Preliminary design of a short-haul regional aircraft

P R E L I M I N A R Y D E S I G N O F A
S H O R T - H A U L R E G I O N A L A I R C R A F T
P R O J E C T E O L O
D E C E M B E R 2 019
U N I V E R S I T A T P O L I T È C N I C A D E C A T A L U N YA
Project Manager: Alex Guerrero Lorente 
Aerodynamics Department: Andrea González Romero, Marc
Lladó Bordàs, Agustí Porta Ko, Pere Valls Badia, Francisco Boira Gual 
Structures Department: Adrià Barja Peláez, Josep Sánchez Sanz 
Propulsion Department: Pol Niño Pol, Gabriel Martín Sánchez,
Adrià Barceló Gregoriano 
Business Department: Juni Choi Bae, Iker Blanco Bravo, Neus
Oliveras Tramunt

Projects Group 04
Table of contents
1 Aim of the project 5
2 Scope 5
2.1 Out of scope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
3 Requirements 6
4 Market Study 7
4.1 Market Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7
4.2 Market’s growth . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8
4.2.1 Types of Regional aircraft . . . . . . . . . . . . . . . . . . . . . 9
4.3 Competition and market share . . . . . . . . . . . . . . . . . . . . . . . 9
4.4 Product’s differentiation . . . . . . . . . . . . . . . . . . . . . . . . . . 11
5 General Configuration 13
5.1 Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
5.2 Wing configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
5.3 Empennage configuration . . . . . . . . . . . . . . . . . . . . . . . . . . 14
5.4 Landing gear type and disposition . . . . . . . . . . . . . . . . . . . . . 14
6 Conceptual design and performance analysis 16
6.1 General outline of the method for estimating Take-Off gross weight, WTO,
empty weight, WE, and mission fuel weight, WF . . . . . . . . . . . . . . 16
6.1.1 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
6.2 Sizing requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
6.2.1 Sizing to take-off distance requirements . . . . . . . . . . . . . . 19
6.2.2 Sizing to landing distance requirements . . . . . . . . . . . . . . 19
6.2.3 Sizing to cruise speed requirements . . . . . . . . . . . . . . . . 20
6.2.4 Sizing to climb requirements . . . . . . . . . . . . . . . . . . . . 21
6.2.5 Matching all the sizing requirements . . . . . . . . . . . . . . . . 22
7 Aircraft’s aerodynamics 25
7.1 Fuselage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25
7.2 Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
7.3 Tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
7.4 Weight components and gravity center calculation . . . . . . . . . . . . . 29
8 Aircraft propulsion 31
8.1 Engine type selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31
8.2 Engine configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
8.3 Number of engines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 33
8.4 Engine selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 34
8.5 Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 35
8.6 Propellant tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37
9 Aircraft systems 39
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9.1 Electric System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39
9.1.1 Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 39
9.1.2 APU (auxiliary power unit) . . . . . . . . . . . . . . . . . . . . 39
9.1.3 Starter Generator Unit . . . . . . . . . . . . . . . . . . . . . . . 40
9.2 Lighting System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
9.2.1 External lights . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
9.2.2 Internal lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
9.2.3 Emergency lights . . . . . . . . . . . . . . . . . . . . . . . . . . 42
9.3 Instrumental system . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42
9.4 Hydraulic System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
9.4.1 Hydraulic fluid . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
9.4.2 Distribution System . . . . . . . . . . . . . . . . . . . . . . . . 44
9.4.3 Main hydraulic system components . . . . . . . . . . . . . . . . 45
9.4.4 Hydraulic System Architecture . . . . . . . . . . . . . . . . . . . 46
10 Structural Design 48
10.1 Fuselage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
10.1.1 Nose structure . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
10.1.2 Cabin Structure . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
10.2 Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
10.3 Tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
10.3.1 Empennage Structure . . . . . . . . . . . . . . . . . . . . . . . . 52
10.3.2 Vertical Tail Structure . . . . . . . . . . . . . . . . . . . . . . . 52
10.3.3 Horizontal Tail Structure . . . . . . . . . . . . . . . . . . . . . . 53
10.4 Materials choice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54
10.5 Landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
11 CFD Analysis 60
11.1 Fuselage Nose Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
11.2 Cruise Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
12 Economic feasibility study 63
12.1 Market Entry Strategy . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63
12.1.1 Implementation . . . . . . . . . . . . . . . . . . . . . . . . . . . 64
12.2 Joint Venture Deal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 64
12.2.1 Financial Agreement . . . . . . . . . . . . . . . . . . . . . . . . 64
12.2.2 Operative Agreement . . . . . . . . . . . . . . . . . . . . . . . . 66
12.3 Pricing Strategy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67
12.4 Cost Benefit Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
12.4.1 Operative Budget & financial statements . . . . . . . . . . . . . . 70
12.4.2 Costing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 76
12.4.3 Ratios . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 77
12.4.4 Internal Rate of Return . . . . . . . . . . . . . . . . . . . . . . . 77
13 Environmental Impact Study 78
13.1 Production . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78
13.2 Usage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 79
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13.3 Dismantling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 80
13.4 Social considerations . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
13.4.1 Ethics code . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
13.4.2 Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . 82
13.4.3 Society . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83
14 Safety considerations 84
14.1 Regulations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84
14.2 Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 85
14.3 Prevention of work risks . . . . . . . . . . . . . . . . . . . . . . . . . . 85
15 Planning and scheduling for the follow-up of the project 86
15.1 Hierarchy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
15.2 Roles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
15.3 Work Breakdown Structure . . . . . . . . . . . . . . . . . . . . . . . . . 88
15.4 Tasks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
15.5 Work tracking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88
16 Quality 89
16.1 Iterative process . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
16.2 Compliance matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 89
16.3 KPIs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90
17 Conclusions 92
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List of Figures
1 Eurostat. International intra-EU air passenger transport by reporting coun-
try and EU partner country. Extracted from [40]. . . . . . . . . . . . . . 8
2 Regional Market. Data extracted from [87]. . . . . . . . . . . . . . . . . 8
3 Market share distribution of the regional aircraft market. Extracted from:
[48]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11
4 Low Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
5 Mid Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
6 High Wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14
7 Iterative process. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
8 Sizing to Take-Off. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
9 Sizing to Landing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
10 Sizing to cruise speed. . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
11 Sizing to climb requirements FAR25.119 and FAR25.121(go-around). . . 22
12 Matching of all the sizing requirements . . . . . . . . . . . . . . . . . . 23
13 Cabin Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
14 Wing planform. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
15 Horizontal stabilizer. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
16 Vertical stabilizer. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
17 Gas turbine efﬁciency [63]. (Note that the x-axis units are in mph, the
cruise speed of the aircraft is between 400-500 km/h ⌘ 250-311 mph). . . 32
18 OWA method Table to select the best propeller. . . . . . . . . . . . . . . 36
19 W-R plot. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 37
20 G-242 Gill Battery [7]. . . . . . . . . . . . . . . . . . . . . . . . . . . . 39
21 APS 3200 Auxiliary Power Unit [89]. . . . . . . . . . . . . . . . . . . . 40
22 Astronics CorePower Brushless Starter Generator Unit [11]. . . . . . . . 40
23 External lights [15]. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
24 Universal codes for each type of system [85]. . . . . . . . . . . . . . . . 45
25 Hydraulic system schematic [85]. . . . . . . . . . . . . . . . . . . . . . . 47
26 Global view. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
27 Upward view. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
28 Internal structure of the nose. . . . . . . . . . . . . . . . . . . . . . . . . 49
29 Global view. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
30 Front view. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
31 Internal structure of the wing. . . . . . . . . . . . . . . . . . . . . . . . . 51
32 Internal structure of the empennage. . . . . . . . . . . . . . . . . . . . . 52
33 Internal structure of the vertical tail. . . . . . . . . . . . . . . . . . . . . 53
34 Internal structure of the horizontal tail. . . . . . . . . . . . . . . . . . . . 54
35 Longitudinal and lateral tip-over criteria [71]. . . . . . . . . . . . . . . . 56
36 Lateral ground clearance criteria [71]. . . . . . . . . . . . . . . . . . . . 56
37 Longitudinal ground clearance criteria [71]. . . . . . . . . . . . . . . . . 57
38 Geometry for static load calculation [71]. . . . . . . . . . . . . . . . . . 58
39 Pressure results. From left to right, geometry proposed in [43], variant 1,
variant 2. Scale is in Pa . . . . . . . . . . . . . . . . . . . . . . . . . . . 60
40 Pressure results. Scale is in Pa. . . . . . . . . . . . . . . . . . . . . . . . 61
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41 Pressure results. Scale is in Pa. . . . . . . . . . . . . . . . . . . . . . . . 61
42 Velocity field XZ plane. Scale is in m. . . . . . . . . . . . . . . . . . . . 62
43 Estimated market share with Kondor Aircraft. . . . . . . . . . . . . . . . 69
44 Sales distribution during the first 20 years. . . . . . . . . . . . . . . . . . 70
45 Available Cash flow. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
46 Treasury. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74
47 Production goals approach [18]. . . . . . . . . . . . . . . . . . . . . . . 79
48 Dismantling operation of an airplane [5]. . . . . . . . . . . . . . . . . . . 81
49 Work Breakdown Structure. . . . . . . . . . . . . . . . . . . . . . . . . . 88
50 Tracking of the indicators. . . . . . . . . . . . . . . . . . . . . . . . . . 90
51 Propulsion department effort by every member. . . . . . . . . . . . . . . 91
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List of Tables
1 Potential competitors. . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13
2 Similar airplanes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24
3 Main fuselage parameters. . . . . . . . . . . . . . . . . . . . . . . . . . 25
4 Main wing geometric parameters. . . . . . . . . . . . . . . . . . . . . . 27
5 Flaps. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
6 Ailerons. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
7 Spoilers. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
8 Relation between engine faliure probability and the number of engines
used. Pef is the probability of engine faliure [71]. . . . . . . . . . . . . . 33
9 Basic engine technical characteristics. . . . . . . . . . . . . . . . . . . . 34
10 Possible propellers that can be attached to the engine. . . . . . . . . . . . 35
11 Technical data of Hamilton Sundstrand 568F-5 propeller. . . . . . . . . . 36
12 Aluminium alloy number and its identification. . . . . . . . . . . . . . . 44
13 Elements of each hydraulic subsystem. . . . . . . . . . . . . . . . . . . . 47
14 Wheels sizing. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 59
15 Structure Sales department. . . . . . . . . . . . . . . . . . . . . . . . . . 70
16 Structure Production department. . . . . . . . . . . . . . . . . . . . . . . 71
17 Structure Administration department. . . . . . . . . . . . . . . . . . . . 71
18 Structure R+D department. . . . . . . . . . . . . . . . . . . . . . . . . . 71
19 Profit and Loss account. . . . . . . . . . . . . . . . . . . . . . . . . . . . 72
20 Treasury Forecast and Cash flow. . . . . . . . . . . . . . . . . . . . . . . 73
21 Compliance matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90
22 Basic summary of the fulfilled requirements . . . . . . . . . . . . . . . . 92
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1 Aim of the project
The main objective of the project is to deliver a preliminary design of a short-haul Re-
gional Aircraft that is capable of achieving several requirements.
From now on, this project will be referred and developed as Project Eolo within the
internal departments of Kondor Aircraft.
2 Scope
The main features that Project Eolo comprises are the execution of:
• Market study and business plan including the background, state of the art, main
alternatives and decision of the best solution
• General configuration and performance analysis
• Aerodynamic analysis which involves the definition of the the main elements: fuse-
lage, wing and tail by using a panel’s method –in the case of the wing– and a CFD
study to evaluate the nose performance as well as the general aerodynamic coeffi-
cients
• Engine choice within the available feasible possibilities
• Study the feasibility of designing the electrical system and control systems and
perform it or, otherwise, outsource this service with the corresponding study of
costs and infrastructure
• CAD design and structural study based on the state of the art materials and archi-
tectures. Internal structure is studied and main aircraft elements are defined
• Internal disposition study including passenger’s seating and overall equipment
• Economic Feasibility study
• Environmental study
• Safety study
• Project Management Plan
• Quality plan involving the evaluation of KPI’s
• Lessons learned document
• Project poster and project presentation
All of the tasks mentioned above will be presented in this same Report document as well
as its Attachments. Also a Budget, specific Drawings and Technical sheets documents
have been elaborated.
2.1 Out of scope
• Design of the engine as well as its other systems (hydraulic, electrical, etc.)
• Structural analysis evaluating the stresses and internal behaviour of forces
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3 Requirements
The aim of this chapter is to present the necessary requirements that the design of the
short-haul regional aircraft must fulfill to meet the needs of the customer.
As a summary, the general requirements are:
• Crew: 2 pilots and 2 TCP
• Capacity: 70 to 80 passengers
• Cruising speed: 450-500 km/h
• Maximum range: longer than 1.600 km with MPL
• Maximum take-off field length of 1.300 meters with MTOW
Moreover, extra requirements and aspects to fulfill are also established. These are:
• Inclusion of a market study to justify the design of the aircraft.
• Compliance of EASA and FAA regulations.
• Inclusion of engine number and type decision and justification.
• Focus on minimizing fuel consumption maintaining a competitive jet sale price (di-
rect operative cost comparison with competitors).
• Delivery of jet drawings including the three views of the plane and its dimensions.
• Definition of materials to be used and the structural concept.
• Deliveries of the passenger cabin design including seating arrangements.
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4 Market Study
The present section is aimed at analyzing the current situation of the Regional Aircraft’s
market to help the reader understand the necessity of this project. This way, in this chapter
several areas such as the Background, State of the art, and the Main alternatives and
decision of the best one are covered.
4.1 Market Analysis
From an economical and statistical point of view, the global aircraft market has grown
significantly over the last years. People’s habits have evolved significantly different since
1980 and all these trends have had an important effect and repercussion on the way people
travel. However, there have been a certain type of aircraft within the industry that have
grown significantly higher than the overall [40], Regional Aircraft.
“The regional market is currently led by non-European players, with the exception of tur-
boprop manufacturer ATR (a 50/50 Joint Venture between Leonardo and Airbus Group).
For Europe‘s aeronautical industry there‘s a clear and urgent need to invest in developing
new technologies in order to recover global leadership.” - CleanSky, 2017.
To get a better approach, a regional aircraft is a commercial aircraft used by some airlines
for short-haul flights and is designed to have a maximum capacity of 80 to 100 passengers.
In some cases, regional airplanes are used by subsidiaries of larger airlines to transport
passengers from one airport to another. The creation of regional aircraft is by any means
a new industry, its origins go back to the end of the Second World War, in 1947. So, the
reader must be asking why it is now an emergent sector within the aerospace industry if
its creation was so long ago.
The following reasons can help to understand it:
• The release of air traffic, first in the United States and later, in the 80’s in Europe
reduced the limitations on the creation of new companies and the demand grew
notoriously.
• The global economic growth between 1980 and 2000’s, as well as the technological
boom that helped reducing fuel consumption while enlarging distances and lower-
ing the prices.
• The need on the part of the regions for a good communications network, since all
the infrastructures, and among the airports, are the main door to grow and generate
jobs. Airports act as magnets for a wide range of companies and industries.
• The saturation of the main airports that forces the usage of secondary airports.
• The progress that regions seek through leisure and tourism. Small airports near a
mountain range, beach or amusement centers have grown considerably. Tourism
has been and still is one of the great opportunities of the 21st century (see Figure
1).
A quite sensible way to support these trends is to accompany them with real data. Ac-
cording to Clean Sky [68], the regional air traffic will generate, in the next 20 years, a
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market demand of 9.000 regional aircraft, with a market value of 360 billion euros (in
average, 18 billion euros per year). Also Boeing estimates that, in the next 20 years the
demand will raise up to 40.000 new aircraft [17] –taking into account that these demands
come from the market growth but also from the need to replace old aircraft–. Therefore it
can be stated that the 22,5% of these 40 thousand new aircraft will be regional.
Figure 1: Eurostat. International intra-EU air passenger transport by reporting country
and EU partner country. Extracted from [40].
4.2 Market’s growth
For a better understanding, there are 2 kinds of regional aircraft in the market: Turboprops
(have a visible propeller) and Jet aircraft. The impact that regional airplanes– in terms of
market share– (considering the total amount of all kind of aircraft) is estimated to be
around 33% [68] . Afterwards, the competitors that exist in this percentage and currently
lead this market will be commented.
Figure 2 can help understand and support the numbers mentioned above so that the reader
can visually understand the big impact and the true necessity that these artifacts have in
the current market situation.
Figure 2: Regional Market. Data extracted from [87].
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It is rather obvious that the trends show a remarkable presence and growth –this graphic
stops at year 2009– from both large and small Regional Jets (RJs) without forgetting about
mentioned Turboprops which clearly show a unique market chance.
4.2.1 Types of Regional aircraft
However, it is not an easy task to understand and summarize the main reasons why Jet
Aircraft are a bit more demanded than Turboprops –some common thinking tends to be-
lieve that Turboprops are old-fashioned vehicles that have been left apart, but that is by no
means a true fact–.
The main differentiation factors between Jet Aircraft and Turboprops reside in its na-
ture: Turboprops represent the ultimate crossover from piston to jet. A turbine engine is
extremely light and produces tremendous power-to-weight as opposed to a comparable
piston engine. Turboprops are also able to land in more difficult situations (steep and
wet tracks, mountainous surfaces...) and that makes them the only suitable choice for
some airports. If it is taken into consideration that turboprops are generally used for short
flights, it is quite reasonable to believe that they win in this fuel efficiency battle.
When it comes to jet engines, it is true that they are able to propel the aircraft to notably
higher flight levels than Turboprops, up to the FL40 (Flight Level) as opposed to the
Turboprop cap of around FL30. Alternatively, jet engines burn a lot more fuel than a
Turboprops but often make up for it by being faster, hence they are able to spend less time
in the air for the same destination.
This way, and assuming the requirements stated in Section 3, the most feasible option is
to choose between the design of a Jet Engine or a Turboprop.
By studying properly these requirements, the speed range which the aircraft must operate
restricts clearly the choice to turboprops. Moreover, if one of the aims of this projects is
based on minimising the fuel consumption, the choice is even more obvious and accurate.
However as this choice is not trivial and naive, a more precise justification is given in
section 8.1, which evaluates the general pros and cons– and makes it up for a solution of
compromise between these 2 configurations– to deliver the most suitable choice that fits
the initial requirements.
4.3 Competition and market share
As stated previously in Figure 2, the fact that airlines are demanding more and more
regional aircraft is mainly because of the new people’s trends and necessities (cheaper
flights, secondary airports, more traffic flow...). A lot of companies might be trying to fill
this necessities, therefore, knowing the competition is key to decide how to differentiate
a product from the others. The following companies are the real and key competitors as
they are producing similar aircraft:
• Embraer E-Jet: its family is a series of medium-range twin-engine jet airliners pro-
duced by Brazilian aerospace conglomerate Embraer. The aircraft is used by both
mainline and regional airlines around the world. In the summer of 2018, Boeing
and the Brazilian Embraer announced a joint venture to take on Airbus-Bombardier
endeavour. Embraer is currently the third largest aircraft manufacturer in the world.
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• Bombardier: CRJ700, CRJ900, and CRJ1000 are regional airliners based on the
Bombardier CRJ200 with its origin in Canada. In 2017, Bombardier had been fac-
ing financial struggles for quite some time and Airbus purchased a majority stake at
Bombardier. That is when the Bombardier CSeries regional jets became the Airbus
A220 and with the financial backing of a huge corporation, the now joint venture of
Bombardier and Airbus A220 is predicted to take up around 60% of the regional jet
market.
• ATR: it is a Franco-Italian aircraft manufacturer located in Toulouse, France. ATR
has sold more than 1,500 aircraft and has over 200 operators in more than 100
countries and they do produce both turbojets and turboprops. Airbus owns 50% of
the company.
• The Mitsubishi Regional Jet (MRJ): Jet aircraft seating 70–90 passengers manu-
factured by Mitsubishi Aircraft Corporation. The project made its maiden flight in
2015 but several problems appeared since then.
• Comac ARJ21: is a twin-engined jet regional airliner. The design incorporates
components from 19 major European and US aerospace suppliers, including Gen-
eral Electric (engine production), Honeywell (fly by-wire system) and Rockwell
Collins (avionics production).
• Sukhoi Superjet 100: it is a fly-by-wire regional jet in the 75- to 95-seat category.
With development starting in 2000, the airliner was designed by the civil aircraft di-
vision of the Initial Research Regional Aircraft Russian aerospace company Sukhoi,
with some co-operation with Boeing in the early stages of the project.
• Antonov An-148: it is a regional jet aircraft designed by the Ukrainian Antonov
company and produced by Antonov itself (74 seats). The development of the plane
was started in the 1990s, and the maiden flight took place on 17 December 2004.
The plane completed its certification programme on 26 February 2007. The An-
148 has a maximum range of 2,100–4,400 kilometres and is able to carry 68–99
passengers, depending on the configuration.
Next, Figure 3 shows the exact market share distribution of all the previously named
aircraft manufacturers and what the future previsions look like:
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Figure 3: Market share distribution of the regional aircraft market. Extracted from: [48].
4.4 Product’s differentiation
Due to the characteristics of aerospace industry there are little to no things that competitors
have not covered, but one of the main problems they face is the increasing workforce costs
and manufacturing delays. One of the main reasons is because of the conservativeness of
the industry and no AI usage on both production processes and supply chain. On the other
hand, there are key aspects that would definitely make a difference on a new product, and
consequently raise the value of Project Eolo if solved appropriately. It is important that
some of them are covered in this project’s design but others are out of the scope due to its
complexity. These are:
• Low weight structural solutions: The scope is to contribute to the reduction of air-
craft weight and simplification of structural parts, thereby lowering fuel consump-
tion and reducing associated environmental impact.
• Reduce external noise: Such as landing gears, doors and bays, high-lift devices
and specific aircraft configurations to cut off external noise. This aspect is relevant
for regional aircraft because their flight departures and landings are more frequent
than larger aircraft in a typical hub airport, thereby affecting airport noise emission.
Also, regional aircraft are able to operate out of smaller city airports as well as
airports closer to urban areas which are more sensitive to noise. As aeroacoustics
is a wide, complex and relatively new field that would require loads of resources, it
will not be explored within this preliminary design.
• Aerodynamics: These are the key techniques to reduce drag and load control to
increase aerodynamic efficiency in cruise and off-design conditions (such as climb
and descent), thus reducing the polluting emissions.
• All electric solutions are addressed because they potentially improve the operational
efficiency of systems and simplify maintenance costs which are critical for regional
aircraft. In turn, they contribute to reduce fuel consumption and emissions because
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electrical systems are more energy efﬁcient and use fewer polluting materials than
traditional solutions.
In conclusion, it can be stated that this company has a potential and powerful oppor-
tunity to operate in the European market with a product that is able to satisfy the
customer’s needs. Big alliances and joint ventures have been historically made to en-
hance the competition and reach as much market share as possible. So, in further sections
there will be a proper explanation of the ﬁnal design solution and the way this company
believes it can be best to enter to the market –business architecture–.
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5 General Configuration
As the aim of this project is to design a regional aircraft it is necessary that, before start-
ing with deeper calculations and mathematical estimations, an overall configuration of
the plane is chosen. This process is about studying the configurations of several similar
aircraft –as mentioned in the market study 4,– that are already flying. After that, it is
possible to start the development of the chosen solution.
The first steep that must be done is a tabulation of the different competitors as seen in
Table 1.
Airplane Type WPL WTO Vcrmax Range
kg kg km/h km
ATR-72 7.500 23.000 510 1.528
Dash 8 q400 8.489 30.482 556 2.040
British Aerospace ATP 7.235 22.930 496 1.825
Table 1: Potential competitors.
Most airplanes which have been built or are being built today are based on a conventional
configuration. Comparing the models of Table 1, it can be stated that in regional aviation,
aircraft use the conventional configuration. This way, this is the one that is going to be
used. However, the deep discussion of the different configuration outlines is presented
in Section 1: General Configuration of the report’s attachments. Here, only a brief sum
up is presented, so that the reader can understand the decision making process.
5.1 Propeller
The decision of using a turboprop and a jet aircraft is mainly based on the general re-
quirements of the project. The aircraft’s requirements that have the most impact in the
propeller decision are:
• Range longer than 1.600 km
• Cruising speed: 450-500 km/h.
It is considered by achieving a compromise between the advantages and disadvantages
that the cruising speed is too low to use turbofans.
On the other hand, the maximum propulsive efficiency for turboprops is reached for
speeds of 300 kts which is equivalent to 550 km/h. Furthermore, the range required
is quite short so, high ceiling will not be reached, making the turboprop the best op-
tion. And finally, for these flight conditions a turboprop engine would burn less fuel per
passenger compared to a turbofan thus, it will maximise efficiency and minimise the
environmental damage.
5.2 Wing configuration
In terms of wing/fuselage disposition, wings can be classified as follows:
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• High wing
• Mid wing
• Low wing
Figure 4: Low Wing Figure 5: Mid Wing Figure 6: High Wing
Once the turboprop engines have been chosen because of its high efficiency for the flight
conditions, the wing/fuselage configuration depends a lot of it. As it can be seen from
Figures 4, 5 and 6 [70], the blades of the propeller and their size have a huge impact on the
way the propeller transforms the power produced by the engine in thrust. However, before
even knowing the diameter of the propeller it is obvious that a high-wing configuration,
with the turboprops arranged below the wing, will guarantee higher safety levels during
take-off and landing because the engine and the blades are far from the landing floor.
5.3 Empennage configuration
The empennage, also known as tail, is the main structure at the rear of an aircraft that
provides stability during flight and allow the aircraft to reach the needed angle of attack
during takeoff and landing. Generally, aircraft incorporate an empennage by having ver-
tical and horizontal stabilising surfaces which stabilise the flight dynamics of yaw and
pitch, as well as housing control surfaces.
From all the configurations studied and considering that a high wing has been selected
for this regional aircraft, the empennage must be as well designed according to it. As can
be checked in Section 1: General Configuration of the Report attachments, the main
reason for choosing a T-tail is because it is kept well out of the disturbed airflow. Every
wing that generates lift creates a wake backwards which means the airflow is disordered
and would drastically reduce the performance of a conventional tail positioned behind.
A T-tail where the horizontal stabilizer is positioned at the top of the vertical stabilizer
would give smoother and faster airflow to the horizontal stabilizer allowing high
performance aerodynamics. It is important to note that T-tail aircraft often suffer a sever
pitching moment instability called deep stall. This phenomenon occurs at high angles of
attack where the horizontal empennage is inside the wing air wake and the aircraft cannot
recover from a stall because the horizontal control areas are outside the clean airflow.
However, a good design of the empennage can solve this problem (See Section 4.3.1:
Overall empennage configuration of the report’s attachments).
5.4 Landing gear type and disposition
Landing gear is the undercarriage of an aircraft and is used for either takeoff or land-
ing. The purpose of the landing gear in an aircraft is to provide a suspension system
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during taxi, take-off and landing. The landing gear supports the craft when it is not fly-
ing, allowing it to take off, land, and taxi without damage. There are two landing gear
configurations:
• Fixed: The whole landing gear system, including wheels, is fixed under the fuselage
or wing.
• Retractable: Landing gear stowed in fuselage or wing compartments while in flight.
The aircraft will be designed to fly at top speeds of 500 km/h. A good way to decrease
parasite drag levels is to use a retractable landing gear configuration, taking the wheels
and the whole landing gear structure out of the airflow and thus, decreasing the drag.
Once the configuration has been selected, the disposition of the wheels under the aircraft
must be discussed. There are two possible configurations for the aircraft that is being
developed.
• Taildragger: Composed by two main wheels towards the front of the aircraft and a
single, much smaller, wheel or skid at the rear. Advantages of this type of landing
gear are:
– It supports a smaller part of the aircraft’s weight allowing it to be made much
smaller and lighter than a nosewheel.
– If a tailwheel fails on landing, the damage to the aircraft will be minimal.
• Tricycle: Undercarriage where there are two main wheels under the wings and a
third smaller wheel in the nose.
– Easier to land.
– Prevents the blast of hot, high-speed gases from causing damage to the ground
surface, in particular runways and taxiways.
– Landing gear can be designed so the fuselage is leveled while the aircraft is
parked and during taxi, increasing passengers comfort.
Nowadays, almost all commercial passenger aircraft have been fitted with tricycle landing
gear. Therefore, this configuration will be used.
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6 Conceptual design and performance analysis
The main purpose of this section is to estimate the basic parameters required for the re-
gional aircraft in order to allow the whole team to start developing the project. In other
words, this section is about the conceptual design and the performance analysis.
6.1 General outline of the method for estimating Take-Off gross weight,
WTO, empty weight, WE, and mission fuel weight, WF
The method presented below is based on an iterative process capable of determining the
basic and most relevant aircraft weights. The full completion and explanation of the
iterative process is found in Section 2: Conceptual design and performance analysis
of the Report’s attachments. This is a sum-up of the method applied to a regional turbo
propeller driven airplane.
The general overview of the iterative process consists of seven steps and is presented
below:
1. Determination of the mission payload weight (WPL).
Mission payload weight (WPL) is speciﬁed in the requirements of the project. This
payload weight consists of the following:
• 80 passengers and baggage
• 2 pilots and 2 TCP.
For passengers and crew in a commercial airplane an average weight of 80 kg per
person and 14 kg of baggage is a reasonable assumption [70].
Wpass+baggage = 80⇤(80+14) = 7.520 kg
WPL = 7.520 kg
Wcrew = 4⇤(80+14) = 376 kg
Wcrew = 376 kg
2. Guessing a likely value of take-off weight, WTOguess
Looking at data for similar airplanes it is found the ATR 72, which is a regional
aircraft with similar speciﬁcations. This similar aircraft is a good example to use in
order to estimate weights. The process starts with a lower take-off weight so that the
weight of similar aircraft is achieved when the WTOguess is increased when iterating.
Thus, an initial WTOguess is considered as 10000 kg and is increased 1 kg for each
iteration.
WTOguess = 10.000 kg
3. Determination of the fuel weight, WF
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To determine the fuel used, the flight is split in 8 stages [70]. In each of them a
relation between starting and ending weight has been found according to statistical
data and Breguet’s equation (see Section 2: Conceptual design and performance
analysis of the Report’s attachments) so in the end, a value of Mf f = 0,8321 is
encountered.
After these 8 stages are completed, the value for mission fuel weight (WF) can
finally be determined from:
WF = (1 Mf f )WTO +WFres (1)
Where WFres accounts for fuel reserves which are assumed to be the 30% of the
total fuel used. The justification of this 30% is explained as well in Section 2:
Conceptual design and performance analysis of the Report’s attachments.
4. Calculation of a tentative operational empty weight value, WOE
The following equation is used as all the variables are known.
WOEtent = WTOguess WF WPL (2)
5. Calculation of a tentative value for empty weight , WEtent
Equation 3 is used in this step:
WEtent = WOEtent Wt fo Wcrew (3)
Wt fo is the weight of all unusable fuel and oil that remains inside the aircraft and it
is considered to be a 0,5% of the WTOguess [70].
6. Calculation of the empty weight value, WE
It is known that there is a linear relationship between log10WTO and log10WE. Thus,
it should be easy to obtain WE from WTO. Considering the calculations from Section
2: Conceptual design and performance analysis of the Report’s attachments, it
is possible to develop a easy relation between log10WTO and log10WE.
WE = invlog10((log10WTO A)/B) (4)
Where A = 0,3774 and B = 0,9647.
7. Comparision between the values of WEtent and WE.
A divergence of 0,5% is usually sufficient at this design level. The process keeps
adjusting the new WTOguess until the values of WEtent and WE agree with each other
within the tolerance limits. When they do, the final weights result is obtained as
presented in Figure 7.
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Figure 7: Iterative process.
6.1.1 Results
It can be seen that the iteration process presented in Figure 7 ﬁnishes after 14.098 itera-
tions and the values obtained are:
WTO = 24.389 kg
WE = 11.070 kg
WF = 5.323 kg
These are the values that will be considered to develop the whole project up until the end.
This preliminary values can be accepted by the fact that the results obtained are indeed
quite similar to the weights that other aircraft –such as the ATR-72– manage. Moreover,
the complexity of the process gives a solid and well-constructed mathematical back-
ground that has been validated by several authors among the aerospace ﬁeld, including
Jan Roskam [70].
6.2 Sizing requirements
In these section and the ones that will precede it, the EASA and FAA regulations will be
followed to set a valid sizing point for which the aircraft can be designed. In this case, the
Part 25 (FAA) and the CS-25 (EASA) must be considered as the aircraft is a propeller-
driven artefact with more than 19 seats and the MTOW will be greater than 8.618 kg.
After that, all the initial requirements (see section 3) regarding the take-off distance, cruis-
ing speed, maximum range, etc. are taken into account to solve the main equations based
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on a standardized procedure [70].
In order not to confuse the reader with unnecessary calculation, all the development of the
sizing requirements is explained completely in Section 2: Conceptual design and per-
formance analysis of the Report’s attachments. This way, only the final graphic results
are presented here.
6.2.1 Sizing to take-off distance requirements
Using mainly the take-off field distance –which is indeed given by the initial require-
ments– it is possible to encounter a relationship between the wing loading and the power
to weight ratio for different maximum lift coefficients.
(W/S)TO/((W/P)TO{CLmaxTO
) = 113,7 lbs/ ft2
Assuming some values of CLmaxTO
, possible values of (W/P)TO are found as a function
of (W/S)TO. These results are plotted in Figure 8.
Figure 8: Sizing to Take-Off.
6.2.2 Sizing to landing distance requirements
Similarly to the procedure presented in 6.2.1, a linear relationship between the wing load-
ing and the maximum lift coefficients is found by considering several requirements such
as the landing distance, approach speed, etc.
(W/S)TO = 33,4·CLmaxL
Supposing different CLmaxL
, wing loads during take off can be found as shown in Figure
9.
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Figure 9: Sizing to Landing.
6.2.3 Sizing to cruise speed requirements
It is shown in [57] that cruise speed is proportional to a parameter IP just as:
Vcr µ IP
where:
IP =
✓
W/S
s(W/P)
◆1/3
(5)
Then, by evaluating Equation 5 with the proper data –air density at 7.500 m and cruise
speed of 500 km/h– it is possible again to ﬁnd a linear relationship wing loading and
power to weight ratio. In order to meet the cruise requirement the relation between (W/P)
and (W/S) must be on the right of the curve in Figure 10.
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Figure 10: Sizing to cruise speed.
6.2.4 Sizing to climb requirements
As no speciﬁc climb requirements are given by the customer, it is considered that the
airplane must meet the FAR25 climb requirements which are deeply explained in Section
2: Conceptual design and performance analysis of the Report’s attachments. Also, the
preliminary drag polars are estimated so that the climb requirements are met by Equation
6. Most of the data required is either calculated by mathematical equations or estimated
according to similar aircraft’s data from [70].
CGR+
⇣
CL
cD
⌘ 1
p
cL
=
18.97hp
p
s
W
P
W
S
1/2
(6)
In order to meet the climb requirements the relation between (W/P) and (W/S) must be
beneath the curve. Therefore, the type of climb that restrict most are FAR 25.119 and
FAR 25.121 (go around), to ease the visualization only the mentioned conﬁgurations are
plotted in Figure 11:
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Figure 11: Sizing to climb requirements FAR25.119 and FAR25.121(go-around).
6.2.5 Matching all the sizing requirements
After having established a series of relations between Weight to Power ratio and Wing
loading (both at take-off), it is now possible to ﬁnd a valid design point that considers
them all. To do so, the mentioned relations are plotted in the same graphic (see Figure
12), so that all the requirements are clearly visualized and taken into account.
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Figure 12: Matching of all the sizing requirements
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Taking only the allowable region (marked with green on Figure 12), the design point
should be as high as possible in order to have the minimum required power and not
”waste” any unnecessary power of the engines. On the other hand, this design point
should also admit a wing loading as large as possible so that the aircraft gains stability
towards turbulence.
Observing Figure 12 and considering the mentioned characteristics about the design point,
it is reasonable to choose point P (shown in Figure 12) as the ”best” design point which
yields to: ✓
W
S
◆
TO
= 74 PSF (7)
✓
W
P
◆
TO
= 11 lbs/HP (8)
(cLmax)TO = 2,2
(cLmax)L = 2,6
From Equation 7 and 8 the following data can be obtained:
S = 67,5 m2
P = 4888 HP
It is also interesting and important to note that the obtained results are compared to similar
aircraft. As can be seen in Table 2 the values obtained for wing surface and power are con-
siderably similar to the ones used by ATR-72 and Dash 8 q400, which gives consistency
and solidity to the results obtained.
Airplane S P
m2 HP
Project Eolo 67,5 4.888
ATR-72 61 4.950
Dash 8 q400 64 5.071
Table 2: Similar airplanes
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7 Aircraft’s aerodynamics
After having obtained a reasonable value of the wing surface in the previous chapter 6,
it is now possible to calculate the main parameters of the aircraft’s aerodynamics as
well as the sizing of its elements such as the fuselage, the wing, the tail and the center of
gravity.
Again, it is important to note that the big amount of calculus and deep discussions which
have been made are detailed in Section 4: Aerodynamics of the Report’s attachments
which is advisable to check, so only the finals results are presented below.
7.1 Fuselage
When talking about the fuselage, it is important to state that the main parameters of the
fuselage’s design are obtained according to the initial requirements stated in 3.
The cross section has been designed to have a circular section mainly due to pressur-
ization and structural reasons. After that, it is possible to calculate the diameter of the
fuselage through statistical data (linear dependency with the width of the seats and its
number).
Regarding the interior, the dimensions of all the elements (seats, aisles, exits, baggage
compartments, galley, lavatory) have been taken into account, and in order to fit all of
them, the necessary length of the airplane has been calculated. To make it easier for the
reader, the final sketch– which can also be found in the Drawings document– is included
below in Figure 13. It must be noted that regulations EASA CS-25 and FAA Part
25 have been followed along the whole process so that the final design can be properly
certified.
Once the external dimensions of the fuselage have been determined, it is important to
calculate the fineness ratio, that will tell if the design is efficient enough in terms of aero-
dynamics (Drag coefficient). The conclusion is that a value of 12,34 is a valid result,
which will infer a low value of the fuselage’s drag, so the target proposed is accom-
plished well enough. The main fuselage parameters are shown in Table 3:
Exterior diameter 2.700 mm
Interior diameter 2.400 mm
Cabin length 21,6 m
Fuselage length 33,31 m
Fineness ratio 12,34
Table 3: Main fuselage parameters.
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C
C
21,60
0,890,75
0,75
0,750,60
0,81
0,99
1,460,51
0,56
Lavatory
FrontBaggage
Kitchen
Wardrobe
TypeIDoor
TypeIDoorTypeIDoor
TypeIIDoor
0,42
1,54
0,93
0,44
0,70
R1,20
0,05
0,45 0,65
0,05
0,45
R0,13
0,19
SECTIONC-C
Scale1:25
AA
BB
CC
DD
EE
FF
8
8
7
7
6
6
5
5
4
4
3
3
2
2
1
1
KONDORAIRCRAFTS
TÍTULO:
ESCALA:1:200HOJA1DE1
EoloRegionalAircraft
Cabindesign
Figure 13: Cabin Design
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7.2 Wing
When it comes to the wing’s design, it is necessary first to determine the preliminary wing
configuration –according to the requirements– which is set to:
• Cantilever configuration
• High wing
After that, the general wing parameters (which are: Quarter Chord Sweep Angle (D25),
Aspect Ratio (AR), Taper ratio (l), Incidence angle (iw), Dihedral angle (G), Wing twist
(ew), the High-lift devices and the Command surfaces) must be determined. This is ac-
complished by studying how each aspect of the wing will affect the overall behaviour
of the wing: checking references, coding analysis (Matlab) and comparing with similar
airplanes.
It is advisable to check all the calculations and the proper development of the main geo-
metric parameters from Section 4: Aerodynamics of the Report’s attachments. However,
the final values are shown in the Table 4 and the graphic wing planform is displayed in
Figure 14.
Sweep angle D25 0
Aspect Ratio AR 12
Taper ratio l 0,45
Dihedral angle G 0
Incidence angle iw 5
Wing Surface Sw 67,5 m2
Table 4: Main wing geometric parameters.
Figure 14: Wing planform.
After the wing’s parameters are calculated, it is possible to study and select the most valid
airfoil that is capable of accomplishing those conditions. This way, after an exhaustive
study the chosen airfoils are the following:
• Root: NACA 23018
• Tip: NACA 23012
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On the other hand, high-lift devices (HLD) are essential in certain aircraft to be able
to perform several maneuvers such as Take-off and landing. The extra lift required to
perform these actions is always extracted from the addition of HLD. This way, it has
been proceeded to calculate the values of the dimensions and placement of the flaps of the
wing. The results obtained are presented in Table 5 .
Table 5: Flaps.
Sf 44,25 m2
bf 17,08 m
c(y = 1,1 m) 3,13 m
c(y = 9,638 m) 2,053 m
dFmax ±40 deg
Then, it is possible to calculate also the values of the dimensions and placement of the
ailerons and spoilers, as shown in Tables 6, 7 .
Table 6: Ailerons.
Sa 4,72 m2
ba 7,68 m
Ca/C 0,23
bai 19,3 m
dAmax ±25 deg
Table 7: Spoilers.
bs 17,08 m
c(y = 1,1 m) 2,64 m
c(y = 9,638 m) 2,05 m
Position Ahead of flaps
7.3 Tail
First, similarly to the wing, an initial choice of the tail configuration must be done. It
has been decided to use a T-tail because of its advantage of being kept well out from the
disturbed flow generated by the propeller and the wing and it is stated that it also provides
longitudinal stability and control. Also, an important point in the T-tail is the deep stall
issue that is also covered. T-tail allows to solve this problem in most flight situations of a
regional aircraft, so it is interesting to check Section 4.3 Tail of the Report’s attachments.
It must be taken into account that the horizontal tail must generate a moment to achieve
longitudinal trim. For this reason, a summation of moments of the the main forces that
interact in the aircraft has been done. Moreover, an analysis via XFLR5 –which can be
checked in Section 4: Aerodynamics of the Report’s attachments– is performed in order
to check stability and trim.
The elevator mounted on the horizontal stabilizer and the rudder on the vertical stabilizer
have also been designed to determine the total tail configuration.
The horizontal stabilizer planform and its dimensions with the control surface is shown in
Figure 15 and the vertical stabilizer is shown in Figure 16.
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Figure 15: Horizontal stabilizer.
Figure 16: Vertical stabilizer.
7.4 Weight components and gravity center calculation
Finally, in order to calculate the center of gravity of the airplane –which is crucial for sev-
eral structural and stability purposes – it must be pointed out the weight of each element
that takes part in the aircraft manufacturing.
To do so as a preliminary calculation, similar aircraft’s data has been used to compute
some element’s weights. After that, the center of gravity can be easily determined by
proceeding with a moment’s summation. The ﬁnal obtained values are:
XGC = 17,018 m
ZGC = 2,256 m
A checking of the solution of the center of gravity has been made a and it has been con-
sidered that the position of the aerodynamic center and the distance that the latter must
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have with the center of masses the dimensions are correct, in concordance with the results
obtained in the Tail section.
The whole developed can be found in Section 4.4: Component Weight Estimation and
Section 4.5: Gravity center of the Report’s attachments.
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8 Aircraft propulsion
The aim of this section is to study, analyze and choose an appropriate propulsion system
for the regional aircraft.
To do so, several previous points must be taken into account, the following are some of
them:
• Required cruise speed and maximum speed.
• Required maximum operating altitude.
• Required range.
• Noise regulations.
• Installed Weight.
• Reliability and maintainability.
• Fuel amount needed, cost and availability.
• Specific customer or market demands.
• Timely certification.
8.1 Engine type selection
Before selecting which engine could satisfy best the requirements of the project there are
some aspects, which are also previously explained in 3, that should be taken into account.
Remembering what it is actually explained, the project aircraft design is focused to flight
at low speed and, as a future airline aircraft it will spend most of its life in a cruise
operation condition. For these airplanes requirements, excess thrust is not as important as
high engine efficiency and low fuel consumption.
So, according to the study, which is detailed in Section 5: Aircraft Propulsion of the
report’s attachments, thrust depends on both the amount of gas moved and the velocity, it
is possible to generate high thrust by accelerating a large mass of gas by small amount,
or by accelerating a small mass of gas by a large amount. That is why the usage of a
Turbofan or a Turboprop could be a good election for the project.
Now, when focusing on the requirements of the project and comparing both turboprop
and turbofan engines, shown in Figure 17, the compromise between advantages and dis-
advantages is sought. As the cruising speed is too low to use turbofans, turboprop is the
final choice.
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Figure 17: Gas turbine efficiency [63]. (Note that the x-axis units are in mph, the cruise
speed of the aircraft is between 400-500 km/h ⌘ 250-311 mph).
8.2 Engine configuration
When it comes to the election of the engine configuration, a comparison between the
advantages and disadvantages of each one should be carried out as well as considering
which fits best with the aircraft design.
It is possible to see each configurations’ advantages and disadvantages from Section 5:
Aircraft Propulsion (Table 14) of the Report’s attachments.
From that comparison it is possible to extract that some configurations must be dismissed
because either the engine cannot be installed or they are not useful according to the design.
So there are only two configurations left, under-wing and leading edge configurations.
Finally, the under-wing configuration is the chosen one. Although not recommended
for unprepared runways, it is considered that a regional aircraft operates in normal and
prepared airports, so that this is not an important aspect to take into account. Furthermore,
this configuration causes minimal lifting effects from an aerodynamic point of view, and it
also has easy maintenance access. Instead, the leading edge configuration directly affects
the aerodynamic efficiency of the aircraft, what may be solved by changing other aircraft
aspects but implies unnecessary costs.
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8.3 Number of engines
It is common to need more than one engine to achieve the required power requirements
and for safeness, for example, in case of engine failure. This fact is better explained below.
Engine failure probability and number of engines used
Airplane with: Failure of: 1 Engine 2 Engine 3 Engine
Two engines 2Pef P2
ef not apply
Three engines 3Pef 3P2
ef P3
ef
Four engines 4Pef 6Pef
2 4P3
ef
Table 8: Relation between engine faliure probability and the number of engines used. Pef
is the probability of engine faliure [71].
Table 8 shows the relationship between the probability of engine failure and the number
of engines used. Having more engines, of course, increases the engine failure probability.
On the other hand, if there are more engines, a systematic failure of all propulsion systems
will be less probable. In addition, it is not desirable to use engines that differ in power (or
thrust) in the same airplane except in special cases.
In conclusion, more engines provide more operational safety, but the maintenance cost
to avoid failures in the propulsion plant is higher. Due to the power, safety, reliability
and maintainability requirements it has been decided to incorporate into the design two
engines (twin-engine conﬁguration), one on each semi-wing.
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8.4 Engine selection
The engine selection has been done by scouting the market in order to take an engine
whose maximum take-off power belonged to a specific power range (2.500-2.900 CV),
provided by the design point.
Two engines were found: AI-24T and the PW127 series. The final selected engine is the
PW127, specifically the PW127G for different factors which are further explained in Sec-
tion 5: Aircraft Propulsion of the Report’s attachment.
The basic technical data of the chosen engine is shown in 9:
Overall length Overall width Dry specific Weight Maximum Take off Power (sea level) 5 min.
Shaft power Max. Air T for Rated Power
[mm] [mm] [kg] [kW] [oC]
PW127G 2130 679 484,4 2178 35
Normal Take off Power Maximum Continuous Power
Shaft power
Maximum
Air Temp for
Rated Power
Shaft power
Maximum
Air Temp for
Rated Power
[kW] [oC] [kW] [oC]
PW127G 1973 35 2178 35
Table 9: Basic engine technical characteristics.
If more information about the global technical data is wished to be checked, it is advisable
to go to Section 5: Aircraft Propulsion of the Report’s attachment.
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8.5 Propeller
This section shows the results obtained in the process of choosing the propeller of the
aircraft. The process is shown in more detail in the 5.3 Propeller study section of the
Report’s attachments. Therefore, the process and its results are brieﬂy described here.
The propeller is basically determined by the following parameters: ”Blade-power load-
ing” (Pbl), Diameter (Dp), Number of blades (nb) and Take-off power (Pmax). The follow-
ing equation is used for the selection of the correct parameters.
Pbl =
4Pmax
pnpD2
p
(9)
*The unit system in this equation is as follows: Pmax [hp], Dp [ft] and Pbl [hp/ft2]
The regulations limit the number of blades the propeller must carry. In this case, np = 6.
According to Roskam II [71] the normal range of blade-power-loading (Pbl) for regional
aircraft is: 3,4-5,2 [hp/ ft2]. If using this equation, the following propeller diameter values
are obtained: 4,12 m (for 3,4 Pbl)- 3,33 m (for 5,2 Pbl).
Table 10 shows some propellers that are close to the necessary requirements.
Propeller options
Dp Pmax rpm mprop nb source
Propeller [m] [kW] [sHP] [rpm] [kg] - -
Dowty R408 4,12 3782 5072 1020 252 6 [6] [79]
Dowty R391 4,12 3505 4700 1020 326 6 [46]
Dowty R381 3,81 2786 3736 1100 227 6 [46]
Hamilton 568F-5 3,89 2178 2920 1200 163,7 6 [45]
Hamilton F568-1 3,93 2050 2749 1200 165 6 [46]
Dowty R410 3,65 1864 2500 1200 172 6 [79]
Dowty R352 3,65 1864 2500 1200 172 6 [79]
Table 10: Possible propellers that can be attached to the engine.
In order to make the right choice, it was believed that a decision criteria would be the
most appropriate way. The propeller decision is made by the ordered weighted average
(OWA) method. In this method certain criteria or decisive factors (n) must be deﬁned.
Each item (pi) is evaluated and the relative weight (pixgi) is calculated.
After that, the following formula can be applied:
OWA =
Ân
i=1 pi ·gi
pmax ·Ân
i=1 gi
(10)
Figure 18 shows the selection of the best propeller following the OWA’s formula:
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Figure 18: OWA method Table to select the best propeller.
The last row of each table shows the OWA value, this is an indicator of the ﬁnal choice.
The propeller with the highest OWA value (value marked in blue) is the best option to take
into account. Through the OWA method it has been determined that the best propeller to
attach to the engine is the Hamilton Sundstrand 568F-5. The PW127G engine with
the Hamilton Sundstrand 568F-5 is an existing conﬁguration, for example, in the CASA
C-295M aircraft, therefore, it is a perfectly feasible option.
The basic technical data of the chosen propeller is shown in 11:
Propeller
Diameter (Dp) Maximum continuous power (Pmax) Take off Mass (mprop) No. Blades (nb)
[m] [kW] [sHP] [rpm] [kW] [sHP] [rpm] [kg] -
3,89 2178 2920 1200 2237 3000 1200 163,7 6
Table 11: Technical data of Hamilton Sundstrand 568F-5 propeller.
More detailed information about the propeller can be found in both section 5.3.3 Techni-
cal data: Hamilton Sundstrand 568F-5 of the Report’s attachments or in the Technical
sheets document.
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8.6 Propellant tank
Performance parameters of the engine chosen are confidential. Therefore the values that
are going to be used in the following calculations are taken from a similar aircraft (ATR-
72), this leads to the calculation of the specific fuel consumption.
Once this fuel consumption is known, the Breguet’s equation (11) can be applied to cal-
culate the range of the plane for different take-off weights.
dW
dt
=
g·Cs ·CD
CL · µ
(11)
Solving this equation a plot of the range of the plane against the take-off mass is done.
The results can be seen in Figure 19.
Figure 19: W-R plot.
The operational empty weight of the aircraft is OEW = 11.546 kg which includes the
empty weight of the plane plus the crew weight. If the aircraft takes-off with maxi-
mum payload in maximum range configuration, the reserves fuel weight is WreservesFuel =
1.229 kg, the trip fuel isWtripFuel = 4.175 kg and the maximum payload weightWMAXPayload =
7.439,4 kg which is the weight of 80 passengers and their luggage. Therefore, it can be
proved using Equation 11 that if the aircraft takes off in this configuration of maximum
operational weight of 24.389 kg it will reach a flight distance of 1.614 km which is a
requirement of the project.
If the flight plan requires more range, payload can be replaced for more fuel to a maximum
of 7.571 kg of used fuel plus the 1.229 kg of the reserves. These means that the tanks
have a maximum capacity of 8.800 kg of fuel. The range would be 4.154 km for this
configuration. Finally, some seats could be empty so the take of weight would increase
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reaching a maximum range of 5.420 km if the plane ﬂights with the tanks full and no
payload.
So in conclusion, it can be stated that in terms of volume, tanks have a maximum capacity
of:
Vf =
Mf
rkerosene
=
8.800kg
800kg/m3
= 11m3
(12)
The volume of the fuel tanks obtained through Equation (12) is 11 m3. The wing of the
aircraft has a capacity for the fuel tanks of 19,81 m3 so, it can be stated that the fuel tanks
are feasible with the current wing design.
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9 Aircraft systems
A proper study of the systems that the aircraft carries could not be forgotten in order to
orientate the reader about this inherently essential equipment. The following section
covers the Electric, lighting, instrumental and hydraulic systems of the airplane.
9.1 Electric System
The Aircraft Electrical System is a self contained network of components that generate,
transmit, distribute, utilize and store electrical energy. In this section, a preliminary elec-
tric system design has been done (the different parts of this system have been selected).
The selection has been done based on the ATR-72 electrical main system elements:
9.1.1 Batteries
The aircraft will need a pair of batteries in order to supply the energy required to the ﬂight
instruments, main systems of the aircraft and passenger services. Based on the ATR-72,
the model selected is the G-242 Gill 24 volt Battery. The technical data of the batteries
can be checked in Section 6: Aircraft Systems of the Report’s attachment.
In Figure 20, the selected battery is shown:
Figure 20: G-242 Gill Battery [7].
9.1.2 APU (auxiliary power unit)
The aircraft will need an auxiliary power unit in order to start the engine or provide en-
ergy while the aircraft is on land. The selected one is the APS 3200. It is an Advanced-
technology modular-design airborne auxiliary power unit. The price range is estimated
between $550.000 and $600.000. The technical data of the APU can be checked in Sec-
tion 6: Aircraft Systems of the Report’s attachment.
In Figure 21, the chosen APU is shown:
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Figure 21: APS 3200 Auxiliary Power Unit [89].
9.1.3 Starter Generator Unit
The aircraft will need a starter generator unit (AC) in order to supply the energy required
to the aircraft. The selected one is the Astronics CorePower Brushless Starter Generator
Unit. The price of this generator is conﬁdential, so for more information it is necessary to
contact with Astronics company. The technical data of the Starter Generator Unit can be
checked in Section 6: Aircraft Systems of the Report’s attachment.
In Figure 22, the selected starter generator is shown:
Figure 22: Astronics CorePower Brushless Starter Generator Unit [11].
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9.2 Lighting System
The aim of this section is to provide a minimum knowledge of which type of lights the
aircraft must be fitted of. Some of them are required by regulation, the FAA AC 20-30B
- Aircraft Position Light and Anti-collision Light Installations [2].
Below, the section is divided into external, internal and emergency lights. It is believed
that each aircraft light and its image is not too relevant to be included in this report.
However, if more data and images about the purpose and regulation of each light is wished
to be seen, see Section 6.2: Lighting systems of the Report’s attachments and Technical
sheets of the project. In addition, it is important to highlight that some of the lights have
been chosen taking Airbus A320 as inspiration.
9.2.1 External lights
It is important to emphasize the main purposes of external lights which are:
• Make the aircraft more visible to other aircraft.
• Improve visibility during some flight performances.
• Provide illumination for some specific purpose.
Some of the lights are a regulatory requirement, following the Minimum Equipment List
(MEL), which is approved by the operator’s national airworthiness authorities.[3]
Figure 23 shows the principal external lights.
Figure 23: External lights [15].
9.2.2 Internal lights
Interior lighting provides illumination for instruments, cockpits, cabins, and other sections
occupied by crew members and passengers. Some of the interior lights required in the
aircraft are presented below.
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9.2.3 Emergency lights
An emergency lighting system includes an emergency light control unit, having an ex-
ternal power input and at least one external control input and a plurality of autonomous
emergency light units. Some previous lights that have en external control input are:
• ”EXIT” emergency lights
• Ceiling emergency lights
• Evacuation path marking lights
• Exterior emergency lights
9.3 Instrumental system
In this section the flight instruments of the aircraft is presented. The description of
these systems and instruments can be found in section 6.3: Instrumental system of the
Report’s attachments.
In order to fulfill the basic requirements of the instrumentation system, the information
of ATR 72 aircraft [44] has been used as an inspiration. The flight instruments [73] are
divided into the sections presented below.
• Air Data System:
– Main systems: Air Data Computers (ADC) (2 units as one is the main one and
the other is the standby unit).
– Standby System.
– Controls:
⇤ Air Speed Indicator. (2 units. Main unit and Standby unit).
⇤ Altimeter. (2 units. Main unit and Standby unit).
⇤ TCAS Vertical Speed Indicator.
⇤ TAT-SAT/TAS Indicator.
• Attitude and Heading Reference System (AHRS):
– Main systems: Two main systems (AHRS).
– Standby Systems:
⇤ Standby Electrical Horizon.
⇤ Standby Magnetic Compass.
⇤ Radio Magnetic Indicator (RMI, compass rose, showing magnetic head-
ing).
– Controls:
⇤ Attitude Heading Reference System (AHRS).
⇤ Radio Magnetic Indicator (RMI).
⇤ Standby Horizon.
• Electronic Flight Instrument System (EFIS):
– Primary Flight Display (PFD).
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– Electronic Attitude Direction Indicator (EADI).
– Electronic Horizontal Situation Indicator(EHSI) (Navigation Display).
• Clocks (Show time, elapsed time and chronometer information.)
• Flight Recoders
– Cockpit Voice Recoder (CVR).
– Digital Flight Data Recoder (DFDR):
⇤ Flight Data Entry Panel (FDEP).
⇤ Flight Data Acquisition Unit (FDAU).
9.4 Hydraulic System
The hydraulic system is one of the fundamental systems to provide the correct handling
of the aircraft. The hydraulic system is used to drive the landing gear, wing flaps, wheel
brakes and, ultimately, all control surfaces. Other systems, such as the fuel system or the
lubrication system, can also be considered low pressure hydraulic systems.
These systems combine the advantages of light weight, ease of installation, simplification
of inspection, and minimum maintenance requirements. The hydraulic systems are very
safe, efficient and have good weight-power relation. Despite this, in recent years in the
world of aviation, the hydraulic system is beginning to be replaced by electrical systems.
This is because the latter can offer faster answers and be somewhat lighter.
This section presents and describes the hydraulic system of the aircraft. However, this
system is described exhaustively in Section 6.4: Hydraulic system of the Report’s at-
tachments.
9.4.1 Hydraulic fluid
The fluid is one of the most important parts of the system since it is responsible for trans-
mitting all the power from the pump to the final actuators.
Not only it is responsible for transmitting power, it also performs the tasks of lubricating,
cleaning and preventing corrosion. The most important properties that must be considered
when choosing the fluid are viscosity, chemical stability, flash point and fire protection.
This last point is very important in the world of aviation. The chosen fluid must be able
to withstand extreme conditions and under no circumstances can it suppose a danger to
the safety of the aircraft. In addition, the chosen fluid cannot damage the non-metallic
materials.
In the civil aircraft market, the following types of fluids can be found [85]:
• Vegetable-base fluid (MIL-H-7644). Is essentially castor oil and alcohol.
• Mineral-base fluid (MIL-H-5606). Is essentially kerosene-type petroleum.
• Synthetic fluid. Non-petroleum base fluid. This type of fluid is fire-resistant, and
it is made for use in high performance applications. The most commonly used fluid
of this type is Skydrol R . It is dyed purple for its identification. It has a wide range
of operating temperatures (between -54 oC and 108 oC).
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More information about each fluid is given in Section 6.4.1: Hydraulic Fluid of the
report’s attachments. Only information of the chosen fluid is shown here.
It is important not to mix the different types of fluids due to the difference in composition
(vegetable base, petroleum base and phosphate ester base). Intermixing different types of
fluid will vary their properties and may affect the integrity of certain system components
(seals and hoses).
Finally, for safety reasons and for its previously presented properties, a synthetic fluid
type (Skydroll R ) is chosen as hydraulic fluid.
9.4.2 Distribution System
The distribution system is responsible for housing and transporting the hydraulic fluid that
transmits the power to the actuators. This part of the installation includes hoses, tubing,
fittings, and connectors.
On the plane a great diversity of hydraulic systems is found and each of them must work
under established conditions (pressure). Also all of them must meet the requirements for
which they are designed to and all must be easily identifiable.
Aircraft fluid lines may be either rigid metal tubing, or any of a variety of flexible hoses.
These lines must be able to withstand the pressure conditions under which the fluid inside
is located. Below, we briefly present the types of lines that we can find on the plane.
Basically, two types of hydraulic lines can be found:
• Rigid fluid lines.
• Flexible fluid lines.
Section 6.4.2: Distribution System of the Report’s attachments document details each
type of distribution system deeply.
The rigid lines pipe is used in stationary applications and where long and straight sections
are possible. The lines are usually made of aluminium-alloy or corrosion-resistant steel
lines. Rigid lines can be bended, but it must be done in a correct way (specified in Section
6.4.2: Distribution system of the report’s attachments).
Each of the different types of tubes is identified by the following color code. See in the
following Table 12:
Aluminium Alloy Number Color of Band
1100 White
3003 Green
2014 Gray
2024 Red
5052 Purple
6053 Black
6061 Blue and Yelloy
7075 Brown and Yellow
Table 12: Aluminium alloy number and its identification.
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Flexible fluid lines will be used to connect stationary parts to moving parts. The flexible
hoses are composed of:
• The inner liner (made of Neoprene, Buna -N, Butyl or Teflon).
• Reinforcement (made of cotton, rayon, polyester, carbon-steel wire or stainless-
steel wire braid).
• Outer cover (made of fabric, rubber or stainless-steel braid).
Obviously, the installation of the hoses must be carried out conscientiously and must be
correctly positioned and subject.
The different aircraft systems must be correctly identifiable. Figure 24 shows the universal
codes for each type of system.
Figure 24: Universal codes for each type of system [85].
9.4.3 Main hydraulic system components
This section presents the most important elements of the aircraft hydraulic system. In
summary, the following components are found:
• Tank.
• Reservoirs.
• Distribution system.
• Valves.
• Actuators.
More specifically, hydraulic power generation and transmission can be produced by:
• Power Transfer Unit.
• Quick Disconnect Couplings.
• Inline AC Motorpump.
• Pressure Switches and Transducers.
• Reinforced Hose Assemblies.
• Engine-Driven Inline Pump.
• Relief Valves.
• RAT Driven Inline Pump.
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• Generator Control Unit.
• Hydraulic Motor-Driven Generator.
• Start Valve.
Some of the hydraulic system components (generation and distribution components) are
shown below. These have been extracted from the EATON R catalog. [24]. For more
information, description or technical data see Section 6.4.3: Main hydraulic system
components of the Report’s attachment.
• Power Transfer Unit (MPHV3-115-1C).
• AC Motorpump (MPEV-032-15).
• Hydraulic Motor-Driven Generator (CMV3-022-EA3C).
• Engine-Driven Pump (PV3-240-10C).
• Ram Air Turbine Pump (PV3-115-EA1D).
• Fuel Boost Pump Type 8410 and Canister Type 8411.
• Park Brake Relief Valve (HR6B9-002-EA5A).
• Stainless Steel Reinforced Hose Assemblies (666/AE641).
9.4.4 Hydraulic System Architecture
This section describes the hydraulic system as a whole and its architecture. It shows
where the hydraulic power is obtained and how it is transmitted to the different elements
and actuators of the plane. The different actuators and control elements of the system
are also shown. It is intended to give an understanding of the system in a simple and
schematic way. For more details, see Section 6.4.4: Hydraulic system architecture of
the report’s attachments.
It should be taken into account that the airplane’s hydraulic system is the system that al-
lows the plane to be controlled, therefore, if it fails, the plane becomes uncontrollable and,
most likely, ends up in a catastrophic accident. Hydraulic systems, fortunately, are usually
very safe systems with a low failure rate. However, comprehensive safety measures must
be applied.
For safety and security reasons, the most common is to have several hydraulic systems in
parallel or redundant. Thus, if one of the systems fails, we have another fully functional
system capable of taking control.
The aircraft’s hydraulic system will consist on three hydraulic subsystems related between
them. Table 13 shows the elements controlled by each subsystem.
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Subsystem 1 (Yellow) Subsytem 2 (Green) Subsystem 3 (Red)
RH Aileron LH Aileron RH and LH Aileron
Rudder Rudder Rudder
RH Elevator LH Elevator RH and LH Elevator
Flight spoiler Flight Spoiler RH and LH MLG Actuator (Landing gear)
RH and LH DN Lock Assist Ground Spoiler RH and LH MLG Uplock (Landing gear)
RH and LH Outbd Brake Flaps NLG Uplock (Lock Door Actuator)
Flaps NLG DN Lock
NLG Steering
NLG Actuator
NLG DP Uplock
RH and LH Inbd Brake
Table 13: Elements of each hydraulic subsystem.
If it is represented schematically, the graph shown in Figure 25 is obtained. As it can be
seen, the architecture of the hydraulic system of the aircraft is based on the explanations
and diagrams of the bibliography (Tucker, Bill L. [85]).
Figure 25: Hydraulic system schematic [85].
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10 Structural Design
Once all the elements for the design of the aircraft are deﬁned, it is possible to deliver
a conceptual proposal for the structure of Project Eolo. This concept is based on the
internal structures of similar airplanes, and a study of the state of the art of the plane
structures can be found in Section 7: Structural state of the art of the Report’s attach-
ments.
This means that a structural analysis covering the loads and stress calculations is out of
this project’s scope. This sort of study would be performed in future stages of the design,
probably, by performing different structural simulations to verify the correct performance
of the structure proposed in the following pages.
First, a view of the complete structure design is presented (see Figures 26 and 27), fol-
lowed by an individual explanation of each part.
Figure 26: Global view.
Figure 27: Upward view.
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10.1 Fuselage
The ﬁrst part considered in the structure is indeed, the fuselage, since it is the center and
main part of the structure and where all the other components are attached.
In order to design the internal structure of the fuselage, two different parts have been
clearly differentiated: Nose and Cabin, each with a different type of structural architec-
ture.
10.1.1 Nose structure
Different geometries for the nose have initially been considered, but thanks to the aero-
dynamics CFD study done in section 11, the most optimized shape of the nose has been
obtained. This proposal for the nose has a complex external shape, and that is why a rather
unusual semimonocoque type of structure is decided to be used for this part (see Figure
28). This means that a complex system of bars– instead of the typical system of frames
and crossbars–, is designed and recovered with plates that are adapted to this unusual
shapes.
Figure 28: Internal structure of the nose.
Using the measurement tool of Solid Works, the volume of skin and bars is determined in
order to be able to specify the amount of material needed in 10.4. So, the results obtained
from this tool are:
• Bars: 0,115 m3
• Skin: 0,218 m3 (43,833 m2)
• Floor: 0,059 m3
10.1.2 Cabin Structure
Since this component does have a constant and rather simple shape, a regular semimono-
coque type of structure has been considered. In this case, the internal structure is formed
by 14 circular frames, 6 T-shaped crossbars and 12 V-shaped crossbars (see Figures 29
and 30). The T-shaped crossbars are positioned in the upwards and downwards area of
the cabin. They are the ones that will take the largest quantity of loads and stress in com-
parison with the V-shaped crossbars, which are situated in between the T-shape ones and
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serve as a support for the main bars. Also, they are a way to reduce weight since their
section is smaller.
Furthermore, the window section has been reinforced by a thin plate rather than using
crossbars, with the window spaces incorporated in the middle of the plate. Finally, in
between this window plates a space for the door has been considered and a ﬂoor has been
incorporated.
Figure 29: Global view.
Figure 30: Front view.
Once the structure is already presented, a calculation of the volume of each part is included
to be able to specify the amount of materials needed in 10.4. Again, the latter tool of Solid
Works has been used to determine the total volume. In the end, the results have been the
following:
• Frames: 0,438 m3
• T-shape crossbars: 0,220 m3
• V-shape crossbars: 0,155 m3
• Floor: 0,722 m3
• Window plate: 0,867 m3
• Skin (external and internal): 1,76 m3 (352,996 m2)
10.2 Wing
Following the concept of the design of the fuselage, the internal structure design of the
wing also uses a semimonocoque type of structure (see Figure 31). This structure is
centered around two main I-shaped crossbars, one straight bar positioned in the middle
line of the wing (so as to be completely straight) and another crossbar situated in the
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quarter chord line, which has a certain angle in order to follow the sweep angle of the
wing. Furthermore, to have the required shape, 26 ribs have been situated across the
length of the wing in different distances, being in the tips and root closer to each other
and more far away in the middle of each semi wing.
It is also important to take into account that the central part of the wing (where the union
with the fuselage is placed) has been modeled as a solid area to consider that this part will
be thicker than the rest of the wing, since at the present stage of the design this union is
yet to be considered and this modeling is meant to take this into account when calculating
the materials needed for building the wing.
Moreover, the introductions of ﬂaps and ailerons has been thought in the total wing design,
but these components have only been modeled as solids since as the union between the
wing- fuselage as well as the systems to deploy them are yet to be considered.
Finally, it is important as well to point out that the ribs have been emptied in some non-
critical zones of their surface in order to reduce weight.
Figure 31: Internal structure of the wing.
Once the internal structural design is known, the volume of each component has been
determined using the same methodology explained previously. So, the results have been
the following:
• Ribs: 0,782 m3
• Crossbars: 0,690 m3
• Skin: 1,820 m3 (127,15 m2)
10.3 Tail
Once the wing has been completely deﬁned, the tail structure is ready to be studied. This
component has been divided into three parts.
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10.3.1 Empennage Structure
This components is responsible for the union between the fuselage and the vertical tail
(since the present design uses a T tail type as mentioned in section 7). So, following
the steps of the cabin design, a semimonocoque type of structure has been used for the
empennage structure (see Figure 32), which consists of 7 ribs that decrease in section held
together by 8 square crossbars.
Figure 32: Internal structure of the empennage.
In order to later determine the amount of material needed, the known software tool is used
again. So, the results have been the following:
• Ribs: 0,109 m3
• Bars: 0,178 m3
• Skin (internal and external): 0,724 m3 (86,25 m2)
10.3.2 Vertical Tail Structure
The next component is the vertical tail, since is the following component attached to the
already studied empennage structure.
Furthermore, for the conﬁguration considered in this design, the vertical tail is also the
union between the horizontal tail and the empennage. So, for this design (see Figure 33),
the same architecture as the one used in the wing has been applied. This means that
the structure has two main crossbars, one positioned in the middle of the wing and the
other positioned in the quarter of the chord, both with certain angle matching the one of
the leading edge.
These bars are used to held together the 6 ribs positioned across the tail to have the shape
needed, all of which has been emptied in non-critical zones of their surface in order to
reduce weight. Also, as occurred in the wing, the rudder has been considered but not
its internal structure, so it has been modeled as a solid.
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Figure 33: Internal structure of the vertical tail.
Once the structure has been described, a calculation of the volume of each component is
presented below to determine the total amount of material needed for the structure of the
plane. The results obtained are:
• Ribs: 0,100 m3
• Crossbars: 0,120 m3
• Skin: 0,398 m3 (25,37 m2)
10.3.3 Horizontal Tail Structure
Finally, the horizontal tail design is presented (see Figure 34). This design also has a
certain similarity with the one used for the wing structure, in a way that both are
formed by a certain number of emptied ribs joint by two main cross bars.
In this case however, while 8 ribs are still emptied, the crossbars are a bit different. For
instance, there is a straight crossbar situated in the middle of the horizontal tail, but the
crossbar that goes in the quarter of the chord is, in this case, also straight and does not go
until the end of the semi tail but to the rib before the end. This is done because the sweep
angle of the tail is a way too large to try to make a crossbar that follows this direction.
Furthermore, the ailerons have been considered but, as in the wing design, their internal
structure has been modeled as a solid.
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Figure 34: Internal structure of the horizontal tail.
For the volume used for each component, the results show:
• Ribs: 0,039 m3
• Bars: 0,070 m3
• Skin: 0,244 m3 (20,78 m2)
10.4 Materials choice
Once all the structure has been presented, the choice of the materials is indeed, a crucial
matter in this section.
In the early stages of the project, the idea of using highly new and advanced materials for
the construction of the aircraft was presented. However, due to the difficulty of finding
reliable and specific data of such materials and considering the fact that most of them are
either conceptual or untested in real life situations, this choice was discarded.
The conceptual idea that has been applied is based on considering materials that are
currently used in the new planes such as the Boeing 787 or the Airbus A320. This fact
will also make this plane better than the rest of its competitors such as ATR, since these
planes were designed several years ago and use older materials. This means that Project
Eolo disposes of lower weight (Project Eolo weights 17% less than ATR-72) and lower
fuel consumption. So, applying this idea it has been stated that 3 main materials will be
used for 3 different parts of the structure.
For the ribs, a material capable of handling certain amount of loads is inherently needed,
but since they will not be the main structural component, the key point of this material
does not have to be its strength but its low weight. For this latter reason, the material that
is going to be used to generate all the ribs, floors and plates is an Aluminum Lithium
alloy. Aluminum has the effective properties –from an aerospace point of view– and these
are increased by the addition of Lithium since it reduces the density of the alloy and in-
creases its Young modulus [56].
The total volume required for this alloy would be 3,115 m3.
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For the crossbars, a strong and quite reliable material that is able to handle the main
loads and stresses is needed. The main reason is because it will be the main structural
component. Therefore, taking all this into account, the material chosen is the titanium
alloy Titanium 5553, which is an emerging alloy used mainly for aerospace purposes
on highly demanding structural applications. This material performs excellently in these
kind of situations thanks to its strength to weight ratio. The reason is because it weights
relatively less compared to others materials that are capable of handling the same amount
of stress [82].
The total volume required of this material is 1,545 m3.
Finally for the skin material, the use of a composite material has been considered since
the beginning in order to reduce weight. After researching among the materials used in
the most recent presented planes, either glass fiber and carbon fiber type of composites
were considered, but after taking into account that carbon fiber can interfere with certain
communication devices, its use was dismissed. Therefore, the material that will be used
for the skin is GFRP (Glass Fiber Reinforced Polymer). This material is used for many
aerospace applications such as the skin of the aircraft thanks to its corrosive resistance
and high strength to weight ratio [77]. The total volume required of this material is
5,163 m3 (656,379 m2).
10.5 Landing gear
Finally, the last structural element of the aircraft which is yet to be defined is the landing
gear. This section is aimed at giving a proper definition of it. As said in 5, two main
configurations are possible:
1. Fixed or non-retractable
2. Retractable
As a general rule, if the cruise speed of the airplane is above 275 km
h , a fixed landing gear
increases drastically the drag. [71]
The customer’s requirements impose a cruising speed of 450 500 km
h , so a retractable
landing gear is chosen.
The usual disposition of the landing gears in regional aircraft is under the wings, near to
the fuselage. This configuration implies that sometimes long landing gear systems cause
several failures [55]. This way, the design of the a landing gear will be under the fuselage
but shorter, and will have an extra wheel under the nose.
Two main criteria are used to assure the aircraft stability and avoid possible failures.
1. Tip-over criteria
2. Ground clearance criteria
Tip-over criteria
Two main criteria must be accomplished
• 15o criteria: It assures the stability of the longitudinal axis. The angle between the
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gravity center and the axis of the landing gear must be of 15o as minimum.
• 55o criteria: It assures the stability of the lateral axis considering the separation
between the main landing gear wheels. The y angle must be less than 55o for all
the load configurations.
Figure 35: Longitudinal and lateral tip-over criteria [71].
Ground clearance criteria
The lateral ground criteria is necessary to comply cross wing landing limitations. It as-
sures there is no risk for the engines to touch the ground. The angle shown at Figure
36 must exceed the 5o. Keeping in mind that our aircraft has high wing configuration, it
won’t be a problem to fulfill the criteria.
Figure 36: Lateral ground clearance criteria [71].
In order to define the longitudinal ground clearance criteria, figure 37 is presented. It
assures that the tail never touches the runway when it starts to lift-off.
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Figure 37: Longitudinal ground clearance criteria [71].
Keeping in mind the criteria presented above, it is possible to define the landing gear
configuration. At Annex A of the report’s attachments, the code of the landing gear
configuration that fulfills the criteria is presented.
The results obtained are the following:
• Nose-Main wheels distance: 16,31 m
• Distance between main wheels: 4,2 m
• Vertical distance between fuselage end and main wheels: 0,4 m
• Longitudinal tip-over criteria: g = 29,46o > 15o
• Lateral tip-over criteria: y = 54,55 < 55o
• Longitudinal ground clearance criteria: q = 11,73o > qLOF
1
• Lateral ground clearance criteria: f = 52o > 5o
After that, the size of the wheels is also a remarkable parameter to consider but to do so,
it is necessary to compute the maximum static load per strut.
1The lift-off angle (qLOF ) cannot exceed 11,73o during the take-off maneuver [71].
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Figure 38: Geometry for static load calculation [71].
Taking the distances shown in Figure 38 and solving Equations (13) and (14), the static
load per strut is obtained.
Pn =
WTO ·lm
lm +ln
(13)
Pm =
WTO ·ln
(lm +ln)·ns
(14)
The following values are used:
WTO = 24.389 kg
ln = 14,81 m
lm = 1,5 m
ns = 2
The value of ns corresponds to the number of struts at the main landing gear.
After solving the equations:
Pm = 11.073 kg
Pn = 2.243 kg
It is obvious that the most part of the load is supported by the main landing gear
(91%), but more precisely the calculus shows that:
Main landing gear
There are two struts with two wheels per strut. Four wheels in total.
Pm
4
= 12.208 lbs
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Nose landing gear
There is only one strut with two wheels.
Pn
2
= 2.473 lbs
It is important to comment that the results have been converted to lbs. because the wheel
catalogue of Michelin Aircraft Wheels presents the maximum wheel load in these units.
The Part II of Roskam’s book, gives an estimation of the wheel dimensions for the main
and the nose landing gear [71, pag. 224].
The wheel selection has been done taking as reference these estimations and choosing
the wheels that fulﬁll the load requirements of the Michelin catalogue. The selection is
presented in Table 14.
Landing gear Diameter (mm) Width (mm) Maximum static load (lbs)
Main 876,3 265,43 13.000
Nose 495,3 171,45 2.500
Table 14: Wheels sizing.
In conclusion, as can be seen in the maximum static load column, the six wheels can
support the load. Also, these tyres have a maximum speed ratio of 190 mph, which do
not exceed by any means the velocity of 115 mph during Take-off.
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11 CFD Analysis
This chapter is aimed at trying to give a critical validation to the aerodynamic calculation
and structural assembly done in previous sections. On the other hand, CFD techniques
also help discuss some design parameters such as the one presented in 11.1.
11.1 Fuselage Nose Analysis
The shape of the nose of the fuselage has a great influence on the aircarft’s drag. A
preliminary shape has been copied from a study [43] which is designed to reduce the drag
compared to the common nose used in aircraft like ATR-72.
A computational fluid dynamics simulation using OpenFOAM, which is an open source
CFD program, has been done to analyse two new fuselage noses that have been designed
to reduce drag.
Figure 39: Pressure results. From left to right, geometry proposed in [43], variant 1,
variant 2. Scale is in Pa
As seen in Figure 39, the geometry proposed in [43] is not as good as it is said because
of the huge pressure drop that occurs at the sides of the nose, where an abrupt change
of angle happens. The second shape has been designed aiming to decrease this pressure
drop. To accomplish it, the nose is sharper and smoother. Finally, another design has been
proposed, maintaining the smoothness of the second but with less sharp-edged contours.
While variant 1 reduces the drag at 3% compared to the one proposed in [43], the last one
is better, as it decreases drag levels at 6%. Therefore, variant 2 is chosen.
11.2 Cruise Analysis
Project Eolo is a regional aircraft designed to fly at cruise speeds of 500 km/h at 24500ft.
For this flight conditions the aircraft has been analysed using OpenFOAM to verify some
parameters and visualize the airflow around the fuselage, wings and empennage. The
results are presented in Figures 40, 41 and 42. Note that the aircraft has been simu-
lated without propeller to avoid using rotational meshes, which could have increased the
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runtime of the whole simulation. However, as it is a ﬁrst estimation, the simulation is
considered acceptable.
Figure 40: Pressure results. Scale is in Pa.
Figure 41: Pressure results. Scale is in Pa.
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Figure 42: Velocity field XZ plane. Scale is in m.
As it can be seen in Figure 40, the wing’s performance is satisfactory. It can be seen
that flow accelerates at the top side of the wing and therefore, a difference of pressure
between the topside and the downside is generating lift. The undercarriage case, where
the main landing gear is folded, is not generating remarkable changes in the velocity field
or pressure. Thus, drag levels are not increased substantially.
Finally, the lift force has been calculated during post-process proving that the aircraft
is capable of generating 245.200 N of lift. That is, the plane has a lift coefficient of
CL = 0,68 at an angle of attack a = 0o. This means that the aircraft could lift 25.000 kg,
which is very close to the total weight of Project Eolo (24.389 kg) –see section 6.1–.
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Preliminary design of a short-haul regional aircraft

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