Jonathon Rowan-Undergraduate Project

The design and analysis of a supercritical aerofoil
for a concept commercial aircraft using CFD
package’s, incorporating materials research
Jonathon Michael Rowan
Supervised by Martin Fiddler
A Final Year Project Report
submitted to the
Faculty of Computing, Engineering and Technology
In partial fulfilment of the requirements for the degree:
Aeronautical Technology BSc.
Staffordshire University
Stoke-On-Trent
April 2014

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Jonathon M Rowan – Final Year Project
“Scientists investigate that which already is; engineers
create that which has never been.”
Albert Einstein
“Strive for perfection in everything you do. Take the best that exists
and make it better. When it does not exist, design it."
Sir Henry Royce

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Abstract
The modern aviation industry is focussed on efficiency and design sustainability. This document
reviews the design and analysis of a supercritical aerofoil section for use in the design of a
conceptual wing arrangement, specifically designed for the new Boeing 777x aircraft. This document
will examine three areas of wing design:
1. Aerofoil cross section.
2. Advanced aircraft wing design.
3. Advanced material application.
The principle focus of this investigation is to improve the lift- to-drag performance and thus
economic efficiency of aircraft though the examination of the three areas of wing design listed
above.
A supercritical aerofoil investigation was completed. The main objective was to design a more
efficient aerofoil section that allows high speed transonic flight, whilst retaining impressive lift-to-
drag ratio.
A concept aircraft investigation was completed. The main objective was to design a conceptual wing
arrangement for the new Boeing 777x aircraft that provides increased efficiency whilst flying in
cruise configuration.
An advanced material selection review was completed. The main objective was to select the
optimum material available to the aviation industry for use in the construction of the aerofoil.
Advanced materials offer significant benefits in efficiency due to reduced weight.
Advanced computational fluid dynamics software ANSYS Fluent was used during the course of this
final year project. This software was used as less sophisticated fluid dynamics software e.g. Cham
Pheonics, cannot run simulations accurately in the transonic speed envelope.
The use of ANSYS Fluent enabled trustworthy testing to show the design of aerofoils that provide
higher lift to drag those current aerofoils. It was also proven that advanced non-planar wing designs
show improved lift to drag to current conventional wing arrangements.
The results presented in this Final Year Project confirm that advanced box wing designs (such as the
Lockheed Martin Box Plane) can offer significant aerodynamic improvements whilst retaining many
conventional parameters of aircraft design. The results show that advanced box wing aircraft offer a
second option in conceptual aircraft design as an alternative to blended wing designs.

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Contents
Abstract...................................................................................................................................................2
Schedule of figures..................................................................................................................................5
Acknowledgements.................................................................................................................................8
Chapter 1 - Introduction .........................................................................................................................9
1.1 - Aims and objectives........................................................................................................11
Chapter 2 - Aerofoil Design – shape & cross-section............................................................................13
2.1 - Aerofoil Design Parameters..................................................................................................14
2.2 Supercritical aerofoil design....................................................................................................15
2.3 - Supercritical aerofoil benefits...............................................................................................18
2.4 - Examples of current transonic aerofoils...............................................................................19
2.5 – Supercritical aerofoil modelling...........................................................................................21
2.6 - Aerofoil families....................................................................................................................22
2.7.1 Initial aerofoil performance testing.............................................................................................23
2.7.2 - Experiment verification......................................................................................................24
2.7.3 - Primary aerofoil testing results..........................................................................................25
2.7.4 - Initial test results ...............................................................................................................31
2.7.5 – Supercritical design phase 2 - aerofoil research ...............................................................32
2.7.6 - Supercritical design phase 2- simulation results ...............................................................33
2.8 Supercritical design phase 2 - Redesign approach..................................................................35
2.8.1 - Supercritical aerofoil testing analysis ................................................................................40
2.8.3 Decision and aerofoil design evaluation..............................................................................42
Chapter 3 - Advanced aircraft wing design...........................................................................................43
3.1 - Concept commercial aircraft research..................................................................................44
3.2 - Initial design concept......................................................................................................48
3.3 - Similar projects/inspiration ..................................................................................................49
3.4 - Current aircraft lift to drag performance..............................................................................52
3.5.1 – Concept design plan and objectives..................................................................................53
3.5.2 - Angle of incidence..............................................................................................................54
Experiment analysis ......................................................................................................................68
Conclusion on advanced box wing designs for commercial use...................................................70
Chapter 4 – Materials selection for advanced box wing design...........................................................71

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4.1 Materials research ..................................................................................................................72
4.2 - Advantages of advancements in material technology..........................................................74
4.3 - Material selection using CES-Edupack..................................................................................75
4.4 Material selection conclusion .................................................................................................81
5.0 - Rendered Images of final design...................................................................................................84
Discussion..............................................................................................................................................86
Conclusion.............................................................................................................................................89
Recommendations................................................................................................................................91
Bibliography ..........................................................................................................................................92
APPENDIX A – Drag terms....................................................................................................................95
APPENDIX B - ‘Technical literature review’..........................................................................................97
APPENDIX C – Aerofoil database links ...............................................................................................101
Appendix D – ‘Boeing 777 Design parameters’..................................................................................103

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Schedule of figures
Figure 1 - Shows the typical parameters that are used define the design of an aerofoil. The
terminology used is explained below. (Cantrell, 2012) ........................................................................14
Figure 2 - An example of a supercritical aerofoil and where the shockwave is produced (Aircraft
Recongnition, 2014)..............................................................................................................................15
Figure 3 - Co-ordinate sketch of National Propulsion Laboratory (NPL) 9510 Transonic Aerofoil
(University, Illinois, 2013) .....................................................................................................................19
Figure 4 - Co-ordinate sketch of the McDonnell/Douglas DSMA-523 Transonic Aerofoil with sharp
trailing edge (University, Illinois, 2013) ................................................................................................20
Figure 5 - Co-ordinate sketch of the NASA SC(2)-0610 supercritical aerofoil ......................................20
Figure 6 - NPL 9510 AEROFOIL - NPL 9510 transonic aerofoil modelled using Creo Parametric design
modeller................................................................................................................................................21
Figure 7 - McDonnell/Douglas DSMA-523 transonic aerofoil with sharp trailing edge modelled to
exact coordinates using Creo Parametric .............................................................................................21
Figure 8 - NASA SC(2)-0610 supercritical aerofoil modelled to exact coordinates using Creo
Parametric.............................................................................................................................................21
Figure 9 - Shows a pressure contour colour map for the NPL 9510 supercritical aerofoil...................25
Figure 10 - The pressure contour map for the Dsma-523A Aerofoil produced by ANSYS fluent post
CFD processor. ......................................................................................................................................26
Figure 11 - The relative pressure contour map for aerofoil SC(2)-0610...............................................27
Figure 12 - Shows the relative contour map for S.U.A.D.1 which shows poorer performance in
comparison to the researched aerofoils...............................................................................................28
Figure 13 - Pressure contour around S.U.A.D.2 which offers the lowest lift to drag of all tested
aerofoils. This is caused by a lack of pressure under the lower surface...............................................29
Figure 14 - Pressure contour map around the best performing S.U.A.D. family aerofoil to this point.
The substantial increase in lift was produced by a much more concave aft section. ..........................30
Figure 15 - Primary aerofoil testing bar chart showing the DSMA-523A aerofoil to have performed
best in terms of lift to drag. ..................................................................................................................31
Figure 16 - NYU/Grumman K-1 transonic aerofoil................................................................................32
Figure 17 - RAE 2822 transonic Aerofoil...............................................................................................32
Figure 18 - Pressure contour map from the simulation using the NYU-Grumman K-1 aerofoil. This
aerofoil shows significant performance improvements.......................................................................33
Figure 19 - Pressure contour map showing the relative pressures around the RAE 2822 transonic
aerofoil. This shows the highest lift to drag result of any researched aerofoil....................................34
Figure 20 - The first redesigned aerofoil using parameters from other aerofoils. This image shows the
pressures around the S.U.A.D.4 project designed aerofoil. .................................................................36
Figure 21 - S.U.A.D.5 pressure contours showing high pressure build up underneath especially in aft
section of the aerofoil...........................................................................................................................37
Figure 22 - Shows the pressure contour map from the highest lift to drag ratio aerofoil tested in the
simulation S.U.A.D.6. ............................................................................................................................38
Figure 23 - The table below shows the results generated from the aerofoil testing completed in
ANSYS Fluent.........................................................................................................................................40

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Figure 24 - Graph – showing the complete results from the aerofoil simulations shown on a bar
chart......................................................................................................................................................40
Figure 25 - One example of a blended wing concept currently being analysed is the Boeing x-48.....45
Figure 26 - Shows potential design ideas for a morphing wing and the structural aerodynamic
changes it would make in flight. (Centre for Mechanics and Control, 2004) .......................................46
Figure 27 - This image shows the performance of vortex drag production for a number a of different
aircraft wings. The picture shows that a box plane produces the least relative vortex drag when
compared to a number of different designs.........................................................................................47
Figure 28 - Potential design ideas include the typical box wing configuration and Boeings fluid wing
configuration.........................................................................................................................................48
Figure 29 - The Lockheed Martin Advanced Box Plane was the main inspiration for this project – it is
often presented as a future aircraft variant. The design is a typical advance box wing. .....................49
Figure 30 - The final design of the Prandtl concept Aircraft. This design was the produced at Pisa
University. The project had the aim of designing an aircraft that would carry over 800 passengers..50
Figure 31 - Shows the comparison in optimum induced drag of a biplane and optimum induced drag
of ‘best wing systems’...........................................................................................................................50
Figure 32 - Shows the A9 Dragonfly - a medium size long haul aircraft design using the advanced box
wing design. ..........................................................................................................................................51
Figure 33 - provides the Lift to drag performance for the commercial aircraft shown: ......................52
Figure 34 - research results - incidence of incidence. ..........................................................................54
Figure 35 - Show the pressure contours around aerofoil S.U.A.D.6 at its optimum angle of attack
which was 2 degrees where its lift-to-drag performance is over 60....................................................55
Figure 36 - Show the velocity contours around aerofoil S.U.A.D.6 at its optimum angle of attack
which was 2 degrees where its lift-to-drag performance is over 60....................................................55
Figure 37 - A 777 replica designed to the same design parameters found on the Boeing technical
information website..............................................................................................................................56
Figure 38 - Pressure 1 contours around the surface of a 777 ..............................................................56
Figure 39 - Pressure 2 contours around surface of a Boeing 777.........................................................57
Figure 40 - Velocity 1 around a Boeing 777..........................................................................................57
Figure 41 - Velocity 2, velocity contours around surface of a Boeing 777 ...........................................57
Figure 42 - 777-300 scale model with engines attached......................................................................58
Figure 43 - Pressure contours around a Boeing 777 with engines.......................................................58
Figure 44 - Pressure contours image 2 .................................................................................................59
Figure 45 - Velocity contours around the surface of a Boeing 777 with engines.................................59
Figure 46 - Velocity contours image 2 ..................................................................................................59
Figure 47 - The first concept wing design for commercial aircraft is a typical box wing design this
design is titles the 777x concept 1........................................................................................................60
Figure 48 - Pressure contours around the surface of 777x concept 1..................................................60
Figure 49 - Pressure contours image 2 .................................................................................................61
Figure 50 - Velocity contours around the surface of 777x1 .................................................................61
Figure 51 - Velocity contours image 2 ..................................................................................................61
Figure 52 - 777x concept 2 is an advanced box design original to the project.....................................62
Figure 53 - Pressure Contours around 777 x concepts 2......................................................................62
Figure 54 - Pressure contours image showing an underside view of the 777x concept2 ....................63

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Figure 55 - Velocity contours around the 777 x concept 2...................................................................63
Figure 56 - Velocity contours showing the under view of the 777x concept 2 ....................................63
Figure 57 - The second designed advanced box wing original to this project titles 777x concept 3 ...64
Figure 58 - Pressure contours around 777 x concepts 3 ......................................................................64
Figure 59 - Pressure contours showing different view of 777x concept 3 ...........................................65
Figure 60 - Velocity contours of 777x concept 3 ..................................................................................65
Figure 61 - Velocity contours image 2 of the 777x concept 3 ..............................................................65
Figure 62 - The Final design is more of a biplane than a box plane this aircraft design is titles 777x
concept 4...............................................................................................................................................66
Figure 63 - Pressure contours around 777 x concepts 4 ......................................................................66
Figure 64 - Second image of pressure contours around 777 x concepts 4...........................................67
Figure 65 - Velocity contours around 777 x concepts 4 .......................................................................67
Figure 66 - velocity contours around 777 x concepts 4........................................................................67
Figure 67 - This table of figures shows the performance of lift to drag ratio of each aircraft design
tested. It is demonstrated that each of the non-planar advanced box style wings offers significant
advantages in lift to drag performance. ...............................................................................................68
Figure 68 - Lift to drag results of each aircraft shown on a line graph.................................................68
Figure 69 - Similar graph to figure 68 however this time the information is represented on a bar
chart......................................................................................................................................................69
Figure 70 - A bar chart comparing a current Boeing 777 against the best performing box wing the
777x concept 3......................................................................................................................................69
Figure 71 - shows the percentage of composite materials in the Boeing fleet as they have evolved
the 787 is now made up of 50% composite with the entire wing structure made using.....................73
Figure 72 - Shows the advancement in material technology and predicts future technology levels the
current technologies being used in the 777 are pre 2000 need a drastic overall. ...............................74
Figure 73 - An image demonstrating the 4 main design needs of aircraft wings.................................75
Figure 74 - Graph showing young’s modulus against density, only the colour circles are materials that
have passed the limit stage...................................................................................................................78
Figure 75 - When the selection line is used the following materials are the best 5 shown in this figure
..............................................................................................................................................................78
Figure 76 - Elastic limit vs density graph only the coloured circles have passed the limit stage .........79
Figure 77 - Yield strength vs density, using a selection line to find the top 5 performing materials...79
Figure 78 - Fracture toughness vs density............................................................................................80
Figure 79 - Fracture toughness vs density using a selection line to select the top 5 materials. ..........80

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Acknowledgements
I would like to take this opportunity to thank to a number of people that helped and encouraged me
in the completion of this Final year Project.
Firstly I would like to thank Staffordshire University for providing excellent facilities throughout my
three years studying at the University. These facilities have helped me achieve the objectives of my
Final Year Project.
I would also like to thank Mr Martin Fiddler who has provided an excellent learning culture
throughout the course and provided sound and reliable support throughout the duration of this Final
Year Project. Mr Fiddler’s guidance has been invaluable and he has helped me to strive to achieve
the projects objectives and to stay focussed.
I would like to express my gratitude to Mrs Debi-Marie Roberts whom I relied upon for a second
opinion and who has given me the knowledge of aerodynamics has allowed me to generate high
quality data for this project. Debi was always there in times when I was struggling with my data
collection.
I would also like to thank my Mother and Father who have provided me with constant support and
who are always a reliable source of help and advice. Their support has allowed me to work to
achieve the best possible outcome from my studies at Staffordshire University. I hope to make them
proud with the result I obtain for this project.
Jonathon Michael Rowan.

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Chapter 1 - Introduction
Overview
The aviation industry is growing at a remarkable rate. This is driven by a Global population increase
and the increase in demand for air travel and transport from emerging and powerful economies in
Countries such as India, China and Brazil. Air travel is becoming increasingly accessible to more and
more people - this means more aircraft, more emissions and more pollution.
It is estimated that passenger and cargo air-traffic will grow by between two and three hundred per
cent in the next two decades. It is expected that medium and long range routes Worldwide will see a
particularly significant increase in air-traffic. (Frediani, 2005)
Given this increasing demand, the aviation industry is striving for efficiencies and giving close
attention to finding ways to make the industry more environmentally friendly and sustainable.
Making progress on these very significant challenges is the task facing all aeronautical design teams
around the World today.
This document details the work undertaken to design and analyse a supercritical aerofoil and wing
arrangement for a concept superjumbo variant aircraft. This concept aircraft would compete in the
aviation marketplace with aircraft such as the Airbus A380, the Boeing 747-700 and the Boeing 777x.
The project centres around 3 mains areas of aircraft wing design with added innovative aerodynamic
features. The chapter headings in the main body of this Final Year Project document will be:
- Aerofoil design - shape / Cross-section for supercritical aerofoils.
- Advanced aircraft wing design (including ‘Aerodynamic improving features’).
- Materials research and Optimum material selection
The project focuses upon innovation in efficiency and sustainable design. Inspiration for the project
came from a number of sources in particular research undertaken on supercritical aerofoils, and the
work completed by Charles Harris (Southampton University 1990) formed the basis of the
investigation into supercritical aerofoil design.
The design of a conceptual large commercial aircraft was inspired by the Lockheed Martin
‘advanced box wing project’ and the Prandtl Plane investigation and design completed at Pisa
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Within each of the three main chapters there will be a detailed explanation of the research
undertaken as the project developed in addition to a clear demonstration of the results achieved
and analysis undertaken.
Project motivation
Justification for the project lies within the evident desire of the aviation industry for fuel efficiency
and the consequent reduction on the cost of air travel. A key aim of this design project was to
greatly improve the aerodynamic performance of commercial aircraft. This is coupled with the aim
of producing this improvement to achieve a key output – the capability to carry more people at the
lowest possible cost. Mass air travel will become increasingly popular with the travelling public and
airline operators will wish to operate the most cost effective service in the delivery of service to their
customers.
The project provides analysis on the complex transitional flows between subsonic and supersonic
speeds and aims to show innovative design in aerofoil shape and wing structure - with an emphasis
placed upon innovation in efficiency.
In aviation there are two main types of drag that can affect the performance of the aircraft thus
increasing fuel costs. These are 1
Induced Drag and 2
Parasitic Drag. A significant objective in the
creation of the concept aircraft design developed during the course of this project was to make the
concept aircraft meet the challenge of a ‘cross-over’ design that delivers improved performance at
low speeds where induced drag is the dominant drag force present and in cruise configuration
where parasitic drag is predominant.
Meeting the challenge of innovation is extremely important in the aviation industry and innovation
provides a strong focus for this project. The commercial success of the innovative Boeing 787
Dreamliner demonstrates the importance of innovations that deliver cost and other benefits to
airline operators and passengers. Boeing has a significant order back-log due to the appreciation of
the benefits the Boeing 787 Dreamliner delivers when compared to other aircraft.
There is no doubt that the Boeing 787 Dreamliner’s design includes a number of innovations that
deliver significant reductions in fuel consumption and total emissions in addition to other benefits.
Whilst recognising the achievements of the Boeing design team, this project has been completed
1
Defined in appendices
2
Defined in appendices

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with firm belief that more radical designs have the potential to deliver more efficient and
economical flight.
1.1 - Aims and objectives
The overall aims of the project were to; analyse a supercritical aerofoil and wing arrangement for a
concept superjumbo variant aircraft; provide valuable research in the design of a suitable
supercritical ‘high speed’ transonic aerofoil design and to demonstrate this design could be
introduced in aircraft manufacture using the most appropriate high performing, low weight
advanced materials.
From the outset, the objectives of the project were to deliver:
 An aerofoil design that offers high cruise speed and that provides improved aerodynamic
performance for improved efficiency in transonic transport.
 A project that has commercial value with emphasis placed on innovation in the three areas
of aircraft wing design noted above.
 A concept wing design for a commercial aircraft that demonstrates innovation that delivers
improved aerodynamic performance offering new levels of efficiency.
 The selection of the optimum performing material for use in the construction of a modern
aircraft.
Time constraints, data collection difficulties and data accuracy were challenges faced in the
completion of this project. Two solutions dealt with these issues:
 Careful time planning and strict adherence to deadlines in the completion of the various
work stages.
 The projects use of advanced computational fluid dynamics ANSYS Fluent software to
produce results that are reliable and can be used to draw significant and meaningful
conclusions.
The projects main conclusions demonstrate that increased aircraft efficiency is possible through
improvements in the three areas of design researched. Improved lift-to-drag results can be achieved
by improvements in aerofoil sections and by advanced non-planar wing design. In addition,

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improvements in aircraft performance can be achieved through the use of advanced materials in
there construction.

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Chapter 2 - Aerofoil Design – shape & cross-section
This chapter of focusses on the design analysis and review of supercritical aerofoils. This research
will be applied later in the project in the design of a conceptual wing. Research will focus on the
design parameters of supercritical aerofoils designed for transonic. This research is important as it
will reveal why the performance of such supercritical aerofoils makes them suitable for commercial
use.
Analysis of the design parameters and the aerodynamic performance of three transonic aerofoils will
be considered and reviewed. These transonic aerofoils have been designed by the some of the most
experienced aviation businesses in the world – it is therefore important to understand them fully
before contemplating design improvements.
New designs will be then be created and undergo the same testing as the commercially designed
aerofoil. This will be followed by performance evaluation analysis of the new aerofoil designs in
different configurations and scenarios.
Following this analysis, the research will lead to a recommendation with justification as to the
aerofoil design that is to be used for the conceptual aircraft.
Chapter 2 objectives
The objectives set for this chapter are to:
 Describe and discuss the parameters used when designing aerofoils.
 Understand the application of parameters in design of transonic supercritical aerofoil
sections and explain the benefits achieved for commercial use.
 Research and understand the design and performance of current aerofoils designed by
important aviation companies.
 Design a selection of transonic aerofoils following examination of other transonic aerofoils.
 Conduct accurate and testing on a selection of the newly designed aerofoils.
 Complete analysis on the best performing newly designed aerofoils.
 Select the best performing aerofoil from those tested and make clear the reasons it will be
used in the design of the advanced box-wing reconfiguration of a Boeing 777x.

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2.1 - Aerofoil Design Parameters
Figure 1 -shows the typical parameters that are used define the design of an aerofoil. The terminology used is explained
below. (Cantrell, 2012)
Leading edge – The frontal point of the aerofoil where first contact with the air occurs.
Trailing edge - Where the upper surface and lower surface meet at the rear of the aerofoil – the final
point of the aerofoils body.
Chord line - The straight line that connects the leading and trailing edges of the aerofoil.
Angle of attack – The angle at which the chord line faces the oncoming air.
Chord - The length of the chord line from leading edge to trailing edge. This is the characteristic
longitudinal dimension of the aerofoil.
Camber – The difference in curvature between the top surface and the bottom surface. This is
known as the asymmetrical difference.
Maximum camber – The point at which the distance between the upper and lower surfaces is at its
maximum.
(Cantrell, 2012)
The remaining design parameters are not shown in Figure 1:
Mean camber line – The mean camber line is a line drawn halfway between the upper and lower
surfaces. The chord line connects the ends of the mean camber line.
Upper camber - The curvature and shape that is between the upper surface and the chord line.
Lower camber – The curvature shape that defines the chord line and the lower surface.
(Cantrell, 2012)

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2.2 Supercritical aerofoil design
A supercritical aerofoil is an aerofoil specifically designed for higher efficiency performance in the
transonic speed range (Skybrary, 2012). When an aircraft with a conventional wing nears the speed
of sound (Mach 1), air flowing across the top of the wing moves faster than Mach 1 and becomes
supersonic. This creates a shock wave on the wing's upper surface despite the fact that the aircraft,
as a whole, has not exceeded Mach 1 (NASA, 2004). The aircraft at this point is flying at what is
called its ‘critical speed’. The shockwave causes the smooth flow of air hugging the wing's upper
surface (the boundary layer) to separate from the wing and create turbulence. Separated boundary
layers are like the wakes left in water behind a speed boat - the air is unsteady and churning and
drag increases. This increases fuel consumption, can lead to a decrease in speed and cause excessive
component vibrations.
A major aim in design of a supercritical aerofoil is to delay the onset of wave drag that becomes
apparent through shockwave formation upon the upper surface as the air breaks the speed of
sound. (NASA, 2004)
Figure 2- An example of a supercritical aerofoil and where the shockwave is produced (Aircraft Recognition, 2014)
There are typical designs features associated with supercritical aerofoils that differ from traditional
aerofoil design parameters. These features are:
 A flattened upper surface which stops the air speeding up as much as it does on a curved
upper surface.

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 A highly curved lift producing rear/aft section. This is to compensate for the lack of lift the
front end of the aerofoil produces.
 A large leading edge radius which slightly slows the air so that it does not reach the critical
speed where the boundary layer separates creating unwanted drag.
These features allow for greater performance in drag reduction and make the aerofoils invaluable to
high performance transonic flight.
Modern aircraft can cruise at a higher subsonic speed well into supercritical range. Less drag at
these speeds is mainly to these supercritical aerofoils. An aircraft designed with supercritical wing
sections will use less fuel than it would otherwise consume. Higher subsonic cruise speeds and less
drag means that aircraft can travel faster using less fuel thus providing increased range and payload.
These factors help reduce the cost of passenger tickets and cargo transport.
Design objectives and challenges
Supercritical sections refer to a special type of aerofoil that is designed to operate efficiently within
the transonic speed range. The design objectives of a supercritical aerofoil are:
 To carry as much lift as is practical on the aft potion of the section where the flow is
subsonic. (Tamkang University , 2012)
 To ensure the aerofoils upper surface are as Flat and smooth as possible - this prevents the
air reaching critical Mach speeds to early over the upper surface of the aerofoil body.
 To ensure the rear/aft section of the aerofoil is suitably cambered to produce high lift - this
is to compensate for the lack of lift that the smooth upper surface creates. (Tamkang
University , 2012)
 To ensure and efficient level of lift is produced on the forward portion of the upper surface.
As the Local airspeed over the top surface increase the pressure near the nose is diminished.
Without additional blunting of the nose extra lift that could be generated here will be lost.
The design challenges associated with supercritical aerofoils are:
 Reduction of shockwave formation. This generally means limiting the minimum pressure
coefficient occurring between the upper and lower surface.

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 Excessive rear production of lift. This can lead to negative pitching moments with trim drag
and structural weight constraints. (Harris, 1990)
 In extreme situations the adverse pressure gradient caused on the rear lower surface can
produce boundary layer separation. (Harris, 1990)
 Extremely thin trailing edges to the aerofoils/ wing section have proved beyond
manufacturing capabilities. This means that certain theoretical aerofoils are impossible to
use into aircraft design. (Harris, 1990)
 Supercritical sections and similar shock-free designs often are very sensitive to Mach and lift
coefficient and can perform poorly in operating conditions for which they were not
designed. (Harris, 1990)
 "Drag creep" can often be a common occurrence. This happens when substantial section
drag increases with Mach number. This can occur even at speeds below the design value.
(Harris, 1990)
It is extremely important to be cautious in the design of supercritical aerofoils. Several major designs
of supercritical wing sections have looked promising initially but have created serious problems
when incorporated into an aircraft design. (Harris, 1990)
The Supercritical design challenge is to create an aerofoil section with high lift without causing
strong shock waves thus reducing drag caused by boundary layer separation. A supercritical aerofoil
can generally tolerate some supersonic flow without drastic increases in the drag. A generalised rule
for supercritical wing sections is that the maximum Mach numbers should not exceed Mach 1.2 to
1.3 to successfully minimise the effect of drag caused by shockwaves. (Tamkang University , 2012)

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2.3 - Supercritical aerofoil benefits
The use of supercritical aerofoils is vital in modern day aviation. There use delivers the following
benefits:
1. Significantly reduced shockwave induced boundary layer separation. This is achieved bcy
keeping flow over the aerofoil smooth and at regulated speeds.
2. The production of a smaller, weaker shock wave at a position further aft on the wing
than tradition aerofoils.
3. The potential for more efficient wing design. The supercritical aerofoil allows for a
reduction in wing sweep or an increase in wing thickness without the corresponding
increase in wave drag that would be associated with a typical aerofoil.
4. Higher lift to drag ratio associated with supercritical aerofoils allows for reduction in fuel
consumption.
(Harris, 1990)
Lift to drag
The main performance parameter used to compare the test supercritical aerofoils will be the results
for lift-to-drag. Lift to drag is a measure of the lift produced divided by the resultant drag forces
acting upon the aerofoil. There are many advantages in maximising the lift to drag of supercritical
aerofoil sections and aircraft wings in general.

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2.4 - Examples of current transonic aerofoils
The first stage in the design of a supercritical aerofoil to be used in combination with the concept
wing for a commercial aircraft is to review the performance of a selection of supercritical and
transonic aerofoils. Following research of a number of supercritical aerofoils I have chosen three
‘candidate’ supercritical aerofoils on which to conduct further research. The aerofoils I have chosen
to review are:
1. National Propulsion Laboratory (NPL) 9510 Transonic Aerofoil
2. McDonnell/Douglas DSMA-523 Transonic Aerofoil with sharp trailing edge
3. NASA SC(2)-0610 Aerofoil
These supercritical aerofoils where remodelled using Creo Parametric 2.0 and then tested, analysed
and compared in an aerodynamic test. This remodelling is undertaken to ensure that precisely the
same test parameters are applied to each aerofoil. This will mean that comparative analysis between
the performance of these supercritical aerofoils and the performance of the project designed
supercritical aerofoil will produce a meaningful comparison and analysis.
1. National Propulsion Laboratory (NPL) 9510 Transonic Aerofoil
This aerofoil was designed by a NASA project for use by aircraft in transonic flight. It is purely
experimental and has not been used in any commercial or military aircraft. However, the design is
interesting including design features that may assist the aims of this project. (Jenkins, 1983)
Figure 3 - Co-ordinate sketch of National Propulsion Laboratory (NPL) 9510 Transonic Aerofoil (University, Illinois, 2013)

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2. McDonnell/Douglas DSMA-523 Transonic Aerofoil with sharp trailing edge
McDonnell/Douglas is a large and successful aviation business that has been responsible for a
number of successful commercial jet aircraft. A supercritical aerofoil designed by such a prestigious
company warrants close examination and investigation.
Figure 4 - Co-ordinate sketch of the McDonnell/Douglas DSMA-523 Transonic Aerofoil with sharp trailing edge
(University, Illinois, 2013)
3. NASA SC (2)-0610 Aerofoil
This supercritical aerofoil is of particular interest as it has been identified as the possible ‘root
supercritical aerofoil’ profile of the supercritical aerofoil used on the Airbus A380 aircraft (University,
Illionois, 2013)
Figure 5 - – Co-ordinate sketch of the NASA SC (2)-0610 supercritical aerofoil (University, Illinois, 2013)
Prior to aerodynamic testing and comparison with the concept aerofoil designs created as part of
this project, these aerofoils were remodelled using Creo Parametric 2.

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2.5 – Supercritical aerofoil modelling
Figure 6 - NPL 9510 AEROFOIL - NPL 9510 transonic aerofoil modelled using Creo Parametric design modeller.
Figure 7 - McDonnell/Douglas DSMA-523 transonic aerofoil with sharp trailing edge modelled to exact coordinates using
Creo Parametric
Figure 8 - NASA SC (2)-0610 supercritical aerofoil modelled to exact coordinates using Creo Parametric

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2.6 - Aerofoil families
When aerofoils are designed they are designated into ‘aerofoil families’. These ‘families’ use the
same initial name (usually an acronym) followed by the allocation of a number to identify the
development stage the aerofoil has reached. There are many famous aerofoil families most notably
the NACA aerofoil series.
The aerofoil family created during this project will use the acronym S.U.A.D.
S.U.A.D. - STAFFORDSHIRE UNIVERSITY AERODYNAMIC DESIGN
S.U.A.D. Aerofoils follow a logical numbering system with the first design known as S.U.A.D.1
followed logically by S.U.A.D.2. The approach provides a simple method of identifying the
development stage of the supercritical aerofoils.

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2.7.1 Initial aerofoil performance testing
Each aerofoil was tested using precisely the same simulation parameters with the aerofoils in cruise
configuration. The test was performed using ANSYS Fluent – this software can perform highly
detailed analysis at transonic speeds. Initially, the tests were completed using the Cham Phoenics
software package. However this software package did not provide accurate information and the
results obtained proved unusable.
The primary objectives of the test were to establish which supercritical aerofoil was best performing
in terms of lift to drag ratio and to study the performance of each supercritical aerofoil in close
detail.
The simulation parameters were:
 Air density = 0.1842 – Representing FL 300
 Alpha (Aerofoil angle of attack) = 4.0
 Airspeed = 0.8 Mach

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2.7.2 - Experiment verification
Prior to testing, the test parameters had to be determined and it was important to verify that the
results that would be obtained from the software package. Verification of the results is vital to
ensure there would be a correct understanding of the simulations results and to ensure correct
conclusions were reached.
Verification was achieved by reference to NASA Contractor Report 166005. This presents the results
of an experiment carried out on the NPL 9510 aerofoil at The University of Southampton by S. W. D.
Wolf. The report logs results for both the coefficient of drag and coefficient of lift and this can be
used to deduce the lift to drag ratio of the supercritical aerofoil tested. The simulation parameters
that were used by S.W.D. Wolf were applied in a simulation using ANSYS Fluent.
Comparative test results:
NPL SECTION ANALYSIS 9510 PROJECT Simulations NPL 9510
ALPHA = 4.0 Alpha = 4.0
MACH NO. =0.8002 Mach number 0.8
LIFT TO DRAG RESULT 32.58 LIFT TO DRAG RESULT 34.3
the results obtained from the simulation completed during the course of this project are extremely
close to the results detailed in NASA Contractor Report 166005. Therefore it is reasonable to
conclude:
 The test parameters of the simulations conducted during the course of this project are
correct.
 The results obtained are verified sufficiently and that accurate conclusions can be reached in
future tests.

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2.7.3 - Primary aerofoil testing results
NPL9510 – LIFT TO DRAG RESULT OF 34.3
PRESSURE
Figure 9 - Shows a pressure contour colour map for the NPL 9510 supercritical aerofoil.
VELOCITY
Figure 9a - Shows the velocity contour colour map around the NPL 9510 aerofoil.

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DSMA-523a - LIFT TO DRAG RESULT OF 35.9
PRESSURE
Figure 10 - The pressure contour map for the Dsma-523A Aerofoil produced by ANSYS fluent post CFD processor.
VELOCITY
Figure 10a – The velocity contour map for the Dsma-523A Aerofoil.

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SC (2)-0610 - LIFT TO DRAG RESULT 28.5
PRESSURE
Figure 11 - The relative pressure contour map for aerofoil SC (2)-0610
VELOCITY
Figure 11a - The relative velocity contour map for the aerofoil SC (2)-0610

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S.U.A.D.1 – LIFT TO DRAG 28.7
PRESSURE
Figure 12 – Shows the relative contour map for S.U.A.D.1 which shows poorer performance in comparison to the
researched aerofoils.
VELOCITY
Figure 12a - Shows velocity contours around the S.U.A.D.1 aerofoil. This image shows significant velocity increase over
the leading edge on the top surface - this is a negative feature for a supercritical section as a desired design features
seeks to limit this.

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S.U.A.D.2 LIFT TO DRAG – 26
PRESSURE
Figure 13 - Pressure contour around S.U.A.D.2 which offers the lowest lift to drag of all tested aerofoils. This is caused by
a lack of pressure under the lower surface
VELOCITY
Figure 13a - A velocity contour map around the S.U.A.D.2 aerofoil. It would appear there has been boundary layer
separation quite early along the top section because the air flow is disturbed and slower.

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S.U.A.D.3 – LIFT TO DRAG RATIO 31
PRESSURE
Figure 14 - Pressure contour map around the best performing S.U.A.D. family aerofoil to this point. The substantial
increase in lift was produced by a much more concave aft section.
VELOCITY
Figure 14a - Velocity contour map for the S.U.A.D.3 There is early boundary layer separation shown on the upper
surface.

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2.7.4 - Initial test results
Figure 15 - Primary aerofoil testing bar chart showing the DSMA-523A aerofoil to have performed best in terms of lift to
drag.
The results from the Initial testing show that the concept supercritical aerofoils created for the
project have failed to generate higher lift-to-drag results than the researched aerofoils. The highest
lift to drag generated was 35.9 by the DSMA-523a aerofoil - this is an impressive result.
The objective at this stage of the project was to design a supercritical aerofoil that performs better
than current aerofoils and this was proving to be a significant challenge.
The main issue in the design of the concept aerofoils was unsmooth sections. This created an early
boundary layer separation which caused a significant increase in drag produced by the aerofoils.
The concept aerofoil did generate reasonable and improved levels of lift however it was the drag
issue that contributed to the reduction in the lift-to-drag results.
0 5 10 15 20 25 30 35 40
SC_6010
NPL
Dsma
SUAD
SUAD2
SUAD3
Primary Aerofoil Testing

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2.7.5 – Supercritical design phase 2 - aerofoil research
Following the unsuccessful initial design stage, further focus was placed on the design effort and
further research was undertaken. Two further supercritical aerofoil designs where selected and
tested using the same test parameter as with the previous experiment.
The secondary aerofoil research focused on conceptual supercritical aerofoil sections that appeared
to offer an opportunity for improved lift to drag ratio results. (University, Illionois, 2013)
The two supercritical aerofoils selected for further analysis where:
Figure 16 - NYU/Grumman K-1 transonic aerofoil
Figure 17 - RAE 2822 transonic Aerofoil

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2.7.6 - Supercritical design phase 2- simulation results
NYU/Grumman K-1 transonic aerofoil – LIFT TO DRAG 46.6
PRESSURE
Figure 18 – Pressure contour map from the simulation using the NYU-Grumman K-1 aerofoil. This aerofoil shows
significant performance improvements.
VELOCITY
Figure 18a - Velocity contour map around the experimental aerofoil section NYU-Grumman K1.

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RAE 2822 transonic Aerofoil- LIFT T0 DRAG 48.5
PRESSURE
Figure 19- Pressure contour map showing the relative pressures around the RAE 2822 transonic aerofoil. This shows the
highest lift to drag result of any researched aerofoil.
VELOCITY
Figure 19a - Shows the velocity contours from the simulation using the RAE 2822.

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2.8 Supercritical design phase 2 - Redesign approach
Instead of further remodelling, it was considered the most appropriate approach to creating an
improved aerofoil section would be to:
 Identify the best performing features of the researched aerofoils
And
 combine these features as far as is practically possible with the features of other aerofoils
that demonstrate excellent performance
The hypothesis was that this combination should produce an aerofoil that has the performance
advantages of both. The target was to improve on the already impressive result from the RAE 2822
aerofoil of a 48.5 lift to drag ratio.
This redesign approach was needed as it was clear that all of the researched aerofoils were designed
with extremely smooth curvature. The Creo Parametric 2.0 design software was unable to replicate
this feature. The lack of extremely smooth curvatures resulted in boundary layer separation over the
upper surface of the test aerofoils. As a result of the turbulent air increased drag figures where
shown in the simulation. These drag figures effected the overall efficiency and performance of the
test supercritical aerofoils sections.
There is specific design parameters needed for a supercritical aerofoil, operating within the
transonic speed range, to perform highly. The optimum parameters are:
1. Flattened smooth upper aerofoil surface which stops the air reaching critical Mach speeds to
early over the upper surface.
2. A highly curved lift producing rear/aft section. This is to compensate for the lack of lift that
the smooth upper surface creates.
3. A large leading edge radius to keep the pressure gradient positive at the front end aiding the
production of lift.

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S.U.A.D.4 – LIFT TO DRAG 41
The first redesign concept was to use the top surface of the NYU-GRUNMAN K1 aerofoil and use a
high lift producing - rear/aft section surface of the McDonnell/Douglas DSMA-523 aerofoil. In the
primary design phase, the DSMA-523 aerofoil was the best performing. The aim was that, through
the use of a smoother flatter upper surface, the S.U.A.D.4 would perform better and show
significantly improved results.
PRESSURE
Figure 20 - The first redesigned aerofoil using parameters from other aerofoils. This image shows the pressures around
the S.U.A.D.4 project designed aerofoil.
VELOCITY
Figure 20a – The velocity contour map showing impressive flow over the top surface of the aerofoil of S.U.A.D.4

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S.U.A.D.5 – LIFT TO DRAG RATIO 45
The same re-design approach was used in the design on S.U.A.D.5. This aerofoil utilises the RAE
2822’s upper surface that is smooth and particularly flat. Once again the rear aft section was
changed in an attempt to generate more lift. The rear section used is from the NPL-9510 aerofoil
from primary testing.
PRESSURE
Figure 21 - S.U.A.D.5 pressure contours showing high pressure build up underneath especially in aft section of the
aerofoil.
VELOCITY
Figure 21a - The velocity contours around S.U.A.D.5

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S.U.A.D.6 - lift to drag 49.5
S.U.A.D.6 is a combination of the NYU-Grunman k1 aerofoil and the RAE 2822 aerofoil - the two
most successful aerofoils that had been tested at this stage.
S.U.A.D.6 generated a lift to drag ratio that exceeded all others. This was achieved through a design
combination of the best design features from the best two aerofoils tested. The NYU-Grunman k1
aerofoil has a larger leading edge frontal radius which is a desired design feature. The NYU-Grunman
k1 aerofoil has the flattest, smoothest and longest upper surface - this allows the air to flow on the
upper surface without boundary layer separation - this is a desired design feature.
The RAE 2822 has a harsher curvature around the aft section - this generates substantial lift. The
aerofoil does however have a very smooth surface that does not produce excessive levels of drag.
PRESSURE
Figure 22 - Shows the pressure contour map from the highest lift to drag ratio aerofoil tested in the simulation
S.U.A.D.6.

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VELOCITY
Figure 22a - The velocity contour map for the S.U.A.D.6 aerofoil shows good flow over the top surface flow that only
separates around the aft section.

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2.8.1 - Supercritical aerofoil testing analysis
Test results:
Aerofoil Lift Drag lift to drag ratio
SC-6010 0.004265900 0.000149790 28.47920422
NPL- 5910 0.00635580 0.00018516 34.32598833
DSMA-523 0.007374800 0.000205420 35.90108071
S.U.A.D. 0.0012187000000 0.0000420460000 28.98492128
S.U.A.D.2 0.006162800 0.000235330 26.18790634
S.U.A.D.3 0.00653310 0.00020891 31.27231822
NYU - Grunman K1 0.028218 0.000605 46.61974623
RAE 2822 0.024993 0.000511 48.10041088
S.U.A.D.4 0.02690400 0.00067028 40.1384496
S.U.A.D.5 0.02628900 0.00058937 44.60525646
S.U.A.D.6 0.02536300 0.00051284 49.55597067
Figure 23 - The table below shows the results generated from the aerofoil testing completed in ANSYS Fluent.
Figure 24 – Graph – showing the complete results from the aerofoil simulations shown on a bar chart.
0 10 20 30 40 50 60
SC_6010
NPL 9510
Dsma 523a
SUAD
SUAD2
SUAD3
NYU k1 grunmen
RAE 2822
SUAD4
SUAD5
SUAD 6
Complete Aerofoil Testing

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Following the testing of three aerofoils from the S.U.A.D. family, no considerable improvement in lift
to drag ratio had been achieved when compared to the commercially researched and tested
aerofoils. This led to further research of different experimental transonic aerofoil. Two aerofoils
(RAE 2822 and NYU K1 grunman) became the basis for the design of S.U.A.D.4 and S.U.A.D.5.
The S.U.A.D.4 and S.U.A.D.5 aerofoils where studied and modified with the aims of drag reduction
and stopping early shockwave formation thus improving performance. This was achieved by
including a sharper trailing edge and a smoother longer top section - this allowed for greater laminar
flow above the top edge these aerofoils where then tested. The outcome was a success and
significantly higher lift to drag results was achieved. S.U.A.D.6 provides a lift to drag of 49.5 - this is
extremely efficient and exceeds the performance of the commercial aerofoils selected for the
simulation.

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2.8.3 Decision and aerofoil design evaluation
The supercritical aerofoil design stage of this project has been a relative success. The result shown
above shows a definite improvement in lift to drag efficiency has been achieved.
From this point, aerofoil S.U.A.D.6 will be the only aerofoil section used in the design of a conceptual
aircraft. The aerofoil makes use of desired design features from other aerofoils and is the most
efficient aerofoil tested. Although this aerofoil is not unique as it uses different sections from other
designs, it does present an improvement.
A comparison of S.U.A.D.6 with the SC (2)-0610 aerofoil (the root aerofoil for the Airbus A380) shows
that the S.U.A.D. 6 provides significant increases in lift to drag. The improvement is a 75% increase in
lift to drag performance. This would deliver the benefits of reduced fuel consumption and overall
efficiency improvement.

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Chapter 3 - Advanced aircraft wing design
This chapter will detail the creation of a concept wing design for commercial aircraft using the
S.U.A.D.6 aerofoil. The aim is to improve the aerodynamic performance of commercial aircraft.
The focus of the aerodynamic improvements will be to improve operational efficiency. Gaining
efficiency improvements is a very important consideration for the aviation industry for the
protection of the environment and other reasons. Achieving improved fuel efficiency is a particular
consideration with fuel costs accounting for approximately 25% of airline operator’s costs. (Lee,
1998)
Improving efficiency is also important for the customers of airline operators – a reduction in running
should result in a reduction in the cost of tickets for passengers and costs charged by the airline
operators for cargo transport.
The lift to drag ratio of an aircraft is a measure of its aerodynamic performance and it is the
improvement in lift-to-drag that is the desired feature. A higher lift to drag ratio delivers the benefits
of improved cost and other efficiencies and improved climb rate.
The area of particular focus for this project is in improving efficiency is cruise flight. Aircraft spend
95% (Lee, 1998) of their flying time in cruise flight - this therefore provides an opportunity for the
introduction of improvement that will deliver significant benefits.
Chapter 2 - Aims and Objectives
These are to:
 Present a unique design of a wing arrangement.
 Incorporate aerodynamic improving design features based on the work completed during
this projects research
 Improve on the lift to drag ratio performance of a current commercial aircraft
 Present the analysis of aerodynamic performance of the wing arrangement using ANSYS
Fluent computation fluid dynamics software.

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3.1 - Concept commercial aircraft research
There is a great deal of research available in the area of conceptual commercial aircraft design that
focuses on aerodynamic improvement. The research undertaken for this project focused on three
aircraft designs:
1. Blended wings.
2. Morphing wing technology.
3. Non planar - Advanced closed/box wing.
Blended wings
This design blends wings into a smooth, wide, flat, tailless fuselage. This wide fuselage is often
shaped like an aerofoil and therefore produces most of the aircraft lift with the wings contributing
lift and overall balance. This configuration enables the entire aircraft to contribute to the lift with
less drag compared to the conventional cylindrical fuselage. The result of this is improved fuel
economy and aircraft range. Blended wings are often referred to as ‘flying wings’ as they are
typically designed using an aerofoil shaped body. (HUANG, 2012)
A great deal of research and development has been undertaken in the area of blended wing design
and to some they represent the future of air travel (Armstrong flight research centre (NASA), 2010).
The aircraft manufacturer Boeing has designed a test aircraft incorporating a blended wing – the X-
48. This aircraft was constructed at the Cranfield Aerospace Centre in the United Kingdom but the
design remains as an unmanned aircraft at present.
The blended wing body aircraft has a smaller frontal area than conventional aircraft design. As a
result of this there is less drag caused by the body. In addition to this there is the added benefit of
increased lift due to the design being based on an aerofoil.
The most significant benefits delivered by the blended wing design are aerodynamic improvements
and improved fuel efficiency. The design also has the advantage of improving the structural integrity
of the aircraft - this due to the integration of the wing structure with fuselage. This integration
means that the maximum wing bending moment and shear are approximately half of that for a
conventional configuration this means that structural weight saving can be achieved. (HUANG, 2012)
The blended wing body design under review features three jet engines mounted on the aircraft and
positioned, so that engine noise is shielded by the aircraft. This can significantly reduce the noise

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level experienced. However mounting engines at the rear can be dangerous as high angles of attack
can limit airflow reaching the engines and result in engine stall. Airports during take-off and
landing
Further challenges arise as the pilot, engines, flight equipment and payload must all fit all within the
depth of the wing section. A wing that is deep enough to accommodate all of these elements will
have an increased frontal area when compared to a conventional wing and fuselage - this results in
higher drag and reduces the drag advantages of the design.
Figure 25 - One example of a blended wing concept currently being analysed is the Boeing x-48
Morphing wings
Aircraft morphing wings affect the aerodynamic characteristics and abilities of aircraft through a
dramatic change in the shape of the aircraft. Several complex aircraft morphing system designs have
exist including rotating, sliding and inflating mechanisms. There are many research projects
underway as researchers have identified the design has the capability to increase versatility and
maximise aircraft efficiency for the duration of the flight. (Min, 2008)
The ‘Defence Advanced Research Projects Agency’ (DARPA)’, considers a morphing aircraft to be an
‘adaptable, time variant airframe, whose changes in geometry influence aerodynamic performance’.
Many conventional aircraft already incorporate features which significantly change the geometry of
the wings to influence there aerodynamic properties e.g. flaps and slats. However these features
cannot create a seamless aerofoil shape when extended as they operate using sliding rails and
hinges. (Center for Mechanics and Control, 2004)

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DARPA's concept of a Morphing aircraft is a multirole aircraft that:
 Changes its state/shape substantially to adapt to the mission environment;
 Provides superior system capability not possible without reconfiguration;
 Uses integrated design of materials, distributed actuators, effectors and mechanisms to
reconfigure in flight. (Center for Mechanics and Control, 2004)
Figure 26 - Shows potential design ideas for a morphing wing and the structural aerodynamic changes it would make in
flight. (Centre for Mechanics and Control, 2004)
The benefits of a morphing aircraft include the ability to change aerodynamic configuration to suit
different conditions – this would enable the aircraft to achieve high levels of efficiency in variable
conditions. The significant challenge of the design is that current material technology used in the
construction of the aircraft would mean a significant weight increases would be necessary to
produce the design. In addition, the design seems to be focused on military applications.
Commercial aircraft fly above the weather in stable conditions and changing the wing configuration
to suit conditions is much less important. In conclusion, this design does not need to be considered
in relation to meeting the aims of this project.
Advanced closed/box wing
Advanced closed/box wings are a type of non-planar wing that provide the deliver reduced induced
drag compared with traditional wings. The design also provides an increase in total lift. However, the
integration and assessment of non-planar wing concepts is complex. (I Kroo, 2005)
The term ‘closed wing’ is used to describe a number of wing designs including annular, joined and
box wings. Whilst there are no aircraft with these wing designs in commercial use, many significant
research projects have taken place, most notably the Lockheed Martin Advanced Box Wing and the
IDINTOS Project. The IDINTOS Project which a research project co-funded by the Regional

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Government of Tuscany (Italy) in 2011 as part of a project to design and manufacture an amphibious
ultra-light Prandtl Plane.
Closed wing surfaces exhibit a number of interesting structural and aerodynamic properties. A box
plane achieves the minimum possible induced drag for a given lift wingspan. A closed wing surface
has no wingtips whatsoever - this greatly reduces or eliminates wingtip drag. Such a design presents
very significant opportunity for the improvement of fuel efficiency in the airline industry.
Figure 27 - This image shows the performance of vortex drag production for a number a of different aircraft wings. The
picture shows that a box plane produces the least relative vortex drag when compared to a number of different designs.
(Frediani, 2005)

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3.2- Initial design concept
Induced drag accounts for 40% of cruise drag and up to 80% of total drag at take-off. If a wing can be
designed and manufactured that reduces these values then it will have an obvious commercial value
(Frediani, 2005). Increasing wing span achieves reduction in induced drag. There is a great deal
research being undertaken in this area, particularly in relation to the performance capabilities of
blended wing design concepts (I Kroo, 2005).
“Induced drag may be easily reduced by increasing the span of a planar wing. A 10% increase in
wing span leads to a 17% reduction in induced drag” (Frediani, 2005)
However, these blended wing concepts represent a significant change in aircraft design. Blended
wing aircraft will cost significantly more in construction and maintenance. In addition the increased
weight of the blended wing aircraft will mean needed longer runways for safe take-off and landing.
Following evaluation of the three wing designs reviewed, the design of the concept aircraft will
include an advanced box non planar wing. As discussed above, this wing design provides for
significantly reduced induced drag. In addition, including this wing into the design of the concept
aircraft does not present significant design changes or challenges – the fuselage will remain
unaltered as will the engines. Indeed there may even be the possibility of considering the inclusion
of this wing design as a ‘retrofit option’ for existing aircraft. Such an option would only be
economically viable should the aerodynamic improvements obtained deliver significant
improvement in fuel economy and reduced emissions.
Potential Design ideas
Figure 28 - Potential design ideas include the typical box wing configuration and Boeings fluid wing configuration.

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3.3 - Similar projects/inspiration
Three aircraft that use an advance box wing design have been researched to help achieve the aims
of this project. These are:
 The Lockheed Martin Advanced Box Plane
 The Prandtl Plane
 The Cranfield A9 Dragonfly
It has not been possible to obtain aerodynamic performance related results for these aircraft.
Therefore their role in the completion of this project is inspirational rather than providing an
opportunity for performance related comparative analysis with aerodynamic performance research.
Lockheed Martin advanced box wing concept aircraft
Figure 29 - The Lockheed Martin Advanced Box Plane was the main inspiration for this project – it is often presented as a
future aircraft variant. The design is a typical advance box wing.

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Prandtl Concept Aircraft – Pisa University
The Prandtl Concept Aircraft is a superjumbo aircraft design proposed by Aldo Frediani, Matteo
Gasperini, Guido Saporito and Andrea Rimondi from the Department of Aerospace Engineering
“Lucio Lazzarino”, Pisa University, Italy. (Frediani, 2005)
The design of the Prandtl Concept Aircraft has been very influential in the project decision to
attempt to reduce lift to drag via the use of an advanced box wing.
A Prandtl Plane aircraft configuration is based on the concept of ‘Best Wing Systems’. Reference is
made to a theoretical result published by Prandtl in 1924, showing that the lifting system with the
minimum induced drag, under certain conditions, is a wing box in the front view.
“In a large transport aircraft during cruise flight, drag is mainly due to friction drag (45-50%) and
induced drag (40-45%) “ (Frediani, 2005)
Figure 31 - Shows the comparison in optimum induced drag of a biplane and optimum induced drag of ‘best wing
systems’.
Figure 30 - The final design of the Prandtl concept Aircraft. This design was the produced at Pisa University. The
project had the aim of designing an aircraft that would carry over 800 passengers.

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A9 Dragonfly – Cranfield University
The A9 Dragonfly is a research project carried out at Cranfield University. The design is a typical box
wing arrangement; the projects objective was to design a medium sized long haul aircraft.
Information regarding the aerodynamic performance of this aircraft is not freely available.
Figure 32 - Shows the A9 Dragonfly - a medium size long haul aircraft design using the advanced box wing design.

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3.4 - Current aircraft lift to drag performance
As this project aims to improve the lift to drag performance of a superjumbo or wide bodied concept
aircraft, it is vital to have the lift to drag performance figures for current wide-bodied commercial
aircraft - in particular the current lift to drag performance data for a Boeing 777.
The Boeing 777 is significant to this project as the concept wing design for commercial aircraft is
being designed around the current design parameters of a Boeing 777x aircraft.
Figure 33 - provides the Lift to drag performance for the commercial aircraft shown:
Aircraft Lift to drag performance
Airbus A330 19
Airbus A340 19
Airbus A380 19
Boeing 747 17
Boeing767 18
Boeing 777 19
Boeing 787 21
(Smith, 2009)

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3.5.1 – Concept design plan and objectives
The Boeing 777x
The design concept this project is following focusses attention on the Boeing 777 aircraft and in
particular the new Boeing 777x which is still at the design stage.
The aim is to build an advanced closed wing design that offers significant increase in lift to drag in
comparison to the old Boeing 777 design.
An advanced box style wing arrangement will be designed for inclusion in the new 777-9x aircraft.
This aircraft will be the largest single floored wide bodied aircraft available. Major airline operators
including Lufthansa and Emirates have already shown a great deal of interest in operating the
aircraft. (Boeing New airplane 777x, 2014)
A 1% reduction of drag for a large transport aircraft saves 400.000 litres of fuel and, 5000 Kg of
emissions per year. In a large transport aircraft during cruise flight 90% of total drag is mainly due to
friction drag and induced drag. (Frediani, 2005)
The advanced box wing aircraft being designed will use the aircraft specifications of the Boeing 777-
300 series aircraft. The new design will share:
1. Fuselage width.
2. Fuselage length.
3. Wing sweep angle.
4. Engine placement and size.
Keeping these test parameters the same allows for an accurate comparison of test results and for
well-informed conclusions to be reached.
A selection of box wing designs will be produced and a conclusion as to the most efficient design for
the 777x can be suggested.

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3.5.2 - Angle of incidence
Chapter 1 describes how an aerofoil was optimised through a combination of features of two
experimental aerofoils. A secondary research exercise will now be completed to establish at which
angle the aerofoil performs to the highest lift-to-drag ratio. This angle is important as it will be used
in the wing design for the ‘angle of incidence’ of the wing.
The angle of incidence is the angle the aerofoil sits at whilst the aircraft is at angle of attack 0. This
means the aircraft fuselage is producing the minimum amount of drag and the wings will be
producing the maximum lift to drag that they can achieve. Determining the most appropriate angle
of incidence will deliver the maximum lift to drag for each design.
Test parameters
S.U.A.D.6 will be tested at angles between 1 and 8 degrees. This range is used as no supercritical
aerofoils maximum lift to drag angle has fallen outside this range (Airfoil investigation database,
2013). The lift to drag ratios of each angle will be logged and the most appropriate angle will be
determined relative to the aims of the project.
Test results:
S.U.A.D. 6 AOA LIFT DRAG LIFT/DRAG
1 0.00322570 0.00005638 57.21659542
1.5 0.00383260 0.00006485 59.09854898
2 0.00440100 0.00007277 60.48154358
2.5 0.00480480 0.00009554 50.29150399
3 0.0052868 0.00011114 47.5688321
4 0.0064413 0.00012963 49.6898866
5 0.0073785 0.00020101 36.707129
6 0.0082537 0.00026035 31.7023238
7 0.0093413 0.00033611 27.7923894
8 0.0099492 0.00043162 23.05083175
Figure 34 - research results - incidence of incidence.

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Angle of attack 2 proved to be the highest lift to drag
Pressure
Figure 35 - Show the pressure contours around aerofoil S.U.A.D.6 at its optimum angle of attack which was 2 degrees
where its lift-to-drag performance is over 60.
Velocity
Figure 36 - Show the velocity contours around aerofoil S.U.A.D.6 at its optimum angle of attack which was 2 degrees
where its lift-to-drag performance is over 60.

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The current Boeing 777-300 aircraft
The current Boeing 777-300 aircraft is a very successful commercial aircraft that is economical to
operate and which has an excellent safety record. Prior to testing the concept wing design for a
commercial aircraft, a scale model of the original Boeing 777-300 created using Creo parametric.
The model was tested to verify the simulations figures 38-46 are accurate. The Target lift to drag was
19, the Project Simulation results for lift to drag were 17.2. These results verify the simulation.
Figure 37 - A 777 replica designed to the same design parameters found on the Boeing technical information website.
Pressure
Figure 38 - Pressure 1 contours around the surface of a 777

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Figure 39 - Pressure 2 contours around surface of a Boeing 777
Velocity
Figure 40 - Velocity 1 around a Boeing 777
Figure 41 - Velocity 2, velocity contours around surface of a Boeing 777

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Boeing 777-300 with engines
Lift-to-drag 12.5
Figure 42 - 777-300 scale model with engines attached.
Pressure
Figure 43 - Pressure contours around a Boeing 777 with engines

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Figure 44 - Pressure contours image 2
Velocity
Figure 45 - Velocity contours around the surface of a Boeing 777 with engines
Figure 46 - Velocity contours image 2

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Boeing 777x concept 1 – typical advanced box wing design
Figure 47 - The first concept wing design for commercial aircraft is a typical box wing design this design is titles the 777x
concept 1.
Pressure
Figure 48 - Pressure contours around the surface of 777x concept 1

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Figure 49 - Pressure contours image 2
Velocity
Figure 50 - Velocity contours around the surface of 777x1
Figure 51 - Velocity contours image 2

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Boeing 777x concept 2 – project designed advanced box wing
Design 2 – lift to drag -14
Figure 52 - 777x concept 2 is an advanced box design original to the project.
Pressure
Figure 53 - Pressure Contours around 777 x concepts 2

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Figure 54 - Pressure contours image showing an underside view of the 777x concept2
Velocity
Figure 55 - Velocity contours around the 777 x concept 2
Figure 56 - Velocity contours showing the under view of the 777x concept 2

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Boeing 777x concept 3 – Project designed advanced box wing 2
Lift to drag result - 15
Figure 57 - The second designed advanced box wing original to this project titles 777x concept 3
Pressure
Figure 58 - Pressure contours around 777 x concepts 3

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Figure 59 - Pressure contours showing different view of 777x concept 3
Velocity
Figure 60 - Velocity contours of 777x concept 3
Figure 61 - Velocity contours image 2 of the 777x concept 3

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Boeing 777x concept 4 –project design advanced box wing 4
Lift to drag result -14
Figure 62 - The Final design is more of a biplane than a box plane this aircraft design is titles 777x concept 4
Pressure
Figure 63 - Pressure contours around 777 x concepts 4

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Figure 64 - Second image of pressure contours around 777 x concepts 4
Velocity
Figure 65 - Velocity contours around 777 x concepts 4
Figure 66 - velocity contours around 777 x concepts 4

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Experiment analysis
Aircraft Lift Drag Lift/Drag rounded
777-300-no engines 0.188590 0.010971 17.189864 17
777-300-with engines 0.173000 0.013932 12.417456 12
777-x1-with engines 0.771970 0.060745 12.708371 13
777-x2-with engines 1.013800 0.074001 13.699815 14
777-x3-with engines 1.140100 0.077171 14.773684 15
777-x4-Boxwing 1.270800 0.090662 14.016898 14
Figure 67 - This table of figures shows the performance of lift to drag ratio of each aircraft design tested. It is
demonstrated that each of the non-planar advanced box style wings offers significant advantages in lift to drag
performance.
Figure 68 - Lift to drag results of each aircraft shown on a line graph
0.000000
2.000000
4.000000
6.000000
8.000000
10.000000
12.000000
14.000000
16.000000
18.000000
20.000000
Lift/Drag
Lift/Drag

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Figure 69 - Similar graph to figure 68 however this time the information is represented on a bar chart.
Figure 70 - A bar chart comparing a current Boeing 777 against the best performing box wing the 777x concept 3
0.000000
2.000000
4.000000
6.000000
8.000000
10.000000
12.000000
14.000000
16.000000
18.000000
Lift/Drag
777-300-noengines
777-300-withengines
777-9x1-withengines
777-9x2-withengines
777-9x3-withengines
777-9x4-Boxwing
0.000000
2.000000
4.000000
6.000000
8.000000
10.000000
12.000000
14.000000
16.000000
777-300-withengines 777-9x3-withengines
Lift/Drag
777-300-withengines
777-9x3-withengines

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Conclusion on advanced box wing designs for commercial use
The results above show that the advanced box wings yield a 30% increase in lift to drag when
compared to a Boeing 777 replica that was subjected to the same test.
The most successful design Boeing 777-x3 delivered a lift to drag ratio of almost 15 - this significantly
higher than the Boeing 777 replica model that delivered a lift to drag ratio of 12.7.
The advanced box wing design successfully improves lift to drag ratio for a commercial aircraft. This
conclusion can be reached as the performance all four concept wing designs for commercial aircraft
was better than the performance of the Boeing 777-300 series replica model in lift to drag ratio.

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Chapter 4 – Materials selection for advanced box wing design.
This is the third and final area of wing design that will be researched. The aim is to determine the
best possible material for the construction of the concept wings having regard to a number of
considerations including performance.
The research will concentrate on current aviation materials. In addition CES Edu-pack resource
software will be used to perform a thorough material selection. The output of the chapter will be a
suggestion as to the top five materials that could be used to deliver optimum performance in the
construction of a concept wing for commercial aircraft. Reference will be made to the cost of the
recommended materials and detailed technical information will be provided.
Attention will be paid to current research into the next generation materials of materials that may
be used the aviation industry.
Given recent developments in the aviation industry, the expected conclusion of this chapter is that
composite carbon fibre materials will prove to be the most appropriate material for the construction
of a concept wing for commercial aircraft.
Chapter 3 – Aims and objectives
 Materials research focusing on current materials in use in the aviation industry.
 Materials selection using CES-Edupack to identify the optimum material for a concept wing
for commercial aircraft.
 Research into the next generation materials of materials that may be used the aviation
industry.

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4.1 Materials research
From the early days of powered flight, the materials used in the aviation industry have been in
constant development. Innovation of materials is vitally important for the aviation industry - the
material used in aircraft construction has significant effect on the performance and cost of operation
of aircraft.
In designing an aircraft wing, a very important consideration is to establish the optimal proportion of
the weight to strength/stiffness. The wing needs to be sufficiently strong and stiﬀ to withstand the
variable operating conditions in which the aircraft will be used. Durability is an important factor.
Also, should a particular part of the wing fail it must not result in the destructive failure of the whole
aircraft.
The design process starts with a speciﬁcation of the requirements and the specification of the
properties the wing will need to meet.
The design output will often be a compromise between material properties and weight. A most
important requirement of the aircraft wing is that it will perform its design function particularly in
critical situations when safety is paramount. (Aerostudents.com, 2013)
The deformation of a material at limit loads must not interfere with the safe operation of the
aircraft. Should the static strength requirement result in a component showing unacceptably high
deflections then the component is said to be ‘stiffness limited design’. (Aerostudents.com, 2013)
The material selected as a result of this investigation has to be able to support ultimate loads
without destructive failure. Further, the material must support limit loads without permanent
deformation of the structure. (Aerostudents.com, 2013)
Aluminum is the most widely used material in the aviation industry, however, should the properties
if aluminum not meets the necessary loads requirements within the size limitations of the wing
design, higher strength materials would be considered (Titanium or Steel). For the purposes of this
investigation Aluminum is too heavy to meet the performance requirements relating to increased
efficiency. Graphite/Epoxy resin based materials or Next Generation Materials will be considered.
An aircraft wing will produce lift because of the unequal pressure between its bottom and top
surfaces. This results in a shear force as well as bending moment, which are at their highest values at
the point where the wing meets the fuselage. The structure at this point needs to be very strong to

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resist the loads and moments but also quite stiﬀ to stop wing bending. The wing will need to be thick
and strong at the point where it meets the fuselage. (Aerostudents.com, 2013)
Other matters for consideration
The advantage of engines mounted under the wing is that their weight is around the area in the lift is
being produced. This reduces the total fuselage weight reducing the shear force and bending
moment that occurs between wing root and fuselage. The rudder and ailerons will also create lift
causing torsion around the fuselage. Since the fuselage is a cylindrical shape it will be able to
withstand torsion very eﬀectively. The landing gear can also generate loads causing torsion on the
fuselage. But the ultimate force caused by the landing gear is the shock produced during landing;
because of this shock absorbers are fitted that absorb the landing energy and thus reducing the
force applied to the structure. (Aerostudents.com, 2013)
(Boeing , 2012)
Figure 71 - shows the percentage of composite materials in the Boeing fleet as they have evolved the 787 is now made
up of 50% composite with the entire wing structure made using
.

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4.2 - Advantages of advancements in material technology
This table shows the advantages of advancements in material technology to 3 different stakeholders.
To the designer To the factory To the airline
Reduce weight Advanced production
techniques
Reduced fuel consumption
Fatigue and corrosion
resistance
Fewer parts Fewer and easier inspections
New design possibilities Reduced production cost Reduce maintenance cost
Increased aerodynamic
ability
Longer flight life
(Boeing , 2012)
Figure 72 - Shows the advancement in material technology and predicts future technology levels the current
technologies being used in the 777 are pre 2000 and need an urgent review

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4.3 - Material selection using CES-Edupack
Material selection will now be undertaken to identify the optimum material for the concept wings
designed in chapter 2. The material selection process is undertaken following a review of the current
materials in use in the aviation industry.
The material selection process will consider:
1. Mechanical properties
2. Thermal properties
3. Electrical magnetic and optical properties
4. Chemical properties.
The aim of this project is to improve the current performance of the Boeing 777 wing so that
recommendations can be made regarding to make recommendations for the design of a concept
wing for a commercial aircraft.
The essential requirements of an aircraft wing are:
A. High stiffness
B. High strength
C. High toughness
D. Low weight
Figure 73 - An image demonstrating the 4 main design parameters of aircraft wings.

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An aircraft wing is a stiffness-limited design at minimum mass (cost, energy, environmental impact)
and strength-limited design at minimum mass (cost, energy, environmental impact).
An aircraft wing is a BEAM (loaded in bending)
- Stiffness, length, shape specified; section area free
- Strength, length, shape specified; section area free
In the completion of the investigation the database used was level 3 aerospace from the academic
version of CES-Edu pack. This software package included details almost 4000 materials which have
applications in aviation industry - This is the most in-depth material investigation that can be
completed as part of this project given the available facilities.
The first stage of a material selection process is to perform a translation. A translation states
function, constraints, objectives and free variables of the design. Translations are always shown
before the main material selection process begins - it allows for clarity of thought regarding the
objectives and constraints.
A Translation performed for optimum material for use in an advanced box wing design for
concept wing design for commercial aircraft.
Function Optimum material for aircraft wing.
Constraints High stiffness
High toughness
High strength
Resistance to corrosion
Ease of maintenance
Objectives Minimum mass
Free variables Choice of material, choice of
manufacturing technique

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Stage two - of the material selection process is to set up the database. Selection of material
classifications to involve in the selection, the material groups that where selected for this selection
where
 Ceramics and glasses
 Fibres and particulates
 Hybrids, composites, foams, honeycombs
 Metals and alloys
 Polymers, plastics, elastomers
This totals around 4000 materials.
Stage three - of the selection process is to limit certain properties – This procedure excludes
materials from the selection process the materials that do not perform the function required.
The limited factors in this selection are as follows.
Density – Limited to a maximum of 3000
Young’s modulus (stiffness) – Minimum requirement of 10 GPA
Yield strength – Minimum requirement of 10 MPA
Tensile strength – Minimum requirement of 15 MPA
Fatigue strength at cycles – Minimum requirement of MPA
Fracture toughness – Minimum requirement of 15 MPA
Stage four in the material selection process is the screening and ranking for the correct material
using a graphs and the suitable selection line. The selection line gradient is controlled by the merit
index. (Ashby, 2009)
A beam in bending – stiffness limited design – Merit index = ρ / E1/2
A beam in bending – strength limited design – Merit index = ρ / σy
2/3
(Ashby, 2009)

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High stiffness – low density
Figure 74 - Graph showing young’s modulus against density, only the colour circles are materials that have passed the
limit stage
Figure 75 - When the selection line is used the following materials are the best 5 shown in this figure
Top 5 materials in rank order are:
1. Cyanate ester/HM carbon fiber, UD composite, 0° lamina
2. Beryllium, grade 0-50, hot isostatically pressed
3. Beryllium, grade I-220B, vacuum hot-pressed
4. Beryllium, grade I-250, hot isostatically pressed
5. Beryllium, grade I-70A, vacuum hot-pressed

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High strength – low density
Figure 76 - Elastic limit vs. density graph only the coloured circles have passed the limit stage
Figure 77- Yield strength vs. density, using a selection line to find the top 5 performing materials.
2. BMI/HS carbon fiber, UD composite, 0° lamina
3. Polyimide/HS carbon fiber, woven fabric composite, biaxial laminate
4. Beryllium (50-127 micron, f)
5. Al-48%B(f), longitudinal

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High fracture toughness – low density
Figure 78 - Fracture toughness vs. density
Figure 79 - Fracture toughness vs. density using a selection line to select the top 5 materials.
1. BMI/HS carbon fiber, UD composite, 0° lamina
2. Magnesium, commercial purity
4. Alumina silicate/Nextel 720, 45Vf - woven fabric
5. Polyimide/HS carbon fiber, woven fabric composite, biaxial laminate

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4.4 Material selection conclusion
Cyanate ester and high modulus carbon fibre composite 0° uni-directional lamina has been selected
as the material that is most suitable for use in a high performance wing arrangement for the new
777x aircraft. Cyanate ester and high modulus carbon fibre composite 0° uni-directional lamina is in
the top five materials in each of the following comparisons:
High stiffness – low density
High strength – low density
High fracture toughness – low density
Cyanate ester and high modulus carbon fibre composite 0° uni-directional lamina is already
commonly used in the aviation industry. Typical uses include high performance spacecraft/aircraft,
missiles, antenna & Randomes.

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4.5 Raw material cost comparisons
The average costs for materials that are commonly used in the aviation industry are shown below:
 Aluminum Plate = $2 - $3 / lb.
 Steel Plate = $5 - $10 / lb.
 Titanium Plate = $15 - $25 / lb.
 Fiberglass/Epoxy Prepare = $15 - $25 / lb.
 Graphite/Epoxy Prepare = $50 - $100 / lb.
(Boeing , 2012)
Aluminium is significantly cheaper than the composite materials; however this investigation focuses
on performance. The highest performing material a Cyanate ester and high modulus carbon fibre
composite 0° uni-directional lamina, will cost around $25/ lb. (Boeing , 2012)
In addition to the raw material costs, there are further costs that need to be considered these are:
1. Detail Fabrication Costs
2. Assembly Costs
3. Life Cycle Costs
4. Cost of Weight (Loss of Payload, Increased Fuel Consumption)
5. Cost of Maintenance
(Boeing , 2012)

Jonathon Rowan-Undergraduate Project

Jonathon Rowan-Undergraduate Project

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