POLYTECHNIC UNIVERSITYOF CATALONIA
AEROSPACE ENGINEERING
Engineering Projects
3 seat Light-Sport Aircraft development | AlOn Contents: Report
Group: G06-AE-2018/19-Q1
Delivery date: 20/12/2018
Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. Contents
1 Introduction 7 1.1 Aim ...... 7 1.2 Scope ...... 7
2 Requirements 9 2.1 Legal and technical specifications ...... 9 2.2 Economical requirements ...... 9
3 Background 10
4 State of the art 11 4.1 Less than 600 kg aircrafts ...... 11 4.1.1 Atec 322 Faeta ...... 11 4.1.2 TL-2000 Sting S4 ...... 12 4.1.3 TL-3000 Sirius ...... 12 4.2 Pipistrel Taurus M ...... 13 4.2.1 Alexander Schleicher ASG 29 E ...... 13 4.3 More than 600kg air crafts ...... 14 4.3.1 ONE Air craft ...... 14 4.3.2 Sling 4 ...... 14 4.4 Review ...... 15
5 Main alternatives selection and development of the best one 16 5.1 Aerodynamics alternatives and selection methods ...... 16 5.1.1 Airfoil ...... 16 5.1.2 Wing ...... 18 5.1.2.1 Rectangular ...... 18 5.1.2.2 Elliptical ...... 18 5.1.2.3 Trapezoidal ...... 19 5.1.2.4 Delta ...... 19 5.1.2.5 Wingtips ...... 19 5.1.3 Wing configuration criteria ...... 19 5.1.3.1 Unweighted Average Method ...... 20 5.1.3.2 Ordered Weighted Average Method ...... 20 5.1.4 Tail ...... 20 5.1.5 Tail configuration selection ...... 21 5.1.6 Decision making ...... 23 5.1.7 Fuselage ...... 23 5.1.8 Fuselage selection criteria ...... 23 5.1.8.1 Unweigthed Average Method ...... 24 5.1.8.2 Ordered Weighted Average Method ...... 24 5.2 Aerodynamics development of the chosen solution ...... 25 5.2.1 Airfoil ...... 25 5.2.2 Wing ...... 27
R 1 G06-AlOn LSA 3 seats | Project report 5.2.2.1 Stall and pre-defined velocities ...... 29 5.2.2.2 Flight condition curves ...... 29 5.2.3 Control surfaces definition ...... 30 5.2.3.1 Elevator ...... 31 5.2.4 Ailerons ...... 32 5.2.4.1 Rudder ...... 33 5.2.4.1.1 Final wing configuration ...... 34 5.2.5 Fuselage ...... 34 5.2.6 Tail design and static stability behaviour ...... 35 5.2.7 Final plane configuration ...... 36 5.2.7.1 Parasite Drag ...... 36 5.2.7.1.1 Wheels parasite drag ...... 37 5.2.7.1.2 Tube parasite drag ...... 37 5.2.7.2 Drag correction ...... 38 5.2.7.3 Design parameters ...... 38 5.2.7.4 Range study ...... 39 5.2.7.5 Flight envelope ...... 39 5.3 Final plane analysis ...... 42 5.3.1 Efficiency ...... 42 5.3.2 Static Stability ...... 42 5.3.3 Dynamic Stability ...... 43 5.3.3.1 Longitudinal Mode 1 ...... 43 5.3.3.2 Lateral Mode 1 ...... 43 5.3.3.3 Lateral Mode 2 ...... 44 5.4 Structures ...... 45 5.4.1 Wing ...... 45 5.4.1.1 Initial approach ...... 45 5.4.1.2 Beam analysis ...... 45 5.4.1.2.1 Twist analysis ...... 47 5.4.1.3 Ribs analysis ...... 48 5.4.2 Fuselage ...... 49 5.4.3 Tail ...... 50 5.4.3.1 Elevator beam ...... 50 5.4.3.2 Fin beam ...... 51 5.4.3.3 Ribs ...... 51 5.4.3.4 Final Result ...... 51 5.4.4 Landing gear ...... 52 5.4.5 Windows ...... 53 5.5 Power Plant ...... 55 5.5.1 Engine ...... 55 5.5.2 Propeller ...... 57 5.5.3 Gear Box ...... 57 5.5.4 Engine Mount ...... 57 5.5.5 Auxiliary Motor Items ...... 58 5.6 Avionics and Systems ...... 59 5.6.1 Basic Equipment ...... 59 5.6.2 Flight and Navigation Instruments ...... 59 5.6.3 Powerplant Instruments ...... 60 5.6.4 Miscellaneous Equipment ...... 61 5.6.5 Safety Belts and Harnesses ...... 61 5.6.6 Controls ...... 62 5.7 Design ...... 63 5.8 Business ...... 65
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5.8.1 Facilities ...... 65 5.8.2 Human Resources ...... 65 5.8.3 Marketing Plan ...... 66 5.8.3.1 Marketing Costs ...... 66 5.8.4 Study of potential costumers ...... 67 5.8.5 Study of advertisement ...... 67 5.8.6 Design of the Marketing Campaign ...... 68 5.8.7 Initial Investment ...... 68
6 Economic feasibility analysis 69 6.1 Break Even Point ...... 69 6.2 Pay-Back Analysis ...... 70 6.2.1 Estimation of decay rate k ...... 70 6.2.2 Pay-Back Time, Return of Investment (ROI) and Internal Rate of Return (IRR) ...... 71
7 Environmental impact analysis 74
8 Safety considerations 75
9 Planning and scheduling for the follow-up of the project 77
10 Conclusions and recommendations 79
R 3 List of Figures
4.1 Atec 322 Faeta ...... 11 4.2 TL-2000 Sting s4 ...... 12 4.5 Alexander Schleicher ASG 29 E ...... 13 4.6 ONE Air craft ...... 14 4.7 Sling 4 ...... 15
5.1 NACA 4415 ...... 16 5.2 NACA 23015 ...... 16 5.3 Davis B24 ...... 17 5.4 FXS 02196 ...... 17 5.5 Seats configuration ...... 23 6 5.6 Airfoil study for different configuration at Re = 1.7 · 10 ...... 26 5.7 Wing efficiency ...... 28 5.8 CL − α ...... 28 5.9 Lift distribution along the wing span ...... 29 5.10 Efficiency for cruise, flap and ailerons configurations ...... 30 5.11 CL for cruise, flap and ailerons configurations ...... 30 5.12 Elevator Airfoil ...... 31 5.13 Hmom Elevator ...... 32 5.14 Aileron Airfoil ...... 32 5.15 Hmom Ailerons ...... 33 5.16 Rudder Airfoil ...... 33 5.17 Hmom Rudder ...... 34 5.18 CL − (α) ...... 35 5.19 E − α ...... 35 5.20 CD experimental data ; Source: researchgate.net ...... 37 5.21 Comparison between both behaviours ...... 38 5.22 Weight-Range diagram ...... 39 5.23 Gust-Airspeed envelope diagram ...... 41 5.24 Efficiency of the final configuration ...... 42 5.25 Static stability of the final configuration ...... 42 5.26 Longitudinal Mode 1 ...... 43 5.27 Lateral Mode 1 ...... 44 5.28 Lateral Mode 2 ...... 44 5.29 Loads distribution and approximations ...... 45 5.30 Moment diagram for n=4 ...... 45 5.31 Square section parameters ...... 46 5.32 Stress analysis ...... 47 5.33 Displacement results ...... 47 5.34 Tip section and twist stress ...... 48 5.35 Sandwich ribs ...... 48 5.36 Beam and ribs ...... 49 5.37 Fuselage structure ...... 50
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5.38 Elevator moment diagram ...... 50 5.39 Elevator FEM analysis ...... 51 5.40 Fin moment diagram ...... 51 5.41 Fin FEM analysis ...... 52 5.42 Beams and ribs ...... 52 5.43 Front leg of the landing gear...... 53 5.44 Half of the rear landing gear...... 53 5.45 Wankel motor IAE50R – AA ...... 56 5.46 Standard frame mount ...... 58 5.47 Dashboard in the cockpit ...... 59 5.48 Initial sizing ...... 63 5.49 Weight and C.G...... 63 5.50 C.G. Position ...... 64 5.51 Floor plan of the chosen office ...... 65 5.52 Map location of the chosen office ...... 65 5.53 Scheme of organization of the members of the project ...... 66 5.54 Division of the Marketing’s Cost ...... 66 5.55 Airplane Shipments Worldwide (1995–2017) ...... 67 5.56 Companies that collaborate in the Advertising Plan ...... 68 5.57 Final Logo for Alpha One ...... 68
6.1 Break-Even Point ...... 70 6.2 Net Present Value for the first seven years of the project ...... 72
8.1 Sequence of BRS parachute ...... 75
R 5 List of Tables
4.1 Atec 322 Faeta main specifications ...... 12 4.2 TL-2000 Sting S4 main specifications ...... 12 4.3 TL-3000 Sirius main specifications ...... 13 4.4 Pipistrel taurus M main specifications ...... 13 4.5 Alexander Schleicher ASG 29 E main specifications ...... 14 4.6 Alexander Schleicher ASG 29 E main specifications ...... 14 4.7 Alexander Schleicher ASG 29 E main specifications ...... 15
5.1 UA for the airfoils ...... 17 5.2 Airfoil OWA ...... 18 5.3 Wing Unweighted Average Method ...... 20 5.4 Wing Ordered Weigthed Average method ...... 20 5.5 Simple Hierarchy Method application ...... 22 5.6 Unweighted Average Method application ...... 22 5.7 Ordered Weighted Average Method application ...... 23 5.8 Unweighted Average Method application(fuselage) ...... 24 5.9 Ordered Weighted Average Method application(fuselage) ...... 24 5.10 Final wing configuration data ...... 34 5.11 Stable distances configuration ...... 36 5.12 Final fin configuration data ...... 36 5.13 Final elevator configuration data ...... 36 5.14 Final wing configuration data ...... 38 5.15 Final fin configuration data ...... 38 5.16 Final elevator configuration data ...... 38 5.17 Control surfaces main sizes ...... 39 5.18 Results of buckling analysis ...... 48 5.19 Rotax 912 Vs AE50R comparison ...... 55 5.20 OWA results ...... 56 5.21 Relative weight table ...... 56 5.22 PRESS results ...... 56 5.23 Custom propeller characteristics ...... 57 5.24 HAUTECLAIRE propeller characteristics ...... 57 5.25 Motor frame characteristics ...... 58
6.1 Variable Costs Specifications ...... 69 6.2 Fixed costs specifications ...... 70 6.3 Cash Flow and Updated Cash Flow for the first 7 years of the project ...... 71 6.4 Return of Investment and Net Present Value for the first 7 years of the project . 72
R 6 Chapter 1
Introduction
1.1 Aim
To design an aircraft with a maximum capacity of 3 people which follows the LSA (Light- Sport Aircraft) regulations and requirements set by regulatory bodies such as the EASA.
1.2 Scope
The scope of the project will be defined in terms of the different fields involved in the de- sign and some manufacture process of the airplane. The project team will focus on each of these fields or working areas, which are aerodynamics, structures, power plant, systems and avionics, CAD design, economics and regulations requirements.
The aerodynamic development consists of obtaining the geometry of the main parts of the aircraft such as wing, tail, fuselage and the main control surfaces. In order to achieve it a drag, lift, static and dynamic analysis will be done.
As for the structural field, the analysis will cover the following areas: beam and wing ribs sizing, buckling body study, beam and tail ribs analysis and landing gear sizing. In addition, a research of different materials will be necessary to select the most appropriate ones.
The power plant tasks will be focused on the accurate sizing and selection of the main ele- ments, which include the study of different propellers, engines, gearboxes and engine benches for a further selection of the most suitable power plant.
The systems and avionics of the aircraft will be chosen after an evaluation of the compulsory and complementary flight instruments. The evaluation will determine which components, peripherals and electric power station will be suitable in terms of weight, complexity and compatibility.
Moreover, a CAD design will show the 3D graphical representation of the main parts of the aircraft. In addition, the blueprints of the main views with the key parameters will be in- cluded. Once the aircraft is finished, a whole aircraft representation will be done in order to create a 3D mock-up. Additionally, the aircraft figure drawing and parts will be printed in a poster.
The economic analysis will cover the estimation and calculus of the project cost, including human resources, materials, software and renting. Furthermore, a feasibility analysis will cover income to cost ratio and payback time.
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The requirements and regulations of each area will be revised in order to ensure that the airplane follows European and American certification standards for an LSA, except for the restriction on passenger number, which is a mandatory requirement from our client.
R 8 Chapter 2
Requirements
2.1 Legal and technical specifications
To develop this project, it is compulsory to list the requirements, both purchaser and the current legislation, to make our product apt to enter the market. First, as a condition from the customer, it must be considered an LSA, except for the number of passengers. Therefore, to follow EASA [1] regulations, it must comply with the following criteria:
• Maximum take-off mass of 600 kg.
• Maximum stalling speed in the landing configuration of no more than 45 knots (83,34 km/h).
• A maximum seated capacity of three people.
• A single, non-turbine engine fitted with a propeller.
• A non-pressurized cabin.
In addition, to make this aircraft also adequate to perform in the United States, the FAA[2] [3] requires additional terms:
• Fixed landing gear (except for a glider or for aircraft intended for operation on water).
2.2 Economical requirements
Despite not having any imposed economical requirement, it is always a must for the econom- ical part of the project to have a budget as small as possible. For instance, the final price of the plane can not exceed that of a conventional LSA. For this measure, the project will have, although not strictly imposed, some economical limitations, in order to make Alpha One eco- nomical feasible. As it will be discussed on subsection 5.8 Business Plan and on section 6 Economical Feasibility, the average LSA price is somewhere between 150.000 eand 200.000 e. Thus, all the departments of the project will have to restrict their work development in such a way so that the final price lays in between these margins.
R 9 Chapter 3
Background
A Light Sport Aircraft (LSA), to be considered as such, must meet the EASA regulations pre- viously mentioned. One of them specifies a maximum seating capacity of no more than two people, including the pilot. In other words, until nowadays all LSAs available in the market is design for a maximum of two passengers, there is not any LSA with bigger capacity. There could be planes for three people but they are not certified as LSAs because, possibly, apart from the maximum capacity requirement, they do not meet the weight requirement estab- lished in a maximum of 600 kg.
Our project pretends to cover this need by designing an aircraft for three passengers, includ- ing the pilot, that follows all the EASA regulations except the maximum capacity. Consid- ering that the plane will fulfill all the other requirements (maximum weight, maximum stall velocity, one non-turbine engine with a propeller and non-pressurized cabin), it could be con- sidered an LSA, but for three people.
The addition of a third passenger involves two main aspects to consider. On the one hand, the fuselage needs to be extended in order to place the additional seat and other elements re- quired, such as a safety belt and other security systems. On the other hand, three passengers instead of two and a bigger fuselage make the maximum weight requirement a critical aspect of the design process. To avoid exceeding the maximum take-off weight of 600 kg, it is spe- cially important to design a high efficiency aircraft. Due to this reason, the design of this new version of LSA is inspired in the aerodynamics of gliders, which provides high efficiency and low structural weight. Besides, it is also necessary to design an optimum structure with light materials but guaranteeing the resistance and security of the aircraft for different flight con- ditions. Finally, the mandatory aviation systems, the non-turbine engine and the propeller should be accurately selected to accomplish both the requirements and their functions in the plane.
R 10 Chapter 4
State of the art
As seen before, the most restrictive requirements are the MTOW of 600 kg and the third pas- senger. With the purpose of giving a quick view on how other manufacturers dealt with these, we will see some aircrafts with a weight below this limit and couple that surpass it but cover the passenger requirement. 4.1 Less than 600 kg aircrafts.
4.1 Less than 600 kg aircrafts
On one hand, we will see three certified LSAs. LSAs price range from 100.000 to 150.000 dol- lars depending on which engine and avionics have been equipped. Since most manufacturers use same engines and avionics, we have decided to remark some technical and performance details of few aircraft to see what common characteristics we can find. Thus, we have chosen two Czech manufacturers: Atec and TL-Ultralight. Atec only has a certified LSA and it is the 322 Faeta. TL-Ultralight has two LSA options with different wing configuration: theTL-2000 Sting S4 and the TL-3000 Sirius.
On the other hand, we think that to accomplish the 600Kg MTOW restriction with 3 people on board, a change in the bases of the aircraft design is needed. One of the heavier parts of the plane is the power plant plus the amount of fuel in the tanks. If the plane was more efficient it would need less power and, in addition, less fuel to maintain a constant speed. Gliders are the most efficient planes so we will also talk about the main characteristics of these planes
4.1.1 Atec 322 Faeta Atec created Faeta looking for a plane that could be certified as light aircraft. The results are the 321 Faeta, certified in UL and S-LSA categories, and the 322 Faeta, certified in LSA cat- egory.
The 322 is a two-seater low-wing plane with tricycle land- ing gear and completely made out of composite materi- als. With this strong but low-weight structure, it achieves an empty weight (EW) of 340 kg and maximum take-off weight (MTOW) of 583 kg, which results in maximum pay- Figure 4.1: Atec 322 Faeta load (MPL) of 243 kg.
Moreover, Atec offers 80, 100 and 115 HP Rotax engines with Fiti propellers. It has also two 50L tanks, achieving a maximum flight range of approximately 800 nm and flight endurance
R 11 G06-AlOn LSA 3 seats | Project report of 8 hours at cruise speed.
Wing span 9,6 m Wing area 10,1 m2 Length 6,2 m Height 2,0 m Stall speed (full flaps/flaps retracted) 32,8 kts / 42,7 kts Cruise speed 114 kts 145,2 Never exceed speed kts
Table 4.1: Atec 322 Faeta main specifications
4.1.2 TL-2000 Sting S4 Sting S4 is the fourth version of TL-Ultralight aircraft. They used their customer’s opinions to improve the previous ver- sion.
It is a two-seater low-wing plane as well, with a tricycle land- ing gear. Its structure is mostly made of carbon fibre, which makes it strong and light. Wing skin and fuselage is made from a sandwich of carbon fibre and foam core, glued with epoxy. The result is an EW of 354 kg and a MTOW of 598 kg. Figure 4.2: TL-2000 Sting s4 Therefore, Sting S4 achieves an MPL of 244 kg, which is al- most the same as the 322 Faeta.
Like Atec, TL-Ultralight offers different engine options, all of them made by Rotax. However, the propeller is provided by PowerMax. It has a 77 L fuel tank with a 22L optional tank in each wing and it reaches a maximum flight range of approximately 750 nm and flight en- durance of 6 hours at cruise speed.
Wing span 9,12 m Wing area 11,1 m2 Length 6,45 m Height 2,0 m Stall speed (full flaps/flaps retracted) 34 kts / 39 kts Cruise speed 115 kts 164 Never exceed speed kts
Table 4.2: TL-2000 Sting S4 main specifications
4.1.3 TL-3000 Sirius The Sirius follows the Sting S4 concept of having a high-wing configuration, instead of a low- wing, to win in-cabin space. It results in an EW of 344 kg and a MTOW of 598 kg, which means the Sirius exceeds the Sting S4’s MPL by 10 kg, with an MPL of 254 kg.
With 128L, it has a slightly bigger fuel capacity than the Sting. Consequently, it overcomes its “brother” and achieves a maximum flight range of approximately 800 nm and flight en-
R 12 G06-AlOn LSA 3 seats | Project report durance of 6.4 hours at cruise speed.
Wing span 9,4 m Wing area 11,26 m2 Length 6,97 m Height 2,3 m Stall,speed,(full,flaps, 34 kts / 43 kts Cruise speed 115 kts 138 Never exceed speed kts
Table 4.3: TL-3000 Sirius main specifications
4.2 Pipistrel Taurus M
The Taurus M is a two-seat motorized glider made by Pip- istrel. The whole structure of the plane is made from com- posite materials. It has a MTOW of 450Kg and an EW of 285 Kg.
It carries a retractable two stroke 53 HP motor with a 30L fuel tank.
Wing span 14.97 m Wing area 12.33 m2 Length 7.27 m Height 1.41 m Stall speed (full flaps/flaps retracted) 34 kts / 38 kts Cruise speed 122 kts 122 Never exceed speed kts
Table 4.4: Pipistrel taurus M main specifications
4.2.1 Alexander Schleicher ASG 29 E The ASG 29 E is a one-seater motorized glider made by Alexander Schleicher. Its structure is made of composite materials. It has a MTOW of 600Kg and an EW of 325 Kg.
It has a retractable two Stroke 24 HP motor with a 10,5L fuel tank.
Thanks to its large aspect ratio (30.9), it is very ef- Figure 4.5: Alexander Schle- ficient, and it can fly with a smaller engine than an icher ASG 29 E LSA.
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Wing span 18 m Wing area 10.5 m2 Length 6,69 m Height - Stall speed (full flaps/flaps retracted) - / 39 kts Cruise speed - kts 146 Never exceed speed kts
Table 4.5: Alexander Schleicher ASG 29 E main specifications
4.3 More than 600kg air crafts
With more than 600 kg and out of the LSA’s definition we can find some aircraft that may cover our purchaser’s necessities. Therefore, we found it convenient to study some of them and see whether it would be possible to modify them to become certified as an LSA.
4.3.1 ONE Air craft ONE has several configurations but the one that we are inter- ested in is the 2+1.
It is a three-seater low-wing plane with tricycle land- ing gear. With a composite structure, like the air- craft mentioned before, it has an EW of 415 kg and a MTOW of 750 kg that leaves us with an MPL of 335 kg. Figure 4.6: ONE Air craft It has a 135L fuel tank and a 100 HP 912 Rotax engine that makes the ONE achieve a maximum flight range of 700 nm and flight endurance of approxi- mately 5 hours at a speed of 120 kt.
Since we are looking for a 600 kg plane with the same capabilities as the ONE, it might be a good reference for our own design. Our target could be to reduce its EW to near 300 kg, like 2-seater LSAs, and keep MPL near 300 kg.
Wing span 9.65 m Wing area 12.47 m2 Length 6,97 m Height 2.2 m Stall speed (full flaps/flaps retracted) 37 / - kts Cruise speed 120 kts Never exceed speed - kts
Table 4.6: Alexander Schleicher ASG 29 E main specifications
4.3.2 Sling 4 The Airplane Factory is the company that designed the Sling. It is a low-wing plane with a tricycle landing gear. It is made of 6061 aluminium alloy with some composite elements in critical places.
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There are 4 versions of the Sling and one of them is certified as an LSA. This one shows similar fea- tures and performances to what we have seen before. Other versions are sold as kits that can be built at home. The Sling 4 is the one that interests us the most. Figure 4.7: Sling 4 It is a four-seater with similar design to the LSA version but adapted for the extra weight. It has a MTOW of 920 kg with an EW and MPL of 460 kg.
It has a 175L fuel tank that feeds a 115 HP Rotax 914 engine to achieve a maximum flight range of approximately 900 nm and flight endurance of 7 hours.
Like the ONE, this might be a good reference. Even though its EW is higher, it could be re- duced by changing materials into light composites. Furthermore, the maximum crew would be reduced to 3 so MPL could be reduced by about 90 kg. Moreover, maybe even more, if the fuel tank was reduced.
Wing span 9.97 m Wing area 12.47 m2 Length 7,70 m Height 2.44 m Stall speed (full flaps/flaps retracted) 42 kts / 47 kts Cruise speed 120 kts 140 Never exceed speed kts
Table 4.7: Alexander Schleicher ASG 29 E main specifications
4.4 Review
After these examples, there are two clear ways to face our problem. First option is using LSAs as reference. We have seen that our 3 certified examples have some common characteristics:
• Composite structures
• EW of near 350 kg and MPL of 250 kg.
• Maximum range of about 750 nm.
• Rotax engines
We want to add a third passenger so we must ensure a crew weight of about 261 kg. This leaves us with fuel and baggage weight of 24 kg or, in other words, a fuel tank of near 36L if we ignore baggage. This would take our plane to a lower maximum flight range than any other aircraft of its category, so we should find a way to convert some EW into PL in order to be competitive in LSAs market.Second option is using gliders as reference. From them, it can be concluded that our airplane needs to be efficient to have a small motor with small fuel con- sumption. To achieve it, the aspect ratio needs to be high and the three passengers must be seated in a row, reducing drag forces. The engine will be fixed on the front to reduce weight, compared to the retractable system of the gliders. In addition, the structure and many other parts must be made from composite materials.
R 15 Chapter 5
Main alternatives selection and development of the best one
5.1 Aerodynamics alternatives and selection methods
5.1.1 Airfoil A study of the following 4 airfoils has been done:
Figure 5.1: NACA 4415
Figure 5.2: NACA 23015
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Figure 5.3: Davis B24
Figure 5.4: FXS 02196
In order to decide which airfoil provides the best perfomance, an Unweigthed Average Method and also and Ordered Weigthed Average Method have been applied. The next tables show both methods.
Unweigthed Average Method
Unweighted Average Method Davis B-24 NACA 23015 NACA 4412 FXS0-2196 CRITERIA P P P P Efficiency 4 3 4,5 5 Cl 3,5 5 4 3 Cd 3,5 4,5 4 3 Cm 5 5 2 2 Stall perf 4 5 5 4 SUM 20 22,5 19,5 17 Max 25 25 25 25 UA 0,8 0,9 0,78 0,68
Table 5.1: UA for the airfoils
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Ordered Weigthed Average Method:
Ordered Weighted Average Method Davis B-24 NACA 23015 NACA 4412 FXS0-2196 Weight CRITERIA P PXG P PXG P PXG P PXG G Efficiency 4 20 3 15 4,5 22,5 5 25 5 Cl 3,5 17,5 5 25 4 20 3 15 3 Cd 3,5 17,5 4,5 22,5 4 20 3 15 1 Cm 5 25 5 25 2 10 2 10 4 Stall perf 4 20 5 25 5 25 4 20 2 SUM 100 112,5 97,5 85 Max(PxG) 125 125 125 125 OWA 0,8 0,9 0,78 0,68
Table 5.2: Airfoil OWA
Both methods show that the NACA 23015 is the best one among the 4 airfoils analysed. Hence, this is the airfoil selected and the one that will be used to do the simulations of the wing.
5.1.2 Wing The wing plant form is a very important aspect in design. It affects the velocity, efficiency, bending moment, fuel tanks, cost... and so on. There are different types of wing plant forms and each of one is used in specific application.They can be classified in 4 major groups used in airplanes:
• Rectangular
• Trapezoidal
• Elliptic
• Delta
5.1.2.1 Rectangular Is the simplest wing plant form and consequently the less expensive to build. However, it has a not desire aerodynamic and structural behaviour. Nevertheless, is widely used in small general aviation because the aerodynamic demands are not very strict and the low cost of fabrication.
5.1.2.2 Elliptical Is the wing plant form with less induced drag. It has a constant Cl distribution which, with an elliptic chord distribution, results in elliptical lift distribution. This behaviour is suitable for aerodynamics and structures. However, it is not very used today because of its price, which is the highest. It also has a not desire stall behaviour due to its constant Cl distribution, at certain angle of attack the whole wing stalls at once. To solve this problem swept was implemented in WWII emblematic planes like the Spitfire.
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5.1.2.3 Trapezoidal Is the most used in heavy transport planes such as commercial planes. With a good design can have almost the same behaviour of the elliptic ones but in a short range of angles of attack, where the cruise angle of attack should be in. In addition, they are more easy to produce and, consequently, less expensive than elliptic ones. However, the stall behaviour is not suitable. With very narrow trapezoidal wings the stall starts at the tip where the ailerons are often located. In commercial planes are used with swept to increase the cruise speed and with some torsion to counter the stall non-desired behaviour.
5.1.2.4 Delta Delta Wing plant form is the trapezoidal concept brought to the limit where the tip chord is 0. It is used in fighter jets in addition to swept because of its ability to increase the airplane speed in exchange of efficiency.
5.1.2.5 Wingtips The wingtips are able to reduce the induced drag by making a softer transition between the higher pressure Intrados and the lower pressure Extrados at the tip of the wing. Two types of wingtips have been studied:
• Winglets: These devices have a complex geometry because most of the time have high inclination angles which made them difficult to build.
• Elliptic wingtips: These devices consist in an elliptic chord distribution at the tip of the wing. The geometry is much simpler than winglets.
5.1.3 Wing configuration criteria Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efficiency for lower weight.
The comparison is based in four main criteria:
• Efficiency: it measures ratio between the functionality provided by the wing and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the wing. A higher value in this criterion indicates a lighter structure.
• Cost: it is related with the difficulty to build each wing, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
• Stall performance: it defines how good the stall performance is in each wing config- uration, when and how it starts and how it evolves. A higher grade in this criterion indicates a better performance.
For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.
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5.1.3.1 Unweighted Average Method
Unweigthed Average Method CRITERIA Trap Ellip Rec Trap +Elip final Trap + Wing Efficiency 3 3 2 5 5 Weight(struc) 4 3 4 3 2 Cost 4 2 5 4 3 Stall perf 5 2 2 5 5 SUM 16 10 13 17 15 Max 20 UA 2 0,5 0,65 0,85 0,75
Table 5.3: Wing Unweighted Average Method
5.1.3.2 Ordered Weighted Average Method
Ordered Weigthed Average Method Trapeziodal Elliptic Rectangular Trapeziodal + wingtip Trapeziodal + winglet CRITERIA P PxG P PxG P PxG P PxG P PxG Efficiency 3 12 3 12 2 8 5 20 5 20 Weight(struc) 4 12 3 9 4 12 3 9 2 6 Cost 4 4 2 2 5 5 4 4 3 3 Stall perf 5 10 2 4 2 4 5 10 5 10 SUM(PXG) 16 38 10 27 13 29 17 43 15 39 Max(PXG) 20 50 50 50 50 50 OWA 2 0,76 0,5 0,54 0,65 0,58 0,85 0,86 0,75 0,78
Table 5.4: Wing Ordered Weigthed Average method
As it can be seen in both tables the best option it he trapezoidal plus a wingtip at the end. Also it can be noticed that the simple trapezoidal is better than the trapezoidal with wingltes when using the OWA. So the final wing configuration is the trapezoidal plus the wingtips.
5.1.4 Tail The first step to design the tail is the evaluation of the different options, which are: conven- tional tail, T tail, cruciform tail, dual tail, triple tail, V tail, inverted V tail, inverted Y tail, twin tail, boom tail, high boom tail and multiple-plane tail.
From all the options available, only conventional tail and T tail are commonly used, the other ones can be rejected due to complexity and, consequently, high cost and weight. A detailed explanation of advantages and drawbacks of every option is presented in the attachments.
In order to select the best tail configuration, three decision-making methods have been ap- plied: Simple Hierarchy Method, Unweighted Average Method and Ordered Weighted Av- erage Method.
The alternatives evaluated are conventional tail and T tail, due to their common application in LSAs and gliders, and cruciform tail, because it is a hybrid between the other two options and it seems to provide also a good performance. A description of these three configurations is described next, before the application of the decision methods.
• Conventional tail design The conventional tail configuration is an aft tail design (rear-mounted tail) which is
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considered as the simplest tail configuration. The horizontal stabilizer does not produce any load on the vertical stabilizer, consequently, weight and complexity can be reduced. Its overall performance (trim, stability and control) is presumably acceptable. It is also one of the lightest configurations. That’s why it is the most common configuration, especially in LSAs. The weak point of this design is the fact that the horizontal stabilizer is located in the wing wake, causing a loss of efficiency.
• T tail design Another usual tail configuration in General Aviation airplanes is the T-tail, which is an aft tail configuration with the horizontal stabilizer located on top of the vertical fin. Its main advantage over the conventional design is that during cruise flight the tailplane remains above of the regions disturbed by the wing wake, downwash, wing vortex and engine prop wash; thus, it provides a higher aerodynamic efficiency. The lower effects from the wings and engine also lead to a diminution in tail vibrations, greatly reducing its fatigue and increasing its durability. Nonetheless, the bending moment created by the elevator is also higher, so the vertical tail structure becomes heavier. Besides, its main disadvantage appears when flying at a high angle of attack, when the turbulent flow separated from the wings might incise upon the elevator and might result in a complete loss of the aircraft longitudinal control, situation known as deep stall.
• Cruciform tail design This tail design is a hybrid variation of the two previous designs. The horizontal em- pennage is located higher than in the conventional tail, so that it is away from the wing wake and the propeller flow, but it is not as high as in T tail configuration. With this de- sign the lower part of the stabilizer and the rudder receive undisturbed airflow due to the lifting force of the horizontal stabilizer. It is important to have undisturbed airflow on the rudder, especially to recover from spins. Although this configuration does not improve significantly the main strengths of its predecessors; it significantly reduces its major drawbacks, especially the deep stall distress.
5.1.5 Tail configuration selection Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efficiency for lower weight. The comparison is based in four main criteria: • Efficiency: it measures ratio between the functionality provided by the tail and the in- crease of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the tail. A higher value in this criterion indicates a lighter structure.
• Cost: it is related with the difficulty to build each tail, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
• Stability and control: it defines the capability of the tail to develop its main function, and it is related to the performance of the plane. It also considers the facility of the tail to be immersed in the turbulent wake coming from other parts of the plane, which the control surfaces less effect. A higher grade in this criterion indicates a better performance. For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.
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Simple Hierarchy Method
The scale used to evaluate each criterion goes from 1 to 10, and the threshold to reject inade- quate alternatives is 6. Then, the grades and results can be seen in the following table:
ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efficiency 4 10 8 Weight 10 6 8 Cost 8 6 4 Stability and control 6 8 6
Table 5.5: Simple Hierarchy Method application
The threshold set in 6 means that the alternatives that have one or more grades under this number are rejected. Analyzing the efficiency, the conventional tail is discarded; considering the cost of manufacturing, the cruciform is also discarded. In this way, the best alternative is the T tail.
Unweighted Average Method The new scale applied goes from 1 to 5, but the grades are proportionally calculated from the previous method in order to follow the same procedure for the different methods. The results obtained are shown in the next table:
ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efficiency 2 5 4 Weight 5 3 4 Cost 4 3 2 Stability and control 3 4 3 SUM 14 15 13 UA 0,7 0,75 0,65
Table 5.6: Unweighted Average Method application
This method considers the sum of all grades for each configuration to calculate an unweighted average; this means that the four criteria are considered equally important. This procedure indicates that the best alternative is the T tail.
Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a specific weight for each criterion, which indicates the importance or weight that will have in the final result. The results obtained can be seen in the following table:
This method is possibly the most accurate among the ones used, due to the fact that each criterion is weighted in order to consider its importance. According to this method, the best alternative is again the T tail.
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ALTERNATIVES Conventional Tail T-Tail Cruciform Tail WEIGHT CRITERIA P PxG P PxG P PxG G Efficiency 2 8 5 20 4 16 4 Weight 5 15 3 9 4 12 3 Cost 4 8 3 6 2 4 2 Stability and control 3 3 4 4 3 3 1 SUM (PxG) 34 39 35 OWA 0,68 0,78 0,7
Table 5.7: Ordered Weighted Average Method application
5.1.6 Decision making In this case, the best alternative is the same for all methods, so undoubtedly the final decision is the T tail configuration. The next step is to determine the main parameters of the tail in order to provide the necessary stability and control with the minimum increment of drag.
5.1.7 Fuselage In order to evaluate all the variables and choose the one with the best performance, the Un- weighted Average Method (UA) and the Ordered weighted Average Method (OWA) have been applied. The pictures below shoe the different alternatives:
(a) (b)
Figure 5.5: Seats configuration
5.1.8 Fuselage selection criteria In order to apply this method it is needed to define the main variables to be evaluated for each configuration. The comparison criteria will be based on:
• Efficiency: it measures ratio between the functionality provided by the fuselage and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
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• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the fuselage. A higher value in this criterion indicates a lighter structure.
• Cost: it is related with the difficulty to build each fuselage, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
• Wing position: it defines how big is the area where the union between the wing and the fuselage is going to be make. A Higher value in this criterion means a higher area, thus a better union.
5.1.8.1 Unweigthed Average Method The scale applied goes from 1 to 5, where a score of 5 in any variable shows the best possible value and a score of 1 shows the worst. This method does not take into account the weight of the criteria evaluated in relation to the other configuration. The obtained results are shown on: Table 5.8:
Unweigthed Average Method Criteria Seats in line One plus two Efficiency 5 4 Weight 4 5 Cost 3 4 Wing position 4 5 SUM 16 18 Max 20 20 UA 0,8 0,9
Table 5.8: Unweighted Average Method application(fuselage)
As can be seen on Table 5.8 the best option in the case is the One plus two configuration which offers better wing position, lower weight and better structural performance although the efficiency is worst.
5.1.8.2 Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a specific weight for each criteria, which indicates the importance or weight that will have in the final result. The results obtained can be seen in the following table:
Ordered Weighted Average Method Seats in line One plus two Weight Criteria P PxG P PXG G Efficiency 5 10 4 8 2 Weight 4 16 5 20 4 Cost 3 3 4 4 1 Wing position 4 12 5 15 3 SUM(PxG) 41 47 Max(PxG) 50 50 OWA 0,82 0,94
Table 5.9: Ordered Weighted Average Method application(fuselage)
As it can be seen in Table 5.9 the best option is again the one plus two configuration. Tak- ing this into account the final fuselage configuration will be the one plus two which is the
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Fig. 5.5(a).
5.2 Aerodynamics development of the chosen solution
The development of the different alternatives have been done using XFLR5, an analysis tool for airfoils, wings and planes which operate at low Reynold Number. Xflr5 allows:
• XFoil’s Direct and Inverse analysis capabilities
• Wing and fuselage design capabilities based on the Lifting Line Theory, Vortex Lattice Method or on 3D Panel Method. It also allow us to define the inertial plane and to introduce extra drag in order to simulate a more real scenario.
The following obtain results have been done using the software mentioned before.
5.2.1 Airfoil As it have been saw on the previous section the selected airfoil was the NA Using a multi- threadbach analysis it have been simulated the next airfoil for a determined range of the Reynolds number:
104 < Re < 5 · 106 (5.1) Following the base airfoil the next modified airfoils in order to fit in all the requirements and get the best aircraft performance.
• NACA 23015:it will be the root chord airfoil, use it until the geometric torsion starts.
• NACA 23013:after the geometric torsion ends.
• NACA 23012: tip and elliptic wingtip.
• NACA 0008: airfoil used mainly on the tail and it’s control surfaces.
Also the deflected airfoils have been analysed in order to simualte the flap and the control surfaces behaviour and them performance.The deflected studied configuration are:
• NACA 23015, flap: 25% of the chord and δ = 20o.
• NACA 23013, aileron: 25% of the chord and δ = 10o
• NACA 0008, fin: 50% of the chord and δ = 10o
The following graphics show the Cl, the polar, the Cm and the Efficiency distribution of the 6 mentioned airfoils at Re = 1.7 · 10 . These graphics show the different behaviours of the airfoils depending on their configuration.
R 25 G06-AlOn LSA 3 seats | Project report 6 10 · 1.7 = e R Airfoil study for different configuration at Figure 5.6:
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As it can be seen on the Fig. 5.6 there is a clear difference between the air foils which do have a deflected flap and those which do not. It can be seen that the ones which have a deflected flap have a bigger Cm from those which do not have it. This is fine and makes sense because that is what it is expected to happen. Also looking at the efficiency diagram it can be seen that the ones without flap have a bigger around the cruise or the normal flight operations α0s. And the last expected behaviour is that the Clmax grows and the αmax decreases when going from the air foil without flap to the one that have one.
5.2.2 Wing Ones the air foil has been chosen and analysed and the wing plant form has been selected, it is time to start the wing study.
There have been analyzed 3 different configurations in order to gather the main flight config- urations which are:
• Cruise
• Take-off
• With the ailerons deflected
These configurations were studied in order to define the max velocity in each option because those had limitations set by EASA.
The wing main parameters that there have to be taken into account are:
• Efficiency: it has to be as high as possible taking into account that it will have to be build and it has structural limitations.
• CLmax it is needed in order to evaluate the stall velocity.
• CLc: it will show the lift distribution along the span, so it can be defined the exact point where the stall starts.
The dihedral angle has not been taken into account in this previous stage because this param- eter only modifies the static equilibrium of the plane. It will be evaluated on a previous stage together with the tail design.
As it can be seen on the before section the trapezoidal wing plus an elliptical wingtip has been selected because of its better performance in front of the others.
The following figures show both the efficiency and the and the CL − α curves:
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Figure 5.7: Wing efficiency
Figure 5.8: CL − α
As it can be seen on the Fig. 5.7 it has a very good maximum efficiency value and a reasonable range of α where the aircraft is able to flight. The wingtip smooths the tip vortex and increase the efficiency. It also makes the construction a little bit more difficult bur the advantages are bigger than the drawbacks.
In order to ensure that the stall starts exactly where it should be. It has been purpose a dif- ferent solution than the ones that can be usually found on other aircrafts. Because of the elliptical wingtip the usual solution of creating a geometrical torsion at the tip it really com- plicated the construction of the wing. So it has been purpose to define a geometrical torsion at the middle of the wing in order to modify the whole lift distribution making sure the stall starts at the middle point of the wing span.
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Figure 5.9: Lift distribution along the wing span
The Fig. 5.9 shows the difference between the wing with and without the geometrical torsion. It can be noticed the increase in the middle point due a torsion of 3o at 3.5m from the root. To ensure a good stall performance it’s really important in order to ensure the security of the passengers and the pilot. Without geometrical torsion the stall starts practically at the same time in the whole wing.
5.2.2.1 Stall and pre-defined velocities Due to EASA limitations there are some velocities that need to be equal or under the limi- tations of the ASTM F2245-15, which defines the specifications and performance of a Light Sport Airplan (LSA). Because of those limitations was needed to increase the wing surface in order to decrease the stall velocity. Initially the wing configuration was:
• Croot = 0.8 m
• Ctip = 0.6 m These main chord sizes were not enough in order to ensure the minimum stall speed which was 44kts. This velocity includes a flap deflection but even though with a 20 o flap deflection was not enough. So in order to decrease the stall velocity the surface had to increase. Be- cause of the wingtip was already defined and changed the tip chord would have supposed to change the wingtip also the modifications were made to the root chord. So, the final con- figuration that ensured the maximum stall performance is:
• Croot = 1 m
• Ctip = 0.6 m
• Hf lap = 0.75c
o • δf lap = 20
5.2.2.2 Flight condition curves The next picture shows the main graphics for the different wing configurations mentioned before.
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Efficiency:
Figure 5.10: Efficiency for cruise, flap and ailerons configurations
CL:
Figure 5.11: CL for cruise, flap and ailerons configurations
5.2.3 Control surfaces definition The control surfaces allow the pilot to govern the plane, this means that angles of yaw, pitch and roll can be controlled. The three control surfaces are:
• Elevator: its task is to control the pitch.
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• Ailerons: their task is to control the roll, however, it induces a moment that also affects the yaw due to the difference of drag produced by the difference of lift that both wings have. This effect is called adverse yaw.
• Rudder: its task is to control the yaw, however, it induces a moment that also affects the roll because the rudder is higher than the center of mass, the rotation point. This effect is called adverse roll.
In order to size the different control surfaces a similarity study has been done. The plane chosen was the glider Taurus from Pipistrel because its similar characteristics.
5.2.3.1 Elevator From the similarity study it can bee seen that the elevator should be a third of the horizontal tail’s surface. This means that Alpha-One should have a 0,5m2 elevator. In order to make it easier to construct, the elevator will have the same span that the horizontal empennage, 4m; so the chord will be 0.125m to fullfil the two conditions.
The hinge moment was calculated assuming that the elevator was a flap of 0,33 chord with 15o of deflection:
Figure 5.12: Elevator Airfoil
Assuming this, XFLR-5 is capable to compute the non-dimensional hinge moment which was the following:
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Figure 5.13: Hmom Elevator
To compute the maximum hinge moment the vne should be used and a Hmom of 0,025:
2 Moment = 0.5 · ρ · vNE · Se · ce · Hmom (5.2)
Moment = 0.5 · 1.225 · 82.32 · 0, 5 · 0, 125 · 0.025 = 6.48Nm (5.3)
5.2.4 Ailerons In the Taurus from Pipistrel, each aileron has a surface of one tenth of the wing. In the de- signed plane this equals to 0.6m2. The span of the aileron will be 4m with 0,15m chord and will be located at the end of the wing.
Figure 5.14: Aileron Airfoil
To obtain the hinge moment the same method will be used. The next figure shows the aileron was a flap of 0,25 chord with 15o of deflection:
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Figure 5.15: Hmom Ailerons
To compute the maximum hinge moment for each aileron, the vne should be used and a Hmom of 0,010:
2 Moment = 0.5 · ρ · vNE · Sa · ca · Hmom (5.4)
Moment = 0.5 · 1.225 · 82.32 · 0.6 · 0.15 · 0.010 = 3.73Nm (5.5)
5.2.4.1 Rudder In the similar plane, the rudder has a surface of one half of the fin. In the designed plane this equals to 0.4m2. The span of the rudder will be the whole fin, 1,2m, and 0,4m chord.
Figure 5.16: Rudder Airfoil
In order to calculate the hinge moment, the same method will be used. Here we can see the airfoil with a flap of 0.5 chord with 15o of deflection:
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Figure 5.17: Hmom Rudder
To compute the maximum hinge moment the vne should be used and a Hmom of 0,065:
2 Moment = 0.5 · ρ · vNE · Sr · cr · Hmom (5.6)
Moment = 0.5 · 1.225 · 82.32 · 0.4 · 0, 4 · 0.065 = 43.15Nm (5.7)
5.2.4.1.1 Final wing configuration Ones all the parameters have been correctly defined the next table shows the final wing con- figuration:
Wing span (m) 16 Wing area (m2) 12,424 Wing load (kg/m2) 48,295 Croot (m) 1 Ctip (m) 0,6 Aspect ratio 20,606 Taper Ratio 10 Efficiency 40,045 Tilt angle (o) 2 v(m/s) 43,6
Table 5.10: Final wing configuration data
5.2.5 Fuselage The main problem about the fuselage configuration is the unusual number of people and also the maximum take off weight. A configuration of three people is an unusual one which does not allow a the typical LSA or glider two seats configuration.So the casuistry is the next:
• 3 seats configuration
• Relative position of the wing and the tail
• Maximum efficiency
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• Costs • Weight and structural facilities
As it has been said before the better solution for the fuselage is the one plus two configura- tion. A picture of the whole airplane can be seen in Fig. 5.5 (a).
Once the fuselage is closed it has been done an anlysis in order to see the performance of the whole air plane. The next figures show both the efficiency and the CL, of the airplane with the final fuselage configuration:
Figure 5.18: CL − (α)
Figure 5.19: E − α
5.2.6 Tail design and static stability behaviour Once the wing and the fuselage has been defined it is time to design the tail and study the static stability behaviour of the plane.
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On the previous section it has already been discussed the different tails and which one was the best one for this case. Now we the tail is going to be sized using some references like the gliders analyzed on the State of the art chapter of this report.
In order the plane to be statically stable in front of a pertubation the plane has to create a negative momentum in order to correct the air plane direction.
To ensure the negative moment the derivative of the momentum coefficient regarding α has to be negative:
∂C M < 0 (5.8) ∂α The thing is, there is not the perfect defined value for each plane. The more negative the slope the more stable will be, but it will also have less maneuverability. And also, the less slope the more unstable will be, but the more maneuverability will have.
Leaving apart the slope, in order to get the best performance the momentum coefficient should be 0 a the maximum efficiency. In order to achieve this, different parameters have to be modified till the the airplane arrive to the equilibrium.
The combination of tail parameters and tail and wing distribution that achieve the equilib- rium is the next one: Wing span (m) 4 Wing area (m2) 1,5 Stable configuration Wing span (m) 2,4 Croot 0,5 2 Wing (x,z) (0.4 , -0,4) m Wing area (m ) 0,8 Ctip 0,25 Elevator (x,z) (6,1.2) m Croot 0,75 Aspect ratio 10,67 Fin (x,z) (5.9,0) Ctip 0,6 Taper ratio 2
Table 5.11: Stable distances configu- Table 5.12: Final fin configu-Table 5.13: Final elevator con- ration ration data figuration data
Finally the wing does not need dihedral, it is enough with the tail surface and the relative position of the wing and the tail, which is a good thing because the dihedral increase the fabrication cost and also difficult the fabrication and makes the structural analysis more com- plicated.
5.2.7 Final plane configuration 5.2.7.1 Parasite Drag In order to get a more realistic performance of the aircraft, the drag coefficients of the landing gear have been estimated. For the estimation of the wheels CD it has been used the following method: • Define the reference distances.
• Get the CD value from a experimental graphics.
• Define the calculated CD on XFLR5 to improve the veracity of the simulations. And for the tube that connects the wheels it has been used the next following system: • Define the reference distances. • Create an equivalent wing using a symmetric air foil.
• Make a simulation in order to get equivalent CD.
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5.2.7.1.1 Wheels parasite drag The reference distance used for the wheel estimation have been: the following ones:
• Radius = 0.4 m
• Width = 0.1 m
CD are not as usual as it might seem, there are not to much forms analyzed and, a disk through the longitudinal face is one of the figures which have not been analyzed.
The figures that might be an approximation to the disk in volume or in some referent distance are:
• Cylinder
• Sphere
• Bullet
Because of its geometry, the bullet its the most realistic way to simulate the wheel, because it has an smooth curve at the beginning and, but there is not as much rounded as the sphere.
The following figure shows the CD distribution through the Re. The Re used has been the one for cruise flight conditions, which is:
Re = 1.7M (5.9)
Thus, using the following figure, the CD for each wheel have been estimated.
Figure 5.20: CD experimental data ; Source: researchgate.net
l Using the bullet body and a d ratio of 1, to be conservative. The extra drag for each wheel is :
−1 CD = 0.85wheel (5.10)
5.2.7.1.2 Tube parasite drag As it has already been said, the tube has been estimated as a wing, a rectangular one. It can be noticed there is a huge difference between a wing and a tube, but taking into consideration that the tube is nor a cylinder neither a wing, it has been decided to estimates as a rectangular wing using a huge symmetric airfoil.
The geometry used to define the wing tube is:
• lcenter = 2m
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• lside = 0.5m (each diagonal longitude).
Using the listed geometry the extra CD obtained is:
CD0 = 0.05 (5.11)
5.2.7.2 Drag correction Applying the drag correction to the final plane configuration on XFlR5, have been obtained the following results:
Figure 5.21: Comparison between both behaviours
As it can be on the Fig. 5.21 the maximum efficiency decreases due to the extra parasite drag caused by the landing gear. It was expected result, even though the addition of the landing gear drag, the efficiency performance of the airplane is still really good, with a maximum efficiency of approximately 35.
5.2.7.3 Design parameters Once everything has been correctly calculated, the main parameters have been summarized in the following tables:
Wing span (m) 16 Wing area (m2) 12,424 Wing load (kg/m2) 48,295 Croot (m) 1 Ctip (m) 0,6 Wing span (m) 4 Aspect ratio 20,606 Wing area (m2) 1,5 Taper Ratio 10 Wing span (m) 2,4 Croot 0,5 2 Efficiency 35,42 Wing area (m ) 0,8 Ctip 0,25 o Tilt angle ( ) 2 Croot 0,75 Aspect ratio 10,67 v(m/s) 43,6 Ctip 0,6 Taper ratio 2
Table 5.14: Final wing configuration Table 5.15: Final fin configu-Table 5.16: Final elevator con- data ration data figuration data
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Fin Elevator Aileron Sw 0,8 1,5 6,212 Croot 0,375 0,5 1 S(control surface) 0,4 0,5 0,6212 C(control surface) 0,375 0,15 0.25
Table 5.17: Control surfaces main sizes
5.2.7.4 Range study Also in order to estimate the different possible ranges, a Weight-range diagram has been ob- tained using the Breguet’s equation.The code developed is attached in the annexes.
Figure 5.22: Weight-Range diagram
The Fig. 5.22 it has been obtained using a code design by our own and it can be found at the annexes of the report. It has to be taken into account the usual range distance is the first one, which is for 2 passenger plus the pilot. The other tow options are an hypothetical ones,and both could be possible to perform by decreasing the payload (in this case the passengers and adding fuel. The maximum range is calculated by decreasing MTOW in one passenger and keeping the extra 70 kg of fuel.
The development of the equation can be read on ht attachments in the same section Range study. There, it can be found the whole explained development from the main consideration to the expression used to calculate the different points of the Fig. 5.22.
5.2.7.5 Flight envelope In order to define the flight envelope where the LSA is going to flight, an airspeed-gust en- velope has been done at sea level. To develop the diagram, the rules and requirements set by
R 39 G06-AlOn LSA 3 seats | Project report the governmental authorities have been taken into account.
The [1] has been followed in order to make sure to meet all the requirements and to calculate the necessary velocities. Following the structural requirement stablished by the ASTM F225, the load factors needed are:
• n1 = 4 : Positive maneuvering load factor
• n2 = −2. Negative maneuvering load factor
• nF1 = 2 :Positive manoeuvring load factor flaps extended
Firstly, the vs is needed in both cases, with and without flaps. Those velocities have been determined by simulations using XFLR5, which already gives the stall velocity for a Cl. In this case the maximum cl is required. The following velocities will allow computing the other ones in order to be able to create the airspeed diagram:
vs = 45.40kts; (5.12)
vs0 = 43.13kts < 44kts −→ EASA requirement; (5.13) Now the needed velocities can be calculated.
Design manoeuvring speed va: √ vA = vs n1 = 90.8kts (5.14)
Flaps maximum operating speed vF:
Should be bigger than 1.4ss and 2vs0. Doing this operation the result is:
vF = 85kts; (5.15)
Design cruising speed vc:
Needs to be bigger than: r wMTOW vCmin = 4.77 = 103kts (5.16) Sw And smaller than: 1.4VH = 121.21kts (5.17)
So, approximately vC can be defined as:
vC = 110kts (5.18)
And finallly 4.7 Design dive speed VD:
vD > 1.4vCmin = 145.34 ≈ 160kts because needs to be higher (5.19) Using the same methodology for the flap and for the negative load factor but, with the re- spective limitations for each case. The next step is to calculate the gust lines for 7.5 and 15 m/s for both vC and vD.
In order to get the load factor increment due to an instant gust the ASTM F2245 has been used. Using the aircraft data and the graphics that can be found in [1], the increments of the load factor for the different cases can be determined.
Using the increments previously calculated from the ASTM F2245, the following diagram has been developed:
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5.3 Final plane analysis
5.3.1 Efficiency The efficiency of the final plane is shown in the following graph:
Figure 5.24: Efficiency of the final configuration
As it can bee seen in Fig. 5.24 the maximum efficiency is 34,5 at 1,5o of angle of attack.
5.3.2 Static Stability The static stability is the immediate response of the plane in front of a perturbation. A static plane is the one that in front of a sudden increase of angle of attack, a restorer moment ap- pears that decreases the perturbation. This means that the variation of the pitching moment coefficient as function of alpha should have a negative slope. Furthermore, the alpha with zero pitching moment coefficient should be roughly the same that the angle of attack of max- imum efficiency. This condition would make the plane to have the angle of attack with the best glide ratio without the pilot’s help. The following figure shows the final plane pitching moment coefficient, Cm, as function of alpha:
Figure 5.25: Static stability of the final configuration
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The slope of the pitching moment coefficient is negative which will make the plane statically stable. The alpha with no Cm is 0,7o, the alpha of maximum efficiency is 1,5o as it is shown in Fig. 5.24.
5.3.3 Dynamic Stability The dynamic stability consists in the behaviour of the plane in front of a perturbation that changes its flight conditions. The desired behaviour is the one that tries to counter the effect of the perturbation and damp the amplitude of it.
In order to evaluate the dynamic stability of the plane a modal analysis have been done. If the Modes have a damped evolution in time, the plane is dynamically stable. The three more important modes are the followings:
• Longitudinal Mode 1
• Lateral Mode 1
• Lateral Mode 2
5.3.3.1 Longitudinal Mode 1 This mode is related to the pitch angle. The next figure shows the angular velocity related to this mode:
Figure 5.26: Longitudinal Mode 1
As it can bee seen, the angular velocity decreases in time so it’s a stable second order response to the first longitudinal mode. It takes about 2s to dissipate the perturbation.
5.3.3.2 Lateral Mode 1 This mode is related to the roll angle. The next figure shows the angular velocity related to this mode:
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Figure 5.27: Lateral Mode 1
As it shows the figure, the angular velocity decreases rapidly in time so it’s a very stable response. This is a first order behaviour with a time constant of τ = 0.1s. This means that in 0.4s the perturbation is almost gone.
5.3.3.3 Lateral Mode 2 This mode is related to the yaw angle. The next figure shows the angular velocity related to this mode:
Figure 5.28: Lateral Mode 2
The figure shows that is a stable mode but is the slowest one. It takes about 8s to dissipate the perturbation and is more oscillating than the other two studied modes.
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5.4 Structures
5.4.1 Wing 5.4.1.1 Initial approach To design the structure of the aircraft, it has been analyzed the loads which must operate with a maximum load factor of 4 and a minimum of -2. In this case, analysis will be focused on stresses produced by bending moment and beam will be designed considering only this moment. After the beam is designed, qualitative twist analysis will be performed in order to see that beam resists. In order to simplify the load distribution first analysis, there will be some hypothesis:
1. Cantilever beam. Semi-wing study
2. Constant weight distribution, fuel included: 750N per wing
3. Elliptical lift distribution divided in four uniform sections
4. Total lift for n=1 will be = 6.000 N approximated
5. Beam length will be 8 m
Figure 5.29: Loads distribution and approximations
After analyzing lift and weight approximation, see Report Attachments section 2.3.1., dia- gram with load factor of 4, shown in Fig. 5.30, has been chosen to analyze the wing, being the most critical between the others.
Figure 5.30: Moment diagram for n=4
5.4.1.2 Beam analysis As the results obtained in the last section, maximum requirements for the beam are deter- mined and its section can be sized. From n=4, maximum momentum at the embedding is 30 kN·m. Two possibilities are proposed for the beam:
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1. Wagner CentrAl beam (I section)
2. Rectangular carbon fiber section
It has been analyzed the two option calculated in the Report Attachments. The Square section might be heavier but it acts as a torsion box, so it doesn’t need a second beam.
Figure 5.31: Square section parameters
Between both possibilities, they show similar results for similar dimensions but first option is heavier. Considering the weight requirement, it is very important to reduce weight. Thus, the rectangular section has been chosen. The dimensions selected for the beam will be those of the lightest beam within regulations. These are:
H = 3mm; B = 100mm; T = 4mm (5.20) Beam’s height will be adapted to wing’s thickness but B will be reduced to 80 mm, reducing weight. Beam has been studied with SolidWorks FEM analysis in order to accept its performance (Load factor of n=4). It is interesting how FEM analysis show lower maximum stress than the initial one. The reason for this difference might be that 3D model used for computational analysis is different from geometry used for the initial approach. To ensure that beam fits in the wing, the beam has been modelled completely tangent to the skin. Then, thickness of the walls of the beam is measured from the nearest point to the center. This way, walls are slightly thicker than previous geometry and, consequently, inertia is higher and stress is lower.
As shown in Fig. 5.32, stress never exceed the maximum allowed of 400 MPa. The maximum stress is found near the embedding as expected. In figure 5.33, the maximum displacement appears at the wing tip. Given that, these are the results for a load factor of 4, a displacement of almost 900 mm is acceptable.
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(a) (b)
Figure 5.32: Stress analysis
Figure 5.33: Displacement results
5.4.1.2.1 Twist analysis This will be a qualitative analysis that will compare magnitude of the momentum required for the beam to crash. Beam will be considered as thin wall profile. Then, stress is considered uniform through the thickness of the profile and can be calculated as:
M τ(l) = x (5.21) 2t(l)A A is the area inside the medium curve of the profile, t(l) is thickness in each point of this curve and Mx is the twist moment.
Maximum stress will appear when A an t are minimum, then, most critical section will be the tip. It is shown in Fig. 5.34.
Since maximum admissible stress is known from regulations [1], critical moment can be cal- culated : Mx = 2τmaxtmin A = 15, 5kNm (5.22) From Fig. 5.30, it can be seen that twist moment needed is half the maximum bending mo- ment. As explained before, this loads only appear in specific maneuvers that are far from the conventional use of the aircraft. Since twist and bending moments have different magnitude orders in aircraft problems, it can be assumed that this critical moment calculated won’t be applied to the wing.
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Figure 5.34: Tip section and twist stress
5.4.1.3 Ribs analysis Ribs are used to avoid buckling and shape the wing. They will be a sandwich composite with a central 2 mm layer made of aramid honeycomb and external 1 mm carbon fiber layers.
Figure 5.35: Sandwich ribs
The ribs will be separated from one to each other depending on an approximation of the critical buckling length for the most critical section of the wing. As it is known that higher inertia helps to avoid buckling: tip’s section will be the most critical one. Procedure and approximations used to calculate critical length are explained in Report Attachments. The results are shown in Table 5.18.
2 4 c [mm] h [mm] p [mm] S [mm ] I[mm ] Pcrit [N] Lcrit [mm] 600 72 1.344 1.344 1.617.508 537.600 1.442
Table 5.18: Results of buckling analysis
Considering that the critical length is the maximum distance between ribs that avoid buck- ling, a lower distance will be used to ensure buckling doesn’t appear despite the approxima- tions, distance chosen: 1.000 mm. The results are shown at table 5.18. The final result for our designed wing is a square section beam made of carbon fiber with composite ribs. An overall image is shown in figure 5.36
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Figure 5.36: Beam and ribs
5.4.2 Fuselage The following structural element designed is the fuselage. It has been chosen a semi-monocoque structure because it is the most used one and it provides the best structural performance.
Hence, starting from this base, different profiles and frame and stringers configurations have been formulated in order to decide which was the best one [2]. For every configuration the material tested has been the central reinforced aluminium, which is selected in the materials attachment and one can see its properties.
These configurations have been tested with three different profiles: I profile, the standard pro- file with two different thickness values; a U profile with a normalized thickness and finally, a rectangular profile with a normalized thickness [3]. Furthermore it has varied the number of frames (mainly situated along the cabin) to avoid buckling and the number of stringers and the possibility of adding shorter stringers to reinforce the main ones.
Finally, five configurations that one can see in the corresponding attachment (Fuselage chap- ter, design section) have been tested. In order to choose the configuration a static analysis has been performed and the main parameters have been evaluated through an OWA method (which can be seen in the attachment). These parameters are: weight of the structure, secu- rity factor and maximum displacement, taking into account that the weight will be the most limiting item, followed by the security factor, which is less important because its main restric- tion is a minimum that has to be compulsory lower; and finally the maximum displacement, which would be even less important because it is not the actual displacement, there would be a skin that will reduce the displacement.
Therefore, after the evaluation it can be concluded that the best option is the fuselage built with a profile in a U shape with a thickness of 3,5mm Fig. 5.37.
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Figure 5.37: Fuselage structure
As one can see, the final fuselage design presents a series of frames in the main cabin in order to prevent the skin from changing its shape and keep the passengers safe and another frame in the tail. Furthermore, there are only two main stringers due to the narrow shape of the tale but it presents a series of shorter stringers in the main cabin. The main advantages of this solution are the weight, which is the lowest of the different one analyzed (41,49 kg) and the existence of the profile in a real catalog which would allow its acquisition instead of its complete manufacture [3].
5.4.3 Tail 5.4.3.1 Elevator beam The results, calculated with STRIAN[4] are shown in Fig. 5.38, based on the analysis done specified in the attachments.
Figure 5.38: Elevator moment diagram
Then, maximum moment at the embedment is 3750 N·m. Apart from this, after some itera- tion, the lightest beam is chosen and its dimensions are:
H = 1mm; B = 50mm; T = 3mm (5.23) Finally, FEM analysis is shown in Fig. 5.39.
Stress from the analysis is below the stress previously calculated, it might be because of the 3D model used is slightly different from ideal geometry used in initial analysis. Besides, us- ing trigonometrical relations, a displacement of 330mm results in about 10 degrees. As it has been considered in the wing, it can be accepted.
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(a) (b)
Figure 5.39: Elevator FEM analysis
5.4.3.2 Fin beam The results, calculated with STRIAN[4] are shown in Fig. 5.40, based on the analysis done in the attachments.
Figure 5.40: Fin moment diagram
Then, maximum moment at the embedment is 2400 N·m. This time, weight barely changes between options compared to MTOW of 600 kg. Then, decision criteria will be two: weight must be lower than 1 kg and maximum stress on the embedment will be the minimum be- tween different options. Thus, beam chosen and its dimensions are:
H = 2mm; B = 45mm; T = 2mm (5.24) Finally, FEM analysis is shown in Fig. 5.41. Results are within regulations. Same conclusions as in elevator have been taken.
5.4.3.3 Ribs Using the same methodology that is used to calculate distance between ribs in the wing, the different results for fin and elevator are shown in attachments. Notice that lower distances are used to ensure that ribs work. Therefore, length chosen for the vertical distance is 1.000 mm and 500 mm for the horizontal.
5.4.3.4 Final Result The final result are two rectangular carbon fiber beams with the same composite ribs used in the wing. An overall image is shown in Fig. 5.42.
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(a) (b)
Figure 5.41: Fin FEM analysis
Figure 5.42: Beams and ribs
5.4.4 Landing gear As it follows, based on the current regulations, which are listed in the corresponding attach- ment, it has been decided to select a landing gear from one of the aircraft introduced through the state of the art in the project charter. Then, each landing gear has been evaluated and analyzed according to the information given in articles and websites for every airplane and evaluating the current normative.
After it has been considered each airplane’s landing gear, whose characteristics are presented in the corresponding attachment, it has been chosen a fixed tricycle undercarriage with a non- steerable nose wheel as the first variant of the TL-3000 Sirius LSA’s landing gear is.
Once it was decided which landing gear would be better for the aircraft in design, the follow- ing step is to calculate if this system met the mechanical requirements that the aircraft will need. The tests that have been developed are the following ones:
1. Analysis of Static Load
2. Buckling test.
3. Drop test (not with the whole aircraft, only with the landing gear but applying the velocity and the loads corresponding to the whole aircraft).
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Even though many analysis have been performed, the choice of the landing gear has been based on the static analysis, which results are presented on the results table of the analysis of the landing gear attachment. This results have been evaluate through an OWA method and it can clearly be seen that the best option is a landing gear in steel 4130 with a void tube. After this choice notwithstand- ing, a further analysis of the front landing gear was performed and it was conclude that the thickness of the tube could have been reduced from 5 mm to 3 mm in benefit of the weight without any inconvenience to its correct function.
After the selection, a buckling test was also performed with the front landing gear and the three analysis mentioned above were also performed with the rear landing gear, considering only one bar. The results of this analysis are presented in the corresponding attachment and they verify that this landing gear can be used.
Hence, the final design of the landing is the following one:
Figure 5.43: Front leg of the landing gear.
Figure 5.44: Half of the rear landing gear.
5.4.5 Windows The only compulsory window that will be installed in the aircraft is a double windshield that will be directly bought from a certified FAA manufacturer. The chosen manufacturer
R 53 G06-AlOn LSA 3 seats | Project report provides windows with the following characteristics [5]:
1. Cooler Cabin Temperatures
2. Reduces IR (up to 62,8 % )
3. Reduces UV (up to 99 %)
4. Reduces Pilot Fatigue
5. Preserves Aircraft Interior
This windshield will be bought and installed in the aircraft.
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5.5 Power Plant
In order to optimally carry out its expected performance, the aircraft needed a suitable power plant. The following elements have been taken into consideration:
• Engine
• Propeller
• Gear Box
• Engine Mount
• Auxiliary Motor Items
5.5.1 Engine The most critical aspects in the designing of a suitable power plant for the aircraft are the overall weight of the engine and the raw power it provides. A thorough research about cur- rent LSA engines has been done and the 2 most convenient engines fulfilling the requirements have been preselected, the Rotax 912 manufactured by Rotax[6] and the AE50R-AA manu- factured by Austro Engines[7]:
Name Rotax 912 AE50R Type: 4- engine stroke Wankel Power(hp): 80 55 Mass(Kg): 55.4 28 Aprox extra Weight(Kg): 11 6 Coolant: Air Water and Glycol Fuel pump: Yes Yes Gear box: includes not included Alternator: 12V/40A 14V/18A RPM: 5800 7750 PWR (max) (hp/Kg): 1.2 1.7 Torque(Nm): 103 52.5 Fuel: AVGas100 RON 95 AVGas 100LL RON 95 Fuel consumtion: 15.8L/h 12L/h Certification: FAR22 JAR33 ESA Price: —– —–
Table 5.19: Rotax 912 Vs AE50R comparison
In order to assess which engine was the most suitable for the project requirements, an evalu- ation using OWA and PRESS methods was undertaken:
Firstly, the OWA method:
Ordered Weigthed Average Method:
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Weight Rotax 912 AE50R Criteria G P PXG P PXG Power 6 3,2 19,2 2,2 13,2 Mass 10 0,5 5 0,99 9,9 Coolant 4 1 4 5 20 Torque 4 0,49 1,96 0,96 3,84 Accessories included 2 5 10 1 2 SUM(pxg) 26 40,16 48,94 OWA 0,309 0,376
Table 5.20: OWA results
Secondly, the PRESS method, with the relative weight table:
Relative Weight using PRESS method:
Criteria Power Mass Coolant Torque Accessories included Weight 6 10 4 4 2 Relative weight 0,231 0,385 0,154 0,154 0,077
Table 5.21: Relative weight table
And the PRESS results:
Alternative Rotax 912 AE50R PRESS 0,344 2,909
Table 5.22: PRESS results
Therefore, according to the results above both methods agree that the IAE50R-AA engine is the most suitable one and will be selected for the power plant:
Figure 5.45: Wankel motor IAE50R – AA
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5.5.2 Propeller To determine an appropriate propeller for the aircraft, some options have been taken into account. On one hand, it could be purchased from amongst several ones that are already manufactured. On the other hand, a custom propeller could be designed specifically for the requirements of the LSA like plane. A prototype of a custom propeller was designed using a Matlab simulation algorithm based on the Blade Element Method and it was compared with an existing one that, for its charac- teristics regarding our needs it outstanded amongst the others:
Custom propeller Number of blades: 2 Diameter: 144 cm Total mass: 1,1 kg RPM: 2000 rpm Price (estimated): >2500 EUR
Table 5.23: Custom propeller characteristics
HAUTECLAIRE propeller[8] Number of blades: 3 Diameter: 155cm Total mass: 1,7Kg Moment of inertia: 1.400 kg·cm2 RPM: 3000rpm Price: 724 EUR
Table 5.24: HAUTECLAIRE propeller characteristics
Although the custom propeller might be more suitable to the power plant of the project, both propellers have similar characteristics. Therefore, it was finally preferred to purchase the Hauteclaire propeller, as it had already been certified and would always be cheaper than manufacturing a custom one.
5.5.3 Gear Box The gear box will make sure that the RPM of the engine match with the RPM required for the optimum performance of the propeller. The selected gear box will be a Helical Gear from the Aeromomentum Company[9].The gear box will convert the engine maximum RPM from 7750rpm to 2994rpm, according to the propeller requirements.
5.5.4 Engine Mount Once the engine is purchased it will need a motor mount so as it can be installed inside the aircraft.The Austro Engines company does not manufacture motor mounts for their engines, so it is going to be readjusted from an existing motor mount. The mount will be adapted from a Rotax 912 motor as this motor has a similar mount attachment and dimensions[10]:
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Figure 5.46: Standard frame mount
Motor mount frame Weight: approx 4,0kg Price: 981,27 eur
Table 5.25: Motor frame characteristics
5.5.5 Auxiliary Motor Items Finally, in order to ensure that the combustion gases produced by the engine are ejected safely, a complementary exhaust system is required. It will be installed once the engine is mounted inside the plane and is composed of the following elements:
• Micro louver air shield
• Exhaust clamps
• Flexible ceramic heat shield
• Muffler Shroud PA-18
• Stainless Steel 90o bend
• Stub exhaust Stacks
Those elements will be bought from specialized manufacturing companies[11].
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5.6 Avionics and Systems
5.6.1 Basic Equipment In this chapter will be presented the different instruments and systems that an aircraft need to provide information to the pilot about flight, engine, systems conditions; as to allow the pilot to control the aircraft through levers, buttons and pedals.
Figure 5.47: Dashboard in the cockpit
The main considerations that will be taken into account are that the equipment chosen must obey the requirements established in [1] and [12] (and all of them have to be LSA approved or certified). Without forgetting that weight is an important factor for design limitation, and also price has to be taken into account. And other important factor is the compatibility between equipment in order to achieve it, instruments have been selected from the same company or group to ensure that fact.
The equipment can be classified into the following groups:
• Flight and Navigation Instruments
• Powerplant Instruments
• Miscellaneous Equipment
• Safety Belts and Harnesses
5.6.2 Flight and Navigation Instruments Light-sport aircrafts (LSA) must be able to fly following VFR rules. Nevertheless, the aircraft will also have IFR instrumentation to be able to fly following instrumental flight rules if it is necessary, safety is improved by giving more useful information to the pilot.
In order to ensure the pilot comfort and safety all systems are doubled. There is an electronic version and an analog one. This duplication is done in order to provide a second system in case one of them fails. For the electronic version, a 7 inch screen displays all the electronic systems and the analog systems have an individual indicator.
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The Flight and Navigation instruments are: • Airspeed Indicator
• Artificial horizon
• Compass/Directional gyro
• Turn Coordinator
• Clock
• Display
• Digital Air Data, Attitude and Heading
• GPS Reciver/Antenna
• Transponder
• Control panel knobs
• VHF COM Radio
• Two-Place Stereo Intercom
• Pitot Probes All the compulsory elements had been chosen considering the less weight and less power required in order to decrease the batteries weight. In addition, most of the systems are from the same manufacturer so the compatibility between them is ensured.
5.6.3 Powerplant Instruments Powerplant instruments are the instruments in the cockpit of an aircraft that provide the pilot with information about fuel and oil temperatures and pressures, revolutions of the engine, voltage of the system, amount of fuel of the tank and so forth. This instrumentation is com- pulsory and brings essential information about the performance. The elements installed are:
• Fuel quantity indicator
• Tachometer (RPM)
• Engine “kill” switch Engine instruments as required by the engine manufacturer:
• Coolant fluid thermometer
• EGT Gauge
• Oil thermometer
• Fuel pressure indicator
• Fuel quantity indicator
• Oil pressure indicator
• Oil temperature indicator
• Voltimeter
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• Electrical Control Unit
The Powerplant Instruments are chosen with the same criteria as in the Flight and Navigation Instruments. In this section, the fuel tanks were selected. The volume of fuel that the tanks of the aircraft can host is 60 litres.
Due to the innovative engine selected, there are not official instruments in the market avail- able yet. To solve this fact the instrumentation selected has been wisely determined to read the parameters in the exact range in which the engine is working. Therefore, the equipment shows perfectly the range of temperatures and pressures in which the engine develops, con- sulting [13].
5.6.4 Miscellaneous Equipment This section conglomerates some compulsory and optional equipment that cannot be in- cluded in Powerplant instruments neither in Flight and Navigation instruments. Most of the equipment is related to ensure aircraft’s safety but there are also some compulsory elements such as lights or the batteries. The Miscellaneous equipment is:
• Aircraft Lights
• Batteries
• Fire extinguishers
• Aircraft’s parachute
There was the possibility of including night aircraft lights, after a brief investigation it was dismissed because in general LSA do not have permission to fly at night and the inclusion of night system means a weight increase.
The battery selection has been done under an OWA and PRESS method to find the most suit- able one. The evaluated factors were Volume, Weight, Pulse Amperage, Power Pulse and Price. These factors had been assigned an evaluation weight and following the decision cri- teria the best option is a 16 cell battery.
On the other hand, the fire extinguisher is a compulsory safety element that must be included. In order to ensure the pilot and occupants’ safety there are two aircraft specialized fire extin- guishers. The criteria to select them was considering the lightest model. In addition, an aircraft parachute has been selected to prevent the death risk if there is a fatal error in mid air. As the safety of our aircraft and occupants must be fundamental, the decision was to include one and search the most suitable in the market for LSA.
5.6.5 Safety Belts and Harnesses This subsection consider the regulated inclusion of Safety Belts and Harnesses for pilots, pas- sengers and baggage. These elements must be aircraft safety certified and the choice is a 5-point System Harness for the three occupants of the aircraft and three polyester strap to fix the baggage.
For selecting this equipment, the main criteria was safety first and then quality and lightness.
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5.6.6 Controls This section will show how the control system has been sized according to the pilot force es- tablished in [1]. The controls are done with mechanical transmission directly from the control surfaces because is the easiest system and much more lighter than the fly-by-wire system. The advantage is that the control surfaces generate a torque value perfectly affordable for the pilot force.
The parameters selected to design the control stick are chosen in order to ensure the pilot comfort and to bring a perfect sensitivity about any plane movement. To do that a torque springs systems must be installed and adjusted to achieve the best performance.
The stick has a length of 14 inches, the allowed movement is 10 inches in the pilot direction, 5 inches in front direction and 10 inches in left and right direction also, the pedal movement is from 0 to 10 inches.
After the design of the control stick the maximum force of the pilot to control the pitch is 188 N (42,5 lbf), the highest force to control the roll is 87 N (19,6 lbf) and the maximum force applied to the pedals to control the yaw is 392 N (88,1 lbf).
The mechanical transmission is done with cable and pulleys, the required cable is 1693in length and this cover all the control surfaces. The global weight of the cabling system is 0,453 kg (1 lb) and it perfectly support all the control surface forces.
To ensure the comfortably of the stick in the pilot hands, a high quality control stick grip has been selected.
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5.7 Design
The main goal to have a CAD model is to make it easier to draw the blueprints needed for the technical sheets. However, since CAD software offers much more possibilities, the model has been used as a support for other departments. In this case, the software used have been Catia v5 and SolidWorks. Furthermore, renders and a mock-up will be done.
The exterior part includes fuselage, wing, tail and landing gear. Fuselage and surfaces have been modeled based on XFLR5 aerodynamic analysis results. Landing gear was chosen ac- cording to LSA requirements. Additionally, the interior part explains the arrangement of the cockpit, instruments, fire extinguishers, parachute seats and baggage. More information about models of each part are presented in report attachments.
It should be pointed out that it has been selected a 1:2 configuration for the seats, where the pilot in the front seat will control each instrument in the cockpit. A lever is between the pilot’s legs, a throttle lever next to them and pedals under the cockpit . Regarding the baggage, it is placed under the seats of each passenger.
On early stages of the project, a first document with general dimensions of the aircraft was given to Structures department in order to start sizing of the structure.
Figure 5.48: Initial sizing
This overview was prepared with the dimensions used by Aerodynamics department during their analysis and weren’t definitive, so there might be minor changes between Fig. 5.48 and the final aircraft.
Once the model is finished, software automatically calculates total weight and center of grav- ity.
Figure 5.49: Weight and C.G.
As seen in Fig. 5.49, CG for 3 passengers isn’t positioned in the symmetry plane YZ. The
R 63 G06-AlOn LSA 3 seats | Project report reason for this deviation might be a wrong position of one element, such as passengers or landing gear protections. Since it is an error of less than 0.5 mm, it has been considered acceptable and hasn’t been corrected. Another important point is that the origin to which are referred these distances is the origin of the model, which is completely arbitrary. For a better understanding of this point, its reference has been changed as it can be seen in Technical Sheets or in Fig. 5.50.
Figure 5.50: C.G. Position
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5.8 Business
5.8.1 Facilities For the project to take place, first a working area is needed. Such working area will consist of an office (for the engineering process), and a workshop, where the plane parts will be pre-constructed. The offices will be located in Barcelona. The reasons to set the offices in Barcelona, despite higher price (average), is having such a big economical nucleus near, and the proximity Barcelona offers to all the company workers (project members). Four different work spaces have been considered, and the decision on the final one will depend on total space, location and monthly rent. After considering 4 options (see Attachments, Business Chapter, subsection 6.1.1), the final office is the following:
Figure 5.51: Floor plan of the chosen office
Figure 5.52: Map location of the chosen office
The total room per person (including common areas) is 14,42 m2, very close to the optimal work conditions. [14] This office is the place where all the work of the project members will be developed. The cost of this office, per month, is 2.068 e. This amount will be useful later on the Economical Feasibility.
5.8.2 Human Resources The Human Resources consist of the organization and work structure of the project members. The team is composed by 14 members. These 14 members will be organized as shown in figure 5.53:
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Figure 5.53: Scheme of organization of the members of the project
5.8.3 Marketing Plan The marketing plan for a new company is something crucial to achieve the maximum possible costumers and to get to know our product better in the market. In the following sections the most relevant issues of the marketing plan will be explained, for further information all the sections are widely explained in the attachments.
5.8.3.1 Marketing Costs As we have previously said, the importance of a good marketing plan is crucial for the de- velopment of a new company, so a good first investment for the marketing plan has to be considered. In our case, the cost estimated for the marketing plan has been approximately the 12% of the total budget, that results in an initial marketing cost of 8.650 e. This budget intended for the marketing campaign is divided as the following graphic shows:
Figure 5.54: Division of the Marketing’s Cost
The capital invested for the second year of the company’s life will depend on the success of the sales, by increasing the marketing costs if the sales are not as expected and slightly decreasing if there is a lot of success of the sales of the product.
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5.8.4 Study of potential costumers The main advantage of our product in front of the standard Light Sport Aircrafts is the larger capacity, so the main costumers we are targeting to are the actual LSA’s owners that may be attracted to this improved version of their aircraft. In the following graphic it shown the the General Aviation Airplane Shipments of Piston- Engine planes between 1995 and 2017. The great decay from 2007 to 2010 is believed to have been caused by the economical crisis Europe and the US suffered on this period of time.
Figure 5.55: Airplane Shipments Worldwide (1995–2017)
Another option could be reaching out to the new pilots of the EASA. The most basic license one needs to be able to fly a LSA is the Light Aircraft Pilot Licence (LAPL). This license has been getting more and more attention each year since it first appeared in 2011, as it gives the user an unified authorization in EU to fly LSAs, and demands less hours of training than the PPL (Private Pilot License).
5.8.5 Study of advertisement The most important thing about the marketing campaign to take into account is the advertis- ing plan. It is essential to create a communication link between the company as the seller and the public that will be our future costumers. It is important not just to show the product itself, but to show the philosophy behind the product. The main media used for aviation industry is printed media, but in the last generations digital media is growing up really fast so there will also be digital advertisement in our company. The final candidates to work with to develop our advertisement plan are the Air Charter Guide publishing our adverts in their authoritative guide, International Air Transport Association (IATA) with its worldwide seen website, Flying Magazine and we will also be assisted by an specialized company in aviation marketing Perfect Landing.
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Figure 5.56: Companies that collaborate in the Advertising Plan
5.8.6 Design of the Marketing Campaign
Having done the research on the advertisers that will help to promote the company’s vis- ibility, the Design Department has brought the logotype that represents our brand. There were different options designed but the final resulting logotype was simple and clear and was considered the best option:
Figure 5.57: Final Logo for Alpha One
5.8.7 Initial Investment Another important issue carried out by the Business Department is to seek for investors. Since the project we are developing requires a big investment, due to the magnitude of any project in this sector, it is really important to find an associate that believes in the start-up and is committed to lend the capital needed to start the company. The company decided to seek for an Angel Investor, this kind of investor is an associate, usually a business man that invest in projects that seems feasible. In the last years these investors have been the main resource for start-up businesses, and the contribution of them is growing every year.
R 68 Chapter 6
Economic feasibility analysis
This chapter consists on the analysis of the economical development of the project, discussing whether it is feasible or not. For this economical analysis, three main methods are going to be used. Firstly, the Break Even Point (BEP) is going to be calculated, and depending on how much profit the team wants to take from each aircraft (this being the difference between the production cost and the selling price), a Selling Price is going to be chosen. Afterwards, the analysis will continue with the Pay-back analysis, considering the operation costs, and the annual revenues. In order to be able to carry out this analysis, beforehand one must calculate the factor of decay (or rate of growth), k. With this final analysis, one will obtain the Pay Back Time (PBT), the Return of Investment (ROI) and the Internal Rate of Return (IRR).
6.1 Break Even Point
The Break Even Point is the point at which the company’s operation costs for building an specific number of aircrafts is matched with the total revenues of the sold aircrafts. Typically, a product has two different types of costs associated with its production, theses are the Variable Costs and the Fixed Costs. The variable costs consist mainly of materials’ costs and production costs. These Variable Costs depend obviously on the number of aircrafts the company has built, so it will increase with each unit produced (see Table 6.1. The Fixed Costs consist mainly on the personnel’s salary, the rent of the facilities, and the marketing campaign costs. These costs are described with detail on the Budget Attachment. These fixed costs do not depend on the number of aircrafts manufactured, and are maintained if the units vary (see Table 6.2. With this two different costs added, one obtains the total cost of production of a set number of aircrafts. Variable Costs Material 55.695 e Production 19.500 e1 Total 75.195 e
Table 6.1: Variable Costs Specifications
Once the Variable Costs and the Fixed Costs, one can calculate the Break-Even Point, as fol- lows. • Selling Price: The selling price must be one that is high enough for us to regain the investment but low enough to be competitive among the other existing LSA’s. The price of the existing LSA’s lies in a range from ≈ 100.000 eto 200.000 e. The chosen selling price is 160.000 e, considering that Alpha One has perks before the other LSA’s, as it gives room for three people instead of two.
1Specified in the Budget Document
R 69 G06-AlOn LSA 3 seats | Project report
Fixed Costs Personnel 210.620,80 e Rent 8.272,00 e Additional 6.720 e Insurance 3.500,00 e Marketing 8.650,00 e Total 237.762,80 e
Table 6.2: Fixed costs specifications
• Fixed Capital of Investment: This capital is the total amount of money the investors have to spend every year in the project for it to keep on going. In this project, it has been estimated as 100.000 eevery year. This cost includes the software licenses, taxes, maintenance cost, etc.
Now one can proceed to graph the Costs vs. Units Sold, that gives the graphic seen on figure 6.1:
Figure 6.1: Break-Even Point
The Break-Even Point as seen in the figure 6.1 is around 4 units. Therefore, once the company sells 5 units, the total costs equal those of the revenues. Considering the market manufactures approximately 1100 new aircrafts every year, one can estimate that the volume our company occupies in the total market is around 1% of the sells, so the amount of planes the company would be likely to sell is 10 every year (see Business Attachment for more information). Al- though this is a good objective for the company, this amount of planes is not achieved in the first year, as the company takes some time to get used to the market, so the estimation of planes sold will start with 7, and increase one unit per year until reaching 10, in 2022. This data will later be useful in the Pay-Back analysis, as it marks the annual Revenues.
6.2 Pay-Back Analysis
The Pay Back Analysis consists on the study of how the Company’s wallet will vary on the forthcoming years.
6.2.1 Estimation of decay rate k The factor of decay k is a crucial fact for calculating the Pay Back Time, the IRR and the ROI. The factor k depends on mainly three aspects, those being:
R 70 G06-AlOn LSA 3 seats | Project report
• Inflation: Inflation is determined by the IPC of the country the company is settled in, in this case Spain. The IPC of spain for 2018 is 1,4%. • Investors’ influence: The money our investors lend to the project is expected to be re- turned in some period in time. This affects the rate at which the value of the money changes. In this case, it has been estimated as 5%. • Risk factor: the risk factor is the probability that the project is not successful. The more new and innovative the project is, the higher the risk factor. Being this project a new concept of aircraft, and the company a brand new one, the risk factor of this project must be high. It has been estimated as 5%.
Once calculated all the dependencies of k, it gives a total value of 11,4%.
6.2.2 Pay-Back Time, Return of Investment (ROI) and Internal Rate of Return (IRR) The Pay-Back Time is the time we need to recover the initial investment. In order to obtain an estimation of it one has to calculate the cash flow of every year since the first investment updating it with the decay rate k calculated in the previous section. The Pay-Back Time is calculated with the following formula: I Pay − Back Time = 0 (6.1) CF¯
With : I0 : InitialInvestment CF¯ : AverageCashFlow
The Average Cash Flow is calculated subtracting the Operation Costs from the annual Rev- enues of each year, and then computing the average throughout the years of study. The Initial Investment is the amount of capital needed to start up the project for the first time. This cap- ital, as discussed in the Business Attachment comes from an angel investor, so this money is expected to be returned in a short period of time. The Initial Investment for this project has been estimated as: 1 M e. The Cash Flow of each year, along with the average, is shown on the following table (considering the units sold per year commented previously):
2018 2019 2020 2021 Revenues - 1.120.000,00 e 1.280.000,00 e 1.440.000,00 e Operation costs - 866.127,80 e 941.322,80 e 1.016.517,80 e CF -1.000.000,00 e 253.872,20 e 338.677,20 e 423.482,20 e UCF -1.000.000,00 e 227.892,46 e 272.907,57 e 306.322,99 e
2022 2023 2024 2025 Average 1.600.000,00 e 1.600.000,00 e 1.600.000,00 e 1.600.000,00 e 1.462.857,14 e 1.091.712,80 e 1.091.712,80 e 1.091.712,80 e 1.091.712,80 e 1.027.259,94 e 508.287,20 e 508.287,20 e 508.287,20 e 508.287,20 e 435.597,20 e 330.041,39 e 296.266,96 e 265.948,80 e 238.733,21 e 276.873,34 e
Table 6.3: Cash Flow and Updated Cash Flow for the first 7 years of the project
For the first even years, as seen in the table above, the Average Updated Cash Flow is 511.668,11 e. This gives a Pay-Back Time of: 1.000.000, 00e Pay − BackTime = = 3, 611 = 3 years and 7 months. 276.873, 34e/year
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Usually for a project to be profitable it is recommended that the PBT is shorter than 6-7 years. With this project’s PBT, we can assume the profitability is high enough to make money suffi- ciently fast for the angel investor, and for the team members to make benefits of it.
Two other parameters calculated are the Net Present Value (NPV) and the Return of Invest- ment (ROI). The NPV is the cumulative Cash Flow over the years, meaning it is the sum of the Cash Flow of a year and all the previous Cash Flows. The Return of Investment is the ratio between the NPV and the initial investment. The higher these values are, the more profitable the investment. For every year, we have computed the associated NPV and ROI, shown in the table 6.4, and graphed on figure 6.2:
2018 2019 2020 2021 NPV -1.000.000,00 e -772.107,54 e -499.199,97 e -192.876,99 e ROI -1 -0,772 -0,499 -0,192
2022 2023 2024 2025 137.164,41 e 433.431,36 e 699.380,16 e 938.113,37 e 0,137 0,433 0,699 0,938
Table 6.4: Return of Investment and Net Present Value for the first 7 years of the project
Figure 6.2: Net Present Value for the first seven years of the project
So, as we can see above, the results of the ROI increase every year and end up almost equal- ing the initial investment on the 7th year (2025).
Finally, we have calculated the discount rate needed to make Net Profit Value zero, this equals the definition of the Internal Rate of Return (IRR) (see formula 6.2). The term of Internal refers to the fact that this parameter excludes external factors, such as inflation, the cost of the capital or other financial risks.
7 C f NPV = t − C (6.2) ∑ ( + )t 0 t=1 1 r As long as a project is feasible and the Net Present Value is positive, the IRR takes on a value higher than the decay rate. In this project, the IRR gives a value of k = 33% which is an acceptable value, as it is higher than the value of k calculated with external factors, 11 %. In this project in particular, this value also has to be around 30 % - 40 % , as discussed in the Business Attachments, since it comes from an angel investor.
R 72 G06-AlOn LSA 3 seats | Project report
In conclusion, the Economical Feasibility analysis of Alpha One project has given a positive result, with a short Pay-Back Time of about 3 years and 7 months, and an Internal Return of Investment high enough to make it profitable for future investments.
R 73 Chapter 7
Environmental impact analysis
The environmental impact analysis consists of all the issues related to the effect our aircraft will have on the environment. These effects must be controlled, and must lay on a range of emissions set by ICAO’s regulations. The regulation Alpha One has to fulfill is ICAO Annex 16, Volume II: Environmental Protection – Aircraft Engine Emissions, or its relative in Europe, EASA’s CS-22. [15] [16]
The only part regarding this limitation in our aircraft is the engine, since the plane has not any turbojets or turbofans to produce thrust. The engine used is the Austro Engine AE50R, as discussed on the Power Plant chapter. This engine is certified according to EASA Part 22 Sub-part H Certification Specifications for Sailplanes and Powered Sailplanes, H referring to the engine sub-part. [17]
Another issue that must be considered when speaking about Environmental Impact, apart from the impact the aircraft will have itself, is the impact of the project development, this being all the resources that will be focused on the project.
Supposing the project will be working once it is finished, and taking into account the com- pany will have to sell several planes per year, one can not wrongfully predict that the pro- duction chain will be carried out non-stop five years per week. This means that the factories manufacturing the plane, as well as the office, but with greater impact, will be on 8 hours per day, 5 days per week. The fact that gives information about a process’ environmental impact is the Carbon Footprint. This value is calculated with different methods, considering all the processes that every part of the plane must go through to give as result the plane itself. One of the most important issues with the Carbon Footprint is the re-usability of the materials used in the plane. By buying these materials to European or USA suppliers, the company (AlphaOne) assures that all the materials used in the manufacturing process can be re-used and comply the European and American laws of pollution and environment. Another significant factor when it comes to the Carbon footprint is the total emissions of carbon our office produces within a year. This value is calculated using estimations from Spain, based on the previous years. [18] [19]. Using some factors relating the Kg of CO2 produced per KWh for electricity, gas, paper and water, these different values have been estimated for a company based in Spain, like ours [20].
Annual Equivalent CO2 Emission (Estimation) Concept Consumption Physic Units Emission Factor kg CO2 / unit Equivalent kg CO2 Electricity 17014 kWh 0,385 6.550 Natural Gas 34968 kWh 0,2016 7.049 Regular Paper 12 kg 3 36 Water 30 m3 0,788 24 Total kg of CO2 13.659
R 74 Chapter 8
Safety considerations
Safety and comfort is a primordial aspect for Alpha One LSA aircraft.
A parachute for the entire aircraft has been included. Ballistic Recovery System (BRS) parachute could lower an entire light aircraft to the ground in the event of loss of control, failure of the aircraft structure, or other in-flight emergencies.
Figure 8.1: Sequence of BRS parachute
The parachute provides more safety to possible mid air incidents. BRS manufacturer has saved hundreds of lives.
In case of fire in the air or on the ground, two Halon fire extinguishers had been installed in the cockpit. The Halon fire extinguishers are light, portable and can extinguish a possible fire in electrical equipment, in cloth, paper, rubber, many plastics, or even in flammable liquids and oils.
Also by structural security, the EASA security rules have been followed in order to design the whole aircraft structure. These structural considerations are the following ones:
• maximum positive load factor of n = 4
• maximum negative load factor of n = −2
• maximum flap load factor of n = 2
The maximum load factors have also been taken into account when defining the flap en- velope. The aerodynamic department have taken into account the maximum top velocities
R 75 G06-AlOn LSA 3 seats | Project report permitted by EASA to design the wing, define the high lift surfaces such as flaps.
There are also some a posteriori safety considerations that should be taken into account when piloting the aircraft:
• Make use of the engine manufacturer manual in order to keep the engine in good con- ditions
• Keep the distance with the propeller once this one it is running.
• Keep a reasonable reserve fuel level.
• Do not exceed the flight conditions.
• Keep always in mind the flight envelope and its maximum velocities.
• Do not exceed the maximum engine speed.
• Follow the maintenance considerations established in the manuals for the engine, parachute system, extinguishers...
R 76 Chapter 9
Planning and scheduling for the follow-up of the project
The planning process had different phases which helped to schedule the tasks and to orga- nize the members of the team.
The first point was to organize the members of the team in different departments. In order to do this, it was compulsory to identify the different aspects of the plane that needed to be sorted out. Those aspects were defined as: 1. Aerodynamics 2. Structure 3. Power Plant 4. Avionics 5. Economics Therefore, the team was divided in these departments in order to organize the work. Aside, a group coordinator, department coordinators and a secretary were chosen in order to keep track and maintain communication between different teams. This organization was kept for all the project duration. This organization was explained in a more detailed manner in the project charter.
Then, each team proposed a list of tasks based on the requirements of each department. These tasks were later the base to build the Gantt diagram that can be seen in the corresponding attachment (Organization, planning and scheduling chapter). With the Gantt diagram, the tasks each member of the group had to complete were scheduled for the completion of the project contents. These contents were distributed along the departments chapters in the re- port assignments.
The project evolution was analyzed and controlled using the schedule and during the weekly meetings it was checked that all the tasks were in time and if there was any delay it was tried to solve it as fast as possible. Weekly meetings were useful also to put in common the in- formation about the project that was generated and helped to organized tasks that involved members of different team groups.
Meanwhile, as the chapters were finished, the same chapters were also redacted in the report, only including the process of choice or design of the principal parts of the aircraft and the most relevant information in order to help the reader to know how the project was devel- oped. Must be said that all the global tasks were submitted in time or even, in advanced.
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The completion of this part took most of the weeks of the project as it was the development of the project itself and provided the content of the rest of the documents almost in its totality.
When the project was enough completed to the team members to assume more tasks, a new planning for the remaining time was done. In this planning were included the rest of the project documents that had not be scheduled yet. These documents were: the drawings, the technical sheets, the budget and the video. As it can seem logic, they were assigned to the dif- ferent departments in the following order: the drawings were prepared by the department of design, the budget was assigned to the economics department and the technical sheets were assigned to the remaining departments.
The documents presentation was due to December 16Th in order to have time to revise ev- erything. At the same time, the members of the group that were assigned to do the presentation pre- pared the media files to support the presentation and also the contents of the same. Re- hearsals for the presentation were scheduled for December 18Th and 20Th.
R 78 Chapter 10
Conclusions and recommendations
It is never easy to pass from the theory to a real practice exercise. The most difficult part of the project has been to develop a three passenger aircraft with a Light Sport Aircraft specifica- tions which restricts the maximum take off weight to 600 Kg. When considering 3 passengers means that more than the half of the weight will be from the passengers and the seats, which only leaves less than 300 kg for structures, engine, fuel, electronics, etc.
Thus, the main objective was to design a high efficient aircraft more similar to a glider in or- der to be able to flight without a fuel an reserve the maximum weight to structures. Because there was no range limitations, to achieve the 600 Kg of MTOW the resultant trip fuel is only 20 kg with a 4 kg of reserve fuel. Also the possible lither materials have been used such as carbon fiber and aluminum alloys.
Once known the limitations it has to be said that all the objectives have been satisfactory achieved. The following point describes the main developed task by each department in or- der to be able to accomplish the project aim.
The aerodynamics departments have developed an LSA/glider, efficient enough to fly with- out an engine, giving the pilot the possibility to flight with or without propulsion.
Structures has developed an air frame lighter enough to keep the MTOW low but also keep the EASA structural requirements by using composite materials both for the airplane skin and the main structure ( beams and ribs).
Power plant has choose an engine two times lighter than the usual ones but with enough power to propel the airplane and achieve the take-off in only 205 m.
Avionic’s has design the control board and the pulleys system that allows the pilot to control the aircraft.
Design has take to reality the aerodynamics and structures design in order to define the blueprint and the renders of the final plane.
And finally economics has defined an economical in order to be able to sell the airplane be- tween the LSA usual values to be competitive. Moreover a 3D printed scale model has been done to show the results of the work done.
It has to be said that there have been some difficulties while developing the project. Mainly at the structures and aerodynamics departments. The LSA requirements has create some hard point to achieve, such as a good cruise performance that did not complicate too much the structures work, or a skin thickness that gives structural solidity but that do not increase the
R 79 G06-AlOn LSA 3 seats | Project report weight too much.
Leaving the technical points apart, it has to be said that some of the theory materials needed were taught after they were needed.This made that some of the taken decision has to be re-taken, but now applying the explained methods, because it was mandatory to use them. The same situation has happened not only when taking decision but when working on some point such as in aerodynamics, there was needed to develop the control surfaces but we have not done the theory at class yet. Finally the team was able to overcome the difficulties and achieve the goal.
A part of these point the whole group has been working really good. Taking into account that we are 14 people which is a quite big group, and not all of us have meet before, the work has been great and the environment too. And spuriously there has been no problem neither between the members of each department nor between them. The projects requirements has been overcome the best way possible and the achieved results are really good.
R 80 Bibliography
[1] ASTM International. Standard Specification for Design and Performance of a Light Sport Air- plane. Tech. rep. ASTM International, 2006. URL: https://www.astm.org/Standards/ F2245.htm. [2] Michael Chun-Yung Niu. Airframe : stress analysis and sizing / Michael Chun-Yung Niu. Dragon Terrance, North Point : Hong Kong Conmilt Press, 1999. ISBN: 9627128082. [3] CÁTALOGO DE PERFILES DE ALUMINIO NORMALIZADOS ALUMINIUM STAN- DARD PROFILES CATALOGUE. Tech. rep. URL: http://www.extrusax.com/imagenes/ descargas/es/12/STANDARDPROFILES-PERFILESNORMALIZADOS.pdf. [4] STRIAN—Structural analysis. URL: http : / / structural - analyser . com/ (visited on 12/15/2018). [5] CoolView Aircraft Windows, Heat Blocking Cabin Windows. URL: https://leeaerospace. com/coolview-aircraft-windows/ (visited on 11/29/2018). [6] Rotax. ENGINE TYPE 912 | 80 hp. URL: https://www.flyrotax.com/files/Bilder/ Produkte20Rotax / Datasheets / Produktdatenblatt _ 912 _ 80hp _ rev . BRP - Rotax _ 20160823.pdf. [7] Austro Engine. AE59R-AA Engine Manual. 2011. URL: http : / / austroengine . at / uploads/pdf/EME10105r6IAE50RAA.pdf. [8] E-PROPS. HAUTECLAIRE carbon propellers for ultralights. URL: http://www.e-props. fr/16/hauteclaireA.php (visited on 12/19/2018). [9] Aeromomentum. Aircraft engine, Gearbox. URL: http://aeromomentum.com/gearbox. html (visited on 12/19/2018). [10] Aircraft Spruce. Rotax engine mount. URL: https://generalaviationnews.com/2009/ 07/13/rotax-engine-mount-at-aircraft-spruce/ (visited on 12/19/2018). [11] Aircraft Spruce. Exhaust System & Components. URL: https://www.aircraftspruce. com/menus/ep/exhaustcomponents.html (visited on 12/19/2018). [12] Cornell Law School. 14 CFR Part 91, Subpart C - Equipment, Instrument, and Certificate Requirements- US Law. 1989. URL: https://www.law.cornell.edu/cfr/text/14/part- 91/subpart-C (visited on 11/13/2018). [13] Engine Manual. “IAE50R – AA”. In: (2014). [14] La oficina ideal: 14m 2 por empleado | Pyme | Cinco Días. URL: https : / / cincodias . elpais.com/cincodias/2014/10/28/pyme/1414500383{\_}553511.html (visited on 12/11/2018). [15] Easa. Annex to EDD 2017/024/R. Tech. rep. URL: https://www.easa.europa.eu/sites/ default/files/dfu/AnnextoEDDecision{\%}28AMC-GMPart-21-Issue2Amdt7{\%}29. pdf. [16] Icao. Doc 9920, Assembly Resolutions. Tech. rep. URL: https://www.icao.int/environmental- protection/Documents/A36{\_}Res22.pdf. [17] Products - Austro Engine. URL: http : / / austroengine . at / en / products (visited on 12/11/2018).
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[18] Energía en edificios de Oficinas. - Enectiva. URL: https://www.enectiva.cz/es/blog/ 2015/06/ideas-energia-edificio-de-oficinas/ (visited on 12/17/2018). [19] Consumo de electricidad medio de una vivienda en España. URL: https://tarifasgasluz. com/faq/consumo-electricidad-hogares/espana (visited on 12/17/2018). [20] Cámara Zaragoza. URL: https://www.camarazaragoza.com/ (visited on 12/17/2018). [21] Federal Aviation Administration. “Advisory Circular”. In: (2006), p. 98. [22] Powered Sailplanes. “Standard Specification for Design and Performance of a Light Sport Airplane 1”. In: (). [23] Atec. 321 Faeta | Light Aircraft DB & Sales. [24] Libice Nad Cidlinou. “ATEC 322 FAETA Flight and Operations Manual”. In: March (2013), pp. 1–54. URL: http://www.atecaircraft.be/dossiers/Manuals/flight- manual-atec-322-faeta.pdf. [25] Ultralight. “TL-2000 Sting S4 PILOTS OPERATING HANDBOOK”. [26] Aircraft Maintenance Manual. “TL-3000 Sirius”. In: (). [28] Alexander Schleichers. “ASG 29 E. Sneak into the Design Process”. [29] Pu Woei Chen, Quen Yaw Sheen, Harn Wen Tan, and Tai Sing Sun. “Fatigue Analysis of Light Aircraft Landing Gear”. In: Advanced Materials Research 550-553 (2012), pp. 3092– 3098. ISSN: 1662-8985. DOI: 10.4028/www.scientific.net/AMR.550-553.3092. URL: https://www.scientific.net/AMR.550-553.3092. [30] Chunyun. Niu. Airframe structural design : practical design information and data on aircraft structures. Hong Kong : Conmilit Press Ltd, 1999, p. 612. ISBN: 9627128090. [31] Adnene Tlili and Sofiene Bouhjar. “Performance Study of a Metal Matrix Composite Alloy for Aircraft Industry Use”. In: June (2015). DOI: 10.13140/RG.2.1.2292.3360. [32] Seats & Seat Covers - Air-Tech Inc. URL: https://air-techinc.com/topic{\_}std{\_ }prods.php?catid=196{\&}pmid=18 (visited on 11/22/2018). [33] Sport Aircraft Seat Company. URL: http://www.sportaircraftseats.com/sportaircraftseats/ Home.html (visited on 11/22/2018).
R 82 POLYTECHNIC UNIVERSITYOF CATALONIA
AEROSPACE ENGINEERING
Engineering Projects
3 seat Light-Sport Aircraft development | AlOn Contents: Technical sheets
Group: G06-AE-2018/19-Q1
Delivery date: 20/12/2018
Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. G06-AlOn LSA 3 seats | Technical sheets
T 2 Contents
1 Specification sheet scope 9
2 General conditions 11 2.1 Weights & Performances ...... 11
3 Parts lists 15 3.1 Fuselage ...... 15 3.2 Wing ...... 15 3.3 Tail ...... 16 3.4 Control surfaces ...... 16 3.5 Propeller ...... 17 3.6 Engine ...... 18 3.7 Landing gear ...... 19
T 3 G06-AlOn LSA 3 seats | Technical sheets
T 4 List of Figures
2.1 W-R diagram ...... 12 2.2 Gust-Airspeed envelope diagram ...... 13
3.1 Root and Chord distribution ...... 17 3.2 Position of the landing gear in the fuselage ...... 19 3.3 Distance between main and front landing gear ...... 19
T 5 G06-AlOn LSA 3 seats | Technical sheets
T 6 List of Tables
2.1 Main weights configuration ...... 11 2.2 AlOn performance data ...... 11
3.1 Fuselage main parameters ...... 15 3.2 Wing main parameters ...... 16 3.3 Fin parameters ...... 16 3.4 Elevator parameters ...... 16 3.5 Control surfaces main sizes ...... 16 3.6 Custom blade design specs ...... 17 3.7 HAUTECLAIRE 3 specs ...... 18 3.8 Custom propeller cruising specs ...... 18 3.9 AE50R technical data ...... 18 3.10 Gear box technical data ...... 18 3.11 Landing gear wheels’ parameters ...... 19
T 7 G06-AlOn LSA 3 seats | Technical sheets
T 8 Chapter 1
Specification sheet scope
The present specification sheet will be focused in the technical specifications, the dimensions, materials, weights, systems and main flight characteristics of the developed project, a LSA 3 seats with a glider configuration.
The technical specification sheet is derived from a preliminary structure, aerodynamic, sys- tem and propulsion design, it only include the designed and choose parts specifically for the the project, it does not include screws or fixing elements.
There is also a list of the main regulation that this projects follows on the general condition section of this document.
Because of the preliminary phase of the design basic technical condition are being shown in this document. Conditions such as machinery, or assemblies will not be taken into account.
T 9 G06-AlOn LSA 3 seats | Technical sheets
T 10 Chapter 2
General conditions
2.1 Weights & Performances
The next table shows the main weight configuration.
Weight (kg) MTOW(Maximum Take Off Weight) 600 MPL(Maximum Pay Load) 261 MFW(Maximum Fuel Weight) 24 MZFW(Maximum Zero fuel Weight) 576 ZFW(Zero Fuel Weight) 340 FW(Fuel weight) 50 TF(Trip fuel) 2 RF(Reserve fuel) 4
Table 2.1: Main weights configuration
Performances Stall with flaps 81.49 (km/h) Stall without flaps 83.33 (km/h) Manouvering speed (km/h) 166.67 max. speed with flaps extended 157.41 (km/h) VNE(velocity never exceed) 243.3 (km/h) Best glide 1:40 Best glide ratio speed (km/h) 167.76 Best glide at 150 km/h 1:37.6 Best glide at 180 km/h 1:38.36 Take off distance (m) 205 Max. positive load factor (n1) 4 Max. negative load factor (n2) -2 Fuel consum. at full power (l/h) 12
Table 2.2: AlOn performance data
Also a weight-range diagram has been done in order to see which are the different possible configuration knowing that the aircraft is able to fly with just the pilot, plus one or plus two passengers. The following figure shows the W-R diagram taking into account the data from
T 11 G06-AlOn LSA 3 seats | Technical sheets
Table 2.1:
Figure 2.1: W-R diagram
In order to determine the range of velocities and load factors where the aircraft is able to fly a gust-airspeed envelope has been done. The following image shows the envelope and all the important velocities:
T 12 G06-AlOn LSA 3 seats | Technical sheets Gust-Airspeed envelope diagram Figure 2.2:
T 13 G06-AlOn LSA 3 seats | Technical sheets
T 14 Chapter 3
Parts lists
3.1 Fuselage
The fuselage is designed to minimize the weight and to resist the loads that will appear dur- ing its performance.
The following table shows the main parameters of the fuselage.
Fuselage Parameter Unit Value Main parameters Fuselage length m 7,917 Cabin maximum height m 1,4 Cabin maximum width m 1,3 Cabin door length cm 92,42 Cabin door width cm 114,12 Beam details Profile dimensions mm 100x15x15 Profile thickness mm 3,5 Stringer details Lower main beam length m 7,317 Upper main beam m 5,910 Cabin stringers m 2,738 Frames position in the main lower beam 1 m 0 2 m 0,913 3 m 1,284 4 m 1,900 5 m 2,911 6 m 4,432 7 m 7,217
Table 3.1: Fuselage main parameters
3.2 Wing
The following table shows the main wing parameters, needed in order to define the wing completely.
T 15 G06-AlOn LSA 3 seats | Technical sheets
Wing Parameter Unit Value Span (b) m 16 2 Wing surface (Sw) m 12,424 Wing load kg/m2 48,295 Root chord croot m 1 Tip chord ctip m 0,6 Aspect ratio (A) 20,606 o Dhiedral (Γw) 0 Tilt angle o 2 Taper Ratio (λ) 10 Efficiency (E) 35.42 Xac mm 225 CL−cruise 0.45 CL−max 1.42 o αcruise 0 o αmax 10.5 CD0 0.008 Geometric torsion o 3 Torsion position m 3,5
Table 3.2: Wing main parameters
3.3 Tail
The following table shows the main tail,both horizontal and vertical parameters, needed in order to define the wing completely.
Wing span (m) 4 Wing area (m2) 1,5 Wing span (m) 2,4 Croot 0,5 2 Wing area (m ) 0,8 Ctip 0,25 Croot 0,75 Aspect ratio 10,67 Ctip 0,6 Taper ratio 2
Table 3.3: Fin parameters Table 3.4: Elevator parameters
3.4 Control surfaces
In this section are defined the control surfaces which allows the plane to maneuver. The ailerons,the fin and the elevator are defined in the next table:
Fin Elevator Aileron Sw 0,8 1,5 6,212 Croot 0,375 0,5 1 S(control surface) 0,4 0,5 0,6212 C(control surface) 0,375 0,15 0.25
Table 3.5: Control surfaces main sizes
T 16 G06-AlOn LSA 3 seats | Technical sheets
3.5 Propeller
The following table (table 3.6) shows the technical specifications of the custom propeller de- sign for the Alpha one:
Custom propeller Number of blades: 2 Diameter: 144 cm Radius of central cone: 10 cm Airfoil used: NACA 4412 Root torsion theta0 = 0.3 rad Root chord chord0 = 0.12 rad Total mass: 1.1 kg RPM: 2000 rpm
Table 3.6: Custom blade design specs
The plot 3.1 illustrates the chord and torsion distribution of the custom blade design:
Figure 3.1: Root and Chord distribution
The table 3.7 illustratesillustrates the technical specifications of the HAUTECLAIRE 3 blade propeller, the alternative already made propeller election:
T 17 G06-AlOn LSA 3 seats | Technical sheets
HAUTECLAIRE propeller Number of blades: 3 Diameter: 155cm Total mass: 1.7Kg Moment of inertia: 1.400 kg·cm2 RPM: 3000rpm
Table 3.7: HAUTECLAIRE 3 specs
The table 3.8 presents the overall performance of the custom propeller design with the wankel motor:
Custom propeller results Power consumed: 14.01 kW Power delivered: 8.33 kW Efficiency: 59% RPM: 2021 rpm Thrust: 147 N
Table 3.8: Custom propeller cruising specs
3.6 Engine
The table 3.9 presents the characteristics of the Wankel AE50R:
Name AE50R Type: Wankel Power(hp): 55 Mass(Kg): 28 Aprox extra mass(Kg): 6 Coolant: Water and Glycol Fuel pump: Yes Gear box: not included Alternator: 14V/18A RPM: 7750 PWR (max) (hp/Kg): 1.7 Torque(Nm): 52.5 Fuel: AVGas 100LL, RON 95 Fuel consumtion: 12L/h Certification: ESA
Table 3.9: AE50R technical data
The following table (tamble 3.10) presents the gear box characteristics description:
Aeromomentum Gear Box Gear ratio: 2.588:1 Weight: approx 3kg
Table 3.10: Gear box technical data
T 18 G06-AlOn LSA 3 seats | Technical sheets
3.7 Landing gear
The landing gear position can be seen in the following drawings, Fig. 3.2 and Fig. 3.3.
Figure 3.2: Position of the landing gear in the fuselage
Figure 3.3: Distance between main and front landing gear
And the parameters of the wheels are the following ones:
Diameter 40 cm Width 10 cm
Table 3.11: Landing gear wheels’ parameters
T 19 POLYTECHNIC UNIVERSITYOF CATALONIA
AEROSPACE ENGINEERING
Engineering Projects
3 seat Light-Sport Aircraft development | AlOn Contents: Budget
Group: G06-AE-2018/19-Q1
Delivery date: 20/12/2018
Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. G06-AlOn LSA 3 seats | Budget
B 2 Contents
1 Units 7
2 Costs 9
3 Budget 11
B 3 G06-AlOn LSA 3 seats | Budget
B 4 List of Tables
1.1 Units of each element of the Aerodynamics department ...... 7 1.2 Units of each element of the Power Plant department ...... 7 1.3 Units of each element of the Avionics department ...... 8 1.4 Units of each element for the Design Department ...... 8
2.1 Unit Cost of each element for the Structures Department ...... 9 2.2 Unit Cost for each element of the Power Plant department ...... 9 2.3 Unit Cost for each element of the Avionics department ...... 10 2.4 Unit Cost of each element for the Design Department ...... 10
3.1 Budget for the Structures Department ...... 11 3.2 Budget for the Power Plant department ...... 11 3.3 Budget for the Avionics and Systems department ...... 12 3.4 Budget for the Design Department ...... 12 3.5 Additional Fixed Costs ...... 13
B 5 G06-AlOn LSA 3 seats | Budget
B 6 Chapter 1
Units
This section contains the number of units needed of each element. All the units are shown in the tables below.
2. Structures ID Content Quantity 2.1 Carbon Fiber (4mm) 2 1 2.2 Pneumatics 3 2.3 Carbon Fiber (3mm) 3 1 2.4 Carbon Fiber (1mm) 4 1 2.5 Honeycomb 1 2 2.6 Carbon Fiber (2mm) 1 1 2.7 Windshield 1 2.8 Epoxy Resin (Skin) 4 2.9 Aluminium (U shape) 8 2.10 Landing Gear 1 2.11 Carbon Fiber Cloth 150 m2
Table 1.1: Units of each element of the Aerodynamics department
3. Power Plant ID Content Quantity 3.1 Motor 1 3.2 Motor Mount 1 3.3 Gear Box 1 3.4 Micro louver air shield 1 3.5 Exhaust Clamp 1 3.6 Ceramic heatshield 1 3.7 Muffer Shroud 1 3.8 Stainless Steel 90o Bend 1 3.9 Homebuilders Exhaust Stacks 1 3.10 Custom Propoeller (Optional) 1 3.11 HauteClaire 3 blade 1
Table 1.2: Units of each element of the Power Plant department
1Units referring to sheets of 0,8 m2 each 2Units referring to sheets of 3 m2 each
B 7 G06-AlOn LSA 3 seats | Budget
4. Avionics ID Content Quantity 4.1 SV-D600/B 7" SkyView SE Display 1 4.2 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 1 4.3 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 1 4.4 SV-GPS-2020 GPS Receiver/Antenna 1 4.5 SV-XPNDR-261 Mode-S Transponder 1 4.6 SV-KNOB-PANEL/H SkyView Knob Control Panel 1 4.7 SV-COM-C25/H SkyView VHF Com Radio 25 kHz 1 4.8 SV-COM-X83 SkyView VHF Com Radio 8.33 kHz 1 4.9 SV-INTERCOM-2S Two-Place Stereo Intercom 1 4.10 SV-NET-3CC Network Cable 1 4.11 AOA - Pitot Probes 1 4.12 AOA - Pitot Mount Bracket 1 4.13 J-3 8-Gal. Wing Tank + Extra Elements 1 4.14 Thermometer Coolant for ROTAX 912/914 1 4.15 IM-554 EGT Gauge for ROTAX 914 1 4.16 IM-584 Fuel Pressure Gauge for Rotax 912 BAR 1 4.17 IAE50R – AA Engine Starter 1 4.18 IM-543 Oil Pressure Gauge for Rotax 912/914 1 4.19 IM-560 Oil Temp Indicator for ROTAX 912S 1 4.20 IM-561 Voltimeter 1 4.21 WINTER 7 FMS 5 Aispeed Indicator 1 4.22 WINTER EBH Altimeter 1 4.23 WINTER Vanetype Variometer 5 STV 5 1 4.24 Kelly MFG RCA22 Artificial Horizon 1 4.25 Kelly RCA15BK-1 Directional Gyro 1 4.26 WINTER Turn Coordinator QM II 1 4.27 WINTER Hours Counter Analog 1 4.28 Westach Dual Fuel Level Gauge Model: 2DA4 1 4.29 Aerovoltz 16 cell battery 1 4.30 Stainless Steel Cable 141 ft. 4.31 Contorl Stick Grip 1 4.32 Simpson 5-Point Harnesses System 1 4.33 H3R Aviation Model A344T - Halon 1211 fire Ext. 1 4.34 Poliester Straps 1 4.35 BRS 7 LSA Canister Parachute System 1 4.36 Aveo Ultra Galactica TM Series lights 1 4.37 Aveo Hercules 30 Landing Light Module 1 4.38 Aerovoltz 16 battery 1
Table 1.3: Units of each element of the Avionics department
5. Design ID Content Quantity 5.1 Seats 3 5.2 Cover 3
Table 1.4: Units of each element for the Design Department
B 8 Chapter 2
Costs
This section contains the unit cost of all the elements the project involves, separated by de- partment.
2. Structures ID Content Unit Cost [e] 2.1 Carbon Fiber (4mm) 740 2.2 Pneumatics 87 2.3 Carbon Fiber (3mm) 630 2.4 Carbon Fiber (1mm) 360 2.5 Honeycomb 74 2.6 Carbon Fiber (2mm) 52 2.7 Windshield 51 2.8 Epoxy Resin 150 2.9 Aluminium (U shape) 80 2.10 Landing Gear 800 2.11 Carbon Fiber Cloth 66,67
Table 2.1: Unit Cost of each element for the Structures Department
3. Power Plant ID Content Unit cost [e] 3.1 Motor 9.000 3.2 Motor Mount 980 3.3 Gear Box 1.412 3.4 Micro louver air shield 54 3.5 Exhaust Clamp 595 3.6 Ceramic heatshield 154 3.7 Muffer Shroud 596 3.8 Stainless Steel 90o Bend 146 3.9 Homebuilders Exhaust Stacks 868 3.10 Custom Propoeller (Optional) 5.000 3.11 HauteClaire 3 blade 724
Table 2.2: Unit Cost for each element of the Power Plant department
B 9 G06-AlOn LSA 3 seats | Budget
4. Avionics ID Content Unit cost [e] 4.1 SV-D600/B 7" SkyView SE Display 1.622 4.2 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 1.052 4.3 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 517 4.4 SV-GPS-2020 GPS Receiver/Antenna 176 4.5 SV-XPNDR-261 Mode-S Transponder 1.930 4.6 SV-KNOB-PANEL/H SkyView Knob Control Panel 220 4.7 SV-COM-C25/H SkyView VHF Com Radio 25 kHz 1.135 4.8 SV-COM-X83 SkyView VHF Com Radio 8.33 kHz 1.925 4.9 SV-INTERCOM-2S Two-Place Stereo Intercom 260 4.10 SV-NET-3CC Network Cable 35 4.11 AOA - Pitot Probes 175 4.12 AOA - Pitot Mount Bracket 84 4.13 J-3 8-Gal. Wing Tank + Extra Elements 601 4.14 Thermometer Coolant for ROTAX 912/914 90 4.15 IM-554 EGT Gauge for ROTAX 914 163 4.16 IM-584 Fuel Pressure Gauge for Rotax 912 BAR 63 4.17 IAE50R – AA Engine Starter 62 4.18 IM-543 Oil Pressure Gauge for Rotax 912/914 62 4.19 IM-560 Oil Temp Indicator for ROTAX 912S 62 4.20 IM-561 Voltimeter 116 4.21 WINTER 7 FMS 5 Aispeed Indicator 455 4.22 WINTER EBH Altimeter 380 4.23 WINTER Vanetype Variometer 5 STV 5 530 4.24 Kelly MFG RCA22 Artificial Horizon 1.033 4.25 Kelly RCA15BK-1 Directional Gyro 970 4.26 WINTER Turn Coordinator QM II 56 4.27 WINTER Hours Counter Analog 373 4.28 Westach Dual Fuel Level Gauge Model: 2DA4 125 4.29 Aerovoltz 16 cell battery 334 4.30 Stainless Steel Cable 0.25 4.31 Contorl Stick Grip 135 4.32 Simpson 5-Point Harnesses System 842 4.33 H3R Aviation Model A344T - Halon 1211 fire Ext. 328 4.34 Poliester Straps 30 4.35 BRS 7 LSA Canister Parachute System 4.274 4.36 Aveo Ultra Galactica TM Series lights 1.199 4.37 Aveo Hercules 30 Landing Light Module 625 4.38 Aerovoltz 16 battery 293
Table 2.3: Unit Cost for each element of the Avionics department
5. Design ID Content Unit Cost [e] 5.1 Seats 114 5.2 Cover 86
Table 2.4: Unit Cost of each element for the Design Department
B 10 Chapter 3
Budget
This section contains the final budget for the whole project, with separate tables for each department, containing the Unit cost, the Units, and the Total Cost of each department. Once calculated this subtotals, the Grand Total can be computed. It has been added an extra table, that does not belong to any department, as it consists on the Fixed Costs. These costs are extra costs, like Insurance, Rent, Marketing..(see Business Attachment for specific information).
2. Structures ID Content Unit Cost [e] Quantity Cost [e] 2.1 Carbon Fiber (4mm) 740 2 1.480 2.2 Pneumatics 87 3 261 2.3 Carbon Fiber (3mm) 630 3 1.890 2.4 Carbon Fiber (1mm) 360 4 1.440 2.5 Honeycomb 74 1 74 2.6 Carbon Fiber (2mm) 52 1 52 2.7 Windshield 51 1 51 2.8 Epoxy Resin 150 4 600 2.9 Aluminium (U shape) 80 8 640 2.10 Landing Gear 800 1 800 2.11 Carbon Fiber Cloth 67 150 10.000 Total 17.288 e
Table 3.1: Budget for the Structures Department
3. Power Plant ID Content Unit cost [e] Quantity Cost [e] 3.1 Motor 9.000 1 9.000 3.2 Motor Mount 980 1 980 3.3 Gear Box 1.412 1 1.412 3.4 Micro louver air shield: 54 4 216 3.5 Exhaust Clamp 595 2 1.190 3.6 Ceramic heatshield 154 2 308 3.7 Muffer Shroud 596 1 596 3.8 Stainless Steel 90o Bend 146 1 146 3.9 Homebuilders Exhaust Stacks 868 1 868 3.10 Custom Propoeller (Optional) 5.000 1 5.000 3.11 HAUTECLAIRE 3 blade 724 1 724 Total 15.440 e
Table 3.2: Budget for the Power Plant department
B 11 G06-AlOn LSA 3 seats | Budget
4. Avionics ID Content Unit cost [e] Quantity Cost [e] 4.1 SV-D600/B 7" SkyView SE Display 1.622 1 1.622 4.2 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 1.052 1 1.052 4.3 SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System 517 1 517 4.4 SV-GPS-2020 GPS Receiver/Antenna 176 1 176 4.5 SV-XPNDR-261 Mode-S Transponder 1.930 1 1.930 4.6 SV-KNOB-PANEL/H SkyView Knob Control Panel 220 1 220 4.7 SV-COM-C25/H SkyView VHF Com Radio 25 kHz 1.135 1 1.135 4.8 SV-COM-X83 SkyView VHF Com Radio 8.33 kHz 1.925 1 1.925 4.9 SV-INTERCOM-2S Two-Place Stereo Intercom 260 1 260 4.10 SV-NET-3CC Network Cable 35 1 35 4.11 AOA - Pitot Probes 175 1 175 4.12 AOA - Pitot Mount Bracket 84 1 84 4.13 J-3 8-Gal. Wing Tank + Extra Elements 601 1 601 4.14 Thermometer Coolant for ROTAX 912/914 90 1 90 4.15 IM-554 EGT Gauge for ROTAX 914 163 1 163 4.16 IM-584 Fuel Pressure Gauge for Rotax 912 BAR 63 1 63 4.17 IAE50R – AA Engine Starter 62 1 62 4.18 IM-543 Oil Pressure Gauge for Rotax 912/914 62 1 62 4.19 IM-560 Oil Temp Indicator for ROTAX 912S 62 1 62 4.20 IM-561 Voltimeter 116 1 116 4.21 WINTER 7 FMS 5 Aispeed Indicator 455 1 455 4.22 WINTER EBH Altimeter 380 1 380 4.23 WINTER Vanetype Variometer 5 STV 5 530 1 530 4.24 Kelly MFG RCA22 Artificial Horizon 1.033 1 1.033 4.25 Kelly RCA15BK-1 Directional Gyro 970 1 970 4.26 WINTER Turn Coordinator QM II 56 1 56 4.27 WINTER Hours Counter Analog 373 1 373 4.28 Westach Dual Fuel Level Gauge Model: 2DA4 125 1 125 4.29 Aerovoltz 16 cell battery 334 1 334 4.30 Stainless Steel Cable 0.25 141 ft 35 4.31 Contorl Stick Grip 135 1 135 4.32 Simpson 5-Point Harnesses System 842 1 842 4.33 H3R Aviation Model A344T - Halon 1211 fire Ext. 328 1 328 4.34 Poliester Straps 30 1 30 4.35 BRS 7 LSA Canister Parachute System 4.274 1 4.274 4.36 Aveo Ultra Galactica TM Series lights 1.199 1 1.199 4.37 Aveo Hercules 30 Landing Light Module 625 1 625 4.38 Aerovoltz 16 battery 293 1 293 Total 22.367 e
Table 3.3: Budget for the Avionics and Systems department
5. Design ID Content Unit Cost [e] Quantity Cost [e] 5.1 Seats 114 3 342 5.2 Cover 86 3 258 Total 600 e
Table 3.4: Budget for the Design Department
B 12 G06-AlOn LSA 3 seats | Budget
Fixed Costs Rent 8.272,00 Additional 6.720,00 Insurance 3.500,00 Marketing 8.650,00 Personnel 210.620,80 Total 237.762,80 e
Table 3.5: Additional Fixed Costs
To be able to approximate more precisely the cost of manufacturing a single airplane, we have to consider the Costs of production, being these the costs relative to the assembly and construction of the different parts of the aircraft. In this project, they have been estimated as a 35% of the Material’s costs. So it gives 0, 35 ∗ 55.695, 25 ≈ 19.500. Now that we have all the costs of the elements needed to produce an aircraft, we can proceed to calculate the Grand Total.
Materials’ Costs = 55.695,25 e Production Costs = 19.500,00 e Fixed Costs = 237.762,80 e Grand Total ≈ 312.957,00 e .
This is the total capital needed to produce a single aircraft. As it is studied in the Economical Feasibility on the Report, the objective of the company is to sell 10 UNITS every year. Since the Fixed Costs do not depend on the number of aircrafts manufactured, only the materials’ costs will grow with each unit, making the project more profitable with each unit sold.
B 13 POLYTECHNIC UNIVERSITYOF CATALONIA
AEROSPACE ENGINEERING
Engineering Projects
3 seat Light-Sport Aircraft development | AlOn Contents: Lessons Learned
Group: G06-AE-2018/19-Q1
Delivery date: 20/12/2018
Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. G06-AlOn LSA 3 seats | Lessons Learned 1 Lessons Learned
LESSONS LEARNED LOG Improvement Category Situation de- Impact in Analysis of recommendations scription project cause Time manage- Difficulties Delays in the Members of Take into ac- ment to complete project sched- the team that count other scheduled tasks ule those tasks responsibilities on time assigned could of the team not complete members and them on time the time they for any reason can spend do- ing the project daily Time man- Preparation Re-scheduling It was not Initially take agement and of additional scheduled a into account Preparation of documents was time to specif- how many doc- the documents not initially ically prepare uments must scheduled the required be presented documents and add to the planning time to do them properly Time manage- Bad planning of Delays and Some tasks Better analysis ment time that tasks blanks on the were not prop- of tasks during took project imple- erly analyzed the planning mentation during the phase planning and had more/less time than needed as- signed Knowledge Tasks required Took more time The lack of Use software transition knowledge to do those knowledge that team mem- team members tasks made some bers dominate did not have tasks slower or search for due to the need specialized of learning how workers to do the task Decision mak- Lack of deci- Delay in the Decision tools Proper study ing sion making project tasks were not given of decision tools finish to the project making tools team members before starting until the project the project was really ad- vanced and some tasks had to be delayed because the impossibility of finishing
LL 1 G06-AlOn LSA 3 seats | Lessons Learned
Table 1 continued from previous page LESSONS LEARNED LOG Category Situation de- Impact in Analysis of Improvement scription project cause recommenda- tions Criteria deci- Difficulties to Delays and Lack of knowl- Have experi- sion select between unnecessary edge on which enced workers different crite- (even incorrect criteria was bet- ria while doing in some cases) ter to analyze the tasks work done due some parts to the usage of the wrong methods or criteria Planning Lack of knowl- Tasks were a bit It was compul- Further study edge on how to changed along sory to define of the project divide an air- the project de- a scope before parts during craft design in velopment the project was the elaboration tasks started and the of the project team members charter and bet- did not know ter definition of how much the scope work they were able to finish in the given time and how to create tasks starting from that Roles and dis- Bad distribu- Delay in the Unknown Better study of tribution tion of team completion of quantity of the tasks before members in some tasks due work required its scheduling some tasks to have had by some tasks assigned less led to a bad members to its assignment of realization than team members needed to do it Economics of Lack of knowl- Problems and Team mem- Provide team the project edge of meth- delays in the bers did not members of ods used to calculation of have enough further infor- calculate eco- economic infor- knowledge on mation in order nomic details mation of the methods of to do this anal- project economic esti- ysis faster and mation until a better really advanced point of the project
LL 2 G06-AlOn LSA 3 seats | Lessons Learned
Table 1 continued from previous page LESSONS LEARNED LOG Category Situation de- Impact in Analysis of Improvement scription project cause recommenda- tions Economics of Lack of in- Unrealistic eco- Many manu- Destine more the project formation on nomic planning facturers did time to search materials prices not provide materials, and form of prices of pieces, prices and costs selling instruments to improve and materials the economic which led to planification an estimation of those costs and made the economic ap- proximations unrealistic
LL 3 POLYTECHNIC UNIVERSITYOF CATALONIA
AEROSPACE ENGINEERING
Engineering Projects
3 seat Light-Sport Aircraft development | AlOn Contents: Report Attachments
Group: G06-AE-2018/19-Q1
Delivery date: 20/12/2018
Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. G06-AlOn LSA 3 seats | Project report
RA 2 Contents
1 Aerodynamics 13 1.1 Wing ...... 13 1.1.1 Airfoil selection ...... 13 1.1.1.1 Polar distribution ...... 14 1.1.1.2 Efficiency versus α ...... 14 1.1.1.3 Momentum versus α ...... 15 1.1.2 General Wing Plant forms ...... 15 1.1.2.1 Rectangular ...... 16 1.1.2.2 Elliptical ...... 16 1.1.2.3 Trapezoidal ...... 16 1.1.2.4 Delta ...... 16 1.1.3 Plant form definition ...... 16 1.1.3.1 Initial efficiency analysis ...... 17 1.1.3.2 Airfoil definition result ...... 17 1.1.3.3 Validation of the initial wing parameters configuration . . . . . 18 1.1.3.4 Wing Efficiency ...... 20 1.1.3.4.1 Aerodynamic Torsion ...... 20 1.1.3.4.2 Wingtips ...... 20 1.1.3.5 Stall Behaviour ...... 21 1.1.3.6 High-lift Device ...... 22 1.1.4 Wing configuration criteria ...... 23 1.1.4.1 Unweighted Average Method ...... 24 1.1.4.2 Ordered Weighted Average Method ...... 24 1.2 Tail ...... 24 1.2.1 Definition of tail ...... 24 1.2.2 Utility of a tail ...... 24 1.2.3 General tail designs ...... 25 1.2.3.1 Conventional tail design ...... 25 1.2.3.2 T-tail design ...... 26 1.2.3.3 Cruciform tail design ...... 26 1.2.3.4 V-tail design ...... 27 1.2.3.5 Twin-tail design ...... 27 1.2.3.6 Boom-tail design ...... 27 1.2.3.7 Dual tail design ...... 27 1.2.3.8 Triple tail ...... 27 1.2.3.9 Other tail configurations ...... 27 1.2.4 Tail configuration selection ...... 27 1.2.4.1 Simple Hierarchy Method ...... 28 1.2.4.2 Unweighted Average Method ...... 28 1.2.4.3 Ordered Weighted Average Method ...... 29 1.2.5 Decision making ...... 29 1.2.6 Tail design ...... 29 1.2.6.1 Parameters determination ...... 29
RA 3 G06-AlOn LSA 3 seats | Project report 1.2.6.1.1 Horizontal tail ...... 30 1.2.6.1.2 Vertical tail ...... 31 1.3 Fuselage ...... 31 1.3.1 Fuselage selection criteria ...... 33 1.3.1.1 Unweigthed Average Method ...... 34 1.3.1.2 Ordered Weighted Average Method ...... 34 1.4 Control Surfaces ...... 34 1.4.1 Elevator ...... 35 1.4.2 Ailerons ...... 36 1.4.3 Rudder ...... 37 1.4.4 Parasite Drag ...... 38 1.4.4.1 Wheels parasite drag ...... 39 1.4.4.2 Tube parasite drag ...... 40 1.4.4.3 Drag correction ...... 40 1.5 Final plane configuration ...... 40 1.5.1 Design parameters ...... 40 1.5.2 Range study ...... 41 1.6 Final Plane Analysis ...... 43 1.6.1 Efficiency ...... 43 1.6.2 Static Stability ...... 43 1.6.3 Dynamic Stability ...... 44 1.6.3.1 Longitudinal Mode 1 ...... 44 1.6.3.2 Lateral Mode 1 ...... 44 1.6.3.3 Lateral Mode 2 ...... 45 1.7 Flight envelope ...... 45
2 Structures 49 2.1 Materials ...... 49 2.1.1 Fuselage and wings internal structure ...... 49 2.1.2 Skin of the aircraft ...... 51 2.1.3 Landing gear ...... 52 2.1.4 Transparent surfaces ...... 53 2.1.4.1 Windshield ...... 53 2.1.4.2 Windows ...... 53 2.1.5 Additional material considerations ...... 53 2.1.5.1 Welding ...... 53 2.1.5.2 Corrosion Prevention ...... 53 2.2 Landing Gear ...... 54 2.2.1 Analysis of the landing gear regulations for an LSA ...... 54 2.2.1.1 Landing gear options ...... 55 2.2.1.2 Atec 322 Faeta ...... 55 2.2.1.3 TL-2000 Sting S4 ...... 55 2.2.1.4 TL-3000 Sirius ...... 55 2.2.1.5 Pipistrel Taurus M ...... 56 2.2.1.6 Alexander Schleicher ASG 29 E ...... 56 2.2.1.7 ONE Aircraft ...... 56 2.2.1.8 Sling 4 ...... 56 2.2.2 Landing gear calculations ...... 57 2.2.2.1 Static analysis ...... 57 2.3 Wing ...... 62 2.3.1 Initial analysis ...... 62 2.3.1.1 Lift approximation ...... 63 2.3.1.2 Weight approximation ...... 63 2.3.1.3 Results ...... 63
RA 4 G06-AlOn LSA 3 seats | Project report 2.3.2 Sizing of the beam ...... 64 2.3.2.1 I section ...... 65 2.3.2.2 Square section ...... 66 2.3.2.3 Results ...... 67 2.3.3 Ribs ...... 68 2.3.4 Final Result ...... 69 2.4 Fuselage ...... 70 2.4.1 Analysis ...... 72 2.4.1.1 Performed analysis ...... 73 2.4.1.2 Results of the analysis ...... 73 2.5 Tail ...... 77 2.5.1 Elevator beam ...... 77 2.5.2 Fin beam ...... 79 2.5.3 Ribs ...... 81 2.5.4 Final Result ...... 81
3 Overall Design 83 3.1 3D design ...... 83 3.1.1 Exterior ...... 83 3.1.2 Interior ...... 84 3.1.2.1 Structure ...... 84 3.1.2.2 Power plant ...... 84 3.1.2.3 Instruments and equipment ...... 85 3.1.2.4 Seats and passengers ...... 85 3.2 Blueprints ...... 87
4 Business 89 4.1 Manufacturing costs ...... 89 4.1.1 Cost of facilities ...... 89 4.1.2 Cost of human resources ...... 91 4.1.3 Additional costs ...... 92 4.1.4 Marketing costs ...... 92 4.1.4.1 Initial marketing campaign costs ...... 92 4.1.4.2 Study of marketing costs over time ...... 93 4.1.4.3 Study of future marketing campaigns ...... 94 4.2 Marketing campaign ...... 95 4.2.1 SWOT analysis ...... 95 4.2.1.1 Research on Market Opportunities ...... 95 4.2.1.2 Research on Market Threats ...... 95 4.2.1.3 Research on Aircraft Strengths ...... 96 4.2.1.4 Research on Aircraft Weaknesses ...... 96 4.2.2 Study of potential customers ...... 96 4.2.3 Set marketing goal ...... 97 4.2.4 Study of advertisement ...... 98 4.2.5 Design of marketing campaign ...... 98 4.3 Initial investment ...... 100 4.3.1 Research on possible investors ...... 100 4.4 Payback Analysis ...... 100 4.4.1 Study of profitability margins ...... 100
5 Organization, planning and scheduling. 103 5.1 Gantt Diagram ...... 103
6 Minutes of the Meeting 105
RA 5 G06-AlOn LSA 3 seats | Project report Bibliography 133
A Code 137 A.1 Weigth-Range Diagram ...... 137 A.2 Gust-Airspeed envelope ...... 138 A.2.1 Second grade polynomial fitting ...... 140 A.2.2 First grade polynomial fitting ...... 140 A.3 Propeller Design ...... 140 A.3.1 Variable declaration ...... 141 A.3.2 Core function ...... 141 A.3.3 importAirfoil ...... 141 A.3.4 discretization ...... 142 A.3.5 nonDymensionalization ...... 142 A.3.6 getLambdaInduced ...... 142 A.3.7 getSpecs ...... 144 A.3.8 Energy ...... 145 A.3.9 Other secondary functions ...... 145 A.4 Gantt diagram ...... 147
RA 6 List of Figures
1.1 Polar distribution of the selected airfoils ...... 14 1.2 Cl versus α of the selected airfoils ...... 14 Cd 1.3 Momentum coefficient of the airfoils selected ...... 15 1.4 Efficiency curve for the N2305 and the B24 airfoils ...... 18 1.5 NACA 23015 ...... 18 1.6 Trapezoidal Wing, S=7,5 m2 ...... 19 1.7 Trapezoidal Wing, S=12 m2 ...... 19 1.8 Efficiency comparison of trapezoidal wings with different surface ...... 19 1.9 Efficiency comparison between wing with and without aerodynamic torsion . . 20 1.10 Winglets ...... 21 1.11 Elliptic Tip ...... 21 1.12 Efficiency comparison between elliptic wingtips, winglets and without wingtips 22 1.13 Cl comparison between torsion and non-torsion wing ...... 22 1.14 Efficiency comparison between torsion and non-torsion wing ...... 23 1.15 Tail configurations ...... 26 1.16 Seats configuration ...... 32 1.17 Efficiency in both configuration ...... 33 1.18 Cm in both configurations ...... 33 1.19 Elevator Airfoil ...... 35 1.20 Hmom Elevator ...... 35 1.21 Cl versus α for the elevator at its maximum deflection angle ...... 36 1.22 Aileron Airfoil ...... 36 1.23 Hmom Ailerons ...... 37 1.24 Rudder Airfoil ...... 37 1.25 Hmom Rudder ...... 38 1.26 Cl versus α fortherudderatitsmaximumdeflectionangle ...... 38 1.27 CD experimental data ; Source: researchgate.net ...... 39 1.28 Comparison between both behaviours ...... 40 1.29 Weight-Range diagram ...... 42 1.30 Efficiency of the final configuration ...... 43 1.31 Static stability of the final configuration ...... 43 1.32 Longitudinal Mode 1 ...... 44 1.33 Lateral Mode 1 ...... 45 1.34 Lateral Mode 2 ...... 45 1.35 Airspeed envelope diagram ...... 47 1.36 Gust-Airspeed envelope diagram ...... 48
2.1 CentrAl configuration ...... 50 2.2 First variant of the TL-3000 Sirius’ landing gear ...... 55 2.3 Second variant of the TL-3000 Sirius’ landing gear ...... 56 2.4 ASG 29 E flying ...... 56 2.5 Front leg of the landing gear...... 57 2.6 Half of the rear landing gear...... 57
RA 7 G06-AlOn LSA 3 seats | Project report 2.7 Stress analysis of the front landing gear...... 59 2.8 Displacements of the front landing gear...... 60 2.9 Stress analysis of half of the rear landing gear...... 60 2.10 Displacements of half of the rear landing gear...... 61 2.11 Tension results of the dropping test...... 61 2.12 Detail of the point where the tension gets its maximum value...... 61 2.13 Displacement results of the dropping test...... 62 2.14 Loads distribution and approximations ...... 62 2.15 Moment diagram for n=1 ...... 63 2.16 Moment diagram for n=4 ...... 64 2.17 Moment diagram for n=-2 ...... 64 2.18 Stress distribution for composite beam ...... 65 2.19 I section parameters ...... 65 2.20 Stress distribution ...... 66 2.21 Square section parameters ...... 66 2.22 Stress analysis ...... 67 2.23 Displacement results ...... 68 2.24 Sandwich ribs ...... 68 2.25 Section approximation ...... 69 2.26 Beam and ribs ...... 70 2.27 Fuselage with a beam thickness of 1 cm...... 71 2.28 Fuselage with a beam thickness of 0.5 cm...... 71 2.29 Fuselage with a beam thickness of 0.5 cm and a reinforcement...... 71 2.30 Fuselage with a U beam...... 72 2.31 Fuselage with a rectangle beam...... 72 2.32 10mm thick I:Results of the stress analysis...... 74 2.33 10mm thick I:Results for the displacement...... 75 2.34 5mm thick I:Results of the stress analysis...... 75 2.35 5mm thick I:Results for the displacement...... 75 2.36 5mm thick I with reinforcement:Results of the stress analysis...... 76 2.37 5mm thick I with reinforcement:Results for the displacement...... 76 2.38 3,5mm thick U:Results of the stress analysis...... 76 2.39 3,5mm thick U:Results for the displacement...... 77 2.40 16mm Rectangle: Results of the stress analysis...... 77 2.41 16mm Rectangle:Results for the displacement...... 77 2.42 Elevator momentum diagram ...... 78 2.43 Elevator FEM analysis ...... 79 2.44 Fin moment diagram ...... 80 2.45 Fin FEM analysis ...... 80 2.46 Beams and ribs ...... 81
3.1 Exterior model ...... 83 3.2 Full structure ...... 84 3.3 Power plant model ...... 84 3.4 Instruments displayed, Cockpit ...... 85 3.5 Passenger model ...... 85 3.6 Seats and prices of Air-Tech Inc.[23] ...... 86 3.7 Measures, technical sheet of the Ultralight seat [23] ...... 86 3.8 Initial sizing ...... 87
4.1 Specifications of Office 1 ...... 90 4.2 Specifications of Office 2 ...... 90 4.3 Specifications of Office 3 ...... 90 4.4 Specifications of Office 4 ...... 91
RA 8 G06-AlOn LSA 3 seats | Project report 4.5 Specific weigh t by category ...... 92 4.6 Scheme of a SWOT analysis...... 95 4.7 Airplane Shipments Worldwide (1995–2017) ...... 96 4.8 Different logo proposals ...... 99 4.9 Net Present Value (years = 7, k = 11,4%) for different selling prices ...... 101 4.10 Pay-Back Time for different selling prices ...... 101 4.11 Internal Rate of Return for different selling prices ...... 101
RA 9 G06-AlOn LSA 3 seats | Project report
RA 10 List of Tables
1.1 Wing parameters of the most similar airplanes and gliders ...... 17 1.2 AlphaOne initial main parameters configuration ...... 17 1.3 Analysis configuration ...... 17 1.4 Initial wing definition ...... 19 1.5 Initial wing definition ...... 19 1.6 Final wing definition ...... 23 1.7 Wing Unweighted Average Method ...... 24 1.8 Wing Ordered Weigthed Average method ...... 24 1.9 Simple Hierarchy Method application ...... 28 1.10 Unweighted Average Method application ...... 28 1.11 Ordered Weighted Average Method application ...... 29 1.12 Unweighted Average Method application(fuselage) ...... 34 1.13 Ordered Weighted Average Method application(fuselage) ...... 34 1.14 Final wing configuration data ...... 41 1.15 Final fin configuration data ...... 41 1.16 Final elevator configuration data ...... 41 1.17 Control surfaces main sizes ...... 41
2.1 Global properties of aluminium ...... 50 2.2 Properties of components of CentrAl ...... 50 2.3 Properties of CentrAl ...... 51 2.4 Properties of T300 ...... 52 2.5 Properties of T400H ...... 52 2.6 Properties of T1000G ...... 52 2.7 Results of the static analysis of the different alternatives of the landing gear. . . 58 2.8 Results of the application of OWA method ...... 58 2.9 Properties of the steel 4130 normalized at 870 ºC...... 59 2.10 Load distribution for n=1 ...... 63 2.11 Lift distribution for each load factor ...... 63 2.12 Beam configurations and results ...... 66 2.13 Beam configurations and results ...... 66 2.14 Results of buckling analysis ...... 69 2.15 Principal results of the different fuselage alternatives analysis ...... 74 2.16 Results of OWA method ...... 74 2.17 Total load and distribution ...... 78 2.18 Dimensions for elevator beam. Referred to Fig. 2.21 ...... 78 2.19 Total load and distribution ...... 79 2.20 Dimensions for fin beam. Referred to Fig. 2.21 ...... 80 2.21 Buckling analysis for fin ...... 81 2.22 Buckling analysis for elevator ...... 81
4.1 Comparative of the 4 possibilities ...... 91 4.2 Basic human resources cost ...... 91
RA 11 G06-AlOn LSA 3 seats | Project report 4.3 Marketing costs by category ...... 93 4.4 Evolution of Airplane Shipments worldwide between 1995 and 2017 ...... 97
RA 12 Chapter 1
Aerodynamics
Due to the complexity that suppose the design of a three seat aircraft which can only weight up to 600Kg, the aerodynamic design is going to be a decisive part on this project. The main idea about the design of this aircraft is to create a hybrid between a traditional LSA (Light- Sport Aircraft) and a glider. The aim is to combine the efficiency of a glider and the main characteristics of the LSA in order to develop an aircraft light enough to be able to transport 3 passengers, including the pilot, without exceeding 600 Kg and having an acceptable performance. Taking into consideration the previous aspects, the aerodynamics study will be divided in the following sections:
1. Wing
2. Tail
3. Fuselage
4. Combined study
1.1 Wing
1.1.1 Airfoil selection The airfoil selection began searching for information about airplanes and gliders that meet the requirements of our configuration or performance characteristics. The possible airfoils that could fit in our model are:
• Davis B-24
• NACA 23015 [1]
• NACA 4412 [2]
• GA-30415 [3]
• FXS0-2196 [2]
The airfoils mentioned above were the ones that could give the best performance to our air- plane. In order to validate our decision, an airfoil analysis with XFLR5 has been done for a range of Reynolds numbers and α.
It could not be found the distribution of the GA-30415 airfoil, which in particular was a char- acteristic one because it was a modification of a 4415 and had a very good performance in terms of efficiency and momentum. The airfoil was found in the book [3] but there was no clue of it neither on the internet nor in some well-known airfoil databases.
RA 13 G06-AlOn LSA 3 seats | Project report 1.1.1.1 Polar distribution
Figure 1.1: Polar distribution of the selected airfoils
As shown in Fig. 1.1 the airfoils can be separated in two groups: 1. NACA 23015 and 4415 2. Davis B-24 and FXS0-2196
The first group has wider range of Cl, the NACAs almost get a Cl = 1, 75 for a Cd = 0, 02; while the second group gets its maximum Cl practically when Cd = 0, both values are really close. As can be seen, the most significant differences are the way the curves evolve with the Cd and the maximum values.
1.1.1.2 Efficiency versus α
Figure 1.2: Cl versus α of the selected airfoils Cd
This graphic is a decisive one because it defines one of the most significant parameters, the efficiency. As can be seen in Fig. 1.2 there is quite a big efficiency difference amongst the four,
RA 14 G06-AlOn LSA 3 seats | Project report so as a first conclusion the best airfoil is thought to be the FXS0.
However, there is also another parameter that should be taken into account, the momentum coefficient, which could cause further problems trying to make up for the wing momentum with the horizontal tail.
So not necessarily the best efficiency airfoil will provide the best performance. On the next section, a balance will be made in order to decide the optimum airfoil.
1.1.1.3 Momentum versus α Another important parameter that could impact on the efficiency of our aircraft is the mo- mentum generated by the wing. In order to design a stable wing it is important to select an airfoil that can provide us an easy way to correct the momentum generated by the wing, which means a low momentum, in terms of the absolute value.
Figure 1.3: Momentum coefficient of the airfoils selected
As seen in Fig. 1.3, the most efficient airfoil analyzed (FXS0) has a huge Cm at α = 5º, which is precisely the angle of attack where it’s efficiency is maximum. A Cm = −0.1 would difficult the tail design, trying to compensate the huge momentum the wing could cause. In order to simplify the tail design and the position of the tail and wing within the fuselage, a wing efficiency analysis has been made using the David B-24 and the NACA 23015. These two airfoils provide the balance between Cm/E desired.
1.1.2 General Wing Plant forms The wing plant form is a very important aspect in design. It affects the velocity, efficiency, bending moment, fuel tanks, cost... and so on. There are different types of wing plant forms and each one is used in a specific application.They can be classified in 4 major groups used in airplanes:
• Rectangular
• Trapezoidal
• Elliptic
• Delta
RA 15 G06-AlOn LSA 3 seats | Project report 1.1.2.1 Rectangular
Is the simplest wing plant form and consequently the less expensive to build. However, it provides an undesired aerodynamic and structural behaviour. Nevertheless, is widely used in small general aviation because the aerodynamic demands are not very strict and the cost of fabrication is very low in comparison to wings with other plant forms.
1.1.2.2 Elliptical
It is the wing plant form with less induced drag. It has a constant Cl distribution which, with an elliptic chord distribution, results in elliptical lift distribution. This behaviour is suitable for aerodynamics and structures. However, it is not very used today because of its price, which is the highest, and also because it has also an undesired stall behaviour due to its constant Cl distribution, which means that at a certain angle of attack the whole wing stalls at once. To solve this problem swept was implemented in WWII emblematic planes, such as the Spitfire.
1.1.2.3 Trapezoidal
Is the most used in heavy transport planes such as commercial planes. A good trapezoidal design can achieve almost the same behaviour as the elliptic wing but in a shorter range of angles of attack, where the cruise angle of attack should be included. In addition, they are easier to produce and, consequently, less expensive than elliptic wings. However, the stall behaviour is not suitable, due to the fact that for very narrow trapezoidal wings the stall starts at the tip where the ailerons are often located. In commercial planes trapezoidal wings are designed with swept to increase the cruise speed and with some torsion to counter the stall non-desired behaviour.
1.1.2.4 Delta
Delta Wing plant form is the trapezoidal concept brought to the limit where the tip chord is 0. It is used in fighter jets in addition to swept because of its ability to increase the airplane speed in exchange for efficiency.
1.1.3 Plant form definition
The definition of the main parameters of the wing plant form is based on a combination of the data obtained from other similar aircrafts and a wing efficiency analysis using the airfoils mentioned in the previous section.
The next table contains the wingspan and area of the airplanes analyzed in order to do an initial sizing of the wing. As can be seen, the wing area is similar for all aircrafts; while the wingspan is divided in two groups:
• LSA < 10 m
• Gliders > 10 m
These values make sense considering the nature of each aircraft. As mentioned above, the plane designed in this project is neither an LSA nor a glider, but it is a combination of both. Due to this reason, the initial wing configuration is the following:
RA 16 G06-AlOn LSA 3 seats | Project report Wing span (m) Wing area (m2) Atec 322 Faeta 9,6 10,1 TL-2000 Sting S4 9,12 11,1 TL-3000 Sirius 9,4 11,26 Pipistrel Taurus M 14,97 12,33 Alexander Schleicher ASG 29 E 18 10,5
Table 1.1: Wing parameters of the most similar airplanes and gliders
Wing span (m) Wing area (m2) AlphaOne 14 11
Table 1.2: AlphaOne initial main parameters configuration
1.1.3.1 Initial efficiency analysis In order to validate the airfoils and the wing parameters selected, an efficiency analysis has been done. The objectives of this first analysis are:
1. Compare and define the best airfoil
2. Validate the wing parameters selected
3. Refinement and analysis of better possibilities.
1.1.3.2 Airfoil definition result In order to compare the two previous airfoils, an analysis with XFLR51 has been done to de- termine the maximum efficiency and the momentum coefficient Cm once the wing is set.
The analysis had the configuration shown in Table 1.3. As a result, shown in Fig. 1.4, the B24 has a higher maximum efficiency than the N23015, it is not a huge difference but could be a good improvement for the wing. But as regards performance of the airplane, it is also important to consider the flight mechanics and the different flight configurations it will have.
α -5º < α < 20º Weight 600 kg Gravity center (0,0) Polar type Fixed Lift Method Horshoe vortex (viscous corrections) kg ρ 1,225 m3 −5 m2 v 1, 5 ∗ 10 s Table 1.3: Analysis configuration
Taking into account the flight mechanics and considering that both top efficiency values are quite high, it would be a better choice to have a wider operative range of angle of attack α
1is an analysis tool for airfoils, wings and planes operating at low Reynolds Numbers.
RA 17 G06-AlOn LSA 3 seats | Project report than a smaller one with a bigger maximum efficiency.
A wider angle of attack will give the airplane a better manoeuvrability and will make possible bigger α0s for those flight configurations that could require them.
Figure 1.4: Efficiency curve for the N2305 and the B24 airfoils
Thus, the NACA 23015 is the airfoil chosen to define the aircraft’s wing. This airfoil can be seen in Fig. 1.5.
Figure 1.5: NACA 23015
1.1.3.3 Validation of the initial wing parameters configuration In order to validate and check if the values we estimated are good enough, an analysis for the parameters defined in Table 1.2 has been developed using the previous analysis configu- ration.
The first objective was to obtain the surface needed to make a horizontal flight with the maxi- mum efficiency. It was also important to consider that the stall speed in landing configuration should be 22,5m/s or less. First, a smaller surface, in comparison with similar planes, has been studied. The first configuration was a trapezoidal plant form with 7,5 m2:
RA 18 G06-AlOn LSA 3 seats | Project report
Wing span (m) 15 Root chord (m) 0,6 Tip chord (m) 0,4 Surface (m2) 7,5 Airfoil 23015 Aspect ratio 30 Maximum efficiency 51,1 Stall speed (m/s) 34,0
Figure 1.6: Trapezoidal Wing, S=7,5 m2 Table 1.4: Initial wing definition
This configuration is able to fly fast and has good efficiency but the stall speed is too high. In order to reduce the stall speed, an increase in the surface is needed. It was decided not to increase the wingspan in order to make the structure and the storing of the plane simpler. For this reason, the following configuration has a chord increase.
Wing span (m) 15 Root chord (m) 1 Tip chord (m) 0,6 Surface (m2) 12 Airfoil 23015 Aspect ratio 18,75 Maximum efficiency 44,7 Stall speed (m/s) 27,5
Table 1.5: Initial wing definition Figure 1.7: Trapezoidal Wing, S=12 m2
With this wider configuration, the stall speed is also higher than the maximum value desired. At this point, it is easy to see that a high-lift device is needed in order to achieve the desired stall speed without losing too much efficiency, because from the first to the second config- uration a reduction of efficiency has already occurred, produced by the enlargement of the surface. In the next figure, both configurations are compared.
Figure 1.8: Efficiency comparison of trapezoidal wings with different surface
The final wing will have these major parameters. From now on, some implementations will be done to improve the aerodynamic characteristics. Specifically, three main aspects need to be revised:
RA 19 G06-AlOn LSA 3 seats | Project report 1.1.3.4 Wing Efficiency The efficiency is a very important parameter in the design. An increment of this value allows producing the same lift with lower drag and, consequently, lower thrust. This makes possible the implementation of a smaller engine with smaller consumption, which involves using smaller fuel tanks. Overall, an increment of efficiency will end in a reduction of weight, which is a hard restriction in this project. In order to improve efficiency, two techniques have been used:
• Aerodynamic Torsion
• Wingtips
1.1.3.4.1 Aerodynamic Torsion The variation of the airfoil along the wingspan can also result in an improvement of efficiency. Since the tip of the wing will have smaller structural stress compared with the root, it can be thinner. This allows the airfoil at the tip to be a NACA 23012 with a linear evolution from the NACA 23015 of the root. In figure 1.9 is shown a comparison in efficiency between the trapezoidal wing and without aerodynamic torsion.
Figure 1.9: Efficiency comparison between wing with and without aerodynamic torsion
As can be seen, the configuration with different airfoils is more efficient, 45,7 in front of 44,7. After these results, the final configuration will implement aerodynamic torsion.
1.1.3.4.2 Wingtips The wingtips are able to reduce the induced drag by making a softer transition between the higher pressure Intrados and the lower pressure Extrados at the tip of the wing. Two types of wingtips have been studied:
• Winglets: These devices have a complex geometry because an almost 90 degrees soft turn is needed with strong narrowing. After several iterations the best results were obtained with the following geometry:
RA 20 G06-AlOn LSA 3 seats | Project report
Figure 1.10: Winglets
Because of its complex geometry, winglets are difficult to study and expensive to build.
• Elliptic wingtips: These devices consist in an elliptic chord distribution at the tip of the wing. The geometry is much simpler than winglets’ as it can be seen:
Figure 1.11: Elliptic Tip
The following graph shows the efficiency of the wing with elliptic wingtips, winglets and without wingtips: As can be seen, the wingtips improve the efficiency of the wing. Winglets and Elliptic wingtips show similar results being the second one a little bit better. As a consequence of the lower cost of production and the performance of the elliptic wingtips, the final wing configuration will incorporate them.
1.1.3.5 Stall Behaviour When the plane reaches the stall speed it is not desired that the whole wing enters the stall at the same time; neither the tip, because the ailerons control would be lost; nor the root,
RA 21 G06-AlOn LSA 3 seats | Project report
Figure 1.12: Efficiency comparison between elliptic wingtips, winglets and without wingtips because the stall wake could affect the tail. To solve this problem, +3º of geometrical torsion has been given to the central part of the wing, at 3,75m from the root. This implementation makes that zone to have more angle of attack and, consequently, more Cl. That results to be the first section to stall. The next figure shows it, in blue the torsion configuration and in red the non-torsion one:
Figure 1.13: Cl comparison between torsion and non-torsion wing
An efficiency study has also been done to know the repercussion of this implementation. It can be seen that it has moved the maximum efficiency angle but the value did not change significantly: This technique allows having a soft stall behaviour without losing efficiency and control of the plane. For these reasons, the final configuration will incorporate geometrical torsion.
1.1.3.6 High-lift Device The stall speed in landing configuration should not exceed 22,5m/s (45 Knots) as the regula- tion specifies. In table Table ?? it can be seen that the stall speed is above the maximum. An increase in the surface would fix it, but then the plane would be less efficient. To solve this
RA 22 G06-AlOn LSA 3 seats | Project report
Figure 1.14: Efficiency comparison between torsion and non-torsion wing problem a high-lift device can be implemented.
The configuration will implement a trailing edge simple flap due to its simplicity. The design has a 0,25 chord flap with a maximum deflection of 20º. The span of the flap is 4m and it is located at the root. This modification makes the stall speed go just under 22,5m/s in landing configuration.
The final wing configuration has these properties:
Wing span (m) 16 Root chord (m) 1 Tip chord (m) 0,6 Surface (m2) 12,5 Root airfoil 23015 Tip airfoil 23012 Aspect ratio 20,6 Maximum efficiency 46,7 Stall speed (m/s) 22,3
Table 1.6: Final wing definition
1.1.4 Wing configuration criteria Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efficiency for lower weight.
The comparison is based in four main criteria:
• Efficiency: it measures ratio between the functionality provided by the wing and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the wing. A higher value in this criterion indicates a lighter structure.
• Cost: it is related with the difficulty to build each wing, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
RA 23 G06-AlOn LSA 3 seats | Project report • Stall performance: it defines how good the stall performance is in each wing config- uration, when and how it starts and how it evolves. A higher grade in this criterion indicates a better performance. For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.
1.1.4.1 Unweighted Average Method
Unweigthed Average Method CRITERIA Trap Ellip Rec Trap +Elip final Trap + Wing Efficiency 3 3 2 5 5 Weight(struc) 4 3 4 3 2 Cost 4 2 5 4 3 Stall perf 5 2 2 5 5 SUM 16 10 13 17 15 Max 20 UA 2 0,5 0,65 0,85 0,75
Table 1.7: Wing Unweighted Average Method
1.1.4.2 Ordered Weighted Average Method
Ordered Weigthed Average Method Trapeziodal Elliptic Rectangular Trapeziodal + wingtip Trapeziodal + winglet CRITERIA P PxG P PxG P PxG P PxG P PxG Efficiency 3 12 3 12 2 8 5 20 5 20 Weight(struc) 4 12 3 9 4 12 3 9 2 6 Cost 4 4 2 2 5 5 4 4 3 3 Stall perf 5 10 2 4 2 4 5 10 5 10 SUM(PXG) 16 38 10 27 13 29 17 43 15 39 Max(PXG) 20 50 50 50 50 50 OWA 2 0,76 0,5 0,54 0,65 0,58 0,85 0,86 0,75 0,78
Table 1.8: Wing Ordered Weigthed Average method
As it can be seen in both tables the best option is the trapezoidal plus a wingtip at the end. Also it can be noticed that the simple trapezoidal is better than the trapezoidal with wingltes when using the OWA. So the final wing configuration is the trapezoidal plus the wingtips.
1.2 Tail
1.2.1 Definition of tail A tail or empennage is a structure located usually at the rear of an airplane which is in charge of providing aerodynamic stability during flight. It can also contain control surfaces used to deliberately change the yaw and pitching moment of the aircraft in order to adjust its trajec- tory. It is commonly composed of a horizontal stabilizer (or tailplane), a vertical stabilizer (or vertical fin) and two control surfaces: an elevator and a rudder.
1.2.2 Utility of a tail In conventional aircraft, tails are generally designed to fulfill a set of concrete roles. These roles are:
RA 24 G06-AlOn LSA 3 seats | Project report • Trim (longitudinal and directional)
• Stability (longitudinal and directional)
• Control (longitudinal and directional)
The first function, trimming, consists in maintaining balance, both longitudinal and direc- tional, during unsymmetrical flight conditions. Its major purpose is to eliminate the need for the pilot to keep constantly focusing on the control surfaces in order to maintain the equilib- rium personally. The trimming function is usually done by trim tab elements or by adjustable stabilizers.
The second function of a tail is to stabilize the plane throughout the flight. Airplane stability is defined as a tendency to recover from undesired transitory perturbations such as, primarily, atmosphere phenomena. The tailplane is in charge of longitudinal stability while the vertical fin is responsible to preserve the directional stability.
The third and most important responsibility of tails is the in-flight control of the aircraft, so as to the pilot can consciously modify its trajectory. The elevator is used to change the pitch moment of the plane and therefore adjust its longitudinal behavior. The rudder is used to change the yaw moment, thus leading to a directional adjustment of its behavior. Combining the rudder and the ailerons (usually located somewhere in the wing surface), the pilot is able to perform a coordinated turn of the plane.
1.2.3 General tail designs The first step to design the tail of our airplane is to decide the configuration to be used. In this case, it is convenient to compare different alternatives in a theoretical way evaluating the advantages and drawbacks for each option.
The most common tail designs used in civil aviation are the following (see figure 1.15):
Conventional tail T tail Cruciform tail Dual tail Triple tail V tail Inverted V tail Inverted Y tail Twin tail Boom tail High boom tail Multiple-plane tail
The main function of the tail of an airplane is to provide both stability and control in pitch and yaw. Taking into account the fact that our airplane is a hybrid between an LSA (Light-Sport Airplane) and a glider, the main aspects to consider to select the optimum tail configuration are not only the stabilization and control provided but also the weight and complexity of the structure required. From all the alternatives mentioned above, it is easy to see that some of them will not provide the desired performance. Moreover, the most common tail configura- tions for both LSAs and gliders are only conventional tail and T tail; the other configurations are rarely used and have become obsolete. In the following, every possible tail configuration will be analyzed.
1.2.3.1 Conventional tail design The conventional tail configuration is an aft tail design (rear-mounted tail) which is consid- ered as the simplest tail configuration. The horizontal stabilizer does not produce any load on the vertical stabilizer, consequently, weight and complexity can be reduced. Its overall performance (trim, stability and control) is presumably acceptable. It is also one of the light- est configurations. That’s why it is the most common configuration, especially in LSAs. The weak point of this design is the fact that the horizontal stabilizer is located in the wing wake, causing a loss of efficiency.
RA 25 G06-AlOn LSA 3 seats | Project report
Figure 1.15: Tail configurations
1.2.3.2 T-tail design
Another usual tail configuration in General Aviation airplanes is the T-tail, which is an aft tail configuration with the horizontal stabilizer located on top of the vertical fin. Its main ad- vantage over the conventional design is that during cruise flight the tailplane remains above of the regions disturbed by the wing wake, downwash, wing vortex and engine prop wash; thus, it provides a higher aerodynamic efficiency. The lower effects from the wings and en- gine also lead to a diminution in tail vibrations, greatly reducing its fatigue and increasing its durability. Nonetheless, the bending moment created by the elevator is also higher, so the vertical tail structure becomes heavier. Besides, its main disadvantage appears when flying at a high angle of attack, when the turbulent flow separated from the wings might incise upon the elevator and might result in a complete loss of the aircraft’s longitudinal control, situation known as deep stall.
1.2.3.3 Cruciform tail design
This tail design is a hybrid variation of the two previous designs. The horizontal empennage is located higher than in the conventional tail, so that it is away from the wing wake and the propeller flow, but it is not as high as in T tail configuration. With this design, the lower part of the stabilizer and the rudder receive undisturbed airflow due to the lifting force of the horizontal stabilizer. It is important to have undisturbed airflow on the rudder, especially to recover from spins. Although this configuration does not improve significantly the main strengths of its predecessors; it significantly reduces its major drawbacks, especially the deep stall distress.
RA 26 G06-AlOn LSA 3 seats | Project report 1.2.3.4 V-tail design In V-tail configuration instead of three surfaces (horizontal and vertical empennages with different variants), there are only two, which intend to serve the same function. The purpose of this tail design is to reduce the total tail area, as both parts of the tail act as horizontal and vertical stabilizers. Despite the fact that this configuration is quite suitable for trimming the aircraft, it is highly inefficient at maintaining its stability, letting the airplane really susceptible to perturbations. It is mainly employed in unmanned reconnaissance aircraft.
1.2.3.5 Twin-tail design This tail configuration is commonly used when a big tail is required but it would be too heavy or unfeasible to make a single regular tail. The big horizontal tail improves hugely the longitudinal performance of the airplane. In addition, all control surfaces remain clear of the fuselage wake region, thus increasing overall controllability, at a cost of a significant weight.
1.2.3.6 Boom-tail design The boom-mounted tail configuration is often used when the aircraft has a prop-driven en- gine installed at the rear of the aircraft, so as to the interference between the prop wash and the tail is reduced.
1.2.3.7 Dual tail design The placement of two vertical stabilizers at the ends of the horizontal one provide a better directional control in low-speed operations and allows a smaller and more aerodynamically efficient horizontal stabilizer. Although the size and consequently the weight of the horizon- tal empennage are reduced, the total weight of the tail is higher due to the addition of two vertical stabilizers instead of one as in other configurations.
1.2.3.8 Triple tail This configuration includes has two vertical stabilizers at the ends of the horizontal one and one more on the fuselage. This fact allows reducing the height of the tail but increases the total weight.
1.2.3.9 Other tail configurations As mentioned above, there are many different tail configurations, but the most common are the ones previously analyzed. Configurations such as inverted V tail, inverted Y tail or multiple-plane tail are rarely used and have been dismissed from the study because they are heavier and less efficient than the most common ones.
1.2.4 Tail configuration selection Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efficiency for lower weight. The comparison is based on four main criteria:
• Efficiency: it measures the ratio between the functionality provided by the tail and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the tail. A higher value in this criterion indicates a lighter structure.
RA 27 G06-AlOn LSA 3 seats | Project report • Cost: it is related to the difficulty to build each tail, which is also related to the com- plexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
• Stability and control: it defines the capability of the tail to develop its main function, and it is related to the performance of the plane. It also considers the facility of the tail to be immersed in the turbulent wake coming from other parts of the plane, which the control surfaces less effect. A higher grade in this criterion indicates a better performance.
For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.
1.2.4.1 Simple Hierarchy Method The scale used to evaluate each criterion goes from 1 to 10, and the threshold to reject inade- quate alternatives is 6. Then, the grades and results can be seen in the following table:
ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efficiency 4 10 8 Weight 10 6 8 Cost 8 6 4 Stability and control 6 8 6
Table 1.9: Simple Hierarchy Method application
The threshold set in 6 means that the alternatives that have one or more grades under this number are rejected. Analyzing the efficiency, the conventional tail is discarded; considering the cost of manufacturing, the cruciform is also discarded. In this way, the best alternative is the T tail.
1.2.4.2 Unweighted Average Method The new scale applied goes from 1 to 5, but the grades are proportionally calculated from the previous method in order to follow the same procedure for the different methods. The results obtained are shown in the next table:
ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efficiency 2 5 4 Weight 5 3 4 Cost 4 3 2 Stability and control 3 4 3 SUM 14 15 13 UA 0,7 0,75 0,65
Table 1.10: Unweighted Average Method application
This method considers the sum of all grades for each configuration to calculate an unweighted average; this means that the four criteria are considered equally important. This procedure indicates that the best alternative is the T tail.
RA 28 G06-AlOn LSA 3 seats | Project report 1.2.4.3 Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a specific weight for each criterion, which indicates the importance or weight that will have in the final result. The results obtained can be seen in the following table:
ALTERNATIVES Conventional Tail T-Tail Cruciform Tail WEIGHT CRITERIA P PxG P PxG P PxG G Efficiency 2 8 5 20 4 16 4 Weight 5 15 3 9 4 12 3 Cost 4 8 3 6 2 4 2 Stability and control 3 3 4 4 3 3 1 SUM (PxG) 34 39 35 OWA 0,68 0,78 0,7
Table 1.11: Ordered Weighted Average Method application
This method is possibly the most accurate among the ones used, due to the fact that each criterion is weighted in order to consider its importance. According to this method, the best alternative is again the T tail.
1.2.5 Decision making In this case, the best alternative is the same for all methods, so undoubtedly the final decision is the T tail configuration. The next step is to determine the main parameters of the tail in order to provide the necessary stability and control with the minimum increment of drag.
1.2.6 Tail design There are two main aspects to take into account from the definition and sizing of the tail. Firstly, it is important to remember its main functions: trim, stability and control. The mov- able surfaces of the tail, typically the rudder and the elevator, provide control; and the hor- izontal and vertical tails, stability, which is what is going to be studied first. The desired tendency of the plane is to automatically recover the initial state after a transitory perturba- tion, which is, in fact, the definition of a stable aircraft. However, the more stable a plane is, the less manoeuvrable. Due to this correlation, to correctly define the size of the tail, a midpoint should be found in order to design a stable plane but also controllable and easy to handle.
The second aspect to consider, especially in T-tail configuration, is the location of the horizon- tal stabilizer. As explained in a previous section, there is a risk of deep stall, which consists in the loss of control of the tail due to its immersion in the wing wake when the plane is flying at high angles of attack, causing the complete loss of the longitudinal control. On the one hand, to avoid this critical situation, the horizontal tail should be high enough to ensure it is not affected by the turbulent wake at high angles of attack. But on the other hand, an increase in the distance between the horizontal tail and the longitudinal plane axis involves a higher bending moment and, consequently, a heavier vertical tail structure is needed. Once again, the optimum design point should be found to avoid both drawbacks.
1.2.6.1 Parameters determination The procedure to determine the main parameters of the tail begins with a preliminary sizing having as a reference similar airplanes, which in this case are both LSAs and gliders. The
RA 29 G06-AlOn LSA 3 seats | Project report main reference has been the glider Pipistrel Taurus, due to its similar weight, fuselage length and wing.
Considering the reference data, an average or an approximated value can be used initially to design the tail and study the plane’s behaviour with XFLR5. Depending on the results, some parameters should be refined to achieve the optimum performance.
Moreover, a generic body has been created in order to study the relative position between the wing and the tail, which is a decisive parameter because it influences the longitudinal stability of the aircraft.
1.2.6.1.1 Horizontal tail
• Aspect ratio The main influences of the aspect ratio are the slope of the lift curve (CLα), the induced drag coefficient (CDi) and the structural weight. Considering the fact that the lift and drag produced by the horizontal tail are significantly smaller than those produced by the wing, it is more important to focus on the reduction of weight rather than on the aerodynamic forces to determine the aspect ratio.
When the aspect ratio decreases, the structural weight does it too. Consequently, the aspect ratio of the tail should be lower than that of the wing. As for the wing the value is about 20, for the tail this parameter is cut approximately by half, ending in a value about 10.
• Taper ratio The taper ratio affects the induced drag, the structural weight and the facility to build the wing and, consequently, the cost. A low taper ratio (λ = 0.3 - 0.5) involves low in- duced drag and low structural weight, but it is easier to build an untapered wing (λ = 1).
The equilibrium point between a good performance (low induced drag and weight) and an affordable cost is a value of 0.5.
• Sweep angle The sweep angle apparently has no direct advantages and, in fact, it has an adverse ef- fect, which results in a reduction of lift (lower CLα and a lower CLmax), an increment of induced drag (higher CDi) and also a higher structural weight. Swept wings are used to increase the drag divergence Mach number, but in this case, it is not necessary, because our plane flights in low Mach numbers. Hence, the sweep angle of the horizontal tail is only about 5º.
• Airfoil selection
Both horizontal and vertical tails have a symmetric airfoil, because the movable sur- faces can deflect on both sides of the undeflected position.
For low Mach number flights, one of the most common airfoils for the horizontal tail is the NACA 0008, which provides the necessary aerodynamic forces.
RA 30 G06-AlOn LSA 3 seats | Project report • Incidence angle The incidence angle of the horizontal tail is the angle between the fuselage reference axis and the reference chord of the horizontal tail. The tail incidence is such that during cruise flight the lift required from the tail to make the pitching moment zero is produced without elevator deflection.
To determine the incidence angle is necessary to know the angle of attack of the plane in cruise flight and the downwash of the tail. This last value is difficult to measure properly. Hence, the tail is built with zero incidence angle and the performance of the airplane will be controlled with the corresponding control surface.
1.2.6.1.2 Vertical tail • Aspect ratio
The influence of the aspect ratio in vertical tails is the same as in horizontal tails, so the points mentioned previously are applicable. Besides, there are additional influences: increment of the height of vertical tail and, consequently, the plane is higher; increment of moment of inertia about the longitudinal axis, which involves lower lateral control; and better directional control due to the increment of the moment arm.
However, the main aspect to consider in a T-tail is the risk of deep stall, so it is important to place the horizontal tail above the wing wake when flying at high angles of attack. This is the criterion applied for our plane to determine the height of the vertical tail.
• Taper ratio
To determine the taper ratio of the vertical tail, the same criterion as in horizontal tail has been applied. As a result, the value selected is 0.67.
• Sweep angle
The adverse effects of sweep explained for the horizontal tail are also applicable for vertical tail. In addition, a swept vertical tail has a higher moment arm.
Considering these effects, a low sweep angle is commonly applied, in this case it is about 7º.
• Airfoil selection
As for the horizontal tail, a symmetric airfoil is commonly used in vertical tails. In this case, the NACA 0008 is selected again, due to its aerodynamic characteristics.
1.3 Fuselage
The main problem about the fuselage configuration is the unusual number of passangers and also the maximum take-off weight. A configuration of three people does not allow the typical LSA or glider two seat configuration. So the casuistry is the next:
RA 31 G06-AlOn LSA 3 seats | Project report • 3 seat configuration
• Relative position of the wing and the tail
• Maximum efficiency
• Costs
• Weight and structural facilities
Analyzing the options, the two different dispositions were studied in order to define which was the best. Fig. 1.16 shows the 2 different configurations.
(a) One plus two (b) In line
Figure 1.16: Seats configuration
Firstly an analysis has been done in order to see which configuration has better behaviour. The main performance graphics are going the be analyzed next.
• Efficiency It is expected that the three seats in line configuration has a better aerodynamic per- formance because of it’s geometry. The results obtained can be seen in the following graphic, which show both efficiency curves. As said before, Fig. 1.17 shows both graph- ics are almost the same. There is not a big difference between the maximum values. So, in this case, the efficiency is not a decisive point.
• Momentum coefficient As can be seen in Fig. 1.18 the difference is not remarkable enough to decide which con- figuration is better. Although the one plus two model has the Cm = 0 best situated than the in line one, the maximum efficiency is more centered in the wider configuration. But in both cases the plane shows the same maneuverability, as both lines are parallel.
RA 32 G06-AlOn LSA 3 seats | Project report
Figure 1.17: Efficiency in both configuration
Figure 1.18: Cm in both configurations
Another important aspect to consider is the union between the wing and the fuselage. Be- cause of its wider configuration, the one plus two design offers a better union and allows a better fit performance.
In order to evaluate all the variables and choose the one with the best performance, the Un- weighted Average Method (UA) and the Ordered weighted Average Method (OWA) have been applied.
1.3.1 Fuselage selection criteria For the application of this method, it is necessary to define the main variables to be evaluated for each configuration. The comparison criteria will be based on:
• Efficiency: it measures ratio between the functionality provided by the fuselage and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.
• Weight: it considers the complexity of the structure required in each configuration and, consequently, the weight of the fuselage. A higher value in this criterion indicates a lighter structure.
• Cost: it is related with the difficulty to build each fuselage, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.
RA 33 G06-AlOn LSA 3 seats | Project report • Wing position: it defines how big is the area where the union between the wing and the fuselage is going to be make. A Higher value in this criterion means a higher area, thus a better union.
1.3.1.1 Unweigthed Average Method The scale applied goes from 1 to 5, where a score of 5 in any variable shows the best possible value and a score of 1 shows the worst. This method does not take into account the weight of the criteria evaluated in relation to the other configuration. The obtained results are shown in Table 1.12:
Unweigthed Average Method Criteria Seats in line One plus two Efficiency 5 4 Weight 4 5 Cost 3 4 Wing position 4 5 SUM 16 18 Max 20 20 UA 0,8 0,9
Table 1.12: Unweighted Average Method application(fuselage)
As can be seen in Table 1.12, the best option in this case is the One plus two configuration, which offers better wing position, lower weight and better structural performance, although the efficiency is worse.
1.3.1.2 Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a specific weight for each criteria, which indicates the importance or weight that will have in the final result. The results obtained can be seen in the following table:
Ordered Weighted Average Method Seats in line One plus two Weight Criteria P PxG P PXG G Efficiency 5 10 4 8 2 Weight 4 16 5 20 4 Cost 3 3 4 4 1 Wing position 4 12 5 15 3 SUM(PxG) 41 47 Max(PxG) 50 50 OWA 0,82 0,94
Table 1.13: Ordered Weighted Average Method application(fuselage)
As can be seen in Table 1.13, the best option is again the one plus two configuration. Hence, the final fuselage configuration will be the one plus two, which is shown in Fig. 1.16(a).
1.4 Control Surfaces
The control surfaces allow the pilot to govern the plane, this means that angles of yaw, pitch and roll can be controlled. The three control surfaces are:
• Elevator: its task is to control the pitch.
RA 34 G06-AlOn LSA 3 seats | Project report • Ailerons: their task is to control the roll, however, it induces a moment that also affects the yaw due to the difference of drag produced by the difference of lift that both wings have. This effect is called adverse yaw.
• Rudder: its task is to control the yaw, however, it induces a moment that also affects the roll because the rudder is higher than the center of mass, the rotation point. This effect is called adverse roll.
In order to size the different control surfaces a similarity study has been done. The plane chosen was the glider Taurus from Pipistrel because of its similar characteristics.
1.4.1 Elevator
From the similarity study it can bee seen that the elevator should be a third of the horizontal tail’s surface. This means that Alpha-One should have a 0.5m2 elevator. In order to make it easier to construct, the elevator will have the same span that the horizontal empennage, 4m; so the chord will be 0.125m to fulfill the two conditions. The hinge moment was calculated assuming that the elevator was a flap of 0,33 chord with 15º of deflection:
Figure 1.19: Elevator Airfoil
Assuming this, XFLR-5 is capable to compute the non-dimensional hinge moment, which is the following:
Figure 1.20: Hmom Elevator
RA 35 G06-AlOn LSA 3 seats | Project report
To compute the maximum hinge moment the vne should be used and a Hmom of 0.025:
2 Moment = 0.5 · ρ · vNE · Se · ce · Hmom (1.1)
Moment = 0.5 · 1.225 · 82.32 · 0, 5 · 0, 125 · 0.025 = 6.48Nm (1.2) In order to determine the influence of the elevator, the lift coefficient at its maximum de- flection angle has been obtained for different values of angle of attack, which can be seen in Fig. 1.21. For further structural calculations, it is necessary to obtain the maximum lift coefficient provided by the elevator, which is approximately 1.2.
Figure 1.21: Cl versus α for the elevator at its maximum deflection angle
1.4.2 Ailerons In the Taurus from Pipistrel, each aileron has a surface of one tenth of the wing. In the de- signed plane this equals to 0.6m2. The span of the aileron will be 4m with 0,15m chord and will be located at the end of the wing.
Figure 1.22: Aileron Airfoil
To obtain the hinge moment the same method will be used. The next figure shows the aileron was a flap of 0.25 chord with 15º of deflection:
RA 36 G06-AlOn LSA 3 seats | Project report
Figure 1.23: Hmom Ailerons
To compute the maximum hinge moment for each aileron, the vne should be used and a Hmom of 0,010: 2 Moment = 0.5 · ρ · vNE · Sa · ca · Hmom (1.3)
Moment = 0.5 · 1.225 · 82.32 · 0.6 · 0.15 · 0.010 = 3.73Nm (1.4)
1.4.3 Rudder In the similar plane, the rudder has a surface of one half of the fin. In the designed plane this equals to 0.4m2. The span of the rudder will be the whole fin, 1.2m, and 0.4m chord.
Figure 1.24: Rudder Airfoil
In order to calculate the hinge moment, the same method will be used. Here we can see the airfoil with a flap of 0.5 chord with 15º of deflection:
RA 37 G06-AlOn LSA 3 seats | Project report
Figure 1.25: Hmom Rudder
To compute the maximum hinge moment the vne should be used and a Hmom of 0,065:
2 Moment = 0.5 · ρ · vNE · Sr · cr · Hmom (1.5)
Moment = 0.5 · 1.225 · 82.32 · 0.4 · 0, 4 · 0.065 = 43.15Nm (1.6) In addition, to determine the influence of the rudder to the aerodynamic forces, the lift coeffi- cient at a maximum angle of deflection as a function of the angle of attack has been obtained using XFLR5. The results can be shown in Fig. 1.26. From this graphic, the maximum lift coefficient can be obtained, which is approximately 1.2.
Figure 1.26: Cl versus α fortherudderatitsmaximumdeflectionangle
1.4.4 Parasite Drag In order to get a more realistic performance of the aircraft, the drag coefficients of the landing gear have been estimated. For the estimation of the wheels CD it has been used the following method:
• Define the reference distances.
RA 38 G06-AlOn LSA 3 seats | Project report
• Get the CD value from a experimental graphics.
• Define the calculated CD on XFLR5 to improve the veracity of the simulations.
And for the tube that connects the wheels it has been used the next following system:
• Define the reference distances.
• Create an equivalent wing using a symmetric air foil.
• Make a simulation in order to get equivalent CD.
1.4.4.1 Wheels parasite drag The reference distance used for the wheel estimation have been: the following ones:
• Radius = 0.4 m
• Width = 0.1 m
CD are not as usual as it might seem, there are not to much forms analyzed and, a disk through the longitudinal face is one of the figures which have not been analyzed.
The figures that might be an approximation to the disk in volume or in some referent distance are:
• Cylinder
• Sphere
• Bullet
Because of its geometry, the bullet its the most realistic way to simulate the wheel, because it has an smooth curve at the beginning and, but there is not as much rounded as the sphere.
The following figure shows the CD distribution through the Re. The Re used has been the one for cruise flight conditions, which is:
Re = 1.7M (1.7)
Thus, using the following figure, the CD for each wheel have been estimated.
Figure 1.27: CD experimental data ; Source: researchgate.net
l Using the bullet body and a d ratio of 1, to be conservative. The extra drag for each wheel is :
−1 CD = 0.85wheel (1.8)
RA 39 G06-AlOn LSA 3 seats | Project report 1.4.4.2 Tube parasite drag As it has already been said, the tube has been estimated as a wing, a rectangular one. It can be noticed there is a huge difference between a wing and a tube, but taking into consideration that the tube is nor a cylinder neither a wing, it has been decided to estimates as a rectangular wing using a huge symmetric airfoil.
The geometry used to define the wing tube is:
• lcenter = 2m
• lside = 0.5m (each diagonal longitude).
Using the listed geometry the extra CD obtained is:
CD0 = 0.05 (1.9)
1.4.4.3 Drag correction Applying the drag correction to the final plane configuration on XFLR5, have been obtained the following results:
Figure 1.28: Comparison between both behaviours
As it can be on the Fig. 1.28 the maximum efficiency decreases due to the extra parasite drag caused by the landing gear. It was expected result, even though the addition of the landing gear drag, the efficiency performance of the airplane is still really good, with a maximum efficiency of approximately 35.
1.5 Final plane configuration
1.5.1 Design parameters Once everything has been correctly calculated, the main parameters have been summarized in the following tables:
RA 40 G06-AlOn LSA 3 seats | Project report Wing span (m) 16 Wing area (m2) 12,424 Wing load (kg/m2) 48,295 Croot (m) 1 Ctip (m) 0,6 Wing span (m) 4 Aspect ratio 20,606 Wing area (m2) 1,5 Taper Ratio 10 Wing span (m) 1,2 Croot 0,5 2 Efficiency 35.42 Wing area (m ) 0,8 Ctip 0,25 Tilt angle (º) 2 Croot 0,75 Aspect ratio 10,67 v(m/s) 43,6 Ctip 0,6 Taper ratio 2
Table 1.14: Final wing configuration Table 1.15: Final fin configu-Table 1.16: Final elevator con- data ration data figuration data
Fin Elevator Aileron Sw 0,8 1,5 6,212 Croot 0,375 0,5 1 S(control surface) 0,4 0,5 0,6212 C(control surface) 0,375 0,15 0.25
Table 1.17: Control surfaces main sizes
1.5.2 Range study Also in order to estimate the different possible ranges, a Weight-range diagram has been ob- tained using the Breguet’s equation.The code developed is attached in the annexes.
The next flight conditions are considered in order to apply the Breguet equation in cruise flight conditions: L = W (1.10) T = D (1.11) Both equation result into the next relationship:
C T = W D (1.12) CL The range can be calculated as:
Z x f Z t f R = dx = dt · v (1.13) xi ti It is also known the relationship between time an weight, which relates the fuel intake with the mass flow:
dW c · dt = − (1.14) g Where: Tv c = c · P = c (1.15) j j η Also applying (1.10) it is obtained the next result: s 2W v = (1.16) CL · Sw · ρ
RA 41 G06-AlOn LSA 3 seats | Project report Now, applying (1.15),(1.14), (1.16) and (1.12) into (1.13) the next result it is obtained: s η 2 Z Wf dW R = − · E 1/2 (1.17) cjg CL · Sw · ρ Wi W Integrating the last equation it is obtain. s 2η 2 1/2 1/2 R = · E · (Wi − Wf ) (1.18) cjg CL · Sw · ρ
Where CL is the optimum one to cruise mode. Applying (1.17) the W-R it is obtained:
Figure 1.29: Weight-Range diagram
The code designed to obtain the Fig. 1.29 can be found in the annexes of the report.
RA 42 G06-AlOn LSA 3 seats | Project report 1.6 Final Plane Analysis
1.6.1 Efficiency
The efficiency of the final plane is shown in the following graph:
Figure 1.30: Efficiency of the final configuration
As it can bee seen in Fig. 1.30 the maximum efficiency is 34,5 at 1,5º of angle of attack.
1.6.2 Static Stability
The static stability is the immediate response of the plane in front of a perturbation. A static plane is the one that in front of a sudden increase of angle of attack, a restorer moment ap- pears that decreases the perturbation. This means that the variation of the pitching moment coefficient as function of alpha should have a negative slope. Furthermore, the alpha with zero pitching moment coefficient should be roughly the same that the angle of attack of max- imum efficiency. This condition would make the plane to have the angle of attack with the best glide ratio without the pilot’s help.
The following figure shows the final plane pitching moment coefficient, Cm, as function of alpha:
Figure 1.31: Static stability of the final configuration
RA 43 G06-AlOn LSA 3 seats | Project report The slope of the pitching moment coefficient is negative which will make the plane statically stable. The alpha with no Cm is 0,7º, the alpha of maximum efficiency is 1,5º as it is shown in Fig. 1.30. It’s important to notice that the position of the center of gravity has a huge impact in stability analysis. In order to achieve the behaviour described in Fig. 1.31 the center of gravity should be at 0,225m behind the leading edge of the wing at the root.
1.6.3 Dynamic Stability The dynamic stability consists in the behaviour of the plane in front of a perturbation that changes its flight conditions. The desired behaviour is the one that tries to counter the effect of the perturbation and damp the amplitude of it.
In order to evaluate the dynamic stability of the plane a modal analysis has been done. If the Modes have a damped evolution in time, the plane is dynamically stable. The three more important modes are the followings:
• Longitudinal Mode 1
• Lateral Mode 1
• Lateral Mode 2
1.6.3.1 Longitudinal Mode 1 This mode is related to the pitch angle. The next figure shows the angular velocity related to this mode:
Figure 1.32: Longitudinal Mode 1
As it can bee seen, the angular velocity decreases in time so it’s a stable second order response to the first longitudinal mode. It takes about 2s to dissipate the perturbation.
1.6.3.2 Lateral Mode 1 This mode is related to the roll angle. The next figure shows the angular velocity related to this mode:
RA 44 G06-AlOn LSA 3 seats | Project report
Figure 1.33: Lateral Mode 1
As it shows the figure, the angular velocity decreases rapidly in time so it’s a very stable response. This is a first order behaviour with a time constant of τ = 0.1s. This means that in 0.4s the perturbation is almost gone.
1.6.3.3 Lateral Mode 2 This mode is related to the yaw angle. The next figure shows the angular velocity related to this mode:
Figure 1.34: Lateral Mode 2
The figure shows that is a stable mode but is the slowest one. It takes about 8s to dissipate the perturbation and is more oscillating than the other two studied modes.
1.7 Flight envelope
In order to define the flight envelope where the LSA is going to flight, an airspeed-gust en- velope has been done at sea level. To develop the diagram, the rules and requirements set by the governmental authorities have been taken into account.
RA 45 G06-AlOn LSA 3 seats | Project report The [4] has been followed in order to make sure to meet all the requirements and to calculate the necessary velocities. Following the structural requirement established by the ASTM F225, the load factors needed are:
• n1 = 4 : Positive maneuvering load factor
• n2 = −2. Negative maneuvering load factor
• nF1 = 2 :Positive manoeuvring load factor flaps extended
Firstly, the vs is needed in both cases, with and without flaps. Those velocities have been determined by simulations using XFLR5, which already gives the stall velocity for a Cl. In this case the maximum cl is required. The following velocities will allow computing the other ones in order to be able to create the airspeed diagram:
vs = 45.40kts; (1.19)
vs0 = 43.13kts < 44kts −→ EASA requirement; (1.20) Now the needed velocities can be calculated.
Design manoeuvring speed va: √ vA = vs n1 = 90.8kts (1.21)
Flaps maximum operating speed vF:
Should be bigger than 1.4ss and 2vs0. Doing this operation the result is:
vF = 85kts; (1.22)
Design cruising speed vc:
Needs to be bigger than: r wMTOW vCmin = 4.77 = 103kts (1.23) Sw And smaller than: 1.4VH = 121.21kts (1.24)
So, approximately vC can be defined as:
vC = 110kts (1.25)
And finallly 4.7 Design dive speed VD:
vD > 1.4vCmin = 145.34 ≈ 160kts because needs to be higher (1.26) Using the same methodology for the flap and for the negative load factor the resultant air- speed diagram is the following:
RA 46 G06-AlOn LSA 3 seats | Project report
Figure 1.35: Airspeed envelope diagram
The next step is to calculate the gust lines for 7.5 and 15 m/s for both vC and vD.
In order to get the load factor increment due to an instant gust the ASTM F2245 has been used. Using the aircraft data and the graphics that can be found in [4], the increments of the load factor for the different cases can be determined.
Using the increments previously calculated from the ASTM F2245, the following diagram has been developed:
RA 47 G06-AlOn LSA 3 seats | Project report Gust-Airspeed envelope diagram Figure 1.36:
RA 48 Chapter 2
Structures
In this chapter, design and analysis of the different structures that conform the aircraft will be presented . Main consideration that will be taken into account is weight.Since LSAs are thought to fly with a maximum of 2 passengers, adding a third one forces the structure to be ligther than convetional LSAs. For this reason, the first that will be done is a materials study in order to find the most suitable options. Another important item in the design is the resistance of the parts, whose analysis will be done during the second part of this chapter, where the design and the analysis if the parts will be presented.
2.1 Materials
As previously introduced, the first thing that needs to be done is an analysis of the materials. In order to do this, it will be presented a choice of materials for: the main structure (fuselage and wings), the skin, the landing gear and the transparent surfaces.
2.1.1 Fuselage and wings internal structure
For both ribs and beams, a light but strong material must be chosen. The most used materials for these applications are aluminium-based composites and, among all of the different pos- sible composites, two of them have been compared: Regular aluminium alloy and Central Reinforced Aluminium (CentrAl).
There are a lot of great advantages when using Aluminium alloys: they are an exception- ally light material with a relatively high strength and with an acceptable electric and thermal conductivity. They are neither magnetic nor toxic while at the same time they reflect light (it means it allows to have a lower heat accumulation), they are water-resistant and they ease the manufacturing process due to their ductility and malleability.
They have really important environmental advantages as well. Current aluminium alloys are completely recyclable and as they reduce the aircraft weight, a reduction of fuel consumption is fulfilled due to the reduction of the needed thrust to propel itself.
The main properties of the chosen aluminium alloy [5] are presented in Table 2.1.
RA 49 G06-AlOn LSA 3 seats | Project report Property Value Density 2.698,4 kg/m3 Fusion point 933,47 K Ce 900 J/K·kg Electric conductivity 37,7.106 S/m Thermal conductivity 237 W/K.m Young modulus 66,6 Gpa Traction resistance 230-570 MPa Mechanical resistance 690 MPa Elastic limit 215-505 MPa Elongation 10-25
Table 2.1: Global properties of aluminium
The second aluminium-based material is called Central Reinforced Aluminium which is a composite consisting of several layers of aluminium and glass fibers glued with epoxy-based adhesives. Between aluminium and the glass fiber composite are layers of a proprietary resin- rich material called “BondPreg” by the developers. The main developer and manufacturer of CentrAl is Alcoa and the structure of the material they distribute is presented in Fig. 2.1.
Figure 2.1: CentrAl configuration
The different materials used to manufacture the CentrAl [6] have diverse properties and when they are glued together, some new values are achieved. This data is presented in Ta- ble 2.2 and Table 2.3.
Unidirectional Lamina Property Aluminum 2024-T3 Unidirectional BondPreg® (S2-Glass Fiber and FM94K Adhesive) UTS (L (material 441,26 MPa 2.213,22 MPa 1.103,16 MPa rolling direction)) UTS (LT (perpendicular to 434,37 MPa 22,06 MPa 21,99 MPa material rolling direction)) Y (L) 324,05 MPa 355,76 MPa 177,88 MPa Y (LT) 289,58 MPa 51,71 MPa 29,03 MPa E (L) 72,39 GPa 53 GPa 27,61 GPa E (LT) - 5,31 GPa 3,12 MPa G12 27,57 GPa 2 GPa 1,17 GPa nu 12 0,33 0,27 0,30 nu 21 - 0,027 0,034 alpha 12 - 0,00000161 F−1 0,00000337 F−1 alpha 21 - 0,0000224 F−1 0,00003904 F−1
Table 2.2: Properties of components of CentrAl
RA 50 G06-AlOn LSA 3 seats | Project report Property Value Density 2.591 kg/m3 Young Modulus 65,2 GPa Poisson 0,33 Elastic Limit 812,2 MPa Y (Yield Strength) 383,8 MPa Deformation failure 0,04761
Table 2.3: Properties of CentrAl
CentrAl provides nearly 25 percent more tensile strength than high-strength aluminum al- loys, is extremely resistant to metal fatigue and is highly damage-tolerant. Taking all those properties into account it is possible to consider that CentrAl, without further analysis, seems a better material for this purpose. The main reasons are:
1. 25 percent more tensile strength than high-strength aluminum alloys
2. Extremely resistant to metal fatigue
3. Highly damage-tolerant
4. So light that a transport-aircraft wing made from a combination of CentrAl and alu- minum – which is better than CentrAl at resisting the compression strains on surfaces such as upper wing-skins – would not only be much stronger than a wing completely made of aluminium, but also could be 20 percent lighter.
So the main conclusion would be that the best material to build the structures is the CentrAl, despite it would have to be combined with aluminium alloys in the structures that have to support a higher compression strain. Nevertheless, it must be mentioned that other possibilities must not be directly discarded because even though CentrAl presents excellent properties its density is not specially low and the main limitation for this aircraft is the weight. Due to this problem fiber and honeycomb composites will be also studied for the wings composition. These materials are also light [7] but enough resistant to some parts of the aircraft due to the combination.
2.1.2 Skin of the aircraft
In order to build a light and effective aircraft skin the materials considered will be different types of carbon fibers. There are a lot of different types of carbon fiber depending on its composition (fiber type, fiber amount, fiber orientation, etc). Some of them are specifically used in aerospace applications, but all of them share certain interesting properties such as high longitudinal tensile strengths, low density, low coefficient of thermal expansion and low thermal and electric conductivity. A properties comparison between different carbon fibers used in aerospace applications is shown in Table 2.4, Table 2.5 and Table 2.6.
RA 51 G06-AlOn LSA 3 seats | Project report T300 (2,7%Epoxy /93% Carbon) Property Value Density 1.760 kg/m3 Tensile Modulus 140 GPa Tensile Strength 1.820 MPa -0,41 alpha Coefficient of thermal expansion ·10-6 ºC−1 Specific Heat 0,777 J/g ºC Thermal Conductivity 0,105 J/cm s ºC Electric Resistivity 1,7 · 10−3 Ω · cm
Table 2.4: Properties of T300
T400H (1,6%Epoxy /94% Carbon) Property Value Density 1.800 kg/m3 Tensile Modulus 145 GPa Tensile Strength 2.250 MPa -0,45 alpha Coefficient of thermal expansion ·10-6 ºC−1 Specific Heat 0,18 J/g ºC Thermal Conductivity 0,0252 J/cm s ºC Electric Resistivity 1,6 · 10−3 Ω · cm
Table 2.5: Properties of T400H
T1000G (0,7%Epoxy /95% Carbon) Property Value Density 1.800 kg/m3 Tensile Modulus 145 GPa Tensile Strength 2.250 MPa -0,45 alpha Coefficient of thermal expansion ·10-6 ºC−1 Specific Heat 0,18 J/g ºC Thermal Conductivity 0,0252 J/cm s ºC Electric Resistivity 1,6 · 10−3 Ω · cm
Table 2.6: Properties of T1000G
The main conclusion would be that Carbon Fiber is such an interesting material to be used in the aircraft skin. The fiber configuration chosen to be used in the aircraft is the T300 because it is able to hold enough tensile strength and has a very low density so the weight of the plane would not specially increased.
2.1.3 Landing gear To choose the materials for the landing gear the main requirements are the weight and the resistance of the materials. Because of their tensile strength the materials that will be evaluated are titanium alloys and also various types of steel.
RA 52 G06-AlOn LSA 3 seats | Project report Initially, titanium [8] was selected for these reasons: 1. Its tensile strength
2. Its low density, which it is considerably lower that any of the steels. But the main problem it presents is its price. Unlike, the steels present a higher density, have a similar tensile strength and its cost are considerably lower. Based on this considerations, when the structure is built, an analysis will be performed to see which material matches better the requirements.
2.1.4 Transparent surfaces 2.1.4.1 Windshield Aircraft windshields [9] may be purchased either from the original aircraft manufacturer or from any of several FAA-PMA sources. These windshields are formed to the exact shape required, but can be slightly larger than necessary so they may be trimmed to the exact size.
2.1.4.2 Windows Are mainly made with two different thermoplastics [9]: 1. Acrylic plastics. They are the current substitute to Cellulose acetates and they are com- monly known with the name of Plexiglas. Two main kinds: MIL-P-5425 for military regular specifications and MIL-P-8184 for military craze-resistant.
2. Cellulose acetate plastics. It was vastly used in the past but was dimensionally unstable and turned yellow after being installed for a long time. It is important to emphasize that transparent thermoplastic sheets soften and deform when they are heated, they must be used and stored where the temperature will never become excessive. Store them in a cool, dry location away from heating coils, radiators, or steam pipes.
2.1.5 Additional material considerations In this subsection there will be added some things that have to be taken into consideration for the manufacturing and the conservation of the aircraft materials [9].
2.1.5.1 Welding Before trying to weld any aluminium alloy it must be tested to determine if it is able to heat treatment or not. In many cases materials are marked to distinguish between heat and non- heat treatable alloys but if in any case it is not marked, there are some ways to identify the materials. First way would be immersing a sample of the material in a 10 % solution of caustic soda. Materials with a high copper content will turn black and would make an evidence of heat tractability. However, if it does not turn black, it is not an evidence of non-heat-treatable so these materials could be tested by chemical or spectre-analysis.
2.1.5.2 Corrosion Prevention There are 3 levels to categorize the severity of corrosion damage (IATA): 1. The damage does not require structural reinforcement or replacement of parts. Also corrosion occurring between inspections that exceeds allowable limits but is local and can be attributed to an event not typical of operator usage of other aircraft of the same fleet.
RA 53 G06-AlOn LSA 3 seats | Project report 2. Corrosion occurring between inspections that requires a single rework/blend -out, which exceeds allowable limits, requiring a repair/reinforcement or complete/partial replace- ment of applicable structure.
3. Corrosion found during first inspection, which is determined to be an urgent airwor- thiness concern requiring expeditious action.
Corrosion prevention [10] involves cleaning the surface and providing a surface finish through layers of coasting. For the surfaces of aluminium alloys, the coating of a corrosion-inhibiting primer is the first coat. Primer is a base protecting used on the outer surface of a substrate before painting. They shield the area and ensure better grid of color to the outer surface. Primers increase the color strength and prevent attack on the substrate by air, water or other elements. Urethane primer is formulated as a high-performance corrosion inhibitor for all types of aircraft. Urethane consists of 2 parts: the base and certain binders with catalysts. Urethane advantages:
• Good inter-coat adhesion
• High durability
• Good chemical resistance
• Abrasion resistant
• High gloss finish
There are some other primers such as polyurethane or epoxy primers. Corrosion-Resistant Alloy can also be used. They are alloys consisting of metals such as chrome, stainless steel, cobalt, nickel, iron, titanium and molybdenum. When combined, such metals can promote corrosion resistance and can offer reliable protection from corrosion, eradicating the need for maintenance and repair. This will make a compulsory point to check if the chosen materials are corrosion-resistant or if they will need an additional concrete product to prevent material from corrosion.
2.2 Landing Gear
2.2.1 Analysis of the landing gear regulations for an LSA The regulations on the landing gear for LSA are not so specific [11], so, basically, it will be the most convenient. The only points that must be considered are the following ones:
1. Depending on the type of the LSA we finally select our landing gear will be approxi- mately conditioned. If it is finally chosen a conventional LSA, the landing gear must be obligatorily fixed. Conversely, if it is decided a glider LSA the landing gear can be fixed or retractile. In order to avoid possible conflicts in a further point of the project it will be chosen a fixed landing gear, which would fit in any case.
2. Depending on the landing gear chosen, it will need structural considerations. If it affects to the wing, the ground loads must be justified in the wing structure.
3. The manufacturer must certificate the landing gear, with its corresponding wheels, brakes and tires, and install all the landing gear system components.
Based on these previous premises, we can conclude that with its required by the norms, the best option for us is to choose an existing fixed landing gear that does not affect to the wings of our plane.
RA 54 G06-AlOn LSA 3 seats | Project report 2.2.1.1 Landing gear options
The evaluated existing landing gears are the following ones:
2.2.1.2 Atec 322 Faeta
The landing gear is a fixed tricycle [12] undercarriage with a steerable nose wheel. The main gear is constructed as a pair of leaf springs of composites. The front leg is made of composites and metal tube suspended with rubber spring. The main gear is a pair of composite springs. Electronic main wheels size is 350x120 mm, front wheel size is 300x100 mm. The main wheels are equipped with hydraulic disc brakes. Fairings are installed on all wheels. Landing Gear (tricycle with front wheel) specifications:
1. Wheel spacing: 6,234 ft / 1,900 m
2. Wheel base: 4,742 ft / 1,445m
3. Front tire: 400x400 / 102 x 102 mm
4. Main tire dimensions: 600x400 / 152 x 102 mm
5. Tire pressure: 23,2 psi / 160 KPa
2.2.1.3 TL-2000 Sting S4
The landing gear is convention a fixed [13], tricycle type with a steerable nose gear and two main landing gears. Hydraulically-actuated brakes are attached on each main landing gear wheel.
2.2.1.4 TL-3000 Sirius
TL-3000 Sirius [14] is fitted with a fixed tricycle undercarriage with main wheels fuselage mounted on composite cantilever spring legs; all of them have brakes. The main gear legs made of carbon and glass fibre composite, which dampens impact and makes the aircraft forbearing in the event of hard landings, are attached to the undercarriage bulkhead trestle located under the pilot seats. The gear is equipped with a shock absorber and attached to the firewall; all wheels are almost completely enclosed in spats. There are two variants of the undercarriage which differ from undercarriage wheels depending on the version of the airplane. First variant has its nose gear and main gear wheels with dimension 400 x 100 mm. The nose gear is non-steerable, and it is equipped with shimmy dumper.
Figure 2.2: First variant of the TL-3000 Sirius’ landing gear
RA 55 G06-AlOn LSA 3 seats | Project report
Figure 2.3: Second variant of the TL-3000 Sirius’ landing gear
Both undercarriage options have different type of laminate wheel covers. It is not possible to combine these two undercarriage variants. First, it is necessary to recognize the undercarriage version of your aircraft to follow all instructions related to your type of undercarriage that are described in this article. Besides, the company has also developed a float installation for the TL-3000 to allow water operations.
2.2.1.5 Pipistrel Taurus M Double retractable main landing gear [15]. However, it is retractable, and it has been decided not to use a retractable landing gear for the project.
2.2.1.6 Alexander Schleicher ASG 29 E There’s no specific information about its landing gear [16]. Nevertheless, regarding to its glider nature it is assumed a one wheel retractable and a rear wheel non-retractable; as it is observed in any of its pictures.
Figure 2.4: ASG 29 E flying
2.2.1.7 ONE Aircraft Fixed tricycle landing gear and a single engine in tractor configuration.
2.2.1.8 Sling 4 The landing gear [17], which is made from composites, is a tricycle landing gear with a steer- able nose wheel. The main landing gear uses a single continuous composite spring section. Landing gear specifications:
1. Wheel track: 1,95 m (6,4 ft).
2. Wheel base: 1,68 m (5,51 ft).
3. Brakes: Hydraulic.
4. Main gear tyres: 15x6.00-6, 6-ply
RA 56 G06-AlOn LSA 3 seats | Project report 5. (2,5 bar (36,26 psi) pressure).
6. Nose gear tyres: 5.00-5, 6-ply
7. (1,8 bar (26,11 psi) pressure).
2.2.2 Landing gear calculations
2.2.2.1 Static analysis This first analysis will be determining for the landing gear because it will be used to choose the material for the landing gear. The first step to proceed with this section has been modeling the whole landing gear with an appropriate software. The results will be presented in the report attachments but for the reader to have a basic idea of what is going to be analyzed, the main parts are shown in Fig. 2.5 and Fig. 2.20.
Figure 2.5: Front leg of the landing gear.
Figure 2.6: Half of the rear landing gear.
Once the models are obtained, the static analysis are performed with the same software. All the analysis were performed over the models meshed with tetrahedral elements and with a quadratic quality. The meshes are created with four Jacobean points and the elements have a size of 6mm.
RA 57 G06-AlOn LSA 3 seats | Project report In order to determinate which was the most appropriate design for the landing gear it has been done a static analysis of the front landing gear and a posterior analysis of the results to make a decision. The front landing gear has two main designs, based on the same structure but one with a solid tube and the other with a void tube to see if the weight reduction compensates the reduction in the resistance. Therefore, the same static analysis has been done over the same front landing gear in six different configurations:
1. Solid tube in a titanium alloy.
2. Solid tube in a steel 4340.
3. Solid tube in a steel 4130.
4. Void tube in a titanium alloy.
5. Void tube in a steel 4340.
6. Void tube in a steel 4130.
Even though in the materials sections titanium is chosen as the best option, different materi- als are evaluated because of its price.
This analysis is performed with a finite element software that builds structured meshes au- tomatically with customized parameters. The parameters chosen for this analysis are: 4 Ja- cobean Points, an element quality of quadratic order, tetrahedral elements and an element size of 10.73mm.
After performing the same analysis, which has consisted in the application of a stress over the wheel axis equal to the weight of the whole aircraft ( 600kg ) - even thought it must be taken into account that in the end this part of the landing gear must only bear with between 8 to 15 per cent of the aircraft gross weight) the results have been those shown in Table 2.7.
Weight (g) Max. Deformation (mm) Material cost ($) Titanium solid 24.364,49 0,1360 901,48 (37$/kg approx) Steel 4340 solid 39.680,76 0.0814 119,04 (3000$/tn approx) Steel 4130 solid 39.680,76 0,0816 79,36(2000/tn approx) Titanium void 8.646,90 0,2837 319,94$ Steel 4340 void 14.082,60 0,1441 42,25$ Steel 4130 void 14.082,60 0,1446 28,16$
Table 2.7: Results of the static analysis of the different alternatives of the landing gear.
Since is not possible to make a decision with only this data, an OWA decision method will be applied to determine which material and structure is more convenient for the case.Results are shown in Table 2.8.
Weight Titanium solid Steel 4340 solid Steel 4130 solid Titanium void Steel 4340 void Steel 4130 void p p*w p p*w p p*w p p*w p p*w p p*w Weight (g) 6 2,97 17,82 1 6 1 6 5 30 4,29 25,74 4,29 25,74 Max. Deformation (mm) 2 3,91 7,82 5 10 4,99 9,98 1 2 3,75 7,5 3,74 7,48 Material cost ($) 1 1 1 4,58 4,58 4,76 4,76 3,66 3,66 4,93 4,93 5 5 Total sum of p*g 45 26,64 20,58 20,74 35,66 38,17 38,22 OWA score 1 0,592 0,457 0,461 0,792 0,848 0,849
Table 2.8: Results of the application of OWA method
RA 58 G06-AlOn LSA 3 seats | Project report As one can see, the model that presents the best result is the void bar model in steel 4130, therefore, this will be the used material.
Hence, the following analysis in both, the front and the rear landing gear will be performed with the parameters from Table 2.9, corresponding to a steel 4130.
Young Modulus 205 GPa Poisson coefficient 0,285 Shear Modulus 80 GPa Density 7.850 kg/m3 Traction limit 731 MPa Elastic limit 460 MPa Thermal conductivity 42,7 W/(m·K) Specific heat 477 J/(kg·K)
Table 2.9: Properties of the steel 4130 normalized at 870 ºC.
And the result of these analysis are the following ones:
First, for the front landing gear, with the parameters mentioned above, results are shown in Fig. 2.7 and Fig. 2.8.
Figure 2.7: Stress analysis of the front landing gear.
RA 59 G06-AlOn LSA 3 seats | Project report
Figure 2.8: Displacements of the front landing gear.
And must be added that it was also performed a buckling test which granted that there would be no buckling on the structure.
And then for the rear landing gear, the analysis is performed with a solid mesh. This mesh is created with four Jacobean points and tetrahedral elements with quadratic quality. The size of these elements is of 24,9mm which makes a total number of elements of 21.439. With this parameters, the results are shown in Fig. 2.9 and Fig. 2.10.
Figure 2.9: Stress analysis of half of the rear landing gear.
RA 60 G06-AlOn LSA 3 seats | Project report
Figure 2.10: Displacements of half of the rear landing gear.
In this case it was also granted that there would be no buckling in the main bar. And finally, the last analysis that was performed was a dropping test. This test gave excellent results because it can be appreciated that the drop does not a have a significant impact on the landing gear. The results are shown in Fig. 2.11 and Fig. 2.13.
Figure 2.11: Tension results of the dropping test.
As the last image Fig. 2.11 is not very clear, detail Fig. 2.12 is added to give more clarity to the reader.
Figure 2.12: Detail of the point where the tension gets its maximum value.
RA 61 G06-AlOn LSA 3 seats | Project report
Figure 2.13: Displacement results of the dropping test.
2.3 Wing
2.3.1 Initial analysis In order to design the structure of the wing, it is necessary to fix design criteria and know the loads that it will deal with.
In this case, analysis will be focused on stresses produced by bending moment and beam will be designed considering only this moment. After the beam is designed, qualitative twist analysis will be performed in order to see that beam resists.
As specified in F2245[18], the aircraft must be able to operate with a maximum load factor of 4 and a minimum of -2. Thus, these will be the design criteria.
Before calculating momentum along the wing, load distribution is needed. In order to sim- plify this first analysis, there will be some hypothesis: 1. Cantilever beam. Semi-wing study 2. Constant weight distribution: 750N per wing (structure+fuel). 3. Elliptical lift distribution approximated as four uniform distributions. Also, from requirements and dimensions: 1. MTOW is 600 kg, total lift for n=1 will be approximately 6000 N 2. Wingspan is 16 m, then beam length will be 8 m.
Figure 2.14: Loads distribution and approximations
RA 62 G06-AlOn LSA 3 seats | Project report 2.3.1.1 Lift approximation
Beam will divided in four regions of 2 m length. Total lift will be divided by a multiple of four and a proportional part will be given to each region. Since each region is a uniform distribution, the value applied will be the force divided by the region length. For n different from 1, each distribution will be multiplied by n. For n=1 it was distributed as show in Table 2.10.
R1 R2 R3 R4 Total Fraction 7/16 5/16 3/16 1/16 16/16 Force [N] 1.312,50 937,50 562,50 187,50 3.000 Distribution [N/m] 656,25 468,75 281,25 93,75 -
Table 2.10: Load distribution for n=1
2.3.1.2 Weight approximation
The distribution must sum a force of 750 N. Since it is uniform, dividing by beam length we could find its value of 750/8=93,75 N/m
2.3.1.3 Results
With this information, moment along the beam was calculated with STRIAN[19].The distri- bution for load factors of interest are shown in Table 2.11 and results are shown in Fig. 2.15, Fig. 2.16 and Fig. 2.17.
Distribution [N/m} R1 R2 R3 R4 n=1 656,25 468,75 281,25 93,75 n=4 2.625 1.875 1.125 375 n=-2 -1.312,5 -937,5 -562,5 -187,5
Table 2.11: Lift distribution for each load factor
Figure 2.15: Moment diagram for n=1
RA 63 G06-AlOn LSA 3 seats | Project report
Figure 2.16: Moment diagram for n=4
Figure 2.17: Moment diagram for n=-2
2.3.2 Sizing of the beam
With the moment diagram, maximum requirements for the beam are determined and its sec- tion can be sized. From n=4, maximum moment at the embedding is 30 kN·m. This will be the requirement used to choose beam’s section. Also, from chord at the embedding of 1.000 mm, maximum thickness of 15 per cent (see Table 1.6) is 150 mm, which determines how big the beam can be. Despite the wing has variable chord, the most critical moment is found in the embedding. Thus, as decision parameter, only embedding sizing will be considered. In order to fit in the wing, beam will have variable size section. After choosing the final section from the embedding parameter, the full beam will be analyzed with FEM methods and if critical stresses appear in any regions, it will be reinforced. From F2245[4], CentrAl maximum stress can be half of its rupture tensile strength of 812 MPA. Then, maximum stress for aluminum can be 406 MPA.For carbon fibers, the maximum is 400 MPA. In order to calculate maximum stress for each material, maximum axial stress produced as result of beam’s bending will be calculated. If this stress is lower than the maximum allowed, the beam can be accepted for the structure.
Also, it will be considered that buckling won’t appear thanks to the ribs. More information about them will be given in Section 2.3.3. Two possibilities are proposed for the beam:
1. Wagner CentrAl beam (I section) with carbon fiber reinforcements
2. Square carbon fiber section with extra thickness in superior and inferior faces.
RA 64 G06-AlOn LSA 3 seats | Project report 2.3.2.1 I section I section is useful for bending stresses but requires a second beam to support torsion. In this analysis only bending will be considered. If the result has a similar weight to the other option, this will be rejected because of the extra weight added by the second beam. In order to reinforce the beam, 1 mm carbon fiber layers will be applied to the superior and inferior faces of the beam. This layer reduces the maximum thickness of the beam. It will be set in 144 mm. The procedure applied to calculate axial stresses is that for ideal composite beams [20]. Beam will be idealized with a young modulus:
∗ E = E0 (2.1)
Where E0 is Young modulus of one of the materials.Then proportional factor will be defined:
E n = i (2.2) i E∗
Where subindex determines the material(consequently, n0=1).Then inertia can be calculated as: ∗ I = ∑ ni Ii (2.3) Finally axial stress: Moment σ∗(y) = (2.4) I∗ Thus, axial stress in each material, shown in Fig. 2.18, will be:
∗ σ(y) = niσ (y) (2.5)
Figure 2.18: Stress distribution for composite beam
In Table 2.12, different dimensions, referred to those shown in Fig. 2.19, for the section and their maximum stress and weight are shown. In all the cases, web will be a 0.5 mm thick plate.
Figure 2.19: I section parameters
RA 65 G06-AlOn LSA 3 seats | Project report E [mm] Y [mm] B [mm] H [mm] Al Stress [MPa] Fiber Stress [MPa] Weight [kg] 20 72 80 4 435,9 441,9 46,0 30 72 80 4 435,4 441,4 52,9 20 72 90 4 387,5 392,9 50,0 20 72 100 4 348,8 353,7 54,0 20 72 90 3 457,0 463,3 38,8 20 72 100 3 411,4 417,1 41,9
Table 2.12: Beam configurations and results
2.3.2.2 Square section Square section might be heavier for same material beams because of its vertical walls but, since it acts as a torsion box, it won’t need a second beam. In this case, axial stress calculations are simplified to: Moment σ(y) = y (2.6) Inertia
Figure 2.20: Stress distribution
Like in the previous case, different dimensions, referred to Fig. 2.21, and its results are shown in Table 2.13.
Figure 2.21: Square section parameters
H[mm] B[mm] Y[mm] T[mm] Stress[MPa] W[kg] 4 80 72 4 421,3 50,5 4 90 72 4 391,4 52,7 4 100 72 4 365,4 55,0 3 90 72 4 430,2 44,6 3 100 72 4 399,0 46,9 3 110 72 4 372,1 49,1
Table 2.13: Beam configurations and results
RA 66 G06-AlOn LSA 3 seats | Project report 2.3.2.3 Results
Both possibilities show similar results for similar dimensions but first option is slightly heav- ier and needs a second beam, which makes it even heavier. Since weight requirement is really restrictive, it is very important to reduce weight as possible. Thus, the rectangular section has been chosen. Dimensions chosen will be those of the lightest beam within regulations. From Table 2.13, these are:
H = 3mm; B = 100mm; T = 4mm (2.7)
Since the wing tip isn’t as required as the embedding, beam’s height will be adapted to wing’s thickness but B will be reduced to 80 mm. This way, section is reduced to a 80x80 mm square beam and weight is decreased.
Beam is studied with SolidWorks FEM analysis, as specified in Section 2.2.2.1, in order to see that it can be used. Loads are those for load factor of n=4. As shown in Fig. 2.22, stress never exceed the maximum allowed of 400 MPa. The maximum stress is found near the embed- ding as expected. In Fig. 2.23, the maximum displacement appears at the wing tip. Given that these are results for a load factor of 4, a displacement of 1.100 mm is acceptable. Using trigonometrical relations, angle of displacement is calculated and results in approximately 10 degrees, which can be considered a minor displacement.
(a) Results
(b) Embedding detail
Figure 2.22: Stress analysis
RA 67 G06-AlOn LSA 3 seats | Project report
Figure 2.23: Displacement results
If this displacement appeared for cruise operation conditions, instead, the beam should be redesigned. Since real shape of the wing would be completely different from the one used for aerodynamic calculations, the beam must be changed in order to reduce displacement to a value that could be considered negligible.
However, operation with a load factor of 4 is limited to few particular maneuvers. Then, this maximum displacement would only appear momentarily.
It is interesting how FEM analysis show lower maximum stress than the initial one. The rea- son for this difference might be that 3D model used for computational analysis is different from geometry used for the initial approach. To ensure that beam fits in the wing, the beam has been modelled completely tangent to the skin. Then, thickness of the walls of the beam is measured from the nearest point to the center. This way, walls are slightly thicker than previous geometry and, consequently, inertia is higher and stress is lower.
Despite it is out of the scope of this project, it is important to mention that further analysis should be done in order to study if aeroelastic effects could take the wing to failure.
2.3.3 Ribs
Ribs are used to avoid buckling and shape the wing. They will be a sandwich composite with a central 2 mm layer made of aramid honeycomb and external 1 mm carbon fiber layers, as shown in Fig. 2.24. With its 48 kg/m3 density, honeycomb support great compression stresses with really low weight.
Figure 2.24: Sandwich ribs
To calculate distance between ribs, an approximation of the critical buckling length will be calculated for the most critical section of the wing. Since higher inertia helps to avoid buck- ling, tip’s section, with the lowest inertia, will be the most critical one.
RA 68 G06-AlOn LSA 3 seats | Project report With the purpose of simplifying calculations, the skin will be approximated as a rectangular thin airfoil beam with length equal to the chord, height equal to the maximum thickness of the section airfoil and thickness equal to skin’s thickness, as shown in Fig. 2.25. Given that thickness of the skin is out of the scope of the project, a mean value of 1 mm will be taken. Further analysis of the skin should be taken for more precise results.
Figure 2.25: Section approximation
With this considerations critical length can be calculated from:
π2EI = Pcrit 2 (2.8) Lcrit s π2EI Lcrit = (2.9) Pcrit E is Young modulus of the material. This value for carbon fiber can be found in Section 2.1. Pcrit can be calculated from critical stress and section’s surface:
Pcrit = σcritS (2.10) Since regulations stipulate a maximum stress of 400 MPa for carbon fiber, this value will be taken as the critical one. Skin can’t receive higher stresses, then it has to be ensured that buck- ling does not appear before this limit.
Section’s surface can be calculated with its perimeter and thickness:
S = pt = (2c + 2h)t (2.11) Inertia for this rectangular section is:
th3 ct3 h I = 2 + 2( + ct( )2) (2.12) 12 12 2 For tip’s section of 600 mm of chord and maximum thickness of 12 per cent (see Table 1.6) the results are shown at Table 2.14.
2 4 c [mm] h [mm] p [mm] S [mm ] I[mm ] Pcrit [N] Lcrit [mm] 600 72 1.344 1.344 1.617.508 537.600 1.442
Table 2.14: Results of buckling analysis
Since critical length is the maximum distance between ribs that avoid buckling, a lower dis- tance will be used to ensure buckling doesn’t appear despite the approximations. Distance chosen will be 1.000 mm.
2.3.4 Final Result The final result is a square section beam made of carbon fiber with composite ribs. An overall image is shown in Fig. 2.26
RA 69 G06-AlOn LSA 3 seats | Project report
Figure 2.26: Beam and ribs
2.4 Fuselage
In order to proceed to design the fuselage a series of models with different beam profiles and different configurations are going to be analyzed.
The first section that was going to be analyzed is a beam with an I shape. This shape is chosen because it is the standard beam shape and it could give an initial approach to the problem solution. In order to choose the ideal profile there will also be analyzed two normalized profiles with different shapes and a more compact section.
Once the sections were chosen, the following step was defining the number of stringers and frames the fuselage would be formed for. Usually, aircraft fuselages are formed by a large number of both of the elements [21], but as the main requirement of this concrete aircraft was a strictly limited weight, hence the main objective will be to use the minimum number of structural components at a minimum thickness.
Therefore, there will be used four main stringers that will go from leading edge to trailing edge. It could have been considered the possibility of using only three stringers (one for the part above, one for the part below and one on one side) but this would have generated an horizontal displacement of the gravity center, which is not convenient. Then, it must also be taken into consideration that the stringer that crosses the windshield will be divided in two parts. For the frames, two main frames will be considered, located the first one where the windshield begins and the other one, formed by only half arch; where it ends, to maintain the structural stability. It will not be considered the addition of more frames because even though the skin is thin, it will have a clamping function. With this distribution, three possible solutions can be raised: one with this exact configuration and a thickness of the profile of 1 cm Fig. 2.27, another with the same structure but a thickness of 0.5 cm Fig. 2.28, and a last one as the previous one but with the two frames with their structure complete, in case reducing the thickness could affect to this specially empty part Fig. 2.29. For the other two cases, there will be used two main stringers that will go from the leading edge to the trailing edge. To complement these ones, two shorter stringers will be added along the cabin. However, to compensate the structural loss of reducing the length of the stringers three additional frames are added. Those two models have a total of seven frames, five along the cabin and two more frames along the tail. These options raise two more possi- ble solutions Fig. 2.30 Fig. 2.31.
The five options can be seen in the following images:
RA 70 G06-AlOn LSA 3 seats | Project report
Figure 2.27: Fuselage with a beam thickness of 1 cm.
Figure 2.28: Fuselage with a beam thickness of 0.5 cm.
Figure 2.29: Fuselage with a beam thickness of 0.5 cm and a reinforcement.
RA 71 G06-AlOn LSA 3 seats | Project report
Figure 2.30: Fuselage with a U beam.
Figure 2.31: Fuselage with a rectangle beam.
To choose the best one between this alternatives, a series of analysis will be performed.
2.4.1 Analysis
For the five options, three main possible problems have to be considered. The principal prob- lem is that the stringers will be working at traction, considering that the thrust will be acting pulling the airplane forward while the drag will be acting pulling the aircraft backwards. Given these forces, another problem that has to be taken into account is the possibility of buckling in the stringers. At the same time, the last problem that can happen is the deforma- tion of the frames.
As it has been chosen a concrete profile to avoid bending moments and buckling the main problem that may concern to the aircraft structure is that the stringers will be working at trac- tion and has to be checked that none of the configurations might generate a configuration of tensions that surpasses the elastic limit of the material.
However, it must be taken into account that due to the choice of a composite material for the fabrication of the fuselage, the Young’s modulus will be different depending on the direction that is analyzed. This could have meant that finite element methods are not valid for this analysis but given that the analysis will be performed with the material working at traction and with the forces applied in a longitudinal direction to the stringers; it can be used the lon- gitudinal Young’s modulus with a negligible error. A slightly higher security factor will be taken although to ensure the security of the structure because the cross Young’s modulus is lower.
RA 72 G06-AlOn LSA 3 seats | Project report Then, provided that the analysis will be acceptable to the case, it can be performed to decide which configuration will be used, and the thickness of the profiles, will be used the following criteria:
1. Weight of the fuselage structure.
2. Maximum tension.
3. Security factor (Elastic limit of 8, 122e + 8).
4. Maximum displacement.
2.4.1.1 Performed analysis To evaluate this, the five analysis have been performed with a finite element method under the same conditions: two distributed loads applied in each edge that will represent the thrust and the drag. The thrust has been estimated based on the engine power and the aircraft characteristic velocity and the drag has been simply calculated for a horizontal straight flight. This way, the applied loads are:
Enginepower Thrust = (2.13) characteristicvelocity
59.656W Thrust = = 1.368, 2569N (2.14) 43, 6m/s and, for a case where the thrust is not equal to the drag (a flight with acceleration) and from equalling the maximum aircraft weight to lift:
1 MTOW = ρv2SC (2.15) 2 L From this equation, using one more time the characteristic velocity, it can be obtained the value of the CL, and furthermore, the value of the CD:
CD = 0, 01016 (2.16)
1 D = ρv2SC = 146, 984N (2.17) 2 D These loads will be applied in a perpendicular direction to the aircraft axis. Aside, it has to be said that Lift and Weight loads will not be considered because they do not affect directly to the fuselage structure. The last thing that must be specified are the mesh parameters. It is used a structured mesh with four Jacobean points. The quality of the tetrahedral elements is quadratic and its size is of 30,6mm.
2.4.1.2 Results of the analysis These results are presented graphically in figures from Fig. 2.32 to Fig. 2.41, and the choice between the alternatives has been taken using an OWA method.
In order to establish the weights, it will be taken into account that the weight is the most restricting item. Then, the second one will be the security factor because it is recommended to be of 2. So, overdimensionated structures will be directly related to extra weight. Finally, the maximum deformation will also have a point, but not so important because the skin will also retain the fuselage structure.
RA 73 G06-AlOn LSA 3 seats | Project report The results are presented in the following table:
Weight Max. tension Security factor Max. displacement 1cm thick I 215,5 kg 1,123e+7 Pa 72,32 7,39 mm 0,5 cm thick I 72,5 kg 2,487e+8 Pa 3,27 18,25 mm 0,5 cm thick I 94,6 kg 3,567e+7 22,77 3,173 mm with reinforcement 0,35 cm thick U 41,49 kg 2,181e+7 37,24 2,01 mm 0,6 cm thick rectangle 72,54 kg 8,06e+6 100,77 1,21 mm
Table 2.15: Principal results of the different fuselage alternatives analysis
Then, in order to choose the better option, an ordered weighted averaging method was used, obtaining the following results.
Weight (kg) Max. Deformation (mm) Security factor Total p*g Score OWA Weight 6 2 4 70 p 1 3,54 2,17 1cm thickness I p*w 6 6,9 8,68 21,58 0,31 p 3,86 1 5 0.5cm thickness I p*w 23,16 2 20 45,16 0,65 p 3,42 4,54 4,20 0,5cm thickness reinforced I p*w 20,52 9,08 16,8 46,4 0,66 p 5 4,81 3,61 0,35cm thickness U p*w 30 9,62 14,56 54,18 0,77 p 4,76 5 1 1,6cm thickness rectangle p*w 28,56 10 4 42,56 0,61
Table 2.16: Results of OWA method
Which led to the conclusion that most adequate alternative was the fuselage with a U profile beam.
Figure 2.32: 10mm thick I:Results of the stress analysis.
RA 74 G06-AlOn LSA 3 seats | Project report
Figure 2.33: 10mm thick I:Results for the displacement.
Figure 2.34: 5mm thick I:Results of the stress analysis.
Figure 2.35: 5mm thick I:Results for the displacement.
RA 75 G06-AlOn LSA 3 seats | Project report
Figure 2.36: 5mm thick I with reinforcement:Results of the stress analysis.
Figure 2.37: 5mm thick I with reinforcement:Results for the displacement.
Figure 2.38: 3,5mm thick U:Results of the stress analysis.
RA 76 G06-AlOn LSA 3 seats | Project report
Figure 2.39: 3,5mm thick U:Results for the displacement.
Figure 2.40: 16mm Rectangle: Results of the stress analysis.
Figure 2.41: 16mm Rectangle:Results for the displacement.
2.5 Tail
2.5.1 Elevator beam
For beam analysis, same procedure applied to the wing will be applied. This time, force dis- tribution will be approximated as uniform. The value for the total lift is calculated with the
RA 77 G06-AlOn LSA 3 seats | Project report maximum lift coefficient from Fig. 1.21.Then:
1 L = ρV S CL = 7.500N (2.18) 2 ne e max Using uniform distribution, the value of the loads are shown in Table 2.17.
Length [mm] Total Force [N] Distribution[N/m] 4.000 7.500 1.875
Table 2.17: Total load and distribution
The results, calculated with STRIAN [19] are shown in Fig. 2.42.
Figure 2.42: Elevator momentum diagram
Then, maximum moment at the embedment is 3.750 N·m. The results for different rectangu- lar beams are shown in Table 2.18.
H [mm] B [mm] Y [mm] T [mm] I [mm4] Stress[MPa] W [kg] 2 45 30 2 223.382 503,6 3,0 3 45 30 2 259.384 433,7 3,8 2 45 30 3 291.316 386,2 3,6 3 45 30 3 327.321 343,7 4,4 3 40 30 3 302.953 371,3 4,2 1 55 30 3 304.047 370,0 3,2 1 50 30 3 284.553 395,4 3,0
Table 2.18: Dimensions for elevator beam. Referred to Fig. 2.21
After some iteration, the lightest beam is chosen and its dimensions are:
H = 1mm; B = 50mm; T = 3mm (2.19) Finally, FEM analysis is shown in Fig. 2.43.
As happened with the wing, stress from the analysis is below the stress previously calculated. The reason might be the same, 3D model used is slightly different from ideal geometry used in initial analysis.
From trigonometrical relations,a displacement of 330mm results in about 10 degrees. As con- sidered in the wing, it can be accepted.
RA 78 G06-AlOn LSA 3 seats | Project report
(a) Stress
(b) Displacement
Figure 2.43: Elevator FEM analysis
2.5.2 Fin beam The procedure and approximations for the fin are the same applied to the elevator. The value for the total lift is calculated with the maximum lift coefficient of 1,2 from Fig. 1.26.Then:
1 L = ρV S CL = 4.000N (2.20) 2 ne e max Now, assuming an uniform distribution, the value of the loads are shown in Table 2.19.
Length [mm] Total Force [N] Distribution[N/m] 1.200 4.000 3.335
Table 2.19: Total load and distribution
The results, calculated with STRIAN[19] are shown in Fig. 2.44.
Then, maximum moment at the embedment is 2.400 N·m. The results of some iterations for beams are shown in Table 2.20.
This time, weight barely changes between options compared to MTOW of 600 kg. Then, decision criteria will be two: weight must be lower than 1 kg and maximum stress on the
RA 79 G06-AlOn LSA 3 seats | Project report
Figure 2.44: Fin moment diagram
H [mm] B [mm] Y [mm] T [mm] I [mm4] Stress[MPa] W [kg] 1 45 45 2 187.381 384,2 0,6 2 45 45 2 223.383 322,3 0,9 1 45 45 1 114.323 629,8 0,4 1 40 45 2 170.561 422,1 0,6 2 40 45 2 206.563 348,6 0,8 2 35 45 2 189.743 379,5 0,8 1 30 45 3 182.210 395,1 0,6
Table 2.20: Dimensions for fin beam. Referred to Fig. 2.21 embedment will be the minimum between different options. Thus, beam chosen and its di- mensions are:
H = 2mm; B = 45mm; T = 2mm (2.21)
Finally, FEM analysis is shown in Fig. 2.45.
(a) Stress (b) Displacement
Figure 2.45: Fin FEM analysis
Results are within regulations. Same conclusions as in elevator can be taken.
RA 80 G06-AlOn LSA 3 seats | Project report 2.5.3 Ribs Using the same methodology that is used to calculate distance between ribs in the wing, re- sults for fin and elevator are shown in Table 2.21 and Table 2.22. For fin, chord is 500 mm and maximum thickness is 12 per cent. For elevator, chord is 250 mm and maximum thickness is 12 per cent.
2 4 c [mm] h [mm] p [mm] S [mm ] I [mm ] Pcrit [N] Lcrit [mm] 500 60 1.120 1.120 936.083 448.000 1.201
Table 2.21: Buckling analysis for fin
2 4 c [mm] h [mm] p [mm] S [mm ] I [mm ] Pcrit [N] Lcrit [mm] 250 30 560 560 117.042 224.000 601
Table 2.22: Buckling analysis for elevator
As explained in Section 2.3.3, lower distances are used to ensure that ribs work. Thus, length chosen for the vertical distance is 1.000 mm and 500 mm for the horizontal.
2.5.4 Final Result The final result are two rectangular carbon fiber beams with the same composite ribs used in the wing. An overall image is shown in Fig. 2.46.
Figure 2.46: Beams and ribs
RA 81 G06-AlOn LSA 3 seats | Project report
RA 82 Chapter 3
Overall Design
The principal reason to have a CAD model is to make it easier to draw the blueprints needed for the technical sheets. However, since CAD software offers much more possibilities than just drawing, the model has been used as a support for other departments. In this case, the software used have been Catia v5 and SolidWorks. Furthermore, renders and a mock-up will be done.
3.1 3D design
Once the materials have been chosen in section 3.1, CAD software gives the center of gravity of the model, which is necessary to know for stability purposes. This model have some exterior parts, which will be used for the 3d printed model, an interior parts, which are only used to find the c.g position. This is why in some cases, they aren’t a real representation, they are solid blocks positioned as required with volume and density of these parts.
3.1.1 Exterior Exterior includes fuselage, wing, tail and landing gear as shown in Fig. ??. Fuselage and surfaces have been imported from XFLR5 aerodynamic analysis. Landing gear was chosen as explained in Section 2.2.2.
Figure 3.1: Exterior model
RA 83 G06-AlOn LSA 3 seats | Project report 3.1.2 Interior Interior includes everything which must be inside the skin: structure, seats arrangement, belts, baggage, parachute, fire extinguishers, instruments and cockpit.
Regarding the seats, it has been decided a 1:2 configuration: 1 in front of the cockpit after the lever and 2 behind. Five-point belts have been placed in each seat and explained in 5.4. The luggage will be placed under the seats, weight depending on the number of passengers. Two fire extinguishers have been placed in the posterior part and a parachute is allocated also in the posterior part.
Furthermore, the cockpit is in the front of the first seat where the pilot maneuver every move- ment of the LSA. A lever is 127 mm in front of the cockpit, separated 254 mm from the pilot between their legs. Finally, it has been placed a throttle lever next to the pilot and pedals under the cockpit.
3.1.2.1 Structure
Used to calculate weights and surfaces. Weight is used as a decision criteria for beams and surface are used to calculate the amount of material needed. After selection, all models were simulated in SolidWorks
Figure 3.2: Full structure
3.1.2.2 Power plant
It has been approximated as a box of 40 kg and dimensions 500x500x420 mm. As propeller modelling only has aesthetical purpose, it was downloaded from GrabCad [22].
Figure 3.3: Power plant model
RA 84 G06-AlOn LSA 3 seats | Project report 3.1.2.3 Instruments and equipment Regarding the instruments and equipment in the inside of our LSA, it has been used belts in each seat and additional ones in the posterior part for the baggage of the passengers.Besides, it has been displayed the different instruments used and explained above in Fig. 3.4
Figure 3.4: Instruments displayed, Cockpit
Which its size is 500x400 mm so it fits in the cabin. The deepest instrument is 141 mm, so it is made all along the cabin covering every instrument and leaving space under it for the pedals placement.
3.1.2.4 Seats and passengers It has been estimated a block of 87 kg, which includes seat, belts, passenger and their luggage. Its dimensions were 1.200mm height and 500mm wide. The model is shown in figure 3.5
Figure 3.5: Passenger model
The seats used in our LSA have been selected after a research of different seats in other LSAs. These seats must be decided according to the sizing of the inside part and an average person and their minimum comfort. The three seats will be the same.
In the following, it is showed two companies which provide seats for sports aircraft: Air-Tech Inc. and Sport Aircraft Seats. The first one gives us a general overview of the dimensions of a seat in this type of airplanes. This company offers, focused on what it is needed, a fiberglass seat structure and a cover seat in black for that structure, besides its technical drawing is given in figures 3.6b and 3.7 [23].
RA 85 G06-AlOn LSA 3 seats | Project report
Figure 3.6: Seats and prices of Air-Tech Inc.[23]
Figure 3.7: Measures, technical sheet of the Ultralight seat [23]
The second company Sport Aircraft Seats, provides a whole seat in “one-piece” between fab- ric and leather[24]. Therefore, it has been chosen a fabric seat instead of a leather one because of its breath-ability as well as how it behaves in cold weather. In this case, it is possible to choose the P110 structure-seat, so it fits for the passenger and our designed aircraft.
RA 86 G06-AlOn LSA 3 seats | Project report 3.2 Blueprints
On early stages of the project, a first document with general dimensions of the aircraft was given to Structures department in order to start sizing of the structure. This overview was prepared with the dimensions used by Aerodynamics department during their analysis and weren’t definitive, so there might be minor changes between figure 3.8 and the final aircraft. Final blueprints are those shown at the technical sheets.
Figure 3.8: Initial sizing
RA 87 G06-AlOn LSA 3 seats | Project report
RA 88 Chapter 4
Business
In this chapter there will be an analysis, study and planning of all the economical and busi- ness points regarding the aircraft itself and the enterprise. A further analysis of the latter will be carried out as well.
4.1 Manufacturing costs
Taking into account all the costs of the project and the fact that there will be a development of an enterprise behind the aircraft, in order to make profit of this product (sold by the com- pany), one must consider all the costs the company will face. All these manufacturing costs have been split into four different categories: • Materials • Facilities • Human resources • Additional costs Some of this categories will be discussed in the next sections.
4.1.1 Cost of facilities To be able to choose an office, some previous calculations must be done to have an order of magnitude of how big the work area must be. Assuming one person takes up approximately 3.5m2 of personal work room, where they have their desk and personal objects. The total space required for one person to work will be evaluated as 14m2 [25]. As there are 14 people working the required room will be approximated as ≈ 14 · 14 = 196m2. Now that an order of magnitude has been set of how big the total room of the office must be, these four different cases can be studied in order to make the final choice. 1. Office 1: C/ Almogàvers 119, Edifici ECOURBAN, 08018 Barcelona. Total room available: 431 - 909 m2. Cost: 16e/m2.
RA 89 G06-AlOn LSA 3 seats | Project report
(a) Floor plant (b) Location
Figure 4.1: Specifications of Office 1
2. Office 2: C/ Calabria 169, 08015 Barcelona. Total room available: 170 - 343 - 418 m2. Cost: 11e/m2.
(a) Floor plant (b) Location
Figure 4.2: Specifications of Office 2
3. Office 3: Av. Diagonal, 309, 08013 Barcelona. Total room available: 113 m2. Cost: 10, 00e/m2.
(a) Floor plant (b) Location
Figure 4.3: Specifications of Office 3
4. Office 4: C/ Bailén, 3, 08010 Barcelona. Total room available: 188 m2. Cost: 11e/m2
With these four options considered, a decision will be taken on the best office on terms of room available and location. Taking into account the room needed is approximately 200m2,
RA 90 G06-AlOn LSA 3 seats | Project report
(a) Floor plant (b) Location
Figure 4.4: Specifications of Office 4
Office 3 falls behind as far as room is concerned. Then, considering the total amount it would cost to rent each office, one gets:
Price per m2 [e] Room [m2] Total rent cost [e] Office 1 16 431 6.896 Office 2 11 170 1.870 Office 3 10 113 1.130 Office 4 11 188 2.068
Table 4.1: Comparative of the 4 possibilities
Now that the total amounts have been computed, an objective and thoughtful choice can be taken. The best option would be Office 4, as it provides enough room (it just lacks 8m2), it is in a very good spot in Barcelona, and the price per squared meter is low enough to keep the budget within the margins. So, finally, one gets a total cost of facilities of 2.068 eper month.
4.1.2 Cost of human resources In the human resources category it will be assessed how much money the company must spend on the people that will work in it. Such quantity is mainly composed of the different salaries of the workers (engineers and coordinators). In order to get the total sum of salaries, the workers wil be divided into two separate cate- gories: nine engineers and four coordinators. The total amount of hours per week will be estimated to be eight hours for the two cate- gories. With this, the salary will be determined by the total amount of hours required to do the project, and the cost per hour of an engineer/coordinator. The remaining team member is the Project Manager, who belongs to a different category. Assuming the project takes a total amount of 11 weeks to be finished, so the total amount of hours worked is:
Workers Weeks Hours/week/worker Salary/hour Total amount [e] Engineers 9 13 8 25 23.400 Coordinators 4 13 8 35 14.560 Project Manager 1 13 8 50 5.200 Total 14 13 8 - 43.160
Table 4.2: Basic human resources cost
With this first estimation one gets the total amount that must be paid to the workers as gross salary. Until this point, any extra hours nor external personnel have been considered. In order to get a better approximation of the total cost of the human resources, one must add a percentage of the gross amount, in case the workers have to work extra-hours. Such per- centage will be estimated as the 10% of the total. So, in conclusion, the total cost of human resources will be estimated at 43160 e+ 4316 e= 47.476 e. This final result consists only of the salaries of internal personnel of the company. It does not
RA 91 G06-AlOn LSA 3 seats | Project report contain any additional personnel costs, as it will be considered in the Additional costs subsec- tion.
4.1.3 Additional costs
Certification of the aircraft: The certification of the aircraft is a process that must be carried out while all the aircraft is be- ing designed. EASA gives the companies the opportunity to certificate its own aircraft model, delivering all the needed documents once it is finished, or delivering each document sepa- rately once that specific part is finished. This system works very well because it gives room for error, so if an specific part of the aircraft does not satisfy EASA’s regulations (Part21A.701 (15)) [26], we can easily go back and correct it before the aircraft is completely finished. So, in order to compute the cost of the certification of the aircraft, one can take the Total Budget of Human Resources (43160 e) and add a 10%, as it is approximately the time an engineer will have to dedicate to check if the part they are working on satisfies the regulations. The total amount of the certification is then: 0.1 · 43.160e = 4.316e.
4.1.4 Marketing costs
One of the main costs of the company is the marketing cost, not in terms of total volume, but in terms that it has to be well defined, as a high percentage of the sells will depend on how good the marketing campaign is. In this section, the cost of the marketing campaign will be studied. The marketing campaign in itself will be thoroughly studied on section 7.2 Marketing campaign.
4.1.4.1 Initial marketing campaign costs
For new companies, the cost of the marketing campaign is set to be ≈ 8 − 12% of the total budget. This amount, though, is lower in this project, since the marketing campaigns of prod- ucts like planes are far more centered in the potential costumers. The total amount invested in marketing is, thus, 8.650 e. With 8.650edesignated to the budget of the marketing cam- paign, this total amount will be split into the marketing subsections that can be seen on Table 4.3: So the total amount spent on marketing is $ 8.650,00, what has been designated. The weight distribution of each subcategory can be seen on Figure 4.5.
Figure 4.5: Specific weigh t by category
RA 92 G06-AlOn LSA 3 seats | Project report CATEGORY QUANTITY UNIT COST SUBTOTAL
National Marketing 250,00 e Banner Ads 1 250,00 e 250,00 e
Local Marketing 1.300,00 e In-Store Marketing 2 400,00e 800,00 e POP 1 00,00 e 500,00 e
Public Relations 700,00 e Public Events 1 500,00e 500,00e Press Releases 2 100,00e 200,00e
Online 400,00e Blog 1 400,00e 400,00e
Advertising 2.350,00e Online 1 1.500,00e 1.500,00e Print 1 850,00e 850,00e
Sales Campaigns 800,00e Campaign A 1 800,00e 800,00e
Market Research 350,00e Surveys 2 175,00e 4.800,00e
Content Marketing 600,00e Sponsored Content 3 100,00e 300,00e Landing Page 2 150,00e 300,00e
Web 750,00e Development 1 750e 750,00e
Other 900,00e Premiums 2 200,00e 400,00e Corporate Branding 1 250,00e 250,00e Business Cards 10 25,00e 250,00e
GRAND TOTAL 8.650,00e
Table 4.3: Marketing costs by category
4.1.4.2 Study of marketing costs over time
Being able to sustain the marketing costs implies the success on sales of the company. Taking into account the high amount that a 1-year marketing campaign costs 8.650 e, and consider- ing the capital assigned to marketing is approximated to be a ≈ 6 − 10% of the company’s benefits, the benefits of the company on the first year should be around $ 86.500. If this initial marketing campaign is sustained over time (see subsection 7.1.5.3 Study of future marketing campaigns), and thus having benefits of over 86.500 e, the company will carry on having this marketing cost. The impact of the initial investment on the annual benefits will be studied in the sections 7.3 Initial Investment and 7.4 Payback Analysis.
RA 93 G06-AlOn LSA 3 seats | Project report 4.1.4.3 Study of future marketing campaigns Once the first exercise is completed (i.e. the first year), the future marketing campaigns may vary from the first one. This differences may reside on the approach, on the scope or on the platforms used in the new marketing campaign. In this prediction, two different situations can occur:
• Success of sales. Assuming that at the end of first exercise, the benefits are high enough (so the sales have been successful, and so the marketing campaign), there would not be any need for a wider marketing campaign. So the marketing campaign for exercise 2 would be equal, or even a reduction of the first one. Considering this situation, the total amount of capital needed for marketing issues on the future can be reduced to a 7-10% of the benefits of the company. This is the best scenario for the company.
• Lack of interest/ low number of sales. Assuming now, that at the end of the first exercise the sales have not been high enough, so not enough profit has been made by the company, there should be some changes on the marketing campaign of exercise 2. Firstly, the amount of capital destined to marketing would rise to 15% of the benefits of the company in exercise 1. Secondly the scope of the campaign could be broadened in order to reach more potential customers and achieve the goal in exercise 2.
RA 94 G06-AlOn LSA 3 seats | Project report 4.2 Marketing campaign
This section is focused on the marketing issues of the project, from the possible customers, to the design of the marketing campaign.
4.2.1 SWOT analysis
Here we explore the Strengths and Weaknesses of our product and the Opportunities and Threats to our company; to better approach decision-making and to more adequately focus our marketing efforts. This kind of analysis is called a SWOT analysis (See Figure 4.6), and it provides a simplified and structured vision of the whole Business project. [27] [28]
Figure 4.6: Scheme of a SWOT analysis.
Source: https://research-methodology.net/theory/strategy/swot-analysis/
4.2.1.1 Research on Market Opportunities
The location in Barcelona gives the company access to a large specialized labour pool. Ties and contacts within Catalonia’s Polytechnic University can yield us cheap interns as well as further access to aerospace engineers, both recent graduates and veterans in the field. Lack of airplane manufacturers in the area only reinforces this access to the labour pool, allowing us to pay lower salaries. Low cost-of-living in Spain vs. the US, Canada and the rest of the EU (where most our competitors are based) further reduces our labour costs relative to the competition.
4.2.1.2 Research on Market Threats
The aforementioned lack of plane manufacturers in the Area also means a less experienced labour pool. Well-established companies in the field could undercut our prices because they benefit from economy of scale. Since the product sold is a sport aircraft without LSA attributions (because of its higher than three seat capacity), it is taken as competitors manufacturers both that sell actual LSA and those that sell three-seated sport aircraft. The main competitors for the company are Flight Design USA with its model CTLS and CTLS Lite,then Czech Sport Aircraftwith the models CLub, Tourer and Professional,and finally CubCrafters with the models Carbon Cub SS and Sport Cub S2. These three companies are known as the leaders of LSA’s sellers. Also One Aircraft has developed a model of LSA with three-seats called ONE 2+1 that will be in the market in 2020.Since this product is basically the same type of LSA model as Alpha One , this could be the main threat for the company, so Alpha One’s aircraft should be in the market before ONE 2+1 Another issue is that volatility of oil prices and EUR-USD exchange rates can affect customers’ demand and harm the company’s long term plans.
RA 95 G06-AlOn LSA 3 seats | Project report 4.2.1.3 Research on Aircraft Strengths
The aircraft offers high fuel efficiency unmatched by any other aircraft capable of carrying three passengers. The aircraft being manufactured out of mostly custom-made parts secures us future funding in the form of repairs and maintenance. Furthermore, the possibility of the airplane entering the category of LSA in a future, makes it even more appealing to Sport Pilots, given they will be able to carry more passengers (and thus more weight) without the need of another license.
4.2.1.4 Research on Aircraft Weaknesses
Limited flying range makes the product unsuitable for the American and Canadian markets, as well as alienating any potential customers who seek medium to long range sport aircraft. The aircraft being manufactured out of mostly custom-made parts makes us more liable to litigation due to malfunction-related accidents. Also the main weakness for any starting company is the lack of experience in the aviation industry, and not having any reputation makes it harder to get into the labour market.
4.2.2 Study of potential customers The main customer market to which the company can center its sells is the one that is already buying conventional two-seats LSAs. With this in mind, the attractive this project’s LSA (AlOn) has over other conventional LSAs is its bigger capacity. Although the total weight is the same, as it is regulated by the OACI (and thus cannot be exceeded), it provides extra room for one person, or to carry more payload if necessary. The market of aviation grew ap- proximately a 6,5% between 2016 and 2017 (see Image 4.7), so a new idea of LSA has enough room in the market to start its own client portfolio. [29] In the next table one can see the Gen- eral Aviation Airplane Shipments of Piston-Engine planes between 1995 and 2017. The great decay from 2007 to 2010 is believed to have been caused by the economical crisis Europe and the US suffered on this period of time. [30]
Figure 4.7: Airplane Shipments Worldwide (1995–2017)
[31]
RA 96 G06-AlOn LSA 3 seats | Project report Airplane Shipments Worldwide (1995–2017) Year Total Single-Engine Piston Multi-Engine Piston Change 1995 666 605 61 - 1996 801 731 70 20,27% 1997 1123 1043 80 40,20% 1998 1606 1508 98 43,01% 1999 1801 1689 112 12,14% 2000 1980 1877 103 9,94% 2001 1792 1645 147 -9,49% 2002 1721 1591 130 -3,96% 2003 1896 1825 71 10,17% 2004 2051 1999 52 8,18% 2005 2465 2326 139 20,19% 2006 2755 2513 242 11,76% 2007 2675 2417 258 -2,90% 2008 2119 1943 176 -20,79% 2009 963 893 70 -54,55% 2010 889 781 108 -7,68% 2011 898 761 137 1,01% 2012 908 817 91 1,11% 2013 1030 908 122 13,44% 2014 1129 986 143 9,61% 2015 1056 946 110 -6,47% 2016 1019 890 129 -3,50% 2017 1085 936 149 6,48%
Table 4.4: Evolution of Airplane Shipments worldwide between 1995 and 2017
4.2.3 Set marketing goal As a start-up, the main goal is to make the company known. Putting itself in the public eye is, however, not a must. The company needs only make itself known to its potential customers, a very specific subset of the general population. Furthermore, an effort to appear professional and trustworthy must be undertaken, since the lack of an established reputation will make clients wary of acquiring the product, fearing to be scammed or given a sub-par purchase. A policy of openness and transparency about the way business is conducted needs be approached, whilst avoiding to harm the company’s own interests. Because of the potential for a profitable after-sales relationship with the customer, the com- pany must strive for a good relationship with them, presenting itself as friendly and close, but without sacrificing the aforementioned professionalism.
Moreover as the marketing goal is set, other concepts need to be defined such as: • Mission: this concept answers the questions: What is our service at the moment? Which is our competitive advantage? A good mission allows a company to reach their vision. In our case our mission is to provide a brand new product, an LSA suitable for an additional passenger, to all the aviation fanatics and LSA lovers. • Vision: this refers to what do we want to become in the future as a company. Since we have come up to the market with a new product, our vision is to expand the company, become the best sellers of this type of light sport aircraft and maybe design a variant of the product. • Values: this is one of the most important aspects to consider. The values of a company
RA 97 G06-AlOn LSA 3 seats | Project report refer to the work philosophy that all the workers of it share such as respect, cooperation, discipline and perseverance.
4.2.4 Study of advertisement To develop the advertising plan that will be developed,it will be done a general study of the best options to invest in advertising, which media is the best one and possible companies that can provide the company this service.
The main media used for advertising in aviation is print media, magazines basically and dig- ital media. Although nowadays digital media seems to be the most powerful way to reach the maximum possible costumers, Alpha One’s target is set on people that are used to print media, so we will focus on that.
Following it is seen the different companies that could provide the company the advertise- ment desired:
• Air Charter Guide : it is the most-read authoritative guide to the air charter industry. This company can provide the company visibility by advertising in the guide pub- lished twice a year, and as long as we commit to advertise the company will promote Alpha One on their website with no extra charge. The Air Charter Guide has 171.738 subscribers all over 148 countries so we would have a wide visibility. The price for a quarter page ad costs around 1.165 $.[32]
• IATA (International Air Transport Association : this large association has half million read- ers worldwide and the website has over 2,3 million views per month. Although the association is more focused in air transport and commercial aviation, more feasible for airline advertising, we could consider advertising with them since they provide flexi- bility to pick and choose our target audience.[33]
• Flying Magazine by Bonnier’s Marine & Aviation Group: it has the largest paid subscription for an aviation magazine, that positions it as the world’s most widely read aviation magazine.[34] Bonnier’s Group offers the possibility to advertise in either the monthly printed publication in any size desired or in their website. Moreover the website can promote our product via digital media not just on their website, also in their social network and encourages the company to have an account for networks like Instagram, Facebook...[35]
• Perfect Landing: this last company is an aviation marketing agency that it would taken into account as an auxiliary help for designing the website or if needed to help us find possible advertisers as they are specialized in this field.[36]
The company has decided to work with all the companies mentioned above. Since all of them can provide us good visibility, we will advertise by printed media and also via online websites with IATA, Air Charter Guide and Flying Magazine with the periodicity set by each company.We will hire the service of Perfect Landing as well to help us develop our website and supervise our marketing plan. Another digital platform that we could use as a way of broaden the possible costumers is using social network as Instagram and Facebook since the both of them are free and would not be an additional cost, a part from designing the Alpha One website.
4.2.5 Design of marketing campaign Basic marketing campaigns are based on a few requirements such as the analysis of the actual market, determining a marketing goal and decide which will be the actions done to set that
RA 98 G06-AlOn LSA 3 seats | Project report goal.
First of all, the analysis of the actual market was made when the SWOT analysis was done, and it was checked who the main competitors in the industry were, companies that sold a similar product to ours and had more experience and fame than our start-up.
On the other hand, the marketing goal and advertisement development was set in previous sections. So in this section one can see the logotypes designed for the brand . Here-under we will show the logotypes designed and the final choice: So finally the last logo was chosen because it was thought that it was simple and serious,
(a) Logo proposal 1
(b) Logo proposal 2
(c) Logo proposal 3
Figure 4.8: Different logo proposals although the other logos are really modern and could be more suitable for start-up. Since our target is mainly business men and not young people, we thought that the simplicity and the colour black would show more seriousness and could be more suitable for our project.
RA 99 G06-AlOn LSA 3 seats | Project report 4.3 Initial investment
4.3.1 Research on possible investors This is a key section of our project. Since the Initial Investment is far larger than what it takes to sustain the company once it has started producing, the project members must seek some investor to be able to make the project a reality. The origin of the money is a very important aspect when it comes to the economical feasibility, as it has a big impact on the factor k, that measures the profitability of the whole project within a period of time. The origin of the money could be: • A bank
• A group of people that invests money (angel investor)
• The members of the project Since we have already discarded the last option, as the Initial Investment is beyond the group’s range, we will focus on the other two options: a bank, or an angel investor. Investing in a brand new project implies risking the money you have invested, as one can not have 100% certainty that the project will be successful. Given this, having a bank investing money in a project like this is usually quite difficult, as they tend not to have a risk factor over ≈ 2%. On the other hand, angel investors demand a high IRR (Internal Rate of Return) in order to earn money in a short period of time. This IRR is usually between 20-30% in the first 5 years of the project. [37][38] The project can find angel investors online, for example at https://angel.co/europe/investors. Last year, in Spain, these angel investors invested a total amount of 21.2 M e. This total amount was a 78% larger than last year. This increase demonstrates the fact that every year the angel investors are more relevant worldwide.[39]
4.4 Payback Analysis
Once the project is all set up, one must carry out an analysis of economic feasibility of the project, regarding if this project is worth developing or not. This economical feasibility anal- ysis is thoroughly carried out in the Chapter 6 Economical Feasibility of the main report. In this section, we will analyze how the price of the plane (set as 160.000,00 e), can vary in order to make it more attractive to the clients or to make it more profitable for the project. This analysis is called the Study of profitability margins
4.4.1 Study of profitability margins As commented previously, in this section the price of the plane will be varied and analyzed in a range from 100.000 to 200.000, computing for each price the NPV, the PBT, and the IRR. This values will be later compared with the values obtained for our project, in the Economical Feasibility analysis, carried out in the Report, Chapter 6. On the figures 4.9, 4.10 and 4.11 one can see the graphics of this computations. It has been highlighted with an orange line the selling price set for our project, in order to have a better visual on where AlphaOne lays in each graphic. As figure ?? shows, the company will not make any profit (in 7 years, with rate of decay k = 11,4%) until the selling price per aircraft is more than 130.000 e. After this price is surpassed, the company can increase its revenues by increasing the price. The intersection of the two curves gives us the NPV calculated for a selling price of 160.000,00 ewith a total value of 938.113,37 e.
On figure 4.10 one can see the evolution of the Pay-Back Time. As it can be seen, it decreases with the price, as expected, tending to 0 as the price increases. The PBT for prices per unit
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Figure 4.9: Net Present Value (years = 7, k = 11,4%) for different selling prices
Figure 4.10: Pay-Back Time for different selling prices
Figure 4.11: Internal Rate of Return for different selling prices less than 120.000,00 eincreases significantly as it tends to ∞. Once more, the intersection of the two curves gives the PBT of AlphaOne, 3,61 years. Finally, on figure 4.11, one can see the evolution of the Internal Rate of Return with the selling prices, like in the two previous figures. This latter figure shows that for prices less than ≈ 130.000,00 e, the IRR turns out to be negative. This indicates us the non-feasibility of the project, without taking into account the decay rate k. If one sets the IRR to be higher than k, the project starts being feasible once the price reaches ≈ 140.000,00 e, and then gains feasibility with higher prices. Two things could be done once we have this results:
• Lowering the selling price: Lowering the selling price might lead us to more sells, but not in the first years, as the company has to adapt itself to the new market (as com- mented on the Report). Even if the company sold one more unit per year, a price of 150.000 ethe PBT would not be any higher (3,74 years), and would have an IRR of 32 %, lower than the current IRR of 33 %.
RA 101 G06-AlOn LSA 3 seats | Project report • Increasing the selling price: Increasing the selling price would turn out to be profitable once the company has enough fame, and a customer wallet big enough to be able to do so without losing sells. Increasing the selling price to 170.000, and considering one unit sold less every year, one gets a PBT of 3,7 years, and an IRR of 32 %, both with economical disadvantages to the current selling price.
With these economical evaluations on how the main economical aspects vary with the price and units sold, it can be concluded that a selling price of 160.000,00 elays in a range where it gives the company a high enough NPV with, an acceptably low PBT, and an IRR higher than k.
RA 102 Chapter 5
Organization, planning and scheduling.
5.1 Gantt Diagram
Due to the complexity of the diagram and in order to keep the quality of the image, the Gantt diagram has been attached on the annexes in its original size. There it can be found the Gantt Annex A.4 that has been followed during the development of the project.
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RA 104 Chapter 6
Minutes of the Meeting
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MINUTES OF THE MEETING 01
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 20th September 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Presentation of the group members and selection of a coordinator and a secretary. Description of the Project Charter and its structure Presentation of the object of the project
ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter with the following content: Aim of the project Scope of the project Requirements of the project Justification 1
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State of art Group Organization o Organization Structure o Roles and responsibilities Planning o Tasks identification from Work Breakdown Structure (WBS) o Brief tasks description o Interdependency relationship among tasks o Human resources and level of effort (hours) to develop each task Budget (initial estimation)
DECISIONS MADE – ASSIGNMENTS: Presentation of the group members and selection of the main roles: o Coordinator: Ariadna Fernández, who will deal with the management tasks and will be responsible for the communications between the customer and the other group members. o Secretary: Itziar Ugartemendia, who will write up the minutes of the meetings and will be in charge of keeping BSCW updated and organized. The working environment will be BSCW. This platform allows organizing files in different folders and subfolders to classify the uploaded documents in the corresponding category. All the members of the group have access and can upload, download or modify the documents. The communication between the customer and the group members will be though ATENEA, where the final documents will be delivered. The communication tool among the group members will be Slack. This application allows having different communication groups. There will be a general group with all the members to discuss the global and most relevant aspects of the project. In addition, there will be a specific group for each department, where only the corresponding members will take part. The aim of the project is to design a Light Sport Aircraft (LSA) for 4 passengers, taking the European regulation into consideration. The next meeting will be the 27th September 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Definition of the aim and scope of the project o Planning of the project and tasks identification o Group organization in different departments
SIGNATURES Team coordinator Supervisor
2
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MINUTES OF THE MEETING 02
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 27th September 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Definition of the aim, scope and justification of the project Analysis and determination of the requirements of the project Group organization in different departments Discussion of the planning and tasks to be done by each department Schedule and time distribution of the tasks Revision of the Project Charter and determination of the remaining contents to be finished
ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter.
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DECISIONS MADE – ASSIGNMENTS: The project management application called Trello will be used to have a real-time organization of the tasks. It allows us to create different blocs, one for each department, and write down the tasks and subtasks to be done by each department in a list called “To do”. When a task is started, it is moved to the list called “In progress” until it is finished and moved to the list called “Verification”, where it is revised by two or three people (that haven’t worked in that task) and moved to the list called “Done”. In this way it is easy to see the evolution of the project and the remaining tasks to be done. Revision of different aspects of the Project Charter. There are some aspects to change: o The first step is the scope definition; afterwards, the tasks determination and the organization in departments; and finally the time estimation and schedule. o The scope defines the amount of work and the level of determination of the project, but it does not include the specific tasks in each department. The department organization is created afterwards according to the scope and the WBS. o The project must meet all the requirements of an LSA, except the number of passengers (three instead of two). Consequently, the aeroplane developed can’t be called LSA. o The State of Art contains the study of planes under 600Kg and planes for 3 passengers, but does not include the planes of two passengers. Discussion of new suggestions and resolution of questions, doubts or problems. The conclusions taken are the following: o The project will be focused on the development of a hybrid between a glider and an LSA in order to reduce the structural weight and increase the efficiency. o The technology used in the plane cannot be determined yet, but it will be considered whether to install Fly-by-wire in order to use electrical actuators instead of mechanical. The weight and the requirements of each configuration will be studied to determine the best option. Development of the BWS and the scope of the project: Definition of tasks and subtasks with a numerical code to organize them clearly. The team organization in departments will be the following:
Departments Members 1. Pol Bernad 2. Edgar Gago Aerodynamics team 3. Pau Nadal 4. Itziar Ugartemendia 1. Eduard Gómez Power plant team 2. Marcel Marín Technical 3. Pau Nadal 1. Xavi Carrillo Design team 2. Alejandro Sans 1. Pol Bernad 2. Ariadna Fernández Structures team 3. Carlos Méndez 4. Pau Nadal
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5. Alejandro Sans 1. Ariadna Fernández Systems team 2. Carlos Medina 3. Carlos Pérez 1. Eduard Gómez Economics team 2. David Rodríguez Non-technical 1. Alexandra Kalina Team management 2. David Rodríguez
It is important to remark that the distribution of the team members could be modified if necessary due to the amount of work or the difficulty of the tasks. There are also the following department coordinators, who are responsible for the coordination and communication among departments: o Aerodynamics coordinator: Pol Bernad o Structures coordinator: Carlos Méndez o Technical coordinator: Pau Nadal o Non-technical coordinator: David Rodríguez Finally, two more important roles, already defined in the first meeting, are: o Main coordinator: Ariadna Fernández o Secretary: Itziar Ugartemendia The next meeting will be the 1th October 2018, from 14:00 to 15:00 h, in a free classroom to be determined (ESEIAAT). Over these days, the following tasks will be carried out: o Correction of the aspects of the Project Charter previously described o Development of the last contents of the Project Charter in order to revise it together during the meeting.
SIGNATURES Team coordinator Supervisor
3
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MINUTES OF THE MEETING 03
Project: Design of a LSA for 3 passengers Participants: Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 1st October 2018, from 14:00 to 15:00 h Place: TR5-3.2 – ESEIAAT
AGENDA: Revision of the WBS Discussion of the remaining contents of the Project Charter Schedule and time distribution of the tasks
ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter.
DECISIONS MADE – ASSIGNMENTS: Once the WBS is finished and the list of tasks for each department is completed, it is time to assign the resources required to develop each task. It is quite difficult to determine how many hours each task will take. For this reason the method used is the following: the measure used instead of hours is the “Story Point” (SP), which considers both time and difficulty. One Story Point is equivalent to the work of a person in a morning/afternoon (that
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is, 2 or 3 hours). Each task can be defined by a multiple of 0,5*SP. In this way it is easy to count the number of Story Points available in order to set the deadlines for each task and create the Gantt Diagram. It will also help to distribute the tasks in an equitable and fair way among the group members. From the revision of the Project Charter the conclusions to be drawn are the following: o The remaining contents are: description of tasks; resources, time and preceding tasks assignment; estimation of the budget; references and bibliography. o The deadline set to conclude these contents is the 2nd October. The following day the Project Charter will be revised and ready to be delivered. The next meeting will be the 4th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Development of the Gantt Diagram and other remaining tasks. o Final revision and delivery of the Project Charter o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 112 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 04
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 4th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Analysis of the progress done and resolution of the problems Delivery of the provisional Project Charter Revision of the Project Charter with the supervisor and analysis of things to be changed Last corrections of the Project Charter
ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter Specific tasks for each department according to the Gantt Diagram
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RA 113 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: Delivery of the provisional Project Charter in BSCW according to the deadline. Revision of the Project Charter with the supervisor and discussion of the following aspects to be modified: o The State of Art: . It should not include specific aspects of our project . It should appear after the requirements, not after the aim . There are some changes that not appear in the last version o Scope: . There are some generic words that should be changed in order to specify more what is going to be done, for example, the words “design”, “analysis”, “study” should describe concretely what is included. Also, when it is said “main parts” it should be specified which parts. o Stakeholders: . It does not give much information in a Project Charter o WBS: . It would be better to turn the page in order to see it clearer . It is also possible to present it in a table o Gantt Diagram: . It must appear at the end of the document, not after the WBS, following the logic organization . It can be included a specific Gantt Diagram for each department to see the tasks clearer . The tasks of the point 7 are not included in the description of tasks o Definition of tasks: . It is very specific but that’s fine if we can deal with it o Design Budget: . We cannot include the annual licenses of Catia and Ansys, only partially because we are working only 4 months in this project . The computers could be include partially or consider that every team member uses his/her own one. . The necessary materials should also appear These aspects have been corrected and modified during the meeting. The next meeting will be the 11th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram. o Evaluate how is the group organization working and determine if it is necessary to make changes in any department
SIGNATURES Team coordinator Supervisor
2
RA 114 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 05
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 11th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram
DECISIONS MADE – ASSIGNMENTS: Develop the corresponding tasks for each department according to the schedule.
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RA 115 G06-AlOn LSA 3 seats | Project report
Each department has explained the progress done and the most important decisions made in order to have an overall view of the project. A brief summary of the tasks developed by each department is given next: o Aerodynamics: . Discussion of possible airfoils and analysis of different studies of airfoils and wings in order to define the basic configuration of the wing, which at first are: airfoil 230XX (the thickness could change according to future analysis); wing span, 16m; root chord, 0.8m; and tip chord, 0.6m. . Study of different tail configurations and selection of two possibilities: Conventional tail or T-tail. The final decision will depend on further analysis of the whole aircraft. . While developing the tasks according to the Gantt Diagram a problem has come up: two blocks of tasks (wing and tail) can’t be done at the same time by different members, as it was at first planned. The department has accorded to work together in both blocks and design all of them in order to study the aerodynamic forces and moments and the stability. o Structures: . Study and sizing of the landing gear according to the most commonly used in similar aircrafts. . Study of the possible materials for the construction of the aircraft and the regulations in this field. o Power Plant: . Study of the different engines and propellers available in the current market, estimation of the power required by the plane and selection of a few possible engines according to the results obtained. . Revision of the Certification Specifications for an LSA by ASTM and the specific regulations for EASA and FAA o Systems and Avionics: . Research about compulsory and complementary systems. . Research on possible control systems. o Economics: . Estimation of costs, including materials, facilities, human resources and other additional costs o Team management: . Coordination and supervision of the progress done. The next meeting will be the 18th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 116 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 06
Project: Design of a LSA 3 for passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 18th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation
1
RA 117 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . Last analysis of plane designs in Xflr5 in order to close the design as soon as possible and start the CAD design. The main wing and tail parameters are already fixed. o Structures: . Final decision of the materials that will be used. . Tests with the landing gear to decide the optimum configuration. o Power Plant: . The engine is already selected and the current work is based in the engine bench. . Development of a Matlab program to analyse the propeller in order to design it and evaluate its viability compared to the other propellers selected. o Systems and Avionics: . The work done by the each member has been joined. It has been discussed what is missing to finish the Flight and Navigation instruments part. . Contact with the Power Plant Department in order to start selecting the power plant instruments for next week. o Economics: . The estimation of the facilities’ and the human resources’ costs is already finished. . Estimation of the marketing campaign costs. . The SWOT analysis has been started. o Team management: . Coordination and supervision of the progress done. The next meeting will be the 18th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 118 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 07
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 25th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation
1
RA 119 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . There are two completed designs of airplane without control surfaces, which are the two possible options for the final plane: the first one with the three passengers located in a row, and the second one with two passengers in a line behind the pilot. In order to select one, the decision methods explained in class will be applied. . Once the final airplane design is selected, the definitive parameters and graphics will be exported. . The current work is focused on the control surfaces (sizing and location), the weight-range diagram and the parasite drag determination. o Structures: . The analysis of the landing gear has already been finished. . Start sizing the structure of the wings and fuselage o Power Plant: . Last steps in the study of the engine bench . Progress in the propeller design and selection of alternative propellers from the market, just in case that it is necessary. . Study of the exhaust fumes and the correct escapement o Systems and Avionics: . All the power plant instruments have been finished . The costs, weight and power of each instrument has been recorded in a table that will be useful for future tasks . The 2D design of the dashboard of the cockpit has been started o CAD design: . Design of the two alternatives for the plane. o Economics: . Finish the SWOT analysis. . Make progress in the estimation of the marketing campaign costs. o Team management: . Coordination and supervision of the progress done. The next meeting will be the 8th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 120 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 08
Project: Design of a LSA for 3 passengers Participants: Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 8th November 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation
DECISIONS MADE – ASSIGNMENTS: The report has been updated to include all the contents that should appear. This document contains all the necessary information about the design of the LSA: the procedures developed and the results obtained from the different analysis.
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RA 121 G06-AlOn LSA 3 seats | Project report
Every department is writing its corresponding part of the report while developing the tasks in order to record all the procedures, analysis, conclusions and decisions made. Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . From the final design it has been necessary to size the tail again and make it bigger to improve its performance. . The current work is focused on the sizing of the control surfaces, which involves the study of the stability and control of the airplane. Besides, some parameters from equivalent airplanes will be used as a first reference. o Structures: . The landing gear design has been finished. . For the fuselage sizing, the momentum equilibrium has been calculated. The same procedure will be applied for the wing and the tail. o Power Plant: . Progress in the design of the propeller in Matlab. . End of the tasks related to the engine mount and the exhaust fumes system. o Systems and Avionics: . All the systems are defined. The task remaining is the determination of the parameters of the controls of the pilot. Some data related to the control surfaces, which the current work of aerodynamics department is required, but it will be available in two or three days. o CAD design: . Make the necessary modifications to the design previously done according to the changes in some aspects of the structure and aerodynamics. o Economics: . Conclude the SWOT analysis . Close the estimation of the marketing campaign costs and the study of the future campaigns. . Begin the study of potential customers and the marketing study. o Team management: . Coordination and supervision of the progress done. The next meeting will be the 15th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 122 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 09
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 15th November 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 29th November: Delivery of a draft of the report. 20th December: Project Delivery and Final Presentation
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RA 123 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: Considering the deadline for the delivery of the draft of the report, one of the main tasks in each department is write all the information required in the document, adding graphs, figures and conclusions. Each department has made progress in their tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . The current work is focused on sizing the control surfaces, including flaps, ailerons and rudder. o Structures: . Structural analysis of the wing and tail separately. An assembly analysis will also be done. o Power Plant: . Selection of the engine and the engine bench. . The current work is based in the design of the propeller and the exhaust fumes system. o Systems and Avionics: . Selection of the most appropriate batteries. . Study of security components, such as fire extinguishers, and selection of the seat belts. . The next tasks are lighting, fuel tank and pilot joysticks for control surfaces. o CAD design: . Selection and design of the seats. . The files of the external part of the plane are ready to print and do the renders. The internal part is in process. . Calculation of the centre of gravity, for which the total mass of the instruments is needed and will be provided by the Systems Department. o Economics: . End the study of potential customers and the advertisement study. . Research on possible investors. . Design of a marketing campaign. . Start the payback analysis. o Team management: . Coordination and supervision of the progress done. The next meeting will be the 22th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.
SIGNATURES Team coordinator Supervisor
2
RA 124 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 10
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 22th November 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 29th November: Delivery of a draft of the report. 20th December: Project Delivery and Final Presentation
1
RA 125 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: General aspects of the oral presentation have been discussed. Each department has made progress in their tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . Design of the control surfaces. . Describe the final results and conclusions in the report. o Structures: . Building of the fuselage structure . Analysis of the wing and fuselage and joint analysis of the whole aircraft. o Power Plant: . Continue the propeller design and exhaust fumes system. . It remains an overall performance evaluation, but previously the propeller design must be finished. o Systems and Avionics: . The selection of batteries and the sizing of the electric system are finished. . The tasks related to the fuel tank are completed. . The compulsory illumination is also finished; and the night flight illumination has been studied and it is not possible to incorporate. . It remains to study of the pilot joysticks, but some data from the control surfaces is needed. o CAD design: . Refinement of the designed aircraft. o Economics: . Payback analysis, income vs. Time analysis o Team management: . Coordination and supervision of the progress done. The next meeting will be the 29th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram. o Try to finish the tasks related to the design and analysis of the aircraft in order to complete the draft of the report. o Start planning and developing the deliverable documents
SIGNATURES Team coordinator Supervisor
2
RA 126 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 11
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 29th November 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise the work done and discuss problems or doubts Delivery of the draft of the report in BSCW Update the Gantt Diagram and organize the following tasks to be done Develop the corresponding tasks for each department according to the schedule
ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation
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RA 127 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS: Some aspects of the contents of the report have been discussed. The Attachments will contain the detailed explanation of every part of the airplane, procedure, graphs, analysis, etc. The Report will be a summary of the necessary information considering that the reader should be able to understand the project without consulting the attachments. Hence, the report will mention every aspect of the final plane configuration (decisions made and method used) and will develop these ideas. The report has initially two main sections: Main alternatives and selection of the best one and Development and design of the chosen solution, but they will probably be merged in only one if necessary. Another important aspect is the fact that the report is thought to be read by engineers, so it should be technical and include de corresponding graphs and calculations. The last tasks have been finished in order to include them in the drafts of the report. Presentation of the drafts of the report by uploading both documents (Technical Report and Report Attachments) in BSCW to be revised by the supervisor. General aspects of the oral presentation have been discussed. Only three or four people will do the presentation, but one person of each department (5 people) will be especially devoted to organise the presentation. Meanwhile, the other team members will be more focused on the deliverable documents. The next meeting will be the 13th December 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Refinement and correction of the corresponding aspects in the report. o Development of the deliverable documents. o Planning and creation of the project Presentation.
SIGNATURES Team coordinator Supervisor
2
RA 128 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 12
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 13th December 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT
AGENDA: Revise and correct the Report and the Report Attachments. Make progress in the deliverable documents. Organise the presentation.
ITEMS ON THE AGENDA: 20th December: Project Delivery and Final Presentation
1
RA 129 G06-AlOn LSA 3 seats | Project report
DECISIONS MADE – ASSIGNMENTS:
Discussion with the project supervisor about some aspects of the conclusions of the project. The content should be not also general conclusions of the project (such as the difficulty of the project and how we have managed to achieve it, how satisfied we are about the results obtained...) but also opinions and comments about the organization, the new competences acquired, the experience of developing such type of project... Another important aspect to comment, although the coordinator of the subject already knows, is the fact that the theory classes are taught simultaneously to the development of the project, so sometimes we have manage aspects that we haven’t seen theoretically in class, such as the project charter or the decision making. During this week, every member of the group will evaluate the others and himself/herself in an Excel document sent by the supervisor. There are different criteria to evaluate with a mark from 1 (minimum) to 4 (maximum). In this session, the 4 members that will do the presentation have continued preparing it, while the others are focused on the deliverable documents and the revision of the report and other documents that are nearly finished. The next meeting will be the 18th December 2018, from 12:00 to 14:00 h, in the Conference Room (ESEIAAT). Over these days, the following tasks will be carried out: o Refinement and correction of all the documents. Everyone will read and revise the whole documents to ensure that all fits. o Correction of the style of the documents in order to standardize the format and style of graphs, images, references, etc. o Finish the presentation and start the rehearsals.
SIGNATURES Team coordinator Supervisor
2
RA 130 G06-AlOn LSA 3 seats | Project report
MINUTES OF THE MEETING 13
Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 19th December 2018, from 8:00 to 10:00 h Place: Classroom 3.5 (TR5) – ESEIAAT
AGENDA: Revision of all documents Rehearsal of the presentation
ITEMS ON THE AGENDA: 20th December: Project Delivery and Final Presentation
DECISIONS MADE – ASSIGNMENTS: General revision of documents and last changes, if necessary. Rehearsal of the presentation with all the team in a classroom with the same projector in order to prove that all slides are properly visible.
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RA 131 G06-AlOn LSA 3 seats | Project report
The full presentation has been done and some important aspects have been discussed: general ways to improve, aspects that should (not) be commented in order to adjust the time of the presentation, recommendations for the people that present, etc. Preparation for possible doubts and questions that could be asked after the presentation.
SIGNATURES Team coordinator Supervisor
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RA 132 Bibliography
[1] Michaels Selig, James J Guglielmo, Andy P Broeren, and Philippe Giguere. “Summary of Low-Speed Airfoil Data Summary of Low-Speed Airfoil Data”. In: 1 (1995), p. 315. ISSN: 1742-6596. DOI: 10.1115/1.1793208. URL: https://m-selig.ae.illinois.edu/ uiuc{\_}lsat/Low-Speed-Airfoil-Data-V5.pdf. [2] David Lednicer. The Incomplete Guide to Airfoil Usage. 2010. URL: https://m-selig.ae. illinois.edu/ads/aircraft.html (visited on 10/12/2018). [3] Harrey Riblett. Ga-airfoils. Sixth Edif. Harry C. Riblett, 1996, p. 142. [4] ASTM International. Standard Specification for Design and Performance of a Light Sport Air- plane. Tech. rep. ASTM International, 2006. URL: https://www.astm.org/Standards/ F2245.htm. [5] Propiedades del Aluminio — Universidad de Cádiz. URL: http://tablaperiodica.uca.es/ Tabla/elementos/Aluminio/Grupo1/Prop.Al (visited on 12/13/2018). [6] Adnene Tlili and Sofiene Bouhjar. “Performance Study of a Metal Matrix Composite Alloy for Aircraft Industry Use”. In: June (2015). DOI: 10.13140/RG.2.1.2292.3360. [7] Nomex aramid honeycomb 2mm. URL: http : / / www . fibermaxcomposites . com / shop / nomex-aramid-honeycombbrthickness-mmbrcell-size-32-mm-p-962.html (visited on 12/15/2018). [8] CÁTALOGO DE PERFILES DE ALUMINIO NORMALIZADOS ALUMINIUM STAN- DARD PROFILES CATALOGUE. Tech. rep. URL: http://www.extrusax.com/imagenes/ descargas/es/12/STANDARDPROFILES-PERFILESNORMALIZADOS.pdf. [9] Advisory Circular TITLE 14 OF THE CODE OF FEDERAL REGULATIONS (14 CFR) GUIDANCE MATERIAL. Tech. rep. URL: https://www.faa.gov/documentLibrary/ media/Advisory{\_}Circular/AC{\_}43.13-1B{\_}w-chg1.pdf. [10] Aviation Coatings for Corrosion Prevention. URL: https://www.corrosionpedia.com/2/ 1823/industries/transportation/aviation-coatings-for-corrosion-prevention (visited on 12/13/2018). [11] Powered Sailplanes. “Standard Specification for Design and Performance of a Light Sport Airplane 1”. In: (). [12] Libice Nad Cidlinou. “ATEC 322 FAETA Flight and Operations Manual”. In: March (2013), pp. 1–54. URL: http://www.atecaircraft.be/dossiers/Manuals/flight- manual-atec-322-faeta.pdf. [13] Ultralight. “TL-2000 Sting S4 PILOT´S OPERATING HANDBOOK”. [14] Aircraft Maintenance Manual. “TL-3000 Sirius”. In: (). [15] Pipistrel Glider Taurus M l Gliding - Pipistrel Andorra. 2003. [16] Alexander Schleichers. “ASG 29 E. Sneak into the Design Process”. [17] Strength Tests of the plane - TomarkAero | Production of UL ultralight / LSA Light sport air- craft Viper SD4 and Skyper GT9. URL: http://www.tomarkaero.com/en/manufacturing/ strength-tests.html (visited on 10/13/2018).
RA 133 G06-AlOn LSA 3 seats | Project report [18] ASTM International. “F2245-15: Standard Specification for Design and Performance of a Light Sport Airplane”. In: (2016), pp. 1–30. DOI: 10.1520/F2245-12D.2. URL: http: //compass.astm.org.proxy.library.carleton.ca/download/F2245.1538.pdf. [19] STRIAN—Structural analysis. URL: http : / / structural - analyser . com/ (visited on 12/15/2018). [20] Chun-yun. Niu. Composite airframe structures : practical design information and data. 1st Ed. Conmilit Press, 1992, p. 664. ISBN: 9627128066. [21] Michael Chun-Yung Niu. Airframe : stress analysis and sizing / Michael Chun-Yung Niu. Dragon Terrance, North Point : Hong Kong Conmilt Press, 1999. ISBN: 9627128082. [22] Stavros Salampoukidis. 3 Blade propeller. 2016. URL: https://grabcad.com/library/3- blade-propeller-2. [23] Seats & Seat Covers - Air-Tech Inc. URL: https://air-techinc.com/topic{\_}std{\_ }prods.php?catid=196{\&}pmid=18 (visited on 11/22/2018). [24] Sport Aircraft Seat Company. URL: http://www.sportaircraftseats.com/sportaircraftseats/ Home.html (visited on 11/22/2018). [25] La oficina ideal: 14m 2 por empleado | Pyme | Cinco Días. URL: https : / / cincodias . elpais.com/cincodias/2014/10/28/pyme/1414500383{\_}553511.html (visited on 12/11/2018). [26] Easa - Rps. AMC and GM to Part 21 Acceptable Means of Compliance and Guidance Material. Tech. rep. 2012. URL: https://www.easa.europa.eu/sites/default/files/dfu/ AnnexItoEDDecision2012-020-R.pdf. [27] Susan E. Jackson, Aparna Joshi, and Niclas L. Erhardt. “Recent Research on Team and Organizational Diversity: SWOT Analysis and Implications”. In: Journal of Management 29.6 (2003), pp. 801–830. ISSN: 0149-2063. DOI: 10.1016/S0149-2063(03)00080-1. URL: https://www.sciencedirect.com/science/article/pii/S0149206303000801. [28] Seyed J. Sadjadi, Maryam Oroujee, and Mir. B. Aryanezhad. “Optimal Production and Marketing Planning”. In: Computational Optimization and Applications 30.2 (2005), pp. 195– 203. ISSN: 0926-6003. DOI: 10.1007/s10589-005-4564-8. URL: http://link.springer. com/10.1007/s10589-005-4564-8. [29] 2017 ANNUAL REPORT General Aviation Manufacturers Association. Tech. rep. URL: www. wfscorp.com. [30] General-Aviation-Manufacturers Association. 2017 ANNUAL REPORT General Aviation Manufacturers Association. Tech. rep. URL: www.wfscorp.com. [31] EASA Anual Safety Review 2017. Tech. rep. URL: https://www.easa.europa.eu/sites/ default/files/dfu/209735_EASA_ASR_MAIN_REPORT_3.0.pdf. [32] Air Charter Guide. URL: http://aircharterguide.com/Product.aspx?Prod=acg (vis- ited on 11/15/2018). [33] IATA. IATA - Advertising. URL: https://www.iata.org/services/advertising/Pages/ index.aspx (visited on 11/15/2018). [34] Flying Magazine. Advertising Specs FLYING MAGAZINE. Tech. rep. URL: http://www. google.com. [35] “Print Specificaions for Flying Magazine”. In: (). URL: http://www.bonniermarinegroup. com/files/{\_}attachments/media{\_}kit/flying{\_}2018{\_}print{\_}specs. pdf. [36] Perfect Landing Media, LLC. - Aviation Marketing. URL: https://www.perfectlandingmedia. com/our-services/aviation-marketing (visited on 11/15/2018). [37] The Importance of Angel Investing in Financing the Growth of Entrepreneurial Ventures. Sba.gov. URL: http://www.sba.gov/advo/research/rs331tot.pdf.
RA 134 G06-AlOn LSA 3 seats | Project report
[38] A Guide to Angel Investors. Entrepreneur. URL: http://www.entrepreneur.com/article/ 52742. [39] 14 business angels in Spain that startups CEOs can’t miss. URL: https://startupxplore. com/en/blog/business-angels-in-spain/ (visited on 11/22/2018).
RA 135 G06-AlOn LSA 3 seats | Project report
RA 136 Appendix A
Code
The following code has been use to develop several parts of the project.
A.1 Weigth-Range Diagram
1 % Data in order to create the weight-range diagram 2 clear all ; 3 clc; 4 5 %Plane data 6 TF = 46 ;%Kg trip fuel 7 RF = 4;% Kg --> reserve fuel 8 FW = TF+RF;% FUEL WEIGHT 9 E = 35 ;%Efficiency for alpha0 10 Sw = 12.424;%[m^2] 11 CL = 0.6 ; 12 rho = 1.225 ;% needs to be re-calculated 13 PL_max = 70 *3; 14 Pl_min = 70; 15 Pl_2 = 70*2; 16 OEW = 340; 17 18 MZFW = OEW + PL_max; 19 ZFW = OEW + Pl_min; 20 ZFW2 = OEW + Pl_2; 21 22 MZFW_R = MZFW + RF; 23 ZFW_R = ZFW + RF; 24 ZFW_2R = ZFW2 + RF; 25 26 %Engine data 27 Rend= 0.85;% average 28 Ce = 75e-8;% micro-grams/J 29 C= 12;%l\h 30 g = 9.81; 31 32 %Configurations 33 MTOW = OEW + PL_max + FW; 34 MLW = OEW + PL_max + RF; 35 36 TOW = OEW + Pl_2 + FW; 37 LW2 = OEW + Pl_min + + RF; 38 LW = OEW + 2*70 + RF; 39 40 %K of breguet 41 K = 2*(Rend * E) / (Ce * g) * sqrt(2 /(CL*Sw*rho)); 42 43 R_MPL = K*log(MTOW/MLW);
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44 R_max = K*log(TOW/LW2); 45 R_MTOW = K*log(MTOW/LW); 46 47 y = [MZFW_R MTOW MTOW TOW]; 48 x = [0 R_MPL R_MTOW R_max]; 49 x=x./1000; 50 OEW_V = [OEW OEW OEW OEW]; 51 MZFW_V = [MZFW MZFW ZFW2 ZFW]; 52 MZFW_RV = [MZFW_R MZFW_R ZFW_2R ZFW_R]; 53 54 55 plot(x,y,'r',x,OEW_V,'g',x,MZFW_V,'b-',x,MZFW_RV,'c'); 56 title('Weight-Range Diagram','Interpreter','Latex'); 57 ylabel('Weight(kg)','Interpreter','Latex'); 58 xlabel('Range(km)','Interpreter','Latex'); 59 %axis([0 95000 600]); 60 legend('MTOW','OEW','MZFW','MZFW+RF','Location','SouthEast');
A.2 Gust-Airspeed envelope
1 clc; 2 clear; 3 close all; 4 5 %% Flight envelope 6 7 %Factores de carga por normativa(LSA) 8 n1 = 4; 9 n2 = -2; 10 11 %% velocidades[kts] para factor de carga4 12 Vs1 = 45.40 ; 13 Va1 = 90.8; 14 Vc = 110; 15 Vd = 160; 16 17 %Ajuste de par bolas 18 X1 = [0 Vs1 Va1]; 19 Y1 = [0 1 4]; 20 [Fit1, gof1] = CurveFitting(X1, Y1); 21 22 %Rectas 23 Rx1=[Va1 Vc Vd]; 24 Ry1=[4 4 4]; 25 26 Vx1=[Vd Vd]; 27 Vy1=[0 4]; 28 29 %% Velocidades para factor de carga -2; 30 Vs2 = 54; 31 Va2 = 76.45; 32 33 %Ajuste de par bolas 34 X2 = [0 Vs2 Va2]; 35 Y2 = [0 -1 -2]; 36 [Fit2, gof2] = CurveFitting(X2, Y2); 37 38 %Rectas 39 Rx2=[Va2 Vd ]; 40 Ry2=[-2 -2]; 41 42 Dx2=[Vd Vd];
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43 Dy2=[0 -2]; 44 %[Line2, gof]= LineFitting(Dx2, Dy2); 45 46 %% Flapn= 2; 47 Vaf = 61.8; 48 Vdf =85; 49 50 Xf = [0 42 Vaf]; 51 Yf = [0 1 2]; 52 [Fitf, goff] = CurveFitting(Xf, Yf); 53 54 %Rectas 55 Rxf=[Vaf Vdf]; 56 Rx3f=[0 Vdf]; 57 Ryf=[2 2]; 58 Ry3f=[0 0]; 59 60 Vxf=[Vdf Vdf]; 61 Vyf=[0 2]; 62 63 %% Gust envelope 64 n3=[1 5.4]; 65 n4=[1 -3]; 66 n3_1=[5.4 4]; 67 n4_1=[-3 -2]; 68 n5=[1 4]; 69 n6=[1 -2]; 70 x=[0 Vc]; 71 x2=[0 Vd]; 72 x1=[Vc Vd]; 73 74 %%Reference lines 75 vs1_v=[Vs1 Vs1]; 76 vs2_v=[Vs2 Vs2]; 77 vc_v=[Vc Vc]; 78 va1_v=[Va1 Va1]; 79 vaf_v=[Vaf Vaf]; 80 vdf_v=[Vdf Vdf]; 81 y=[-3 5.4]; 82 83 %% Plotting th fligth envelope 84 figure; 85 %n=4; 86 plot(Fit1,'b',X1,Y1); 87 hold on; 88 plot(Rx1,Ry1,'b'); 89 plot(Vx1,Vy1,'b'); 90 %n=-2; 91 plot(Fit2,'b',X2,Y2); 92 plot(Rx2,Ry2,'b'); 93 plot(Dx2,Dy2,'b'); 94 %n=2 --> flap 95 plot(Fitf,'r',Xf,Yf); 96 plot(Rxf,Ryf,'r'); 97 plot(Rx3f,Ry3f,'r'); 98 plot(Vxf,Vyf,'r'); 99 % gust envelope 100 plot(x,n3,'--g'); 101 plot(x,n4,'--g'); 102 plot(x1,n4_1,'--g'); 103 plot(x1,n3_1,'--g'); 104 plot(x2,n5,'--g'); 105 plot(x2,n6,'--g'); 106 %Reference velocities 107 plot(vs1_v,y,'--k'); 108 plot(vs2_v,y,'--k');
RA 139 G06-AlOn LSA 3 seats | Project report
109 plot(vc_v,y,'--k'); 110 plot(va1_v,y,'--k'); 111 plot(vaf_v,y,'--k'); 112 plot(vdf_v,y,'--k'); 113 114 axis([0 Vd+10 -3.5 5.5]); 115 grid on; 116 title('Airspeed\& Gust envelope','Interpreter','Latex'); 117 ylabel('Load factorn[]','Interpreter','Latex'); 118 xlabel('Airspeed[kts]','Interpreter','Latex'); 119 legend off;
A.2.1 Second grade polynomial fitting
1 function [fitresult, gof] = CurveFitting(X1, Y1) 2 %% Fit:'untitled fit 1'. 3 [xData, yData] = prepareCurveData( X1, Y1 ); 4 5 % Set up fittype and options. 6 ft = fittype('poly2'); 7 8 % Fit model to data. 9 [fitresult, gof] = fit( xData, yData, ft ); 10 11 % Plot fit with data. 12 % figure('Name','untitled fit 1'); 13 %h= plot( fitresult, xData, yData); 14 % legend(h,'Y1 vs. X1','untitled fit 1','Location','NorthEast'); 15 %% Label axes 16 % xlabel X1 17 % ylabel Y1 18 end
A.2.2 First grade polynomial fitting
1 function [fitresult, gof] = LineFitting(Dx2, Dy2) 2 %% Fit:'untitled fit 1'. 3 [xData, yData] = prepareCurveData( Dx2, Dy2 ); 4 5 % Set up fittype and options. 6 ft = fittype('poly1'); 7 8 % Fit model to data. 9 [fitresult, gof] = fit( xData, yData, ft ); 10 11 %% Plot fit with data. 12 % figure('Name','untitled fit 1'); 13 %h= plot( fitresult, xData, yData); 14 % legend(h,'Dy2 vs. Dx2','untitled fit 1','Location','NorthEast'); 15 %% Label axes 16 % xlabel Dx2 17 % ylabel Dy2 18 end
A.3 Propeller Design
The following code has been used to design the propeller. Not all code is presented, since most of it deals with making plots and similar tangential functions. Only the chore of the
RA 140 G06-AlOn LSA 3 seats | Project report computation is shown, as well as the variable declaration so as to show the input variables.
A.3.1 Variable declaration
1 % Solver setup 2 n_el = 100;%number of elements in blade 3 4 % Physical info 5 U = 56.5;% Speed[m/s] 6 height = 2000;% Altitude[m] 7 R = 0.72;% Radius[m] 8 N = linspace(1600,2300,501);% Rotation(rpm)(max 2400 rpm) 9 nb = 2;% Number of blades 10 r_min = 0.1;% Radius of central cone 11 AirfoilFile ='NACA4412.csv';% Airfoil 12 theta0 = 0.3;% Root torsion 13 chord0 = 0.12;% Root Chord
A.3.2 Core function This function is sufficient to obtained the desired results, and its inputs are those declared in the previus snippet of code.
1 function [T,P_consumed, P_useful, efficiency] = ... propeller(n_el,U,height,R,omega,nb,r_min,AirfoilFile,theta0,chord0) 2 [Cl_mat,Cd_mat] = importAirfoil(AirfoilFile); 3 [dr,r, Theta, Chord, Sigma] = discretization(nb, n_el, r_min, R,theta0, ... chord0); 4 [lambda_forwards, lambda_sound] = nonDymensionalization(U,omega,R,height); 5 [Lambda_induced] = ... getLambdaInduced(n_el,lambda_forwards,lambda_sound,r,Theta,Sigma,Cl_mat,Cd_mat); 6 [T,M] = getSpecs(n_el,nb,r,dr,R,Theta,Chord, omega, Lambda_induced, ... lambda_forwards, lambda_sound,height, Cl_mat, Cd_mat); 7 [P_consumed, P_useful, efficiency] = Energy(T,M,U,omega); 8 end
In here we see that several functions are sequentially called, this was done to keep the code tidy and errors better localized.
A.3.3 importAirfoil Tihs function imports the angle of attack vs. lift and drag coefficient curves from any csv file.
1 function [Cl_mat,Cd_mat] = importAirfoil(AirfoilFile) 2 sourceFile = readtable(AirfoilFile); 3 4 %Airfoil data 5 Polar(:,1) = table2array(sourceFile(:,1))*pi/180; 6 Polar(:,2) = table2array(sourceFile(:,2)); 7 Polar(:,3) = table2array(sourceFile(:,3)); 8 9 Polar_extra = [%Approximated values for large AoA 10 -pi/2 0 0.5; 11 -3/4*pi 0.5 0.2; 12 pi 0 0.1; 13 3/4*pi -.5 0.2; 14 pi/2 0 0.5]; 15
RA 141 G06-AlOn LSA 3 seats | Project report
16 Polar = [Polar; Polar_extra]; 17 sort(Polar); 18 19 Cl_mat = get_coefMat(Polar(:,1),Polar(:,2)); 20 Cd_mat = get_coefMat(Polar(:,1),Polar(:,3)); 21 end
The get_coefMat function is used to interpolate the curves more precisely, with cubic polyno- mial interpolation instead of linear interpolation. This is also necessary for the solver, since it uses a method that requires the function to be differentiable everywhere (which this method guarantees but linear interpolation does not).
A.3.4 discretization This function creates some variables necessary to work with blade elements.
1 function [dr,r, Theta, Chord, Sigma] = discretization(nb, n_el, r_min, R, ... theta0, chord0) 2 dr = (1-r_min/R)/(n_el); 3 r = r_min/R + dr/2 : dr : 1 - dr/2; 4 Theta = thetaDistribution(r,theta0); 5 Chord = chordDistribution(r,chord0); 6 Sigma = Chord*nb / (pi * R); 7 end
We see two functions being called here. These define the geometry of the blade regarding torsion and chord as a function of the distance from the center.
1 function [ Theta ] = thetaDistribution(x, theta0) 2 Theta = theta0./x; 3 end
1 function [ Chord ] = chordDistribution(x, chord0) 2 Chord = chord0*(1-x/2); 3 end
A.3.5 nonDymensionalization This function creates some nonDymensional variables. Working with nondymensional mag- nitudes simplifies the code.
1 function [lambda_avance, lambda_sonido] = ... nonDymensionalization(U,omega,R,height) 2 lambda_avance = U/(omega*R); 3 [T,¬,¬] = atmosphere(height); 4 lambda_sonido = sqrt(1.4*287*T)/(omega*R); 5 end
A.3.6 getLambdaInduced This function is the heart of the program. It has only one output: Lambda_induced. This is an array which equals the nondymensional downwards airspeed caused by the propeller at each distance form the center. This is essential to obtain the angle of attack of each blade element.
RA 142 G06-AlOn LSA 3 seats | Project report The objective is to solve for lambda induced in the blade element equation at each blade element: 2 2 8(λ f + λi)λir = (r + (λ f + λi) )(Clcos(φ) − Cdsin(φ)) (A.1) To do so an initial lambda induced is proposed and both sides of the equation calculated. The two are subtracted and, if the proposed value for lambda induced was correct the result will be zero. Otherwise, we’ll obtain some non-zero value used as a measure of the error. This is done in the BEMequation function. The objective is, therefore, to minimize this error.
2 2 Error = |8(λ f + λi)λir − (r + (λ f + λi) )(Clcos(φ) − Cdsin(φ))| (A.2)
To do so a method called gradient descent is used. This method consists in calculating the derivative of some unknown function (in this case error as a function of lambda induced), and to push the independent variable towards the direction where its value decreases (increased when the derivative is negative, decreased otherwise; in proportion to the magnitude of this derivative).
Because calculating a derivative with solely two data points is problematic (heavily distorted by noise), three of them are obtained and the minimum value is used. Among the methods tested, this was the quickest to converge.
Gradient descent is performed on the Error function until it reaches a threshold (in this case a 0.1% error). Because Lambda_induced(r) is a continuous function, we can use the value ob- tained in a blade element to initialize the value for the following element, thus propagating the results forwards and considerably reducing the time to converge.
1 function [Lambda_induced] = ... getLambdaInduced(n_el,lambda_forwards,lambda_sound,r,Theta,Sigma,Cl_mat,Cd_mat) 2 %Lambda is calculated witha gradient descent algorithm 3 Lambda_induced = zeros(1,n_el);%Starting value 4 dLi = 1E-4;%Size of the step to compute derivative 5 steps = zeros(1,n_el);%Number of steps per blade element(for debugging) 6 for e=1:n_el%Algorithm ran per each blade element 7 for i=1:1E4 8 %A first attempt at guessing value is computed 9 [error] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e), ... lambda_forwards,lambda_sound,Cl_mat,Cd_mat); 10 11 if abs(error)<1E-3 12 %If the error is within bounds, lambda_induced is propagated to ... the next element and the program moves on 13 if(e6=n_el) 14 Lambda_induced(e+1) = Lambda_induced(e); 15 end 16 steps(e) = i; 17 break 18 else%If the error is too high, we slightly change the value of ... lambda_induced to aproximate the derivative 19 [error0] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e)-dLi,lambda_forwards, ... lambda_sound, Cl_mat,Cd_mat); 20 [error1] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e)+dLi,lambda_forwards, ... lambda_sound, Cl_mat,Cd_mat); 21 22 %The derivative is calculated thrice to smooth out the noise 23 dE_dLi(1) = (abs(error) - abs(error0)) / dLi; 24 dE_dLi(2) = (abs(error1) - abs(error)) / dLi; 25 dE_dLi(3) = (abs(error1) - abs(error0)) /(2*dLi);
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26 27 [¬,j] = min(abs(dE_dLi));%The most conservative value is ... chosen 28 k = dE_dLi(j)*dLi; 29 Lambda_induced(e) = Lambda_induced(e) - k;%The program ... descends down the error function 30 end 31 end 32 end 33 end
The function that obtains the Error is the following:
1 function [error] = BEMequation(r,Theta, ... Sigma,Lambda_induced,lambda_forwards,lambda_sound,Cl_mat,Cd_mat) 2 %Obtaining relevant angles 3 Phi = atan((Lambda_induced + lambda_forwards)/r); 4 Alpha = Theta - Phi; 5 6 %Obtaining compressibility correction factor 7 Mach = sqrt((Lambda_induced + lambda_forwards)^2 + r^2)/lambda_sound; 8 9 %Obtaining dC_Fz(e) 10 Cl = eval_coefMat(Alpha, Cl_mat) / sqrt( 1 - Mach^2); 11 Cd = eval_coefMat(Alpha, Cd_mat); 12 dC_Fz = Cl*cos(Phi) - Cd*sin(Phi); 13 14 %BEM theory tells us that F1= F2: 15 F1 = (r^2+ (lambda_forwards + Lambda_induced)^2)*dC_Fz.*Sigma; 16 F2 = 8*(lambda_forwards + Lambda_induced)*Lambda_induced*r; 17 error = F1 - F2; 18 end
A.3.7 getSpecs Once the airflow is well defined, this function obtains the the thrust(T) and torque on the shaft (M).
1 function [T,M,Alpha] = getSpecs(n_el,nb,r,dr,R,Theta,Chord, omega, ... Lambda_induced, lambda_forwards, lambda_sound, height, Cl_mat, Cd_mat) 2 %Computation of relevant angles 3 Phi = atan((Lambda_induced + lambda_forwards)./r); 4 Alpha = Theta - Phi; 5 6 %Computation of aerodynamic coefficients 7 Cl = zeros(1,n_el); 8 Cd = zeros(1,n_el); 9 for e=1:n_el 10 Cl(e) = eval_coefMat(Alpha(e),Cl_mat); 11 Cd(e) = eval_coefMat(Alpha(e),Cd_mat); 12 end 13 14 %Glauert compressibility factor 15 Mach = sqrt((Lambda_induced + lambda_forwards).^2 + r.^2)/lambda_sound; 16 if max(Mach) > 1 17 error('Supersonic flow achieved'); 18 end 19 Cl = Cl./sqrt(1 - Mach.^2); 20 21 E = Cl./Cd; 22 23 %Nondymensional forces 24 dC_Fz = Cl.*cos(Phi) - Cd.*sin(Phi);%Vertical force
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25 dC_Fx = Cl.*sin(Phi) + Cd.*cos(Phi);%Resistive forces 26 27 %Computation of speed and density to re-dymesionalize 28 [¬,¬,rho] = atmosphere(height); 29 V = sqrt(((Lambda_induced + lambda_forwards).^2 + r.^2)) * omega.*r; 30 31 %Computation of Thrust(T) and Torque(M) gradients along the radius 32 dFz = 0.5*rho*V.^2.*Chord .*dr .* dC_Fz; 33 dMz = 0.5*rho*V.^2.*Chord .*dr .* dC_Fx .*r; 34 35 %Integration of gradients 36 T = nb*sum(dFz); 37 M = nb*sum(dMz); 38 end
A.3.8 Energy This function calculates power used and consumed, as well as its ratio (the efficiency).
1 function [P_consumed, P_useful, efficiency] = Energy(T,M,U,omega) 2 P_consumed = M*omega; 3 P_useful = T*U; 4 if P_consumed<0 || P_useful<0 5 efficiency = 0; 6 else 7 efficiency = P_useful / P_consumed; 8 end 9 end
A.3.9 Other secondary functions The outputs of the last two functions are enough for the core function to return what it was re- quested to: Thrust, Used power and Power consumed. The program can therefore end here. However, there are some functions that are more tangential to the topic (yet still essential to the program). These are exposed here.
The atmosphere function obtains various air properties as a function of height, according to the International Standard Atmosphere.
1 function [T,P,rho] = atmosphere(h)%h in meters 2 T = 288.15 - 0.0065*h;%T inK 3 P = 101325 * (T/288.15)^-5.2586;%P in Pa 4 rho = 1.225*(T/288.15)^-6.2586;% rho in kg/m3 5 end
The following two functions are used to interpolate using third degree polynomials. Matlab has a built-in function already, but one better suited for periodic functions was written. They both are heavily commented so they are not further detailed.
1 function [coefMat] = get_coefMat(x,y) 2 % This funcion has as inputs: 3 %-x is any angle between-infinity and+infinity. 4 %-y the dependent variable ofa periodic functionf(x) with period2pi 5 % 6 % There is just one output 7 %- An array with the following columns: 8 % X0|A|B|C|D 9 % --> such thaty=A+ Bx+ Cx^2+ Dx^3 within the domainx(i):x(i+1) 10 % --> X0 isx(i), to map each polynomial to its lower bound
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11 % --> The last element wraps around the first one, so the 12 % properties of continuity and differntiablity are mantained for the 13 % entirety of the period 14 % 15 % To use this output array one must use the function 16 %- eval_coefMat(X,coefMat) 17 positions = [mod(x,2*pi), y]; 18 19 %This prelimnary array is sorted alongx: 20 positions = sortrows(positions); 21 x = positions(:,1)'; 22 y = positions(:,2)'; 23 24 % Obtaining gradients at lower and upper bounds: 25 dydx = zeros(1,length(x)); 26 for i=2:length(x)-1 27 dydx(i) = (y(i+1) - y(i-1))/(x(i+1) - x(i-1)); 28 end 29 dydx(1) = (y(2) - y(end))/(x(2)+2*pi - x(end)); 30 dydx(end) = (y(1) - y(end-1))/(x(1)+2*pi - x(end-1)); 31 32 % Coefficients obtained fromy and dy/dx at x_lowerBound and x_upperBound 33 coefMat = zeros(length(x),5); 34 for i=1:size(coefMat,1)-1 35 coefMat(i,:) = get_coefRow(x(i),x(i+1),y(i),y(i+1),dydx(i),dydx(i+1)); 36 end 37 coefMat(end,:) = ... get_coefRow(x(end),x(1)+2*pi,y(end),y(1),dydx(end),dydx(1)); 38 end 39 40 41 function [coefRow]=get_coefRow(X1,X2,Y1,Y2,dY1,dY2) 42 %This function adjusts the polynomialy=A+ Bx+ Cx^2+ Dx^3 43 % such that it passes through(X1,Y1) and(X2,Y2) with derivatives 44 % dY1 dY2 at these points 45 if X1 == X2 46 coefRow = [X1, (Y1+Y2)/2, 0, 0, 0]; 47 else 48 Xdif = X2-X1; 49 % The equations to adjust this can become quite cumbersome, but in 50 % matrix form they are somewhat more bearable 51 A = 1/Xdif^2*[ 52 (X2^3 - 3*X2^2*X1)/Xdif -X2^2*X1 (3*X2*X1^2 - X1^3)/Xdif ... -X2*X1^2; 53 6*X1*X2/Xdif X2^2+2*X1*X2 -6*X1*X2/Xdif ... X1^2+2*X1*X2; 54 -3*(X1+X2)/Xdif -2*X2 - X1 3*(X1+X2)/Xdif ... -X2-2*X1; 55 2/Xdif 1 -2/Xdif ... 1; 56 ]; 57 B = [Y1, dY1, Y2, dY2]'; 58 Polynomial = A*B; 59 coefRow = [X1, Polynomial']; 60 end 61 end
Now the one that used the output:
1 function [coef] = eval_coefMat(x, coefMat) 2 % This function recieves two inputs: 3 %-x independent variable 4 %- coefMat is the matrix generated by get_coefMat that contains the 5 % polynomial coefficients 6 %
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7 % There's only one output: 8 %-y is the dependent variable 9 % 10 11 x = mod(x,2*pi); 12 13 if x
A.4 Gantt diagram
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