

My first task as Airframe lead was to establish a design manual for the Terrasoar aircraft

so that all design activities were conducted to the same standard and design intent could

be understood by manufacturing and assembly personnel as well as the customer and

certification and airworthiness authorities. Because this aircraft had to meet real - world

airworthiness and certification requirements, the basis used for the design manual was the

BAE Systems design standards. This manual is covered in the next section to demonstrate

the level of control afforded to this design.

1.3:- Terrasoar Design methodology (Design Manual).

1. Introduction.

This Design Manual is intended only for use on the Terrasoar project and serves as

the reference for designers to the BAE Systems Technical Standards Manuals, and

CATIA design procedures within BAE Systems produced and maintained by DPA,

and accessed through the BAE Systems Intranet /Airframe Engineering / Design /

User Guides / Approved Methods.

This Design Manual states how the design schemes will be produced and co-

ordinated in a systematic methodology in accordance with Technical Standards

Manual 01.04.27. All of the methods covered within satisfy the requirements of ISO

9001, JAR, Defence Standards 00-970 and 05-123, for the production and

management of Design Schemes.

The procedures laid out and referenced in this manual will apply to all significant

design tasks (except for Repair Schemes) and its application has been agreed by the

Airframe Design Lead and the Terrasoar IPT Chief Engineer.


15

2. Design scheme control philosophy.

The Terrasoar Airframe is broken down into the following major airframe

components which are designed and controlled by the processes and procedures

covered by this manual. These major airframe components are listed below:-

Fuselage: - This includes the internal structure; skins; internal and external

systems integration; powerplant integration; fuel system integration, and

recovery system integration.

Wing: - This includes the internal structure; skins; flight control surface design;

and systems integration.

Tail Booms and Empennage: - This includes boom and empennage structure;

skins; control surface design; and systems integration.

These major airframe components will use a common axis system known as the

AIRCRAFT AXIS all component axes will be defined relative to this aircraft

axis and no component designer will disregard this rule.

Key Datum Models: - Will record the location of all structural elements of each

component and will be defined at component level and integrated with the

Aircraft Axis, (an example of the starboard wing substructure Key Datum

Model is shown in figure 8 and datum planes in figure 9). The primary function

of these models is to enable analysis of the structure of the aircraft and

determine the load paths, and interfaces between components.

Assembly Model: - This is a master model into which the component models are

located. This model demonstrates the maturity and integrity of the aircraft

structural layout; systems integration; fastener layout; and skins design. This

gives a timely indication of fowling / clash detection. Each component would in

an aircraft of high complexity and substance have an individual assembly model

but this is not deemed necessary for a modest UAV.

3. Design scheming.

All design schemes will be formally numbered in accordance with the Terrasoar

project rules with numbers issued by the Configuration Control Authority,

reference 1, using the relevant title block.

Schemes status will be identified adjacent to the title block prior to release outside the

Airframe Team responsible for the scheme‟s creation.


16

An Item Release Schedule (IRS) will be prepared and issued, which will detail:-

The scheme numbers

Brief description of content

A, B and C Status dates in accordance with the Design Programmes

Management Procedures TSM 01.09.15.

There are three levels of scheming:-

A Status: - Design concept / feasibility.

B Status: - Definition of basic structure / system and other specialist

requirements (Preliminary Design Review).

C Status: - Comprehensively engineered and sized drawing depicting the

„frozen‟ agreed configuration standard for the product (Critical Design

Review). The C status scheme will contain the following:

1. Pictorial views or solid models of the assembly with detail views and balloon

references including the components with sufficient dimensional data to

enable engineering drawing to commence and stress weights and production

engineering requirements to be satisfied.

2. Jigging hole positions.

3. Control dimensions, datum‟s and ICY requirements identified.

4. Special notes / reminders.

5. A Bill of Materials (standard parts / equipment) material specifications and

used on drawing numbers.

6. A „Circulation Box‟ with names of functional representatives working on the

changed area.

7. Change effect statement in the title block (according to project needs).

8. The detail parts which will NOT require a Structures Team signature on the

manufacturing drawings prior to production will be identified using the flag

note system relevant to the project. This will be indicated on the part number

on the face of the drawing and in the Bill of Materials. The note itself WILL

reference the Design Review Meeting Minutes at which endorsement of this

status was given.


17

Detail manufacturing drawings indicated thus will be identical to the „C‟ status scheme

with NO DEVIATION allowed.

All detail manufacturing drawings created will reference the „C‟ status scheme.

Frozen „C‟ status schemes once agreed cannot be changed without recourse to the formal

change control procedure.

The „C‟ status scheme forms the basis for the creation of Engineering / Manufacture

drawings.

It should be noted that the „C‟ status scheme can be used by Departments downstream of

the Airframe Team to carry out advanced tooling, preliminary planning activities etc., in

advance of the receipt of manufacturing drawings. It is NOT an authority to proceed

with manufacture. Any advance manufacturing activity from „C‟ status schemes

will be strictly „at risk‟ of scrap.

Note 1: Guidelines on scheme content to be used in the preparation of each status of

scheme are given in APPENDIX 1 of TSM 01.04.27 (Volume 01 – General Design

Procedures).

Note 2: At B Status consideration may be given to advance ordering of material, long

lead items and major components in conjunction with Production Engineering.

Note 3: At C Status, advance ordering requirements may be specified in accordance with

TSM 01.09.08 (Engineering Advance Material Release) project rules, and Design Visual

Examination (DVE) in accordance with TSM 01.09.16.

The use of A and B status schemes within this project will give a design history and

allow assessment of the level of maturity each component has reached over the span of

the design activities. It is mandatory that the final scheme status is always C for all

project design tasks.

Initial issue and each raise of issue of a scheme will be by sequential numbering, as per

normal practice. However, any changes of status requires a new scheme number and a

raise of issue of the previous status scheme to record cross – reference to new scheme

number. Each issue will be controlled by the relevant site documentation which will act

as the signature collection document for the scheme in accordance with normal Drawing

approval procedures. The equivalent site documentation is: -


18

Design Change Note: (DCN) – Warton / Samlesbury.

Design Department Instruction: (DDI) – Brough.

The DCN / DDI will contain a written statement of changes to be made to the scheme

from the previous issue. DCN / DDI numbers will be taken from the project register /

scheme control sheet (Sheet 1 normally).

Circulation at the initial issue must be specified above the scheme title block. All

subsequent issues will receive the same circulation. Changes in circulation are permitted

providing they are recorded in the scheme.

In considering the circulation of schemes, selection should be from the following:

Technical – Relevant departments affected:

Cost Engineering:

Process Engineering:

Manufacturing, including Jig & Tool Design and NC Programming:

Other BAE Systems sites (checkers):

Subcontractors when applicable.

To avoid delay or non – availability of schemes whilst being passed through the normal

issuing system, printing and distribution will be permitted ONLY WITH: -

PRELIMINARY DRAFT clearly written on the drawing boarder or tagged to a solid

model. Final schemes and all DCN / DDI‟s issued must be filed in the Terrasoar Design /

Model Store and stored electronically by Configuration Control PDM system, until

project hand over in 2006.

All Design / Stress programmes and Drawing release schedules for significant design

tasks will carry dates for A, B, and C status schemes as appropriate.

The Bill of Materials will be of the Manual type based on Figure 10, but will initially be

based on an excel spreadsheet owned by configuration control this will then be

transposed onto the PDM system and will use the same Eurofighter format as on the pre –

printed detail sheets.

The following Technical Standards Manuals will be followed for the production all

design schemes and engineering drawings for the Terrasoar Project: -

TSM 01 General Design Procedures.

TSM 08 CATIA Procedures.

TSM 12 Lofting Procedures.

TSM 13 Structures Design Procedures.


19

TSM 14 Mechanical Systems Design Procedures.

TSM 15 Electrical Bonding / Lightning Strike / EMP Design Procedures.

TSM 16 Electrical Installation Design Procedures.

TSM 19 Design Quality Procedures.

TSM 42 Metallic Material Processing Standards.

The following BAE Systems CATIA user guides will be used for all Terrasoar CATIA

design activities: -

BAE – WDO – GDE – GEN – DPA – 150832 General Solid Modelling.

BAE – WDO – GDE – GEN – DPA – 150827 Modelling Machined Parts.

BAE – WDO – GDE – GEN – DPA – 150826 Modelling Sheet Metal Parts.

BAE – WDO – GDE – GEN – DPA – 150829 Modelling Carbon Fibre Composite

Parts.

BAE – WDO – GDE – GEN – DPA – 150885 Interference Analyses.

BAE – WDO – GDE – GEN – DPA – 150896 Drawing Productions.

BAE – WDO – GDE – GEN – DPA – 150833 CATIA Database Assemblies

Scheming.

4 Scheme Approvals.

All schemes will be fully checked before final issue, and final approval of a scheme for

issue will be an acknowledgment that due consideration has been given to: -

The design is fit for purpose.

All reasonable design paths have been explored.

All relevant factors have been considered.

The design can be cost effectively produced, inspected, tested, installed, operated

and maintained in a satisfactory manner.

There is adequate documentation to support the design.

Target masses have been achieved („C‟ Status).


20

Figure 8: - Terrasoar Starboard Outboard Wing primary structure Key datum

model.

Figure 9:- Key datum planes for: - Composite (green) and Metallic (blue) details.

(A) Composite. (B) Single sided

metallic. (C) Double sided

metallic.

LEADING EDGE SPAR

DATUM X= -1144mm

TRAILING EDGE SPAR

DATUM X= -1302.7mm

OUTBOARD INTERFACE JOINT RIB DATUM Y= -563.6mm

TIP RIB DATUM Y= -1810mm

KEY DATUM PLANES


21

Figure 10:- Standard Bill of Materials. TSM 01.04.27.(Note after to prevent

confusion with other projects on the Eurofighter PDM system the Terrasoar Project

aircraft B o M was compiled as an excel spreadsheet and maintained by design and

controlled by configuration management).


22

Note: Schemes will only be submitted to functional specialists for technical approval

after they have been checked by the Airframe Team they will be reviewed by the

customer and signed off because for this project only Cranfield University is the design

authority. Scheme and models that do not carry a checking signature or are not supported

by the sign off document must be returned to design.

Functional specialists will not carry out a check on drawing quality. If spotted however,

any errors are to be notified to the Airframe Team for correction.

2: - Requirements capture.

The objective of the Terrasoar project was to produce a single engine pusher propeller

high wing monoplane uninhabited air vehicle (UAV), which is capable of flying at

10,000ft for a duration of 5 hours, with a payload of 5kg‟s in unrestricted airspace.

The GTOW of the aircraft was not to exceed 40kg‟s which impacted on both the

selection of a structural design philosophy for the aircraft, the materials to be used, and

the manufacturing processes employed to produce the airframe components.

The requirement for the airframe to be dismantled for transportation to and from the

flight test location and rapidly reassembled for the flight trials, with subsequent rapid

brake down for transport after trials, and the dimensions of the doors, with the need to

ensure adequate clearance for extraction of the aircraft from the Ford transit transport

vehicle dictated the size of major airframe components. Therefore types of structural

mate joint for the airframe components needed to be robust and easy to assemble.

The need to embody and integrate multiplex redundant systems which could be accessed

prior to flight, made provision of quick release access panels which could then be sealed

for flight a priority and posed significant challenges for the wing, fuselage design, and

empennage design and for the wing this is covered within this thesis.

The key parameters of the 5th

Intakes concept were presented within AVT-0215 and

formed the starting point for the 9th

Intakes design studies for the revised Terrasoar, and a

brief specification of this vehicle (which resulted from the 5th

intake work) is as follows: -

MTOW 40 Kg

Maximum Altitude 10000ft

Payload Capacity 5 Kg

Wing Span 5.5m

Wing Area 1.21m2

Aspect Ratio 25

Outer Wing Dihedral 5 deg


23

Mission Duration Approx 5 hours with 1 hour loiter at max altitude

Powerplant MOKI 215cc 4 stroke 5 cylinder Radial engine

(13.5hp at 4300rpm) www.moki.co.uk

2.1: - Aircraft specification document divergence.

The Aircraft specification document was incorporated within AVT-0215 and was very

prescriptive with a view to promoting the 5th

Intakes design solution this was later

modified to reflect the research findings of the 9th

Intake and their impact on the

Terrasoar aircraft, the initial airframe description and the resultant changes are given

below.

(a) Payload requirements: -

The aircraft was to be capable of carrying a maximum payload of up to 5 Kg. Where this

mass was to include all provision for power and / or cooling of the payload.

In addition, the aircraft nose was to be transparent to the payload sensor frequency

enabling for forward, up, down, and sideways field of regard for the payload. In addition,

provision of a continuous airflow through the payload bay during the flight of the aircraft

for environmental, for sampling payloads, must considered.

The MicroSAR, which is a low cost Synthetic Aperture Radar, was considered as a

possible payload for the Terrasoar aircraft, and was being used for land and sea detailed

observation. Market sectors include agriculture, forestry, geology, and maritime. Details

of this sensor were available at the web sites given below: -

www.imicrosensors.com www.astrium-space.com

The Cranfield contact for MicroSAR was Dr. Steve Hobbs, Astronautics and Space

Engineering (telephone 01234-750111 ext 5121 e-mail [email protected] ).

BAE SYSTEMS had a requirement for a low cost vehicle for airborne testing of one of a

radar system which is covered within reference 10.

A further payload considered was the Hyper Spectral Camera, which had applications for

monitoring water quality and changes to the coastline and estuaries. One problem of any

such camera is that it may have to be mounted on a small stabilised platform to remove

angular movements of the aircraft in flight. This is a trade off with the Flight Control

System.

http://www.moki.co.uk/

http://www.imicrosensors.com/

http://www.astrium-space.com/

mailto:[email protected]


24

Ultimately to reduce the manufacturing cost of the aircraft the current which is

considered as a concept demonstrator and manufacturing process learning tool, the nose

covering the payload bay was made from hand layed carbon fibre cloth and as such is

opaque to all of the above payloads, and is shown in figure 11, and the design and

manufacture of this article is covered within references 6 and 7.

Figure 11: - Terrasoar nosecone / payload bay cover manufactured by Man Tech

BAE Systems Samlesbury. (Source: - Man Tech BAE Systems).

The Terrasoar in its current form will now carry ballast representative of an inclusive

payload mass of 5kg.

(b) Airframe Construction: -

The airframe mass and physical size was kept to that required for the missions outlined

within this thesis and covered in depth within reference 1, although maturation weight

gain due to conservative safety factors due to the novel nature of the construction

materials selected for the airframe have been a consistent challenge to achieving the

performance and payload capability required.


25

Careful consideration was given in the airframe design to ease of: -

Transportation:

Assembly and Dis-assembly:

Manhandling.

The above areas are covered in sections 4 and 5 of this thesis.

Materials used in the construction are limited only by the need to avoid those which are

dangerous to individuals and the environment.

It was expected that use would be made of composite materials to achieve the mass and

stiffness targets of the airframe, and indeed extensive use of low temperature (800 C to

1000 C) curing glass pre-preg cloth was intended for the wing and empennage with resin

infusion intended for the fuselage of this aircraft to reduce manufacturing costs and

provide manufacturing process “learning” within BAE Systems.

Economic construction methods which minimised costs of airframe manufacture were

used in conjunction with the technology development requirements within BAE Systems

to develop methodologies to build small low cost UAV‟s, and this is covered within

section 7 of this thesis in relation to the wing and in reference 6 with respect to the

fuselage. A suitable manufacturer was selected by selecting BAE Systems Man Tech and

the reasons for this are covered within references 8, 6, and 1, Man Tech is consistent with

the quality requirements for the route to certification.

The sensitivity of any payload sensor requirements to the fuselage material surrounding it

was investigated, but for the reasons given above the payload has been omitted from this

configuration.

(c) Propulsion: -

Intake 5 purchased a MOKI 215cc 4 stroke 5 cylinder radial engine. This engine was to

be utilised for the 9th

Intakes GDP. Some limited engine with propeller testing was

conducted by Intake 5 at Woodford.

This testing was evaluated and the engine performance was confirmed by additional

testing which was conducted at ground level (taken as sea level) ambient conditions using

a ground test rig designed and constructed specifically for Terrasoar testing, and is

covered within reference 8.

The engine was integrated into the aft fuselage and this is covered within references 6

and 8. Studies were instigated to determine if the engine gearbox could be removed and

replaced with a simple toothed belt drive, in order to remove cost and reliability

associated with the gearbox and these are covered within reference 8.


26

(d) Performance: -

1) Height and Altitude:-

The maximum normal operating altitude will be 10000ft. The MOKI engine performance

needs to be confirmed at this altitude. The funds provided for this project of £100,000

were just sufficient to cover the aircraft manufacture, materials, systems, and

qualification testing, as well as Category B flight clearance to 400ft, full certification to

fly the aircraft above 400ft would have required additional in the order of £50,000 and

after consultation with the customer additional funding was ruled out. Therefore this

requirement was deferred and removed form the 9th

Intake, thus there was no need to test

the engine at altitude for the flight test program currently selected for this stage of the

Terrasoar project, which are outlined below in section 2.2, and covered in detail in

references 1 and 8.

2) Endurance:-

The vehicle was to be capable of a 1 hour loiter at the maximum normal operating

altitude. The maximum total mission time will be 5 hours. This requirement was also

changed to reflect the missions outlined in section 2.2 and detailed in references 1 and 8.

3) Velocity:-

No limitation on vehicle velocity was set. Consideration however must be given to the

ability to maintain position in typical wind speeds at altitude. The aircraft is deemed

capable of meeting the stability and speed requirements of the missions outlined in

section 2.2, for further information the reader is referred to references 1 and 8.

4) Environmental Conditions:-

Initially defined performance characteristics which were to be available under ISA

conditions up to 10000ft were modified to take account of the revised mission

requirement and reduce the complexity an hence cost of the aircraft systems: -

Although it should be possible to operate the aircraft under temperature variations from a

minimum operating temperature of -15 0C to a maximum operating temperature of +115

0C, (for manufacture) no funds were available to qualify the adhesive joints which

constitute the majority of the airframes interfaces within this temperature range especially

the operating minima.

Operation in icing conditions was not required, but consideration was to be given to the

use of de-icing fluids on the vehicle. The compatibility of the materials and adhesives

with de-icing fluids was not tested but should constitute a further work package on a

Terrasoar envelope expansion program should one be considered by the customer.


27

Provision was also to be provided to permit possible operation of the aircraft in moderate

rainfall conditions. All joints and gaps are to be sealed using glass composite closure

plates for the control surfaces and end of part, as well as PRC sealant, and aero - tape

(used on gliders and large model aircraft, which will enable the aircraft to operate in

conditions of light to moderate rainfall. However the FCS was not tested to establish if

loss of signal would occur in theses conditions and the reader is referred to reference 9

for further details on this aspect of the FCS operational test and verification.

The system was also to have been able to operate, including launch and recovery in

surface wind speeds of up to 10m/s, and structurally the airframe was capable of

operating within these parameters however the reader is directed to references 1, 8, and 9

for clarification of the FCS, and engine capabilities under these conditions.

(e) Systems and operational aspects: -

The following requirements were read in conjunction with CAP722 „CAA UAV

Operations in UK Airspace‟ (ref 2).

1) Launch and Recovery:-

Launch and Recovery was to be accomplished by whatever means are deemed

appropriate. Operation, however, must be possible from areas such as small airfields. If

any physical facilities are required, then these must be defined. As stated above the 9th

Intake rejected the proposals of intake 5 and a fixed tricycle undercarriage was used with

nose wheel steering and brakes, this is covered within references 6 and 7.

Consideration was to be given to propeller clearance during launch and recovery. This

was a risk which when considering using the Intake 5 configuration. Because of the lower

altitude requirements of the initial specification the propeller was resized accordingly,

and an undercarriage sized to meet the propeller clearance requirements during aircraft

rotation was designed for Terrasoar. The sizing of the propeller is covered in references 7

and 8.

2) Range:-

The air vehicle was to be able to operate up to at least 20Km from the Ground Control

Station (GCS). Category B clearance was given to line of site range to ensure adequate

manual revision this is covered within reference 1. An overview of the missions slated for

the Terrasoar aircraft is given in section 2.2 of this thesis and are covered in detail in

reference 1.


28

4) Vehicle Avionics Package:-

The original 5th

Intake Terrasoar was designed to use the „XRAE‟ avionics crate, which

was shared with Eclipse and Total Eclipse, neither of which has ever flown or are likely

so to do, hence because of its availability and capability it was proposed for the Terrasoar

small UAV. Eclipse and Total Eclipse were both unstable aircraft configurations.

However because Terrasoar is a stable aircraft, the question was raised about whether the

XRAE crate and its Flight Control System (FCS) were necessary. After inspection of the

XRAE crate by the 9th

Intakes FCS expert Mr Robert Currie it was suggested that the best

location for this device was a Science Museum exhibit, and after much research and cost /

capability trades the Micropilot FCS system from Canada was selected and this system is

covered in detail within reference 9, and its integration with the airframe is covered in

reference 6.

The Cranfield contact for FCS is Mike Cook, Flight Dynamics Group (telephone 01234-

750111 ext 5337 or e-mail [email protected] ).

The avionics requirements for UAV flight in shared airspace were also to be investigated,

and the cost of achieving this needed to be determined. Two possible choices were

proposed by the Cranfield University which were as follows: -

(a) Could the Terrasoar be given the capability to fly in shared airspace?

(b) Could the Terrasoar be limited to flying over a range within the confines of controlled

airspace?

The Barrow-in-Furness range in Cumbria was originally proposed by the customer but

this proved to be too expensive and the range selected for all planned Terrasoar flights

was Cambeltown Airport (previously known as Machrihanish Airport) in Argyll,

Scotland under the operational umbrella or BAE Systems Regional Aircraft Prestwick,

this is covered in detail in references 1 and 8.

(f) Design requirements: -

The route to certification was to be established by following the guidelines of ref 1

„Aircraft Airworthiness Certification Standards for Civil UAV‟s issued by the CAA.

Terrasoar was required to meet the guidelines in ref 1 and the airworthiness requirements.

These were to be either: - Microlight aircraft (section S), JAR VLA (very light aircraft) or

JAR VLA (very light aircraft) or JAR 23. In the event the JAR VLA (very light aircraft)

requirements were selected as the route to certification and the implementations of these

requirements are covered in references 1 and 8.

mailto:[email protected]


29

2.2. Revisions mission requirement specification.

The mission requirements for the Terrasoar aircraft: - airframe, flight control system,

avionics, propulsion, and fuel systems clearance (i.e. the complete Terrasoar platform as

an integrated system, for handover to the customers Cranfield University and BAE

Systems had changed from a single flight to 10,000ft which was determined to be

unrealistic within the Terrasoar budget and timescales, to four modest missions as

described below, to be flown from Cambeltown Airport (previously known as

Machrihanish Airport) in Argyll, Scotland.

The selection of Cambeltown Airport was due to its proximity to the sea permitting

emergency diversion, and the airports current low operational usage. Flight Testing was

to be performed in Uncontrolled/Sanitised Airspace and would strictly adhere to an

issued Flight Plan sanctioned and Approved by BAE Systems – Regional Aircraft,

Prestwick, who were the flight certification authority for the Terrasoar aircraft.

All Take-Off and Landings was to be performed by the Mission Commander; who would

be a certified pilot with a valid Private Pilots Licence (PPL).

Figure 12: - Cambeltown Airport illustrating its layout and proximity to the sea.


30

In order to minimise any damage to personnel, buildings, vehicles etc all flight plans

devised by Mr Alan Barnes and detailed in reference 1 were chosen to be where possible

predominately over the sea. However this would not be possible in the early flight test

stages because the range of the radio control signal was limited to 400 metres.

The four flight test plans selected for each were as follows and are presented in overview

here: -

S1 Initial Airworthiness evaluation – Pilot In Control PIC / 400 meter radius /100

metre altitude:

S2 FCS Evaluation – As S1 (Micropilot FCS) autopilot engaged:

S3 Out of Radio Control Range evaluation – 1 kilometre range over sea/700metre

altitude:

S4 Endurance evaluation – 3km range over sea/700 metre altitude/1Hour endurance.

Mission S1: - Initial Airworthiness evaluation.

Certification S1 was to be the initial flight of Terrasoar after acceptable pre-flight ground

testing including fast Taxi-runs. S1 would demonstrate a preliminary level of aircraft

flight capability by performing a takeoff, single circuit of the airfield at an altitude of 100

metres. Followed by controlled approach and landing. This mission was to be performed

under full Pilot In Control (PIC) mode. The level of Documentation identified as

Documentation S1 in reference 1, was devised to satisfy the signatory that the air vehicle

had been designed to achieve a safe and controlled flight. Successful achievement of

Certification S1 would allow progression to the next level of flight testing and would be

part of the formal aircraft Certification and Qualification documentation supplied to the

customers.

The flight plan for the S1 mission identified the proposed flight parameters including the

intended altitude and flight speeds.

Conditions: - Clear sky to XXX metres: high cloud base: no rain.

Winds: - of less than XX m/s (near still air conditions) measured wind speed at sea level.

Fuel: - XX Litres (With Ballast Fitted XX grams).

Takeoff Direction: - into wind.


31

Figure 13: - Micropilot representation of the S1 mission flight plan. (Source: -

Reference 12: - Document Number: - T-CONFIGM-0009-Iss DRAFT (19/09/05)).

Table 2: - Mission S1 Test documentation. (Reference 12)

AVT-0215-ISS 2 Customer Specification 2

DBS-TT-0001 Design Build Standard 1

DS-TT-0001 Design Standard 1

T-CERT-0001 A/C Flight Certificate (C of AW) Statement for S1/S2 DRAFT

T-CERT-0003 A/C Level SQS 0

T-CERT-0004 Equipment Prelim Hazard Analysis 0

T-CERT-0006 Fuel PSQS Statement for S1/S2 DRAFT

T-CERT-0007 Powergen PSQS Statement for S1/S2 DRAFT

T-CERT-0009 FCS PSQS Statement for S1/S2 DRAFT

T-CERT-0010 NWS/Braking PSQS Statement for S1/S5 DRAFT

T-CERT-0010 Risk Register 1

T-COMPM-DOC Compliance Matrix 1

T-CONFIGM-0004 Terrasoar Flight Manual DRAFT

T-CONFIGM-0008 Change Control Procedure 1

T-CONFIGM-0019 Terrasoar Flight Clearance DRAFT

T-DESIGN-TS-0006 TERRASOAR AIRFRAME STRUCTURE SIGN WING 1

T-DESIGN-TS-0007 TERRASOAR AIRFRAME STRUCTURE SIGN FUSELAGE 1

T-DESIGN-TS-0008 TERRASOAR AIRFRAME STRUCTURE SIGN EMPENNAGE 1

T-DESIGN-TS-0009 U/C Design Sign-off 0

T-DESIGN-TS-0010 Design Statement for S1/S2 0

T-DESIGN-TS-0011 Stress Statement S1/S2 DRAFT

T-EDR-AVS/FCS-C FCS EDR Compliance 0

T-EDR-FUEL-C Fuel System EDR Compliance 0

T-EDR-PAYLOAD-C Payload EDR Compliance 0

T-EDR-POWERGEN-C Power Generation EDR Compliance 0

T-EDR-POWERPLANT-C Powerplant EDR Compliance 0

T-PERM-PAR-0010 Performance Statement for S1/S2 DRAFT

T-SYST-DOC-0002 System Safety Programme Plan DRAFT

T-SYST-DOC-0003 Flight Limitations 1

T-TEST-0002 Flight Testing 1

T-TEST-0005 Test Statement for S1/S2 DRAFT


32

Mission S2: - FCS Evaluation.

Certification S2 was to be the second flight of Terrasoar after successful completion of

S1 testing. The S1 mission having demonstrated the preliminary aircraft handling in

flight by performing: - a takeoff: a circuit of the airfield at 100metres altitude: and a

controlled approach and landing. The S2 mission was to consist of a circuit of the airfield

at 100 metres altitude performed using the Micropilot FCS and PIC used only for Takeoff

and Landing. The level of Documentation identified in Documentation S2 would satisfy

the signatory that the air vehicle had been designed to achieve a safe and controlled flight

and will provide confidence testing for the FCS system within a safe window for manual

override. Successful completion of Certification S2 would allow progress to the next

level of flight testing and would be part of the aircraft Certification and Qualification

documentation.










33


AVT-0215-ISS 2 Customer Specification 2

DBS-TT-0001 Design Build Standard 1

DS-TT-0001 Design Standard 1

T-CERT-0001 A/C Flight Certificate (C of AW) Statement for S1/S2 DRAFT

T-CERT-0003 A/C Level SQS 0

T-CERT-0004 Equipment Prelim Hazard Analysis 0

T-CERT-0006 Fuel PSQS Statement for S1/S2 DRAFT

T-CERT-0007 Powergen PSQS Statement for S1/S2 DRAFT

T-CERT-0009 FCS PSQS Statement for S1/S2 DRAFT

T-CERT-0010 NWS/Braking PSQS Statement for S1/S5 DRAFT

T-CERT-0010 Risk Register 1

T-COMPM-DOC Compliance Matrix 1

T-CONFIGM-0004 Terrasoar Flight Manual DRAFT

T-CONFIGM-0008 Change Control Procedure 1

T-CONFIGM-0019 Terrasoar Flight Clearance DRAFT

T-DESIGN-TS-0006 TERRASOAR AIRFRAME STRUCTURE SIGN WING 1

T-DESIGN-TS-0007 TERRASOAR AIRFRAME STRUCTURE SIGN FUSELAGE 1

T-DESIGN-TS-0008 TERRASOAR AIRFRAME STRUCTURE SIGN EMPENNAGE 1

T-DESIGN-TS-0009 U/C Design Sign-off 0

T-DESIGN-TS-0010 Design Statement for S1/S2 0

T-DESIGN-TS-0011 Stress Statement S1/S2 DRAFT

T-EDR-AVS/FCS-C FCS EDR Compliance 0

T-EDR-FUEL-C Fuel System EDR Compliance 0

T-EDR-PAYLOAD-C Payload EDR Compliance 0

T-EDR-POWERGEN-C Power Generation EDR Compliance 0

T-EDR-POWERPLANT-C Powerplant EDR Compliance 0

T-PERM-PAR-0010 Performance Statement for S1/S2 DRAFT

T-SYST-DOC-0002 System Safety Programme Plan DRAFT

T-SYST-DOC-0003 Flight Limitations 1

T-TEST-0002 Flight Testing 1

T-TEST-0005 Test Statement for S1/S2 DRAFT

Mission S3: - Out of Radio Control Range evaluation.

Certification S3 was to be the third flight of Terrasoar after acceptable S2 testing. S3

would demonstrate an extended level of aircraft flight capability by performing a takeoff,

agreed flight plan at up to a distance of 11 kilometres from the mission controller (Base

Station) and at an altitude of 700 metres. Followed by controlled approach and landing.

The mission will be performed utilising the FCS Micropilot functionality with PIC for

Takeoff and Landing.


34

The distance from the Base Station was to be greater than the range for communication

within the Micropilot specification and therefore would test the reversionary aspects of

the system. The level of Documentation identified in Documentation S3 would satisfy the

signatory that the air vehicle has been designed to achieve a safe and controlled flight and

will provide confidence testing for the FCS system within a safe window for manual

override. Successful completion of Certification S3 would allow progression to the next

level of flight testing and would form part of the formal aircraft Certification and

Qualification documentation.










35


Document Name Issue Status Owner

Design Build Standard 0 A. Barnes

Design Standard 0 A. Barnes

Aircraft PSQS 0 D. Baird

Fuel PSQS 0 V. Hind

Payload PSQS 0 Ra. Cunliffe

Power Gen PSQS 0 Ro. Cunliffe

FCS PSQS 0 B. Currie

Powerplant PSQS 0 D. Adams

Wing EDR Compliance (covering) 0 G. Wardle

Fuel EDR Compliance 0 V. Hind

Payload EDR Compliance 0 Ra. Cunliffe

Power Gen EDR Compliance 0 Ro. Cunliffe

FCS EDR Compliance 0 B. Currie

Fuse EDR Compliance 0 P. Gilligan

U/C EDR Compliance 0 R. Sneddon

Tail & Booms EDR Compliance 0 J. Pennington

Powerplant EDR Compliance 0 D. Adams

Equipment Documentation 0 Ro. Cunliffe

Hazard Analysis (Equipment & Aircraft) 0 D. Baird

Mass/Performance & C of G Statement 0 A. Barnes/J. Pennington

Concession Document 0 TBD

Test Documentation 0 E. Bradley

Mission S4: - Endurance evaluation.

Certification S4 was be the final certification flight of Terrasoar after acceptable S4

testing. S4 would demonstrate an endurance flight capability by performing a takeoff,

agreed flight plan at up to a distance of 3 kilometres from the mission controller (Base

Station) and at an altitude of 700 metres. Followed by controlled approach and landing.

The mission was to be performed using the Micropilot FCS for the sustained airborne

flight with PIC for Takeoff and Landing, with the aircraft repeating the circuit for a 1

hour duration thus demonstrating the Terrasoar‟s endurance capability. This mission was

to conclude flight certification and acceptance trials upon which ownership of the aircraft

and all documentation including the aircraft manuals could be formally handed over to

the customers namely Cranfield University and BAE Systems.

These four missions constituted a much more structured flight clearance program than

that originally envisioned in the first draft of AVT-0215 and were agreed with BAE

Systems Regional Aircraft to enable clearance of the Terrasoar Unmanned Air Vehicle,

and this concludes the requirements capture section of this thesis.


36

3.0: - Airframe Conceptual Design (Phase 1).

The selection of the best configuration for the mission of an aircraft is fundamental to all

aircraft projects, and when there is no clear wining configuration it is necessary to

identify two or more possible configurations for independent study and subsequent

comparison this invariably involves trade studies and the final configuration could be an

combination of best features of two or more configurations that meet specific

requirements, as was the case with the F-35 Joint Strike Fighter.

Three design options were considered as stating points to meet the Terrasoar

requirements which were as follows:-

1) An adaption or specialised derivative of an existing design: - This option is

invariably in response to systems changes and / or mission enhancements and role

expansion, for example specialised ECM aircraft derived from combat aircraft

airframes such as the EF-111 Raven derived from the F-111A tactical fighter

bomber, and the defence suppression F/A-18G Growler derived from the F/A-18F

tactical fighter. These changes are usually quite modest for example a new nose to

house optical and / or infra red sensors for a reconnaissance aircraft derivative of

a fighter or bomber or a new fin (vertical tail) for an Electronic Countermeasures

Measures derivative to house the RF jamming antenna. This is considered as a

low risk low cost option.

2) A major modification or direct development of an existing aircraft: - This

option usually involves expensive major changes to the airframe for example the

fuselage extensions and new composite wing of the Harrier AV-8B / GR-5 to

accommodate heavier bomb loads and more capable systems, over the original

Harrier AV-8A / GR-3, and the evolution of the F/A-18A into the F/A-18E with a

new wing, fuselage stretch, engines, and systems. This option has a relatively high

risk compared with option 1 depending on the extent of the development

undertaken where as the development of the Tornado Air Defence Variant (ADV)

stand off interceptor, from the original Tornado Interdictor Strike (IDS) bomber

(the reverse of the development of the F-15E bomber from the F-15D fighter)

would be considered as low risk with fuselage plug inserts, the development F/A-

18E would be considered as high risk, with the high possibility of adversely

impaction on the existing types capabilities in areas other than those addressed by

the specific need. Therefore careful trade studies are required to avoid a point

design and enable the modifications to enhance capabilities rather than substitute

one at the expense of others, still required to realise mission objectives.


37

3) A completely new design: - This is clearly the most expensive of the three

options and carries the greatest risk, however there are circumstances when this

option is necessary such as in the development of: - stealthy (low – observable)

manned combat aircraft, both fighters and bombers: UCAV‟s and novel UAV‟s

where imparting of completely new capabilities which cannot be retro-fitted to

conventional aircraft. For example during the Cold War it was determined by the

USAF that the Radar Cross Section (RCS) of the B-52‟s and B-1 bombers in Air

Combat Commands (ACC) inventory could not be significantly reduced by

application of radar absorbing materials, and only a completely new design which

incorporated signature reduction from the outset of the configuration, would yield

a significant RCS reduction enabling ACC to complete its mission of deep

penetration into hostile airspace this design became the Northrop B-2 Sprit

bomber. Although undertaking a completely new aircraft design is much less in

the first decade of the 21st century than it was in say the early 1950‟s in order for

an aircraft manufacturing nation to maintain its design, manufacturing, and

operational capabilities it is essential to periodically develop a new aircraft

usually as part of a multinational team, as with the Airbus A380, and the Joint

Strike Fighter F-35 projects.

In view of the change in mission requirements of the Terrasoar UAV over those

originally stipulated for the 5th

Intakes aircraft namely the operational altitude being

reduced from 40,000 feet to 10,000 feet, and the launch and recovery problems associated

with their design the scope of this phase was widened from a review of the 5th

Intakes

design with some minor modifications (or option 1 above), to a clean sheet design study

(or option 3 above). The objective of this phase was to propose a solution which more

closely matched the new requirements covered in section 2 above. The first part of this

process was to produce configurations and match them against the mission and systems

requirements, and where possible improve on the 5th

Intakes configuration. This was

done by concept generation, and hand analysis calculation, only CAD models were

produced and no full parametric analysis beyond the level of the M1 Cranfield University

module material was conducted. Down selection was based on lest risk configuration

until the final selection meeting when major modification and development of the

original Terrasoar configuration (or option 2 above) was selected as the lowest risk

option for the 9th

Intake by the customer.

When considering the clean sheet design the following factors highlighted in reference 12

(Page 16: Aircraft Conceptual Design Synthesis: by Howe. D: Published by Professional

Engineering Publishing Ltd 2000), should be considered in terms of their effect on the

reasoning behind design choices. These factors and their influence on configurations in

general terms are detailed below: and are as follows:-

(a) Technical developments:

(b) Radical innovative configurations:

(c) Conservative or well established configurations:

(d) Optimisation and risk reduction.


38

A. Technical developments: - The technical developments which would have the

greatest impact on a small UAV of the Terrasoar‟s scale were confined to the field

of large radio controlled flying models, were reliable and relatively cheap scale

turbofan and turbojet engines have come to be readily available, and the authors

thoughts turned to the possibility of designing a scaled U-2 type aircraft of long

range high altitude (not possible with the MOKI engine), and high endurance with

a more comprehensive fuel system than the 5th

Intakes plastic bag, this would also

obviate the propeller ground clearance requirements, and the very dubious

indexing system.

An other major development was in the field of Flight Control System Avionics,

the XRAY crate was clearly dated from initial inspection and a much more

compact and light weight system could found and was purchased for this project,

although throughout the concept phase the BAE Systems customer representative

Mr Peter Hamilton insisted that space be available for the XRAY crate in all

configurations studied which was primarily due to his involvement in the non-

flying Eclipse program.

Further technical developments which were to play a key role in Terrasoar design

decisions were innovations in manufacturing processes and materials, such as

resin infusion, and a foam sheet material R63.80 which was low temperature

curing foam in sheets 3mm thick and could be used in making wing spars, ribs,

and skins so a light but strong wing could be produced providing tooling costs

could be controlled, as will be seen in section 4 of this thesis manufacturing

tooling cost estimates eventually mitigated against this innovative material

although it has been used successfully on other UAV and manned trainer aircraft

projects more complex than Terrasoar, and could have reduced the weight

problems that were to stalk the Terrasoar in design maturation.

B. Radical innovative configurations: - these are often seen as possessing

significant advantages over more conservative designs for example the radical

Northrop YF-23 Advanced Tactical Fighter (ATF) design compared with the

conservative Lockheed YF-22 ATF design, but as experience demonstrated the

former aircraft (YF-23) was an excellent point design as a stealthy interceptor, but

the latter aircraft (YF-22) was more manoeuvrable and applicable to a wider range

of air combat situations than the YF-23, and it is the Lockheed F/A-22A which

will form the backbone of the USAF ACC fighter arm for the next 40 years and

not the more radical Northrop aircraft. Therefore, radical configurations should

only be considered when their theoretical advantages are relevant to the dominant

requirement in the specification. An other example taken directly from reference

13, is the Wright‟s Flyer which had a foreplane configuration, which had major

advantages in terms of lift generation, but later experience demonstrated that this

configuration had adverse secondary effects which off set the lift advantages for

most subsonic aircraft, and it is more suitable for supersonic flight, or when the

FCS can accommodate a naturally unstable aircraft.


39

C. Conservative or well established configurations: - These exist for some aircraft

types such as large subsonic transports both military and civilian, for example the

Airbus range of airliners form the A-300 to the A-380 all have moderately swept

wings, with pod turbofan engines hung underneath the wings a wide body

fuselage with the exception of the A-320 and its derivatives, and a single vertical

fin, and all moving horizontal elevator tail, and this configuration with the

exception of the Lockheed Tristar, DC-10, and MD-11 is the prevalent airliner

configuration on both sides of the Atlantic. Military transports follow a

conservative configuration as a functional requirement all have rear loading ramps

and T-tails, or high mounted tails, the former type is seen on the Lockheed C-5

Galaxy, the Airbus 400M, and the Boeing C-17, the latter is seen on the Lockheed

C-130, the AN – 124, and the Transal, all have pod engines either turbofans or

turboprops, and all have high mounted wings to clear the cargo bays. Therefore

configuration selection is straightforward, and the only time two or more different

configurations would be considered would be if there were compelling

requirements to incorporate innovative technological developments, such as the

blended wing body, which is still under research study for both of these

applications, or propfans studied in the late 1970‟s.

D. Optimisation and risk reduction: - The final configuration submitted as the

potential customer solution which eventually progresses through the design

process to detail design and final manufacture, is not necessarily the one which is

determined to be the best form the optimisation process. Indeed a radical solution

may have many attributes to recommend it but the very novelty of the

configuration may mitigate against it on grounds of risk reduction or cost in

changing the operating infrastructure to support the aircraft, this is seen as one of

the reasons the blended wing body airliner has still to catch on, where as the

Airbus A380 still looks conventional and can operate from major airports without

their needing to undertake any major changes. The true cost of the final

configuration will also be a major factor for example a truly mission adaptive

wing currently is still a very expensive option and is confined to research aircraft,

because of their exotic materials, sensors, and actuators, and shear complexity

compared with conventional three section wings (leading edge flap: torsion box:

trailing edge flap).

3.1: - Airframe configuration submissions and design decisions.

With the above points in mind although Phase 1 was opened to consider a completely

new configuration the most likely outcome from this Phase would be a final decision to

rework the existing 5th

Intakes configuration on grounds of cost, time-scale, and facilities

to verify any completely new concept. However even this pessimistic outlook should not

have dampened an opportunity to demonstrate creativity, but unfortunately this was not a

view shared by the majority of the 9th

Intake. An open forum session was held 13th

August 2003, to present alternative configurations to that of the 5th

Intake, to which there

were four original respondents including the author and three who submitted pictures of

existing designs out of thirteen and these are presented below.


40

Figure 16: - Damian Adams proposed configuration for alternative Terrasoar.

General Notes

• Schematic layout only - sizing to be

established

• Configuration capable of supporting

existing XRAE crate

• Single Propeller mounted in blended

structure above fuselage which may

need narrowing in this region to

place propeller in free-stream air.

– Negates issue around propeller

orientation upon landing

• ‘V’ Tail can be configured (angled)

to reduce the yawing affects caused

by the propeller

• Low wing enables use of ‘preferred’

carry through structure for wings,

this also improves internal

packaging

• Traditional flaps, elevators etc.

employed

• Configuration allows ease of

position of the payload

• Low wing eases use of an

undercarriage - could be located

above wing root in pods

• Anhedral wing to aid landing

• LE sweep in the order of 5-10

degrees

• Assumes propeller diameter

reduction to circa 500-600mm

diameter

Figure 17: - Chart of supporting notes for Damian Adams submission.

„ V ‟ Tail Configuration - Taking Surfaces out of Prop Wash (Alternative is High „ T ‟ Tail)

Single Boom Joining Fuselage to Fin - Carrying Services and Signal

High Mounted Engine with Cowling Blended into Fuse

Propeller Positioned Clear of Fuselage and Boom - no need for Indexing Mechanism

Medium Aspect Ratio, Low Wing Supporting Fuselage Carry Through

Hemispherical, Rotary, Multiple Cameras Blended into Front Fuse Sampling Also to be adjacent Nose

Main Fuselage Section Sized for Fuel, Avionics, Payloads and Potentially U ‟ Carriage


41

Figure 18: - Mr Alan Barnes submission for an alterative Terrasoar configuration

which appears to be the Hindustani Aircraft Industries (HAI) UAV concept.

Figure 19: - James Pennington‟s submission for an alternative Terrasoar

configuration for which data and additional views are shown in figure 20 below.


42

Figure 20: - James Pennington‟s submission for an alternative Terrasoar

configuration appeared to be a high drag configuration with little internal volume.

Figure 21: - Robert Currie‟s submission for an alternative Terrasoar configuration

more details are covered in figures 22 through 23, a good looking design.


43

Figure 22: - Robert Currie‟s alternative Terrasoar configuration dimensions.

Figure 23: - Robert Currie‟s alternative Terrasoar configuration notes.


44

TACTICAL RECON UAV CONFIGURATION ONE.

High tail clears prop wash.

Long fuse balances

engine mass gives space

envelope for fuel and

FCS, with best field of

regard for sensors.

Ball turret for optical sensor.

Figure 24: - The author‟s first alternative Terrasoar configuration 1.

TACTICAL RECON UAV CONFIGURATION ONE.

Flaperon wing control surfaces.

Single elevator and rudders.

Figure 25: - The author‟s first alternative Terrasoar configuration 1.


45

GENERAL FEATURES OF CONFIGURATION 1 NEW BUILD.

Wing span:- 4.0 m

Root Chord:- 1.0 m

Tip Chord:- 300 mm

L/E Sweep:- 15 deg

O/A Length:- 3.5 m

Fuse width max:- 300 mm

Propeller diameter:- 700 mm

High wing, high tail, twin boom configuration with single propeller. Tandem

undercarriage units mounted in fuselage with wing tip outriggers or

reinforced wing tip snubbers fitted. Terrasoar engine and flight control crate

fitted and wing joints at the boom interfaces.

A smaller wing taper ratio in this case 0.3, decreases the moment arm

from the root to the centre of pressure maintaining the same lift, and

decreases the bending moment at the wing root, resulting in a lighter wing

structure.

Figure 26: - Alternative Terrasoar configuration 1 key design feature notes.

Figure 27: - The author‟s second alternative Terrasoar configuration 2.

Low drag slender fuselage configuration.

Main undercarriage stored in rear of wing pods mounted on

removable wing.

High lift wing reducing landing speed and a

lighter wing structure.


46

Figure 28: - The author‟s second alternative Terrasoar configuration 2.


Wing span:- 5.0 m

Root Chord:- 1.0 m

Tip Chord:- 500 mm

L/E Sweep:- 15 deg

O/A Length:- 4.7 m

Fuse diameter :- 300 mm


High wing, medium tail, long fuselage configuration with single propeller.

Main undercarriage units mounted in wing pods nose unit in fuselage, uses

Terrasoar engine and flight control crate, with sensors in nose and wing

pods.

A larger wing taper ratio in this case 0.5, than configuration one but still

decreases the moment arm from the root to the centre of pressure

maintaining the same lift, and decreases the bending moment at the wing

root, resulting in a lighter wing structure.


TACTICAL RECON UAV CONFIGURATION TWO

The twin tail offers the same amount of lateral stability for a smaller size than a single tail fin.

Engine air cooling from meshed annulus.


47

TACTICAL RECON UAV CONFIGURATION THREE.

Figure 30: - The author‟s third alternative Terrasoar configuration 3.

TACTICAL RECON UAV CONFIGURATION THREE.

Figure 31: - The author‟s third alternative Terrasoar configuration 3.

High mounted jet engine

out of ground FOD field.

Larger fuselage configuration enabling a larger fuel tank to be

incorporated.

The twin tail keeps the rudders out of the jet wash.


48


Wing span:- 5.0 m

Root Chord:- 1.0 m

Tip Chord:- 500 mm

L/E Sweep:- 15 deg

O/A Length:- 4.6 m



Pod jet version of configuration 2, this high wing, medium tail, large

fuselage configuration has twin tails. Main undercarriage units mounted in

wing pods nose unit in fuselage, uses jet engine and flight control crate,

with sensors in nose and wing pods.

A wing taper ratio of 0.5, decreases the moment arm from the root to the

centre of pressure maintaining the same lift, and decreases the bending

moment at the wing root, resulting in a lighter wing structure.


TACTICAL RECON UAV CONFIGURATION FOUR.

Figure 33: - The author‟s fourth alternative Terrasoar configuration 4.

Main undercarriage stored in rear of wing pods mounted on

removable faired in wing.

The large fuselage configuration enabled a larger fuel tank to be incorporated and a 25kg payload

to be carried.


49

TACTICAL RECON UAV CONFIGURATION FOUR.

Figure 34: - The author‟s fourth alternative Terrasoar configuration 4.


Wing span:- 5.0 m

Root Chord:- 1.0 m

Tip Chord:- 500 mm

L/E Sweep:- 15 deg

O/A Length:- 4.6 m



Large fuselage section version of configuration 2, this high wing, medium

tail, long fuselage configuration with single propeller. Main undercarriage

units mounted in wing pods nose unit in fuselage, uses Terrasoar engine

and flight control crate, with sensors in nose and wing pods.





The twin tail offers the same amount of lateral stability for a smaller size than a single tail fin.

The original Terrasoar engine with extended prop shaft is

incorporated in this design.


50

TACTICAL RECON UAV CONFIGURATION FIVE.

Figure 36: - The author‟s fifth alternative Terrasoar configuration 5.

TACTICAL RECON UAV CONFIGURATION FIVE.

Figure 37: - The author‟s fifth alternative Terrasoar configuration 5.

Imbedded jet engine to reduce drag fitted with short side pod ducts.

Fuselage blanking of the short side pod ducts would be a problem.

Empennage control actuators near hot exhaust would be a problem.


51


Wing span:- 5.0 m

Root Chord:- 1.0 m

Tip Chord:- 500 mm

L/E Sweep:- 15 deg

O/A Length:- 4.5 m



Single tail jet version of configuration 4, this high wing, medium tail, long

fuselage configuration. Main undercarriage units mounted in wing pods

and nose unit in fuselage, uses jet engine in rear fuselage with aft intakes,

flight control crate, with sensors in nose and wing pods.





Author’s configurations general notes.

• All configurations have flaps, ailerons, and elevators.

• Only one conventional tail configuration was considered.

• Aircraft configurations 3 to 5 have tricycle undercarriage arrangement.

• All configurations have been sized based on Terrasoar overall

dimensions, and have been modified to meet the new mission

requirements and need parametric analysis.

• The prop size has been reduced to diameters of 700 mm - 800 mm to

reflect the new mission requirements and needs parametric analysis, also

no sizing has been done on the jet configurations.

•These are all general arrangement CATIA V5 models and are only

intended for configuration concept capture and not intended to form the

basis of surface models for manufacture.

Figure 39: - Overview design notes on all of the authors alternative Terrasoar

configurations.


52

Figure 40: - Mr Craig Carr‟s alternative Terrasoar configuration a picture of the

Pioneer light UAV.

Figure 41: - Mr Robert Sneddon‟s alternative Terrasoar configuration the Boeing

Scan Eagle light UAV.


53

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55

3.1.1: - Authors notes on configuration generation: - In order to develop the

configurations presented above, the author had to make some basic decisions in advance

of defining the first lines of the CATIA V5 CAD model and conduct basic trade studies

as to the form which the final submissions would take, and I am sure the reader will find

similar consideration in the thesis of the other team members who made submissions at

this stage of the design process. These decisions were based on the following choices: -

(1): - The best location for the propeller and engine, either a tractor (with the propeller

and engine at the nose of the aircraft) or a pusher (with the propeller and engine at the

rear of the aircraft):

(2): - The type of engine giver a free choice, either a piston engine as per the original

requirements of AVT-0215, a jet (turbojet or turbofan), a turboprop: or even a ducted fan:

(3): - The number of engines given the requirements for a safety case was it really a good

idea to rely on a single engine, on the other hand multi engines would add cost

complexity, and asymmetric flight cases should one of two wing mounted engines fail:

(4): - The position of the wing would the best solution be a low - wing, mid – wing, or

high – wing, or indeed would a biplane configuration be appropriate, and should the wing

be swept or straight, and the airfoil shape required for low drag and high lift:

(5): - Empennage location, would this be boom mounted, aft fuselage mounted, or front

mounted canard, or wing tip mounted fins, or indeed would be needed at all (e.g.

UCAV‟s or Dark Star UAV).

These design choices were made by recourse to literature both theoretical and practical

experience based and examination of existing UAV‟s, developed for reconnaissance

purposes.

1):-Engine / propeller location: - Selection of a single or multiple engine propeller

power plant for the UAV give rise to a choice of locations for the propeller i.e. either

tractor or pusher configurations to determine the best location consideration was given as

to the relevant advantages and disadvantages of either location for the Terrasoar aircraft

these are described below and examples of each are shown in figures 43 and 44

respectively:-

Tractor configuration Advantages: - (1) The heavy engine would be in the nose

of the aircraft which would help to move the centre of gravity forward thus

allowing a smaller empennage for stability: (2) The propeller would be working in

an undisturbed free stream of air: (3) The engine especially an air-cooled engine

of the type that the MOKI was would receive much more effective cooling: (4)

The propeller ground clearance problems on rotation of the aircraft are reduced

for a tricycle undercarriage layout and are eliminated in a tail dragger (tail

wheeled) layout.


56

Tractor configuration Disadvantages: - (1) The propeller slipstream would

disturb the quality of the airflow over the fuselage and wing root: (2) The

increased velocity and flow turbulence over the fuselage due to the propeller

slipstream would increase the local skin friction on the fuselage increasing drag:

(3) The field of regard for any sensors in a single engine configuration would be

reduced because these would need to be mounted either above, or below, the

fuselage or in a wing nacelle and all of these positions would have resulted in

some obscuration from the fuselage itself.

Pusher configuration Advantages: - (1) A clean high quality airflow would

predominate over the wing and fuselage: (2) The field of regard for sensors would

be vastly improved by mounting them in the nose of the aircraft: (3) The aft

mounting of the propeller in the single engine configuration arrests flow

separation at the rear of the fuselage by drawing air in to the rear fuselage and

thus producing a pressure gradient favourable flow retention, this results in a

steeper closure angle for the rear fuselage, which can be made shorter reducing

the wetted surface area and drag: (4) Reduced noise and vibration for the payload

and avionics equipment: (5) Simplification of the fuel system as the fuel tank

could be positioned immediately in front of the engine (separated by a fire wall)

and fuel pipes run directly to the engine instead of convoluted paths under the

avionics bay which would need screening for its entire length.

Pusher configuration Disadvantages: - (1) A heavy engine at the rear of the

aircraft would result in a rearward shift in the centre of gravity (c.g.) of the

aircraft thus reducing longitudinal stability of the configuration, which would

require countering with more of the fuselage ahead of the wing or a heavy

payload / systems in the forward fuselage, or even additional / larger longitudinal

control surfaces: (2) Long undercarriage is usually required enable the clearance

of the propeller from the ground during aircraft rotation on take-off, and to protect

the propeller from damage from FOD on landing: (3) Ensuring a clear flow of

cooling air to the engine would be more difficult with a rear mounted pusher

configuration which would require ducting, complicating the design, of the

fuselage, and this would be a problem on the Terrasoar with the powerful air –

cooled MOKI engine (4) Location and integration of empennage control surface

actuators where the empennage is located at the rear of the fuselage, would be

difficult therefore on low cost UAV‟s the twin boom empennage configuration

predominates, with the exception of the Scan Eagle and a few other very small

UAV‟s.

In spite of the disadvantages the pusher configuration was selected for all propeller

versions of alternative Terrasoar proposed by the author because the field of regard for

the sensor payload was deemed a paramount requirement by the author.


57

Figure 43: - Raptor Tractor propeller UAV configuration used by NASA.

(AeroDYNAMIC Version 3.0: AIAA EDUCATION SERIES CD-ROM 2003)

Figure 44: - Predator Pusher propeller UAV configuration used by the USAF.

(AeroDYNAMIC Version 3.0: AIAA EDUCATION SERIES CD-ROM 2003)


58

2):-Engine type jet or piston: - Selection of the type of engine would have a profound

effect on the aircraft configuration and the reasons for considering a jet for this aircraft

were based upon the altitude of operation originally required for the Terrasoar UAV

which could not be met by the MOKI engine (intended for large model aircraft for World

War 1 films) and the availability of small turbofan engines.

A jet engine housed inside the fuselage as in configuration 5 shown in figures 36, and 37

would have resulted in duct and fuel tank integration problems if the fuel could not be

stored in inboard wing tanks with a dry outboard wing as well as within a annular fuel

tank between the side wall of the fuselage and the duct. A single duct could be

incorporated above the fuselage but this would be susceptible to loss of airflow in

manoeuvres due to obscuration by the fuselage and wings, although this is has not been a

problem on Global Hawk, on the other hand a single intake below the fuselage would

result exposure to Foreign Object Debris (FOD). One solution to jet engine integration

proposed in configuration 3 shown in figures 30 and 31 was to pod the engine above the

fuselage with pressure fed fuel from a fuselage tank below the engine this configuration

was similar to the He-162 of 1944-45 and was flight proven by the German Air Force in

World War 2, and the Boeing YQM-94A Gull UAV in the 1970‟s the engine considered

for this configuration would have been a commercially available high bypass ratio

turbofan or a ducted fan. In this location the engine would be clear of fuselage and wing

flow obscuration, as well as FOD, and could be easily maintained in service.

A different piston engine was also considered with a self starter and a two stroke cycle, or

even an all electric engine, but these were also rejected and the team on grounds of risk

and they elected to stay with the MOKI engine, so as best one could the author

incorporated this into the proposed piston engine configurations.

3):-Engine number: - As stated above a single engine aircraft runs the greatest risk of

loss if the engine fails so the author considered briefly multi engine configurations such

as the Boeing Condor twin engine HALE of 1988 (covered in reference 13: - Page 47:

Attack of the Drones A History of Unmanned Aerial Combat: by Yenne. Bill: Zenith

Press 2004), which was a large aircraft with a 200ft wing span and a 68ft fuselage, of

20,00lbs. In this configuration the engines were placed close to the aircraft centreline to

minimise asymmetry from the loss of a single engine. However multiple engines in the

author‟s view would have complicated the overall design with the requirements for

multiple fuel systems and feeds as well as increased aircraft weight and maintenance in

the field, therefore no multi engine configurations were proposed for alternative

Terrasoar.

4):-Wing type and position on the fuselage: - There were two major considerations

here namely the geometric shape of the wing and its location on the fuselage: - The first

major consideration was the shape of the wing which is described by (a) aspect ratio: (b)

wing sweep: (c) taper ratio: (d) variation of airfoil shape and thickness along the span

aerodynamic twist: and (e) geometric twist (change in airfoil chord incidence angle along

the span).


59

(a) The aspect ratio AR of the wing was given by dividing the total reference wing span

i.e. tip to tip squared b2 by the wing area S and an aspect ratio between 6 and 10 was

considered adequate for the missions envisaged for this UAV, because this falls in the

range for low speed subsonic transports, where as the aspect ratio for a fighter is between

2 and 5, and for gliders its 10 to 15.

(b) There were two sweep angles of importance when considering wing geometry,

namely the leading edge sweep angle LE and the sweep angle of the quarter-chord line

c / 4 as shown in figure 45 below.

Figure 45: - The major parameters of wing definitions. (Reference 14: - Whitford R:

Fundamentals of Fighter Design: Airlife Publishing Ltd: 2000.)

The leading – edge sweep angle is of most relevance to supersonic aircraft designs

because in order to reduce wave drag at supersonic speeds, the leading edge should be

swept behind the Mach cone, and the quarter-chord line is of relevance to high-speed

subsonic aircraft flying in the transonic region neither case was applicable to the

Terrasoar aircraft so no in depth study was made for these concepts, or in the as built

Terrasoar design.

c /4

S

b / 2

cr


60

(c) Taper ratio is the ratio of the tip chord to the root chord ct / cr shown in figure 45,

and influences span efficiency e given by 1/1+ where is the induced drag factor and

is calculated from lifting line theory and is a function of aspect ratio and taper ratio as

shown in figure 46 below.

Figure 46: - Induced drag factor as a function of taper ratio for wings of different

aspect ratios. (Reference 15:- Page 110: Anderson. J. D.: Aircraft Performance and

Design: McGraw Hill: 1999.)

For the minimum induced drag the best lift distribution will be elliptical for a straight

finite wing, which implies an elliptical plan form, like the Supermarine Spitfire of World

War 2 with curved leading and trailing edges, however this would be expensive to

manufacture and was not justified for the Terrasoar aircraft in view of the much lower

costs associated with manufacturing a wing with straight leading and trailing edges. Also

by selecting the correct taper ratio an elliptical lift distribution could be approximated.

For wings in general (i.e. those not necessarily having an elliptical lift distribution) the

Span Efficiency factor referred to above is introduced so that the induced drag coefficient

can be expressed as: -

Equation 3.1.1

ARe

CC L

Di

2


61

Where: - e = 1 for elliptical loading:

e < 1 for non elliptical loading:

(typically 0.85 < e <0.95)

Therefore total wing drag is composed of two components which are profile drag form

skin friction plus pressure, and induced drag so the drag coefficient can be expressed as

follows: -

Equation 3.1.2

Referring to figure 45 for an aspect ratio of for example 7.0, in the bottom half of the 6 to

10 aspect ratio range for this type of aircraft, the minimum value of = 0.005 at a taper

ratio of approximately 0.3, therefore a tapered wing of AR = 7 and a taper ratio of 0.3

would have an induced drag only 0.5% greater than would an elliptical wing of the same

physical size. This justified to the author further consideration of tapered wings on all of

the configurations submitted for both the first and second stages of the concept design

phase of this project.

The choice of taper ratio (defined as = ct / cr = 1 for a rectangular wing) for each

configuration was based on the following factors (covered in depth in reference 13 pages

423 – 424) namely: -

(1) Reduced wing weight over that of a simple rectangular wing: - because as the

taper ratio is reduced the centre of pressure moves closer to the wing root, and therefore

the moment arm from the wing root to the centre of pressure decreases, and thus the

bending moment at the root decreases but the lift remains the same. The net result is that

the wing structure could be made lighter by using a taper ratio less than 1.

(2) Flow separation behaviour: - low taper ratio wings exhibit undesirable flow

separation behaviour and stall behaviour. In the case of a rectangular wing = 1, flow

separation occurs in the root region which has the advantage of creating flow trails down

stream of the wing root resulting in buffeting of the horizontal tail which gives a stall

warning to the aircraft pilot. Also for the rectangular wing the flow at the tip region is

still attached therefore the ailerons usually located in the region of the wingtips would

have full authority.

InducedProfileTotal

ARe

CcC L

dD

2


62

As the taper ratio decreases the region where flow separation first develops moves

outwards towards the wing tips, and where = 0.5 the region of separation is about half

of the semi – span, and clear of both the wing root and tip ailerons. But when the taper

ratio is reduced to = 0 the stall region is initiated at the tip resulting in loss of aileron

control authority which would be unacceptable for a manned or unmanned aircraft.

Therefore a compromise between the structural benefit a small taper ratio and the

aerodynamic benefits of a large taper ratio was employed in all configuration studies in

this phase of the project, however this concept was not adopted on the as built Terrasoar

design at the insistence of the manufacturing and aerodynamic leads, on grounds cost and

complexity.

(d) A common feature for many aircraft including general aviation aircraft and some

UAV‟s such as Global Hawk is a variation of airfoil section along the wing span, where

one airfoil section is used at the wing root, and a different one is used at the wing tip with

the sections in between the root and tip being a linear interpolation between the root and

tip sections, and this is called aerodynamic twist. Where this is applied the root sections

are relatively thick (in the order of 15% to 17%), and the wing section tapers to a thinner

section at the tip (usually about 12%). The reasons for imparting this aerodynamic twist

are as follows: - (1) structural: (2) aerodynamic: (3) and wing internal capacity.

(1) The structural reasoning for this is that the bending moment is greatest at the wing

root and a thicker root airfoil section permits the design of thicker wing skins and spars at

the root imparting grater structural strength to this region of the wing.

(2) The aerodynamic reasoning is that a 17% airfoil section will stall at a lower angle of

attack than a 12% airfoil therefore if the wing airfoil section tapers from 17% thickness at

the root to 12% thickness at the tip the wing will stall first at the root, maintaining

attached flow at tip retaining the effectiveness of the ailerons at the tips because they are

still in the attached flow region. Also buffeting occurs at the wing root in stall conditions

for such an aerodynamically twisted wing and this buffeting act‟s as an early indicator to

the pilot that the aircraft is stalling enabling recovery to be attempted with all control

surfaces available.

(3) The capacity benefit comes from the thicker wing root section being employed for

internal wing fuel tanks, increasing range without losing cabin volume.

Although these benefits were explained to the team and adopted for the author‟s initial

configuration studies the manufacturing, and aerodynamic leads felt this was too much

effort would be required to produce such a wing and this concept was not adopted on the

as built Terrasoar design.


63

(e) Geometric twist is shown in figure 47 defines the situation where the chord lines for

the spanwise distribution of do not all lie on the same plane. This results in a spanwise

variation in the geometric angle of incidence for the sections. The chord of the root

section of the wing shown figure 47 is inclined at 50 relative to the aircraft axis. The

chord of the tip section, however, is parallel to the longitudinal axis of the aircraft. In this

case the incidence of the wing airfoil sections relative to the aircraft axis decrease toward

the wing tip, and the wing has “wash out”. This geometric twist is used on many subsonic

aircraft to create wash out to control the spanwise lift distribution, and hence boundary

layer separation and hence the stall characteristics. Where the angle of incidence

increases toward tip, the wing has “wash in”.

Figure 47: - Unswept trapezoidal tapered wing with geometric twist (wash out):

(This model was created by the author for IRP).

The airfoil section distribution, the aspect ratio, the taper ratio, the twist, and the sweep

angle of the wing planform are the principle factors that determine the aerodynamic

characteristics of the wing and are influential in the stall, structural weight and stiffness

of the wing. As will be seen later rejection of some of the wing characteristics promoted

in this conceptual phase would have consequences for the as built Terrasoar design, in

terms of weight, and stiffness requirements.

The second major consideration was the location of the wing relative to the fuselage, for

which there were three basic vertical location choices to be evaluated namely: - (1) high -

wing: (2) mid – wing: (3) low - wing, and these are sketched in figure 48 below, and their

relative merits and disadvantages were reviewed as discussed below.

Chord of the wing

root section

Chord of the wing

tip section

c/4 (unswept

Aircraft longitudinal axis Parallel to aircraft

longitudinal axis

y = 0

y = - (b/2)


64

Figure 48: - Sketches for the comparison of high – wing, mid – wing and low – wing

configurations after Anderson .D. John reference 13.

(a) High – wing position: - This position for the wing is commonly found on large civil

and military cargo transports and small commuter airliners such as the BAe 146, and

enables the fuselage to be placed lower to the ground which simplifies loading in the case

of transports. This configuration enables the whole fuselage section to be used for cargo

stowage without interruption from the wing box, passing through the cabin this was

partly why the Lockheed C-5 was selected over the Boeing 747 as the USAF‟s heavy lift

cargo aircraft. The high – wing also has greater lateral, rolling stability. For low wing

configurations as in figure 48(c) a dihedral upwards slope is usually built into the wing to

increase lateral, rolling stability.

The reason for this is that when an aircraft rolls the lift vector tilts away from the vertical,

and the aircraft starts to sideslip in the direction of the lowered wing. Where a dihedral

has been incorporated into the wing design of the extra flow velocity component

generated by the sideslip creates an increasing lift on the lowered wing, hence tending to

restore the wings to a level equilibrium position, and this is the basis of lateral stability

for naturally stable aircraft design of low wing aircraft. High – wing aircraft on the other

hand are much more stable in this regard requiring no dihedral, this is because the extra

flow velocity component generated by the sideslip when the aircraft rolls creates a region

of higher pressure in the flow interaction region between the fuselage side and the bottom

surface of the lowered wing at the wing root. This increased pressure under the lowered

wing has the effect of rolling the wings back to the level equilibrium position.

(a) High - wing

(b) Mid - wing

(c) Low - wing

Anhedral

Dihedral


65

Indeed the high wing position can be too stable in roll and has to be countered to improve

aircraft manoeuvring as in the case of the Lockheed C-5, C141, and BAe 146, where an

anhedral or downward slope shown in figure 48(a) was adopted to reduce over stability in

roll manoeuvres. For all of the first stage phase 1 configurations for alternative Terrasoar

generated by the author, the high – wing configuration was selected because of its high

natural stability which the author reasoned would reduce control demands on the FCS

and help a safe recovery of the aircraft in the event of an aileron actuator failure. The

author did not feel that any of the configurations 1 through 5 would suffer from over

stability therefore no anhedral slope was applied to the wing designs, as could complicate

their manufacture.

(b) Mid – wing position: - The mid – wing position used on the Lockheed U-2 spy plane

and the F-16 jet fighter shown in figure 48(b), was originally favoured by the author

because of its low drag, due to the fact that of all three options the mid – wing

configuration has the minimum wing – body interface, and unlike the high – wing and

low – wing positions requires no fillet to decrease wing – body interference, and neither

anhedral or dihedral for stability refinement. However the mid – wing has a major

structural disadvantage namely the bending moment due to wing lift must be carried

through the fuselage, and unlike the case of high – wing, and low – wing positions where

the wing torsion box can be extended across the fuselage, the mid – wing requires heavy

ring frames attached to the leading edge, and trailing edge, and intermediate wing torsion

box spars. The more spars running the length of wing results in a greater number of

fuselage ring frames carrying the bending moment across the fuselage in the in the case

of the F-16 four wing panel root attachment fish plates form the spar to frame interface

on each side of the fuselage, where four frames are attached to nine spars in each wing. In

the case of the high aspect ratio (10.6) U-2 which had only three spars, the bending

moment was distributed trough twelve wing attachment joints (six each side of the

fuselage) which mated the wings to the wing root attachment ribs and hence to four main

fuselage frames which carried the bending moment across the fuselage. These heavy

carry through frames add considerably to the empty weight of the aircraft and mid - wing

position was considered to be too heavy and complex for such a small aircraft as

Terrasoar, although this configuration has been successfully employed in large scale

flying jet powered models of both the F-16 and U-2.

(c) Low – wing Position: - The major advantages of the low – wing position over both

high – wing, and mid – wing positions is in reducing undercarriage physical size and

weight, and when retractable undercarriage is considered this can be retracted into the

wing torsion box, which for most aircraft is the strongest component of the airframe

structure. Although for both the high - wing and mid – wing can employ a main wheel

undercarriage bow attached to the bottom of the fuselage for fixed undercarriage or

centreline bicycle undercarriage for retractable units. Additionally the low – wing

configuration requires dihedral for lateral stability as shown in figure 48(c) and a fillet at

the wing body interface to minimise drag inducing aerodynamic interference, and the

reasons for filleting are covered below.


66

Until the early 1930‟s most mono-planes were of the high – wing configuration which

was largely the result of the aerodynamic interference at the wing to fuselage junction

which was found to be worst in low – wing configuration. Putting a circular fuselage on

top of the wing has the effect of producing a pair of rapidly diverging surfaces which

steepens an already adverse pressure gradient almost guaranteeing flow separation and

inducing drag, and reduced lift, furthermore the separated flow could impinge on the

empennage horizontal tail resulting in further stability and control problems for the low –

wing configuration (covered in detail in reference 13). It was only through the discovery

of the beneficial aerodynamic effects of mounting a fillet at the wing body junction at the

California Institute of Technology CalTech in the USA that largely overcame these flow

separation problems that enable the low – wing configuration became widely adopted in

modern aircraft designs, and what reaming inferiority there was in the low – wing

configuration could be addressed by dihedral slopping for roll stability and was

compensated for by the reduction in undercarriage length and weight and the ground

cushioning effect on landing with the Wing In Ground WIG effect.

On balance though the high – wing configuration was felt to be a better solution for

alternative Terrasoar configurations submitted in stage one of this concept design phase

because the need for fairings would add to the complexity and cost of manufacture

requiring accurate tooling and assembly, and for the reasons stated in consideration of the

high – wing position.

The airfoil shape considered for these studies was a NACA four – digit airfoil NACA

4412, as most general aviation aircraft employ either four – digit, five – digit or 6-serise

airfoil sections – the laminar – flow series airfoils the choice is large, although the initial

choice made by the author seemed well suited in terms of depth for structure, and ease of

manufactured. The characteristics of an airfoil section are shown below in figure 49.

AERO 315

Airfoil Characteristics

Mean camber line

Chord line

Chord

x=0 x=c

Max thickness

Max camber

Leading edge Trailing edge

x

z

Figure 49: - Characteristics of an airfoil section (created for the authors IRP).


67

The characteristics of the airfoil section are defined by several shape parameters of which

the most significant are shown in figure 49 and include: -

1. The maximum thickness to chord ratio and its chordwise location:

2. The nose radius, which should be relatively large to give good maximum lift

coefficient CLmax:

3. The degree and distribution of camber, if employed, some degree of camber is

common for wing sections to enhance lift characteristics:

4. The trailing edge angle, which is usually made as small as possible within

handling and manufacturing constraints.

The NACA nomenclature is used to describe a wide range of airfoil sections in use today

these were developed in the 1930‟s through to the 1950‟s and below is the descriptive

nomenclature for the four digit airfoil section series:-

4 digit code used to describe airfoil shapes:

1st digit - maximum camber in percent chord:

2nd digit - location of maximum camber along chord line (from leading edge)

in tenths of chord:

3rd and 4th digits - maximum thickness in percent chord:

For example: NACA 2412 with a chord of 4 feet:

A max camber: 0.08 ft (2% x 4 ft):

Location of max camber: 1.6 ft aft of leading edge (0.4 x 4 ft):

Max thickness: 0.48 ft (12% x 4 ft).

For the preliminary design phase the most critical design parameters for this low speed

aircraft were the maximum lift coefficient, the drag coefficient, and the moment

coefficient, which were obtained from the NACA airfoil data charts, although

consideration was given to the following characteristics as advised in reference 12: -

Aircraft Conceptual Design Synthesis: Dr Howe. Denis: Published by: -Professional

Engineering Publishing Ltd: 2002, namely: -

A. The stalling characteristics where a gentle loss of lift is preferable, for light

aircraft and UAV‟s:

B. The airfoil pitching moment characteristics which may be particularly important

at high speeds causing a significant drag penalty:

C. The depth and shape of the airfoil with respect to the effect on structural design,

ease of manufacture, and possible fuel storage:


68

D. The slope of the lift curve as a function of incidence in that it affects the overall

aircraft attitude, especially at high values of lift coefficient, such as required on

landing.

Lift coefficient: -

Drag coefficient: -

Moment coefficient: -

Note: no dimensional coefficients!

Where: - L, D, and M are the actual lift, drag, and moment (positive nose up) acting on

the airfoil respectively, S is the airfoil reference area and c is mean chord (S divided by

the span b), V is the flight velocity, and is the local air density.

The choice of airfoil section is broadly based on the need to obtain the best aerodynamic

efficiency in the primary operating conditions of the aircraft which in the case of the

Terrasoar was cruising flight.

The maximum lift coefficient (CLmax) is considerably influenced by the nose radius of

the airfoil, decreasing as radius decreases. In the case of Terrasoar a low speed airfoil

would have a maximum lift coefficient in the region of 1.6, which the selected airfoil

with a CLmax of 1.4 obtained from the NACA charts was reasonably close to for initial

selection.

The forces, moments, centre of pressure, and aerodynamic centre on a typical wing are

illustrated in the following figures, 50 through 53 to enable the reader to appreciate the

significance of the defining airfoil coefficients: -

2

2

2

2

2

2

VcS

MC

SV

DC

SV

LC

pM

pD

pL


69

Figure 50: - Airfoil forces and moments (created for authors IRP).

Figure 51: - Airfoil centre of pressure (created for authors IRP).


70

Figure 52: - Aerodynamic centre 1 illustrates how the moment changes with location

(created for authors IRP).

Figure 53: - Aerodynamic centre 2 illustrates the point on the airfoil where the

moment is independent of the angle of attack (created for authors IRP).


71

Figure 54: - NACA 4412 chart analysis 1 (a) lift curve (b) drag curve.

In order to gain an appreciation of the airfoils behaviour the following parameters were

obtained from the NACA lift and drag curves for the NACA 4412 section at 120 angle of

attack and a Reynolds Number of 5.9x106 as shown in figures 54 to 56.

cl = 1.3

cl= ao = ( cl / ) = 0.083/0

cd = 0.018

cm c/4 =-0.025

cm a.c. = -0.04

cLmax= 1.4

stall = 150

l=0 = -20

(a) (b)


72

Figure 55: - NACA charts 4412 analysis 2 (a) lift curve (b) drag curve.

The variations of lift cl, drag cd, and momentum cm coefficients with angle of attack

and Reynolds number Re are shown in the NACA charts as lift and drag curves shown in

figures 54 to 56 (the same chart) which are based on actual experimental data obtained by

NACA in the 1940‟s (note in the majority of aerodynamic literature the lift, drag and

moment coefficients are in lower case when dealing with two-dimensional shapes as in

the case of airfoils).

The variation of lift coefficient with angel of attack is shown chart (a) in figures 54

through 56, note the lift curve of cl versus is linear over the most practical angle of

attack range, this is of great importance and is called the lift slope and is designated ao,

for thin military airfoils, a theoretical value for the lift slope is 2 per radian, or 0.11 per

degree is used in performance analysis (page 62: reference 13). This reportedly also the

case for most experimentally measured lift slopes of conventional airfoils, the lift slope

for NACA 4412 was measured as 0.083/0 from chart (a), which is reasonably close to the

theoretical value. Note from chart (a) at = 0 there is a positive value for the lift

coefficient i.e. cl = 0.1 and the airfoil must be pitched down to a negative angle of attack

denoted by l = 0 (in this case -2

0) for the cl lift coefficient to equal 0.

(a) (b)


73

Figure 56: - NACA charts 4412 analysis 3 (a) lift curve (b) drag curve.

This negative angle of attack is because NACA 4412 is a positively cambered airfoil, a

symmetrical airfoil would have a l = 0 value equal to 0

0 and an airfoil with negative

camber would have a positive value for l = 0, (negative cambers have not been used in

practical aircraft aerodynamics to date).

At high angles of attack the linearity of the lift curve diminishes due to the diversion of

the lift coefficient as it reaches its maximum value cl max in figures 54(a) through 56(a),

as increases further. The drop in cl is the result of separation of the airflow over the top

surface of the airfoil causing a loss of lift, but over the linear portion of the lift curve the

flow is attached over the majority of the airfoil surface. The lift curve variation of cl with

Re are shown for Reynolds numbers over a range of 1.0 x 106 to 8.9 x 10

6 in figures 54

through 56, and there appears to be very little effect at these high Reynolds numbers

associated with normal size wing chord lengths for normal flight.

(b) (a)


74

However at the much lower Reynolds numbers associated with model aircraft and many

small UAV‟s (in the order of 100,000) there is a substantial Reynolds number effect on

the lift curve which is reduced below the values shown here, according to reference 15.

However there was insufficient time within this project for the author to explore this

further and a full and detailed treatment of this issue should be given in reference 7.

The variation of momentum coefficient cm is shown to be essentially linear over most of

the practical angles of attack in figures 54(a) through 56(a) demonstrating that for this

airfoil the moment coefficient curve mo = dcm / d is essentially constant over the

practical angles of attack considered. Although the curve is positive for this airfoil this is

not always the case and for some airfoils it can be negative. As with the lift coefficient

the slope becomes non – linear at high angels of attack due to flow separation on the top

surface of the airfoil and also at low angles of attack due to flow separation on the bottom

surface of the airfoil. In common with the lift curve the linear portion of the momentum

curve is essentially independent of the Reynolds number values considered in these

charts.

The variation of cd with the lift coefficient is shown in figures 54(b) through 56(b) and

from these charts it can be seen that for a positively cambered airfoils of the NACA 4412

type the minimum value of the drag coefficient cd min dose not necessarily occur at zero

angle of attack but rather at some finite small angle of attack. This is because at very low

angles of attack for example -20 to +2

0 the drag is the result of friction and pressure drag

over the airfoil surface. At higher angles of attack the drag coefficient rises very steeply

(exponentially) due to flow separation over the airfoil which creates rapidly large

pressure drag values. Also shown in these charts is the effect of Reynolds number on cd

which unlike cl (at the Reynolds number range considered here) is sensitive to Re and is

larger at lower Reynolds number values, also because the Reynolds number influences

the extent and character of the separated flow region cd at high is also influenced by

the Re value considered.

Also shown in figures 54(b) through 56(b) is the variation of the moment coefficient

about the aerodynamic centre cm a.c. which is essentially constant over the range of lift

coefficients and hence the range of angles of attack considered.

Although this airfoil section was considered adequate for initial concept definition and

was carried through to the second stage configuration studies by the author the airfoil

selected for the final flying aircraft was the Wortmann FX 63-137 was intended for use in

applications where compressibility was not a problem which was the case with the

Terrasoar. But although maximum lift coefficients in the order of two were quoted for

this airfoil these were associated with a high zero lift pitching moment in excess of -0.1,

and the very thin reflexed trailing edge of the Wortmann airfoil was to cause difficulties

in structural design and manufacturing requiring some modification, additionally this

section was very sensitive to surface roughness at the scales used on Terrasoar

complicating manufacture still further.


75

The Wortmann FX63-137 airfoil is fully described in reference 7 including all data

charts, and is shown below in figure 57 for comparison with the NACA 4412 selected for

the concepts only this concludes discussion of the author‟s concept wing configurations.

Figure 57: - WORTMANN FX 63 – 137 Airfoil.

5): - Empennage configuration: - The conventional configuration for the empennage

unit would have been a horizontal fixed tail plane and a fixed vertical fin each being

fitted with a hinged rear flap acting as an elevator for pitch control and a rudder for yaw

control respectively, as shown in figure 58, and in some cases a dorsal fairing is

employed at the base of the fin to arrest the possibility of fin stall. However due to the

decision made above to adopt a pusher propeller / engine configuration for all piston

engine configurations this conventional layout was not possible without limiting the

effectiveness of the control surfaces as airflow drawn in by the propeller would

considerably restrict there movement in response to actuator inputs. Also in the case of

the jet powered configuration three the conventional fin would be in constant jet wash

and be rendered ineffective. These considerations left only two options open to the

author, of which: - one was the twin tail fin replacing the conventional single vertical fin

with a pair of fins located at the extremities of the horizontal tail, although this tends to

have mass (stiffening of the horizontal tails) and aerodynamic interference penalties (due

to fuselage obscuration in sideslip): and the second was the twin boom layout used in

configuration one, with twin fins located on the booms, which incurred a wing mass

penalty, and greater interference drag.


76

Figure 58: - Aircraft reference axes and corresponding aerodynamic moments.

Taking all of the above into account the author generated the four configurations GW 1

through 4 described above and submitted these for evaluation, as shown in figure 42 and

table 5 the results of which were generated in the author‟s absence due to sickness, by the

M1 module runner up team.

3.2: - Airframe configuration down selection.

The outcome of this initial concept was that only three of the original designs fared well

but these were compared with photographs of current flying UAV‟s which had hundreds

of man hours and millions of dollars spent on them therefore the decision matrix could

not be deemed a reliable indicator of the three original designs merit. The jet options

proposed by the author were always going to be high risk options on grounds of

endurance and in the field maintainability, but in the authors view the piston engine

configurations were worthy of further study and design maturation.

Of these three the authors Configuration 1 was judged to have the least risk associated

with the overall concept, however to reduce risk further for a fixed or retractable tricycle

undercarriage version two low wing configurations were explored by the author with a

reduced drag fuselage cross – section and were configurations 6(A) / (B) and 7

respectively described below. In the final analysis these would be compared with a

modified Intake 5 airframe developed by Mr James Pennington.


77

TACTICAL RECON UAV CONFIGURATION SIX (A).

Enlarged twin booms max 250 mm diameter.

Wing flight control surfaces, are

combined flaps and ailerons.

Tail group control surfaces are

conventional rudders and single

elevator.

Figure 59: - Stage 2 authors configuration 6(A) forward isometric view.

TACTICAL RECON UAV CONFIGURATION SIX (A).

Aerodynamic fins to support duct.

High tail of configuration 1.

Fixed undercarriage 450

mm legs, with 200 mm

diameter wheels.

Figure 60: - Stage 2 authors configuration 6(A) aft isometric view.


78

Figure 61: - Stage 2 authors‟ configuration 6A supporting data.

Configuration 6(A)/(B):- This configuration was offered by the author in a fixed

undercarriage (A) version and a retractable undercarriage (B) version, and with the

exception of the retractable undercarriage feature in all other respects the two

configurations were identical. A tubular fuselage of 400mm diameter with a keel and two

side longerons, four bulkhead frames for:- the engine attachment: fuel tank fire wall:

wing pick-up attachment: and nose wheel attachment: two intermediate frames separating

avionics and payload bay and power modules, was proposed. This fuselage skin would

have two sections a lower integral skin attached to the frames longerons and keel to

distribute the fuselage bending loads, and a top cover which started at the payload bay

frame and ran to the engine bay as shown in figures 59 and 60. The empennage booms

were sized to permit the undercarriage to be retracted into them on the 6(B) version but

due to the heavy landing case and the wing joint being at the boom attachment point the

size was kept common for both 6(A) and (B) versions. The wing was to be a two spar

structural wing (containing rubs rather than a block of foam) with the leading edge spar at

15% chord and the trailing edge spar at 75% chord, with attachment ribs for the ailerons

and flaperons, and fuselage attachment ribs with lugs which would be pinned through

longeron „H‟ extensions and covered by the fuselage / wing blend panels. These wings

would have top and bottom wing skins forming the torsion box broken at the inboard

wing / outboard wing joint, with a „D‟ – nose attached to the leading edge spar and „V‟

trailing edge cover attached to the trailing edge. A propeller FOD ring was also fitted to

prevent foreign object debris impacting on the propeller which could have been thrown

up form a rough field take-off and landing. The wide track robust undercarriage layout

promoted off airfield operations, and although the cylindrical fuselage would have

required avionics repackaging the form is used in many UAV‟s including the new Watch

Keeper Royal Army UAV‟s, but this was considered too radical and configuration 7

studies began.

GENERAL FEATURES OF CONFIGURATION 6(A) NEW BUILD.

Wing span: - 4.0 m Wing area: - 2.764 m2

Root Chord: - 1.0 m Tip Chord: - 300 mm L/E Sweep: - 11 deg O/A Length: - 2.6 m Fuse diameter: - 400 mm Propeller diameter: - 650 mm Tail plane area: - 0.525 m2

Fin area (each one): - 0.307 m2

Main wheel track: - 1.0 m Nose wheel to main wheel: - 1.720 m


79

ALTERNATIVE TERRASOAR CONFIGURATION SEVEN ISOMETRIC VIEW.

Figure 62: - Stage 2 authors configuration 7 forward isometric view.

ALTERNATIVE TERRASOAR CONFIGURATION SEVEN EXTERNAL FEATURES.

Tapered wing.

Split aileron.

Low – wing.

Split aileron.

Tricycle undercarriage

fixed (non – retracting.

Tapered fuselage

to reduce drag.

Tail booms built into

wing combined with

wing interface.

Wing root blend indicated shown

but feature not modelled.

Figure 63: - Stage 2 authors‟ key external features of configuration 7 forward

isometric view.


80

ALTERNATIVE TERRASOAR CONFIGURATION SEVEN EXTERNAL FEATURES.

Tail group control surfaces

are conventional rudders

and single elevator.

Pusher prop with rear mounted

engine with smaller propeller

diameter 800mm.

Thinner tail boom then

configuration 6 to reduce

drag.

Figure 64: - Stage 2 authors‟ key external features of configuration 7 aft isometric

view.

Configuration 7 external features:- This aircraft shown in figures 62 trough 64 was a

lager aircraft than configuration 6, with a wing span of 5m and a length of 3.1m, and a

higher aspect ratio wing of 9.8, compared with 5.8 for configuration 6, but was still of

tapered planform. The undercarriage was fixed for this configuration but retained a wide

track for rough field service use, and empennage boom diameter had dropped to 50mm

reflecting the use of filament wound CFC tube which would be much stronger and stiffer

than the structural boom proposed in configuration 6, although like the former

configuration they formed part of the inboard / outboard wing interface. The empennage

layout was the same basic arrangement as configuration 6 with a single elevator attached

to the horizontal tail, and two rudders attached to the two vertical fins mounted on the

booms, controlling pitch and yaw respectively. Roll was controlled by pairs of ailerons

mounted on each of the outboard wings and driven by independent actuators one for each

aileron, this provided a degree of control surface redundancy and dispensed with the need

for the flaperons proposed for configuration 6. The undercarriage was not faired on this

configuration to reduce weight but required proper sizing, the propeller diameter had also

been increased to 800mm on advice from Mr James Pennington who was responsible for

the propeller sizing activity, and clearance in take-off rotation still required examination.

As with configuration 6 there was a wing / fuselage blend cover which was removable for

wing attachment bolt removal. The fuselage was optimised for a 25kg payload proposed

by BAE SYSTEMS North America, and tapered over the wing to reduce drag.


81

ALTERNATIVE TERRASOAR CONFIGURATION SEVEN INTERNAL FUSELAGE STRUCTURE.

Payload bay.

XREA avionics bay.

Fuel tank bay.

Transponder and nav bay.

Power modules.

Engine bay.

Wing attach full frames.

Fire wall frames.

Split frames.

Split frame.

Figure 65: - Stage 2 authors‟ key internal layout details of configuration 7 fuselage

aft isometric view.

Configuration 7 fuselage internal features: - The configuration 7 fuselage was to

employ structural skin technology with a two part stiff skin carrying the fuselage bending

loads between frames. These frames would be a nix of horizontally split frames and full

depth frames with the latter being mechanically fastened to the bottom skin but not to the

top skin, and the split frames being fastened to both. The bottom fuselage would have

been fitted with a reinforced belt strip at the waterline interface joint between the top and

bottom skins and the top skin would be mechanically fastened through this strip using

anchor nuts and countersunk bolts. This top skin would be additionally stiffened by the

top half of the split frames which would have been mechanically fastened to it, this would

have enabled the top skin to be removed an one piece during ground handling for battery

charging, payload servicing, or avionics modification. Also of note is the provision for

the XRAY avionics crate although an alternative was under investigation, the customer

had requested this provision as a fall back solution. The nose wheel attaches to the

forward payload bay split frame, and the wing attaches to the three centre fuselage full

frames. Analysis may have shown the stiff skins to be too heavy, so ample provision for a

keel and longerons to be built into the design running under the systems, power, and

avionics trays, as shown in figure 65 above.


82

ALTERNATIVE TERRASOAR CONFIGURATION SEVEN INTERNAL WING STRUCTURE.

Aileron actuator bays.

Wing box leading edge ‘C’ spar.

Wing box trailing edge ‘C’ spar.

Split aileron.

Bottom wing skin.

Top wing skin.

Wing inboard / outboard joint

and boom interface.

Wing ‘D-nose’ locked by T tie.

Figure 66: - Stage 2 authors‟ key internal layout details of configuration 7 wing

forward isometric view.

Configuration 7 wing internal features: - This configuration also had a structural wing

like that proposed for configuration 6, (containing spars and ribs as apposed to a block of

foam), the wing torsion box had a leading edge „C‟ section spar at 25% chord and a

trailing edge „C‟ section spar at 75% chord with 6 outboard wing ribs and three inboard

wing ribs (wing to fuselages attachment ribs obscured by fuselage), as shown in figure 66

above. The wing to fuselage attachment ribs were to be full chord ribs attaching to the

full depth frames shown in figure 65, by means of six outboard / inboard pins three each

side passing through fork end lugs in the frames, and single tongue lugs on the ribs.

The inboard to outboard wing joint was achieved by the boom effectively hooking over

metal fittings that form slots for the overlapping torsion box spar webs, and the

undercarriage leg being screwed into a metal fitting at the attachment end of the boom.

Although this attachment methodology requires much further investigation the author

considered it as a good starting point.

Separate bays were provided for the aileron actuators bounded by the ribs to which the

ailerons attached to the wing spars, as shown above.


83


Wing span:- 5.0 m

Wing area:- 2.552 m2

Root Chord:- 750 mm

Tip Chord:- 300 mm

L/E Sweep:- 5 deg

O/A Length:- 3.1 m

Fuse diameter max:- 568.744 mm


Tail plane area:- 1.609 m2

Fin area (each one):- 0.402 m2

Boom Length:- 2.54 m

Boom diameter:- 50 mm

Figure 66: - Stage 2 authors‟ configuration 7 supporting data.

Figure 67: - General configuration notes for both configurations 6 and 7.

This concludes the descriptions of configurations 6 and 7 as presented for the final

selection review on the 4th

September 2003.

General configuration notes.

• All configurations had ailerons, rudders, and elevators.

• Both aircraft configurations 6(A) / (B) and 7 had tricycle undercarriage arrangement.

• All three configurations had been sized based on Terrasoar overall dimensions, and had been modified to meet the new mission requirements and need parametric analysis.

• The prop size had been reduced to diameters of 650 mm - 800 mm to reflect the new mission requirements and needed parametric analysis.

•These are all general arrangement CATIA V5 models and were only intended for configuration concept capture and not intended to form the basis of surface models for manufacture.

•Configuration 7 had the same twin boom layout with a different fuse and

a five meter wing span, sized for a 25 kg payload.


84

Figure 68: - A sketch of Mr James Pennington‟s final submission for the Terrasoar

redesign showing reconfigured 5m wing and fixed undercarriage this is fully

described in reference 7.

The final outcome of all of the aircraft concept design phase activity was that the 9th

Intake were asked to chose between developing the redesign of the existing Terrasoar

aircraft or developing the configuration 7 proposal of the author at a final down selection

meting on the 4th

September 2003 given that a meeting had already taken place between

BAE SYSTEMS and Cranfield University and decided that as much of the original

Terrasoar as possible would be embodied in the new aircraft. The final decision by the 9th

Intake was to peruse a reconfigured version of the 5th

Intakes Terrasoar design, but not to

just produce detail designs of the 5th

Intakes concepts as originally proposed by the

customer. Instead a compromise solution was found in an agreement to use the basic

configuration shown in sketch form in figure 68, as the basis for an evolved Terrasoar

which is at time of writing in manufacture.

4.0: - Terrasoar Wing Design (Phases 2 to 4).

With the final selection of a viable configuration the Terrasoar project entered phase 2 of

the design process framework as described in section 1.1 of this thesis, and the author

assumed responsibility for the wing design which was to be broken down into three

sections for transportation, and its interfaces with the booms, fuselage, and ailerons, as

well as the control surface actuator integration design this section covers the design of the

wing itself and section 5 covers the interface joint design, and section 6 covers the control

surface design and systems integration.


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4.1: - Structural definition for wing design.

The wing configuration selected was rectangular with a taper ratio of ct / cr = 1.0 with

initially with no twist, and the maximum design velocity of the Terrasoar UAV in normal

flight was estimated to be in the region of 85 miles / hour - far below the transonic

regime; and hence there is no aerodynamic requirement for any sweep in the wing. From

reference 13 the minimum induced drag would be produced from with an elliptical

spanwise lift distribution, which implies an elliptical planform shape. However as stated

above the higher production costs associated with a wing with curved leading and trailing

edges in the planform view could not justified in view of the cheaper costs of

manufacturing wings with straight leading and trailing edges. Moreover, by choosing the

correct taper ratio, the elliptical lift distribution can be closely approximated, however as

stated above this has not been considered necessary in the down selected design.

The rectangular wing chosen would have been heavier than a tapered wing of the same

span because as the taper ratio decreases the centroid of the lift distribution (centre of

pressure) moves closer to the root of the wing. In turn, the moment arm from the root to

the centre of pressure decreases, and the bending moment at the root decreases, and the

lift stays the same. As a result, the wing structure can be made lighter. This trend is a

benefit obtained from using a small taper ratio, and as will be seen in the next section this

was selected by the 5th

Intake.

However wings with low taper ratios (pointed wing tips) exhibit undesirable flow

separation and stall behaviour, because as the taper ratio decreases the region where flow

separation first develops moves out towards the tip and as a result aileron control is

adversely affected.

The rectangular wing of taper ratio = 1.0 selected here would develop flow separation

first in the root region. This location for flow separation has two advantages: - (1) If a

tractor configuration had been selected then the separated, turbulent flow trails

downstream from the root region causing buffeting as it flows over the horizontal tail,

thus serving as a dynamic stall warning to the controller, however with the twin boom

layout this indicator was lost. (2) The wing - tip region still has attached flow, and

because the ailerons (for lateral control) are located in this region, the pilot still has full

aileron control.

The conceptual design of aircraft structure required assessment of where major structural

members would be positioned within the aircraft. Careful positioning of structural

members could save significant structural weight and can greatly simplify manufacturing,

operation, and maintenance of the aircraft.

This thesis deals with the wing alone and the following overview covers the wing

structural layout for a conventional aircraft or large scale flying model which the author

intended to employ in the design of the Terrasoar wing.


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The main function of the wing is to pick up the air loads and transmit them to the

fuselage. The wing cross - section takes the shape of an airfoil, which is designed based

on aerodynamic considerations. The wing as a whole performs the combined functions of

a beam and a torsion member. It consists of axial members in stringers, bending members

in spars and shear panels in the cover skin and webs of spars.

The spar is a heavy beam running spanwise to take transverse shear loads and spanwise

bending. It is usually composed of a thin shear panel (the web) with a heavy cap or flange

at the top and bottom to take bending, this forms spars of either an „I‟ or „C‟ cross -

section as shown in figure 69.

Wing ribs are planar structures capable of carrying in-plane loads. They are placed

chordwise along the wing span. Besides serving as load re-distributors, ribs also hold the

skin stringers to the designed contour shape. Ribs reduce the effective buckling length of

the stringers (or the stringer - skin system) and thus increase their compressive load

capability. Figure 70 shows a typical rib construction. Note that the rib is supported by

spanwise spars.

The cover skin of the wing together with the spar webs forms an efficient torsion

member. For subsonic aircraft like the one considered here, the skin is relatively thin and

may be designed to undergo post buckling. Thus the thin skin can be assumed to make no

contribution to bending of the wing box, and the bending moment is taken by the spars

and stringers. Figure 71 presents two typical wing types used for subsonic aircraft the

bottom section consists only of spars to take bending and is the concentrated flange type,

the top section uses both spars and stringers to take bending and is the distributed flange

type which is the subject of this assignment.

1. Figure 69: - A typical basic spar construction (Reference 16:- Introduction to

Aeronautics: A design perspective, authors Brandt, Steven, A: Stiles, Randall

.J: Bertin, John, J: Whitford, Ray, published by American Institute of

Aeronautics and Astronautics Inc 1997).


87

Figure 70: - A typical rib construction (Reference 16).

Figure 71: - Typical wing cross - sections for subsonic aircraft: - (A) concentrated

flange type: and (B) distributed flange type, note front and rear spars are common

to both (Reference 16).

Every structural design problem is different, but the following general guidelines

suggested in reference 14 highlight the pitfalls which should be avoided and the goals to

strive for when laying out an aircraft structure.

(A)

(B)


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1) Never attach anything to skin alone. Even thick aluminium skin has relatively

little strength against point loads perpendicular to its surface. Pylons, landing

gear, control surfaces, etc., must be attached through the skin to major structural

components (spars, ribs, bulkheads, keels, etc.) within the structure.

2) Structural members should not pass through air inlets, passenger cabins, cargo

bays, etc.

3) Major load - bearing members such as spars should carry completely through a

structure. Putting unnecessary joints at the boundaries of fuselages, nacelles, etc.,

weakens the structure adds weight.

4) Whenever possible, attach engines, equipment, landing gear, systems, seats,

pylons, etc., to existing structural members. Adding structures to „beef up‟

attachment points adds weight. Plan the positions of major structural members so

that as many systems as possible can be attached to them and so that the

structures can carry as many different loads as is practical.

5) Design redundancy into the structure so that there are multiple paths for loads to

be transmitted. In this way, damage or failure of a structural member will not

cause the loss of the aircraft.

6) Mount control surfaces and high - lift devices to a spar, not just the rear ends of

ribs.

7) Finally structural layout is a very creative process and innovation can often save

weight, complexity and cost.

Initial thoughts for the Terrasoar wing structural members were as follows: - spars were

to be „C‟ section composite spars with integral flanges and webs, as these would be easier

to manufacture requiring a simple block tool draping tool and would be in keeping with

the scale of this aircraft, the intermediate ribs could also be back to back „C‟ sections

composite components adhesively bonded together at their common interface, and the

main fuselage / wing interface and aileron attachment ribs would be double sided

aluminium machining‟s, and the wing tip close out ribs would be single sided aluminium

machining‟s, the skins would be two piece with the top skin having an integral „D‟-nose.

As will be seen in subsequent sub – sections some of this initial concept survived into the

final wing design, but before the author started initial layout of the new wing an

examination was made of the 5th

Intakes wing to see if any lessons could be learnt from it

and employed in the new wing design.


89

4.2: - Review of the 5th

Intakes wing design.

The general configurations the fixed aerodynamic surfaces are shown in figures 72

through 77 from reference 3 and structural layout data from Holmes Graham MSc GDP

Thesis (not submitted): and the CATIA models of the 5th

Intake were used for this

review.

Figure 72: - Major fixed aerodynamic surface airfoil selection (reference 3).

Figure 73: - Wing configuration showing dihedral on outboard panels (reference 3).


90

Figure 74: - Aerodynamic twist built into the wing panels (reference 3).

Figure 75: - Major fixed aerodynamic surfaces (reference 3).


91

Figure 76: - Proposed wing manufacturing methods (reference 3).

Figure 77: - Transportation brakes and wing dimensions (reference 3).

1680mm 3200mm

1315mm

246mm 123mm


92

From the aerodynamic analysis and propeller diameter analysis undertaken and captured

in reference 7 for the alternative Terrasoar (which will from here on be called

Terrasoar New Build 1) the 5th

Intake wing planform and physical dimensions shown in

figures 72 through 77 were judge as incompatible with the revised mission requirements

covered in section 2 of this thesis. Indeed by the end of preliminary phase 2 design

studies for Terrasoar New Build 1 the wing span had been reduced from the 5m span of

the phase 1 configuration shown in figure 68 to 3.620m, which remained constant for the

rest of the project, this compared with the 5th

Intakes 6.560m, also the chord was

increased to a constant 0.362m over the entire span of the wing compared to the 5th

Intakes 0.246m centre panel chord, and 0.123m tip chord. These dimension changes

combined with the adoption of a conventional recovery on fixed undercarriage enabled

the ailerons and their actuators to be moved into the outboard wing panels to provide a

greater aerodynamic roll moment giving these control surfaces more authority.

The aerofoil selected was the Wortmann FX 63-137 because of its high potential lift

coefficient and experience gained with this airfoil section within Cranfield University on

UAV‟s of a similar scale to Terrasoar New Build 1, effectively nullifying the 5th

Intakes

airfoil selection of DAE 11 as shown in figure 72. The reason for the dihedral in the

outboard wing panels of the original Terrasoar design shown in figure 73 was not clear

because as stated above in section 3 this feature is usually associated with low wing

configurations to aid roll stability, high wing configurations are inherently roll stable and

although not found with Terrasoar New Build 1 often require a slight anhedral to over

come a high degree of roll stability, the author found it difficult to see the benefit of the

figure 73 configuration given the comparatively small roll moment arm and physical size

of the 5th

Intakes ailerons.

Given the 50 dihedral of the outboard wing panels the geometric twist of -2.5

0 to create

some degree of „wash out‟ was probably necessary to control the lift distribution and

hence the boundary layer separation and stall characteristics, but was initially judge

unnecessary on Terrasoar New Build 1, during design maturation described below a -50

geometric twist was incorporated in the outboard wing panels for the same reason.

Initially a similar separated inboard to outboard wing joint and boom to wing joint

philosophy to that of the 5th

Intake shown in figure 75 was considered for Terrasoar New

Build 1, because initially the combined bending moment of the wing and the twisting

moment of the booms was considered too high for a single joint but during maturation

described below a single joint was developed to handle both interfaces in one assembly.

The wing construction shown in figure 76 was not considered adequate for the Terrasoar

New Build 1 wing by the author and when the detail drawings (reference 3 sub ref: -

Appendix 4: Richard Halliburton: GDP MSc Thesis Terrasoar HALE UAV Wing and

Tail Boom Design: Cranfield University 2001) were examined the design philosophy

although possibly adequate for the 5th

Intake wing was not appropriate larger wing

developed for the Terrasoar New Build 1.


93

Because of the changes in wing span, propeller diameter, and boom separation the

location of the 5th

Intakes transportation joints shown in figure 77 were not appropriate

for the wing of Terrasoar New Build 1.

The final conclusion of this review of the work of Mr Richard Halliburton the wing and

tail boom designer of the 5th

Intake, was that although a good degree of design maturity

was presented in his thesis, design drawings and PDR presentation unfortunately non of

this could be carried across to the Terrasoar New Build 1 wing design.

4.3 (A) Status wing layout and structural definition for PDR.

From aerodynamic analysis covered in reference 7 the PDR (A) Status wing design was a

3.620mm straight wing, with an aspect ratio of 10, the transportation requirements being

met by breaking the wing into three sections as shown in figure 78, with load transfer

accomplished through forward and aft spars at 15% chord and 65% chord respectively.

The wing airfoil co-ordinates were sourced from the University of Illinois Airfoil Data

Site: - http://www.aae.uiuc.edu/m-selig/ads/coord/fx63137.dat the recommended data

site.

Figure 78: - (A) Status wing general arrangement model at PDR exit.

The outstanding issues from PDR were as follows: - mate joint philosophy, centre to

outboard wing sections, and centre section to fuselage: the manufacturing methods for the

wing and the rest of the airframe: aileron size: FCS integration: navigation light

integration: structural sizing.

http://www.aae.uiuc.edu/m-selig/ads/coord/fx63137.dat


94

The first choice for the wing structure following the work of intake 5 was the use of a

foam core for the wing sections to which the spars were to be bonded, this would have

imparted torsional rigidity under aerodynamic and manoeuvre loads, and ground

handling, which would be covered by a glass fibre skin for aerodynamic cleanness of the

wing outer mould lines, and to seal the foam from moisture ingress. Ribs were required

for attachment of wing to the fuselage and where the cavities for the aileron control

actuators were located (areas with on foam supporting them), as well as for attachment of

the booms for the empennage, and the structural key features of this wing are shown in

figure 79 below.

Figure 79: - (A) Status internal structural arrangement at PDR exit.

The attachment to the fuselage was to be achieved by two ribs with integral fore and aft

lugs aligned with the fwd and aft spars as detailed in section four, these lugs located into

the channel section and bolted through as descried in section five.

The outboard wing structure is shown some detail below in figure 80 and all of these

initial models were produced as CATIA V5 part concept models of CATIA V4 models

which did not permit making individual components transparent, and for further detail of

the models generated at PDR the reader is referred to the models CATIA V4 held by both

Cranfield University and BAE SYSTEMS New Business Samlesbury.


95

Figure 80: - (A) Status outboard wing internal structural arrangement at PDR exit.

The central wing section structural layout is shown below in figure 81 illustrating the

location of the fuselage attachment and wing mate joint ribs.

Figure 81: - (A) Status centre wing internal structural arrangement at PDR exit.


96

The interface joint between the centre wing and the outboard wing sections was to be

achieved by either a tong and groove joint or a pin and lug system both of which are

detailed in section five of this thesis, and the proposed rib arrangement drawing is shown

below in figure 82.

Figure 82: - (A) Status Wing / fuselage interface attachment rib at PDR exit.

4.4 (A) Status wing layout and structural definition post PDR maturation.

Maturation of the PDR structural concept continued but the resulting wing was repeatedly

evaluated as over the parametric weight target, using foam core material that wound not

degrade making the aircraft un airworthy in a year or two after first flight. In the Design

Week of 07 – to – 11 June 2004 (during this week visits were made to a number of

manufacturing facilities to assess their ability to produce the airframe within the

Terrasoar budget and time scale), on a visit to Slingsby Aviation on 08/06/2004 (by

Manufacturing Engineering / Chief Engineer / Airframe Lead (Myself) and

Aerodynamics Lead / Finance Lead) the design and manufacturing teams were

introduced to a new light weight tough / high strength / flexible / long life foam core

material which could be shaped in complex curvature and cured at lower temperatures

thus not requiring an autoclave.


97

This material in thin sheet form had been used in sandwich skin construction with two

layers of glass pre preg each side to form wings and fuselage structures unsupported by

ribs or frames, of Classified Military UAV‟s, undertaking much demanding missions than

the Terrasoar aircraft, as well as in manned light combat training aircraft.

The Slingsby approach offered a considerable opportunity for weight reduction for the

wing / fuselage / and empennage major build components of the airframe, although the

costs of manufacturing at Slingsby was considered to be beyond the Terrasoar budget.

Therefore upon our return from this assessment visit Manufacturing was tasked by

Airframe to research the possibility of sourcing this material or similar with mechanical

properties data sheets, within BAE Systems, and to assess the viability of an in house

build of the vehicle. Airframe took the action of developing a post PDR (A) Status level

design study employing this build philosophy and to supply all design and materials data

to Stress Engineering. This was the most constructive outcome of the Design Week.

One of the most unfortunate outcomes of the design week was the insistence by Peter

Hamilton that the wing joint philosophy be changed, to his design, this was later to be

formally challenged by Senior BAE Systems stress engineers and found woefully

inadequate, and is addressed in section five.

During the post PDR (A) Status design study the presentation of which on the 17 / 11 /

200 is attached as Appendices 1, the following decisions were made by the author in the

role of the wing designer, in liaison with: - fuselage design, fuel team, flight control

systems, manufacturing and aerodynamics: -

1. Reduce landing speed and aid rotation by imparting a downward twist in the

outboard wing resulting twist in the outboard wing sections between 00 at the

inboard / outboard joint increasing to 20 at the outboard wing leading edge tips.

2. The wing would be raised on the fuselage to lie across the top of the fuselage box

and the fuselage / attachment should be integral to the wing, this provided a

substantial increase in the fuel tank volume if required.

3. The two spar structure would be retained as per PDR which is the only continuous

load path through the wing, and provide load paths and attachment points for the

ailerons.

4. The empennage booms would be attached inboard of the outboard wing section

mate joints and attached through retained inserts in the fwd and aft spars covered

in section four.

5. The four aileron layout would be retained from PDR, each with an individual

actuator this provided duplex redundancy in the aileron system, as a critical flight

safety case, actuator bay access cut outs would be made in the top outboard wing

skins and covered by removable panels.


98

6. The ailerons would be attached to the rear spar with pairs of discrete lugs as the

wing twist made the original piano hinge arrangement of PDR impractical, also

the author personally preferred discrete lugs as the piano hinges are prone to

jamming as the wing bends in flight. These would be covered by rubber seals.

7. The spars would be made of the same foam sandwich construction as the wing

and fuselage skins, and would be of C - section. These spars would then be

adhesively bonded to the wing skins and additionally secured with countersunk

insurance fasteners into internally mounted nut plates for the top skin and anchor

nuts for the bottom skin. The sub structure would be assembled in basic jigs prior

to skin attachment, and the skins would be two part top and bottom skins with an

integral leading edge as part of the top skin panels, landing on the bottom of the

leading edge spar.

8. Navigation lights would be fitted in the outboard wing section leading edges.

9. The wing tip ribs would be machined aluminium details.

10. Ribs of the same construction as the wings would be used to strengthen aileron

actuator bays by local stiffening, and form a landing for the access panels. Ribs

would also be used as torsion close outs for the three wing sections.

The resulting these studies design were presented at GDP review of the 17 / 11 / 04 and

the detailed presentation was included as Appendices 1 in the GDP Interim report, and

the external configuration is shown below as figure 83, with the internal structure shown

below as figure 84 for reference, as well as the common wing skin and spar construction,

shown for the outboard wing in figures 85 and 86.

At the close of Phase 1 the Airframe design team still lacked any stress data on which to

produce reliably sized detailed parts and material data was being obtained. For design

graphical representation a total sandwich thickness of 4mm was considered realistic.

Following the GDP meeting of the 17 / 11 / 2004 as Airframe Lead the author chaired an

internal design review of the post PDR (A) Status design study wing design with the

Aerodynamics Lead and the Fuselage Design / Systems integration Lead, this covered the

design issues raised in the minutes of the formal GDP meeting namely: -

1. The clip attachment to the fuselage of the wing:

2. The attachment methods for the outboard wing sections:

3. The attachment methods for the empennage booms to the wing:

4. Weight reduction of the airframe in its current non stressed state:

5. No issues were raised with the material or basic structural layout of the wing:


99

Figure 83: - The Post PDR (A) Status design study exit wing, external configuration.

Figure 84: - The Post PDR (A) Status design study exit wing, internal configuration.


100

Figure 85: - The Post PDR (A) Status design study 0utboard wing, and external skin

configuration detail.

Figure 86: - The Post PDR (A) Status design study 0utboard wing and internal

structure spar configuration detail, common to all spars inboard and outboard.


101

4.5: - (B) Status wing layout and structural definition for CDR.

To address the issues raised from the Post PDR (A) Status wing design study the

Aerodynamics Lead suggested a wide rectangular single spar structure would eliminate

the need for the additional ribs, but the author cautioned that a weight trade study would

be required when initial sizing data was available as this spar would have a greater

weight than two modest C – section spars, and would require a reinforced lap joint, as the

spar would be wrapped around a male mould tool. Stress Engineering was been actioned

to study this option, with no result.

Additionally weight reduction could be achieved by adopting a common interface joint

for both the outboard wing sections and the empennage booms although this had been

examined before with a common rib, a fresh study was required as the method of wing

construction had radically changed over the primitive solid foam block method favoured

by Cranfield University representatives and intake 5.

The wing / fuselage attachment philosophy was wrong and required urgent review, as

well as the fuel requirements, as parametric analysis undertaken to that date appeared to

be very conservative, and the Airframe team still had no test data of any kind to support

the Post PDR (A) Status design study fuel tank volume, subsequent testing covered in

reference 8 supported the view of the Airframe team that the initial figures were over

conservative, therefore the decision made to lower the wing back into the fuselage was

correct. The suggestion was made that the wing could be attached to the engine fire wall

and the forward fuel tank frame by two continuous sleeved bolts running forward to aft,

through the spar, and the Aerodynamics Lead and Fuselage Design Lead were charged

with investigating this.

Airframe Design was charged with producing models to reflect the outcomes of this

meeting, and this resulted in the Interim Maturation concept.

The author was on a Cranfield Module S02 and had an extensive F-35B Phase 1 exit /

Phase 2 entry Vertical Tail Customer Presentation / Design Review commitments

increasing an already full F-35 workload, and could offer advice but not design time to

this modelling effort, and this role fell to James Pennington the Aerodynamics Lead, who

was given a two week design window by his IPT Lead namely Eurofighter Typhoon.

Following the authors return from the Christmas holiday the author instigated an “ALL

HAND‟S” Airframe Review meeting on the 7th

January 2005, minutes of which were

submitted as Appendices 4 of the GDP interim report. The complete airframe was studied

at this meeting. No problems with the wing design were foreseen at this meeting though a

discussion of the use of the inboard ailerons as flaperons with FCS Lead / Wing Design

Lead / and Aerodynamics Lead (from whom the request came) to decrease the landing

speed and aid rotation, resulted in an initial enlargement of these surfaces. This Interim

Maturation design study wing layout and internal structural designs shown in figures 87

and 88 for reference.


102

The design of the double duty inboard / outboard wing and wing / boom joint required

some refinement as shown in section five but the general concept was approved to go

forwards to the next GDP review of the 19 / 01 / 2005 and the full presentation for this

meeting was submitted as Appendices 2 of the GDP interim report.

Figure 87: - The Interim Maturation wing, external configuration.

The key features of the Interim Maturation wing were as follows: -

The inboard ailerons had been extended to the inboard / outboard wing interface and use

the same attachment philosophy as employed in the Post PDR (A) Status design study,

and the outboard wing sections were also been extended to the common interface joint.

The aileron actuator wires were to be run through forward and aft tubes in the rectangular

spar as shown in section five, with interface connectors held in place at the wing interface

joints (aiding rapid assembly at the test site) this effectively separated the inboard and

outboard control circuits as a safety case, as proposed for the previous wing design. Also

this eliminates the requirement for a cut out in the spar the fuselage interface.


103

Forward and Aft sleeve bolts were used to attach the wing to the fuel tank frames, and

hence affect the wing / fuselage interface. The aerodynamic interface and fuselage seal

was affected by the use of rubber seals as shown in figure 5, these seals must be resistant

to the effects of fuel spillage and a suitable material was being investigated by the

manufacturing team.

Figure 88: - The Interim Maturation wing, internal configuration.

As per the previous study rubber aerodynamic blade seals were to be used to cover the

aileron / flaperon / wing interfaces to maintain the clean wing airflow and hence reduce

drag. Also following the previous design the aileron actuators were be accessed through

reinforced cut outs in the wing top surface and top spar flange immediately above the

actuator, the access panel would then be clipped into place by inserting a forward tongue

into the forward skin mounting and secured by a rotating clip at the aft end of the panel

would be affected by a counter sunk fastener as shown in figure 89.

The integration of the navigation lights was ongoing and a presentation was to be

available at the end of February 2005.

Box spar

Box spar

Box spar

Combined wing / boom joint ribs (2 off)

Fuselage interface rubber seals.

Sandwich structure used in both spar and skin as per post PDR design.

Fwd /Aft sleeved bolts (2 off)


104

Interim Maturation configuration design study was concluded with a wing structural

review on the 31st January 2005, and however a fundamental redesign of the wing

substructure was required after the manufacturing maturation meeting on the 15th

February 2005 which evaluated the tooling requirements for the Interim Maturation wing

configuration as being unsustainable within the GDP funding cap of £60,000 maximum,

and this lead to the crash Pre-CDR revision design program to meet the Terrasoar Critical

Design Review set for the 6th

April 2005.

Figure 89: - Aileron integration and actuator access panel.

The Pre-CDR revision design study was launched by the author and looked to develop an

outboard wing sections that could be manufactured using uncomplicated tooling enabling

manufacturing within the GDP budget constraints, devise a systems integration solution

for the aileron actuators, and wiring, and a less complex interface for aileron attachment.

At this stage manufacturing believed that the wing centre section box spar could be

produced with a collapsible mandrel because this spare unlike the outboard wing sections

had no twist, using the foam core sandwich construction shown in figure 86 and this Pre-

CDR revision wing is shown below in figure 90. As a manufacturing easement the

section skin panels were split, to form a „D‟-nose and a rear wing unitised section as

shown in figure 91 below.

Top skin Top skin

Aileron actuator

Tongue

Rotating clip

Access panel

Spar

Bottom skin

Counter sunk bolt

FWD


105

Figure 90: - (B) Status Pre CDR Revision study wing, external configuration.

Figure 91: - (B) Status Pre CDR Revision study wing, internal configuration.


106

Figure 92: - (B) Status Pre CDR Revision outboard wing, external configuration.

Figure 93: - (B) Status Pre CDR Revision outboard wing, internal configuration.


107

Figures 91 through 94 show the authors redesign of the outboard wing sections in which

the lessons learnt from both the PDR wing and Post PDR (A) Status design study were

applied.

The wing spar reverted back to separate „C‟-section forward and rear spars which had the

same consolidated sandwich structural construction as shown in figure 86 above,

consisting of two glass cloth plies (each ply being 0.25mm thick) each side (4 plies in all)

of an Alcan Airex R63.80 foam sheet (3mm thick), to give a total spar thickness of 4mm,

for both web and flanges.

These spares were adhesively bonded with flanges inward to a Styrofoam core housing

the actuator mounting plate this core had been machined to the Wortmann FX63-137

airfoil section, and the outboard interface joint, and tip rib. The Styrofoam „D‟-nose was

then adhesively bonded to external face of the front spar web, and the trailing edge

Styrofoam section with the aileron spigot and ailerons was then bonded to the rear spar

external face web, and all interface gaps would be filled with standard aircraft filler. The

resulting structure is shown in figure 93 above.

Finally a 2 ply skin was wrapped over the resulting structure and adhesively bonded to it

under vacuum bag pressure, which would give the skin a bagged finish which at this time

was deemed aerodynamically acceptable, and is shown in figure 92 above.

Figure 94: - (B) Status Pre CDR Revision outboard wing, aileron integration.


108

Shown in figures 92 and 94 are the changes made to reduce the complexity of the wing /

aileron interface, the multiple hinge approach was dropped by the author in favour of a

continuous steel spigot covered with a plastic sleeve, which was mounted in a potted

insert in the trailing edge foam and ran to a location cup machined into the tip rib, prior to

fitting the ailerons were slipped onto the spigot (each aileron detailed in section 6 having

a hole through its foam core) and then the assembly was adhesively bonded into place as

explained above.

Figure 94 also shows the method selected by the author to mount the aileron actuators as

a combined unit on a single aluminium single side minimum thickness machined plate

each actuator being screwed into spacer blocks prior to mounting on the aluminium plate

securing brackets as detailed in section 6 of this thesis. The physical dimensions for the

ailerons being supplied by Mr James Pennington and Miss Rachel Cunliffe, based on

there detailed analysis which is reported in references 7 and 10. The cover plate for

access to the aileron actuators was intended to land on a picture frame but this had not

been modelled by CDR, and as explained below this was later abandoned in favour of the

cover landing directly on to the spars which were given local reinforcement.

During the Pre CDR Revision design process the author with Mr Paul Gilligan the

fuselage designer held manufacturing briefing meetings at Mellor House the metallic

structures machining centre for BAE Systems Samlesbury to discuss the wing interface

joints, boom attachments, tip ribs, and fuselage frames, and their advice was used to

mature the design of these components. For the fibreglass sandwich structural

components and general fibreglass manufacture and assembly, similar manufacturing

briefings were held with Mr Bob Cross and Mr Craig Carr of the Man Tech (advanced

manufacturing facility) at BAE Systems Samlesbury, and their advice was used in

maturing these components.

The degree of maturity in the design at this point still in the absence of any stress data

lead to the upgrading of this design to (B) Status under the Terrasoar Design manual

guidelines detailed in 1.3. This design was submitted for the Critical Design Review of

21st April 2005, and this design was passed by Cranfield University and BAE Systems

representatives for funding for manufacture subject to detail stressing and pre-production

modification.

Subsequent to the CDR the date of 23rd

May 2005 was set as the stress data release date

from Mr Phil Stocking and Mr Robert Sneddon to the Airframe Engineering team

represented by the author as Airframe Lead and Wing Designer, and Mr Paul Gilligan as

the Fuselage Designer, and Mr James Pennington as the Empennage Designer.

At this meeting Mr Phil Stocking met each designer in turn to discuss the structural sizing

he and Mr Robert Sneddon had produced and the design changes he felt necessary in the

CDR design so that the final design could be signed off for manufacture at the next

meeting on the 31st August 2005.


109

4.6: - (C) Status revised CDR wing incorporating stressing.

At this meeting detail sizing and ply orientations were released as shown in table 6 for the

wing, the fuselage and boom / empennage structures being covered in references 6 and 7

respectively. In addition to these certain design modifications to the centre wing section

were requested, as well as a support lug for the aileron spigot, and to the wing boom

interface boom connectors. The wing sizing, centre wing changes and the addition of the

spigot lug, and their implications are the subject of the following sections, the boom

interface changes are covered in reference 7.

Table 6: - Stress data for the Terrasoar post CDR wing using wrapped skin.

Component Number of plies Orientation

Outboard wing skins 6 Tool face: - 0/90/0/90/0/90

Outboard wing spars 4 Tool face: - -45/+45 foam -45/+45

Inboard wing skin core 9 Tool face: - 0/90/0/90/0/90/0/90/0

Inboard wing skin fwd 6 Tool face: - 0/90/0/90/0/90

Inboard wing skin aft 6 Tool face: - 0/90/0/90/0/90

Inboard wing spars 4 Tool face: - -45/+45 foam -45/+45

The centre wing section changes requested were basically a major redesign eliminating

the box spar and substituting for it the same forward and rear spar and three piece foam

structural philosophy as used in the outboard wing sections. The skin panels would also

have to be bridged to carry the torsion loads which would not be carried by the „C‟

section spars, and this bridge or core skin section would require the thickest ply lay-up of

9 plies. In order to achieve this localised reinforcement the skin would need to ramp up at

1 in 20 shallow ramp gradient, with a resulting impact on the spars of this new centre

section, also these spars required local reinforcement where the wing / fuselage

attachment bolts penetrated the spar. By contrast the proposed changes to the outboard

wing consisted only of changes in skin thickness and localised spar reinforcement around

the spigot lug rear spar penetration. The final wing design that was signed off on the 31st

August 2005 is shown in figures 95 to 99 below.


110

Figure 95: - (C) Status Post CDR wing, with stress sizing at sign off.

Figure 96: - (C) Status Post CDR centre wing skin, at sign off.


111

Figure 97: - (C) Status Post CDR centre wing internal structure, at sign off.

Figure 98: - (C) Status Post CDR centre wing outboard external, at sign off.


112

Figure 99: - (C) Status Post CDR centre wing outboard internal, at sign off.

Following the wing sign – off the wing skins and structural components were weighed

using the CATIA V4 model and the weights function the results of which are presented in

section 7. The Outer Mould Line surfaces were then released to Man Tech by the auth so

that they could commence tool design for the skin tools, and the pre - manufacturing (C)

Status model review was held on the 4th

October 2005.

4.7 Detailed design for BAE Systems manufacturing.

The major design decisions that resulted from the Man Tech (C) Status model review of

the 4th

October 2005 were as follows:-

Elimination of the ramps in the skin and spars of the centre section of the wing by

making the 9 ply zone continuous between the port and starboard wing joints:

Elimination of packers between the wing interface joints and the skin glue line

IML by locally growing out the interface joints:

Elimination of a separate foam packer to fill the gap under the actuator mounting

tray by use of an integral packer built into the outboard bottom wing skins:

Changing wing skin manufacture form the wrapped philosophy of the GDP wing

to two piece moulded skins to improve OML surface finish:

Wing / fuse attachment bolts to run normal to the attachment frames to ease

assembly:

Increase size of actuator cover plate to land on locally reinforced spars

eliminating the need for a separate landing frame:


113

Figure 100: - Manufacturing maturation of Centre wing section structure.

Figure 101: - Manufacturing maturation of Centre wing top skin.

9 PLY CONTINUOUS ZONE TO SKIN LIMITS.

6 PLY AFT CONTINUOUS ZONES TO SKIN LIMITS.

6 PLY FWD CONTINUOUS ZONES TO SKIN LIMITS.

SKIN RAMPS


114

Figure 102: - Manufacturing maturation of Centre wing bottom skin.

Figure 103: - Manufacturing maturation of Outboard wing bottom skin.

9 PLY CONTINUOUS ZONE TO SKIN LIMITS.

6 PLY FWD CONTINUOUS ZONES TO SKIN LIMITS.

6 PLY AFT CONTINUOUS ZONES TO SKIN LIMITS. SKIN RAMPS

SEALS

6 PLY OUTBOARD WING SKINS

INTEGRAL ACTUATOR PLATE PACKER


115

Figure 104: - Manufacturing maturation of Outboard wing top skin.

Figure 105: - Manufacturing maturation of Outboard wing section structure.

6 PLY OUTBOARD WING SKINS

SIZE OF ACTUATOR COVER PLATE INCREASED TO LAND ON SPARS.


116

Figure 106: - (C) Status development aileron structure.

Extension of the tip foam core to the tip rib:

Simplification of the tip rib navigation light integration:

And finally an initial definition of the aileron structure.

These modifications were duly incorporated by the author into the signed – off model and

are shown in figures 100 through 105 above, and the (C) Status structural configuration

for the ailerons is shown in figure 106 above, and the final manufacturable foam core

sandwich spar model was submitted to the Man Tech design review meeting on the 13th

October 2005. At this meeting the author was informed that the results draping tests on

the Alcan Airex R63.80 sandwich core foam had revealed problems in achieving the bend

radius required to manufacture the spars and a 4 - ply class only spar was proposed as a

solution to this problem. This resulted in a further IML modification of the spars for all

three sections of the wing and changes to the foam core OML‟s the resulting spar

thickness having decreased by 3mm, these modifications were conducted by the author,

and a final model was submitted to Man Tech on 12th

November 2005, for IML tool

design, which was the responsibility of Man Tech.

The responsibility for further design maturation of the aileron was formally passed to

BAE SYSTEMS New Business, Samlesbury Design Team, headed by Mr Peter

Hamilton, as was the responsibility for any minor wing modifications, on the 14th

November 2005.

4 PLY GLASS AILERON SKIN

4 PLY GLASS AILERON CLOSEURE SKIN IS NOT SHOWN FOR CLARITY

FOAM AILERON CORE

COMPOSITE AILERON DRIVE HORN

PLASTIC AILERON SPIGOT SHEATH


117

5.0 Wing interface joint concept assessment and selection.

5.1: - PDR Trade Studies and (A) Status schemes.

As stated in section 4.3 above the major structural interface joints were a major

unresolved issue at PDR and the initial PDR and the author focused on the following

concepts for joint types: -

1. Wing to fuselage interface joint: - rib lugs with rose bearings attached by 5mm

diameter bolts through fuselage „U‟ section longerons which was the PDR

Submission Presentation, and is shown as figure 82 page 96, in this thesis.

2. Outboard wing sections to wing centre section interface joint: - aluminium tongue

extensions bonded and mechanically fastened into outboard wing spars and the

Rohacell foam core of the PDR wing these would then be slotted into cavities

between the centre wing spars and the foam core of the centre wing and would be

secured in place by counter sunk retaining bolts through the top and bottom wing

skins in to potted threaded inserts in the foam core.

3. The boom to wing interface joint: - initial consideration of this joint was to use

potted threaded inserts through the forward and aft spar flanges and into the

Rohacell foam core, to secure bolts which would attach through flanges in the

boom ends. A secondary concept was to use a common rib as both the outboard

wing section / wing centre section interface onto which the booms were pinned,

this concept was rejected out right because advisors felt skin stresses could not be

transferred across the joint however all of the stresses would have been carried by

the spars anyway.

5.1.1: - Post PDR (A) Status design study.

As stated in section four, after the visit to Slingsby Aviation the structural build

philosophy moved away from the heavy Rohacell core wing, and toward the much lighter

and more practical foam sandwich philosophy with interface ribs and sandwich skins.

This in turn impacted the interface joint philosophy and was also seen as an opportunity

to increase the fuel reserves to a more conservative level, than that which could be

accommodated in the PDR design.

The studies perused in the design week were as follows: -

1. Wing / Empennage Boom interface joint (author and James Pennington).

2. Wing / Fuselage interface joint (author and Paul Gilligan / James Pennington).

3. Centre wing / Outboard wing interface joint (author).

The results of these studies were carried over into the Post PDR (A) Status design study

detailed above in section four.


118

5.1.2: - (A) Status Wing / Empennage Boom interface joints.

At this stage the Wing to Empennage / Boom interface was still a discrete joint inboard of

the wing interface and the Concept 1 pre – PDR study was a rib with lugs through the

forward and aft spar flanges and the bottom wing skins to which the Boom‟s were bolted

in a similar way to that originally proposed for the wing to fuselage interface joint at

PDR. This proposal in spite of being simple and practical was considered as too heavy for

this aircraft.

(a)

(b)

(c)

Figure 107: - Concept 2 / Wing to Empennage Boom interface joint, (a) lugs

attached to forward and aft spars: (b) channel section boom attached to lugs: (c)

boom skin covering joint.


119

A lighter alternative to Concept 1 was proposed and is shown above in figure 107 this

was Concept 2 and utilised the same Boom arrangement as Concept 1 but the rib in the

wing was eliminated and instead the lugs were bolted directly through the spars which

would be locally reinforced. Plastic blocks filling the local spar „C‟ –section, would

house threaded inserts for the lug bolts to screw into, an alternative to this block

arrangement would have been anchor nuts adhesively bonded to the inside face of the

bottom flanges of each spar. This concept was carried through to Post PDR (A) Status

design study.

5.1.3: - (A) Status Wing / Fuselage interface joint.

Early in the design week Mr Peter Hamilton insisted that the longerons should be

eliminated from the fuselage design so the largest possible fuel tank could be housed

(whether required or not), and reliance should be placed in a stressed 2mm thick CFC

skin, although this skin had a large removable access panel, which did not bode well for

structural integrity.

After the visit to Slingsby Aviation and the identification of the new high rigidity

sandwich construction, elimination of the longerons which only spanned one bay

appeared to be a real possibility. The wing could then be raised above the fuselage so as

not to restrict the fuel tank growth (as stated in section three the fuel tank sizing was

based on an estimate because we still had no SFC data). James Pennington suggested a

hood attachment clipped to the fuselage sides with quick release latches which Paul

Gilligan agreed to source, I proposed that we would need a very good into wind lap joint

with blade or P - seals to cover the joint. The major problem would be drag and the

author and Mr James Pennington spent some considerable effort to create a blend, the

Globe Master 1 and the Shorts Sky Van had similar problems, but Terrasoar was

complicated by the need to blend over a short distance between the wing leading edge

and the forward fuel tank frame which was also the landing for the FCS removable access

panel. Neither the author nor Mr James Pennington were satisfied with the end result but

time was limited so this Concept 1 was presented at the GDP design review of the 17 / 11

/ 2004 as shown in figure 108 below.

Initially the Concept 1 hood was to be integral to the bottom wing skin but the complex

curvature and requirement for large bend radii rendered this solution impossible,

therefore the author proposed bonding the hood to the bottom wing skin and reinforcing

this joint with bolts through the forward and aft spars.

This concept was rejected by the Customer: - Cranfield University so no further work

was done on this concept study which was closed. This has lead to the greatly improved

and more efficient interface joint in the Interim Maturation design study described below.


120

Wing / fuselage interface bonded to lower

skin and bolted to spars. Fitted with 4 quick

release clips to attach to the fuselage.Leading edge C section spar

broken at wing joints.

Trailing edge C section spar

broken at wing joints.

Torsion box closure ribs 2 places.

Wing section interface

joint (4 places).

Figure 108: - Post PDR (A) Status Wing / Fuselage attachment Concept 1.

5.1.4: - (A) Status Outboard wing / Centre wing interface joints.

After the Slingsby Aviation visit the author considered the PDR interface joint proposal

to be un - workable and started to study alternatives, at this time in the authors F-35 role

he was responsible for the wing fold joint, so an initial thought was to employ some of

the philosophies used in this key responsibility to create not a folding wing of course but

a multi lug pinned interface joint this is shown in figure 109. Ribs mounted in both the

outboard and centre wing sections have double staggered lugs for top and bottom pins

which are inserted through a cut out in the leading edge and screwed home into retained

nuts mounted in the trailing edge, and the skins extend over the interface with the

outboard skin slotting under the inboard skin to form a lap interface. A panel is then fitted

over the leading edge and secured with aero tape. This was Concept 1, however on the

scale at which we are working this solution although big and workable on the CATIA

screen becomes more like watch making in practice, and the integrity of the leading edge

cover panel would be in doubt, so this concept was shelved pre – Phase 1.


121

Figure 109: - (A) Status Outboard Wing / Centre attachment Concept 1, in which

the outboard wing and centre wing lugged ribs are connected with two common

interface pins inserted through the leading edge (Starboard wing interface shown).

Concept 2 was to form a plug and slot interface between ribs mounted at the ends of the

torsion boxes of both outboard wing and the centre wing, this is shown in figure 110, a

quick release latch across the bottom tension skin of the interface would ensure there was

no risk of separation in flight. Because the top wing surface in flight is in compression

and the plug is nearly full chord a latch across to top skin would not be required.

Although this solution would have increased the wing weight by a small margin

minimum machining would have kept this low and it could have been traded off against

other areas of the structure. This concept was loosely based on some joints the author had

designed for the F-35 Horizontal Tail test box.

Centre wing section

rib with two lugs.

(Starboard)

Outboard wing sections rib

with three attachment lugs.

(Starboard)

Pin insertion from

wing leading edge.

Attachment pins.

Interface


122

Figure 110: - (A) Status Outboard Wing / Centre attachment Concept 2, in which

the aluminium outboard wing rib has a tongue which slots into the aluminium

centre wing rib, and the joint is secured by a latch across the bottom skin.

(Starboard wing interface shown).

At the insistence of Peter Hamilton who believed there would be major corrosion

problems with the Concept 2 between the aluminium ribs and the “Carbon” skins, but the

author was not and never have proposing carbon fibre for the wing. He brought over one

of his „designers‟ from Nimrod and between them a wing joint I was unhappy with was

concocted and is shown here as Concept 3 in which bars pinned and wire locked through

„aluminium‟ inserts with class cloth backing (the wing spar skins are glass cloth but

neither of them took any notice) inserts carried the wing torsion loads, and wing bending

and were to be installed in both the leading forward and aft spar interfaces.

Concept 3 unfortunately was the only one to see visibility at the GDP meeting of the 17 /

11 / 2004 and was rejected by a much more experienced stress engineer who was present

to offer his assistance. This concept is shown below as figure 111.

Centre wing slotted rib

(Starboard).

Outboard wing tongued rib

(Starboard).

Latch across bottom tension

skin (omitted for clarity).

Hard mate joint at

rib interface.


123

Interface attachment pin 4 parts size TBD.

Locking pins 8 parts size TBD.

Outboard spar

sizing TBD.

Outboard aluminium

inserts 4 parts size TBD.

Inboard spar

sizing TBD.

Inboard aluminium

inserts 4 parts size TBD.

Outboard wing joint Concept 3.

Figure 111: - (A) Status scheme of Peter Hamilton for the Outboard Wing / Centre

attachment Concept 3 (rejected by all who have seen it).

5.2: - (B) Status Interface Joint Design Studies.

As a result of the post GDP wash up meeting revisions were made to all of the interface

joints, resulting in the combining of the boom interface and the wing interface joints into

a common interface, and the lowering of the wing into the fuselage, as stated above the

author was on a Cranfield Module S02 and had an extensive F-35B Phase 1 exit / Phase 2

entry Vertical Tail Customer Presentation / Design Review commitments increasing an

already full F-35 workload, and could offer advice but not design time to this modelling

effort, and this role fell to Mr James Pennington the Aerodynamics Lead, who was given

a two week design window by his IPT Lead namely Eurofighter Typhoon.


124

5.2.1: - (B) Status Wing / Fuselage interface joint development for CDR.

The wing / fuselage interface joint has been considerably simplified and relies on two

sleeved bolts through the fuel tank frames as shown below in figure 112 and has been

approved by BAE Systems stress engineers and Terrasoar manufacturing, this joint still

requires assessment from the Fuel Lead to determine impact on fuel capacity.

Figure 112: - (B) Status Wing / Fuselage interface joint Concept 1, reviewed at the

Terrasoar „ALL HANDS‟ meeting of the 07 / 01 / 2005.

5.2.2: - (B) Status Wing / Empennage Boom and Outboard Wing interface joint for

CDR.

The current double duty inboard / outboard wing and wing / boom joint was devised by

James Pennington required some refinement however the general concept is sound and

was approved to go forward to be presented at CDR. This consists of the Outboard wing

and the Centre wing torsion box ribs forming a common interface, with lugs which are

force mated by a slotted Empennage boom attachment which is located by pinning

through both the attachment and the rib lugs, as shown in figure 113. The recommended

improvement was to add the tongue and slot interface from the (A) Status studies as

shown in figure 114.

Fwd /Aft sleeved bolts (2 off)

Fuel tank frames


125

Figure 113: - (B) Status Common Wing / Empennage boom and Outboard Wing

interface joint Concept 1, reviewed at the Terrasoar „ALL HANDS‟ meeting of the

07 / 01 / 2005.

This would require a double sided machining of the centre wing ribs, which is possible

using the facilities at 25 Hanger BAE Systems Warton, even with the outboard tongue

having the same external dimensions as the internal ones of the outboard wing rib pocket.

There could however be some tolerance issues with the tongue and slot interface and the

alignment of the attachment pins but the author believes these can be resolved by

manufacturing tongue mate joint first and then drilling off the Empennage boom

attachment pin holes. After the post CDR manufacturing briefing meeting with Mellor

House Machining Centre representatives the author was assured that the joint in the form

shown in figure 114 could be manufactured on their machines with a wall thickness of

2.5mm and internal radii of 8mm.

This enhanced joint would eliminate the droop in the outboard wing sections which are

now 27% larger than the Post PDR (A) Status design study outboard wing sections,

during ground handling and manoeuvring prior to flight, and this has been verified by

studies showing that the current wing will only have a 1.5mm inboard / outboard wing

gap in ground manoeuvring. Once airborne off course these gaps will be sealed by the lift

on the wing as shown in figure 115.

Empennage boom skin

Centre wing box spar

Centre wing rib

Outboard wing rib

Empennage boom slotted end fitting

Attachment pins


126

Figure 114: - (C) Status Common Wing / Empennage boom and Outboard Wing

interface joint Concept 2 presented at CDR.

Figure 115: - The basic forces acting on an aircraft in flight (Reference 13).


127

5.3: - (C) Status Interface Joint Design Maturation.

Immediately post CDR wing sign – off Mr Peter Hamilton insisted that a series of joint

weight reduction studies be conducted to lighten the ignoring the fact that most of the

wing weight was in the skins, and his calculations were based on the assumption that the

author had used 17 glass cloth plies in the design of the wing spars which was a totally

wrong assumption because at every meeting the wing spar structure had been explained

as being a foam cored sandwich. One of the weight reduction studies he was most

insistent was that shown in figure 116, which was actually heavier and more difficult to

machine than the CDR fixture, and was rejected by manufacturing, at Mellor House, and

by Man Tech.

Figure 116: - Post CDR weight reduction joint configuration proposed by Mr Peter

Hamilton and rejected by manufacturing.

The only real changes requested to aid assembly were those covered above in section 4.7,

namely the: -

1. Grow out of the top and bottom of the common wing / empennage interface joint

surfaces between the fwd and rear spars to meet the wing skin glue line IML and

modification to interface with the all glass 4 ply spars shown in figure 117:

2. The realignment of the wing / fuselage interface joint pins as shown in figure 118.


128

Figure 117: - Post sign – off Common Wing / Empennage boom and Outboard Wing

interface joint manufacturing maturation.

Figure 118: - Post sign – off Wing / Fuselage attachment pin realignment

manufacturing maturation.

Top and bottom local surface grow out to meet skin IML.

Top and bottom local surface grow out to meet skin IML.


129

Wing joint maturation in the form of detail OML change was handed over to BAE

SYSTEMS New Business, Samlesbury Design Team, headed by Mr Peter Hamilton, as

was the responsibility for any minor wing modifications, on the 14th

November 2005, and

this concluded joint maturation by the author.

6.0 Wing flight control surfaces and systems integration.

Aileron interface attachment, aileron design and integration of the aileron actuation

system and associated wiring will be covered in this section.

6.1: - (A) Status aileron attachment interface studies.

Initial studies by the author for PDR proposed that the wing fight control surfaces cold be

attached to the rear spar and supporting ribs by two hinges per aileron surface, and two

per flap. The main points to consider when mounting there surfaces are: -

1) The bending distortion of the control surface relative to the fixed surface of the wing

must be limited to prevent aileron reversal and the risk of the knuckle of the control

surface fouling the gap seals.

2) The control hinge loads and the resulting shear forces and bending moments should

be equalised as far as is possible.

3) Structural failure of a single hinge should be tolerated unless each hinge is of fail safe

design and can tolerate cracking in one load path.

Where only two hinge points are used as in the case of the Pre and Post PDR concepts

shown in section four and in figure 119, span -wise distortion or misalignment can be

overcome by designing one of the hinges so that it can rotate about a vertical axis as

shown in figure 120. This is called the floating hinge concept and is used to best effect on

two hinge systems. Considering the control surfaces to be uniform beams the

minimum deflection is experienced when the hinges support the beam at 55% of the

control surface span i.e. each hinge at 27.5% outboard of the control surface span-

wise centre as shown in figure 120.

The flight control surface attachment using the lug and clevis system which the author

devised in the Post PDR (A) Status design study in which equally spaced lugs on the

aileron were pinned into clevises mechanically fastened to the box spar using pinned

plate fitted from the inside of the box spar so that the pins pass through the box spar web,

was carried through to the Interim Maturation study prior to CDR.

The basic arrangement is shown in figure 119 had limitations and was complex to

assemble so a more workable solution was required for CDR. The intake 5 proposal of

installing piano hinge interface attachments would not work on the revised design due to

the wing twist of 20 additionally such hinges are prone to jamming under flight

conditions.


130

Figure 119: - Post PDR (A) Status design study Wing / Aileron interface joint

(aileron design is covered in section 6.4 of this thesis).

Figure 120: - General arrangement of a two hinge plane flap illustrating the floating

hinge concept (Reference 17: - Howe. Denis, PhD: Aircraft Loading and Structural

Layout: Professional Engineering Publishing Ltd: 2004)

Knuckle

Flap CL


131

6.2: - (B) Status aileron attachment interface studies for CDR.

As stated in the previous section the (A) Status wing / aileron interface attachment was

considered too complex for an aircraft of the Terrasoar size, but could be made to work

on a structural wing as conceived for the Post PDR (A) Status design study wing,

however this arrangement would have been difficult to implement on the Interim

Maturation design study wing with the integral box spar. However the Pre – CDR

revision outboard design wing offered new possibilities for a simpler wing / aileron

interface attachment therefore after consultation with Mr Robert Currie FCS lead and

keen large radio controlled aircraft model builder the philosophy of mounting the ailerons

on a single continuous spigot was proposed which considerably reduced complexity of

the interface and simplified the design of the aileron, as shown in figure 106 above and

described in section 6.4 below.

Figure 121: - (B) Status design study Wing / Aileron interface presented at CDR.

The authors resulting design is shown above as figure 121 and consisted of a threaded

insert adhesively bonded into a hole in the trailing edge foam, a 10mm diameter stainless

steel spigot threaded at one end, two plastic sleeves which one was to be bonded into

each aileron, and spigot receiving cup which was machined into the tip rib. The design as

shown in figure 121 was presented at CDR, were it was proposed by Mr Phil Stocking

that the spigot would require some additional support at mid – span to reduce the risk of

bending under flight control actuation loads.

Threaded insert potted into trailing edge foam.

Aileron spigot with plastic sleeve

Aileron mounted on spigot

Machined spigot cup in tip rib


132

6.3: - (C) Status aileron attachment interface studies manufacture.

Further to the Critical Design Review the author designed a spigot support lug which

passed through the locally reinforced web of the rear spar to hook on to a backing plate

this mechanical fastening was reinforced with adhesive securing the lug to the plate and

the plate to the rear spar.

Figure 122: - Post CDR Aileron interface maturation with support lug integration.

Figure 123: - Support lug attachment with rear spar removed for clarity.

Threaded insert potted into trailing edge foam.

Spigot support lug

Spigot support lug

Spigot lug attachment plate

Rear spar

Tip rib spigot insertion hole

Tip rib spigot insertion hole

Spigot


133

The only other modification required post CDR was a request that the spigot be inserted

through the tip spar so that the tip spar could be used as the datum, and to that an

insertion hole was designed in through the bottom of the spigot tip rib location cup.

6.4: - (A) Status aileron design studies.

The initial pre –PDR control surface concept design studies were conducted by Mr David

Baird and Mr John Baggaley with Mr Baird being responsible for aileron design studies,

and these are included here for reference. After the departure of Mr Craig Carr from the

team Mr Baird took up his responsibilities and the design of the ailerons passed to the

author, and at the same time Mr Baggaley passed his elevator and rudder design

responsibilities to Mr James Pennington.

Figure 124: - David Baird‟s Pre – PDR (A) Status aileron design, shows a multi spar

design.

The first design concept from David Baird shown in figure 124 was a multi spar design

which although possibly possessing high rigidity would be complex to manufacture at the

Terrasoar scale. The second design concept from David Baird was both multi spar and

multi ribbed honey comb core aileron which appeared more complex than the first

concept.


134

Figure 125: - David Baird‟s second Pre – PDR (A) Status aileron design, shows a

multi spar and multi rib design.

Figure 126: - David Baird‟s third PDR (A) Status aileron design, replaced the ribs

with foam.


135

Mr David Baird‟s third and final concept removed the ribblets which were replaced with

higher density foam, but still had machined metallic closing ribs which appeared to the

author to be still too heavy and complex for the Terrasoar aircraft in view of its small

size, although this final design was considered suitable to be presented by Mr Baird at the

PDR review.

6.5: - (B) Status aileron design studies.

The Post PDR (A) Status wing design study gave an opportunity for the author to start

afresh with the aileron design as this responsibility had been passed over to him as stated

above. The first consideration was to maintain the aileron‟s stiffness whilst reducing its

weight and complexity. Form the first pre – PDR wing studies detailed in section four,

the intention was to mount two aileron pairs on each outboard wing section, therefore a

heavy complex solution would increase manufacturing time and could adversely impact

the aircrafts overall weight. The author therefore proposed the foam cored single spar

aileron, with no nose cell, wrapped in four ply fibreglass cloth at 00 / 90

0 / 0

0 / 90

0

orientations outward from the foam core which would act as the tool face. The spar

would be a „C‟ section of the same two ply fibreglass cloth / foam core / two ply

fibreglass cloth sandwich structure as the wing spars with the aluminium attachment lugs

and aileron drive horn recessed and adhesively bonded into the aileron core foam, and

this aileron is shown below in figure 127.

Figure 127: - The authors Post PDR (B) Status aileron design based on dimensions

supplied by Mr James Pennington and Miss Rachel Cunliffe (control sizing).

4 Ply glass cloth wrapped aileron skin

4 Ply glass cloth aileron closure skin is not shown for clarity

‘C’ Section spar imbedded in the foam core

Foam core

Recessed lugs

260.5mm

30.8mm

227.8mm


136

6.6: - (C) Status aileron design study for CDR.

The changes to the outboard wing structural layout made to accommodate manufacturing

detailed above in section four lead to a reassessment of the aileron to wing attachment

philosophy as detailed in section five, and resulted in the single spigot method being

chosen by the author. In parallel this lead to a further aileron redesign which simplified

manufacture still further. This aileron consisted of a foam core airfoil shape with „D‟-

nose which was wrapped in four plies of fibreglass cloth, with a horizontal spigot hole

bored through housing a plastic sleeve to protect the core from abrasion from the spigot.

Initially the aileron drive horn was aluminium and recessed and bonded in place as was

the case with the previous (B) status aileron detailed in the above section, and this

configuration was presented at the CDR review. Post CDR studies demonstrated the

availability of composite actuation horns as COTS items shown in section 6.7 and these

could be more easily integrated into the foam core and cover skin. This matured (C)

status design is shown below as figure 128.

Figure 128: - The authors Post CDR (C) Status aileron design based on dimensions

supplied by Mr James Pennington and Miss Rachel Cunliffe (control sizing).

This final design has been submitted to Man Tech for OML foam machining and

manufacturing maturation. All further aileron development is the responsibility of BAE

SYSTEMS New Business at Samlesbury.

4 Ply glass cloth wrapped skin

Foam core

4 Ply glass cloth aileron closure skin is not shown for clarity

Composite horn

Plastic sheath liner

Spigot hole

129.6mm

253.7mm

33.9mm


137

6.7: - (C) Status integration of the COTS FCS systems.

The Terrasoar FCS required integration of the actuators and wiring through out the wing,

also the centre section of the wing had to carry the empennage wiring to the booms, the

FCS schematic is shown below in figure 129.

Figure 129: - Schematic of the Terrasoar Flight Control System.

Due to the location of the fuel tank directly beneath the wing centre section the above

schematic dose not truly reflect the Terrasoar wiring layout for the wing, which was

devised by the author, and is shown diagrammatically below overlaid on the wing model

in figures 130, and 131 below. The wiring for the FCS was 5mm COTS wiring and the

electrical connections over the transportation joints were made by COTS Binder

Steceverbinder interface connectors of the 710 series, which have an external diameter of

12mm, and an overall length of 36mm, these connectors would fit within the wing

interface joint. The layout submitted to manufacturing and signed – off by the customer

was as follows: - (1) The aileron actuators wires would run from the FCS bay located

forward of the forward fuel tank frame and enter the wing leading edge passing down the

wiring tunnel in the „D‟-nose to the interface connector for the outboard actuator, and

through the leading edge spar to a second connector in the centre of the interface joint for

the inboard actuator, the wires would then run down front and rear channels cut into the

outboard wing foam as shown in figure 132, entering the actuator bay from the forward

and rear spars respectively. (2) The empennage wire would follow the same route as the

outboard aileron wire exiting the wing on the inboard side of the interface joint to a

connector in the boom, covered in reference 7.


138

Figure 130: - FCS Actuator wiring and connections at the Terrasoar wing interface.

Figure 131: - FCS actuator wires in the Terrasoar outboard wing.

Inboard Actuator wiring route

Outboard Actuator

wiring route Empennage Actuator

wiring route

Inboard Actuator wiring route

Outboard Actuator wiring route Foam core

Forward spar

Rear spar


139

Figure 132: - Actuator wiring channels in the Terrasoar outboard wing.

The wiring route for the navigation lights has still to be defined at the time of writing but

the author‟s current proposal is to run the wing tip lights down the same route as the

inboard actuator exiting at the wing tip rib.

Figure 133: - Aileron actuator installation.

Inboard Actuator wiring channel

Outboard Actuator wiring channel

Forward spar

landing

Rear spar landing


140

The aileron actuators are mounted on a single aluminium tray which is machined to

minimum thickness as shown in figure 133 the actuators themselves are screwed into

spacer blocks which are then push filled into securing grow outs from the mounting plate

this gives the actuators an interference fit and secures them in place though out the flight

envelope. These actuators drive the ailerons through adjustable push pull rods as shown

in figure 134 below these are COTS items and are shown with aileron drive attachment

horns in figure 135 below.

Figure 134: - Aileron actuator drive mechanism

.

Figure 135: - Aileron horns and actuator drive rods

Adjustable push – pull rod

Aileron horn

Actuator horn


141

These proposals and schemes concluded the author‟s work on Terrasoar FCS integration

and all future integration studies have been passed over to BAE SYSTEMS New

Business at Samlesbury.

7.0 Materials and manufacturing.

7.1: - Materials selection and aircraft weight.

As stated above in section four at the time of writing the wing skin material and the wing

spar materials are being re-evaluated to save weight and reduce materials costs, this has

been initiated by BAE SYSTEMS New Business Design Lead Mr Peter Hamilton, who

incorrectly assumed the foam core spars to be 17 ply solid glass cloth. Current studies

reflecting the changes in component sizing have been conducted by the author and tow

proposals from Mr Robert Sneddon our stress engineer are shown below in figures 136

though 139, but this is now the province of New Business. The weight of the aircraft has

grown through out design maturation which is common with all aircraft designs in the

authors experience, although the original empirical wing weights initially quoted as

targets were unrealistic and this issue is covered in reference 1.

Figure 136: - First outboard wing skin wing re-stressing with new material, figure

supplied by Robert Sneddon Hide Group Ltd.


142

Figure 137: - First inboard wing skin wing re-stressing with new material figure


Figure 138: - Second outboard wing skin wing re-stressing with new material figure



143

Figure 139: - Second inboard wing skin wing re-stressing with new material figure


A comparison can be made with the sign - off stress data presented in table 6 on page 109

above, these material changes did not affect the OML of the wing but did impact on the

IML of the skins and all of the internal structure which had to be continually re-modelled

to the new IML surfaces, this has to date been done five times by the author in the last

three weeks, as the materials selection now appears to be in a continuous state of change,

in addition to the skin material changing the foam core material has also changes from

Styrofoam to Tricast 6 a denser foam with higher temperature resistance. Further

responsibility for changing part sizing has been transferred to BAE SYSTEMS New

Business as it cannot be sustained by the author who has been under constant pressure

from line managers within the F-35 (JSF) IPT to leave the course. The weight tables for

the CDR and Structural sign – off wing designs are given in appendices A-1, the initial

target weight of 2.85kg which was always unrealistic can be compared with these figures,

and a weight target of 6kg would have been a more realistic starting point for this aircraft,

in the view of the author.

7.2: - Manufacturing methods and tool design.

As stated above the Terrasoar aircraft is in an advanced state of tool design and the

current wing centre section assembly jig is shown below in figure 140. The wing skins

are to be produced as top and bottom sections in female mould tools to provide the best

surface finish, and the internal foam cores are to be machined from Tricast 6

Polyurethane foam rather than the original Styrofoam used in the sign – off design.


144

The wing spars will be formed around simple block male tools and will not now be of the

sandwich construction but of solid glass or carbon and a comparison between the sign –

off spars and the current all glass spar is shown in figure 141.

Figure 140: - The Terrasoar wing centre section assembly jig model supplied by Bob

Cross Man Tech BAE Systems Samlesbury.

Figure 141: - changes made to the sign – off spar design to aid manufacturing.

Sign-off front spar New rear spar


145

All further changes are now the responsibility of BAE SYSTEMS New Business Design

Lead Mr Peter Hamilton.

7.3: - Structural test and qualification proposal.

Airframe structural design requires a continuing assessment of structural function to

determine whether or not the requirements have been satisfied.

The expected service performance must be satisfied before the structure enters the service

environment. This assessment is the structural testing which will ensure and substantiate

structural integrity per certification for either civil or military requirements. The basic

“building block” approach, shown in figure 142, for testing of anisotropic laminate

structures should be established at the early stages of development because the validation

process for composite structures is very dependant on testing of all levels of the

fabrication process.

Composite structural testing is similar to most metallic structural testing (the majority of

metallic testing procedures are applicable to composite structures) in that it requires

knowledge of design and analysis. The difference is that composites behave

anisotropically and need thorough experimental testing, not only of the structure as a

whole, but also of test specimens at the coupon, element, and component levels.

Design with composite materials requires knowledge of lamination theory and

appropriate failure criteria as well as related analysis. These analyses must deal with the

new set of material properties that result from the making of the laminate. Laminate

properties test results are not useful to the engineer until the data is reduced, and

translated into design allowables, and then reported in a standard format that can be

clearly understood with no ambiguity.

Due to the constraints of this thesis it is only practical to present an overview of the

structural testing and inspection required to meet the airworthiness type certification

requirements for a composite wing box structure, there is neither space nor time to detail

each and every test required or inspection procedure, but this can be found in the

reference material used to compile text.

At the outset this is a proposal and not a definitive statement of work, and is also

dependent on the type of material ultimately selected for the Terrasoar wing manufacture,

which may ultimately be a better known CFC composite with which BAE Systems has

enough experience to make coupon testing unnecessary.


146

Figure 142:- The Building Block Testing Approach which should be used for the

composite wing box validation programme.

The purpose of a structural test program is to establish failure modes, demonstrate

compliance with criteria, and correlate test results with theoretical predictions and thus

assure confidence in the part or overall airframe structure that it will perform

satisfactorily throughout its service life. The correct approach would therefore be to adopt

the Building Block Approach as illustrated above in figures 142.

Coupon Testing:-

Coupon testing is fundamentally important in that a structures constituent components

and materials are studied under an encompassing range of service conditions before a

program is locked into a production design. For example, expensive redesigns may be

avoided by an early screening of matrix materials to assess moisture degradation effects.

A broad range of material and component characterisation tests should be completed to

establish lamina material properties and establish lamina design allowables (design

criterion varies for particular applications.

A large number of tests are required to satisfy these requirements. It is vital that emphasis

is placed on accurate material property characterisation, as modern computer design

techniques e.g. FiberSim TM and FEA used in analysis of composite anisotropic materials

are extremely dependant on and sensitive to the quality of the material property data

parameters which are furnished from coupon testing results, directed to establish lamina

material properties and establish lamina design allowables (design criterion varies for

particular applications).

COUPONS

1 Modulus:

2 Static strength:

3 Fatigue:

4 Damage

sensitivity:

5 Moisture &

temperature

effects

ELEMENTS

1 Joints:

2 Crippling:

3 Shear webs:

4 Structural

details

COMPONENTS FULL SCALE

1 Stiffened

Panels:

2 Major

Joints:

3 Full Scale

sections:

4 Structural

details

1 Static:

2 Fatigue:

3 Damage

Tolerance

The Building block test sequence.


147

Single ply (lamina: tape or fabric) properties are obtained experimentally from multi-ply

unidirectional laminate specimens where all plies have the same orientation. For tape

laminates with all fibres aligned in the same direction (also tests on cross – plied

laminates can be considered to determine unidirectional properties), the ply properties

needed for design are: -

Ultimate strength values:

Elastic constraints

Poisson‟s ratio values.

Test coupons that are designed to be weighed during the conditioning process should be

weighed immediately after fabrication. All of the coupons are then stored in a dry

desiccated chamber prior to conditioning. It is vitally important that the fibre volume and

void content of each coupon is known. Moisture is absorbed by the matrix, so the

percentage of matrix in a given coupon will affect the amount of moisture absorbed. The

size and concentration of voids present in the coupon must also be known. The relative

humidity in the conditioning environmental chamber will determine the maximum

moisture content of the conditioned test coupons in this conditioning. Table 7 illustrates

the effect of Fibre Volume Percentage (FVP) on the mechanical properties of laminate

test coupons.

There are several basic coupon tests which would form the basis of a building block test

program aimed at validating a composite wing box structure, and these would deliver an

adequate design database, for establishing the design properties of the material system

and identify the most critical environmental exposures including humidity and

temperature. These tests are outlined below: -

1. Tensile tests: - The strain measurements of composite coupons can be sensitive to

the coupon configurations (configuration dose not effect such testing in metals).

Therefore unidirectional test coupons for modulus tests only, require special load

introduction tabs, but multidirectional test coupons do not. (Test coupon

configurations are detailed in reference 18 Composite Materials for Aircraft

Structures 2nd

Edition: Baker. A: AIAA Education Series: 2004).

2. Compression tests: - The measured strain for compression coupons is sensitive to

coupon configuration, and the fixture used for loading. The coupon must be

constrained from buckling, and the recommended test fixtures detailed in

reference 1 are based on those developed for unidirectional coupon testing.


148

Table 7: - Effect of FVP on mechanical properties of test laminates.

Properties Effect of FVP on laminate mechanical properties*.

[0]nt [90]nt [+/-45]ns [(+/-45)5/016/904]c

Ultimate

Strength.

Varies directly

with FVP

Not sensitive

to FVP

Varies directly

with FVP

Varies directly with

FVP

Ultimate

strain

Not sensitive

to FVP

Not sensitive

to FVP

Not sensitive

to FVP

Not sensitive to

FVP

Prop. Limit

Stress

Varies directly

with FVP

Not sensitive

to FVP

Varies directly

with FVP


FVP

Prop. Limit

Strain

Varies directly

with FVP

Not sensitive

to FVP

Varies directly

with FVP


FVP

Poisson‟s ratio Not sensitive

to FVP

Not sensitive

to FVP

Varies directly

with FVP

Not sensitive to

FVP

Modulus of

elasticity

Varies directly

with FVP

Varies directly

with FVP

Varies directly

with FVP


FVP

*The above deductions are valid for both tensile and compressive properties.

3. Shear tests: - There are numerous shear test methods which are covered in detail

in reference 16, some of the most important ASTM (American Society for Testing

and Materials) methods are listed below in table 8

Table 8: - ASTM Shear coupon composite shear tests.

ASTM Type Description

D2344 Interlaminar Short beam shear (3-point)

D3846 Interlaminar Short beam shear (4-point)

D3518 Inplane +/- 45o tensile test

D4255 Inplane Rail shear

4. Flexural tests: - These are not considered an intrinsic property, but the test is

inexpensive to run and is considered a good quality control test Reference 16: -

Anon ,”Standard Test Methods for Flexural Properties of Unreinforced and

Reinforced Plastics and Electrical Insulating Materials”, D790, Annual Book of

ASTM Standards, ASTM, is the most quoted source for test methods.


149

5. Short Beam tests: - This test is useful for evaluating the interlaminar shear

behaviour of the laminate matrix. This test is generally accepted as a method for

obtaining a qualitative measure of the laminate matrix condition rather than a

procedure for generating valid design data. Processing variables which affect the

test results are: -

Elevated temperature

Matrix moisture content

General matrix condition

6. Moisture and temperature tests: - At ply – level or coupon level there is a need to

test due to the physical compliance of the laminate which is usually accounted for

by ply – level properties. These moisture and temperature effects are accounted

for in stress analysis of the structure by using a reduced allowable.

7. Notch testing: - The notch tension and compression tests are conducted to

determine the most damaging combinations of stress raiser, temperature, and

moisture. The effect of the hole diameter size (fasteners) on residual tensile and

compression strength are evaluated based on given test requirements. Tests are

conducted for several for several hole sizes larger than baseline hole e.g. 0.25inch

(6.35mm) is the most common baseline with a w/d ratio of 6. The tests are

conducted for the most critical environmental condition determined from the

tension and compression tests. The ratio of hole diameter to coupon width will be

the same for all coupons. A reduction factor as a function of increasing hole

diameter is determined from these tests.

8. Impact testing: - Impact testing of coupons in which the coupon is subjected to a

single impact and then compression tested does not give an accurate indication of

suitable design properties and is merely used to compare material characteristics.

The impact testing to determine design values should be done on components and

/ or full scale test structures.

9. Fastener bearing and pull – trough tests: - (A) Fastener bearing strength for tape

composites is a function of the lay - up

The 100% 0o ply laminate would fail by shear tear out, and strength would be

essentially a function of shear strength of the matrix and the cross – sectional area

to the edge of the coupon:

The 100% 90o ply laminate would fail by net section tension and strength would

be a function of the matrix tensile strength and the net cross – sectional area:

Both of these are comparatively weak failure modes, multi – directional

reinforcement is required if appreciable bearing strength is to be obtained.

The bearing strength allowable can be determined from the following equation: -


150

Fbru = Fbr x Ke x Kcsk (equation 1.1)

Where: - Fbr = “B” allowable for room temperature dry (RTD) bearing strength of

non – countersunk holes:

Ke = Environmental correction factor:

Kcsk = Countersunk correction factor.

The “B” bearing allowable for a non – countersunk hole RTD condition is

modified by Ke and Kcsk to account for environment, t/d, w/d, single lap shear

etc. and countersunk thickness for flush fasteners. To determine the Fbr allowable,

tests would be conducted on double lap shear coupons, made from several

different laminate thicknesses. These test results would then be pooled to establish

a “B” bearing allowable. The test for extreme environmental conditions, are used

to provide the Ke environmental correction factors. Each coupon used to test

environmental conditions would be cut from the same location in the laminates.

Then the ratio of the environmental bearing strength to the RTD bearing strength

can be used for statistical analysis. The bearing tests for fabric laminates would be

similar to the tests for tape laminates.

(B) Pull – through tests: - The pull – through test, is conducted to determine the

load required to pull fasteners through composite laminates. This property is

important for structures subjected to internal pressure loads such as integral wing

fuel tanks, or to permit buckling of skins (or webs) without failure at fastener

attachments. The majority of failures of secondary supports, (e.g. secondary tension

due to diagonal tension shear buckling effect) are in the form of fasteners pulling

through laminates, particularly countersunk fastener heads. Composites are

generally weak in pull – through strength. Pull – through strength is a function of: -

Laminate thickness:

Fastener diameter:

Configuration of fastener head type:

Laminate deflection.

Therefore, tests need to be conducted for several thicknesses, fastener diameters,

and fasteners head types. Laminate support should supply rigidity of the same order

as that expected for the structure.

10. Process Control Testing: - The above testing is primarily qualification and

acceptance testing, were as process control testing is the next step in the test

program, (see figure 143 and is derived to ensure that the fabrication process is

working as it is supposed to work. The concept is derived from the early days of

metal bonded structures when adhesive qualification coupons were run along with

bonded assemblies. As they were processed together, if the coupon tests were

successful this was taken as evidence that the process was being adequately

performed. This concept was extended to composites and travellers (tag end

coupons) are manufactured with composite components with the following

considerations: -


151

a. The test specimen must reflect the process (cure cycle) that is occurring at critical

locations (often more than one on large complex parts). This means the significant

factors in the process must be reflected in the fabrication of the test coupon.

b. The coupon should be made along with the part (same material, same exposure to

contamination in addition to the cure cycle).

c. The coupon should be a standardised type so that acceptance criteria can be

established with statistical relevance.

d. The method should be simple enough to be systemised to the point that only

manufacturing quality control interface is involved. Airframe engineering should

only become involved during set – up and when deviations to the requirements

occur.

Tests must satisfy items (a) and (b) above and should be conducted on pieces trimmed

from production parts. However, this approach is contrary to items (c) and (d)

requirements and therefore requires a trade – off study for cost effectively.

Figure 143: - Materials Qualification Tests V‟s, Batch Acceptance Tag – End Tests,

for the coupon test program.

Candidate material

Submitted.

Materials lab.

Qualification test

performed.

Airframe Engineering.

Test results reviewed.

(a) Material qualification tests.

(b) Batch acceptance and tag end test.

Material lot

received.

Materials lab.

Batch acceptance

tests performed.

Production.

Parts and tag ends

layed up.

Materials lab.

Tag end tests

performed.

Production.

Clearance to continue

fabrication.


152

Some elevated temperature moisture coupon testing data would be used in support of

element testing to meet certification requirements.

Element and component tests: -

Element and component testing represents the second and third blocks in the building

block composite certification test program shown in figure 142, although there is a

growing opinion that these tests can be reduced and substituted by computer analysis

tools based on FEA and structural optimisation software. However currently and for the

foreseeable future representative structural element testing will play a key role in

composite airframe certification programs. Figure 144 shows typical structural elements

and components used for allowables verification, and fulfilment of both static and fatigue

/ damage tolerance structural integrity requirements. Such elements contain detail

features such as holes, notches, stringer run – outs, joggles, etc, and the objective of

element and component testing is to determine what effect these features have on the total

structure, for example:-

An access hole through a skin structure may drastically alter the stress

concentration and redistribution in the surrounding area:

A fastened bonded and / or fastened joint may also produce significant stress

perturbations in the joints immediate vicinity:

These sections of components may induce large stress perturbations in the

constitutive material and induce failure modes very different from those predicted

by laminate theory:

In addition to inplane axial and shear loads, concentrated normal tension load on a

composite integrally stiffened panel, can be used to determine the flatwise tension

and peel strength between the skin and stiffener which are much lower than

inplane laminate strengths, hence stiffener pull – off strength tests would be

conducted as part of the wing structure qualification program.

Element and component testing will require much more instrumentation and have more

complicated load introduction and test fixtures than coupon testing therefore this form of

testing is more expensive, but yields a much more accurate picture of structural

behaviour.

This element and component testing would be used to cover the element and component

level testing for both the Proof of Structure: - Static and Fatigue / Damage Tolerance

requirements.


153

Figure 144:- Component Tests (DC-10 Vertical Composite Fin Box) (reference 16).

Joint Design evaluation: - One of the most difficult aspects of joint testing is inducing

the loads into the joint in a fashion which is representative of the boundary conditions of

a test article. For example, it may be difficult or virtually impossible to determine, much

less duplicate in a test, the stiffness boundary conditions which are present at the joint in

actual service. The choice of boundary conditions which are readily reproducible in most

tests consist of either free or fixed supports, which usually have a very high reserve factor

on them for BAE Systems STF in the order of 4. Based on previous testing on legacy

aircraft information may be available as to the procedure and gripping hardware which

would be most appropriate for approximating in situ conditions, such as historical tests on

Airbus A320‟s empennage which could be applied to the A400M wing testing. The

service stress distribution in the components which border the joint would then have to be

predicted by analytical methods probably FEA modelling. Then it is possible to

approximate the same stress proportions by using boundary control techniques which are

related to an active feedback signal from the component under test. Such a test would be

expensive, but the application may be critical enough to warrant resorting to such a

technique, for example: - the adhesively bonded spar to bottom wing skin joints. These

tests would be conducted in the hot / wet condition as described for coupon testing in

section above.


154

Cut – outs: - Obviously, small coupon test specimens are inappropriate for evaluating the

effects of cut – outs, or large flaws unless the imperfection being assessed is small

compared to the coupon width; coupons can also be adversely affected by free edge stress

effects. Therefore panel tests with major and minor dimensions close to that of the actual

structure would be used for notch, cut – out or imperfection tests.

In composite structures, a large cut – outs such as access holes, or fuel transfer holes, will

present significantly different stress redistribution around the edge of the cut – out and

therefore an array of strain gauges would be used to quantifying the strain distribution.

But the tips of cracks cause steeper strain gradients which would be measured by photo –

elastic coatings or Moiré fringe analysis.

Free Edge Effects: - The delamination problem which is associated with free edges in

cross – ply laminates will be more severe in laminates with cut – outs because large stress

concentrations exist in the vicinity of cut - outs. Therefore measurements of through – the

– thickness deformation should be made at the cut - out edge since this may be the most

relevant measurement to support analytical characterisation studies. Also strain gauges,

displacement sensors, and optical methods could be used for delamination strain

characterisation.

Damage Tolerance testing: - Damage tolerance testing is significantly different for

composites than for metal. Damage tolerance in metals is related to the rate of

propagation of a crack of a given size and location, where as damage tolerance in

composites is primarily dependent on resistance to impact. Composite material structures

must be designed to support design loads after an impact that has a reasonable probability

of occurring during fabrication or during the service life of the structure. To define a

strain allowable to account for impact damage compression stress is similar to defining a

fatigue allowable for metal structure (tensile stress is critical). The fatigue allowables are

selected based on limited tests and previous design experience. However, final fatigue

substitution is based on durability (fatigue) tests conducted on full – scale components or

the complete airframe. Compression tests are conducted on impact damaged coupons to

select preliminary compression design stress allowables, and then compression tests of

impact damaged structural panels and subcomponents are conducted to substantiate the

design allowable.

To define design allowables for impact damage, tests would be conducted on flat

laminates loaded in compression. These may have varying amounts of impact damage,

dependent on panel thickness and damage tolerance requirements for damage visibility

and maximum impact energy. The panels must be large enough to nullify size effects,

e.g., 25.4cm x 30.48cm. The results being representative of impact damage to areas of the

structure between reinforcements (e.g. stiffeners). The effect of impact damage where

reinforcements are attached to the skin or the effect on the reinforcements themselves

would be determined by tests on reinforced structurally representative panels.


155

Because strength and damage sustained can vary as a function of lay - up configuration,

several variations of each laminate would be tested. The effects of environmental

degradation would also be evaluated with tests at given moisture content and

temperature, with pre – conditioned structural panels, tested in environmentally

controlled test chambers. Some tests would also be conducted with higher impact

energies to determine the trend of data for wider damage widths. It would also be

necessary to conduct sufficient cyclic tests to ensure that no detrimental damage growth

will occur during the structures expected service life.

The damage requirements vary considerably, depending on mission and life – time

requirements. For example the requirements for a typical military composite structure (as

this aircraft may be used as a military transport) are as follow: -

(a) Low level impact damage: -

An impact of 8.4J from an impactor with a 12.7mm diameter hemispherical head:

The damage laminate should have the capability of carrying static ultimate loads.

(b) High level impact damage: -

An impact of 140J from an impactor with a 25.4mm diameter hemispherical head:

or an impactor which would not cause a dent deeper than 2.54mm:

The damaged laminate should have the capability of carrying static limit load.

Durability (Fatigue) testing: - Durability testing in composites must consider the effects

of environmental exposure on static and dynamic behaviour. Therefore the durability

testing of the composite wing components becomes a function of load cycling and

environmental exposure. Airframe durability testing would be would be accomplished

using a flight by flight real – time loading spectrum based on the aircrafts life – time and,

concurrently, environmental exposure based on flight temperatures and ground based

moisture environments. In addition, accelerated flight spectrum loading and accelerated

moisture / temperature environments could be used to simulate real – time testing but

care would need to be taken in correlation of these accelerated tests with real – time

loading and environmental conditioning.

Fibre – dominated laminates are considerably more efficient in load – carrying ability

than are matrix – dominated laminates: however the latter are sometimes needed, for

multidirectional loadings and damage tolerance requirements. It is generally assumed that

matrix – dominated laminate design is governed by durability strength, where as fibre –

dominated laminate design is governed by static strength. Therefore, durability testing for

structural integrity verification of matrix – dominated laminates such as those for the

bottom wing skins would have to include bonded joints to the spars.


156

Element and component tests would address requirements for Proof of Structure by static:

fatigue: damage tolerant: and fail safe evaluations.

Full – Scale testing (FST): -

Full – scale testing (FST) of the complete airframe, or the testing of a major structural

component, in this case the wing, is the major test in an airframe structural test program,

and is the fourth and final building block in figure 142. FST is one of the primary

methods of demonstrating that the airframe or major structural component can meet the

structural performance requirements and is extremely important because it tests all of the

related structures in the most realistic manner.

Typical FST include: - static: durability (fatigue): and damage tolerance. The use of FST

must take into account the unique characteristics of composite structures and their

response to the expected service conditions as simulated by the test.

FST is necessary check in the process of developing satisfactory structural systems,

although analytical techniques have significantly improved in recent years with more

capable computer analysis techniques and the wide – spread use of finite element

analysis, the complexity of composite structural systems still requires FST verification

programs.

Test requirements such as limit and ultimate loads are often established on the basis of

material test scatter derived from coupon testing, and composites usually exhibit higher

scatter than metallics therefore this raises difficulties in establishing values for the test.

Also, composite laminates exhibit relative brittleness, low interlaminar strength and

differences in coefficient of thermal expansion (CTE) in contact with metal parts, and all

of these factors would present serious problems for the FST program.

There are three considerations which would need to be addressed when choosing the size

of the FST article for the wing test but are equally applicable to all FST programs: -

The test article must be large enough to allow for proper complex loading and for

the load interactions at interfaces that would otherwise would be difficult to

simulate:

If the component is small enough it is less expensive to use a FST environmental

test to certify the structure, (an example of scale is the environmentally controlled

cabin pressure and bending test conducted by myself at BAE Systems Brough

STF on the Eurofighter Typhoon single seat test article in 1992 which was a

complete forward fuselage contained within a purpose built chamber under load):

Structural configuration also has an important role in the environmental condition

test: - Primary or secondary structure: Type and complexity of loading.


157

The test article size for the wing test would be a production representative half span

article with a dummy counter balance as in the case of the Nimrod MR4 test at BAE

Systems STF. This would suffice as the wing / fuselage would be included in the test

article and the port wing design features would be mirrored in the starboard wing, if there

were any non – symmetrical details these would be tested as component test articles.

The major FST objectives are as follows: -

To verify analysis with actual internal load distribution were a test article is used

which may not representative of the final production configuration of the

structure, (as was the case with the JSF Horizontal Tail Test box for which I was

responsible for designing).

To observe any unexpected discrepancies occur:

To evaluate whether durability and damage tolerance have been adequately

assessed:

To evaluate the durability of combinations of composite and metal parts,

particularly in interface areas where glass cloth packers are required due to

galvanic corrosion and to investigate differential thermal expansion problems.

Instrumentation (all data would be electronically recorded and controlled by computer

data logging and control system) used on the FST structures would include: -

Strain gauges:

Deflection indicators:

Accelerometers:

Stress coatings:

Acoustic emission detectors:

Evener systems.

Pre – test prediction of the wing FST structural failure loads, locations and mechanisms

are important as they will profoundly influence the test loadings, rig design and load

application. These would be based on minimum margin of safety calculations and the

known statistical variation of the material allowable developed from coupon tests and

used in analysis.

Appropriate “knock – down” factors are applied to test margins after completion of the

mechanical property and environmental testing program covered above. These results

would be verified by long – term aging tests on critical structural components, which are

subjected to real – life environments and tested at various intervals throughout the

duration of the test program. This would clear the FST article of the requirement to be

environmentally conditioned because perfect duplication of moisture / temperature / and

time histories for such a large and complex structure would impossible and even

attempting it would be unacceptably costly, and component testing is considered

validation under airworthiness requirements.


158

Careful consideration of the method of inducing loads into the FST of the wing would be

required, generally: -

a. In tension testes – The mating structures must be sufficiently strong that they

must not fail before the structure under test.

b. In compression tests – The mating structure must be simulated and the loads

applied to it in such that the rotational characteristics are approximated. This

subjects components which are in buckling critical to appropriate end – fixity

conditions and ensures adequate load diffusion into the test structure.

Static FST of the wing: - The static FST a most important test in the qualification of

composite airframe structures because of their brittleness and sensitivity to stress

concentrations compared with the same structures in metal therefore the following

methodology would be applied to the test article described above:

1. The parameters considered for the static test would be:

Type of test structure

Type and number of load conditions

Usage environment to be simulated

Type and quantity of data to be obtained.

As stated above the environmental effects would be addressed at the analysis,

coupon, structural element, and component level “building block” stages (see

figure 2.2(b)). The sums of these tests would be consolidated to validate and

satisfy the consideration of environmental testing.

2. The method of loading the FST article requires careful consideration due to the

composites weak through the thickness strength (tension) and sensitivity to stress

concentrations, possible methods for the wing test are outlined below:

a) Tension – patches method (see figure 145:

Offers uniform load distribution with a closer representation of the real

structure load but is expensive:

Involves a more complex test set – up (higher cost and longer set – up

time):

Introduction of load directly into a composite bonded surface must be done

more carefully than with metal surfaces because of their inherent through –

the thickness weakness.


159

b) Loading frame method (see figure 146:

Less complex loading set – up and less costly method:

All loads are converted into numerous compressive concentrated loads

(this is not as effective as the tension – patches method but it is acceptable:

The attachment of substructures such as spars, ribs etc, at locations of

concentrated loads needs careful investigation to make sure there is

sufficient strength reserves are present in the substructure.

3. The following FST sequence would be followed in accordance with reference 1:

a) Checking of the test set – up, which would involve functional testing of:

Loading jacks and evener system:

Instrumentation:

Data recording:

Real – time data displacement (this check would be accomplished by

applying a simple load case at low levels to ensure that the loads are

induced as expected.

b) A strain and deflection survey would be run to determine whether the strain

distribution and deflections are as predicted.

c) The lowest of the loads to be certified are applied first i.e. the conditions for

which there is the highest confidence are run first and the conditions with the

highest risk of premature failure are run last.

d) The early test results could be extrapolated to the predicted design ultimate

load level for analysis validation.

e) If a risk of failure before design load is determined then the test would be

stopped and a careful review and investigation would need to be conducted.

4. Ultimate load requirements – i.e. the type of load required by the qualifying or

certification agencies to meet their validation requirements includes:

a) U.S. FAA requires the structure to the limit load (same as that governing

metallic structures):

b) U.S. Military requires testing to the ultimate load.

c) AMC No 1 to CS 25.603 requirements call for testing to ultimate load for the

article like the U.S. military requirement.


160

5. The final step is a review of data obtained the test and supporting evidence from

element and sub – component testing and evaluation of its correlation with the

analytical stress analysis.

Figure 145: - Tension - patch methodology of loading a wing test: - this multi pad

and loading tree method is being replaced by multi actuator loading on military

aircraft testing programs (reference 16).

The structure should be able to withstand static loads to be expected during completion of

a flight on which damage resulting from obvious discrete sources occurs.


161

Figure 146: - The loading frame method of loading the wing test article (reference

16).

Durability FST of the wing: - Cyclic Full Scale Testing of airframe structures used to

evaluate metal structures is also applied to composite structures. In general, FST cycle

testing is limited to 2 to 4 lifetimes of spectrum loading (2 for civil aircraft) in the

presence of BVID, including a spectrum load enhancement factor such as environmental

effects. Periodic inspections must occur during FST durability testing at specific intervals

between the limits of detection and the time when limits of residual strength capability

have been reached. These inspections are conducted to determine whether any damage is

progressing due to cyclic loading in order to: -

Obtain the durability performance of the structural details:

Detect any critical damage whose growth would result in failure of the test article

during the durability test.


162

For example stiffness changes in a composite structure has been found to be an indication

of fatigue damage, hence crack and delamination (very difficult to detect) inspections are

conducted at intervals throughout the test, after a given number of cycles which would be

based on coupon, element and sub - component level testing. The inspection plan would

use the minimum detectable damage / defect size established in the materials

qualification and manufacturing development coupon, element, and sub - component test

level of the building block test program and would determine: - the frequency, and extent

of the inspections, the methods employed, intervals, inspection for zero growth, and the

residual strength associated with assumed damage. Non – Destructive Inspection

techniques likely to be employed are ultrasonic C – scan, x-ray, acoustic detection by

microphones in the structure to listen for delaminations. Finally a post – test inspection of

the test article after the FST durability test would be conducted to ensure that no damage

had occurred that would threaten the structural integrity of the composite wing box.

Damage Tolerance FST of the wing: - Testing composite FST structures for damage

tolerance is especially important because it addresses the concerns associated with both

the static and durability FST‟s. The damage tolerance test, like the static test, is a

qualification requirement to meet the Proof of Structure requirements of AMC CS

25.603, and is also required by the U.S. FAA, and military regulatory authorities. The

load specified by civil and military requirements varies )both specify a residual strength

requirement which in this case is equal to or greater than the strength required for the

specified design loads considered as the ultimate load) and requirements also vary

depending on: -

Ability to inspect damage:

Type of service inspection used:

Type of aircraft.

As in durability tests the critical flaw or damage may be associated with either its initial

state or its growth after cyclic loading. The environmental effect during the cyclic test is

not easily defined but the load enhancement of the spectrum as recommended for the

durability test would be the best option. Because the FST damage tolerance test has many

similarities to the static, and durability tests, all the testing considerations which apply to

them are also applicable to this test.

If the residual strength test is successfully passed the structure can then be loaded to

failure to further evaluate its damage tolerance capability.

The flutter proof of structure requirement would be met by sub – component testing.

The test program outlined above would meet the damage tolerance / environmental

degradation / impact evaluation requirements for military aircraft composite wing box

certification criteria.


163

However although the above test program or something like it would be required should

BAE SYSTEMS seek to build the Terrasoar aircraft in quantity and gain unrestricted

airspace certification, for a one off aircraft flying in a sanitised range only minimal

structural proof testing would be required. This concludes the authors‟ coverage of

materials issues, manufacturing, and testing for the Terrasoar UAV.

8.0 Conclusions.

The objective of the Terrasoar project was to design, manufacture, ground test, and fly a

High Altitude Long Endurance small low cost Unmanned Air Vehicle with CAA

certification for flight at 10,000ft (3,048m) in controlled airspace, with an endurance of

5hours, and payload of 5kg, which. However these objectives were too ambitious for the

resources available to the team in terms of funding manpower and timescale.

The final solution has been to design and manufacture an aircraft which can meet a set of

more realistic missions criteria up to an altitude of 400ft (121.92m), and an endurance of

4.5hours, which was within the financial limits of the project, with the capability of being

modified with additional systems, new outboard wing, and a new engine, to fly at the

original target altitude.

The wing weight at CDR was 8.056kg and after stress maturation this grew to 11.318kg a

net gain of 3.262kg in the immediate future the material selection must be finalised so

that the final iteration of weight analysis can be undertaken based on the new detail part

sizing.

The major advances in the current wing design over that proposed by the 5th

intake are

summarised below:-

1. The fuselage to wing joint, proposals aimed at producing a robust interface joint

which could withstand the moreover, and fight loads of a much larger chord wing

than was proposed by the 5th

intake, yet facilitated the removal of the wing for

transportation to and from the test site. This involved several concepts: - The first

concept was a conventional lug and clevis design with the lugs being integral to

machined wing root ribs, these would then be slotted into „H‟ section longerons in

the fuselage and bolted into place, but although this was good in principal and is

used on the Harrier jet fighter, access for bolt insertion and securing mitigated

against this method, also in maturation the longerons were deleted. The second

concept was a hood which was to be integral to the bottom wing skin but the

complex curvature and requirement for large bend radii rendered this solution

impossible, therefore the author proposed bonding the hood to the bottom wing

skin and reinforcing this joint with bolts through the forward and aft spars. The

hood itself was to bee attached to the fuselage by quick release locks of the type

used to secure aircraft radomes. This concept was deemed to complex both to


164

manufacture and structurally analyse. The third and final solution was a greatly

improved and more efficient interface in which joint the wing / fuselage interface

joint was been considerably simplified and relies on two sleeved bolts through the

fuel tank frames, with ample access on either of the frames to insert and secure

the bolts so that the wing could be secured for fight and removed for

transportation. The major advantages over the 5th

intake design are that the whole

wing loads are distributed along the sleeve bolts and are not point loads, and the

bolts along the chord line are in shear as the wing tries to lift off from the fuselage

and not in tension which was the case for the 5th

intake concept. Their concept

which consisted of screwing in two bolts at right angles to the chord plane into the

rear spar, and using a machined channel with integral lugs adhesively bonded to

the front spar would possibly have failed in tension both at bond interface

between the spar and the channel, and then in tension at the bolts through the rear

spar, this could not happen with the final 9th

intake joint design.

2. The inboard to outboard wing joints were another major improvement which also

had several conceptual iterations initially for two independent joints, one for the

wing interfaces and another for the boom interfaces, before the final combined

wing and boom design was selected. The first concept for a wing only interface

joint was based on ribs mounted in both the outboard and centre wing sections

have double staggered lugs for top and bottom pins which are inserted through a

cut out in the leading edge and screwed home into retained nuts mounted in the

trailing edge, and the skins extend over the interface with the outboard skin

slotting under the inboard skin to form a lap interface. A panel is then fitted over

the leading edge and secured with aero tape. This concept was however on the

scale of the Terrasoar like watch making, and the integrity of the leading edge

cover panel would be in doubt, so this concept so it was dropped. The second

concept for a wing only interface joint was to form a plug and slot interface

between ribs mounted at the ends of the torsion boxes of both outboard wing and

the centre wing, a quick release latch across the bottom tension skin of the

interface would ensure there was no risk of separation in flight. Because the top

wing surface in flight is in compression and the plug is nearly full chord a latch

across to top skin would not be required. Although this solution would have

increased the wing weight by a small margin minimum machining would have

kept this low and it could have been traded off against other areas of the structure.

This concept was loosely based on some joints the author had designed for the F-

35 Horizontal Tail test box, there was also no risk of galvanic corrosion as the

composite used in the skin and spars was glass cloth and not carbon fibre, and this

philosophy was further developed into the combined wing and boom interface

joint. Initial studies for the boom interface centred on the lug rib concept similar

to that proposed initially for the wing to fuselage joint, and subsequent to this lugs

on tabs fitted through the spar flanges had been proposed, but neither of these

were satisfactory, and a way of combining the lugged rib of the boom interface

with the wing plug and slot interface was devised. The third concept for a

combined joint was achieved by putting integral boom attachment lugs onto the

underside of the plug and slot wing interface ribs. The booms would then be


165

slotted over lugs and pinned in place using a channel section end fitting securing

both ribs together, and making the need for a bottom skin latch un-necessary, and

this concept was carried through into the production aircraft as explained in the

bulk of this thesis. This was a much better system from both a manufacturing and

assembly view point and gave one common load path for wing bending and boom

loads, and once again the attachment bolts were in shear. The 5th

intake interface

joints were still immature when they finished the course, and resembled a

complex hook over attachment over the front spar for the boom to interface with

an attachment which was not defined, and as for the inboard to outboard wing

interface this consisted of a complex pinning arrangement which was equally

vague, and compared with this the authors combined wing and boom arrangement

has the clear advantages stated above.

3. The first concept for the flight control surface attachment using the lug and clevis

system which the author devised post PDR in which equally spaced lugs on the

aileron were pinned into clevises mechanically fastened to the rear spar using

pinned plate fitted from the inside of the rear spar so that the pins pass through the

spar web, was carried through to the Interim Maturation study prior to CDR. The

basic arrangement had limitations and was complex to assemble so a more

workable solution was required for CDR. The 5th

intakes proposal of installing

piano hinge interface attachments would not work on the revised design due to the

wing twist of 20 additionally such hinges are prone to jamming under flight

conditions. The second concept was developed after consultation with Mr Robert

Currie FCS lead and keen large radio controlled aircraft model builder which was

the philosophy of mounting the ailerons on a single continuous spigot which

considerably reduced complexity of the interface and simplified the design of the

aileron itself, and consisted of a threaded insert adhesively bonded into a hole in

the trailing edge foam, a 10mm diameter stainless steel spigot threaded at one

end, two plastic sleeves which one was to be bonded into each aileron, and spigot

receiving cup which was machined into the tip rib. With the addition of a support

lug at the mid span of the spigot in between the two ailerons this arrangement was

carried through to manufacture and unlike the 5th

intakes proposal was resistant to

jamming.

4. The structural layout of the wing and the locating of the ailerons and their

actuators in the outboard wing panels offered clear advantages over the 5th

intakes

wing in that the linkages could be more direct, i.e. no build up of excessive

backlash in the mechanism, the shorter the link, the less chance of control surface

flutter. Also the structural layout of two „C‟ section spars clamping over the front

and rear of the foam core gave the wing enhanced stiffness forming with the skin

a continuous large section filled box spar. By comparison the 5th

intake wing was

a lose collection of disassociated parts which did not act concertedly to impart

stiffness to the wing.


166

Finally and in fairness to the 5th

intake wing designer their wing was half the chord of

that selected for the 9th

intake Terrasoar design which placed server constraints on the

location of systems and the structural layout of the wing, and the design was still at

Preliminary Design Review stage when they finished the course and many issues

remained unresolved.

The authors‟ wing design represents a full design evolution taking onboard inputs from

other team members and technical specialists in manufacture and structures to produce a

design solution which meets the operational requirements, and can be produced cost

effectively within BAE Systems, Man tech facility.

The structural test and qualification program for the wing detailed above in the main

body of the thesis is ambitious but represents a test program or something like it that

would be required should BAE SYSTEMS seek to build the Terrasoar aircraft in quantity

and gain unrestricted airspace certification.

However for a one off aircraft flying in a sanitised range only minimal structural proof

testing would be required. This concludes the authors‟ coverage of materials issues,

manufacturing, and testing for the Terrasoar UAV.

Currently the Terrasoar program is in full manufacture with completion of all major

structure scheduled for April, with final assembly and systems installation in May for a

planned first flight in July 2006.

This has been a rewarding project and the author looks forward to its successful

completion with the flight test program.

9.0 Further work required.

1. Establish a final wing skin and spar material and re-stress to give final structure

seizing:

2. Complete aileron design of the drive horn integration:

3. Complete systems integration with respect to the navigation lights:

4. Stress the ailerons and the spigot designs.

5. Proof test the as built wing:

6. Build and fly the Terrasoar aircraft.


167

10 References

1): - UAV Certification Management, Certification / Qualification and Control of Mass /

C of G / Inertia: Mr Alan Barnes GDP thesis: Cranfield University 2006.

2): - Terrasoar Unmanned Aerial Vehicle Group Design Project Thesis: Miss Emma

Bradley GDP thesis: Cranfield University 2006.

3): - Preliminary Design Review (PDR) for “Terrasoar” High Altitude Long Endurance

Unmanned Air Vehicle: 5th

Intake PDR presentation: Cranfield University 2001.

4): - Authors private collection of Terrasoar data.

5): - AVT-0215 9th

Intake Group Design Project Frame Work document: School of

Engineering Cranfield University: March 2003.

6): - Fuselage design and integration for a High Altitude Long Endurance Aerial Vehicle

UAV: Mr Paul Francis Gilligan GDP thesis: Cranfield University 2006.

7): - Terrasoar Aerodynamic Configuration, Performance, and Design: Mr James

Pennington GDP thesis: Cranfield University 2006.

8): - Terrasoar Chief Engineer, Powerplant and Manufacturing Group Design Project: Mr

Damian Adams GDP thesis: Cranfield University 2006.

9): - Development of a Cost Effective UAV Avionics and Flight Control System: Mr

Robert Currie GDP thesis: Cranfield University 2006.

10): - Terrasoar UAV Payload, Stability and Flutter Group design Project: Miss Rachel

Cunliffe GDP thesis: Cranfield University 2006.

11): - Terrasoar Structural Analysis and Detail Stressing: Mr Robert Sneddon GDP

thesis: Cranfield University 2006.

12): - Page 16: Aircraft Conceptual Design Synthesis: Howe. D. PhD: Professional

Engineering Publishing Ltd 2000.

13): - Page 47: Attack of the Drones a History of Unmanned Aerial Combat: Yenne. B.:

Zenith Press 2004.

14): - Fundamentals of Fighter Design: Airlife Publishing Ltd 2000.

15): - Page 110: Aircraft Performance and Design: Anderson. J. D.: McGraw Hill

Publishing 1999.


168

16): - Introduction to Aeronautics (A design perspective): Drandt. S.: Stiles. R.: Bertin. J.

J.: Whitford .R. : AIAA Education Series 1997.

17): - Aircraft Loading and Structural Layout: Professional Engineering Publishing Ltd

2004.

18): - Composite Materials for Aircraft Structures 2nd

Edition: Baker A.: AIAA

Education Series 2004.


169

Appendices

A-1: - Wing weight tables.

Table 9: - CDR wing submission wing weight table based on Phenolic fibreglass and

R63.80 foam core and Styrofoam and Al 2024-T351 densities.

Component Number for

T-WINGPORT-401.

CATIA

Weight kg

Design change: Reason for

change.

T-WRAPPEDSKINPORT-001 0.916 Skin thickness 2 plies current

T-WINGPORTDNOSE-001 0.095 None: Weight of Styrofoam

T-WINGPORTBOX-001 0.243 None: Weight of Styrofoam

T-WINGPORTTRAILINGEDGE-

001 0.069 None: Weight of Styrofoam

T-PORTWINGLESPAR-001 0.285 None: Weight of Sandwich

T-PORTWINGTESPAR-001 0.271 None: Weight of Sandwich

T-PORTWINGCUTOUT-001 0.115 Thickness change from 2 plies to

6 plies: stress calculations

T-PORTWINGACTPLATE-001 0.156 Provisional open for reduction

T-PORTWINGTIPRIB-001 0.285 Provisional open for reduction

T-WINGIBATTACHRIB-001 0.367 Provisional open for reduction

T-WINGOBATTACHRIB-001 0.713 Provisional open for reduction


T-WINGSTBD-401.

CATIA

Weight kg


change.

T-WRAPPEDSKINSTBD-001 0.916 Skin thickness 2 plies current

T-WINGSTBDDNOSE-001 0.095 None: Weight of Styrofoam

T-WINGSTBDBOX-001 0.243 None: Weight of Styrofoam

T-WINGSTBDTRAILINGEDGE-


T-STBDWINGLESPAR-001 0.285 None: Weight of Sandwich

T-STBDWINGTESPAR-001 0.271 None: Weight of Sandwich

T-STBDWINGCUTOUT-001 0.115 Thickness change from 2 plies to

6 plies: stress calculations

T-STBDWINGACTPLATE-001 0.156 Provisional open for reduction

T-STBDWINGTIPRIB-001 0.285 Provisional open for reduction

T-WINGIBATTACHRIB-001 0.367 Provisional open for reduction

T-WINGOBATTACHRIB-001 0.713 Provisional open for reduction


170

Table 9 continued: -

The total wing weight after CDR for Centre box spar and foam spared wing in its

final form was = (Outboard wing weight 3.315kg x 2 = 7.030) + Centre wing weight

1.026kg = 8.056kg.

Densities Used:-

Aluminium 2024-T351 Plate = 2768kg/m3

Fibreglass = 300g/m2 = 1200kg/m

3

Styrofoam = 40kg/m3

Sandwich foam R63.80 core = 90kg/m3

These values were produced using a CATIV V5 transposed model copy of the original

CATIA V4 model, with densities in kg/m3.


T-WINGCENTRE-401.

CATIA

Weight kg


change.

T-WINGCENTSKINNOSEPORT-

001 0.053

Skin thickness 4 plies plus

foam.

T-WINGCENTSKINREARPORT-

001 0.193


foam.

T-WINGCENTSKINNOSESTBD-

001 0.053


foam.

T-WINGCENTSKINREARSTBD-

001 0.193


foam.

T-CENTWINGBOXSPAR-001 0.534 None: Weight of Sandwich


171

Table 10: - Post CDR sign - off wing redesign weight table based on revised

fibreglass and R63.80 foam core and Styrofoam and Al 2024-T351 densities


T-WINGPORT-401.

CATIA

Weight kg


change.

T-WINGOBTOPPORTSKIN-001 0.916

Skin thickness change from 2

plies to 6 plies: stress

calculations.

T-WINGOBLOWERPORTSKIN-

001 0.965



calculations / maturation

T-WINGPORTDNOSE-001 0.095 None: Weight of Styrofoam

T-WINGPORTBOX-001 0.243 None: Weight of Styrofoam

T-WINGPORTTRAILINGEDGE-


T-PORTWINGLESPAR-001 0.285 None: Weight of Sandwich

T-PORTWINGTESPAR-001 0.271 None: Weight of Sandwich

T-PORTWINGCOVEPLT-001 0.115 Thickness change from 2 plies

to 6 plies: stress calculations

T-PORTWINGACTPLATE-001 0.156 Provisional open for reduction

T-PORTWINGTIPRIB-001 0.291 Maturation

T-WINGOBATTACHRIB-001 0.552 Maturation


T-WINGSTBD-401.

CATIA

Weight kg


change.

T-WINGOBTOPSTBDSKIN-001 0.916



calculations.

T-WINGOBLOWERSTBDSKIN-

001 0.965



calculations / maturation

T-WINGSTBDDNOSE-001 0.095 None: Weight of Styrofoam

T-WINGSTBDBOX-001 0.243 None: Weight of Styrofoam

T-WINGSTBDTRAILINGEDGE-


T-STBDWINGLESPAR-001 0.285 None: Weight of Sandwich

T-STBDWINGTESPAR-001 0.271 None: Weight of Sandwich

T-STBDWINGCOVEPLT-001 0.115 Thickness change from 2 plies

to 6 plies: stress calculations

T-STBDWINGACTPLATE-001 0.156 Provisional open for reduction

T-STBDWINGTIPRIB-001 0.291 Maturation

T-WINGOBATTACHRIB-001 0.552 Maturation


172

Table 10 continued: -

The total wing weight after stressing and manufacturing maturation for the foam

spared wing in its final form was = (Outboard wing weight 3.958kg x 2 = 7.916) +

Centre wing weight 3.402kg = 11.318kg.

Densities Used:-

Aluminium 2024-T351 Plate = 2768kg/m

3

Fibreglass = 300g/m2 = 1200kg/m

3

Styrofoam = 40kg/m3 (although TRICAST 6 has currently been selected at time of

writing this is in conjunction with a different skin and spar material and ply lay up

therefore because Styrofoam was the material at sign – off this has used in this

analysis).

Sandwich foam R63.80 core = 90kg/m3

These values were produced using a CATIV V5 transposed model copy of the original

CATIA V4 model, with densities in kg/m3.

The total weight gain due to stressing sizing and maturation over the CDR

presentation was 3.262kg indicating that an initial starting weight of 6.00kg for the

wing in parametric analysis would have been more viable than the 2.85kg estimate.


T-WINGCENTRE-401.

CATIA

Weight kg


change.

T-WINGCENT TOPSKIN-001 0.847


plies and foam to 9 plies inner/ 6

plies outer: stress calculations.

T-WINGCENT LOWERSKIN-001 0.820


plies and foam to 9 plies inner/ 6

plies outer: stress calculations.

T-WINGCENTDNOSE-001 0.024 None: Weight of Styrofoam

T-WINGCENTTBOX-001 0.061 None: Weight of Styrofoam

T-WINGCENTTRAILINGEDGE-


T-CENTWINGLESPAR-001 0.143 None: Weight of Sandwich

T-CENTWINGTESPAR-001 0.134 None: Weight of Sandwich

T-WINGJOINTPORT-001 0.678 Manufacturing maturation

T-WINGJOINTSTBD-001 0.678 Manufacturing maturation


173

A-2: - Major wing component sketch book for part identification.

T-WINGOBTIPRIB-001/-002

T-WINGOBTOPSKIN-001/-002


174

T-WINGOBLOWERSKIN-001/-002

T-WINGOBDNOSE-001/-002


175

T-WINGOBLESPAR-001/-002

T-WINGOBTESPAR-001/-002


176

T-WINGOBTIPCORE-001/-002

T-WINGOBMAINCORE-001/-002


177

T-WINGOBACTPLATE-001/-002

T-WINGOBJOINTRIB-001/-002


178

T-WINGOBACTCOVER-001/-002

T-WINGOBTE-001/-002


179

T-WINGCTJOINTRIB-001/-002

T-WINGCTTOPSKIN-001


180

T-WINGCTLOWERSKIN-001

T-WINGCTDNOSE-001/-002

- 001

- 002


181

T-WINGCTTE-001/-002

T-WINGCTLESPAR-001

- 002

- 001


182

T-WINGTESPAR-001

T-WINGCTCORE-001

Due to the current material changes the author deemed the generation of a drawing set as

inappropriate to this thesis as the drawings would be out of date when the thesis was read

therefore the current F-35(JSF) policy has been adopted by producing the above sketch

book of all major structural wing components for identification sizing can be obtained

from interrogation of the Cranfield University solid model in the Terrasoar data pack NB

-001 is a PORT component -002 is a STARBOAD component, Blue = Aluminium, Green

= Glass cloth: Yellow = Foam.


183

A-3 Wing sign-off document.

TERRASOAR DEMONSTRATION AIRCRAFT

AIRFRAME STRUCTURE SIGN – OFF DOCUMENT

MAJOR AIRFRAME COMPONENT: - WING

Prepared for: - BAE SYSTEMS

Cranfield University

Prepared by: - G. Wardle.

Airframe Lead.


184

Document scope:-

The purpose of this document is to facilitate the release for manufacture of the major

airframe structural components of the Terrasoar UAV demonstration aircraft by formal

signature certifying that the major airframe component assembly and the constituent

substructure have been designed to meet the design intent.

This has been achieved by a visual analytical examination of CATIA solid models of the

major aircraft component and the constituent substructure, to ensure accurate sizing of

structural members, layout of systems, manufacturability and assembly of parts, and

maintainability by the customer, the structures engineer, and the airframe lead.

Document authority:-

This document hereby certifies the release for manufacture of all of the attached

structural members of this major aircraft component. After which all modifications will

be raised in issue, and will require re - certification.


185

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187

A-4 Wing drawing pack.

TERRASOAR WING MAJOR COMPONENT

DIMENSIONED DRAWINGS.

Contents: - 2-D drawings for all metallic parts, and dimensioned isometric

drawings for all composite and foam parts, for the complete wing.

Terrasoar 9th intake thesis Geoff Wardle

Documents

Transcript of Terrasoar 9th intake thesis Geoff Wardle