JOHNSON_BENJAMIN_11379847_Appendix Collection

Project Report
BEng
The Conceptual Design of a Two Seater
Electrically Powered Training Aircraft
Appendices
Name: Benjamin James Johnson
Supervisor: Liz Byrne
May 2015
SCHOOL OF ENGINEERING AND TECHNOLOGY

School of Engineering and Technology BEng Final Year Project Report
BACHELOR OF ENGINEERING DEGREE WITH HONOURS IN
AEROSPACE ENGINEERING
BEng Final Year Project Report
School of Engineering and Technology
University of Hertfordshire
The Conceptual Design of a Two Seater Electrically Powered
Training Aircraft
Appendix 1
Research
Report by
Benjamin James Johnson
Supervisor
Liz Byrne
Date
20 APRIL 2015
The Conceptual Design of a Two Seater Electrically Powered Training Aircraft
i

ABSTRACT
Appendix 1 to the main report, this document details the way in which data from other aircraft
was found and analysed so that the initial design for the concept aircraft can be specified,
contained is the data sheets created for the other aircraft. This document also contains the
market research around electrical and training aircraft, the current and near-future electric
aircraft and research into the Cessna 152 aircraft.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
TABLE OF CONTENTS ................................................................................................................iii
LIST OF FIGURES........................................................................................................................iv
1 Market................................................................................................................................... 1
1.1 Global Warming............................................................................................................. 1
1.2 Energy Prices ................................................................................................................ 2
1.3 Electric Energy .............................................................................................................. 4
2 Electric Aircraft...................................................................................................................... 6
2.1 Solar Impulse 1 ............................................................................................................. 6
2.2 Sunseeker 1 .................................................................................................................. 7
2.3 Sunseeker II .................................................................................................................. 7
2.4 Sunseeker Duo.............................................................................................................. 8
2.5 E-FAN 2.0...................................................................................................................... 9
3 Future Electric Aircraft........................................................................................................ 10
3.1 Solar Impulse 2 ........................................................................................................... 10
3.2 SUNSTAR ................................................................................................................... 11
3.3 E-FAN 4.0.................................................................................................................... 12
4 Training Aircraft History...................................................................................................... 13
5 Aircraft Data Sheets ........................................................................................................... 14
6 Development of Aircraft Requirements .............................................................................. 56
7 Cessna 152......................................................................................................................... 59
7.1 Cessna Aircraft Company History............................................................................... 59
7.2 Cessna 152 Specification............................................................................................ 59
REFERENCES............................................................................................................................ 61
iii

LIST OF FIGURES
Figure 1 - Jet Fuel and Crude Oil Price - [6] ................................................................................. 3
Figure 2 - Crude Oil Worldwide Distribution - [7] .......................................................................... 4
Figure 3 - Solar Impulse 1 - [15].................................................................................................... 6
Figure 4 - Sunseeker 1 - [16] ........................................................................................................ 7
Figure 5 - Sunseeker 2 - [18] ........................................................................................................ 8
Figure 6 - Sunseeker Duo - [20].................................................................................................... 9
Figure 7 - E-FAN 2.0 - [21]............................................................................................................ 9
Figure 8 - Solar Impulse 2 - [23].................................................................................................. 10
Figure 9 - SUNSTAR - [24] ......................................................................................................... 11
Figure 10 - E-FAN 4.0 - [21]........................................................................................................ 12
Figure 11 - Cessna 152 3 View Sectional Drawing - [27] ........................................................... 60
Table 1 – American Aviation AA-1 Yankee Clipper .................................................................... 15
Table 2 – Aero Ltd. AT-3............................................................................................................. 16
Table 3 - Aeronca L-3 Grasshopper............................................................................................ 17
Table 4 – Aeronca Model 7 Champion........................................................................................ 18
Table 5 – Beechcraft Aircraft Corporation Model 77 Skipper ..................................................... 19
Table 6 – Mustang Aeronautics Bushby Mustang II ................................................................... 20
Table 7 – Cessna Aircraft Company 140 .................................................................................... 21
Table 8 - Cessna Aircraft Company 150..................................................................................... 22
Table 9 – Cessna Aircraft Company 152 .................................................................................... 23
Table 10 – Cessna Aircraft Company 162 Skycatcher ............................................................... 24
Table 11 – Czech Aircraft Works Sport Cruiser .......................................................................... 25
Table 12 – Denney Aerocraft and Kitfox Aircraft Denney Kitfox Model 2 ................................... 26
Table 13 – Diamond Aircraft DA20 ............................................................................................. 27
Table 14 – Flight Design CT ....................................................................................................... 28
Table 15 – Glasair Aviation GlaStar............................................................................................ 29
Table 16 – Grob Aircraft G115 Tutor........................................................................................... 30
Table 17 – Grob Aircraft G120 .................................................................................................... 31
Table 18 – Jeffair Barracuda....................................................................................................... 32
Table 19 – Liberty Aerospace XL2.............................................................................................. 33
Table 20 – North American Aviation T-6 Texan.......................................................................... 34
Table 21 - Piper Aircraft J-3 Cub................................................................................................. 35
Table 22 - Piper Aircraft PA-18 Super Cub................................................................................. 36
Table 23 - Piper Aircraft PA-38 Tomahawk................................................................................. 37
Table 24 - Polikarpov Po-2.......................................................................................................... 38
Table 25 - RagWing Aircraft Designs RagWing RW11 Rag-A-Bond.......................................... 39
iv

Table 26 - Rans Inc. S-19 Venterra ............................................................................................ 40
Table 27 - Slingsby Aviation T67 Firefly...................................................................................... 41
Table 28 - Stoddard-Hamilton Aircraft Glasair I .......................................................................... 42
Table 29 - Stoddard-Hamilton Aircraft Glasair II ......................................................................... 43
Table 30 - Stoddard-Hamilton Aircraft Glasair III........................................................................ 44
Table 31 - Symphony Aircraft Industries Symphony SA-160...................................................... 45
Table 32 - Eklund Engineering Thorp T-18................................................................................. 46
Table 33 - IndUS Aviation Thorp T-211 ...................................................................................... 47
Table 34 - Van's Aircraft RV-4..................................................................................................... 48
Table 39 - Van's Aircraft RV-12................................................................................................... 53
Table 40 - Vultee BT-13 Valiant.................................................................................................. 54
Table 41 - Yakovlev Yak-18 ........................................................................................................ 55
Table 42 - Excel Comparison table............................................................................................. 57
Table 43 – Cessna 152 Technical Specification - [26]................................................................ 60
v

1 Market
For the initial development of any product an investment must be made, this investment is
time and money. The return from this investment is generally money or knowledge and
therefore a market or sector must be identified in which the product will fill a niche. This target
market sets the product aside and makes it desirable therefore offering a return on the
investment, the larger the target or the more important the larger the return. Therefore the
initial stage of any development project is the identification of the market.
1.1 Global Warming
It is widely acknowledged that global warming is having a negative impact upon the planet,
the problems caused by rising sea levels and changing climate are costing organisations both
time and money. To stop these problems global warming must be reversed or at least slowed,
this can only be accomplished through massive innovation across all sectors. The most
accepted cause of global warming is the increase in greenhouse gases and the ‘greenhouse
effect’, the increase in the blanketing of the earth by gases which trap heat within the Earth’s
atmosphere which would otherwise be radiated into space. Without this effect Earth would not
be able to support life; however man’s effect upon the atmosphere has increased the amount
of greenhouse gases and caused the atmosphere to retain too much heat therefore warming
the planet. The Intergovernmental Panel on Climate Change stated that; “Continued emission
of greenhouse gases will cause further warming and long-lasting changes in all components
of the climate system, increasing the likelihood of severe, pervasive and irreversible impacts
for people and ecosystems. Limiting climate change would require substantial and sustained
reductions in greenhouse gas emissions which, together with adaptation, can limit climate
change risks.” [1] The currently recognised effects associated with climate change are;
“Glaciers have shrunk, ice on rivers and lakes is breaking up earlier, plant and animal ranges
have shifted and trees are flowering sooner…loss of sea ice, accelerated sea level rise and
longer, more intense heat waves.” [2] However, other unknown effects may be seen which
haven’t been predicted including economic and social effects.
The main gases that contribute to the greenhouse gases are; water vapour, Carbon Dioxide,
Methane, Nitrous Oxide and Chlorofluorocarbons. Each of these gases has a particular effect
upon the Earth’s atmosphere and each come from a particular source:
• Water Vapour; the most abundant greenhouse gas and increases as the Earth’s
atmosphere warms but does not actively effect global warming itself.
• Carbon Dioxide; produced by respiration, the burning of fossil fuels and certain
natural events such as volcanic eruptions is the most stable and therefore most
persistent greenhouse gas. Humans have increased the concentration of Carbon
Dioxide in the atmosphere by 33% since 1760.
The Conceptual Design of a Two Seater Electrically Powered Training Aircraft 1

• Methane; produced by human activities as well as natural sources, is a more
problematic greenhouse gas, however is in much less abundance.
• Nitrous Oxide; is produced by burning fossil fuels and using commercial and organic
fertilizers.
• Chlorofluorocarbons; are the only gas in the atmosphere that are entirely of human
creation, as well as being a greenhouse gas they destroy the ozone layer causing
more of the suns radiation to heat the atmosphere.
To combat the heating of the atmosphere and the increases in greenhouse gases much of
the research and development in industry has been aimed at reducing the use of fossil fuels.
This has either been through using renewable or sustainable energy sources, creating
recyclable products or increasing the efficiency of existing systems. For the European aviation
industry the European Commission released a report entitled; Flightpath 2050 Europe’s
Vision for Aviation, stating; “Environmental protection has been and remains a prime driver in
the development of air vehicles and new transport infrastructure. In addition to continuously
improving fuel efficiency, the continued availability of liquid fuels, their cost impact on the
aviation sector and their impacts on the environment have been addressed as part of an
overall fuel strategy for all sectors.” [3].
This report lays out the European Commission’s goals for the aviation industry in 2050: [3]
• In 2050, technologies and procedures available allow a 75% reduction in CO2
emissions per passenger kilometre to support the Air Transport Action Group (ATAG)
target (10), and a 90% reduction in nitrogen oxide (NOx) emissions. The perceived
noise emission of flying aircraft is reduced by 65%. This is relative to the capabilities
of typical new aircraft in 2000.
• Aircraft movements are emission-free when taxiing.
• Air vehicles are designed and manufactured to be recyclable.
• Europe is established as a centre of excellence on sustainable alternative fuels,
including those for aviation, based on a strong European energy policy.
• Europe is at the forefront of atmospheric research and takes the lead in the
formulation of a prioritized environmental action plan and establishment of global
environmental standards.
1.2 Energy Prices
Alongside the problems with atmospheric changes by the increase in greenhouse gases is
the problem presented by the reduction in remaining fossil fuel reserves. “There are an
estimated 1.3 trillion barrels of proven oil reserve left in the world’s major oil fields, which at
present consumption rates will be sufficient to last 40 years…it is likely by then that the
world’s population will be twice as large, more industrialization” [4], this suggests that oil
based fuels cannot be relied upon unless there is a dramatic decrease in the consumption of

oil or more oil is discovered. The reduction in oil and its impending rarity has also driven the
price of oil up, “a barrel that cost $10 in 1998 cost $64 in 2007 and today costs $135” [4] that
is an increase of 1250% in less than 15 years. This increase has massive economic impacts;
the direct impact of rising oil prices is a rise across all forms of fuel created from crude oil, in
JAN 2007 the UK’s average price for a litre of unleaded petrol was 90.8 pence in OCT 2014
this had risen to 126.7 pence [5], over the same period the price of Jet fuel rose from $50 a
barrel to $100 (Figure 1) this is an 100% increase in fuel costs for aircraft operators.
Figure 1 - Jet Fuel and Crude Oil Price - [6]
However the increased price of fuel is not the only effect, increased fuel prices increases the
cost of using machinery to harvest crops, this in turn increases the price farmers charge for
their crop and the price the final vendor charges for the product. In the aerospace industry the
increased cost of aviation fuel increases the cost of the flight, this increased cost is reflected
as an increase in ticket price, charter cost or freighter charges. These in turn can lead to
customers seeking alternate options to those given by the aerospace industry, due to the
relatively higher cost the industry becomes less popular and profits fall. Alongside services
provided by the aerospace industry its pilots must also be trained, as simulation is not
completely true to reality training and flying hours must be maintained on an airframe, this
means that pilots must regularly fly, this requires fuel and therefore if fuel costs more it
increases the cost of pilots maintaining their qualifications. The same approach applies to
training new pilots, for a Private Pilot’s License it’s expected that between 45 and 60 hours
flying is required, therefore for a Cessna 152 flying 45 hours it will use approximately 1518.75
litres of fuel, as a Cessna 152 uses MOGAS, unleaded petroleum, at the current price in fuel
alone the PPL costs £1924.26 a fuel cost increase of 10 pence increases the total PPL fuel
cost by £151.87 a 7.3% increase. The Cessna is a relatively typical training aircraft but 45
hours is the minimum time required it can typically take up to 60 hours to complete the PPL
and these costs increase relatively. These costs increase massively as the aircraft fuel
consumption increases especially with commercial pilot training and airline transport pilot

(ATP) training requiring a minimum of 1500 hours flying, in a Boeing 767 this equates to
8176500 litres of fuel used, at a cost of 41 pence per litre overall costs £3,352,365, a 10
pence increase in jet fuel would cost an extra £817,650. This assumption is not entirely valid
however; if fuel prices could be lowered or a sustainable suitable, cheaper alternative to
current fuels found, this massive cost to the aerospace industry could be lowered
substantially.
1.3 Electric Energy
A widely recognized alternative to fossil fuels is electrical energy; generated from burning
fossil fuels, nuclear fission or fusion, solar energy harvesting or chemical reaction, electrical
energy can be suited to most applications that a fossil fuel is currently the only solution.
Energy is invaluable to everyone, it is required for all of life but it can be quantified, stored and
sold, the form that it is sold in can be more or less valuable to a customer and so energy
prices are varied. This is due to the differences in energy density for different storage
methods, three of the most recognized forms of energy are Oil, Natural Gas and Coal, these
energy forms are then refined and used or transferred into a different more usable energy
form. However each of these energy forms must be mined or harvested, due to the value of
the energy being harvested these sites are often the focus of huge contest from company to
country level. As can be seen from (Figure 2) the location of oil is focused in several places,
this presents a problem for those countries that rely on oil but have either no or little oil
themselves; this problem is energy security and a lack of. Fossil fuels by their very nature are
only found in large quantities in fixed locations; however renewable energy sources tend to be
available to all countries. Electrical energy can be generated in many different ways and
therefore offers a high energy security as long as the ability to generate it is available; this
makes it a desirable form of energy as, along with its high security, it also has many uses.
Figure 2 - Crude Oil Worldwide Distribution - [7]

Electrical energy however is currently hard to store, 1 litre of unleaded petrol has
approximately 8.5 kWh of energy in it [8], an average sized car battery can store around 2
kWh. This means that to store the same amount of energy on an aircraft that fuel has using
car batteries you would need around 4.25 times the fuel capacity in batteries. A Cessna 152
has a fuel capacity of 98 litres meaning that it would require 416.5 car batteries for the same
energy, along with this batteries will only typically last 12 to 15 years unlike a fuel tank which
unless damaged will last the aircraft lifetime [9]. However an engine specific fuel consumption
of anything less than 100% will mean that an engine isn’t turning all the available energy in
the fuel into power, thus it is storing fuel that isn’t converted into propulsive force. A typical car
engine has an SFC of 30% to 40% [10] meaning that less than half of the stored energy is
transferred into power, where as an electric motor has an efficiency of around 80%-90%
meaning that the energy storage is around 4 times the size when converting to electrical
energy but the motor efficiency is double so only half the energy is required.
Most importantly however the use of electrical energy by motors produces zero tail pipe
emissions, therefore if the electricity is generated in a zero emission way the whole cycle can
have zero effect upon the atmosphere. The tail pipe emissions are not the only form of
pollution caused by a fossil fuel engine, noise has always been an issue whenever aircraft are
concerned, be it expanding airports or low flying aircraft the noise from a large or particularly
loud aircraft can cause problems. Along with the disruption the noise also represents
inefficiency, the energy used to create the noise must come from the fuel used by the engine
and thus the engine is not running at 100% efficiency. Electrical motors transfer energy in a
much more efficient manner, generally on a small motor the only sound heard is that of the
bearings on the main shaft and the machine that is attached to the motor. On larger motors
these do become more apparent along with other noises but they are still much quieter than
relative conventional fossil fuel engines. Therefore the advantages and disadvantages of
electric energy use can be summarised into this table:
Advantages Considerations
Zero emissions Battery energy density much lower than fuel
More efficient Cost
Lower noise pollution
Greater energy security

2 Electric Aircraft
Currently compared to conventional aircraft, successful electrical aircraft are few and far
between, however the concept has been explored since 1884. The La France airship was the
first aircraft to fly using an electric motor and the first fully controlled flight of any aircraft, the
flight lasted approximately 23 minutes and the aircraft flew 8 kilometres returning to the start
point it had left from. [11] The first flight of a manned electrical aeroplane was on 21 OCT
1973 with the flight of the MB-E1; it flew for 9 minutes and 5 seconds and marked the first
ever manned flight by a solely electric powered aircraft. [12] 1979 marked the first flight of a
solar powered manned aircraft, that being the flight of the Mauro Solar Riser, this flight
covered 800m at heights of around 3m. [13] The next achievement marked by an electrically
powered aircraft was that set by the NASA Environmental Research Aircraft and Sensor
Technology Program (ERAST), the Pathfinder, Pathfinder Plus, Centurion and Helios were
solar powered unmanned aircraft and through their research, development and flights set the
altitude records for solar powered, electric powered, propeller driven and FAI class U-1.d
aircraft. [14] Since these achievements and advancements in electric propulsion and storage
technologies electrical aircraft have become more abundant with several being available as
kit aircraft for private flying. Some of the most notable are mentioned below:
2.1 Solar Impulse 1
Description: “With its huge wingspan equal to that of an Airbus A340, and it’s proportionally
tiny weight – that of an average car - the HB-SIA prototype presents physical and
aerodynamic features never seen before. These place it in a yet unexplored flight envelope.”
[15]
Mission: “It was not built to fly round the world. Its purpose was rather to demonstrate the
feasibility of the program by making the first ever whole day-and-night flight without fuel” [15]
Weight: 1600 kg
Power Plant: 4 x 10hp brushless, sensor less electric engines
Energy Storage: Lithium Polymer Batteries 240Wh/kg total weight 400kg
Figure 3 - Solar Impulse 1 - [15]

2.2 Sunseeker 1
Description: Built between 1986 and 1989, Solar Flight’s first airplane, Sunseeker I, was
designed for a mission to cross America, a feat it accomplished during the summer of 1990.
The expedition began in the Southern California desert and with 21 flights ended in North
Carolina, in a field near where the Wright Brothers first flew. It was the first crossing of the
United States made by a solar-powered airplane; an affirmation of the technology's potential
and a milestone in aviation history. [16]
Mission: Fly across America
Figure 4 - Sunseeker 1 - [16]
2.3 Sunseeker II
Description: “After successfully crossing the United States in Sunseeker I, a long series of
modifications and refinements led to an almost entirely new airplane. New wings were
constructed with a different plan form, more surface area for solar cell coverage, and a new
technique for integrating the latest generation of solar cells into the actual wing structure
rather than bonding them to the surface. The new aircraft features a unique teetering
propeller, which drastically reduces vibration. In 2006, a new motor was constructed for the
airplane that is twice as powerful as Sunseeker I's motor. An improved tail was fitted to the
aircraft in addition to a new set of control electronics designed by Alan Cocconi for the
batteries and solar arrays. The new electronics greatly increase the system's efficiency. The
new aircraft is fitted with four packs of advanced lithium polymer batteries to increase power
for take-off and climb.” [17]
Mission: “Sunseeker II completed a vast flying tour of Europe. The tour began with the first
crossing of the Alps ever made by a solar powered airplane and continued down the length of
Italy to Sicily, followed by a route along the Dolomites through Austria and Slovenia, and

finally a journey through the South of France and Spain ending at Spain’s southern coast.”
[17]
Weight: 120kg
Power Plant: 5kW Electric Motor
Energy Storage: 48 x Lithium Polymer Batteries
Figure 5 - Sunseeker 2 - [18]
2.4 Sunseeker Duo
Description: “The Sunseeker Duo is the most advanced solar powered airplane in the world.
It is Solar Flight’s third solar powered airplane. It has a wingspan of 22 meters; an empty
weight of 280 kg and 1510 solar cells with 23% efficiency. The airplane is able to cruise
directly on solar power with two people on board. The structure must be incredibly light and
aerodynamically efficient to perform well with only the power from integrated solar arrays. It
uses a battery pack located in the fuselage to store energy harvested from the solar cells
which line its wings and tail surfaces. The undercarriage is retractable tricycle gear, fully
sprung, with a steerable nose wheel and ensures that the Duo will operate normally at any
airport in the world. The folding wings give the airplane a hanger footprint no larger than a
conventional light plane. If necessary, the Sunseeker Duo can also be disassembled and
packed into a trailer. First flown under power in December 2013, it has now logged several
hundred hours in the air, and carried more than a few passengers. Irena Raymond became
the second pilot of the DUO, and has made 10 solo flights in it.” [19]
Mission: First Two Seater Solar Powered Aircraft
Weight: 270kg
Power Plant: 20kW Direct Drive Motor
Energy Storage: Battery pack

Figure 6 - Sunseeker Duo - [20]
2.5 E-FAN 2.0
Description: “It is as clean as a butterfly and hums like a bee: with a 600-kilogram weight
and maximum speed of 160 km/h, E-Fan is the first aircraft with fans to have fully electric
propulsion. The plane has zero carbon dioxide emissions in flight and is significantly quieter
than a conventionally powered aircraft. Lower noise levels of electric propulsion would
potentially benefit airport operations by allowing extended flight operation times and therefore
allowing increases in air traffic.” [21]
Mission: The E-Fan, a fully electrically-powered aviation training aircraft
Weight: 600kg
Power Plant: 2x 30kW Electric Ducted fans
Energy Storage: 2x 250V Lithium Ion Polymer Batteries made by KOKAM
Figure 7 - E-FAN 2.0 - [21]

3 Future Electric Aircraft
3.1 Solar Impulse 2
Description: “Solar Impulse is the only airplane of perpetual endurance, able to fly day and
night on solar power, without a drop of fuel. After the Solar Impulse prototype’s 8 world
records, when it became the first solar airplane ever to fly through the night, between two
continents, and across the United States, it is time for Bertrand Piccard and André
Borschberg to move on to the final phase of the adventure: the 2015 round-the-world flight.
What better way to demonstrate the importance of the pioneering, innovatory spirit than by
achieving “impossible” things with renewable energy and highlighting new solutions for
environmental problems?” [22]
“The chances of succeeding at the first attempt to build a solar airplane capable of flying
around the world were judged to be slim, so a more rudimentary prototype, HB-SIA, was first
constructed. Lessons learned from this prototype are incorporated in Solar Impulse
2, the Round-The-World Solar Airplane.” [23]
Mission: Solar Impulse is the only airplane of perpetual endurance, able to fly day and night
on solar power, without a drop of fuel. [23]
Weight: 2300kg
Power Plant: 4 x motors producing 17.5 CV
Energy Storage: Lithium Batteries weighing 633kg
Figure 8 - Solar Impulse 2 - [23]

3.2 SUNSTAR
Description: “Using extensive laminar flow techniques, the SUNSTAR takes advantage of
sailplane aerodynamic design philosophy to achieve the lowest possible power requirement to
maintain flight at high altitudes. To enable solar powered flight in the widest range of
conditions, the SUNSTAR has the best coverage of solar cells ever achieved for flight times
running into months or even years. For maximum power at low sun angles some solar arrays
are mounted on the sides of the aircraft. A three motor configuration was chosen for
maximum reliability. The front mounted motors and propellers are optimized for lower
altitudes, for take-off and climb. After the SUNSTAR reaches its operational altitude, these
motors are shut down, and the propellers fold back, out of the airstream. Station holding is
done with the single pusher motor, centrally mounted with a large diameter propeller, optimal
for high altitudes. This central motor is designed for the low power cruise condition, for
minimal power consumption while on station. The SUNSTAR will be test flown initially with a
pilot on board. From the beginning, all the controls will be “fly by wire”. Optionally manned will
be the first step toward fully autonomous operation. The inclusion of a manned cockpit in the
prototype allows much more freedom in testing, considering the restrictions placed on un-
manned aircraft over populated areas. The SUNSTAR concept is a modular system which is
configurable for a variety of missions. The central pod is interchangeable and options include
a multi seat cockpit, or an un-manned instrument pod. A pressurized cockpit for the
occupants is also in the planning stage. The wingspan can be changed for different missions,
by eliminating some wing sections. Unlike some other drone projects, the SUNSTAR has
conventional landing gear, so it can use airports and taxiways normally. Prototypes of the
systems for the SUNSTAR are already flying in Solar Flight's flagship, the SUNSEEKER
DUO. Strategic partners are invited to help define mission specific optimization and bring the
project to completion.” [24]
Mission: The SUNSTAR is Solar Flight's design for the HALE mission. (High Altitude Long
Endurance) [24]
Figure 9 - SUNSTAR - [24]

3.3 E-FAN 4.0
Description: “The 2.0 version will be followed by the E-Fan 4.0, a four-seater plane targeted
for full pilot licensing and the general aviation market. A company wholly owned by Airbus
Group, named Voltair SAS, will develop, build and offer service for the two E-Fan production
versions. The final assembly facilities will be located at Bordeaux-Mérignac Airport in the
framework of French government-backed projects for the country’s future industrialisation,
called La Nouvelle France Industrielle.” [21]
Mission: 4 Seater Training Aircraft
Figure 10 - E-FAN 4.0 - [21]

4 Training Aircraft History
Ever since man first took flight it has been known the pilots need training and that an aircraft
specially designed for this purpose will allow a pilot to be trained faster and more effectively,
some recognize the first trainer aircraft as the Curtiss JN-4D Jenny produced for the US Army
in 1915 it used the modern technologies of current aircraft and based them in a robust and
easily adaptable structure, its estimated that 95% of all WW1 Allied pilots trained in a JN-4.
During WW2 and with further advances in aerodynamic understanding and technology aircraft
such as the de Havilland Tiger Moth and North American T-6 Texan emerged, both were
primary trainers showing simple but robust structures with predictable flying characteristics
and cheap maintenance. After WW2 and the invention of the jet engine and its application in
aircraft there was a split into prop and jet trainers, with primary learning staying with propeller
aircraft due to their relatively lower maintenance costs and slower, more easily controlled
flying characteristics. With the huge spending in technology and defence during the Cold War
many new ideas and innovations came to life as company budgets were near unlimited,
nearly any imaginable aircraft configuration was designed, created and tested creating a huge
array of aircraft which all required more training and research. In line with the advances in
military aviation after WW2 and still to the present civil aviation, particularly passenger flight
advanced tremendously. Older air frames and old technologies became available to the
civilian market as military organizations modernized and looked to sell older aircraft, these
aircraft were then used by entrepreneurs to advance airlines and freight businesses, as these
companies became more proliferate; aircraft manufacturers began to design aircraft
especially for them. The advances and the increased spending in the aviation industry also
lead to new methods and decreased costs in manufacturing which allowed smaller companies
with niche markets to develop, one of these was the Cessna Aircraft Company.

5 Aircraft Data Sheets
Initially research revolves around analysing current aircraft used in general aviation and
training roles, this information can then be used to make assumptions around the initial
design of the aircraft.
Primary data required includes:
• Wingspan
• Range
• Maximum Take-Off Weight
• Total Empty Weight
• Power
• Thrust to Weight Ratio
• Wing Area
• Wing Loading
The following aircraft data was taken from [25] and allows the designer to start the design
process. Pictures taken from [20] for an analysis of general layout.

Manufacturer American Aviation Name AA-1 Yankee Clipper
Year 1968
Wingspan 7.46 m Range 785 km
Maximum Take-
off Weight
680 kg Total Empty Weight 461 kg
Power 80.6 kW Thrust to Weight Ratio 0.119
Wing Area 9.11 m
2
Wing Loading 74.6 kgm
-2
Table 1 – American Aviation AA-1 Yankee Clipper

Manufacturer Aero Ltd Name AT-3
Year 1997
Maximum Take-
off Weight
Power 75 kW Thrust to Weight Ratio 0.129
Wing Area 9.30 m
2
-2
Table 2 – Aero Ltd. AT-3

Manufacturer Aeronca Name L-3 Grasshopper
Year 1941
Maximum Take-
off Weight
Wing Area 15.60 m
2
-2
Table 3 - Aeronca L-3 Grasshopper

Manufacturer Aeronca Name Model 7 Champion
Year 1944
Maximum Take-
off Weight
Wing Area 15.80 m
2
-2
Table 4 – Aeronca Model 7 Champion

Manufacturer Beechcraft Aircraft Corporation Name Model 77 Skipper
Year 1978
Maximum Take-
off Weight
Wing Area 12.10 m
2
-2
Table 5 – Beechcraft Aircraft Corporation Model 77 Skipper

Manufacturer Mustang Aeronautics Name Bushby Mustang II
Year 1966
Maximum Take-
off Weight
Wing Area 9.00 m
2
-2
Table 6 – Mustang Aeronautics Bushby Mustang II

Manufacturer Cessna Aircraft Company Name 140
Year 1946
Maximum Take-
off Weight
Wing Area 14.80 m
2
-2
Table 7 – Cessna Aircraft Company 140

Year 1957
Maximum Take-
off Weight
Wing Area 15.00 m
2
-2
Table 8 - Cessna Aircraft Company 150

Year 1977
Maximum Take-
off Weight
Wing Area 14.90 m
2
-2
Table 9 – Cessna Aircraft Company 152

Manufacturer Cessna Aircraft Company Name 162 Skycatcher
Year 2006
Maximum Take-
off Weight
598.7 kg Total Empty Weight 376.5 kg
Wing Area 11.14 m
2
-2
Table 10 – Cessna Aircraft Company 162 Skycatcher

Manufacturer Czech Aircraft Works Name SportCruiser
Year 2006
Maximum Take-
off Weight
Wing Area 13.60 m
2
-2
Table 11 – Czech Aircraft Works Sport Cruiser

Manufacturer
Denney Aerocraft
Kitfox Aircraft
Name Denney Kitfox Model 2
Year 1984
Maximum Take-
off Weight
Wing Area 12.28 m
2
-2
Table 12 – Denney Aerocraft and Kitfox Aircraft Denney Kitfox Model 2

Manufacturer Diamond Aircraft Name DA20
Year 1992
Maximum Take-
off Weight
Wing Area 11.61 m
2
-2
Table 13 – Diamond Aircraft DA20

Manufacturer Flight Design Name CT
Year 1996
Maximum Take-
off Weight
Wing Area 9.94 m
2
-2
Table 14 – Flight Design CT

Manufacturer Glasair Aviation Name GlaStar
Year 1994
Maximum Take-
off Weight
Wing Area 11.90 m
2
-2
Table 15 – Glasair Aviation GlaStar

Manufacturer Grob Aircraft Name G 115 Tutor
Year 1985
Maximum Take-
off Weight
Wing Area 12.20 m
2
-2
Table 16 – Grob Aircraft G115 Tutor

Manufacturer Grob Aircraft Name G 120
Year 1999
Maximum Take-
off Weight
Wing Area 13.29 m
2
-2
Table 17 – Grob Aircraft G120

Manufacturer Jeffair Name Barracuda
Year 1975
Maximum Take-
off Weight
Wing Area 11.15 m
2
-2
Table 18 – Jeffair Barracuda

Manufacturer Liberty Aerospace Name XL2
Year 2008
Maximum Take-
off Weight
Wing Area 10.41 m
2
-2
Table 19 – Liberty Aerospace XL2

Manufacturer North American Aviation Name T-6 Texan
Year 1935
Maximum Take-
off Weight
Wing Area 23.60 m
2
-2
Table 20 – North American Aviation T-6 Texan

Manufacturer Piper Aircraft Name J-3 Cub
Year 1938
Maximum Take-
off Weight
Wing Area 16.58 m
2
-2
Table 21 - Piper Aircraft J-3 Cub

Manufacturer Piper Aircraft Name PA-18 Super Cub
Year 1949
Maximum Take-
off Weight
Wing Area 16.58 m
2
-2
Table 22 - Piper Aircraft PA-18 Super Cub

Manufacturer Piper Aircraft Name PA-38 Tomahawk
Year 1978
Maximum Take-
off Weight
Wing Area 11.59 m
2
-2
Table 23 - Piper Aircraft PA-38 Tomahawk

Manufacturer Polikarpov Name Po-2 Kukuruznik
Year 1927
Maximum Take-
off Weight
Wing Area 33.20 m
2
-2
Table 24 - Polikarpov Po-2

Manufacturer RagWing Aircraft Designs Name RagWing RW11 Rag-A-Bond
Year 1996
Maximum Take-
off Weight
Wing Area 11.50 m
2
-2
Table 25 - RagWing Aircraft Designs RagWing RW11 Rag-A-Bond

Manufacturer Rans Inc. Name S-19 Venterra
Year 2007
Maximum Take-
off Weight
Wing Area 11.79 m
2
-2
Table 26 - Rans Inc. S-19 Venterra

Manufacturer Slingsby Aviation Name T67 Firefly
Year 1974
Maximum Take-
off Weight
Wing Area 12.60 m
2
-2
Table 27 - Slingsby Aviation T67 Firefly

Manufacturer Stoddard-Hamilton Aircraft Name Glasair I
Year 1979
Maximum Take-
off Weight
Wing Area 7.55 m
2
-2
Table 28 - Stoddard-Hamilton Aircraft Glasair I

Manufacturer Stoddard-Hamilton Aircraft Name Glasair II
Year 1989
Maximum Take-
off Weight
Wing Area 7.55 m
2
-2
Table 29 - Stoddard-Hamilton Aircraft Glasair II

Manufacturer Stoddard-Hamilton Aircraft Name Glasair III
Year 1990
Maximum Take-
off Weight
Wing Area 7.55 m
2
-2
Table 30 - Stoddard-Hamilton Aircraft Glasair III

Manufacturer Symphony Aircraft Industries Name Symphony SA-160
Year 2001
Maximum Take-
off Weight
Wing Area 11.90 m
2
-2
Table 31 - Symphony Aircraft Industries Symphony SA-160

Manufacturer Eklund Engineering Name Thorp T-18
Year 1963
Maximum Take-
off Weight
Wing Area 8.00 m
2
-2
Table 32 - Eklund Engineering Thorp T-18

Manufacturer IndUS Aviation Name Thorp T-211
Year 1945
Maximum Take-
off Weight
Wing Area 9.67 m
2
-2
Table 33 - IndUS Aviation Thorp T-211

Manufacturer Van’s Aircraft Name RV-4
Year 1979
Maximum Take-
off Weight
Wing Area 10.20 m
2
-2
Table 34 - Van's Aircraft RV-4

Year 1986
Maximum Take-
off Weight
Wing Area 10.20 m
2
-2

Year 2001
Maximum Take-
off Weight
Wing Area 11.20 m
2
-2

Year 1995
Maximum Take-
off Weight
Wing Area 10.80 m
2
-2

Year 2002
Maximum Take-
off Weight
Wing Area 11.50 m
2
-2

Year 2006
Maximum Take-
off Weight
Wing Area 11.80 m
2
-2

Manufacturer Vultee Name BT-13 Valiant
Year 1939
Maximum Take-
off Weight
Wing Area 22.20 m
2
-2
Table 40 - Vultee BT-13 Valiant

Manufacturer Yakovlev Name Yak-18
Year 1946
Maximum Take-
off Weight
Wing Area 17.80 m
2
-2
Table 41 - Yakovlev Yak-18

6 Development of Aircraft Requirements
Using this data, assumptions can be made around the wing loading, structural weight,
propulsion required and general dimensions of the aircraft. Useful information can also be
gleaned from the year of first production, as research in the aerospace industry increases so
does technical knowledge and manufacturing methods, these allow for specialised materials
or manufacturing techniques to be used and the efficiency of aircraft structures, propulsion
units and wings increased. This can mean that older aircraft have misleading or very
conservative properties which when applied to modern aircraft create an inefficient result.
Microsoft Excel is an extremely useful tool for this application, utilising the functions built into
the software, by inputting the data, the assumptions for required aircraft parameters can be
found and an initial design specification can be created. The data can be plotted to find the
correlations between aircraft maximum take-off weight and; range, thrust to weight ratio,
wingspan and wing loading. These can give the designer an insight into the initial
requirements for the aircraft design. It can also be used to identify a market niche in terms of
aircraft ability; this can be of particular interest if the aircraft being designed is a cargo or
freight aircraft for maximum take-off weight or for a passenger aircraft for increased range.
The data was thus input into an excel table, Table 42, and several graphs were created to
create an initial design specification. For this aircraft the most useful comparisons are shown
in Graph 1 and Graph 2 giving an estimated wing loading for an aircraft of this type and an
estimate of range and thrust to weight ratio.

AircraftRange
(km)
Wingspan
(m)
MaxTake-Off
Weight(kg)
TotalEmpty
Weight(kg)
Power
(kW)
ThrusttoWeight
Ratio(kW/kg)
WingArea
(m^2)
Wing
Loading
(kg/m^2)
Mass
Ratio
Aspect
Ratio
AeroAT-37177.55582350750.1299.3062.60.6016.13
AeroncaChampion74010.70533325500.09415.8033.70.6107.25
AeroncaL-335010.67572379480.08415.6036.70.6637.30
Alpha20007968.3310005751190.11913.0076.90.5755.34
BeechcraftSkipper7649.14760499860.11312.1062.80.6576.90
BushbyMustang26927.376804201200.1769.0075.60.6186.04
Cessna14072410.16658404630.09614.8044.50.6146.97
Cessna15067810.20730504750.10315.0048.70.6906.94
Cessna15276810.20757490820.10814.9050.80.6476.98
Cessna162Skycatcher8709.14598.7376.574.60.12511.1453.70.6297.50
CZAWSportCruiser10208.65600335730.12213.6044.10.5585.50
DennyKitfox12729.76544295600.11012.2844.30.5427.76
DiamondDA20101310.87750529930.12411.6164.60.70510.18
FlightDesignCT12668.50600318750.1259.9460.40.5307.27
GlasairGlaStar231510.678895441200.13511.9074.70.6129.57
GrobG115115010.009906851390.14012.2081.10.6928.20
JeffairBarracuda7247.5410436781640.15711.1593.50.6505.10
LibertyXL29268.72794526930.11710.4176.30.6627.30
PiperJ-3Cub35410.74550345480.08716.5833.20.6276.96
PiperPA-1873510.737944221120.14116.5847.90.5316.94
PiperPA-38Tomahawk86710.3675751283.50.11011.5965.30.6769.26
RagWingRW11Rag-A-Bond4518.53386191390.10111.5033.60.4956.33
RansS-19Venterra9338.53599372750.12511.7950.80.6216.17
SlingsbyT67Firefly75310.6911577941940.16812.6091.80.6869.07
Stoddard-HamiltonGlasairI18947.429986211500.1507.55132.20.6227.29
Stoddard-HamiltonGlasairII28157.109536351300.1367.55126.20.6666.68
Stoddard-HamiltonGlasairIII20927.0910897032240.2067.55144.20.6466.66
SymphonySA-16066010.769736571190.12211.9081.80.6759.73
ThorpT-188756.357254541350.1868.0090.60.6265.04
ThorpT-2117647.62575339750.1309.6759.50.5906.00
Van'sAircraftRV-128428.21600340740.12311.8050.80.5675.71
Van'sAircraftRV-411707.016804101100.16210.2066.70.6034.82
Van'sAircraftRV-611597.017264381300.17910.2071.20.6034.82
Van'sAircraftRV-712397.708155041190.14611.2072.80.6185.29
Van'sAircraftRV-815137.328165081500.18410.8075.60.6234.96
Van'sAircraftRV-911438.507944661200.15111.5069.00.5876.28
AVERAGE1029.08.88751.88470.65102.700.1311.7368.000.626.84
Table 42 - Excel Comparison table

Graph 1 - Comparison of Range Against Maximum Take-off Weight and Thrust to
Weight Ratio
Graph 2 - Comparison of Wing Loading and Maximum Take-off Weight
0.000
0.050
0.100
0.150
0.200
0.250
0
500
1000
1500
2000
2500
3000
350 450 550 650 750 850 950 1050 1150
ThrusttoWeightRatio
Range(km)
Maximum Take-Off Weight (kg)
T/W and Range against MTOW
Range (km) Thrust to Weight Ratio (kW/kg) Linear (Range (km)) Linear (Thrust to Weight Ratio (kW/kg))
0.0
20.0
40.0
60.0
80.0
100.0
120.0
140.0
160.0
350 450 550 650 750 850 950 1050 1150
WingLoading
Maximum Take-off Weight (kg)
Wing Loading (kg/m2)

7 Cessna 152
From analysis of the data found in section 5 and section 6 it can be found that the Cessna
Aircraft Company 152 is the most successful aircraft of this type, therefore it will be the
benchmark for the aircraft development. By aiming the aircraft to be a similar but improved
aircraft to the Cessna 152 it can fill the same market sector as a modern replacement.
7.1 Cessna Aircraft Company History
Opening in 1911 Cessna began building test aircraft and in 1929 certified its first aircraft, with
the certification occurring on the same day as the 1929 stock market crash the Cessna DC-6
sold less than 25 airframes and the company closed in 1932. In 1934 it reopened and began
manufacturing for the US Army in 1940, in 1956 Cessna released the Cessna 172 the most
popular aircraft in aviation history selling over 43000 airframes and still in production. The
Cessna 172 as a 4 seat aircraft was developed and in 1958 the Cessna 152 was created, a 2
seat variant of the Cessna 172 with much the same airframe, over 22500 Cessna 152 have
been manufactured. With both these aircraft being recreational aircraft and aimed solely at
the civilian market it naturally became the primary trainer of choice for many flying schools,
with many still being used by flying schools today.
7.2 Cessna 152 Specification
To benchmark the designed aircraft against the Cessna 152 the releveant benchamrk data is
required therefore the aircraft specification is needed, this is shown in Table 43 with a 3 view
sectional drawing shown in Figure 11.
The Cessna 152 is an all-metal high-wing two seat aircraft widely used as a trainer. It was
introduced in 1978 as a successor of the popular 150. The 152 strongly resembles its
predecessor but has some significant changes. The Continental 80 octane engine was
replaced by a 100 octane Lycoming O-235-L2C and the propeller was replaced by a
McCauley-design. This gives the 152 a bit more power than the 150. [26]
Cessna 152
Parameter English Metric
Dimensions
Overall Height (max) 8' 6"
Overall Length 24' 1"
Wing
Span (overall) 33' 4"
Area 159.5 sq ft
Wing Loading 10.5 lb/sq.in 51 kg/msq
Baggage Allowance 120 lbs 54kg
Capacities
Total Fuel Capacity (standard tanks) 26.0 US gal 98 liters
Fuel Capacity (standard tanks, useable) 24.5 US gal 92.3 l
Total Fuel Capacity (long range tanks) 39.0 US gal 147 l
Fuel Capacity (long range tanks, useable) 37.5 US gal 141.3 l
Oil Capacity 7 qts

Weights
Maximum Weight 1670 lbs 757 kg
Standard Empty Weight 1081 lbs 490 kg
Max. Useful Load 589 lbs 267 kg
Range
Cruise: 75% power at 8,000ft
Time (standard tanks) 3.4 hrs
Range (standard tanks) 350nm 648 km
Cruise: 75% power at 8,000ft
Time (long range tanks) 5.5 hrs
Range (long range tanks) 415nm 769 km
Service Ceiling 14,700ft 4480 m
Engine
Avco Lycoming O-235-L2C 110BHP at 2,550
Power Loading 15.2 lbs/hp 6.88 kg/hp
Propeller: Fixed Pitch, diameter 69" (max)
Take Off Performance
Ground Roll 725ft 221m
Total distance over 50' obstacle 1340ft 408m
Landing Performance
Ground Roll 475ft 145m
Total distance over 50' obstacle 1200ft 366m
Speeds
Maximum at sea level 110 kts 204 km/hr
Cruise, 75% power at 8,000ft 107 kts 198 km/hr
Climb Rate
Rate of Climb at Sea Level 715 fpm 218 m/min
Best Rate of Climb Speed 67 kts 124 kph
Stall Speed
Flaps up, power off 48 kts 89 kph
Flaps down, power off 43 kts 80 kph
Max. Demonstrated Crosswind 12 kts 22 kph
Table 43 – Cessna 152 Technical Specification - [26]
Figure 11 - Cessna 152 3 View Sectional Drawing - [27]

REFERENCES
[1] Intergovernmental Panel on Climate Change, “IPCC Fifth Assessment Synthesis Report -
Approved Summary for Policy Makers,” 2014.
[2] NASA, “NASA Global Climate Change,” [Online]. Available:
http://climate.nasa.gov/causes/. [Accessed 05 NOV 2014].
[3] European Commission, “Flightpath 2050 Europe's Vision for Aviation,” Publications Office
of the European Union, Luxembourg, 2011.
[4] Institution of Mechanical Engineers, “When will oil run out?,” [Online]. Available:
http://www.imeche.org/knowledge/themes/energy/energy-supply/fossil-energy/when-will-
oil-run-out. [Accessed 11 NOV 2014].
[5] PetrolPrices.com, “The Price of Fuel,” 2014. [Online]. Available:
http://www.petrolprices.com/the-price-of-fuel.html#j-1-1. [Accessed 11 NOV 2014].
[6] Platts, “Platts Jet Fuel,” Platts, OCT 2014. [Online]. Available:
http://www.platts.com/jetfuel. [Accessed OCT 2014].
[7] NationMaster, “Energy > Oil > Reserves: Countries Compared,” [Online]. Available:
http://www.nationmaster.com/country-info/stats/Energy/Oil/Reserves. [Accessed 09 DEC
2014].
[8] Alternative Fuels Data Center, “Alternative Fuels Data Center - Fuel Properties
Comparison,” 29 OCT 2014. [Online]. Available:
http://www.afdc.energy.gov/fuels/fuel_comparison_chart.pdf. [Accessed 20 NOV 2014].
[9] U.S Department of Energy, “Benefits and Considerations of Electricity as a Vehicle Fuel,”
[Online]. Available: http://www.afdc.energy.gov/fuels/electricity_benefits.html. [Accessed
09 DEC 2014].
[10] Libralato, “Libralato engine for hybrid vehicles,” 2013. [Online]. Available:
http://www.libralato.co.uk/technology/hybrid.html. [Accessed 20 NOV 2014].
[11] T. Sharp, “The First Powered Airship | The Greatest Moments in Flight,” Space.com, 17
JUL 2012. [Online]. Available: http://www.space.com/16623-first-powered-airship.html.
[Accessed 03 APR 2015].
[12] R. Moulton, “An electric aeroplane,” FLIGHT International, p. 946, 6 DEC 1973.
[13] A. Noth, “History of Solar flight,” Autonomous Systems Lab, Swiss Federal Institute of
Technology Zürich, Zürich, 2008.
[14] NASA, “NASA Armstrong Fact Sheet: Helios Prototype,” NASA, 28 FEB 2014. [Online].
Available: http://www.nasa.gov/centers/armstrong/news/FactSheets/FS-068-DFRC.html.
[Accessed 03 APR 2015].
[15] SolarImpulse, “Solar Impulse 1,” [Online]. Available: http://www.solarimpulse.com/en/our-

adventure/hb-sia/#.VIbmLTGsWSo. [Accessed 09 DEC 2014].
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flight.com/projects/sunseeker-i/. [Accessed 09 DEC 2014].
[17] SolarFlight, “Sunseeker II,” [Online]. Available: http://www.solar-
flight.com/projects/sunseeker-ii/. [Accessed 09 DEC 2014].
[18] Gizmag, “Sunseeker II & III on show in Paris,” 2010. [Online]. Available:
http://www.gizmag.com/sunseeker-solar-powered-aircraft-in-paris/15512/. [Accessed 09
DEC 2014].
[19] SolarFlight, “Sunseeker Duo,” [Online]. Available: http://www.solar-
flight.com/projects/sunseeker-duo/. [Accessed 09 DEC 2014].
[20] Wikipedia, “Wikipedia,” Wikipedia, [Online]. Available: en.wikipedia.org.
[21] Airbus, “The future of e-aircraft,” [Online]. Available:
http://www.airbusgroup.com/int/en/story-overview/future-of-e-aircraft.html. [Accessed 09
DEC 2014].
[22] SolarImpulse, “The First Round the World Solar Flight,” [Online]. Available:
http://www.solarimpulse.com/en/our-adventure/the-first-round-the-world-solar-
flight/#.VIbe4jGsWSo. [Accessed 09 DEC 2014].
[23] SolarImpulse, “Solar Impulse 2,” [Online]. Available: http://www.solarimpulse.com/en/our-
adventure/solar-impulse-2/#.VIblyTGsWSo. [Accessed 09 DEC 2014].
[24] Solar Flight, “Sunstar,” [Online]. Available: http://www.solar-flight.com/projects/sunstar/.
[Accessed 09 DEC 2014].
[25] Jane's Information Group, Jane's All the World's Aircraft, Jane's Information Group.
[26] D. A. Durbin, “AIRCRAFT SPECIFICATION SHEET,” [Online]. Available:
http://www.excelsiorscastle.com/dand/aviation/n89773/c152_specs.html.
[27] G. E. J. C. R. Gallery, “Cessna 152,” [Online]. Available:
http://www.generationv.co.uk/ejcgallery/displayimage.php?album=21&pid=458.
[28] International Air Transport Association, “Jet Fuel Price Development,” 2014. [Online].
Available: http://www.iata.org/publications/economics/fuel-monitor/Pages/price-
development.aspx. [Accessed 11 NOV 2014].

Training Aircraft
Appendix 2
Initial Technical Design
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 2 to the main report, this document details the way in which data from other aircraft
found in Appendix 1 can be used so that the initial design for the concept aircraft can be
specified and how the initial parameters for the aircraft are specified.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Initial Design Specification.................................................................................................... 1
Design Specification ............................................................................................................. 2
1.1.1 Purpose and Role.................................................................................................. 2
1.1.2 Dimensions............................................................................................................ 2
1.1.3 Payload.................................................................................................................. 2
1.1.4 Performance.......................................................................................................... 2
1.1.5 Handling ................................................................................................................ 2
1.1.6 Equipment ............................................................................................................. 2
1.1.7 Structural ............................................................................................................... 2
2 Matching Plot........................................................................................................................ 3
2.1 Estimations.................................................................................................................... 3
2.2 Stall Speed .................................................................................................................... 4
2.3 Max Speed .................................................................................................................... 5
2.4 Take-Off Run................................................................................................................. 5
2.5 Rate of Climb................................................................................................................. 6
2.6 Absolute Ceiling ............................................................................................................ 7
2.7 Matching Plot Analysis .................................................................................................. 8
REFERENCES............................................................................................................................ 10
iii

LIST OF FIGURES
Equation 1 - General Lift Equation - [1]......................................................................................... 4
Equation 2 - Matching Plot Stall Speed - [1] ................................................................................. 4
Equation 3 - Matching Plot Maximum Speed - [1]......................................................................... 5
Equation 4 - Matching Plot Take-Off Run - [1] .............................................................................. 6
Equation 5 - Matching Plot Rate of Climb - [1].............................................................................. 6
Equation 6 - Matching Plot Absolute Ceiling - [1].......................................................................... 7
Table 1 - Design Aims................................................................................................................... 1
Table 2 - Design Specfication ....................................................................................................... 2
Table 3 – Estimations.................................................................................................................... 3
iv

1 Initial Design Specification
From analysing the data found in Appendix 1 a selection of design aims can be chosen and a
design specification can be created, the design specficiation will drive all design decisions
and the final aircraft should fulfill all requirements layed out by it. In most cases, such as this,
the design specification can be used as a benchmark for the final aircraft, where if the aircraft
exceeds the requirements of the design specification it is more desirable. However in some
other cases, by exceeding the design specification given by a customer the aircraft may
become less desirable as it may become more costly, may fall into a category it wasn’t
intended for or may be less efficient such as carry more cargo than available.
The data given in Appendix 1 was sorted and Table 1 was created, this table lists the average
values for the data and several design aims were selected, these design aims are selected to
beat the competitor aircraft and therefore offer a more capable aircraft.
Design Aims
AverageCompetitorValues
Average2SeaterValues
DesignAimsfromAverages
DesignAimsfromGraphs
DesignAims
Range (km) 821.25 1026.95 1000 1000
Wingspan (m) 10.45 9.12 10 9
Max Take Off Weight (kg) 761.75 831.67 775 750 750
Total Empty Weight (kg) 512.25 530.74 500 450 500
Power (kW) 90.50 113.60 90 100 90
Power Loading 8.42 7.32 8.61 7.50
Wing Area (m^2) 13.82 12.72 13.5 13
Wing Loading (kg/m^2) 64.75 69.17 65 62.5 60
Table 1 - Design Aims
By using the values in Table 1 and using the Cessna 152 data in Appendix 1 a final design
specification can be created,Table 2, this design specifcation will be the minimum acceptable
specification for the final aircraft.

Design Specification
1.1.1 Purpose and Role
A 2 seater aircraft for primary flight training and air experience flying, to be used as a basic,
entry level trainer for pilots with very little to no experience up to trainee pilots taking solo
flight tests. The aircraft should also appeal to private owners for utility and personal pleasure
flying.
1.1.2 Dimensions
• Wing Span <10m
• Height <3m
• Length <8m
1.1.3 Payload
• A minimum of 2 adults with headset, parachutes and 25kg of baggage each
• A maximum take-off weight of 750kg
1.1.4 Performance
• The aircraft should be able to fly at least 6 hours
• The aircraft should be able to take off from grass strips in light rain
• The aircraft should be electrically powered with a power source that is easily
interchangeable
1.1.5 Handling
• A very predictable aircraft with stable and soft flying qualities
• Easy and natural stall recovery
• Large areas for pilot error and harmonic, gentle control movements
• Good ground handling with independent braking system
1.1.6 Equipment
• Basic Flight instrumentation, possibility for glass cockpit and yoke controls
• Excellent view forwards in flight and when taxiing
• The aircraft will have fixed undercarriage and stowing areas behind the seats
• Minimum Forward View <10m
1.1.7 Structural
• Composite construction with lightweight, modern techniques.
• Able to endure rough landings and general mishandling.
• The aircraft should protect the pilot and occupant in the event of a crash.
• Simple to repair and maintain.
Table 2 - Design Specfication

2 Matching Plot
To begin the design process a matching plot will be created, this uses a series of estimations
against the design aims to find the most critical design consideration for the aircraft, this gives
the most important reqiremnet for the aircraft and thus the wing loading and power loading so
that the design process can begin. There are several parts to the matching plot all of which
are plotted and can be analysed, these are:
• Stall Speed
• Max Speed
• Take-Off Run
• Rate of Climb
• Ceiling
To begin the creation of the matching plot each of these is calculated in line with the design
aims.
2.1 Estimations
The first stage of the matching plot creation is to estiamte the values of several parameters
which are unknown at this time, this is to account for factors surrounding aerodynamic
efficiency of the aircraft, these estimations are made using data from current aircraft and
choosing averages. From analysing several other aircraft and selecting average values the
following set of estimations are made:
CLMAX 1.6 Max Lift Coefficient
CD0 0.0386 Zero-Lift Drag Coefficient
K 0.037229226 Induced Drag Factor
CDTO 0.061002522 Drag Coefficient at Take-off Configuration
CD0TO 0.0476 Zero-Lift Drag Coefficient at Take-off
CD0LG 0.006 Landing Gear Drag Coefficient
CD0HLD_TO 0.003 High Lift Devices Drag Coefficient
CLC 0.3 Coefficient of Lift at Cruise
CLFLAPTO 0.3 Coefficient of Lift at Take-off Flap Configuration
CLTO 0.6 Coefficient of Lift at Take-off Configuration
CLR 1.32231405 Coefficient of Lift at Take-off Rotation
AR 9 Aspect Ratio
e 0.95 Oswald Efficiency
L/D MAX 15 Lift/Drag Ratio
ηT 0.5 Propeller Efficiency at Take-off
ηP 0.8 Propeller Efficiency
μ 0.3 Runway Friction Coefficient
Table 3 – Estimations

2.2 Stall Speed
The first parameter to be calculated for the matching plot is the stall speed, stall speed is the
speed in which the aircraft in clean configuration with no power stops generating enough lift to
offset the weight of the aircraft and thus starts to fall towards the ground. Therefore a slower
stall speed allows for slower flight which is considered a safer condition as it allows for the
pilot to react more easily to unfavourable conditions during flight, it also allows the aircraft to
land at lower speeds therefore reducing ground run and increasing safety.
𝐿𝐿 =
1
2
𝜌𝜌𝑉𝑉2
𝑆𝑆𝐶𝐶𝐿𝐿
Equation 1 - General Lift Equation - [1]
�
𝑊𝑊
𝑆𝑆
�
𝑉𝑉𝑆𝑆
=
1
2
𝜌𝜌𝑉𝑉𝑆𝑆
2
𝐶𝐶𝐿𝐿 𝑀𝑀𝑀𝑀𝑀𝑀
Equation 2 - Matching Plot Stall Speed - [1]
The calculation of the stall speed is done using Equation 2, this is derived from Equation 1 the
calcualtion of lift for an aircraft. By carrying out this calculation for a stall speed of 45 knots the
following results are obtained:
Graph 1 - Matching Plot Stall Speed
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot for Stall Speed
Stall Speed
Acceptable Region

2.3 Max Speed
The second parameter to be calculated is the aircraft maximum speed, the aircrafts maximum
speed is the speed at which the thrust the aircraft can create becomes equal to the drag
created by the aircraft, this drag is a combination of parasitic and induced drag calculated with
the aircraft in cruise configuration.
�
𝑊𝑊
𝑃𝑃
�
𝑉𝑉 𝑀𝑀𝑀𝑀𝑀𝑀
=
𝜂𝜂𝑝𝑝
�
0.5𝜌𝜌𝐶𝐶𝐷𝐷0 𝑉𝑉𝑀𝑀𝑀𝑀𝑀𝑀
3
(𝑊𝑊 𝑆𝑆⁄ )𝑉𝑉 𝑀𝑀𝑀𝑀𝑀𝑀
�
�
+ �
{2𝐾𝐾 𝜌𝜌𝜌𝜌⁄ }
𝑉𝑉𝑀𝑀𝑀𝑀𝑀𝑀
� �
𝑊𝑊
𝑆𝑆
�
𝑉𝑉 𝑀𝑀𝑀𝑀𝑀𝑀
Equation 3 - Matching Plot Maximum Speed - [1]
The calculation of the maximum speed is done using Equation 3, calculating the thrust and
drag produced by the aircraft. By carrying out this calculation for a stall speed of 121 knots
the following results are obtained:
Graph 2 - Matching Plot Maximum Speed
2.4 Take-Off Run
The third parameter to be calculated is the aircraft take-off run, the aircrafts take-off run is the
distance from 0 knots to airborne over a 10.7m obstacle, this is taken in take-off configuration
on the worst runway surface for friction.
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot Maximum Speed
Max Speed
Acceptable Region

�
𝑊𝑊
𝑃𝑃
�
𝑇𝑇𝑇𝑇
= �
1 − 𝑒𝑒�𝜂𝜂𝑇𝑇 𝜌𝜌𝜌𝜌𝐶𝐶 𝐷𝐷𝐷𝐷 𝑆𝑆𝑇𝑇𝑇𝑇�1
𝑊𝑊
𝑆𝑆
� ��
𝜇𝜇 − �𝜇𝜇 +
𝐶𝐶𝐷𝐷𝐷𝐷
𝐶𝐶𝐿𝐿𝐿𝐿
� 𝑒𝑒�𝜂𝜂𝑇𝑇 𝜌𝜌𝜌𝜌𝐶𝐶 𝐷𝐷𝐷𝐷 𝑆𝑆𝑇𝑇𝑇𝑇�1
𝑊𝑊
𝑆𝑆
� ��
� × �
𝜂𝜂𝑝𝑝
𝑉𝑉𝑇𝑇𝑇𝑇
�
Equation 4 - Matching Plot Take-Off Run - [1]
The calculation of the take-off run is done using Equation 4, again calculating the thrust and
drag produced by the aircraft. By carrying out this calculation for a take-off speed of 54 knots
and a take-off run of 350m the following results are obtained:
Graph 3 - Matching Plot Take-Off Run
2.5 Rate of Climb
The fourth parameter to be calculated is the aircraft rate of climb, the aircrafts rate of climb is
the rate at which the aircraft gains height in any configuration, the rate of climb is effected by
the amount of availble thrust and its’ ratio over drag.
�
𝑊𝑊
𝑃𝑃
�
𝑅𝑅𝑅𝑅𝑅𝑅
= 1
𝜂𝜂𝑝𝑝
+
�
2
𝜌𝜌�3𝐶𝐶𝐷𝐷0
𝐾𝐾
�
𝑊𝑊
𝑆𝑆
�
�
1.155
{𝐿𝐿 𝐷𝐷⁄ }𝜂𝜂𝑝𝑝
��
Equation 5 - Matching Plot Rate of Climb - [1]
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot Take-Off Run
Take-off Run
Acceptable Region

The calculation of the aircraft rate of climb is done using Equation 5, again calculating the
thrust and drag produced by the aircraft. By carrying out this calculation for a rate of climb of
300m/min the following results are obtained:
Graph 4 - Matching Plot Rate of Climb
2.6 Absolute Ceiling
The fifth and final parameter to be calculated is the aircraft absolute ceiling, the aircrafts
absolute ceiling is the maximum height at which the aircraft can maintain straight and level
flight, the aircraft ceiling is the point at which the drag of the aircraft is equal to the thrust of
the aircraft and the lift produced by the aircraft is equal to the weight.
�
𝑊𝑊
𝑝𝑝
�
𝐴𝐴𝐴𝐴
=
𝜎𝜎𝐴𝐴𝐴𝐴
�
2
𝜌𝜌𝐴𝐴𝐴𝐴�3𝐶𝐶𝐷𝐷0
𝐾𝐾
�
𝑊𝑊
𝑆𝑆
�
𝐴𝐴𝐴𝐴
�
1.155
{𝐿𝐿 𝐷𝐷⁄ }𝜂𝜂𝑝𝑝
��
Equation 6 - Matching Plot Absolute Ceiling - [1]
The calculation of the aircraft absolute ceiling is done using Equation 6Equation 5, again
calculating the thrust and drag produced by the aircraft. By carrying out this calculation for an
absolue ceiling of 7500m the following results are obtained:
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot Rate of Climb
Rate of Climb
Acceptable Region

Graph 5 - Absolute Ceiling Matching Plot
2.7 Matching Plot Analysis
From calculation of all the required parts the matching plot can be constructed and analysed.
Graph 6 - Matching Plot
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot Absolute Ceiling
Ceiling
0
0.05
0.1
0.15
0.2
0.25
0 100 200 300 400 500 600
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot
Stall Speed Max Speed Take-off Run Rate of Climb Ceiling
Acceptable Region
Acceptable Region

As can be see from the matching plot the critical condition is aircraft stall speed and aircraft
maximum speed, this is because for a prop driven aircraft all parameters must be as low as
possible.
Graph 7 - Matching Plot Interception
From the matching plot the intercept between the critical conditions is analysed giving values
for both Power and Wing loading, the intercept is chosen due to it allowing for the minimum
condition for both conditions, this is due to power loading being defined as N/W, therefore as
the weight of the aircraft is fixed and the power increased the power loading will decrease
becoming more favourable. This is also true of the wing loading, N/m
2
, as the weight is fixed
and the wing area increases the wing loading will become more favourable.
From the initial design specification the maximum take-ff weight is selected at 750kg, this
gives the aircraft a wing loading of 525N/m
2
and a power loading of 0.0625N/W, and therefore
a wing area of 14m
2
and a required power of 117kW. Using this information the technical
development of the aircraft can begin.
0.04
0.06
0.08
0.1
0.12
0.14
480 490 500 510 520 530 540 550
PowerLoading(N/W)
Wing Loading (N/m2)
Matching Plot
Stall Speed Max Speed Take-off Run Rate of Climb Ceiling

REFERENCES
[1] M. H. Sadraey, Aircraft Design: A Systems Engineering Approach, John Wiley & Sons,
2012.

Training Aircraft
Appendix 3
Concept Design and Design Development
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 3 to the main report, this document details the way in which a concept for an aircraft is
designed and developed into a concept that will be taken into the design process. This
document also details the design development process used throughout this project and the
salient points and milestones throughout.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Concept Generation ............................................................................................................. 1
2 Concept Analysis.................................................................................................................. 3
2.1 Concept 1 ...................................................................................................................... 4
2.2 Concept 6 ...................................................................................................................... 5
2.3 Concept 7 ...................................................................................................................... 6
2.4 Concept 8 ...................................................................................................................... 7
2.5 Concept 9 ...................................................................................................................... 8
2.6 Concept 10 .................................................................................................................... 9
2.7 Design Development................................................................................................... 10
3 Development Process ........................................................................................................ 11
3.1 First Estimate .............................................................................................................. 11
3.2 Fuselage Design ......................................................................................................... 12
3.3 Wing Design ................................................................................................................ 12
3.4 First Layout Sketch...................................................................................................... 12
3.5 Second Estimate ......................................................................................................... 12
3.6 Centre of Gravity Analysis........................................................................................... 12
3.7 Tail Design .................................................................................................................. 13
3.8 Second Layout Sketch ................................................................................................ 13
3.9 Third Estimate ............................................................................................................. 13
3.10 Landing Gear Design .................................................................................................. 13
3.11 Structural Design......................................................................................................... 14
3.12 Drag and Thrust Analysis............................................................................................ 14
3.13 Control Surface Design ............................................................................................... 14
3.14 Third Layout Sketch .................................................................................................... 14
3.15 Final Weight and Centre of Gravity............................................................................. 14
3.16 Final Performance Analysis......................................................................................... 15
3.17 Final Stability and Control Analysis............................................................................. 15
3.18 Final Specification ....................................................................................................... 15
3.19 Final Assembly............................................................................................................ 15
REFERENCES............................................................................................................................ 16
iii

LIST OF FIGURES
Figure 1 - Initial Concepts 1-4....................................................................................................... 1
Figure 2 - Initial Concepts 5-8....................................................................................................... 2
Figure 3 - Initial Concepts 9-12..................................................................................................... 2
Figure 4 - Design Development Process - [2] ............................................................................. 11
Table 1 – Initial Concept Analysis Scoring.................................................................................... 3
Table 2 - Concept 1....................................................................................................................... 4
Table 3 - Concept 6....................................................................................................................... 5
Table 4 - Concept 7....................................................................................................................... 6
Table 5 - Concept 8....................................................................................................................... 7
Table 6 - Concept 9....................................................................................................................... 8
Table 7 - Concept 10..................................................................................................................... 9
Table 8 - Final Concept Evaluation Scoring................................................................................ 10
iv

1 Concept Generation
To begin the design process a view for the aircraft must be created, this gives the designer a
view of the final product and can help to rectify discprepencies in the theoretical design,
therefore it is inperitive that throughout the design process the sketch or multiple sketches are
updated in line with any changes made to the design. However the designer must first
produce an initial sketch as a start point for the aircraft, this is done by creating and analysing
several different designs and choosing the most favourable. In this instance 12 initial
concepts are created and analysed.
Figure 1 - Initial Concepts 1-4

2 Concept Analysis
The initial concept generation produced 12 different concepts, these concepts are then
graded using a set of 5 questions which scored from 0 to 10. The questions used are;
• How would you rank the Aesthetic appeal of this concept?
• How simple do you believe the concept would be to design?
• How innovative do you think the concept is?
• How much research do you believe has been done on the concept?
• How easy would the interaction between trainer and trainee be?
Concept
Aesthetics
Simplicityof
Design
InnovativeIdea
Current
Researchinto
Aircraft
Configuration
Trainer-
Instructor
Interaction
Score
1 5 10 2.5 10 5 32.5
2 7.5 7.5 5 5 5 30
3 5 7.5 5 7.5 5 30
4 5 2.5 10 2.5 5 25
5 2.5 7.5 2.5 7.5 10 30
6 5 7.5 5 10 10 37.5
7 5 7.5 2.5 7.5 10 32.5
8 5 7.5 5 7.5 10 35
9 7.5 5 10 5 5 32.5
10 10 5 10 5 5 35
11 10 2.5 10 2.5 5 30
12 10 0 10 2.5 5 27.5
Table 1 – Initial Concept Analysis Scoring
The highest scoring 6 concepts are then developed and a 3 view drawing created for each,
again these concepts are scored and ranked using questions. The same 5 questions from the
first analysis are used along with 5 additional; the scoring is also changed for an overall score
out of 100.
The 5 additional questions:
• How good would the forward view from the aircraft to the ground be for both
passengers?
• How good would the view sideways towards the ground be from the aircraft in straight
and level flight for one passenger?
• How good would the view sideways towards the sky from the aircraft in straight and
level flight for one passenger?
• How much clearance behind the main gear is there for take-off and landing roll?
• How easy is it to access to the battery storage compartment?

2.1 Concept 1
Pros
Slim profile reduces drag
Relatively high tail allows good ground clearance on take-off roll
Good view forward for front pilot
Good view sideways and up for both pilots
Conventional layout
Cons
Tandem cockpit means lack of interaction between trainer and trainee
Tandem cockpit also limits rear pilots view forward
Low wing limits view to ground
Table 2 - Concept 1

2.2 Concept 6
Pros
Good view from cockpit forward, up and to each side
High wing allows for easy access to battery compartment
Relatively high tail allows for large take-off rotation with lots of clearance
Conventional layout allows for trainee to become familiar with other aircraft more easily
Cons
Short fat fore section and cockpit means a lot of drag
Wing position means tail may need to be longer than anticipated
Wing position creates structural considerations especially with interaction of cockpit canopy
CG positioning limited
Table 3 - Concept 6

2.3 Concept 7
Pros
Good view from cockpit forward and up
Conventional layout allows for trainee to become familiar with other aircraft more easily
Low wing offers relatively easy placement of wing structure
Tail can become shorter as CG can be placed in the optimum position
Cons
Battery compartment requires leaning over aircraft
Very low design means take-off roll limited
Limited view sideways to ground
Not very aesthetically appealing
Table 4 - Concept 7

2.4 Concept 8
Pros
Good view from cockpit forward, up and to each side
High wing allows for easy access to battery compartment
Cons
Short fat fore section and cockpit means a lot of drag
Wing position means tail may need to be longer than anticipated
Wing position creates structural considerations especially with interaction of cockpit canopy
CG positioning limited
Very limited take-off roll due to inverted tail
Table 5 - Concept 8

2.5 Concept 9
Pros
Aesthetically pleasing
Excellent view forwards and up
Motor placed behind pilot means in the event of failure pilots still have clear forward view
Cons
Battery compartment very hard to reach
Limited view sideways to ground
Twin boom design increases complexity and limits airflow to propeller
Table 6 - Concept 9

2.6 Concept 10
Pros
Aesthetically pleasing
Excellent view from cockpit forward, up and to each side
Relatively high tail allows for large take-off rotation with lots of clearance
Cons
Battery compartment very hard to reach
Twin boom design increases complexity and limits airflow to propeller
Table 7 - Concept 10

Concept
Aesthetics
SimplicityofDesign
InnovativeIdea
CurrentResearchinto
AircraftConfiguration
Trainer-Instructor
Interaction
ForwardViewfromAircraft
LateralViewfromAircraftto
Ground
LateralViewfromAircraftto
Sky
ClearanceonAircrafton
Take-off,LandingRoll
AccesstoBatteryStorage
Score
1 6 9 5 10 2 4 4 10 7 8 65
6 6 7 7 8 10 8 8 9 8 9 80
7 3 9 5 10 10 8 6 10 4 4 69
8 7 6 8 6 10 8 8 9 2 8 72
9 8 4 9 4 10 9 4 10 4 2 64
10 9 3 9 3 10 9 8 9 8 6 74
Table 8 - Final Concept Evaluation Scoring
After the final evaluation of the concepts Concept 6 is chosen to be taken through to the
development stage.
2.7 Design Development
From the concept design the final aircraft will be developed, Concept 6 has been chosen for
development due to its favourable characteristics, throughout the development however the
concepts will be returned to and assessed if any features on them could benefit the
development of the chosen concept. The concept drawing will be developed into a set of
sketches to roughly estimate the geometry of the fuselage and other features, this will then be
input into the development process and after several iterations the final aircraft will be
created, specified and then taken further. The development process initially involved several
estimates as shown in Appendix 2, these estimates will eventually be changed to calculate
the true values and the ranges of values the aircraft can perform under.

3 Development Process
The development process adopted is a combination taken from [1] and [2], with the theoretical
methods used taken from [1] and a basis fro the development process taken from [2]. This
development process will take the concept aircraft from initial concept to full 3D model with an
in depth analysis of critical characteristics and flying ability.
The development process is shown below:
Figure 4 - Design Development Process - [2]
3.1 First Estimate
The first estimates are the most important as they lay the foundation for the work moving
forwards, the first estimations of MTOW, Wing Area, Fuselage Drag and Cruise thrust specify
the aircraft aerofoil shape, fuselage shape and engine size. From these estimates the initial
aircraft will be specified and shaped. Most importantly for this first estimate is the weight and
First Estimate
• MTOW
• Wing Area
• Drag Estimate
• Thrust at Cruise
Fuselage Design Wing Design First Layout Sketch
Second Estimate
• Drag
• Thrust Centre of Gravity Analysis Tail Design Second Layout Sketch
Third Estimate
• Drag
• Thrust Landing Gear Design Structural Design Drag and Thrust Analysis
Control Surface Design Third Layout Sketch
Final Weight and Centre of
Gravity
Final Performance
Analysis
Final Stability and Control
Analysis
Final Specification Fianl Assembly

size of the fuselage, all other parameters are unknown apart from the weight and size of the
payload, the fuselage must be able to carry this payload so therefore must be designed with
these limits in place. From these the initial drag of the aircraft can be calculated, this drag
drives the thrust requirement, this drag and weight also drives the lift requirements of the
aircraft. From here the process of iteration begins and the dimensions can be whittled down
as more information comes apparent.
3.2 Fuselage Design
The design of the fuselage is driven by the payload the aircraft will have to carry, in this case
as it is a training and general aviation aircraft the payload is passengers and possibly light
baggage. The chosen concept also shows that the fuselage will contain the electric motor and
power source for the aircraft therefore this also must be accounted for, this stage also
includes the layout of the cockpit and instrumentation.
3.3 Wing Design
The first major decision stage is that of wing design, the weight of the aircraft has been
specified and thus the lift required can be found, also flying qualities such as stall
characteristics and stall angles can be chosen. This stage involves the selection of a wing
aerofoil, using the data belonging to this aerofoil and various wing lifting theories all
parameters for the wing can be designed.
3.4 First Layout Sketch
The first sketch of the aircraft is a milestone in the design process allowing the designer the
first glimpse of the aircraft being designed, little more than a fuselage and wing it nevertheless
allows intuitive design to take place. If features of the previously designed parts interfere or
conflict then the sketch will highlight them, allowing the designer to check the processes being
used, also the sketch can be used to verify the initial estimations.
3.5 Second Estimate
With the wing and fuselage designed the drag, and as a function of this the thrust, can be
updated and estimates made again for missing parts. During this stage any of the initial
estimates can also be updated in light of the wing a fuselage having been previously
designed.
3.6 Centre of Gravity Analysis
This stage can be argued as one of the most critical, errors in centre of gravity or weight
estimations and calculations can cause major problems in later stages and therefore it
requires checking and verifying regularly. During this stage the weights and locations of all
components ideally, but all major components are computed. Through a mixture of

estimation, analysis and research the values of mass and location in the three aircraft axis for
each component is input into a table and the centre of gravity location for the aircraft is
calculated. This stage must be referred to for all component information and marks the start of
the iteration process.
3.7 Tail Design
The third design stage is for the tail section and shares many of the same methods as that for
the wing, the relationship between the wing and centre of gravity is analysed and the effect
they have upon each other is negated by the tail, therefore any changes to the wing or centre
of gravity necessitate a change to the tail section. This stage is the first analysis of how the
aircraft will handle and behave in several flight conditions and as such gives some estimates
of the aircraft performance.
3.8 Second Layout Sketch
Much like the first layout sketch this stage allows the verification of the designers work up to
this point and can be regarded as the second major milestone. With the inclusion of the tail
section the aircraft design will become a lot easier to understand and again intuitive design
can take place. Visual analysis of the aircraft sketch can allow verification of the design
process used and can also indicate future design problems or considerations.
3.9 Third Estimate
Again this stage allows for the modification of existing estimated parameters and the inclusion
of drag for the tail section, it is also a chance to verify existing estimations to check for values
which may have been mistakenly high or low.
3.10Landing Gear Design
During this stage the specification of the landing gear takes place, analysis of the weight,
centre of gravity and design requirements allows the designer to design the landing gear for
the aircraft and find component information for the parts used in it. During this stage however
the designer must be aware of the customer and consumer for the aircraft; skill, environment
and location must be considered as each of these may have a bearing on what is one of the
first structural components. Different pilots have different abilities and such the landing gear
will have to withstand multiple loading scenarios, the environment in which the aircraft is
landing may require specialist landing gear components or design to function and different
countries have different ways of analysing and certificating aircraft which may include landing
gear configuration or strength.

3.11Structural Design
This stage is very similar to that of the landing gear design and therefore must be done in
unison with it, the structural analysis and design of the aircraft involves the designer sizing,
positioning and specifying every major structure throughout the aircraft. It also involves
designing the fixtures, mounts and fittings each of these components as to best suit their role
and design. Throughout this stage more than any other the centre of gravity for the aircraft
must be continually analysed as the major structures will form a large majority of the weight
component for the aircraft, also information gathered during the wing and tail design stages
will allow the analysis of bending and shear force diagrams for the wings and tail.
3.12Drag and Thrust Analysis
This stage involves the analysis of the drag and therefore thrust values for the aircraft,
allowing the selection of a propulsion method and appropriate sizing of it. During this stage
the final major centre of gravity changing parts are analysed, specified and input into the
design and such marks the end of the iteration process involving the centre of gravity
specifically. It also marks the final use of estimations in the design process as all thrust and
drag calculations are completed.
3.13Control Surface Design
This stage involves specification of all major control surfaces, like the tail design stage it is
closely linked to the centre of gravity value, however it does not have a great deal of effect
upon the centre of gravity itself as generally they are light relative to the other aircraft
components. Again an acute knowledge by the designer of the customer and consumer is
required due to the regulation surrounding aircraft controllability.
3.14Third Layout Sketch
The third layout sketch stage marks the culmination of all the design work and is the third
major milestone in the design process. Like the other two sketch stages this too is a chance
for intuitive design and a verification of the values used up to this stage. The final layout
sketch however shows an overall view of the entire aircraft and therefore should be studied
much more intently by the designer as it will show much more detail and therefore much more
opportunity for error.
3.15Final Weight and Centre of Gravity
This stage begins the final stages of the design process; it involves an analysis of every
weight for every component input into a table to give the final values for the aircraft centre of
gravity and centre of gravity variation. This stage also allows the designer to specify fully the
aircraft weights such as maximum take-off, manufacturers empty, fuel empty to name a few.

3.16Final Performance Analysis
This stage like the previous is a coming together of all aircraft parameters, it also fully
specifies the aircraft performance across flight phases and conditions, producing items such
as payload range graphs and specifying aircraft max speed, cruise speed, endurance and
range values.
3.17Final Stability and Control Analysis
This stage again is an analysis of all other stages, it produces items such as gust and
manoeuvre envelopes, it also can be used to analyse SPPO, phugoid, dutch roll, roll
convergence and spiral mode values to check for stability throughout all flight phases and
conditions.
3.18Final Specification
This stage involves the gathering of all data, all graphs and all values for the entire aircraft
and presenting them to the designer. This stage marks the end of any iteration processes and
therefore the designer must refer to the initial design specification and ensure that the aircraft
meets this fully, if not the offending area must be examined and the iteration process begun.
3.19Final Assembly
The final assembly marks the last major milestone the designer will be encountering using
this design process. The final assembly shows the entire aircraft and all components giving a
like for like representation of what the aircraft would look like in the real world.
This design develpoment process will be used to develop the concept previously shown into a
full specification and final assembly; however it does not mark the end of the development
process. Further analysis into the structure, aerodynamic properties and flying qualities using
CFD, FEA and other simulations will be used to fully understand and improve the aircraft
characteristics which may not have shown during the design process. After this manufacturing
limitations will have to be assessed and finally several prototype aircraft would have to be
built and tested to verify the entire process. Even after the aircraft is manufactured however,
advances in technology and manufacturing may allow further development of the aircraft
technologies, and different but similar requirements may encourage development of different
aircraft variations upon the same initial design.

REFERENCES
2012.
[2] D. Stinton, The Design of the Airplane, Reston: American Institute of Aeronautics and
Astronautics, 2001.

Training Aircraft
Appendix 4
Aerofoil and Wing Design
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 4 to the main report, this document details the way in which an aerofoil was chosen
and how the wing was designed for the concept aircraft, including all previous versions of the
wing.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Wing Aerofoil Selection ........................................................................................................ 1
1.1 Aircraft Flight Profile...................................................................................................... 1
1.2 Lift Coefficient Requirements ........................................................................................ 2
1.2.1 Wing Aerofoil Cruise Lift Coefficient, 𝑪𝑪𝑪𝑪𝑪𝑪31T .............................................................. 2
1.2.2 Wing Aerofoil Gross Maximum Lift Coefficient, 𝑪𝑪𝑪𝑪 𝑴𝑴𝑴𝑴𝑴𝑴 𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮31T .......................... 3
1.2.3 Wing Aerofoil Net Maximum Lift Coefficient, 𝑪𝑪𝑪𝑪 𝑴𝑴𝑴𝑴𝑴𝑴31T ........................................... 3
1.3 Aerofoil Selection .......................................................................................................... 3
2 Wing Design ......................................................................................................................... 9
2.1.1 Taper Ratio.......................................................................................................... 11
2.1.2 Twist .................................................................................................................... 11
2.1.3 Resulting Wing .................................................................................................... 12
3 High Lift Device Design ...................................................................................................... 15
4 Wing Technical Specification.............................................................................................. 17
5 Iterations............................................................................................................................. 19
REFERENCES............................................................................................................................ 24
iii

LIST OF FIGURES
Figure 1 - Aircraft Flight Profiles.................................................................................................... 1
Figure 2 - Aerofoil Profile for NACA 652-415 ................................................................................ 8
Figure 3 - 3D Aerofoil Profiles ..................................................................................................... 14
Equation 1 - Wing Area ................................................................................................................. 2
Equation 2 – General Lift Equation - [1]........................................................................................ 2
Equation 3 - Cruise Lift Coefficient - [1] ........................................................................................ 2
Equation 4 - Wing Aerofoil Gross Maximum Lift Coefficient - [1].................................................. 3
Equation 5 - Aspect Ratio - [1] ...................................................................................................... 9
Table 1 - Initial Aerofoil Selection – Data from [2]......................................................................... 4
Table 2 - Final Aerofoil Selection – Data from [2] ......................................................................... 5
Table 3 - Aerofoil Coordinates for NACA 652-415 ........................................................................ 8
Table 4 - Wing Dimensions ......................................................................................................... 13
Table 5 - 3D Aerofoil Coordinates............................................................................................... 14
Table 6 - High Lift Device Dimensions........................................................................................ 16
Table 7 - Wing Technical Specification....................................................................................... 18
Table 8 - Wing Iterations ............................................................................................................. 23
Table 9 - High Lift Devices Iterations .......................................................................................... 23
Code 1 - Wing Lift Distribution - [1] Modified by Benjamin James Johnson ............................... 10
Code 2 - Wing Lift Distribution Inputs.......................................................................................... 10
Code 3 - Final Wing Inputs.......................................................................................................... 12
Code 4 - High Lift Device Lift Coefficient - [1] Modified by Benjamin James Johnson ............... 15
Code 5 - High Lift Device Lift Coefficient Inputs ......................................................................... 15
iv

1 Wing Aerofoil Selection
To begin the technical design of the aircraft the main lifting surface or wing must be designed,
the wing is made from an aerofoil cross section or multiple aerofoils and may have a twist,
camber, sweep and tapor, each effecting the way it generates lift across its span. From
Appendix 2 only one parameter for the wing is known and this is the wing loading, a measure
of how much force is upon each unit area of the wing.
1.1 Aircraft Flight Profile
To begin the wing design process the aircraft flight profile must be analysed, the flight profile
is a plotted flight for the aircraft giving the altitude and range or endurance of single flight, in
the design process the flight profile is an idealised flight of the aircraft to allow for design
decisions to be made such as cruise altitude, cruise speed, range, endurance and climb
rates.
Figure 1 - Aircraft Flight Profiles
The aircraft flight profiles are created and shown in Figure 1, it is then clear that the aircraft
will cruise at a height of 4500m for approximately 7 hours with reserve fuel left. Therefore the
aircraft wing must be able to produce lift at an altitude of 4500m, therefore the requirement for
wing lift can be analysed and the wing can be designed.
-500
500
1500
2500
3500
4500
5500
0 1 2 3 4 5 6 7 8 9
Altitude(m)
Time (hours)
Aircraft Flight Profiles
Cruise No Reserve Cruise With Reserve 30 min Training Flights
2 hr Training Flights Aerobatics

1.2 Lift Coefficient Requirements
The wing design requires an aerofoil or several to create the wing, as the aircraft flight profile
is now available the lift coefficients required of the wing can be found and a suitable aerofoil
can be designed or selected. Initially for this process 3 parameters are required;
• Ideal wing aerofoil cruise lift coefficient, 𝐶𝐶𝑙𝑙𝑙𝑙, the lift coefficient required of the aerofoil
to maintain straight and steady level flight.
• Wing aerofoil gross maximum lift coefficient, 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺, the lift coefficient required of
the aerofoil at take-off with flaps.
• Wing aerofoil net maximum lift coefficient, 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀, the lift coefficient required of the
wing aerofoil at take-off without flaps.
With these three parameters calculated an aerofoil can be selected from those already
designed or a completely new aerofoil can be designed, due to the low cost market that this
aircraft is targetting an existing aerofoil will be selected as this reduces development costs.
1.2.1 Wing Aerofoil Cruise Lift Coefficient, 𝑪𝑪𝒍𝒍𝒍𝒍
The calculation of 𝐶𝐶𝑙𝑙𝑙𝑙 requires three already chosen parameters, maximum take-off weight,
wing loading and aircraft cruise speed, these two can be input into the general lift equation
and 𝐶𝐶𝑙𝑙𝑙𝑙 can be calculated.
𝑆𝑆 = 𝑊𝑊 �
𝑊𝑊
𝑆𝑆
�
−1
Equation 1 - Wing Area
𝐿𝐿 =
1
2
𝜌𝜌𝑉𝑉2
Equation 2 – General Lift Equation - [1]
𝐶𝐶𝐿𝐿𝐿𝐿 =
2𝑊𝑊𝐶𝐶
𝜌𝜌𝑉𝑉𝐶𝐶
2
𝑆𝑆
Equation 3 - Cruise Lift Coefficient - [1]
For the designed aircraft the cruise speed is selected as 110 knots, from the wing loading
requirement the wing area is calculated using Equation 1, 𝑊𝑊𝐶𝐶 or the aircraft cruise weight
must also be known, for a conventional aircraft the weight will change as fuel is used
therefore for some aircraft this change in weight will drastically change the amount of lift
required of the wing compared to the maximum take-off weight, however for this aircraft as an
battery power source is being used which does not greatly change weight during flight the
maximum take-off weight is used.
Also to be considered is that the aircrafts wing is not 100% efficient, the fuselage and other
parts of the aircraft effect the airflow over the wing and thus the aerofoil does not produce all
the lfit it is capable of, therefore the wing lift is taken as 85.5% thus this must be factored into
the ideal cruise lift coefficient calculation. Therefore using the aircraft crusie speed of 110

knots, mass of 750kg, wing area of 14m
2
and a cruise altitude of 4500m it is found that the 𝐶𝐶𝑙𝑙𝑙𝑙
required is approximately 0.5.
1.2.2 Wing Aerofoil Gross Maximum Lift Coefficient, 𝑪𝑪𝒍𝒍 𝑴𝑴𝑴𝑴𝑴𝑴 𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮𝑮
The calculation of 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 again requires parameters layen out in section 1.2.2, however
for this calculation the stall speed is used instead of crusie speed, this gives the worst flying
condition required of the wing and thus the greatest amount of lift it must produce with flaps.
As stated in section 1.2.2 the wing however is not 100% efficient and thus another 85.5% is
factored into the caculation of 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺. Using a stall speed of 45 knots, mass of 750kg,
wing area of 14m
2
and altitiude of 0m it is found that the 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 required is approximately
1.87.
1.2.3 Wing Aerofoil Net Maximum Lift Coefficient, 𝑪𝑪𝒍𝒍 𝑴𝑴𝑴𝑴𝑴𝑴
The calculation of 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 is the calcualtion of the maximum lift coefficient of the wing without
the effect of flaps, this is calcualted by analysing the lift coefficient of similar aircraft with flaps
and substituting this from the 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 in accordance with Equation 4.
𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 = 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 + ∆𝐶𝐶𝑙𝑙 𝐻𝐻𝐻𝐻𝐻𝐻
Equation 4 - Wing Aerofoil Gross Maximum Lift Coefficient - [1]
A general aviation aircraft of this weight generally has a ∆𝐶𝐶𝑙𝑙 𝐻𝐻𝐻𝐻𝐻𝐻 of around 0.7 and therefore
the aircraft 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 is approximately 1.17.
With this calculation complete all required lift coefficients have been found for the aircraft and
thus an aerofoil can be selected. For benefits in manufacturing and development the wing will
consist of a single aerofoil profile across its length therefore reducing development time and
costs and reducing manufacturing complexity, time and cost.
1.3 Aerofoil Selection
When selecting the aerofoil there are several parameter that must be considered;
• Lift coefficients, 𝐶𝐶𝑙𝑙𝑙𝑙, 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀 𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺𝐺 and 𝐶𝐶𝑙𝑙 𝑀𝑀𝑀𝑀𝑀𝑀, all of which are calculated.
• Drag coefficient, 𝐶𝐶𝑑𝑑 𝑚𝑚𝑚𝑚 𝑚𝑚, the miniumum drag condition of the aerofoil at the ideal lift
coefficient, this must be as small as possible to reduce the amount of drag produced
by the wing at cruise.
• Pitching moment coefficient, 𝐶𝐶𝑚𝑚0, the pitching moment of the aerofoil at 0° alpha, this
must be as small as possible to reduce the pitching moment produced by the wing at
cruise and thus reduce horizontal stabiliser size.
• Stall angle, ∝𝑆𝑆, the stall angle of the aerofoil at both 0° and 60° flap extension, this
must be as high as possible therefore allowing lift at higher angles of attack and
increasing flight safety.

• Stall quality, the qualities of the aerofoil after the stall, due to the requirement for the
aircraft to be a docile primary trainer and general privation aviation aircraft the stall
quality of the aerofoil must be moderate to soft to reduce the danger of the stall upon
the aircraft flight.
As stated already the aerofoil will be selected from those already designed, these are
available in several texts such as [2], the available aerofoils can then be placed into a table,
Table 1, and analysed for their suitability.
Profile Cdmin Cm0 αS Flaps 0° αS Flaps 60° Cli ClMAX Cl MAX GROSS Stall Quality
64(1)-212 0.0045 -0.025 15 11 0.4 1.55 2.4 Moderate
64(2)-415 0.005 -0.07 14 12 0.7 1.45 2.65 Moderate
65(2)-415 0.005 -0.06 16 11 0.7 1.45 2.6 Soft
64(1)-412 0.005 -0.075 14 12 0.6 1.55 2.5 Moderate
66(3)-418 0.005 -0.07 18 9 0.5 1.4 2.6 Moderate
747A415 0.005 -0.02 16 12 0.4 1.2 2.55 Soft
65(3)-618 0.0055 -0.1 18 10 0.7 1.4 2.6 Soft
63(2)-615 0.0055 -0.11 13 12 0.6 1.45 2.8 Moderate
63(1)-412 0.0055 -0.075 14 11 0.6 1.55 2.5 Moderate
63(2)-415 0.0055 -0.07 14 12 0.6 1.5 2.65 Moderate
65(3)-418 0.0055 -0.06 16 11 0.6 1.35 2.7 Soft
2410 0.0055 -0.05 15 11 0.5 1.7 2.5 Moderate
1410 0.0055 -0.02 14 11 0.2 1.5 2.3 Moderate
63(3)-618 0.006 -0.1 12 12 0.7 1.4 2.85 Soft
64(3)-618 0.006 -0.08 16 11 0.7 1.35 2.75 Soft
64(3)-418 0.006 -0.06 16 12 0.7 1.35 2.8 Soft
63(3)-418 0.006 -0.07 13 13 0.4 1.4 2.8 Soft
1412 0.006 -0.025 15 11 0.4 1.6 2.5 Moderate
0012 0.006 0 16 10 0.2 1.5 2.4 Sharp
64(4)-421 0.0065 -0.07 18 10 0.2 1.35 2.75 Soft
4412 0.007 -0.09 13 11 0.7 1.7 2.7 Soft
4415 0.0075 -0.09 12 12 0.7 1.4 2.7 Soft
4418 0.0075 -0.08 14 9 0.6 1.4 2.65 Soft
4421 0.0085 -0.08 14 9 0.3 1.3 2.7 Soft
Table 1 - Initial Aerofoil Selection – Data from [2]

From an initial selection of aerofoils which are sorted in accordance to the aerofoil parameters
the best 5 are selected and placed into Table 2.
Profile Cdmin Cm0
αS Flaps
0°
αS Flaps
60° Cli
ClMA
X
ClMAX
GROSS
Stall
Quality
64(1)-
212
0.004
5
-
0.025 15 11
0.
4 1.55 2.4 Moderate
64(2)-
415 0.005 -0.07 14 12
0.
7 1.45 2.65 Moderate
65(2)-
415 0.005 -0.06 16 11
0.
7 1.45 2.6 Soft
64(1)-
412 0.005
-
0.075 14 12
0.
6 1.55 2.5 Moderate
66(3)-
418 0.005 -0.07 18 9
0.
5 1.4 2.6 Moderate
Table 2 - Final Aerofoil Selection – Data from [2]
Again the aerofoils are sorted and it is found that NACA Profile 652-415 is the most suitable
due to its appropriate lift coefficients, low drag coefficients, low pitching moment, high stall
angles and soft stall qualities. The aerofoil graphs shown in Graph 1, Graph 2, Graph 3 and
Graph 4.
Graph 1 - Coefficient of Drag against Coefficient of Lift for NACA 652-415
0.00000
0.00500
0.01000
0.01500
0.02000
0.02500
0.03000
0.03500
-1 -0.5 0 0.5 1 1.5
Cd
Cl

Graph 2 - Pitching Moment Coefficient against Coefficient of Lift for NACA 652-415
Graph 3 - Coefficient of Lift against Angle of Attack for NACA 652-415
-0.200
-0.180
-0.160
-0.140
-0.120
-0.100
-0.080
-0.060
-0.040
-0.020
0.000
-0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1 1.2 1.4
Cm
Cl
-1
-0.5
0
0.5
1
1.5
2
2.5
3
-10 -5 0 5 10 15 20 25 30
Cl
Alpha (°)
Cl FLAPS 60° Cl FLAPS 0°

Graph 4 - Lift/Drag Ratio against Angle of Attack for NACA 652-415
Along with the aerodynamics qualities of the aerofoil the physical properties must also be
found, this is to ensure that the aerofoil coordinates and thickness are available when
modelling and analysing other aerodynamic properties.
Upper Surface Lower Surface Thickness
0 0 0 0 0
0.313 1.208 0.687 -1.008 2.216
0.542 1.48 0.958 -1.2 2.68
1.016 1.9 1.484 -1.472 3.372
2.231 2.68 2.769 -1.936 4.616
4.697 3.863 5.303 -2.599 6.462
7.184 4.794 7.816 -3.098 7.892
9.682 5.578 10.318 -3.51 9.088
14.679 6.842 15.303 -4.15 10.992
19.726 7.809 20.274 -4.625 12.434
24.764 8.55 25.236 -4.97 13.52
29.807 9.093 30.193 -5.205 14.298
34.854 9.455 35.146 -5.335 14.79
39.903 9.639 40.097 -5.335 14.974
44.953 9.617 45.047 -5.237 14.854
50.000 9.374 50 -4.962 14.336
0.00
20.00
40.00
60.00
80.00
100.00
120.00
140.00
160.00
-4.00 -2.00 0.00 2.00 4.00 6.00 8.00 10.00 12.00
Cl/Cd
Alpha (°)

55.043 8.91 54.957 -4.53 13.44
60.079 8.26 59.921 -3.976 12.236
65.106 7.462 64.894 -3.342 10.804
70.124 6.542 69.876 -2.654 9.196
75.131 5.532 74.869 -1.952 7.484
80.126 4.447 79.874 -1.263 5.71
85.109 3.32 84.891 -0.628 3.948
90.080 2.157 89.92 -0.107 2.264
95.040 1.058 94.96 0.206 0.852
100.000 0 100 0 0
Table 3 - Aerofoil Coordinates for NACA 652-415
Figure 2 - Aerofoil Profile for NACA 652-415
With the selection of the aerofoil complete the design of the wing can be completed.
-50
-40
-30
-20
-10
0
10
20
30
40
50
0 20 40 60 80 100
65(2)-415
Calibration
Upper Surface
Lower Surface
Aerodynamic Centre

2 Wing Design
With the selection of an aerofoil the wing can be designed, to begin an aspect ratio and
setting angle must be found for the wing. The setting angle, 𝑖𝑖𝑤𝑤, is set at the angle for the ideal
cruise lift coefficient, this is to ensure that during the cruise the fuselage is at 0° and the wing
is still creating the required lift. The aspect ratio however must be either selected through an
iterative process of wing design or as in this case is given by the design specification,
Appendix 2. From the aerofoil data in section 1 the wing setting angle is selected to be 4° and
from the design specification the wing span is selected to be 9m, this gives and aspect ratio of
5.7857 from Equation 5.
𝐴𝐴𝐴𝐴 =
𝑏𝑏2
𝑆𝑆
Equation 5 - Aspect Ratio - [1]
To analyse the 3D properties of the wing Pradtl lifting line theory is used in MatLab, Pradtl’s
lifting line theory is generally accurate and offers an excellent insight into how a lifting surface
will perform for a given set of parameters. The base wing is then turned into several variables
and an iterative process can be started to maximise the efficiency of the wing and make sure
its suitable for its intended application.
N = 9; % (number of segments - 1)
b = sqrt(AR*S); % wing span (m)
MAC = S/b; % Mean Aerodynamic Chord (m)
Croot = (1.5*(1+lambda)*MAC)/(1+lambda+lambda^2); % root chord (m)
theta = pi/(2*N):pi/(2*N):pi/2;
alpha = i_w+alpha_twist:-alpha_twist/(N-1):i_w;
% segment's angle of attack
z = (b/2)*cos(theta);
c = Croot * (1 - (1-lambda)*cos(theta)); % Mean Aerodynamics Chord at each segment (m)
mu = c * a_2d / (4 * b);
LHS = mu .* (alpha-alpha_0)/57.3; % Left Hand Side
% Solving N equations to find coefficients A(i):
for i=1:N
for j=1:N
B(i,j) = sin((2*j-1) * theta(i)) * (1 + (mu(i) * (2*j-1)) / sin(theta(i)));
end
end
A=Btranspose(LHS);
for i = 1:N
sum1(i) = 0;
sum2(i) = 0;
for j = 1 : N
sum1(i) = sum1(i) + (2*j-1) * A(j)*sin((2*j-1)*theta(i));
sum2(i) = sum2(i) + A(j)*sin((2*j-1)*theta(i));
end
end
CL = 4*b*sum2 ./ c;
CL1=[0 CL(1) CL(2) CL(3) CL(4) CL(5) CL(6) CL(7) CL(8) CL(9)]
y_s=[b/2 z(1) z(2) z(3) z(4) z(5) z(6) z(7) z(8) z(9)]
plot(y_s,CL1,'-o')
grid

CL_wing = pi * AR * A(1)
Code 1 - Wing Lift Distribution - [1] Modified by Benjamin James Johnson
clc
clear
S = 14 ;
AR = 5.785714286 ;
lambda = 1.000001 ;
alpha_twist = -0.000001 ;
i_w = 4 ;
a_2d = 6.332274577 ;
alpha_0 = -2.5 ;
Wing_Lift_Distribution
Code 2 - Wing Lift Distribution Inputs
Graph 5 - Base Wing Lift Distribution – CL=0.5121
As can be seen from Graph 5 the lift distribution across the wing is non-elliptical, this has
several non-desirable consequences but most importantly for this aircraft the non-elliptical
distribution will promote tip stall, this condition is when the tip of the wing stalls at the same
time as, or before, the root of the wing. This causes a loss of roll control and makes recovery
from the stall more difficult, in a training aircraft this condition is entirely undesirable and
therefore must be designed out. There are several ways this condition can be designed out,
these include the introduction of taper, twist, sweep and a change in aspect ratio, as the
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

aspect ratio is fixed and sweep is uneccesary due to the sweep being more important in
transsonic and supersonic aircraft the change in twist and taper must be analysed.
2.1.1 Taper Ratio
Graph 6 - Wing Lift Distribution with Taper
of 1 – CL=0.5121
Graph 7 - Wing Lift Distribution with Taper
of 0 – CL=0.4086
As the taper ratio increases the lift generated at the tip of the aerofoil increases, however so
does the lift across the entire surface, it can be seen that the rectangular wing has a good lift
distribution where as a wing with a taper ratio of 0 has a very undesirable wing lift distribution
for a training aircraft.
2.1.2 Twist
Graph 8 - Wing Lift Distribution with
Twist of 0° – CL=0.5121
Graph 9 - Wing Lift Distribution with Twist
of -5° – CL=0.3693
As the twist of the wing increases the lift generated at the tip of the aerofoil decreases,
however so does the lift across the entire surface, it can be seen that as the wing increases
twist the lift distribution becomes more elliptical and thus more suitable, however this is at the
expense of lift.
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

2.1.3 Resulting Wing
Through an iterative process, comprising many wing configurations the most suitable
configuration is selected, this wing offers a good compromise between the parameters whilst
maintaining its necessary requirements. The final wing is described in Code 3 and Graph 10.
clc
clear
S = 14 ;
AR = 5.785714286 ;
lambda = 0.850001 ;
alpha_twist = -2.000001 ;
i_w = 4 ;
a_2d = 6.332274577 ;
alpha_0 = -2.5 ;
Wing_Lift_Distribution
Code 3 - Final Wing Inputs
Graph 10 - Final Wing Lift Distribution – CL=0.4793
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

Graph 11 - Wing Lift Distribution Comparison
This wing when compared to the initial design has a much more suitable lift distribution and
also has a overall lift coefficient closer to the ideal lift coefficient for the wing, from Code 1 the
dimensions of the wing can also be found.
MAC 1.5556 m
CROOT 1.7838 m
b 9 m
CTIP 1.51623 m
Table 4 - Wing Dimensions
These dimensions can then be used to specify the size of the root and tip aerofoils shown in
Table 5 and Figure 3.
Root Aerofoil Tip Aerofoil
Upper Surface Lower Surface Upper Surface Lower Surface
0.000 0.000 0.000 0.000 0.000 0.000 0.000 0.000
5.583 21.548 12.255 -17.981 4.746 18.316 10.417 -15.284
9.668 26.400 17.089 -21.406 8.218 22.440 14.525 -18.195
18.123 33.892 26.472 -26.258 15.405 28.808 22.501 -22.319
39.797 47.806 49.393 -34.534 33.827 40.635 41.984 -29.354
83.785 68.908 94.595 -46.361 71.217 58.572 80.406 -39.407
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 1 2 3 4 5
CL
y/S
3D Wing Lift Distribution
Modified Wing Base Wing

128.148 85.515 139.422 -55.262 108.926 72.688 118.509 -46.973
172.708 99.500 184.052 -62.611 146.801 84.575 156.445 -53.220
261.844 122.048 272.975 -74.028 222.567 103.740 232.029 -62.924
351.872 139.297 361.648 -82.501 299.092 118.402 307.400 -70.126
441.740 152.515 450.160 -88.655 375.479 129.638 382.636 -75.357
531.697 162.201 538.583 -92.847 451.943 137.871 457.795 -78.920
621.726 168.658 626.934 -95.166 528.467 143.360 532.894 -80.891
711.790 171.940 715.250 -95.166 605.021 146.149 607.963 -80.891
801.872 171.548 803.548 -93.418 681.591 145.816 683.016 -79.405
891.900 167.213 891.900 -88.512 758.115 142.131 758.115 -75.235
981.857 158.937 980.323 -80.806 834.578 135.096 833.275 -68.685
1071.689 147.342 1068.871 -70.924 910.936 125.241 908.540 -60.285
1161.361 133.107 1157.579 -59.615 987.157 113.141 983.942 -50.672
1250.872 116.696 1246.448 -47.342 1063.241 99.192 1059.481 -40.241
1340.187 98.680 1335.513 -34.820 1139.159 83.878 1135.186 -29.597
1429.288 79.326 1424.792 -22.529 1214.894 67.427 1211.074 -19.150
1518.174 59.222 1514.286 -11.202 1290.448 50.339 1287.143 -9.522
1606.847 38.477 1603.993 -1.909 1365.820 32.705 1363.394 -1.622
1695.324 18.873 1693.896 3.675 1441.025 16.042 1439.812 3.123
1783.800 0.000 1783.800 0.000 1516.230 0.000 1516.230 0.000
Table 5 - 3D Aerofoil Coordinates
Figure 3 - 3D Aerofoil Profiles
-1000.000
-800.000
-600.000
-400.000
-200.000
0.000
200.000
400.000
600.000
800.000
1000.000
0.000 500.000 1000.000 1500.000 2000.000
[mm]
Configuration
Upper Surface
Lower Surface
Upper Surface
Lower Surface

3 High Lift Device Design
With the completition of the wing design and its optimisation for cruise the ability for the
aircraft to take off must be analysed, again using Equation 2 the lift coefficient at take-off
speed can be calculated. Again Wing Lifting Line theory and MatLab is utilised with a variation
in variables and the high lift devices are designed through an iterative process. For this
aircraft only flaps will be employed due to the complexity and unecessary features associated
with slats.
N = 9; % (number of segments-1)
S = 14; % m^2
AR = 5.785714286; % Aspect ratio
lambda = 0.85; % Taper ratio
alpha_twist = -2; % Twist angle (deg)
a_2d = 6.332274577; % lift curve slope (1/rad)
b = sqrt(AR*S); % wing span
MAC = S/b; % Mean Aerodynamic Chord
Croot = (1.5*(1+lambda)*MAC)/(1+lambda+lambda^2); % root chord
alpha=i_w+alpha_twist:-alpha_twist/(N-1):i_w; % segment's angle of attack
for i=1:N
if (i/N)>(1-bf_b)
alpha_0(i)=a_0_fd; %flap down zero lift AOA
else
alpha_0(i)=a_0; %flap up zero lift AOA
end
end
c = Croot * (1 - (1-lambda)*cos(theta)); % MAC at each segment
mu = c * a_2d / (4 * b);
for i=1:N
for j=1:N
B(i,j) = sin((2*j-1) * theta(i)) * (1 + (mu(i) * (2*j-1)) / sin(theta(i)));
end
end
A=Btranspose(LHS);
for i = 1:N
sum1(i) = 0;
sum2(i) = 0;
for j = 1 : N
sum1(i) = sum1(i) + (2*j-1) * A(j) *sin((2*j-1)*theta(i));
end
end
CL_TO = pi * AR * A(1)
Code 4 - High Lift Device Lift Coefficient - [1] Modified by Benjamin James Johnson
clc
clear
i_w = 10 ;
a_0 = -2.3 ;
a_0_fd = -4.8 ;
bf_b= 0.3 ;
WLD_HLD
Code 5 - High Lift Device Lift Coefficient Inputs

Through the iterative process the high lift devices are found to be;
bf/b 35 % HLD Span to Wing Span
cf/c 20 % HLD Chord to Wing Chord
αTO WING 10 ° Wing Angle of Attack at Take-off
δf TO 15 ° HLD Deflection at Take-off
α0FLAP -3.45 ° Zero-Lift Angle of Wing with Flaps
Down
CL WING TO 1.1408 Wing Lift Coefficeint at Take-off
αTO FUSELAGE 6 ° Fuselage Angle of Attack at Take-off
bf 3.15 m HLD Span
cf 0.31112 m HLD Chord
Table 6 - High Lift Device Dimensions
With the design of the wing and high lift devices complete the first stage of the technical
design is complete, this allows the designer to continue to design the fuselage and analyse
the drag of the aircraft.

4 Wing Technical Specification
Wing Design
S 14 m
2 Wing Area
AR 5.785714286 Aspect Ratio
λ 0.85 ° Wing Taper Ratio
ct 1.51623 m Wing Tip Chord
cr 1.7838 m Wing Root Chord
c 1.5556 m Wing Mean Aerodynamic Chord
b 9 m Wingspan
t/c 0.15 Thickness to Chord Ratio
αt -2 ° Wing Twist Angle
Λ 0 ° Wing Sweep Angle
Γ 0.00 ° Wing Dihedral
iw 4 ° Wing Setting Angle
iwi 4 ° Ideal Wing Incidence
CLα 4.696 Wing Lift Curve Slope
CL CRUISE 0.4793 Cruise Lift Coefficient
Aerofoil Design
Design Parameters
CLC 0.400633864 Ideal Lift Coefficient
CLCW 0.421719857 Wing Cruise Lift Coefficient
CLi 0.468577619 Ideal Wing Aerofoil Cruise Lift Coefficient
CLMAX 1.601287062 Aircraft Maximum Lift Coefficient
ClMAX 1.172850365 Wing Aerofoil Net Maximum Lift Coefficient
ClMAX GROSS 1.872850365 Wing Aerofoil Gross Maximum Lift Coefficient
CLMAX W 1.685565328 Wing Maximum Lift Coefficient
Aerofoil Parameters
Profile 65(2)-415 Wing Aerofoil Profile
Cdmin 0.005 Minimum Drag Coefficient
Cl/Cd MAX 140.00 Wing Aerofoil Maximum Lift to Drag Ratio
Cl0 0.275 Wing Aerofoil Lift Coefficient at Zero Angle of Attack
Cli 0.7 Wing Aerofoil Ideal Lift Coefficient
ClMAX 1.45 Wing Aerofoil Maximum Lift Coefficeint

ClMAX GROSS 2.6 Wing Aerofoil Net Maximum Lift Coefficeint
Clα 6.332274577 1/rad Wing Aerofoil Maximum Lift to Drag Ratio
Cm0 -0.06 Wing Pitching Moment Coefficient at Aerodynamic Centre
LE Radius 1.505 Wing Aerofoil Leading Edge Radius
Stall Quality Soft Wing Aerofoil Stall Qualities
α0 -2.5 ° Wing Aerofoil Zero Lift Angle of Attack
αli 4 ° Wing Aerofoil Angle of Attack for Ideal Lift Coefficient
αS Flaps 0° 16 ° Stall Angle at 0° Flap Deflection
αS Flaps 60° 11 ° Stall Angle at 60° Flap Deflection
HLD Design
bf/b 35 % HLD Span to Wing Span
cf/c 20 % HLD Chord to Wing Chord
αTO WING 10 ° Wing Angle of Attack at Take-off
δf TO 15 ° HLD Deflection at Take-off
α0FLAP -3.45 ° Zero-Lift Angle of Wing with Flaps Down
cf 0.31112 m HLD Chord
Table 7 - Wing Technical Specification

5 Iterations
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0.9
1
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0.8
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

Table 8 - Wing Iterations
clc
clear
i_w = 10 ;
a_0 = -2.3 ;
a_0_fd = -4.8 ;
bf_b= 0.3 ;
WLD_HLD
CL WING 1.0172
clc
clear
i_w = 10 ;
a_0 = -2.76 ;
a_0_fd = -5.26 ;
bf_b= 0.3 ;
WLD_HLD
CL WING 1.0537
clc
clear
i_w = 10 ;
a_0 = -3.45 ;
a_0_fd = -5.95 ;
bf_b= 0.3 ;
WLD_HLD
CL WING 1.1085
clc
clear
i_w = 10 ;
a_0 = -3.45 ;
a_0_fd = -5.95 ;
bf_b= 0.35 ;
WLD_HLD
CL WING 1.1408
Table 9 - High Lift Devices Iterations
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7
0 0.5 1 1.5 2 2.5 3 3.5 4 4.5 5
0
0.1
0.2
0.3
0.4
0.5
0.6
0.7

REFERENCES
2012.
[2] I. H. A. a. A. E. V. Doenhoff, Theory of Wing Sections Including a Summary of Aerofoil
Data, New York: Dover Publications Inc, 1959.

Training Aircraft
Appendix 5
Fuselage Design and Drag Analysis
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 5 to the main report, this document details the way in which the fuselage is designed
and how the drag for the aircraft is analysed including the wing data from Appendix 4, the
stabiliser data from Appendix 8 and the undercarriage data from Appendix 7.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Fuselage Design................................................................................................................... 1
2 Drag Analysis........................................................................................................................ 2
2.1 Parasitic Drag................................................................................................................ 2
2.1.1 Skin Friction Drag Calculation............................................................................... 3
2.1.2 Pressure Drag Calculation .................................................................................... 4
2.2 Induced Drag................................................................................................................. 6
2.3 Total Aircraft Drag ......................................................................................................... 7
2.3.1 Total Drag at Cruise .............................................................................................. 7
2.3.2 Total Drag at Take-Off........................................................................................... 9
2.4 Minimum Drag Condition............................................................................................. 10
REFERENCES............................................................................................................................ 11
iii

LIST OF FIGURES
Figure 1 - Cockit Elevation Sketch................................................................................................ 1
Figure 2 - Induced Drag - [4] ......................................................................................................... 6
Equation 1 - General Drag Equation - [2]...................................................................................... 3
Equation 2 - Reynolds Number - [1].............................................................................................. 3
Equation 3 - Skin Friction Coefficient - [3]..................................................................................... 3
Equation 4 - Form Factor - [3] ....................................................................................................... 4
Equation 5 - Aerodynamic Surface Profile Drag Coefficient - [3].................................................. 4
Equation 6 - Non-Aerodynamic Surface Profile Drag - [3] ............................................................ 5
Equation 7 - Induced Drag - [4]..................................................................................................... 7
Equation 8 - Drag Coefficient - [4]................................................................................................. 7
Table 1 - Reynolds Number at Cruise Calculations ...................................................................... 3
Table 2 - Parasitic Drag Coefficient at Cruise............................................................................... 5
Table 3 - Parasitic Drag Coefficient at Take-off............................................................................ 6
Table 4 - Induced Drag Coefficient at Cruise................................................................................ 7
Table 5 - Induced Drag Coefficient at Take-Off ............................................................................ 7
Table 6 – Total Aircraft Drag at Cruise.......................................................................................... 8
Table 7 - Total Aircraft Drag at Take-Off....................................................................................... 9
iv

1 Fuselage Design
The fuselage design centres around the design specification and the drag of the aircraft, it
encompasses the design of all major fuselage components including the cockpit layout,
engine compartment layout, landing gear and wing box layout and any required compartment
or cargo space required. Like all other process involved in the design development the
fuselage will be subject to iterations to maintain the required specifications and reduce drag
for the aircraft. Initially for the fuselage design the most important requirements must be
analysed; in this case, for a two seater training aircraft and using the design specification in
Appendix 1 the most important requirements are:
• Two seats Side by Side
• Storage for Baggage
• Storage for Removable Fuel Source
• Good Fore and Lateral View
From the concept analysis in Appendix 3 there are several more requirements:
• High Wing
• Tricycle Undercarriage
• Fore Mounted Motor
From these requirements the most important and largest is the cockpit section and thus it
begins the design process, using a modelling tool such as Dassault Systems CATIA software
the aircraft is 3D modelled however this will be discussed in Appendix 9, the initial fuselage
design is done in a manner such that changes can be quickly and easily made. Initially 2
elevation sketches are done so that the cockpit can be sized around the occupants thus
reducing size and drag.
Figure 1 - Cockpit Elevation Sketch

2 Drag Analysis
Using data taken from Appendix 4, 7 and 8 the drag analysis can begin, the drag upon an
aircraft is the force exerted by the air the aircraft is travelling through due to the mass
component of air. However due to the density of air changing with altitude drag forces
decrease as aircraft gain altitude, along with decreased drag however the less dense air
causes decreased lift therefore limiting the height aircraft can fly and the drag reduction they
can exploit. Along with the physical mass effect of air against the motion of the aircraft,
parasitic drag, is the induced drag created by the wing lift, these both will be discussed in
detail through the following sections. For a full analysis of the drag upon the aircraft the
theoretical methods can be used however they rely on assumptions and thus are not 100%
accurate, experimental methods can be used to analyse the aircraft drag further and are more
successful however they are generally costly and time inefficient thus for this analysis only
theoretical methods will be used.
2.1 Parasitic Drag
Aircraft parasitic drag is the resistance to the aircraft movement caused by all components of
the aircraft and their contact with the air, parasitic drag comes in several forms and can
account for most of the drag generated by a light general aviation aircraft such as the aircraft
being designed, also due to its mechanical nature is constantly varying with changes in air
density, speed, area and Reynolds number. The forms of parasitic drag are;
• Profile drag comprised of:
o Pressure Drag, the effect of the pressure field within the boundary layer of air
around the component.
o Skin Friction Drag, the mechanical effect of the air particles against the
surfaces of the aircraft within the boundary layer.
• Interference Drag, the effect of the interaction between the boundary layers and
pressure distributions between components of an aircraft that are in close proximity to
one another.
• Cooling Drag, the effect of ducting air through heat exchangers and cooling
components and the pressure drop associated.
• Wave Drag, the effect of shock waves associated with supersonic and hypersonic air
flow.
For the calculation of the parasitic drag of the aircraft, cooling and interference drag will be
assumed as negligible, this is due to the complex nature of their calculation and due to the
contribution of other drag forms being much greater, to account for this the drag will be
assumed as low and thus power plant selection in Appendix 6 will reflect this. Along with
these two wave drag can also be omitted, this is due to the effects of wave drag only being

apparent at transonic and supersonic speeds therefore it is not applicable to the designed
aircraft.
2.1.1 Skin Friction Drag Calculation
Profile drag comprises of both pressure and skin friction drag across the aircraft, however for
skin friction drag comprises a large proportion of it and thus pressure drag can be considered
by applying a factor to skin friction drag. The calculation of profile drag begins with the
calculation of skin friction, to makes sure the power plant selected is powerful enough the
drag is assumed at the worst case therefore the air is assumed as fully turbulent even though
it will be a mixture of laminar and turbulent flow. Initially the calculation of the Reynolds
number for each body is required, the Reynolds number is the, “ratio of inertial forces to
viscous forces and describes the degree of laminar or turbulent flow”. [1] This is required so
that the airflow can be analysed for each component.
𝐷𝐷 =
1
2
𝜌𝜌𝑉𝑉2
𝑆𝑆𝐶𝐶𝐷𝐷
Equation 1 - General Drag Equation - [2]
𝑅𝑅𝑒𝑒 =
𝜌𝜌𝜌𝜌𝜌𝜌
𝜇𝜇
Equation 2 - Reynolds Number - [1]
𝐶𝐶𝑓𝑓 =
0.455
(log 𝑅𝑅𝑒𝑒 𝑥𝑥)2.58
Equation 3 - Skin Friction Coefficient - [3]
By using Equation 2 and data from the relevant appendices the Reynolds number and skin
friction coefficient for each component at cruise is calculated, Table 1.
Component Length Symbol Width Symbol Wetted
Area
Symbol Re Cf
Fuselage 7.50 xf 1.35 yf 10.12 Sf WET 20933625 0.002675738
Wing 1.55 c 9 b 28 SWET 4340339 0.003445369
Nose Gear 0.7 HNG 0.1 DNG 0.07 SNG WET 1953097 0.00395698
Main Gear 1.1 HMG 0.1 DMG 0.11 SMG WET 3069152 0.0036554
Horizontal
Stabiliser
0.699 ch 3.4228 bh 4.786 Sh WET 1950865 0.003957786
Vertical
Stabiliser
0.659 cv 1.4 bv 2.8 Sv WET 1839535 0.003999514
Table 1 - Reynolds Number at Cruise Calculations

The skin friction coefficient for each component can be input into Equation 1 and the skin
friction drag can be calculated for each component, with the calculation of skin friction drag
complete the pressure drag for each component calculation, this must be split into two parts
segregating the aircraft components between aerodynamic surfaces and non-aerodynamic
surfaces.
2.1.2 Pressure Drag Calculation
The pressure drag calculation for the aerodynamic surfaces begins with the calculation of the
correction factor for the skin friction drag, it must be noted that for an aerodynamic surface as
the thickness increases a larger pressure gradient is generated at the rear of the aerofoil this
causes an increase in the boundary layer thickness and increases the pressure drag.
𝐾𝐾 = 1 +
𝑇𝑇
𝑐𝑐
�(2 − 𝑀𝑀0
2) cos Λ𝑐𝑐 4⁄ � + 100 �
𝑇𝑇
𝑐𝑐
�
4
Equation 4 - Form Factor - [3]
𝐶𝐶𝐷𝐷0 = 𝐾𝐾𝐶𝐶𝐹𝐹
𝑆𝑆𝑤𝑤𝑤𝑤𝑤𝑤
𝑆𝑆𝑟𝑟𝑟𝑟𝑟𝑟
Equation 5 - Aerodynamic Surface Profile Drag Coefficient - [3]
The correction factor is then applied to Equation 5 and the profile drag coefficient for the
aerodynamic surface is calculated. For calculation of the pressure drag of the non-
aerodynamic surfaces such as the fuselage, undercarriage and nacelles a fineness ratio is
employed, this is the ratio of length to maximum thickness and the reason that most
commercial passenger aircraft are long thin tubes due to a shorter fatter body producing a
higher fineness ratio and therefore more pressure drag. The value for this correction factor is
selected of a graph relating fineness ratio to correction factor shown in Graph 1.

Graph 1 - Non-Aerodynamic Correction Factor - [3]
𝐶𝐶𝐷𝐷0 = 𝐾𝐾𝐶𝐶𝐹𝐹
Equation 6 - Non-Aerodynamic Surface Profile Drag - [3]
This correction factor is then applied to Equation 6 and the profile drag coefficient for each
non aerodynamic component at cruise is calculated, inputting the relevant data, the correction
factor and the profile drag coefficient is calculated and input into Table 2.
Component Cf Fineness Ratio K CD
Fuselage 0.002675738 5.55757037 1.25 0.003345
Wing 0.003445369 1.313125 0.009048
Nose Gear 0.00395698 7 1.6 0.006331
Main Gear 0.0036554 11 1.0575 0.003866
Horizontal Stabiliser 0.003957786 0.891161 0.007054
Vertical Stabiliser 0.003999514 1.101599 0.013365
Table 2 - Parasitic Drag Coefficient at Cruise
Using this data the parasitic drag for each component can be calculated and each
components contribution to parasitic drag can be calculated, however as mentioned before
this drag is directly related to the speed and altitude of the aircraft and therefore this
calculation must be undertaken for the most extreme or most informative flight conditions,
with this in mind the calculation is undertaken again for the aircraft at take-off, Table 3.

Component Cf Fineness Ratio K CD
Fuselage 0.002839448 5.55757037 1.25 0.003549
Wing 0.003678876 1.313125 0.009662
Nose Gear 0.004240671 7 1.6 0.006785
Main Gear 0.003909163 11 1.0575 0.004134
Vertical Stabiliser 0.004287504 1.101599 0.014328
Table 3 - Parasitic Drag Coefficient at Take-off
With this calculation complete the parasitic drag for the aircraft has been calculated, therefore
the induced drag must be calculated to for a full drag analysis of the aircraft.
2.2 Induced Drag
Induced drag is the drag caused as a result of the aerodynamic lift created by the wing and
the vortex systems behind the aircraft that this creates, as shown in Figure 2 the effect of the
wing upon the airflow causes it to be pushed in a slight downwards direction, this causes the
lift to be produced at an angle behind perpendicular to the aerofoil and thus a drag
component is introduced into the lift production.
Figure 2 - Induced Drag - [4]
This induced drag factor increases and decreases with the amount of lift created by the
aerofoil and similarly to parasitic drag decreases with altitude, however due to the high
amount of lift required when an aircraft is flying slowly induced drag is very high when an
aircraft is at take-off and can cause dangerous conditions at the stall.
The calculation of induced drag begins with making two assumptions;
• Oswald efficiency factor, 𝑒𝑒, the value of 𝑒𝑒 relates to the configuration of the aircraft
and is also known as the airplane efficiency factor and is the relationship between
wing aspect ratio, sweep and wing position relative to fuselage.
• Correction factor, 𝛿𝛿, the value of 𝛿𝛿 is a correction factor used depending upon the lift
distribution across the wing and its relationship to the ideal elliptical distribution.
Incident
airflow
Lift
Net direction of airflow past aerofoil
Net direction
of airflow past aerofoil
Incident
airflow
Induced drag
Lift

Wing induced drag coefficient is then calculated using the aspect ratio of the wing and the lift
coefficient input into Equation 7.
𝐶𝐶𝐷𝐷𝐷𝐷 =
𝐶𝐶𝐿𝐿
2
𝜋𝜋𝜋𝜋𝜋𝜋𝜋𝜋
(1 − 𝛿𝛿)
Equation 7 - Induced Drag - [4]
Much like the calculation of the parasitic drag coefficient in section 2.1 this calculation must
be made for the most extreme aircraft conditions, therefore again cruise and take-off is
selected, 𝛿𝛿 is chosen to be 1.05 and 𝑒𝑒 to be 0.9 giving values for induced drag coefficient of
Table 4 and Table 5.
Component AR K CDi
Wing 5.785714286 1.313125 0.014043181
Vertical Stabiliser 2.123467766 1.101599003 0
Table 4 - Induced Drag Coefficient at Cruise
Component AR K CDi
Wing 5.785714286 1.313125 0.079555399
Vertical Stabiliser 2.123467766 1.101599003 0
Table 5 - Induced Drag Coefficient at Take-Off
2.3 Total Aircraft Drag
With the calculation of parasitic and induced drag complete the total drag for the aircraft can
be analysed for the most extreme aircraft conditions, this calculation is undertaken using the
coefficients previously found, Equation 8 and Equation 1.
𝐶𝐶𝐷𝐷 = 𝐶𝐶𝐷𝐷0 + 𝐶𝐶𝐷𝐷𝐷𝐷
Equation 8 - Drag Coefficient - [4]
This is calculated for both the aircraft take-off condition and the aircraft cruise condition giving
the results shown below in; Table 6
2.3.1 Total Drag at Cruise
Component Wetted Area CD CDi D
Fuselage 10.128672 0.003344672 0 61.42373

Wing 28 0.0090484 0.014043181 225.41721
Nose Gear 0.07 0.006331167 0 116.26966
Main Gear 0.11 0.003865586 0 70.990128
Horizontal Stabiliser 4.786158517 0.007054045 0.003982331 146.34604
Vertical Stabiliser 2.8 0.013365291 0 245.44886
Table 6 – Total Aircraft Drag at Cruise
Graph 2 - Comparison of Induced and Parasitic Drag at Cruise
Graph 3 - Comparison of Component Drag at Cruise
CD0
CDi
Fuselage
Wing
Nose Gear
Main Gear
Horizontal Stabaliser
Vertical Stabaliser

2.3.2 Total Drag at Take-Off
Component Wetted Area CD CDi D
Fuselage 10.128672 0.00354931 0 23.483752
Wing 28 0.009661648 0.079555399 748.95979
Nose Gear 0.07 0.006785073 0 44.892945
Main Gear 0.11 0.00413394 0 27.351917
Horizontal Stabiliser 4.786158517 0.007559818 0.001155098 59.965292
Vertical Stabiliser 2.8 0.014327675 0 94.798022
Table 7 - Total Aircraft Drag at Take-Off
Graph 4- Comparison of Induced and Parasitic Drag at Take-Off
Graph 5 - Comparison of Component Drag at Take-Off
CD0
CDi
Fuselage
Wing
Nose Gear
Main Gear
Horizontal Stabaliser
Vertical Stabaliser

2.4 Minimum Drag Condition
Along with the drag analysis requirement for power plant selection it can also be used to find
the minimum drag condition, this is the condition at which the aircraft flies at its most efficient
and therefore has its greatest endurance, it can be found through analysis of Equation 8 or
can be seen on a graph, therefore the induced and parasitic drag is plotted, Graph 6, so the
minimum drag speed can be found for cruise altitude.
Graph 6 - Total Aircraft Drag at 4000m
0
500
1000
1500
2000
2500
3000
3500
4000
4500
5000
0 10 20 30 40 50 60 70 80 90
DragForce(N)
Aircraft Speed (knots)
Parasitic Drag Induced Drag Total Drag

REFERENCES
[1] Airfoil Tools, “Reynolds number calculator,” 2015. [Online]. Available:
http://airfoiltools.com/calculator/reynoldsnumber. [Accessed APR 2015].
2012.
[3] D. J. Knight, Drag, Hatfield: University of Hertfordshire, 2014.
[4] D. J. Knight, Induced Drag, Hatfield: University of Hertfordshire, 2014.

Training Aircraft
Appendix 6
Propulsion Systems Design and Performance Analysis
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 6 to the main report, this document details the way in which the drag analysis data
from Appendix 5 is used to select a propulsion system for the aircraft and how the aircraft will
perform with the chosen system.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Power plant Design .............................................................................................................. 1
1.1.1 Propulsion System Type Selection ....................................................................... 1
1.1.2 Fuel System Type Selection.................................................................................. 1
1.2 Thrust Requirements..................................................................................................... 2
1.3 Power Requirements..................................................................................................... 2
1.4 Motor Selection ............................................................................................................. 2
1.5 Propellor Design............................................................................................................ 3
2 Performance Analysis........................................................................................................... 5
2.1 Take-Off Performance................................................................................................... 5
2.2 Aircraft Climb Performance ........................................................................................... 8
3 Aircraft Power Source........................................................................................................... 9
3.1 Energy Requirement ..................................................................................................... 9
3.2 Battery Specifications.................................................................................................. 11
REFERENCES............................................................................................................................ 12
iii

LIST OF FIGURES
Figure 1 - Aircraft Flight Profile ..................................................................................................... 9
Equation 1 - Cruise Engine Power - [1]......................................................................................... 2
Equation 2 - Engine Power Required for Take-Off - [1] ................................................................ 2
Equation 3 - Propellor Diameter - [1]............................................................................................. 3
Equation 4 - Propellor Tip Static Speed - [1]................................................................................. 3
Equation 5 - Propellor Required RPM - [1].................................................................................... 4
Equation 6 - Take-Off Ground Distance - [3] ................................................................................ 6
Equation 7 - Distance to Screen Height - [3]................................................................................. 7
Equation 8 - Aircraft Climb Angle - [3]........................................................................................... 8
Equation 9 - Rate of Climb - [3]..................................................................................................... 8
Equation 10 - Energy Required................................................................................................... 10
Table 1 - Motor Selection - [2]....................................................................................................... 3
Table 2 - Propellor Assumptions - [1]............................................................................................ 3
Table 3 - Aircraft Take-Off Speeds - [3] ........................................................................................ 5
Table 4 - Take-Off Ground Run for Maximum Take-Off Weight ................................................... 7
Table 5 - Aircraft Energy Usage.................................................................................................. 10
Table 6 - Battery Capacity........................................................................................................... 11
iv

1 Power plant Design
The aircraft power plant is the system the aircraft uses to produce thrust, offsetting the drag of
the aircraft and producing forward velocity and thus lift, the power plant is selected based on
the thrust requirements of the aircraft at cruise and take-off, for this aircraft a sole electric
propulsion system is selected using a removable fuel source and an electric motor.
1.1.1 Propulsion System Type Selection
Initially a propulsion method is selected, for a conventional aircraft this would be a selection
between a prop driven or jet aircraft, and then a selection between turbo-prop, conventional
prop, turbo fan, turbo jet, ram jet or a combination of these or others. However the designed
aircraft is not conventional, the selection of an electric fuel source limits the current available
technology to an electric motor and thus a prop driven aircraft, however electric jet engines
are in development using the same principles as conventional jet engines however currently
these are highly inefficient for the application proposed, mostly being used as propulsion for
model aircraft or spacecraft during orbital manoeuvres. Therefore as the aircraft would be
aimed at targeting a near future customer the electric motor is selected with a prop driven
aircraft configuration.
1.1.2 Fuel System Type Selection
With the propulsion system type selected a power source is required, within the design
specification laid out in Appendix 2 the power source is required to be removable, this limits
the available types of power source that can be used. Most simply a battery could be used to
store the electric energy and this could be ducted to the motor much like a conventional
aircraft, also conceivable is a mixture of solar and battery power, much like that used on some
solar aircraft today, the combination of battery and solar ‘recharge’ would work much like a
conventional aircraft fuel system with the batteries being topped up through the flight. Other
modern technologies that could be exploited are Formula 1’s kinetic energy recovery system
or ram air turbines exploiting the wasted energy created in braking and through flight however
neither could be the sole provider of power for the aircraft. Also conceivable is the use of
hydrogen power cells to generate the required power working in an almost identical way to
conventional fuel aircraft however this would require the storage of hydrogen on the aircraft
which may not be easily removable. Less conceivable but still a concept possibility is the use
of nuclear fission or fusion reactors, if these could be created in a small enough format but
still produce the required output this may be a possible fuel type however again due to the
near future market of this aircraft a battery system will be developed first, however it will be in
a format that could be used for several other fuel types and thus could be easily changed.
1

1.2 Thrust Requirements
The initial stage of the technical design of the propulsion system is the calculation of the
thrust requirement; this is initiated at the cruise condition with the requirement for steady
flight. At steady flight the aircraft is not accelerating nor decelerating, it is also not climbing or
falling thus both thrust and drag, and lift and weight are equal respectively, using this
condition it can be seen that the required thrust for steady flight is equal to the drag at steady
flight. Therefore from the analysis of Appendix 5 it is clear that the thrust required for steady
flight is 865.9N thus the aircraft power plant must be able to produce 865.9N of thrust at
4500m.
1.3 Power Requirements
As the propulsion system type has been selected as an electric motor a more conventional
unit of measurement is required so that a motor can be selected, also due to the prop driven
nature of the aircraft a correction factor is required due to the efficiency of the propeller, as
the propeller is an aerodynamic surface it is not 100% efficient and thus the motor will require
more power to negate the efficiency losses.
𝑃𝑃𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶 𝐶𝐶 =
𝑇𝑇𝑉𝑉𝐶𝐶
𝜂𝜂𝑝𝑝
Equation 1 - Cruise Engine Power - [1]
𝑃𝑃𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶 = 𝑃𝑃𝑀𝑀𝑀𝑀𝑀𝑀 𝜎𝜎1.2
Equation 2 - Engine Power Required for Take-Off - [1]
Through the use of Equation 1 and Equation 2 using data from Appendix 2 and section 1.2
the power required by the motor can be calculated, using a cruise altitude of 4500m and a
propeller efficiency of 0.8 [1], an average for modern aircraft propellers, the cruise power
required is found to be 61.24kW and the take-off power required is 99.23kW. These power
requirements allow a motor to be selected or designed, for similar reasons as those used in
Appendix 4 and the aerofoil selection the motor is chosen to be selected from an existing
manufacturer rather than developing a new unit, this is to reduce development time and costs
for the aircraft.
1.4 Motor Selection
With the required power from the motor calculated the power plant can be selected, as
previously stated the motor selected will be of current design to fulfil the low cost and low
development time requirements for the aircraft. Several motors are selected for evaluation
from UQM Technologies due to the good availability of information for their products, plus
their suitability for the project, the selected motors are listed in Table 1.
2

Name Production
Company
Max
Power
Continuous
Power
Weight Required V Supply
PowerPhase
Select 145
UQM 145 85 50 340-420 DC
PowerPhase
Pro 135
UQM 135 60 50 270-425 DC
PowerPhase
Select 125
UQM 125 45 15.9 300-420 DC
PowerPhase
Pro 100
UQM 100 60 50 270-425 DC
Table 1 - Motor Selection - [2]
From the calculated requirements it is seen that the PowerPhase Pro 100 would provide the
required max power but would not be able to continuously produce the required continuous
power for steady flight, therefore the PowerPhase Select 145 is selected and detailed in
Appendix 6a.
1.5 Propellor Design
To accompany the motor a propeller is designed, again the propeller would be selected to
reduce costs and development time however in this text only the propeller requirements are
calculated using assumptions of propeller performance, this is to both size the propeller for
the landing gear requirement in Appendix 7 and to size a gearbox for the aircraft. The
propeller design begins with calculating a propeller diameter using Equation 3.
𝐷𝐷𝑃𝑃 = 𝐾𝐾𝑁𝑁𝑁𝑁�
2𝑃𝑃𝑀𝑀𝑀𝑀𝑀𝑀 𝜂𝜂𝑃𝑃 𝐴𝐴𝐴𝐴𝑃𝑃
𝜌𝜌�0.7𝑉𝑉𝑡𝑡𝑡𝑡𝑡𝑡_𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐�
2
𝐶𝐶𝐿𝐿𝐿𝐿 𝑉𝑉𝐶𝐶
Equation 3 - Propellor Diameter - [1]
ηP 0.8
ARP 9
Vtip_cruise 250
CLP 0.3
Table 2 - Propellor Assumptions - [1]
𝑉𝑉𝑡𝑡𝑡𝑡𝑡𝑡_𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐𝑐 = �𝑉𝑉𝑡𝑡𝑡𝑡𝑡𝑡_𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠
2
+ 𝑉𝑉𝐶𝐶
2
Equation 4 - Propellor Tip Static Speed - [1]
3

𝑉𝑉𝑡𝑡𝑡𝑡𝑡𝑡_𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠𝑠 =
𝐷𝐷𝑃𝑃 𝜋𝜋𝜋𝜋
60
Equation 5 - Propellor Required RPM - [1]
Using the assumptions in Table 2 and the aircraft cruise data from Appendix 4 the propeller
diameter is calculated using Equation 3 at 2.21m, using this and Equation 4 the propeller tip
static speed is calculated as 243.51ms
-1
and using Equation 5 the required RPM is 2100.06.
From analysing the data in Appendix 6a it can be seen that for the motor the most efficient
power application is 85kW, as stated in Appendix 5 the drag calculations for the aircraft are
too conservative and therefore it can be expected that the aircraft will require more power in
the cruise, also to be noted is the gearbox, this will take some of the required power from the
motor due to friction and other resistive forces so therefore it is prudent to assume that the
motor will be required to produce around 85kW of power thus the most efficient RPM for the
motor at this power output is around 4000 RPM therefore the gearbox will be required to half
the output RPM of the motor and thus has a gear ratio of around 2:1.
4

2 Performance Analysis
The performance analysis of the aircraft is required to calculate the ability of the aircraft
through all flight phases, for this concept aircraft the take-off run, rate of climb and speeds of
the aircraft are to be calculated to size the correct fuel source.
2.1 Take-Off Performance
To begin the analysis of take-off performance data is taken from Appendices 4 and 5 and
section 1, to begin the analysis the relevant speeds for the aircraft must be determined:
• Minimum Control Speed, 𝑉𝑉𝑀𝑀𝑀𝑀, the speed at which the aircrat control surfaces start to
become effective.
• Stall Speed, 𝑉𝑉𝑆𝑆, the speed at which the aircraft stalls.
• Critical Engine Failure Speed, 𝑉𝑉1, the speed at which the pilot can safely carry out the
take-off in the event of engine failure.
• Rotation Speed, 𝑉𝑉𝑅𝑅, the speed at which the aircraft begins rotation to increase wing
angle of attack.
• Minimum Unstick Speed, 𝑉𝑉𝑀𝑀𝑀𝑀, the speed at which the aircraft can take-off even with
one engine inoperative.
• Lift-Off Speed, 𝑉𝑉𝐿𝐿𝐿𝐿𝐿𝐿, the speed at which the aircraft lifts off the ground.
• Take-Off Climb Speed, 𝑉𝑉2, the speed at which the aircraft has achieved 10.7m in
altitude and begins climb away.
From the analysis in Appendix 8 due to the aircraft only having one engine, the one engine
inoperative conditions during take-off are to be considered as total engine failure conditions,
therefore it would not be expected for the pilot to continue take-off with total engine failure and
thus 𝑉𝑉1 and 𝑉𝑉𝑀𝑀𝑀𝑀 are not applicable to the aircaft, however will be used to estimate the relevant
speeds. To begin the speed analysis the design appendicies are reffered to, such as
Appendix 4 for aircraft stall speed and Appendx 8 for minimum control speed, using this data
and Table 3 the relevant take-off speeds for a 0m altitude runway are calculated.
𝑽𝑽 𝑴𝑴𝑴𝑴 Appendix 8 36 knots
𝑽𝑽𝑺𝑺 Appendix 4 45 knots
𝑽𝑽𝟏𝟏 𝑉𝑉1 ≥ 1.05 𝑉𝑉𝑀𝑀𝑀𝑀 37.8 knots
𝑽𝑽𝑹𝑹 𝑉𝑉𝑅𝑅 ≥ 1.05 𝑉𝑉𝑀𝑀𝑀𝑀 43.2 knots
𝑽𝑽 𝑴𝑴𝑴𝑴 𝑉𝑉𝑀𝑀𝑀𝑀 = 𝑉𝑉𝑅𝑅 43.2 knots
𝑽𝑽𝑳𝑳𝑳𝑳𝑳𝑳 𝑉𝑉𝐿𝐿𝐿𝐿𝐿𝐿 ≥ 1.10 𝑉𝑉𝑀𝑀𝑀𝑀 54 knots
𝑽𝑽𝟐𝟐 𝑉𝑉2 ≥ 1.2 𝑉𝑉𝑆𝑆 54 knots
Table 3 - Aircraft Take-Off Speeds - [3]
5

Using this data the take-off performance of the aircraft can be analysed, for its lightest and
heaviest take-off cases, to calculate the distance required for the aircraft to reach speed 𝑉𝑉𝐿𝐿𝐿𝐿𝐿𝐿,
Equation 6 can be utilised.
𝑠𝑠1 =
𝑊𝑊
𝜌𝜌𝜌𝜌𝜌𝜌(𝜇𝜇𝜇𝜇𝐿𝐿 − 𝐶𝐶𝐷𝐷)
ln �1 +
𝜌𝜌𝜌𝜌(𝜇𝜇𝜇𝜇𝐿𝐿 − 𝐶𝐶𝐷𝐷)
2(𝑇𝑇 − 𝜇𝜇𝜇𝜇)
�
Equation 6 - Take-Off Ground Distance - [3]
By using this and data form Appendix 4 and Appendix 5 the calculation the take-off distance
can be undertaken, the results shown in .
Time Speed Lift Drag Distance
0 0 0 0 0
0.181846 1 0.051506 0.524857 0.140313
0.363812 2 0.39753 1.952646 0.421091
0.545966 3 1.325328 4.231623 0.889384
0.728371 4 3.122155 7.339768 1.545612
0.911091 5 6.075269 11.26307 2.39032
1.094182 6 10.47193 15.99144 3.424173
1.277701 7 16.59938 21.51708 4.647946
1.461701 8 24.74489 27.83368 6.062518
1.646235 9 35.19572 34.93597 7.668871
1.831351 10 48.23911 42.81944 9.468082
2.017097 11 64.16233 51.48019 11.46132
2.203519 12 83.25263 60.91476 13.64983
2.390659 13 105.7973 71.12006 16.03495
2.578561 14 132.0835 82.0933 18.61809
2.767262 15 162.3986 93.83197 21.40073
2.956801 16 197.0298 106.3337 24.38441
3.147214 17 236.2643 119.5965 27.57073
3.338535 18 280.3895 133.6182 30.96137
3.530794 19 329.6926 148.3971 34.55802
3.724022 20 384.4607 163.9315 38.36244
3.918246 21 444.9813 180.2197 42.37641
4.11349 22 511.5415 197.2603 46.60176
4.309778 23 584.4286 215.0518 51.04033
4.507074 24 663.9299 233.0562 55.69328
4.705506 25 750.3326 252.8822 60.56385
4.905034 26 843.924 272.9186 65.65321
6

5.10567 27 944.9913 293.701 70.96323
5.307422 28 1053.822 315.2282 76.49575
5.510297 29 1170.703 337.4992 82.25259
5.714298 30 1295.921 360.513 88.23551
5.919423 31 1429.765 384.2686 94.44625
6.125669 32 1572.521 408.765 100.8865
6.333029 33 1724.476 434.0015 107.5578
6.54149 34 1885.918 459.9771 114.4617
6.751037 35 2057.134 486.691 121.5996
6.96165 36 2238.411 514.1424 128.9728
7.173307 37 2430.037 542.3306 136.5825
7.385978 38 2632.298 571.2547 144.4299
7.599632 39 2845.483 600.9142 152.5157
7.814231 40 3069.877 631.3082 160.8408
8.029733 41 3305.769 662.4362 169.4057
8.246093 42 3553.446 694.2974 178.2108
8.463259 43 3813.195 726.8913 187.2564
8.681176 44 4085.304 760.2171 196.5423
8.899781 45 4370.059 794.2744 206.0684
9.11901 46 4667.747 829.0625 215.834
9.338792 47 4978.657 864.5808 225.8386
9.559051 48 5303.075 900.8288 236.0811
9.779708 49 5641.289 937.806 246.5602
10.00068 50 5993.585 975.5119 257.2743
10.22187 51 6360.251 1013.946 268.2216
10.4432 52 6741.575 1053.107 279.3999
10.66455 53 7137.844 1092.996 290.8068
10.88831 54 7549.344 1133.612 302.5086
Table 4 - Take-Off Ground Run for Maximum Take-Off Weight
The second part of the calculation for Take-Off run is the climb to screen height, this is the
distance covered as the aircraft travels from lift off to 10.7m altitude, utilising Equation 7 this
distance can be found and the entire take-off run for the aircraft can be found.
𝑠𝑠𝐴𝐴 =
𝑊𝑊
(𝑇𝑇 − 𝐷𝐷)
�
(𝑉𝑉2
2
− 𝑉𝑉𝐿𝐿𝐿𝐿𝐿𝐿
2 )
2𝑔𝑔
+ 10.7�
Equation 7 - Distance to Screen Height - [3]
Through calculation of both these distances the entire ground run can be calculated and
plotted in Graph 1.
7

Graph 1 - Take-Off Ground Distance
2.2 Aircraft Climb Performance
The second required performance statistic is the aircraft climb angle and rate, the aircraft
climb performance is analysed by finding the excess thrust that the aircraft has available and
utilising this to climb. The calculation of the aircraft climb angle is initiated using Equation 8
and the speed of climb is found using Equation 9.
𝐶𝐶𝐶𝐶𝐶𝐶 𝐶𝐶𝐶𝐶 𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴 = sin−1
�
(𝑇𝑇 − 𝐷𝐷)
𝑊𝑊
�
Equation 8 - Aircraft Climb Angle - [3]
𝑅𝑅𝑅𝑅𝑅𝑅 =
𝑉𝑉(𝑇𝑇 − 𝐷𝐷)
𝑊𝑊
Equation 9 - Rate of Climb - [3]
Utilising the data from Appendix 5 it is found that the aircraft will climb at a 14.6° angle, with
the wing setting angle at 4° the aircraft will climb at a fuselage angle of 10.6° at a rate of
127.7m/min however this is a very conservative calculation and would require further analysis
of the thrust and drag of the aircraft to find the climb performance of the aircraft more
accurately.
0
50
100
150
200
250
300
350
400
0 2 4 6 8 10 12 14
Distance(m)
Time (s)
Two Pilots Full Baggage One Pilot No Baggage
8

3 Aircraft Power Source
With the crucial performance characteristics of the aircraft found and the flight profile data
available from Appendix 4 the aircraft fuel source can be specified, as stated in section 1.1.2
the fuel type to be used is a battery bank, this is due to the development of battery technology
in recent years with the research and development of the Airbus E-FAN 2.0 and other aircraft
plus the interest in green technologies for the motorsport and automotive industries as stated
in Appendix 1.
3.1 Energy Requirement
From analysis of the aircraft flight profile the worst flight situation for the aircraft is the cruise
no reserve, this condition should never be encountered however it must be considered as the
worst case, the battery must be designed to idle, climb, cruise, descend, land and idle again
for a 9 hour period, this is a huge difference to the 1 hour endurance of the E-FAN 2.0
however with advances in technology this concept may be possible in the near future.
Figure 1 - Aircraft Flight Profile
The calculation for the required power begins with the calculation of the power required for
each stage of flight, using the power required multiplied by the time required for the required
energy for each flight stage can be found.
0
500
1000
1500
2000
2500
3000
3500
4000
4500
0 1 2 3 4 5 6 7 8 9
Altitude(m)
Time (hrs)
Cruise No Reserve Cruise With Reserve 30 min Training Flights
2 hr Training Flights Aerobatics
9

𝐸𝐸 = 𝑃𝑃 × 𝑡𝑡
Equation 10 - Energy Required
By using Equation 10 the energy required for each flight stage can be found, 2 kW is added to
run any axillary systems that the aircraft requires.
Stage Power (kW) Time (h) Energy (kWh)
Idle 2 0.75 1.5
Taxi 35 0.25 8.75
Take-Off 117 0.00416667 0.4875
Climb 117 0.51666667 60.45
Cruise 87 6 522
Descent 77 0.5 38.5
Landing 52 0.08333333 4.333333
Taxi 35 0.25 8.75
Idle 2 0.5 1
Table 5 - Aircraft Energy Usage
Graph 2 - Comparison of Aircraft Flight Stage Energy Usage
For a flight with a cruise of 6 hours at 4500m it is found that the batteries are required to
provide at least 645.7708kWh of energy or 2324.775MJ, this is equivalent to around 76 litres
of petrol.
Idle
Taxi
Take-Off
Climb
Cruise
Descent
Landing
Taxi
Idle
10

3.2 Battery Specifications
The motor and controller however requires a voltage of 340V to 430V DC, with a power of
145kW giving a maximum current of 453.125A reducing to 265.625A at cruise therefore the
battery capacity must be calculated.
Stage Power (kW) Current (A) Time (h) Capacity
Idle 2 6.25 0.75 4.6875
Taxi 35 109.375 0.25 27.34375
Take-Off 117 365.625 0.004167 1.523438
Climb 117 365.625 0.516667 188.9063
Cruise 87 271.875 6 1631.25
Descent 77 240.625 0.5 120.3125
Landing 52 162.5 0.083333 13.54167
Taxi 35 109.375 0.25 27.34375
Idle 2 6.25 0.5 3.125
Table 6 - Battery Capacity
Therefore the battery is found to need a capacity of 2018.035Ah, however as the transfer
cannot be 100% efficient the battery is chosen to hold 2500Ah giving an efficiency of
approximately 80% in line with Appendix 1.
The battery has been chosen to weigh 30kg each through Appendix 10 and the design
requirement for easy handling in Appendix 2, therefore the specific energy of each battery is
required to be around 10.76kWh/kg or 77.4925MJ/kg.
11

REFERENCES
2012.
[2] UQM Technologies, “Innovative Solutions for Electrifying Vehicles,” UQM Technologies,
2015. [Online]. Available: https://uqm.com/products/full-electric/prototype/commercial-
vehicles/. [Accessed 2015].
[3] D. K. Hart, Aircraft Performance, Hatfield: University of Hertfordshire, 2010.
12

Training Aircraft
Appendix 6a
UHQ PowerPhase Select 145
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015

PAGE INTENTIONALLY BLANK

for electric, hybrid electric, and fuel cell powered vehicles
PowerPhase®
145
Key Features:
• 400 Nm peak torque
• 145 kW peak, 85 kW continuous motor power
• 145 kW peak, 85 kW continuous generator power
• Full Power at 340-430 VDC
• EV/HEV traction drive or HEV starter/generator system
• Efﬁcient, power dense, brushless permanent magnet motor
• Microprocessor-controlled inverter with sine wave drive
• Application-friendly graphical user interface
• Regenerative Braking
Beneﬁts:
Tight voltage regulation
Improved braking and extended range
Suitable for automotive applications
Enhanced thermal management
Torque, speed, and voltage control modes
Rugged, weatherproof enclosure
Liquid cooling
Light weight
Driver Electronics Incorporate:
Serial communication
CAN bus compatibility
Diagnostic capability
Temperature sensing/alarm
Speed sensing
Graphical user interface

SPM218-143-3 Motor/Generator
PowerPhase®
145
Dimensions
Length 10.987 in 279 mm
Diameter 11.00 in 280 mm
Weight 110 lb 50 kg
Performance
Peak power 194 hp 145 kW
Continuous power at 5,000 rpm 114 hp 85 kW
Peak torque 295 lbf•ft 400 N•m
Continuous torque 184 lbf•ft 250 N•m
Maximum speed 8000 RPM
Maximum efﬁciency 94%
Power density (based on 145 kW) 1.76 hp/lb 2.90 kW/kg
www.uqm.com sales@uqm.com 303.682.4900 4120 Specialty Pl., Longmont CO 80504

DD45-500L Inverter/Controller
PowerPhase®
145
Dimensions
Length 14.96 in 380 mm
Width 14.37 in 365 mm
Height 4.69 in 119 mm
Weight 35.0 lb 15.9 kg
Operating Voltage
Nominal input range 340 to 430 VDC
Operating voltage input range 240 to 430 VDC
Minimum voltage limit 240 VDC (with derated power output)
Input current limitation 500 A
Inverter Type
Control type PWM & phase advance,
3-Phase Brushless PM
Power device IGBT module half bridge × 3
Switching frequency 12.5 kHz
Standby power consumption 17 W (inverter and microprocessor)
Liquid Cooling System
Minimum coolant ﬂow 8 l/min (50/50 water/glycol mix)
Max. inlet temp of controller 131° F 55° C
Inner diameter of hose 5/8 in 16 mm
Max. inlet pressure 10 psig 0.7 bar
TI2812 Digital Signal Processor (internally packaged)
Nominal input voltage 12 VDC
Input supply voltage range 8 to 15 VDC
Input supply current range 0.3 to 0.5 A

PowerPhase®
145
Testing Conditions
Continuous Output: 85 kW at 5,000 rpm with 340 VDC input, 55•
C coolant
Peak Output: 145 kW with 340 VDC input, 55•
C coolant, duration 30-90 seconds
Motoring Efﬁciency Map
Includes controller and motor
Speed (rpm/100)

PowerPhase®
145
Testing Conditions
C coolant
Motoring Efﬁciency Map
Includes controller and motor

PowerPhase®
145
Testing Conditions
C coolant
Generating Efﬁciency Map
Includes controller and generator

Training Aircraft
Appendix 7
Landing Gear and Structural Design and Analysis
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 7 to the main report, this document details the way in which the landing gear is
designed and positioned and how the structure is designed and analysed.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Landing Gear Design............................................................................................................ 1
2 Structural Design .................................................................................................................. 3
2.1 Flight Critical Components ............................................................................................ 3
2.2 Failure and Crash Critical Components ........................................................................ 3
REFERENCES.............................................................................................................................. 5
iii

LIST OF FIGURES
No table of figures entries found.
Equation 1 – Aircraft Engine Centreline Clearance - [1] ............................................................... 1
Equation 2 - Aircraft Landing Gear Height - [1]............................................................................. 1
Equation 3 - Nose Gear Force - [1]............................................................................................... 2
Table 1 - Landing Gear Requirements.......................................................................................... 1
Table 2 - Undercarriage Loading .................................................................................................. 2
Table 3 - Structural Breakdown..................................................................................................... 3
iv

1 Landing Gear Design
The landing gear for an aircraft is the components on which the aircraft stands, designed to
hold the aircraft off the ground for engine or propeller clearances and a means of landing the
aircraft without damaging aircraft components, landing gear may be of many forms, with
wheels being common but other forms such as skids, skies, floats or keels can also be used.
From the concept development in Appendix 3 the aircraft has been chosen to use a tricycle
undercarriage arrangement utilising a fixed wheeled landing gear configuration, the landing
gear design process begins with the ranking of the landing gear requirements so that the
worst condition for the landing gear can be identified.
Condition Rank
Wing Surface Clearance 4
Fuselage Surface Clearance 2
Propeller Clearance 1
Stabiliser Clearance 5
Take-Off Rotation 3
Table 1 - Landing Gear Requirements
From analysing the concept in Appendix 3, Table 1 is created to identify the requirements of
the undercarriage and as shown the propeller clearance is the worst case scenario for the
undercarriage and thus the propeller clearance will dictate the length of the undercarriage.
𝐻𝐻𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶 𝐶𝐶𝐶𝐶 = Δ𝐻𝐻𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶 +
𝐷𝐷𝑃𝑃
2
Equation 1 – Aircraft Engine Centreline Clearance - [1]
𝐻𝐻𝐿𝐿𝐿𝐿 = 𝐻𝐻𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶𝐶 𝐶𝐶𝐶𝐶 −
𝐷𝐷𝑃𝑃
2
Equation 2 - Aircraft Landing Gear Height - [1]
From Appendix 6 the data for the wing can be found and using Appendix 5 the fuselage
diameter is available and thus using Equation 1 and Equation 2 the aircraft landing gear
height is calculated as 0.77m from the aircraft fuselage and 1.36m from the aircraft centreline.
With the height of the landing gear selected the aircraft track and base must be defined, the
landing gear track is the distance between the main gear laterally and the base is the distance
between the main and nose or tail gear. For an aircraft with tricycle landing gear around 85%
of the aircraft weight is required on the main gear and to maintain control during the taxi
around 15% of the aircraft weight is required on the aircraft nose gear [1]. Through iteration of
the elevator requirement in Appendix 8 the main gear position is found to be at 0.2m behind
1

the foremost aircraft centre of gravity, using the tricycle undercarriage loading requirement the
nose gear position can be found using Equation 3.
𝐹𝐹𝑁𝑁𝑁𝑁 =
𝐵𝐵𝑀𝑀𝑀𝑀 𝑊𝑊
𝐵𝐵
Equation 3 - Nose Gear Force - [1]
From the data available the force on the main gear is found to be 6475N and the force at the
nose gear is found to be 883N, utilising Equation 3 it is found that the aircraft requires a base
of 1.67m placing the main gear at 2.66m from the nose and the main gear 1m from the nose.
However the landing gear must be specified for landing, with the downward velocity of the
aircraft causing the dynamic loading upon the aircraft to be greater than the static loading. To
account for this velocity component a factor of 1.5 – 2 can be applied to the force upon the
landing gear and thus the maximum expected loading upon each wheel is shown in Table 2.
From the maximum static and maximum dynamic load expected upon the undercarriage
component a wheel and tyre can be specified, again for decreased cost and development
time an existing component is selected, specified in Appendix 7a and Appendix 7b.
Position Static Force (N) Max Force (N) Wheel Tyre
Nose Gear 882.9 1766 Grove 51-1A Dunlop DA13822
Left Main Gear 3237.3 6475 Grove 51-1A Dunlop DA13822
Right Main Gear 3237.3 6475 Grove 51-1A Dunlop DA13822
Table 2 - Undercarriage Loading
The landing gear is also used for braking during landing, due to Appendix 4 the aircraft will
land between 54knots and 45 knots, causing at maximum 144583Nm of Kinetic Energy per
wheel, the brakes will consist of two Kevlar based brake pads clamping onto a steel brake
disc by means of a hydraulic brake system.
2

2 Structural Design
The structure of the aircraft has two main functions, one to hold all components of the aircraft
together and prevent structural failure of any component and two, to protect the passengers in
the event of a failure or crash. Therefore the structure must be strong enough the both
maintain structural integrity during all flight conditions and strong enough to protect the pilot
and co-pilot in the event of a crash, however due to its relatively high weight component it
must also be as light as possible, the aim of the structural design is to fulfil both these
conditions in the most efficient way possible. Therefore the aircraft structure is split into two
sections, failure and crash critical and flight critical, the breakdown of components is shown in
Table 3, with failure and crash critical components being those that are critical to the survival
of passengers during a crash or failure and flight critical being those components that are
critical to the flight of the aircraft.
Flight Critical Failure Critical
Wing Cockpit
Horizontal Stabiliser Fore Firewall
Vertical Stabiliser Aft Firewall
Tail Arm
Engine Bay
Table 3 - Structural Breakdown
2.1 Flight Critical Components
The flight critical components are the components which the aircraft requires to fly, the wings
of the aircraft are considered initially due to the similarity of the structure to those of the
horizontal and vertical stabiliser, using Pradtl’s lifting line theory again the wing lift distribution
of the aerodynamic surface is analysed and the force upon several sections is calculated, the
structure in the wing will be required to offset this force at its maximum, each wing structure
will consist of a main and rear spar and several ribs. The tail arm is required to resist the force
of the horizontal and vertical stabiliser as its corrects the aircraft pitching moment and thus
must be strong enough in both the lateral and vertical motion, the engine bay must also be
strong enough to hold all major engine components throughout the flight and resist the torque
effect of the motor throughout the flight.
2.2 Failure and Crash Critical Components
The failure and crash critical components are the components which the aircraft requires to
maintain structural integrity in the event of failure or a crash scenario, again this category can
be split into two sub-categories being crash condition and catastrophic failure condition with
3

the cockpit structure being required in the crash condition and firewall structure being required
in a catastrophic failure such as engine fire or battery fire. The most important of these is the
survival of the cockpit section in a crash situation and thus the structure in this section must
be built to a suitable standard. The design of the aircraft structures for failure and crash
critical components is discussed in Appendix 9; the design however has been built to
withstand a force of around 29430N which represents a 4g crash.
4

REFERENCES
2012.
5

Training Aircraft
Appendix 7a
Grove Aircraft 50-102 501-A
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015

http://www.groveaircraft.com/5series.html

Training Aircraft
Appendix 7b
Grove Aircraft 50-102 501-A
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015

Part
Number
Tyre Size Aircraft Main or AUX
DA13822 5.00-4.5 MB326 AUX
DA13822 5.00-4.5 MB329 AUX
Characteristics
Chined No
Ply Rating 6
Tubed/Tubeless TT
Aspect Ratio 0.85
Speed MPH 160
Max Load lbs 1650
Inflation And Dimensions
Inflation_Pressure_Unloaded_psi 78
Inflation_Pressure_Loaded_psi 81
Inflation_Pressure_Type Standard
Typical_Weight_lbs 7.10
Inf_Dim_Width_Shoulder
Inf_Dim_Width_Max 5.30
Inf_Dim_Width_Min 5.00
Tread_Type TC
Skid_Depth
Inf_Dim_OD_Min 12.95
Inf_Dim_OD_Max 13.45
Inf_Dim_OD_Shoulder
Max_Load_lbs 1650
Loaded_Radius 5.75
Rim_Dim_Width_Between_Flanges4.75
Rim_Dim_Ledge_Diameter 4.50
Rim_Dim_Flange_Height 0.66
Rim_Dim_Min_Ledge_Width 0.96
Approvals
QTR No 768
Test Spec MIL-T-5041F
CAA EASA Approval
FAA Approval
MOD Approval Y
NSN Approval
Civ/Mil M
http://www.dunlopaircrafttyres.com/products/part-search.aspx

Training Aircraft
Appendix 8
Stabiliser Design, Control Surface Design and Stability and
Control Analysis
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
Appendix 8 to the main report, this document details the way in which the stability of the aircraft
is designed and analysed with reference to the horizontal and vertical stabilisers, pitch, yaw and
roll control surfaces and theoretical and simulated analysis of the flight characteristics of the
aircraft.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
LIST OF FIGURES.........................................................................................................................v
1 Centre of Gravity .................................................................................................................. 1
1.1 Centre of Gravity Analysis............................................................................................. 1
2 Longitudinal Stability............................................................................................................. 5
2.1 Longitudinal Static Stability ........................................................................................... 5
2.1.1 Pitching Moment.................................................................................................... 5
2.1.2 Stabiliser Moment Arm.......................................................................................... 6
2.1.3 Aerofoil Selection .................................................................................................. 6
2.1.4 Horizontal Stabiliser Design .................................................................................. 7
2.1.5 Horizontal Stabiliser Vertical Position ................................................................... 8
2.1.6 Horizontal Stabiliser Setting Angle........................................................................ 9
2.1.7 Stick Fixed Static Longitudinal Stability of Aircraft ................................................ 9
2.1.8 Neutral Point Analysis ......................................................................................... 10
2.2 Longitudinal Dynamic Stability .................................................................................... 10
2.2.1 Phugoid Motion ................................................................................................... 11
2.2.2 Phugoid Approximation ....................................................................................... 11
2.2.3 Aerodynamic Derivatives..................................................................................... 12
2.2.4 Phugoid Calculation ............................................................................................ 13
2.2.5 Short Period Pitching Oscillation......................................................................... 14
2.2.6 Short Period Pitching Oscillation Approximation................................................. 14
2.2.7 Short Period Pitching Oscillation Calculation...................................................... 15
2.2.8 Flying Characteristics .......................................................................................... 15
2.2.9 Elevator Design ................................................................................................... 16
2.2.10 Elevator to Trim ................................................................................................... 16
3 Lateral Stability ................................................................................................................... 18
3.1 Static Directional Stability............................................................................................ 18
3.1.1 Vertical Stabiliser Design .................................................................................... 18
3.1.2 Static Directional Stability Derivative Calculation................................................ 19
3.2 Lateral Dynamic Stability............................................................................................. 19
3.2.1 Crosswind Requirement...................................................................................... 21
3.2.2 Rudder Design .................................................................................................... 21
3.2.3 Aileron Design ..................................................................................................... 22
3.2.4 Spiral Mode Analysis........................................................................................... 23
3.2.5 Spiral Mode Approximation ................................................................................. 23
3.2.6 Spiral Mode Calculation ...................................................................................... 24
iii

3.2.7 Roll Convergence Analysis ................................................................................. 24
3.2.8 Roll Convergence Approximation........................................................................ 24
3.2.9 Roll Convergence Calculation............................................................................. 25
3.2.10 Dutch Roll............................................................................................................ 25
3.2.11 Dutch Roll Approximation.................................................................................... 25
3.2.12 Dutch Roll Calculation......................................................................................... 26
3.2.13 Flying Characteristics .......................................................................................... 26
REFERENCES............................................................................................................................ 28
iv

LIST OF FIGURES
Figure 1- Centre of Gravity Variation in Longitudinal Axis ............................................................ 4
Figure 2 - Aircraft Dimensions....................................................................................................... 6
Figure 3 - Elevator Angle to Trim at Cruise Altitude.................................................................... 17
Figure 4 - Control Surface Angle of Attack Effectiveness Parameter – [1]................................. 21
Equation 1 – Centre of Gravity Equation - [1] ............................................................................... 1
Equation 2 - Pitching Moment Equation - [2] ................................................................................ 5
Equation 3 - Volume Coefficient Equation - [2] ............................................................................. 5
Equation 4 - Lift Equation - [2]....................................................................................................... 6
Equation 5 - Horizontal Stabiliser Height Equation - [1]................................................................ 9
Equation 6 - Horizontal Stabiliser Incidence - [2] .......................................................................... 9
Equation 7 - Stick Fixed Longitudinal Stability Equation - [2] ....................................................... 9
Equation 8 - Neutral Point Equation - [2]..................................................................................... 10
Equation 9 - X Equation - [4]....................................................................................................... 10
Equation 10 - Z Equation - [4] ..................................................................................................... 11
Equation 11- Pitching Moment Equation - [4] ............................................................................. 11
Equation 12 - Phugoid Approximation X Equation...................................................................... 11
Equation 13 - Phugoid Approximation Z Equation...................................................................... 11
Equation 14 - Phugoid Approximation Determinant.................................................................... 12
Equation 15 - Phugoid Approximation Characteristic Equation .................................................. 12
Equation 16 - General Characteristic Equation - [4] ................................................................... 12
Equation 17 - Motion Period Equation ........................................................................................ 12
Equation 18 - Motion Time to Half Amplitude Equation .............................................................. 12
Equation 19 - SPPO Approximation Z Equation ......................................................................... 14
Equation 20 - SPPO Approximation Pitching Moment Equation ................................................ 14
Equation 21 - SPPO Approximation Determinant....................................................................... 15
Equation 22 - SPPO Approximation Characteristic Equation ..................................................... 15
Equation 23 - Elevator Angle to Trim Equation - [2].................................................................... 16
Equation 24 - Static Directional Stability Derivative Equation - [1] ............................................. 18
Equation 25 - Sideslip Equation - [8]........................................................................................... 19
Equation 26 - Roll Equation - [8] ................................................................................................. 19
Equation 27 - Roll Moment Equation - [8] ................................................................................... 20
Equation 28 - Rudder Control Derivative Equation - [1].............................................................. 22
Equation 29 - Required Rudder Deflection for OEI Equation - [1] .............................................. 22
Equation 30 - Aileron Rolling Moment Coefficient Equation - [1]................................................ 22
Equation 31 - Steady State Roll Rate Equation - [1]................................................................... 22
Equation 32 - Moment of Inertia in X - [7] ................................................................................... 23
v

Equation 33 - Bank Angle for Steady Roll Rate Equation - [1].................................................... 23
Equation 34 - Time to Bank Equation - [1] .................................................................................. 23
Equation 35 - Spiral Mode Approximation Characteristic Equation - [7]..................................... 23
Equation 36 - Roll Convergence Approximation Equation - [8] .................................................. 24
Equation 37 – Roll Convergence Approximation Characteristic Equation.................................. 25
Equation 38– Dutch Roll Approximation Sideslip Equation ........................................................ 25
Equation 39 – Dutch Roll Approximation Roll Moment Equation................................................ 25
Equation 40 - Dutch Roll Approximation Determinant ................................................................ 26
Equation 41 - Dutch Roll Approximation Characteristic Equation............................................... 26
Table 1 – Component Centre of Gravity Analysis......................................................................... 2
Table 2 - Component Moment Analysis........................................................................................ 3
Table 3 - Aircraft Centre of Gravity ............................................................................................... 3
Table 4 - Load Considerations ...................................................................................................... 4
Table 5 – NACA 0009 Aerofoil Data ............................................................................................. 7
Table 6 - MatLab Script Inputs ...................................................................................................... 8
Table 7 - Horizontal Stabiliser Parameters ................................................................................... 8
Table 8 - Longitudinal Aerodynamic Derivatives......................................................................... 13
Table 9 - Phugoid Approximation Results................................................................................... 14
Table 10 - SPPO Approximation Results.................................................................................... 15
Table 11 - Longitudinal Flying Characteristics - [7]..................................................................... 16
Table 12- Vertical Stabiliser Parameters..................................................................................... 19
Table 13 - Lateral Aerodynamic Derivatives - [7]........................................................................ 20
Table 14 - Spiral Mode Approximation Results........................................................................... 24
Table 15 - Roll Convergence Approximation Results ................................................................. 25
Table 16 – Dutch Roll Approximation Results ............................................................................ 26
Table 17 - Lateral Flying Characteristics - [7] ............................................................................. 27
Code 1 - MatLab Tail Lift Script - [1] – (Modified by Benjamin James Johnson).......................... 7
vi

1 Centre of Gravity
An aircraft’s centre of gravity is the datum from which all calculation of stability comes;
therefore defining an aircraft’s most extreme centre of gravity limits is one of the most
important parts of designing one. If a consumer was to load an aircraft such that the centre of
gravity fell outside the fore or aft limits it could not fly in a stable condition therefore the first
stage in analysing the stability of an aircraft is to find these limits, a process of computing and
analysing the centre of gravity variation for different load cases and conditions that the design
requires. The calculation of the centre of gravity of an aircraft requires only the weight and
location of each component, for a small general aviation aircraft where component weights
are relatively similar each component must be considered as each effect the centre of gravity
greatly. For a larger transport aircraft relatively light components may be omitted during the
initial design stages, however once the components weight is defined it must be considered
due to the importance of the aircraft centre of gravity.
𝑥𝑥𝑐𝑐𝑐𝑐 =
∑ 𝑚𝑚𝑖𝑖 𝑥𝑥𝑐𝑐𝑐𝑐𝑖𝑖
𝑛𝑛
𝑖𝑖=1
∑ 𝑚𝑚𝑖𝑖
𝑛𝑛
𝑖𝑖=1
Equation 1 – Centre of Gravity Equation - [1]
The calculation of aircraft centre of gravity is completed using Equation 1; this equation can
be manipulated to calculate the centre of gravity in the y and z axis also. By calculating the
centre of gravity for the aircraft the desired range for the centre of gravity can be found and
therefore the design of the tail can begin. It can be seen however that an aircraft’s centre of
gravity will change as the weight of the components it is made up of change, fuel is an
example of this, as fuel is used through the flight the total weight of the aircraft lowers; this in
turn changes the centre of gravity. If this change is not accounted for this can cause the
aircrafts centre of gravity to move outside the allowed limit and the stability of the aircraft to be
compromised mid-flight, the change can be addressed in two ways however, as fuel is a liquid
and relatively stable it can be pumped around the aircraft to compensate for the change in
centre of gravity or the aircraft can be designed to account for all changes in aircraft weight
throughout the flight envelope.
1.1 Centre of Gravity Analysis
To start the design process information gathered during the initial research stages is input into
a table, Table 1; this table serves as the foundation of the centre of gravity analysis.
Whenever a component is updated or changed the table must be updated to account for this,
the main data required is the location and weight of each component.

Component X (m) Y (m) Z (m) Weight (kg)
Pilot 2.0438 0.3200 1.143 100.00
Co-Pilot 2.0438 -0.3200 1.143 100.00
LH Seat 2.0438 0.3200 1.143 30.00
RH Seat 2.0438 -0.3200 1.143 30.00
LH Wing 3.0950 0.0000 2.2301 24.927
RH Wing 3.0950 0.0000 2.2301 24.927
LH Landing Gear 3.2000 0.0000 0.6000 25.00
RH Landing Gear 3.2000 0.0000 0.6000 25.00
Nose Landing Gear 0.5000 0.0000 0.5000 15.00
Fuel Source 3.3000 0.0000 1.3000 60.00
Electrical Engine 0.2600 0.0000 1.3000 50.00
Propellor 0.1000 0.0000 1.3000 10.00
Main Spar 3.0950 0.0000 2.2301 40.00
Rear Spar 3.0950 0.0000 2.2301 20.00
Keel 4.0000 0.0000 2.2301 40.00
Horizontal Tail Main Spar 6.2680 0.0000 2.2301 10.00
Horizontal Tail Rear Spar 6.5000 0.0000 2.2301 5.00
Vertical Tail Main Spar 6.2680 0.0000 2.4000 10.00
Vertical Tail Rear Spar 6.5000 0.0000 2.4000 5.00
Rudder 6.6000 0.0000 2.4000 2.00
Aileron 3.9034 0.0000 2.2301 4.00
Flap 3.9034 0.0000 2.2301 4.60
Elevator 6.6000 0.0000 2.2301 3.00
Cockpit Frame 0.9642 0.0000 1.1000 60.00
Payload 2.2500 0.0000 1.4875 50.00
Table 1 – Component Centre of Gravity Analysis
Equation 1 is then applied to the table to calculate the centre of gravity in each axis, as shown
in Table 2 each of the components weights has been multiplied by gravitational acceleration
to give weight in newton’s, this is then multiplied by the distance in each axis to produce a
moment in the x, y and z axis for each component in a more conventional format.
Component Weight (N) Moment X (Nm) Moment Y (Nm) Moment Z (Nm)
Pilot 981.0000 2004.9678 313.9200 1121.6774
Co-Pilot 981.0000 2004.9678 -313.9200 1121.6774
LH Seat 294.3000 601.4903 94.1760 336.5032

RH Seat 294.3000 601.4903 -94.1760 336.5032
LH Wing 244.5339 756.8323 0.0000 545.3350
RH Wing 244.5339 756.8323 0.0000 545.3350
LH Landing Gear 245.2500 784.8000 0.0000 147.1500
RH Landing Gear 245.2500 784.8000 0.0000 147.1500
Nose Landing Gear 147.1500 73.5750 0.0000 73.5750
Fuel Source 588.6000 1942.3800 0.0000 765.1800
Electrical Engine 490.5000 127.5300 0.0000 637.6500
Propellor 98.1000 9.8100 0.0000 127.5300
Main Spar 392.4000 1214.4780 0.0000 875.0912
Rear Spar 196.2000 607.2390 0.0000 437.5456
Keel 392.4000 1569.6000 0.0000 875.0912
Horizontal Tail Main Spar 98.1000 614.8908 0.0000 218.7728
Horizontal Tail Rear Spar 49.0500 318.8250 0.0000 109.3864
Vertical Tail Main Spar 98.1000 614.8908 0.0000 235.4400
Vertical Tail Rear Spar 49.0500 318.8250 0.0000 117.7200
Rudder 19.6200 129.4920 0.0000 47.0880
Aileron 39.2400 153.1694 0.0000 87.5091
Flap 45.0868 175.9915 0.0000 100.5480
Elevator 29.4300 194.2380 0.0000 65.6318
Cockpit Frame 588.6000 567.54 0.0000 647.4600
Payload 490.5000 1103.6250 0.0000 729.6188
Table 2 - Component Moment Analysis
Longitudinal Stability
Weight
(N)
Moment X
(Nm)
Centre of in X
(m)
7342.29 18032.2833 2.4559
Lateral Stability
Weight
(N)
Moment Y
(Nm)
Centre of in Y
(m)
7342.29 0.0000 0.0000
Vertical Stability
Weight
(N)
Moment Z
(Nm)
Centre of in Z
(m)
7342.29 10452.1691 1.4236
Table 3 - Aircraft Centre of Gravity
For Table 1 and Table 2 the data gave an output of Table 3, however the datum and axis
system being used must be noted, if an inconsistent system is used the centre of gravity

location may be incorrect. In this case the system used took the origin at the foremost point of
the aircraft, through the centreline at ground level. Each payload is considered in the same
way as the aircraft components using a table such as Table 1, and the centre of gravity is
calculated again using Equation 1. This variation must also consider that payloads can be
placed in various positions and therefore all positions in which the payload can be placed
must be analysed and designed for, or the analyser must produce a document showing the
safe limits for aircraft loading to ensure stable flight. When undertaking the analysis the
designer must also consider the impact upon the performance of the aircraft, a relatively
heavy load will lower the available weight of fuel and therefore reduce range and endurance
whereas a relatively light load may not limit the amount of fuel which may instead be limited
by fuel storage volume. These two design stages must interact to find the most efficient and
effective use of aircraft weight to fit the design specification and design limits.
Status XCG (m) YCG (m) ZCG (m) Weight (kg)
Two Pilots Full Baggage 2.4559 0.0000 1.4236 748.45
One Pilot Full Baggage 2.5195 0.0493 1.4668 648.45
Two Pilots No Baggage 2.4707 0.0000 1.4190 698.45
One Pilot No Baggage 2.5420 0.0000 1.4650 598.45
Empty Aircraft 2.5519 0.0000 1.5610 438.45
Table 4 - Load Considerations
Figure 1- Centre of Gravity Variation in Longitudinal Axis
0.0000
0.5000
1.0000
1.5000
2.0000
2.5000
0.0000 1.0000 2.0000 3.0000 4.0000 5.0000 6.0000 7.0000
Z(m)
X (m)
One Pilot Full Baggage Two Pilots Full Baggage Two Pilots No Baggage
One Pilot No Baggage x0 xn
Empty Aircraft

2 Longitudinal Stability
Longitudinal stability is the stability in the XZ, or longitudinal axis of the aircraft. The main
effectors upon longitudinal stability are the centre of gravity, aerodynamic centre and
horizontal stabiliser. The horizontal stabiliser is a second lifting device used to offset the
moment created by the wings lift about the centre of gravity. This infers that if the wing centre
of lift is forward of the centre of gravity the horizontal stabiliser will produce lift in an upwards
direction for an aft mounted horizontal stabiliser, tail, or lift in a downwards direction for a fore
mounted horizontal stabiliser, canard and vice versa. For a centre of lift that is far from the
centre of gravity a larger moment is produced by the wing, therefore a larger restoring force
would be required by the horizontal stabiliser, therefore having the centre of gravity close to
the wing centre of lift or aerodynamic centre is a more desirable condition as it will minimise
horizontal stabiliser size and weight, therefore reducing cost.
2.1 Longitudinal Static Stability
2.1.1 Pitching Moment
Longitudinal stability is defined as; “the tendency of a body (or system) to return to equilibrium
when disturbed.” [2]. The moment created by the wing aerodynamic centre upon the centre of
gravity of the aircraft is called the pitching moment or 𝐶𝐶𝑚𝑚𝑚𝑚𝑚𝑚, as stated this is negated by the
horizontal stabiliser making the aircraft longitudinally statically stable. Therefore for straight
and level, steady flight the pitching moment must be equal to 0.
𝐶𝐶𝑚𝑚𝑚𝑚𝑚𝑚 = 𝐶𝐶𝑚𝑚0 + 𝐶𝐶𝐿𝐿(ℎ − ℎ0) − 𝐶𝐶𝐿𝐿
′
𝑉𝑉�ℎ
Equation 2 - Pitching Moment Equation - [2]
𝑉𝑉� = �
𝑙𝑙𝑆𝑆′
𝑐𝑐𝑐𝑐
�
Equation 3 - Volume Coefficient Equation - [2]
It can therefore be seen from Equation 2 that to have a pitching moment of 0 the pitching
moment created by the horizontal stabiliser must be equal to the pitching moment about the
aerodynamic centre, 𝐶𝐶𝑚𝑚0, and the pitching moment created by the difference in centre of
gravity and the aerodynamic centre, where the aircrafts centre of gravity and aerodynamic
centre are denoted by ℎ and ℎ0 respectively. This is the distance of each with respect to the
wings leading edge at mean aerodynamic chord, non-dimensionalised by the mean
aerodynamic chord. It can be inferred that the horizontal stabiliser therefore must be large
enough so that it produces a suitable restoring force; this is shown by the horizontal stabiliser
volume coefficient or 𝑉𝑉�ℎ, calculated using Equation 3.

2.1.2 Stabiliser Moment Arm
From the analysis of the centre of gravity the designing of the horizontal stabiliser can begin,
initially data from the wing and data from the centre of gravity analysis is used alongside the
aircraft design to find the key dimensions, of which the most important are the aerodynamic
centre of the wing, centre of gravity and horizontal stabiliser arm. The centres of gravity
parameters are available from previous analysis; however the wing aerodynamic centre must
be found using a combination of aerofoil data and wing analysis. For a wing the aerodynamic
centre is generally located at 25% of the mean aerodynamic chord however it be found in
aerofoil summary books such as Theory of Wing Sections by [3], this measurement along with
the centre of gravity are non-dimensionalised by the mean aerodynamic chord, the horizontal
stabiliser arm is designed through iteration and physical limitation of the aircraft and design
specification.
Figure 2 - Aircraft Dimensions
From this process the tail arm is chosen to be 2.730m placing it at 4.849m from the nose of
the aircraft, using an analysis of existing stable aircraft of this type it is found that for a light
general aviation aircraft 𝑉𝑉�ℎis typically 0.3. [1]. Using this value an area for the tail is found, this
area is important as lift is a function of area as shown in Equation 4.
𝐿𝐿 =
1
2
𝜌𝜌𝑉𝑉2
Equation 4 - Lift Equation - [2]
2.1.3 Aerofoil Selection
These values are then input into Equation 2, the output being a required tail lift coefficient or
𝐶𝐶𝐿𝐿
′
of -0.179 for cruise; this value allows the designer to fully design the remaining parameters
of the horizontal stabiliser. First an aerofoil section must be chosen for the horizontal stabiliser
as it is a lifting surface. There are several given parameters when designing this lifting
surface; the aerofoil must be symmetrical, this is because it will need to counter pitching
moments both nose up and nose down, it is also desirable for the stabiliser aerofoil to have
no pitching moment at its aerodynamic centre which is a feature of all symmetrical aerofoils. It
ℎ
𝑙𝑙ℎ =2.73017
ℎ0

is also desirable to have as low a minimum coefficient of drag ,or 𝐶𝐶𝑑𝑑𝑑𝑑𝑑𝑑𝑑𝑑, and as high a stall
angle, or 𝛼𝛼𝑠𝑠, as possible. NACA profile 0009 is chosen in line with these aims and the data for
the aerofoil is taken, Table 5.
Profile Cdmin Cm0 αS Flaps 0° ClMAX Clα
0009 0.005 0 13 1.3 6.7
Table 5 – NACA 0009 Aerofoil Data
2.1.4 Horizontal Stabiliser Design
A value for the aspect ratio of the tail is estimated at around
2
3
the aspect ratio of the
horizontal stabiliser, now using MatLab the lift produced by the horizontal stabiliser can be
analysed. This is done using a script utilising Pradtl’s lifting line theory. Pradtl’s lifting line
theory is generally accurate and offers an excellent insight into how a lifting surface will
perform for a given set of parameters.
N = 9; % (number of segments-1)
b = sqrt(AR*S); % tail span
MAC = S/b; % Mean Aerodynamic Chord
Croot = (1.5*(1+lambda)*MAC)/(1+lambda+lambda^2); % root chord
alpha=a_h+alpha_twist:-alpha_twist/(N-1):a_h; % segment's angle of attack
c = Croot * (1 - (1-lambda)*cos(theta)); % Mean Aerodynamics chord at each segment
mu = c * a_2d / (4 * b);
for i=1:N
for j=1:N
B(i,j)=sin((2*j-1) * theta(i)) * (1+(mu(i) *(2*j-1))/sin(theta(i)));
end
end
A=Btranspose(LHS);
for i = 1:N
sum1(i) = 0;
sum2(i) = 0;
for j = 1 : N
sum1(i) = sum1(i) + (2*j-1) * A(j)*sin((2*j-1)*theta(i));
end
end
CL_tail = pi * AR * A(1)
Code 1 - MatLab Tail Lift Script - [1] – (Modified by Benjamin James Johnson)

S = 2.393079258
AR = 4
lambda = 0.850001
alpha_twist = -0.000001
a_h = -2.23022134
a_2d = 4.895475362
alpha_0 = 0
Table 6 - MatLab Script Inputs
By utilising this tool a coefficient of lift for the horizontal stabiliser is found of -1.328, this is
lower than required and will negatively impact the static stability of the aircraft causing too
large a restoring moment and not allowing the aircraft to return to equilibrium. Through
changing the incidence the horizontal stabiliser is found to produce the required 𝐶𝐶𝐿𝐿 at −3.02°
.
S 2.393079 m2
AR 3.857143
λ
0.85
°
αt
0
°
i
-3.02
°
b
3.4228
m
c
0.6992
m
croot
0.7542
m
ctip
0.64107
m
Table 7 - Horizontal Stabiliser Parameters
It must be noted that the sweep angle and taper ratio of the horizontal stabiliser are selected
to be the same as that of the wing, this is to ensure similar benefits of this lifting surface as
that of the wing, reducing the bending moment and structure of the tail. However there is no
twist upon the horizontal stabiliser, this is because there is no requirement for elliptical lift
distribution across the stabiliser as it should never stall and therefore tip stall is not a problem.
2.1.5 Horizontal Stabiliser Vertical Position
Now the effect of the wing upon the horizontal stabiliser must be analysed, the aircraft is
chosen to have a high wing and conventional tail, this however means that the horizontal tail
will be in the wake region of the wing causing it to lose effectiveness at the stall, this must be
avoided because, as the aircraft loses lift it starts to slip backwards as the horizontal stabiliser
can no longer pitch the nose down, as the velocity of the slip increases the stabiliser is pulled
further down into the stall and the aircraft cannot be recovered. To ensure this condition never

occurs the horizontal stabiliser must be at least partially out of the wing wake region at the
wing stall angle, it must have a lower aspect ratio than the wing and it must produce a nose
down pitching moment at wing stall.
𝑙𝑙ℎ tan(𝛼𝛼𝑠𝑠 − 𝑖𝑖𝑤𝑤 − 3) > 𝑧𝑧𝑡𝑡 > 𝑙𝑙ℎ tan(𝛼𝛼𝑠𝑠 − 𝑖𝑖𝑤𝑤 + 3)
Equation 5 - Horizontal Stabiliser Height Equation - [1]
Equation 5 is therefore used to ensure that at wing stall the horizontal stabiliser is within the
required region to maintain effectiveness throughout the stall. It is found that the horizontal
stabiliser must be located between 0.732m and 0.432m above the wing chord line.
2.1.6 Horizontal Stabiliser Setting Angle
Although the horizontal stabiliser is within the requirement for deep stall elimination it will not
be outside the wing downwash region, this region is created by the wing trailing edge vortices
and causes an effect upon the airflow behind the wing, and therefore the airflow on the
horizontal stabiliser. This effect changes the lift generated by the horizontal stabiliser but can
be accounted for by setting the horizontal stabiliser to produce the required lift coefficient for
static stability.
𝛼𝛼′
= 𝛼𝛼 �1 −
𝛿𝛿𝜀𝜀
𝛿𝛿𝛿𝛿
� + 𝛿𝛿 − 𝜀𝜀0
Equation 6 - Horizontal Stabiliser Incidence - [2]
The effect of the wing can then be calculated by Equation 6 and then negated using the
incidence in Table 7 to calculate the setting angle.
2.1.7 Stick Fixed Static Longitudinal Stability of Aircraft
Finally for the horizontal stabiliser design the static stability for the entire aircraft must be
analysed, throughout the design process each stage has been aimed at ensuring the final
product will be stable, however it must be proven analytically once all parameters are
available. For static stability it is required that the aircraft return to equilibrium after a
disturbance, mathematically this can be shown as
𝛿𝛿𝐶𝐶 𝑚𝑚𝑚𝑚𝑚𝑚
𝛿𝛿𝐶𝐶𝐿𝐿
< 0.
𝛿𝛿𝐶𝐶𝑚𝑚𝑚𝑚𝑚𝑚
= −ℎ0 − ℎ + 𝑉𝑉�
𝑎𝑎1
′
𝑎𝑎1
�1 −
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
�
Equation 7 - Stick Fixed Longitudinal Stability Equation - [2]

Where 𝑎𝑎1 is the lift curve slope of the wing or,
𝐶𝐶𝐿𝐿
𝐶𝐶𝛼𝛼
and 𝑎𝑎1
′
is the lift curve slope of the horizontal
stabiliser or,
𝐶𝐶𝐿𝐿
′
𝐶𝐶𝛼𝛼
. For the entire aircraft it is found that
𝛿𝛿𝐶𝐶 𝑚𝑚𝑚𝑚𝑚𝑚
= −1.07 … this fits into the
requirement for longitudinal static stability.
2.1.8 Neutral Point Analysis
As discussed previously the centre of gravity can change, meaning that the effect of the wing
aerodynamic moment about the centre of gravity will also change and therefore the required
restoring moment by the tail will change, this requirement for stability is called elevator angle
to trim and will be discussed in section 2.2.10. As the range for centre of gravity is increased
so too is the stability in the defining axis this however means that to control the aircraft larger
control inputs are needed which require larger control surfaces or more force upon the control
surface meaning they require more structure creating other design challenges, this range is
the stability margin. This margin is bounded from the foremost centre of gravity location to the
aircraft neutral point or ℎ𝑛𝑛, this point is the aft-most point at which static stability is possible,
as can be seen on Figure 1 all load cases are in front of the aircraft neutral point therefore the
aircraft is stable in all flight phases.
ℎ𝑛𝑛 = ℎ0 + 𝑉𝑉�ℎ
𝑎𝑎1
′
𝑎𝑎1
�1 −
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
�
Equation 8 - Neutral Point Equation - [2]
The neutral point gives the aircraft loadmaster a limit to work to, however it is not as helpful
for stability analysis, more useful is the stability margin or 𝐻𝐻𝑛𝑛. The stability margin is the range
from the aircraft centre of gravity to the neutral point, and due to the requirement of Equation
8, 𝐻𝐻𝑛𝑛 > 0 for statically stable flight. Completing these calculations yields a stability margin of
0.260 and a neutral point at 𝑥𝑥 = 2.85966 …m or 55.2% mean aerodynamic chord.
2.2 Longitudinal Dynamic Stability
An aircraft flying in equilibrium that experiences a longitudinal disturbance may experience
two types of motion, Phugoid and Short Period Pitching Oscillation. For an aircraft to be
longitudinally dynamically stable it must be positively damped in both motions, for the aircraft
to have good flying qualities, the combination of damping and natural frequency must be
conducive to reducing the workload upon the pilot. Longitudinal dynamic stability can be
approximated from the aircraft longitudinal equations of motion by considering the effect they
have upon the aircrafts flight.
�
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑋𝑋𝑢𝑢
𝑚𝑚
� 𝑢𝑢 + �−
𝑋𝑋𝑤𝑤
𝑚𝑚
� 𝑤𝑤 + 𝑔𝑔𝑔𝑔 =
∆𝑇𝑇
𝑚𝑚
Equation 9 - X Equation - [4]

�−
𝑍𝑍𝑢𝑢
𝑚𝑚
� 𝑢𝑢 + �
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑍𝑍𝑤𝑤
𝑚𝑚
� 𝑤𝑤 + �−𝑈𝑈
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜃𝜃 = 0
Equation 10 - Z Equation - [4]
�−
𝑀𝑀𝑤𝑤
𝐵𝐵
−
𝑀𝑀𝑤𝑤̇
𝐵𝐵
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝑤𝑤 + �
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝑀𝑀𝑞𝑞
𝐵𝐵
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜃𝜃 =
𝑀𝑀𝜂𝜂
𝐵𝐵
𝜂𝜂
Equation 11- Pitching Moment Equation - [4]
2.2.1 Phugoid Motion
Phugoid motion is described as; “a low frequency, lightly damped oscillation characterised by
a change in forward velocity and pitch angle at nearly constant incidence.” [5]. The Phugoid
motion is characterised by low damping and a long period, inferring a gentle but continuous
motion. When the aircraft experiences Phugoid motion it begins to oscillate, pitching up and
increasing airspeed as it is initially disturbed, this then manifests as the aircraft climbs and
loses speed it begins to pitch down until it reaches its maximum amplitude. As the aircraft
reaches maximum amplitude it reaches minimum speed, fully pitching nose down, the weight
component of the aircraft then takes precedence and the aircraft flies towards the ground
increasing its airspeed, this increase in airspeed causes the lift to increase over the wing and
the aircraft to pitch nose up. As the aircraft reaches its minimum amplitude it reaches its
maximum speed, this causes the aircraft to begin to climb again and repeat the cycle. As the
damping on this motion increases the length of period for this motion decreases along with
the amplitude of this motion.
2.2.2 Phugoid Approximation
Phugoid motion can be approximated from the aircraft longitudinal equations of motion, as it
is deemed a change in forward velocity, 𝑢𝑢, and pitch angle, 𝜃𝜃, with little to no change in
incidence, 𝑤𝑤, it can be calculated by removing the pitching moment equation, Equation 11,
and setting 𝑤𝑤 = 0, it must also be noted that there is assumed no pilot input. This leaves only
two equations:
�
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑋𝑋𝑢𝑢
𝑚𝑚
� 𝑢𝑢 + 𝑔𝑔𝑔𝑔 = 0
Equation 12 - Phugoid Approximation X Equation
�−
𝒁𝒁𝒖𝒖
𝒎𝒎
� 𝒖𝒖 + �−𝑼𝑼
𝜹𝜹
𝜹𝜹𝜹𝜹
� 𝜽𝜽 = 𝟎𝟎
Equation 13 - Phugoid Approximation Z Equation
It can then be assumed that the solution of each variable is of the form; 𝑥𝑥 = 𝑥𝑥0 𝑒𝑒 𝜆𝜆𝜆𝜆
where 𝑥𝑥0 is
the value of 𝑥𝑥 when 𝑡𝑡 = 0, from this assumption each variable can be replaced and the

equations can be analysed by taking the determinate and calculating the characteristic
equation.
�
𝜆𝜆 −
𝑋𝑋𝑢𝑢
𝑚𝑚
𝑔𝑔
−
𝑍𝑍𝑢𝑢
𝑚𝑚
−𝑈𝑈𝑈𝑈
� = 0
Equation 14 - Phugoid Approximation Determinant
𝜆𝜆2
+ �−
𝑋𝑋𝑢𝑢
𝑚𝑚
� 𝜆𝜆 +
𝑔𝑔
𝑈𝑈
�−
𝑍𝑍𝑢𝑢
𝑚𝑚
� = 0
Equation 15 - Phugoid Approximation Characteristic Equation
From this characteristic equation the main features of the motion can be found by comparing
it with the general characteristic equation, Equation 16.
𝜆𝜆2
+ 2𝜁𝜁𝜔𝜔𝑛𝑛 + 𝜔𝜔𝑛𝑛
2
= 0
Equation 16 - General Characteristic Equation - [4]
By comparing the approximated solution with the general equation the damping and natural
frequency of the motion can be found, and therefore the period,𝑇𝑇, and time to half
amplitude,𝑡𝑡1
2
, can be calculated using Equation 17 and Equation 18 respectively.
𝑇𝑇 =
2𝜋𝜋
��(2𝜁𝜁𝜔𝜔𝑛𝑛) − (4𝜔𝜔𝑛𝑛
2�
Equation 17 - Motion Period Equation
𝑡𝑡1
2
=
− ln 2
(−𝜁𝜁𝜔𝜔𝑛𝑛)
Equation 18 - Motion Time to Half Amplitude Equation
2.2.3 Aerodynamic Derivatives
To calculate these parameters the relevant aerodynamic derivatives must also be calculated,
an aerodynamic derivative is: “The rate of change of any aerodynamic force or aerodynamic
moment with respect to one of the disturbance quantities, all other disturbances being
assumed zero.” [6] The aerodynamic derivatives are partial derivatives of all forces and
moments the aircraft experiences with respect to the disturbance values, this means that
there are many aerodynamic derivatives, however of the 60 possible derivatives almost half of
them can be assumed as zero due to aircraft symmetry.
In the longitudinal axis there are only 8 critical derivatives of a possible 15, shown in Table 8
are the derivatives required to calculate the Phugoid approximation in 2.2.4 and Short Period
Pitching Oscillation approximation in 2.2.7. For calculation of the Phugoid the flying conditions

must be accounted for, as can be seen in Table 8 each aerodynamic derivative relies upon
the variable 𝑄𝑄 or dynamic pressure. Dynamic pressure is the product of density and velocity
measuring the kinetic energy per unit volume of the fluid, in this case the air, and changes
with atmospheric conditions and speed. Therefore for a full stability analysis of an aircraft, all
flight phases must be taken into account and all flight conditions, however of all conditions;
max speed/max altitude, cruise and take-off are the most critical flight phases and max fore
and max aft centre of gravity positions are the second set of critical phases, therefore if the
stability is analysed in these flight phases it can be assumed to be covering all others.
Derivative Calculation
𝑋𝑋𝑢𝑢 −
�𝐶𝐶𝐷𝐷𝑢𝑢
+ 2𝐶𝐶𝐷𝐷0
�𝑄𝑄𝑄𝑄
𝑚𝑚𝑢𝑢0
𝑍𝑍𝑢𝑢 −
�𝐶𝐶𝐿𝐿𝑢𝑢
+ 2𝐶𝐶𝐿𝐿0
�𝑄𝑄𝑄𝑄
𝑚𝑚𝑢𝑢0
𝑍𝑍𝑤𝑤 −
�𝐶𝐶𝐿𝐿𝛼𝛼
+ 2𝐶𝐶𝐷𝐷0
�𝑄𝑄𝑄𝑄
𝑚𝑚𝑢𝑢0
𝑀𝑀𝑤𝑤 −
𝐶𝐶𝑚𝑚𝛼𝛼
𝑄𝑄𝑄𝑄
𝑢𝑢0 𝐼𝐼𝑦𝑦𝑦𝑦
𝑀𝑀𝑤𝑤̇ −
𝐶𝐶𝑚𝑚𝛼𝛼̇
𝑄𝑄𝑄𝑄𝑐𝑐2
2𝑢𝑢0
2
𝐼𝐼𝑦𝑦𝑦𝑦
𝑍𝑍𝛼𝛼 𝑢𝑢0 𝑍𝑍𝑤𝑤
𝑀𝑀𝛼𝛼 𝑢𝑢0 𝑀𝑀𝑤𝑤
𝑀𝑀𝛼𝛼̇ 𝑢𝑢0 𝑀𝑀𝑤𝑤̇
𝑀𝑀𝑞𝑞 −
𝐶𝐶𝑚𝑚𝑞𝑞
𝑄𝑄𝑄𝑄𝑐𝑐2
2𝑢𝑢0 𝐼𝐼𝑦𝑦𝑦𝑦
Table 8 - Longitudinal Aerodynamic Derivatives
2.2.4 Phugoid Calculation
For the aircraft, the parameters found through the horizontal stabiliser design, centre of
gravity analysis and aerodynamic analysis are applied and the Phugoid can be approximated.
For the initial analysis of the Phugoid motion the aircraft will be set in cruise condition, this is
because the aircraft and pilot will spend most of its flying time in this condition therefore pilot
workload must be as low as possible. The aircraft cruise conditions are shown in Technical
Specification and the aircraft longitudinal aerodynamic derivatives are shown in Table 8
therefore the analysis of the Phugoid motion can begin, from the analysis of the Phugoid
motion the following results are obtained:

𝜔𝜔𝑛𝑛 0.277889745
𝜁𝜁 0.114498306
𝑇𝑇 11.38002
𝑡𝑡1
2
21.78481518
Table 9 - Phugoid Approximation Results
2.2.5 Short Period Pitching Oscillation
Short period pitching oscillation or SPPO is described as; “a short period heavily damped
oscillation characterised by changes in pitch angle and incidence … with little variation in
forward speed”. The SPPO is characterised by very high damping and a short period,
inferring a sharp and short movement with no continuous motion. When the aircraft
experiences SPPO motion it tends to pitch quickly, the pitching movement is restored by the
horizontal stabiliser, however as the aircraft returns to its equilibrium position it will still have a
component of pitch rate causing the aircraft to overshoot equilibrium, if underdamped this
motion will continue. Along with this component the horizontal stabiliser adds a damping
effect, the pitching moment causing an up or downwash on the horizontal stabiliser increasing
the incidence on lifting surface and hence a lifting moment is caused in the opposite direction
to the movement. If the damping on this component is not suitable the horizontal stabiliser
can oscillate about the equilibrium position, causing the tail section to rise and fall, if this were
to happen at take-off or landing the tail could strike the ground causing unfavourable
circumstances.
2.2.6 Short Period Pitching Oscillation Approximation
Similar to the Phugoid motion the SPPO can be approximated from the aircraft longitudinal
equations of motion, as it is deemed as a change in 𝜃𝜃 and 𝜔𝜔 with little or no change in 𝑈𝑈 it
can be calculated by removing the X equation, Equation 9, and setting 𝑢𝑢 = 0, again assuming
no pilot input. This leaves only two equations:
�
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑍𝑍𝑤𝑤
𝑚𝑚
� 𝑤𝑤 + �−𝑈𝑈
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜃𝜃 = 0
Equation 19 - SPPO Approximation Z Equation
�−
𝑀𝑀𝑤𝑤
𝐵𝐵
−
𝑀𝑀𝑤𝑤̇
𝐵𝐵
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝑤𝑤 + �
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝑀𝑀𝑞𝑞
𝐵𝐵
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜃𝜃 = 0
Equation 20 - SPPO Approximation Pitching Moment Equation
Again the variables can be assumed as; 𝑥𝑥 = 𝑥𝑥0 𝑒𝑒 𝜆𝜆𝜆𝜆
, replaced and the determinate can be
taken to calculate the characteristic equation for the SPPO motion.

�
𝜆𝜆 −
𝑍𝑍𝑤𝑤
𝑚𝑚
−𝑈𝑈
−
𝑀𝑀𝑤𝑤
𝐵𝐵
− 𝜆𝜆
𝑀𝑀𝑤𝑤̇
𝐵𝐵
𝜆𝜆 −
𝑀𝑀𝑞𝑞
𝐵𝐵
� = 0
Equation 21 - SPPO Approximation Determinant
𝜆𝜆2
+ �−
𝑀𝑀𝑞𝑞
𝐵𝐵
−
𝑈𝑈𝑈𝑈𝑤𝑤̇
𝐵𝐵
−
𝑍𝑍𝑤𝑤
𝑚𝑚
� 𝜆𝜆 + �−
𝑈𝑈𝑈𝑈𝑤𝑤
𝐵𝐵
+
𝑍𝑍𝑤𝑤 𝑀𝑀𝑞𝑞
𝑚𝑚𝑚𝑚
� = 0
Equation 22 - SPPO Approximation Characteristic Equation
Then again by comparing Equation 22 to Equation 16 the period and time to half amplitude
can be calculated.
2.2.7 Short Period Pitching Oscillation Calculation
For the aircraft, the parameters found through the horizontal stabiliser design, centre of
gravity analysis and aerodynamic analysis are applied and the SPPO can be approximated.
For the initial analysis of the SPPO motion the aircraft will be set in cruise condition for the
reasons stated in section 2.2.4. The aircraft cruise conditions are shown in Technical
Specification and the aircraft longitudinal aerodynamic derivatives are shown in Table 8
therefore the analysis of the SPPO motion can begin, from the analysis of the SPPO motion
the following results are obtained:
𝜔𝜔𝑛𝑛 6.516564
𝜁𝜁 0.439088
𝑇𝑇 0.536587
𝑡𝑡1
2
0.242245
Table 10 - SPPO Approximation Results
2.2.8 Flying Characteristics
Given the values in Table 9 and Table 10 the aircraft can be compared to a flying
characteristics table such as Table 11, from this table a range of values is given for each
motion and a score for the flying characteristic can be found, this level indicates the workload
upon the pilot for a given flying characteristic’s parameters. If the aircraft being analysed does
not fall within the required limits then either redesign or stability augmentation may be
required.
As can be seen by comparing the three tables the aircraft is a level 1 in both the Phugoid and
SPPO modes it can also be noted that the aircraft is stable in both modes, this can be verified
by the positive damping constant for both modes.

Phugoid Mode
Level 1 ζ > 0.04
Level 2 ζ > 0
Level 3 T2 > 55
Short Period Mode
Category A and C Category B
ζ ζ ζ ζ
Level min max min max
1 0.35 1.3 0.3 2
2 0.25 2 0.2 2
3 0.15 --- 0.15 ---
Table 11 - Longitudinal Flying Characteristics - [7]
2.2.9 Elevator Design
The elevators are the control surface used to manoeuvre the aircraft in the pitch about the
lateral axis; they are generally positioned on the trailing edge of the horizontal stabiliser, the
elevator design is dictated by the elevator trim requirement. The horizontal stabiliser has been
designed to keep the aircraft stable in the cruise condition with the most extreme centre of
gravity, however as discussed in section 1Error! Reference source not found. and section
2.1 the centre of gravity of the aircraft changes over the flight, this results in the horizontal
stabiliser having to provide different lift values through the flight, as the size of the horizontal
stabiliser on a conventional aircraft cannot be changed through the flight an elevator is
employed to change the horizontal stabiliser lift. Along with the trim requirement a more
critical employment of the elevator is pitch control at low speeds such as at take-off and
landing, the aircraft’s elevator must allow it to change the aircraft’s pitch at take-off to allow
take-off rotation and to stop ground looping.
Thus initially the design begins with calculating all aircraft moments about the main gear; this
gives a required lift force for the elevator to be able to achieve to rotate the aircraft, from this
lift force a desired lift coefficient for the elevator can be calculated and therefore angle of
attack effectiveness is given. This can be analysed against Figure 4 and an elevator chord to
horizontal stabiliser chord can be found, this allows the designer to then specify a span for the
elevator, then through further MatLab analysis a suitable deflection angle can be found for the
elevator.
2.2.10 Elevator to Trim
After analysing the requirement for take-off rotation the trim condition must be considered,
again this can utilise tools such as MatLab, the basis of this analysis is Equation 7 but now
considering the elevator effect as well.
0 = 𝐶𝐶𝑚𝑚0
− 𝑉𝑉� 𝑎𝑎1
′ (𝛿𝛿 − 𝜀𝜀0) − 𝑉𝑉� 𝑎𝑎2
′
𝜂𝜂 − 𝐶𝐶𝐿𝐿(𝐻𝐻𝑛𝑛)
Equation 23 - Elevator Angle to Trim Equation - [2]

Equation 23 can be used to calculate the angle required by the elevator to maintain steady
flight and this can be plotted for various altitudes, Figure 3.
Figure 3 - Elevator Angle to Trim at Cruise Altitude
100 150 200 250 300 350 400 450 500
-10
-9
-8
-7
-6
-5
-4
-3
-2
-1
0
Speed (knot)
δE
(deg)
Most aft cg
Most forward cg

3 Lateral Stability
Lateral stability is the stability in the XY, or lateral axis of the aircraft. The main effectors upon
lateral stability are the centre of gravity, aerodynamic centre, thrust location and vertical
stabiliser. The vertical stabiliser is a third lifting device used to offset the moment created by
offset thrust about the centre of gravity, crosswind or prop rotation. The vertical tail is
designed to maintain directional stability in two critical situations, the first as previously
remarked is the crosswind condition most importantly at take-off and landing speed with a
maximum 90° crosswind, this condition is most critical for aircraft with propulsion mechanisms
along or very close to the centre line of the aircraft. The second critical situation is the one
engine inoperative condition, which is of increasing importance the further the propulsion
mechanisms is from the aircraft centreline. Therefore to reduce the criticality of these
situations firstly the aircraft side profile must be as small as possible as to reduce crosswind
effect, however this is not always practical as aircraft are designed to carry a payload and this
payload may need to be housed inside the fuselage. For the one engine inoperative condition
the propulsion mechanisms must be mounted as close to the centreline as possible as to
negate the moment created by only one about the aircraft centre of gravity, however for some
aircraft it is not practical or efficient to mount the engine inside or against the fuselage due to
the reduction in fuselage or wing space or the increase in fuselage to engine interference
drag.
3.1 Static Directional Stability
For the static directional stability it is generally intended by the designer that the aircraft will
be symmetrical along the longitudinal axis, meaning that any moment created by any part
along one side of the aircraft will be restored by the component on the other. This is an ideal
case but generally it can be applied even on aircraft where gear retraction is done one side at
a time or other such cases due to the ability of the vertical stabiliser to negate any temporary
effects upon static directional stability. However, the designer may not be able to effectively
reduce the effects of one engine inoperative conditions or crosswind, therefore the vertical
stabiliser is designed to negate these conditions. This implies that the static directional
stability derivative, Equation 24, must be positive as to return the aircraft to equilibrium.
𝐶𝐶𝑛𝑛 𝛽𝛽
= 𝐾𝐾𝑓𝑓1 𝐶𝐶𝐿𝐿 𝛼𝛼_𝑣𝑣
�1 −
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜂𝜂𝑣𝑣
𝑙𝑙𝑣𝑣 𝑆𝑆𝑣𝑣
𝑏𝑏𝑏𝑏
Equation 24 - Static Directional Stability Derivative Equation - [1]
3.1.1 Vertical Stabiliser Design
To size the vertical stabiliser an analysis of other aircraft is initially required, this analysis
allows the designer to choose a vertical tail coefficient for the aircraft, and by choosing a
similar value from a similar aircraft it can generally ensure that the final design will be stable.

A value of 𝑉𝑉�𝑣𝑣 = 0.02 similar to that of the Cessna 152 is chosen, and to reduce the structural
penalty of the tail the span of the vertical stabiliser is set at 1.4m to ensure the horizontal
stabiliser can be as close to its minimum allowable position to reduce heavy structure at the
top of the vertical stabiliser. This sets the vertical stabilisers parameters to those in Table 12
with the NACA 0009 profile being used again for commonality:
S 0.923018 m
2
AR 2.123468
Λ 70 °
b 1.4 m
c 0.659299 m
Table 12- Vertical Stabiliser Parameters
3.1.2 Static Directional Stability Derivative Calculation
With the design of the vertical stabiliser complete 𝐶𝐶𝑛𝑛 𝛽𝛽
can be calculated, the only unknown
being 𝐾𝐾𝑓𝑓1, this constant is the contribution of the aircraft fuselage and is generally between
0.65 and 0.85 [1] with larger fuselages contributing more. The vertical stabiliser efficiency can
be approximated at 𝜂𝜂𝑣𝑣 = 0.98 and the vertical stabiliser side wash gradient assumed as,
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
= 0. This gives a value of 𝐶𝐶𝑛𝑛 𝛽𝛽
= 0.071229 … therefore the aircraft is statically directionally
stable.
3.2 Lateral Dynamic Stability
An aircraft flying in equilibrium that experiences a lateral disturbance may experience three
types of motion, roll convergence, spiral mode and Dutch roll mode. For an aircraft to be
laterally dynamically stable it must be positively damped in all motions, for the aircraft to have
good flying qualities, the combination of damping and natural frequency must be conducive to
reducing the workload upon the pilot. Lateral dynamic stability can be approximated from the
aircraft lateral equations of motion by considering the effect they have upon the aircrafts flight.
�
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑌𝑌𝑣𝑣
𝑚𝑚
� 𝑣𝑣 − 𝑔𝑔𝜙𝜙 + 𝑈𝑈
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
= 0
Equation 25 - Sideslip Equation - [8]
−
𝐿𝐿𝑣𝑣
𝐴𝐴
𝑣𝑣 + �
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝐿𝐿𝑝𝑝
𝐴𝐴
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜙𝜙 −
𝐿𝐿𝑟𝑟
𝐴𝐴
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
=
𝐿𝐿𝜉𝜉
𝐴𝐴
𝜉𝜉 +
𝐿𝐿𝜁𝜁
𝐴𝐴
𝜁𝜁
Equation 26 - Roll Equation - [8]

−
𝑁𝑁𝑣𝑣
𝐶𝐶
𝑣𝑣 −
𝑁𝑁𝑝𝑝
𝐶𝐶
𝛿𝛿
𝛿𝛿𝛿𝛿
𝜙𝜙 + �
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝑁𝑁𝑟𝑟
𝐶𝐶
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜓𝜓 =
𝑁𝑁𝜉𝜉
𝐶𝐶
𝜉𝜉 +
𝑁𝑁𝜁𝜁
𝐶𝐶
𝜁𝜁
Equation 27 - Roll Moment Equation - [8]
Derivative Calculation
𝑌𝑌𝛽𝛽
𝑄𝑄𝑄𝑄𝐶𝐶𝑦𝑦 𝛽𝛽
𝑚𝑚
𝑌𝑌𝑝𝑝
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑦𝑦𝑝𝑝
2𝑚𝑚𝑢𝑢0
𝑌𝑌𝑟𝑟
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑦𝑦𝑟𝑟
2𝑚𝑚𝑢𝑢0
𝑌𝑌𝛿𝛿𝑟𝑟
𝑄𝑄𝑄𝑄𝐶𝐶𝑦𝑦𝛿𝛿𝛿𝛿
𝑚𝑚
𝑌𝑌𝛿𝛿𝐴𝐴
𝑄𝑄𝑄𝑄𝐶𝐶𝑦𝑦𝛿𝛿𝛿𝛿
𝑚𝑚
𝐿𝐿𝑝𝑝̇
𝑄𝑄𝑄𝑄𝑏𝑏2
𝐶𝐶𝑙𝑙𝑝𝑝
2𝑢𝑢0 𝐼𝐼𝑥𝑥𝑥𝑥
𝐿𝐿𝑟𝑟̇
𝐶𝐶𝑙𝑙𝑟𝑟
2𝑢𝑢0 𝐼𝐼𝑥𝑥𝑥𝑥
𝐿𝐿𝛽𝛽
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑙𝑙𝛽𝛽
𝐼𝐼𝑥𝑥𝑥𝑥
𝐿𝐿𝛿𝛿𝛿𝛿̇
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑙𝑙𝛿𝛿𝛿𝛿
𝐿𝐿𝛿𝛿𝛿𝛿̇
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑙𝑙𝛿𝛿𝛿𝛿
𝑁𝑁𝛽𝛽̇
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑛𝑛 𝛽𝛽
𝐼𝐼𝑧𝑧𝑧𝑧
𝑁𝑁𝑝𝑝
𝐶𝐶𝑛𝑛𝑝𝑝
2𝑢𝑢0 𝐼𝐼𝑧𝑧𝑧𝑧
𝑁𝑁𝑟𝑟
𝐶𝐶𝑛𝑛𝑟𝑟
2𝑢𝑢0 𝐼𝐼𝑧𝑧𝑧𝑧
𝑁𝑁𝛿𝛿𝛿𝛿̇
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑛𝑛𝛿𝛿𝛿𝛿
𝑁𝑁𝛿𝛿𝛿𝛿̇
𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑛𝑛𝛿𝛿𝛿𝛿
Table 13 - Lateral Aerodynamic Derivatives - [7]

3.2.1 Crosswind Requirement
The vertical stabiliser is also required to negate OEI and crosswind, due to the aircraft only
having one engine only the crosswind is analysed. The moment produced by a crosswind
about the centre of gravity must be negated by the vertical stabiliser arrangement, this
requirement is analysed by calculating the centre of the wetted side area and applying the
crosswind force at 90° to the fuselage centreline. It is found for the aircraft to be able to
maintain directional stability in a 20knot crosswind at take-off the vertical stabiliser must be
able to produce at least 321.03N of lifting force to counteract the moment created by the
crosswind.
3.2.2 Rudder Design
From these requirements to maintain directional stability the rudder can be designed, the
rudder controls the aircraft in the vertical axis, allowing the pilot to change heading by yawing
the aircraft. The rudder is also used in crosswind conditions to maintain heading. By analysing
the vertical stabiliser and these two conditions the rudder can be designed for safe flight at
the most critical conditions; these include take-off, landing and cruise flight phases with fore
and aft extreme centre of gravity positions. The condition of most importance for crosswind
performance is that at take-off, this condition is when the aircraft is travelling at its slowest
and therefore the vertical stabiliser and rudder are both at their least effective. For the aircraft
crosswind is identified as worst inhibitor of directional stability and therefore requires the
largest restoring moment from the vertical stabiliser. Initially the rudder is sized as a
proportion of the vertical stabiliser, in this case it is decided that the rudder will occupy 80% of
the vertical tail span and 30% of the vertical tail chord using this information a control surface
effectiveness value is selected from Figure 4.
Figure 4 - Control Surface Angle of Attack Effectiveness Parameter – [1]

𝐶𝐶𝑛𝑛𝛿𝛿𝛿𝛿
= −𝐶𝐶𝐿𝐿𝛼𝛼𝑣𝑣
𝑉𝑉�𝑣𝑣 𝜂𝜂𝑣𝑣 𝜏𝜏𝑅𝑅
𝑏𝑏𝑅𝑅
𝑏𝑏𝑣𝑣
Equation 28 - Rudder Control Derivative Equation - [1]
𝛿𝛿𝑟𝑟 =
𝑇𝑇𝑂𝑂𝑂𝑂𝑂𝑂 𝑌𝑌𝑇𝑇
−𝑄𝑄𝑄𝑄𝑄𝑄𝐶𝐶𝑛𝑛𝛿𝛿𝛿𝛿
�
Equation 29 - Required Rudder Deflection for OEI Equation - [1]
Using Equation 28 and the value taken from Figure 4 for the specified rudder, the rudder
control derivative can be calculated which can then be used in Equation 29 to calculate the
required deflection by the rudder to counteract the yawing moment of asymmetric thrust, the
speed used for this calculation is the minimum manoeuvre speed, typically 80%-100% of stall
speed. It is found for the aircraft that at a minimum manoeuvre speed of 36 knots a rudder
deflection of 23.2° is required to offset the crosswind, it is also noted that the minimum
manoeuvre speed is 18 knots lower than the lift off speed of the aircraft meaning that at take-
off, the most critical condition for crosswind operation the rudder will satisfy the control
requirements with a 25% safety margin.
3.2.3 Aileron Design
The ailerons are the control surface used to manoeuvre the aircraft in the roll about the
longitudinal axis; they are generally positioned on the trailing edge of the wing at the
outermost available position. The aileron design is dictated by the time to bank requirement,
this is the time allowed for the aircraft to roll through a certain angle within a required time.
Initially as in section 3.2.2 values for the aileron dimensions are chosen, in this case as the
high lift devices require 70% of the wing span and the fuselage requires around 12% of the
wing span the ailerons are chosen to take 40% of the wingspan, their positions is chosen from
70% to 90% of the half wingspan meaning that they have as high as possible moment arm
but do not introduce large bending forces at the wing tips when they are deflected. The
ailerons are also chosen to occupy 20% of the wing chord allowing space in front of them for
connectors and actuators to be attached to the wings rear spar, again the control surface
effectiveness is taken from Figure 4.
𝐶𝐶𝐿𝐿𝛿𝛿𝛿𝛿
= �
2𝐶𝐶𝐿𝐿𝐿𝐿 𝜏𝜏𝐴𝐴 𝑐𝑐𝑟𝑟
𝑆𝑆𝑆𝑆
� × �
𝑦𝑦2
2
+
2
3
�
𝜆𝜆 − 1
𝑏𝑏
� 𝑦𝑦3
�
𝑦𝑦1
𝑦𝑦0
Equation 30 - Aileron Rolling Moment Coefficient Equation - [1]
𝑃𝑃𝑠𝑠𝑠𝑠 = �
2𝐿𝐿𝐴𝐴
𝜌𝜌(𝑆𝑆𝑤𝑤 + 𝑆𝑆ℎ + 𝑆𝑆𝑣𝑣)𝐶𝐶𝐷𝐷𝑟𝑟
𝑌𝑌𝐷𝐷
3
Equation 31 - Steady State Roll Rate Equation - [1]

𝐼𝐼𝑋𝑋𝑋𝑋 =
𝑏𝑏2
𝑚𝑚𝑅𝑅�𝑥𝑥
2
4
Equation 32 - Moment of Inertia in X - [7]
𝜙𝜙1 = �
𝐼𝐼𝑋𝑋𝑋𝑋
𝜌𝜌(𝑆𝑆𝑤𝑤 + 𝑆𝑆ℎ + 𝑆𝑆𝑣𝑣)𝐶𝐶𝐷𝐷𝑟𝑟
� ln(𝑃𝑃𝑠𝑠𝑠𝑠
2 )
Equation 33 - Bank Angle for Steady Roll Rate Equation - [1]
𝑡𝑡2 = �
2𝜙𝜙𝑑𝑑𝑑𝑑𝑑𝑑.
(𝑃𝑃𝑠𝑠𝑠𝑠
2 2𝜙𝜙1⁄ )
Equation 34 - Time to Bank Equation - [1]
These values are then input into Equation 30, Equation 31, Equation 32, Equation 33 and
Equation 34 respectively giving a value of 𝑡𝑡2 = 1.269𝑠𝑠, this value falls within the required time
to bank and thus the ailerons are suitable. It must be noted that the value for 𝑅𝑅�𝑥𝑥 = 0.246, this
approximation has been shown to give good correlation to real world values for moment of
inertia and the constant 𝑅𝑅�𝑥𝑥 is 0.246 for light general aviation aircraft.
3.2.4 Spiral Mode Analysis
Spiral mode is the aerodynamic effect upon the wings caused by a yawing moment. As the
aircraft is disturbed in the vertical axis, the vertical stabiliser restores the aircraft due to the
static directional stability, as the aircraft yaws back towards its initial condition the fore moving
wing increases in speed, this causes an increase in lift upon this wing, whilst symmetrically
the aft moving wing slows and produces less lift. This causes an unbalance in the lift created
across the wing and the aircraft experiences a rolling moment. As the aircraft rolls it begins to
sideslip towards the lower wing, this movement causes an up wash against the vertical
stabiliser and decreases the incidence thus decreasing the generated lift causing the nose to
fall further into the sideslip, this increases the sideslip angle and causes the aircraft to fly in an
increasingly tight spiral.
3.2.5 Spiral Mode Approximation
To approximate spiral mode a method similar to that used in section 2.2.2 and 2.2.6 is used.
As spiral mode is assumed as only a change in heading angle and yaw angle with little or no
change in sideslip velocity, therefore Equation 26 and Equation 27 are considered and 𝛽𝛽 = 0
along with all control inputs. This gives a characteristic equation for spiral mode of:
𝐿𝐿𝛽𝛽 𝑁𝑁𝑟𝑟 − 𝐿𝐿𝑟𝑟 𝑁𝑁𝛽𝛽
𝐿𝐿𝛽𝛽
= 𝜆𝜆
Equation 35 - Spiral Mode Approximation Characteristic Equation - [7]

3.2.6 Spiral Mode Calculation
For the aircraft, the parameters found through the aileron design are applied and the spiral
mode can be approximated. For the initial analysis of the spiral mode motion the aircraft will
be set in cruise condition for the reasons stated in section 2.2.4. The aircraft cruise conditions
are shown in Technical Specification and the calculation of the aerodynamic derivatives is
shown in Table 13.
𝑇𝑇 13.24
Table 14 - Spiral Mode Approximation Results
3.2.7 Roll Convergence Analysis
Roll convergence is a lateral stability phenomena created by the aerodynamic effect of the
wing during the roll, it is distinguished as a non-oscillatory heavily damped motion comprising
of a change in roll angle with little or no change in yaw angle or lateral velocity. It is seen
when an aircraft is disturbed and moved into a rolling motion, the rolling motion is then
opposed by the motion of air over the wing. As the aircraft rolls a downwash is created on the
rising wing, this causes a reduction in incidence and thus a reduction in lift, symmetrically on
the falling wing an up wash is created, this causes an increase in incidence and thus an
increase in lift, these two aerodynamic effects oppose the initial rolling moment and thus
equilibrium is reached. As the motion is non-oscillatory it does not return to the initial
conditions and thus demonstrates that conventional aircraft are not bank angle stable.
3.2.8 Roll Convergence Approximation
To approximate roll convergence a method similar to that used in section 2.2.2 and 2.2.6 is
used. As roll convergence is assumed as only a change in roll angle with little or no change in
yaw angle or sideslip velocity, therefore only the roll equation, Equation 26 is considered and
𝑣𝑣 and 𝜓𝜓 are set to 0 along with all control inputs.
�
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝐿𝐿𝑝𝑝
𝐴𝐴
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜙𝜙 = 0
Equation 36 - Roll Convergence Approximation Equation - [8]
taken to calculate the characteristic equation for the roll convergence motion. As there is only
the one equation the characteristic equation can be found without calculating the determinant.

𝜆𝜆2
− �
𝐿𝐿𝑝𝑝
𝐴𝐴
� 𝜆𝜆 = 0
Equation 37 – Roll Convergence Approximation Characteristic Equation
Then again by comparing Equation 37 to Equation 16 the period and time to half amplitude
can be calculated.
3.2.9 Roll Convergence Calculation
For the aircraft, the parameters found through the aileron design are applied and the roll
convergence can be approximated. For the initial analysis of the roll convergence motion the
aircraft will be set in cruise condition for the reasons stated in section 2.2.4. The aircraft cruise
conditions are shown in Technical Specification and the calculation of the aerodynamic
derivatives is shown in Table 13.
𝑇𝑇 0.87
Table 15 - Roll Convergence Approximation Results
3.2.10 Dutch Roll
Dutch roll is an oscillatory motion combining yaw and roll, this motion is caused due to the
effect of the vertical stabilisers restoring yaw moment, as the aircraft is disturbed in the
vertical axis the incidence at the vertical stabiliser generates lift in the opposing direction. This
change in lift causes the tail to create a moment opposing the initial disturbance moment and
a change in lift across the wings, this change in lift causes a roll moment upon the aircraft. As
the aircraft returns to equilibrium the yawing motion of the aircraft causes the vertical
stabiliser to pass through equilibrium and an opposite lift is created, the aircraft oscillates
through this motion until it is damped due to the directional stability of the vertical stabiliser,
however it will not return to the initial heading and thus aircraft are not heading stable.
3.2.11 Dutch Roll Approximation
The Dutch roll is distinguished as a change in yaw angle, 𝜓𝜓 and side slip velocity, 𝜐𝜐 with little
or no change in roll angle, 𝜙𝜙. Therefore it is approximated by setting 𝜙𝜙 = 0 and removing the
roll equation.
�
𝛿𝛿
𝛿𝛿𝛿𝛿
−
𝑌𝑌𝑣𝑣
𝑚𝑚
� 𝑣𝑣 + 𝑈𝑈
𝛿𝛿𝛿𝛿
𝛿𝛿𝛿𝛿
= 0
Equation 38– Dutch Roll Approximation Sideslip Equation
−
𝑁𝑁𝑣𝑣
𝐶𝐶
𝑣𝑣 + �
𝛿𝛿2
𝛿𝛿𝛿𝛿2
−
𝑁𝑁𝑟𝑟
𝐶𝐶
𝛿𝛿
𝛿𝛿𝛿𝛿
� 𝜓𝜓 = 0
Equation 39 – Dutch Roll Approximation Roll Moment Equation

taken to calculate the characteristic equation for the Dutch roll motion.
�
𝜆𝜆 −
𝑌𝑌𝑣𝑣
𝑚𝑚
𝑈𝑈𝑈𝑈
−
𝑁𝑁𝑣𝑣
𝐶𝐶
𝜆𝜆2
−
𝑁𝑁𝑟𝑟
𝐶𝐶
𝜆𝜆
� = 0
Equation 40 - Dutch Roll Approximation Determinant
𝜆𝜆 �𝜆𝜆2
+ �−
𝑌𝑌𝑣𝑣
𝑚𝑚
−
𝑁𝑁𝑟𝑟
𝐶𝐶
� 𝜆𝜆 + �
𝑌𝑌𝑣𝑣 𝑁𝑁𝑟𝑟
𝑚𝑚𝑚𝑚
−
𝑈𝑈𝑈𝑈𝑣𝑣
𝐶𝐶
�� = 0 = 𝜆𝜆2
− �
𝑌𝑌𝛽𝛽 + 𝑢𝑢0 𝑁𝑁𝑟𝑟
𝑢𝑢0
� 𝜆𝜆 +
𝑌𝑌𝛽𝛽 𝑁𝑁𝑟𝑟 − 𝑁𝑁𝛽𝛽 𝑌𝑌𝑟𝑟 + 𝑢𝑢0 𝑁𝑁𝛽𝛽
𝑢𝑢0
Equation 41 - Dutch Roll Approximation Characteristic Equation
Producing a characteristic equation, Equation 41 that can again be compared to Equation 16
to find the period and damping.
3.2.12 Dutch Roll Calculation
For the aircraft, the parameters found through the aileron and rudder design are applied and
the Dutch roll can be approximated. For the initial analysis of the Dutch roll motion the aircraft
will be set in cruise condition for the reasons stated in section 2.2.4. The aircraft cruise
conditions are shown in Appendix 4 and the calculation of the aerodynamic derivatives is
shown in Table 13, therefore the analysis of the Dutch roll motion can begin, from the analysis
of the Dutch roll motion the following results are obtained:
𝜔𝜔𝑛𝑛 2.4763672
𝜁𝜁 0.6718621
𝑇𝑇 1.712799
𝑡𝑡1
2
0.4166106
Table 16 – Dutch Roll Approximation Results
3.2.13 Flying Characteristics
Given the values in Table 14, Table 15 and Table 16 the aircraft can be compared to a flying
characteristics table such as Table 17 for the same reasons as stated in section 2.2.8. As can
be seen by comparing the tables the aircraft is a level 1 in the roll convergence and Dutch roll
mode, it can also be noted that the aircraft is stable in all modes, this can be verified by the
positive damping constant for all modes.

Spiral Mode
Class Category Level 1 Level 2 Level 3
I, IV A 12s 12s 4s
B, C 20s 12s 4s
II, III All 20s 12s 4s
Roll Convergence
Class Category Level 1 Level 2 Level 3
I,IV A 1.0s 1.4s 10s
II,III A 1.4s 3.0s 10s
All B 1.4s 3.0s 10s
I,IV C 1.0s 1.4s 10s
II,III C 1.4s 3.0s 10s
Dutch Roll Mode
Level Category Class Min 𝜁𝜁 Min 𝜁𝜁𝜔𝜔𝑛𝑛 Min 𝜔𝜔𝑛𝑛
1 A I,IV 0.19 0.35 1.0
1 A II,III 0.19 0.35 0.4
1 B All 0.08 0.15 0.4
1 C I,II-C,IV 0.08 0.15 1.0
1 C II-L,III 0.08 0.15 0.4
2 All All 0.02 0.05 0.4
3 All All 0.02 --- 0.4
Table 17 - Lateral Flying Characteristics - [7]

REFERENCES
2012.
[2] University of Hertfordshire, Static Stability, Hatfield, Hertfordshire, 2014.
[3] I. H. A. a. A. E. V. Doenhoff, Theory of Wing Sections Including a Summary of Aerofoil
Data, New York: Dover Publications Inc, 1959.
[4] University of Hertfordshire, Approximations to the Longitudinal Natural Modes, Hatfield,
Hertfordshire, 2014.
[5] University of Hertfordshire, Introduction to Aircraft Stability and Control, Hatfield,
[6] University of Hertfordshire, Determination of Aerodynamic Forces and Moments, Hatfield,
[7] R. C. Nelson, Flight Stability and Automatic Control.
[8] University of Hertfordshire, Solution of the Lateral Equations of Motion, Hatfield,
[9] University of Hertfordshire, DAG Lecture Slides 1, Hatfield, Hertfordshire`, 2014.
[10] Massachusetts Institute of Technology, Lecture AC 2, Cambridge, Massachusetts, 2003.

Training Aircraft
Appendix 9
Aircraft Modelling
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015
i

ABSTRACT
This document is an appendix to the main report, it describes in detail how to aircraft was
modelled using the Dassault Systems CATIA software and contains technical drawings for all
components designed.
ii

TABLE OF CONTENTS
ABSTRACT ....................................................................................................................................ii
1 Aircraft Modelling.................................................................................................................. 1
1.1 Aircraft Modelling Process............................................................................................. 1
1.2 Aircraft Sketch Design................................................................................................... 2
1.3 Modelling Software........................................................................................................ 2
1.4 Aircraft 3D Modelling Techniques ................................................................................. 2
1.4.1 Part Design............................................................................................................ 2
1.4.2 Surface Design...................................................................................................... 2
1.4.3 Assembly Design................................................................................................... 3
1.4.4 Rendering.............................................................................................................. 3
1.4.5 Drafting.................................................................................................................. 3
1.5 Model Comparison to Initial Sketch............................................................................... 3
2 Final Aircraft Design and Specification................................................................................. 5
2.1.1 General Arrangement............................................................................................ 5
2.1.2 3 View Render....................................................................................................... 5
2.1.3 Section and Detail Renders................................................................................... 5
2.1.4 Technical Specification.......................................................................................... 5
iii

LIST OF FIGURES
Figure 1 - Model Creation ............................................................................................................. 1
Figure 2 - Aircraft General Arrangement....................................................................................... 6
Figure 3 - Aircraft 3 View Renders ................................................................................................ 7
Figure 4 - Aircraft Detail and Section Renders ............................................................................. 8
iv

1 Aircraft Modelling
A description of the how the aircraft was modelled, which conventions and programs were
used and how the aircraft moved from a drawing to a fully rendered 3D drawing, for further
information refer to Appendix 9 and for A3 versions refer to Appendix C.
Aircraft sketching and modelling is an integral part of any design process for any product,
having a 2D or 3D representation for a product is an excellent tool for intuitive design and
allows and individual designer or a design team a view of all components for a product
making clashes between design aims visible and more easily understood. A 2D or 3D
representation is also a necessary for marketing a product, giving a customer a view of the
project and if used at design meetings can allow the customer to review the design for
considerations that the design specification may not have considered. For an aerospace
application the 2D and 3D representations can also be used for evaluation purposes, using a
CAD model for Finite Element Analysis, Computational Fluid Dynamics and other simulation
techniques.
1.1 Aircraft Modelling Process
The aircraft modelling process begins with the concept sketch and ends with a 2D or 3D
model, the model can take any form however different models are useful for different
applications. In this project it is suitable to create a final 3D CAD model for the aircraft as this
can be used further with evaluation of the aircraft, simulation and creation of a physical 3D
model for aircraft wind tunnel testing or marketing purposes, the process involved in the
development of the model from concept sketch to 3D model is shown in, Figure 1.
Figure 1 - Model Creation
Concept Sketch
Generation
Concept Sketch
Evaluation
Final Concept
Sketch
Part Concept
Sketches
Part Design
Sketches
3D Part
Creation
3D Assembly 3D Rendering

1.2 Aircraft Sketch Design
As can be seen the first stage of the modelling process is the creation of the initial sketch,
along with generating a concept for the technical design of the aircraft the concept for the
aircraft model is created, the initial sketches are dimensionless representations and thus may
not be scale however the initial sketches are designed to distinguish major design
considerations and to implement initial design decisions. Along with an initial sketch for the
entire aircraft sketches of individual major components are also generated, these sketches
are used when selecting or designing component parts for the aircraft, components such as
aircraft major structures, engines, fuel sources, landing gear systems and cockpit, these
sketches are shown in Appendix C.
1.3 Modelling Software
The modelling software used for this project is the Dassault Systems CATIA software
package, CATIA is an industry standard CAD and CAE software that contains programs for
modelling using a sketching tool for part or surface design, also contained are programs for
rendering and drafting models along with some analysis and evaluation tools for applications
such as FEA.
1.4 Aircraft 3D Modelling Techniques
With the software used and the programs contained within several modelling techniques are
utilised, this is due to the advantages and disadvantages of some techniques for the
applications required during this modelling process.
1.4.1 Part Design
Part Design utilises a combination of simple geometries to create complex parts, part design
typically involves the creation of sketches which are then extended through planes to create
solid parts and then hollowed and shaped to create the desired product. Part design is useful
for creating basic shapes such as rectangles and cylinder and can be used to create complex
parts with simple geometrical features such as straight edges. Therefore part design is used
for the creation of the motor, controller, spars and other simple parts, part design is also
utilised to create simple geometries upon complex parts, such as pipe fittings, and to convert
surfaces into parts for material analysis.
1.4.2 Surface Design
Surface Design utilises complex geometries to create complex parts through a combination of
sections and guides using a mathematical solution to compute how the surface behaves,
surface design is useful for creating complex objects from a series of curves such as aerofoils
and aircraft surfaces and as such all the aircraft surfaces including cockpit, wing and stabiliser
surfaces were created using surface design. Surface design can also be utilised to extrude

basic shapes along complex curves such as those required for the creation of the aircraft
structural components, landing gear structures and pipe work. The major limitation of surface
design however is that it cannot be used for material analysis and therefore these surfaces
must be converted into parts so that material properties can be assigned to them.
1.4.3 Assembly Design
Assembly design uses a system of constraints between parts to create assemblies,
assemblies are a combination of parts which could represent the final product or that can be
used to ease the design process, where many parts are required assemblies can be split in
several subassemblies, these subassemblies maintain the constraints assigned and act as
parts in a larger assembly, the assembly design program is utilised during the project in both
applications, the creation of sub-assemblies such as the motor, aircraft structure, battery
compartment and undercarriage and the assembly of the final aircraft. The assembly program
allows for the combination of many small or complex parts reducing the work required when
creating parts; however it cannot create parts and thus relies on the other programs.
1.4.4 Rendering
The rendering tool uses a mathematical representation of light and light sources combined
with the material properties of the model to create a realistic representation of the product
generally for marketing purposes such as promotion of the product, the rendering tool
computes how rays of light interact with a surface and the material it’s been assigned
including the direction and intensity of any reflected light, the rendering tool also contains
scenes in which the product can be input and thus represented. The rendering tool however
relies completely on the model input into it and thus requires a combination with either part or
assembly design.
1.4.5 Drafting
Much like the rendering tool the drafting program generates images of the product, the
drafting program however is used to create a dimensioned technical drawing of the product,
the drafting program takes the part and surface design features creating a technical diagram,
Like the render this can be used to market the aircraft, giving the customer a technical
diagram for the entire aircraft or individual parts. Again like the rendering tool the drafting
program completely relies upon a model created by part, surface or assembly design.
1.5 Model Comparison to Initial Sketch
With the creation of the 3D model a comparison can be made to the original concept sketches
to ensure that the initial concepts have been adhered to and thus the design specification
from an aesthetic point of view has been fulfilled.

2 Final Aircraft Design and Specification
The specification and design of the final version of the aircraft with a final render and final
technical specification compiling all information around the aircraft refer to Appendix C for A3
versions.
With the technical and 3D design of the aircraft complete the final concept of the aircraft can
be presented, this is done in several ways with the most suitable mentioned below:
2.1.1 General Arrangement
The general arrangement for the aircraft gives potential customers the major dimensional
data, allowing them to immediately see the size, weight and geometry of the aircraft, the
general arrangement also allows a customer to identify quickly whether the aircraft will be
suitable for their chosen application.
2.1.2 3 View Render
A 3 view render shows a potential customer another general arrangement however all
dimensions must be estimated as the render is not dimensioned, although not as technically
useful as the general arrangement the 3 view render is an excellent marketing tool and can
be used to show potential liveries, paint schemes and scenarios giving a more appealing
view.
2.1.3 Section and Detail Renders
Sections can be used to show individual details to a customer and market unique selling
points for an aircraft, for this aircraft details such as the motor, battery compartment and
cockpit view can be shown again to increase marketability of the aircraft.
2.1.4 Technical Specification
The technical specification for the aircraft gives a customer all the salient points surrounding
the aircrafts, performance, statistics and other details which may be hard to visualise or
impossible to show in any other form giving the customer a detailed numerical comparison to
other aircraft.

Figure 2 - Aircraft General Arrangement

Figure 3 - Aircraft 3 View Renders

Figure 4 - Aircraft Detail and Section Renders

Training Aircraft
Part Catalogue
Report by
Supervisor
Liz Byrne
Date
20 APRIL 2015

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1/1
GENERALARRANGEMENT
UH145-T
1:80
2389
3651.66
8998.83
9001575
1794.31
6521.07
1518.44
826.08
643.22
3422.8
1625.8
426
2000
1453.6
1118.88
847.52

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:25
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
HT0014.40
A4
1/1
BenJohnson
13/04/2015
UH145-T
B
B
C
C
SectioncutB-B
SectioncutC-C

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:5
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
1.96
A4
1/1
BenJohnson
13/04/2015
INSTRUMENTPANEL
UH145-T

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:4
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
2.64
A4
1/1
BenJohnson
13/04/2015
RADIATOR
UH145-T

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:15
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
VT0025.16
A4
1/1
BenJohnson
13/04/2015
UH145-T

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALEWEIGHT(kg)DRAWINGNUMBERSHEET
SIZE
CHECKEDBY:
BenJohnson
DATE:
13/04/2015
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
31.94
A4
1/11:15
UH145-T
SUBBATTERY
560
608
250

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:25
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
19/03/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
SUBMOTOR113.99
A4
1/1
UH145-TBenJohnson
13/04/2015
970
950476

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:80
SIZE
CHECKEDBY:
BenJohnson
DATE:
23/03/2015
DESIGNEDBY:
BenJohnson
DATE:
19/03/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
SUBSTRUCTURE376.42
A4
1/1
UH145-T

AD
B
C
A
D
3
3
2
2
4
4
1
1
SCALE
1:80
SIZE
CHECKEDBY:
DATE:
DESIGNEDBY:
BenJohnson
DATE:
13/04/2015
A_
B_
C_
D_
E_
F_
G_
H_
I_
WS00116.04
A4
1/1
UH145-TBenJohnson
13/04/2015
9000.72
1794.31
1529.04
A
A
B
B
SectioncutA-A
SectioncutB-B

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