EP-UAV Project 2009

iSpy We watch over you…

Kate Rietdyk

Komal Sidhu

Lim Ee Wei

Tanmay Bhat

Abhiram Ramesh

Executive Summary 1

Design Requirements and Objectives 2

Conceptual Design 3

Synthesis of concept 3.1

Configuration design 3.1.1

Initial Sizing 3.1.2

Fuselage layout 3.1.3

Airfoil design 3.1.4

Wing planform design 3.1.5

Empennage design 3.1.6

Geometry for landing gear 3.1.7

Control surface sizing 3.1.8

Pod Layout 3.1.9

Analysis of concept 3.2

Analysis of propulsion system 3.2.1

Weight estimation and Centre of Gravity (cg) estimation 3.2.2

Stability and Aerodynamic performance analysis 3.2.3

Aircraft performance analysis 3.2.4

CAD definition of the concept 3.2.5

Preliminary and Detailed Design 4

Structure layout, initial sizing, and internal layout Design 4.1

Wing structure 4.1.1

Fuselage structure 4.1.2

Empennage structure 4.1.3

Control surface structure 4.1.4

Landing gear 4.1.5

Integration of propulsion and control system 4.1.6

Pod Structure 4.1.7

Payload arrangement 4.1.8

Structural analysis 4.2

Fabrication 5

Preparation for materials 5.1

Fuselage fabrication 5.2

Wing fabrication (inner section) 5.3

Wing fabrication (outer section) 5.4

Empennage fabrication (inner section) 5.5

Landing gear fabrication 5.6

Installation propulsion system in to airframe and test 5.7

Control system installation and test 5.8

Control surface fabrication 5.9

Assembly and test 5.10

Measurement of Weight and CG location 5.11

Tests 6

Ground tests 6.1

Ground test 1 6.1.1

Ground test 2 6.1.2

Flight tests 6.2

Air test 1 6.2.1

Air test 2 6.2.2

Air test 3 6.2.3

Appendix 7

Appendix A – AVL files

Appendix B – Airfoil data

Appendix C – Preparation of materials

Appendix D – Control surface sizing

Appendix E – Sketches from Team notebook

Executive Summary 1

iSpy is a twin fuselage EP-UAV powered by two electric motors and two 4-cell batteries

with aerial photography and video capturing capabilities from a range of altitudes. The

design requirements set out by the course outline was set as the minimum performance

requirement for this project.

The initial design ideas that were discussed involved various configurations, which

ranged from complicated, challenging concepts to extreme simplicity which involved

considering nothing outside the course requirement for this project. Some of these

sketches are shown below in figures 1.01 to 1.06

Figure 1.01 Figure 1.02

Some of these designs were scratched at initial discussions itself due to their lack of

ingenuity or design brilliance. Sketches shown in figures 1.04 and 1.05 are examples of

these. Others were left out due to the lack of time, money and man hours required to

complete this project, such as the design shown in figure 1.06.

The final design that was chosen, however, was the idea based on creating the Virgin

Galactic’s Global Flyer replica. The idea blossomed with discussion and soon the basic

sketches grew into many drawings with minor details and the final decision was made.

A twin fuselage design with a high aspect ratio wing and an H-tail that can carry a pod in

the middle that can be modified to be used as spy plane with an onboard camera was

the team’s favourite. The payload may also be this pod with additional space within to

hold metal blocks or sand bags.

Figure 1.03 Figure 1.04

The aircraft design phases went fairly smooth and there were no issues that could not

be solved with advice from professors and senior students. Analyses on all aspects were

conducted in various ways as progress into the design phases continued using

numerical methods during initial stages and then more complex methods were used.

CATIA was a major component during design and it was also used to estimate weights

and CG position. Applications such as AVL had to be used to estimate the aircraft’s

stability and aerodynamic performance in later stages of design where equations were

not satisfactory due to design complexities.

Fabrication of the aircraft was done in a shorter time frame than design itself but

workload was manageable and the results were rewarding. A few obstacles we did

come across were easily overcome with determination and hard work by the team.

Weight estimations did not pan out as it was initially calculated and some errors in

calculations were later brought to our attention. Landing gears gave some trouble

during tests due to lack of strength and stability of the front right wheel but it was not a

major issue and easily overcome by the pilot’s experience.

Figure 1.05 Figure 1.06

Wing sweep and other complications were discarded due to the fact that they do not aid

in increasing performance and the only real value they had was to increase aesthetics. It

was in everyone’s interest to add a quarter circular shaped edges to the wing to ensure

that the design stayed above other simple twin fuselage designs.

The project was undertaken and completed successfully in the given time frame with

the allowed materials completing all given tasks, although the weight limit exceeded.

This was expected due to the design itself and the fact that there were two batteries,

extra wiring, two motors and two fuselages which was a key point in this design.

Design Requirements and Objectives 2

The EPUAV project had a set of guidelines and mission objectives as a pathway for the

successful completion of the project.

Objectives

To understand the process of aircraft design and developments

To integrate or motivate to gain the knowledge of aerodynamics, structures,

propulsion, flight dynamics etc.

To develop student skills solving engineering problems

To cultivate engineering teamwork

Requirements

Performance

Endurance T >11 minutes

Maximum level flight speed Vmax>18 m/s

Minimum level flight speed Vmin<9 m/s

Takeoff distance (ground run) Sto<18 m

Gross weight Wto<2.8 kg

Payload Wpl>0.5 kg

Operation

Wing span <2.3 m

Fuselage length <2.2 m

Propulsion

One motor

Battery: Li

Propeller

Speed controller

Cost

Airframe material cost <¥2000

Conceptual Design 3

Synthesis of concept 3.1

Configuration design 3.1.1

The group discussed and shortlisted a few possible design ideas for the EPUAV

High performance and manoeuvrability

Seaplane

Flying wing

Conventional

Biplane

Twin fuselage

At the time of the beginning of the project, Virgin galactic was in the news and the idea

of making a near replica of the global flyer brought the group to a consensus.

Figure 3.1.101

The initial idea was to create a near replica of the global flyer as shown in figure 3.1.101.

The basic idea was not changed throughout the project however due to an enormous

array of limitations; the final product wasn’t a resemblance of the global flyer.

The twin fuselage, large aspect ratio wing idea with a pod in the middle was what later

became the iSpy.

Initial Sizing 3.1.2

Estimating aircraft weight at the start of a project is purely based on historical data. The

maximum weight allowed was 2.8kg. However, since the iSpy was to be a twin fuselage

design, rough estimates of additional components were added.

Given allowed weight 2.8kg (single engine, single fuselage design)

iSpy

Additional battery 400g

Additional motor 250g

Second fuselage 150g

Takeoff weight was therefore, estimated to be 3.6 kg.

The first step in terms of sizing was to calculate wing size. Fuselage size was to be

determined later as fuselage size is inversely proportional to the empennage size. CAD

visualisation of the aircraft would give a better idea about the size of the empennage

and hence the fuselage and a reasonable visually pleasing size would be chosen.

As the aircraft had to have a large aspect ratio, maximum allowed wingspan of 2.3m was

to be used. The chord length was determined from wing loading from historical data.

Typical wing loading suggested in various resources is 4.6 kg/m2 but this would’ve

resulted in a chord length of 34cm. The group realised that this does not fit in with the

idea of a slender thin wing. After discussions with the professor, a wing loading of 7-8

kg/m2 was chosen which resulted in a chord length of around 22cm and an aspect ratio

of over 10.

Fuselage layout 3.1.3

As the whole idea of the plane was to imitate the Global Flyer, we decided to have twin

fuselages. The fuselages were made to be thin and slender with a slight taper at the nose

of the fuselage, a constant body and a slightly bigger taper at the back to go with the

glider concept.

The initial design of the fuselage was to have a complete flat top surface on a circular

body. The circular body was chosen because a circle can take load from every direction.

The height of the top surface from the bottom of the circular fuselage is 80% of the

diameter of the circle. This method is used throughout the design of the fuselage for

consistency purposes (e.g. Nose section, tapered tail section). The idea to have a

complete flat surface on the top of the fuselage was to make sure the connection of the

wing and empennage of the plane to the fuselage would be really easy and quick. This is

really important as our aircraft had to be dissembled and assembled on the spot of the

flight testing ground as it is too big to be carried around assembled.

The nose section of the fuselage was tapered at the bottom to make the aircraft more

aerodynamic. The first design was to have a flat top surface at the nose section but was

later found that that would be impossible as the motor mount would stick out from the

top surface at that given dimension of the nose. We then changed the design slightly to

make up for this. We pushed the top section of the nose down into a little taper and we

increased the nose rib size slightly to ensure the motor mount fits in properly.

The body of the fuselage was made to a constant shape so that the structure is strong

enough to withstand the load from the wing. It is also to ensure a spacious and even

situation inside the fuselage to slide the battery back and forth for cg estimation.

The whole length of the fuselage is 1390mm which is 1.39m. This consists of the nose

section being 100mm, body 760mm and the tail section 530mm. Calculations were

being done to pinpoint the exact length of the aircraft. This will be discussed in detail

later in the report.

Airfoil design 3.1.4

Wing: Before the design layout could start, the airfoil geometry had to be chosen. Parameters such as the cruise speed, takeoff and landing distances, stall speed, handling qualities and overall aerodynamic efficiency during all phases of flight were taken into consideration while selecting the most suitable airfoil geometry as it has direct effects to the parameters mentioned above. It was decided that the airfoil should have a large camber and a thickness to chord ratio in the range of 12-16%. Having kept that in mind, the airfoil also had to meet the desired lift coefficients and have low drag. Due to the small size of the aircraft, it was vulnerable to rapid changes in pitching moments. A smaller value of moment coefficient was desirable for the airfoil as it reduces the nose down moment produced by the wing. A secondary requirement of the team was to have a flat bottom airfoil which makes it easier to fabricate. In order to compare the two dimensional airfoil sections, a user friendly version of Xfoil (ProfiliV2) was used. From the requirements mentioned above, 10 of the most suitable airfoils were short listed and compared. Refer to Appendix B for tables that list the aerodynamic parameters that depend on the airfoil geometry. It provides values of the lift, drag and moment coefficients at different angles of attack. The graphs show the coefficients of lift and drag plotted against different angles of

attack for five airfoils. It also shows the lift to drag ratio and moment coefficients for the

selected airfoils.

Figure 3.1.401 showing cl vs cd values for different airfoils.

Figure 3.1.402 showing values of Cl at different angles of attack.

Figure 3.1.403 showing cl/cd characteristics along with moment coefficients for different angles of attack

From the graphs and tables above, NACA 4415 and Wortman FX 77-153 suited the best

for our purpose. Wortman FX 77-153 showed better lift coefficient characteristics at

higher angles of attack and moment coefficient values. But this was compromised with

higher values of drag. NACA 4415 had better L/D characteristics.

Hence, to achieve the best outcome, NACA 4415 with an angle of incidence of 2 degrees

was selected. The angle of incidence gave us higher values of lift coefficients at small

angles of attack. Compromises on stall characteristics had to be made in order to

achieve this.

Tail:

A symmetric airfoil with a thickness ratio of fewer than 10% was considered suitable

for the tail. A symmetric airfoil produces no lift at zero angle of attack. Adding

symmetric airfoils also avoids trim drag. A thin airfoil was considered to reduce as

much drag as possible. The following graphs compare different airfoils selected for the

tail.

Figure 3.1.404 shows values of cl at different angles of attack.

Figure 3.1.405 showing moment coefficient values at different angles of attack

MM010 was selected as the most suitable airfoil for the horizontal tail due to its superior L/D

and moment coefficient characteristics.

Wing planform design 3.1.5

The wing shape had to be a new idea. The group decided against a rectangular wing as it

would have made the iSpy look ordinary. A wing with curved edges was most suitable to

make the iSpy more visually pleasing. A high wing configuration was chosen due to ease

of fabrication. It was decided not to have sweep, dihedral or winglets due to the added

complexity with fabrication and also due to the fact that the additions to lift would be

negligible in low speed flight. According to the initial plan, the wing chord had to be as

small as possible. A wing loading of 7 kg/m2 gave a chord length of 22cm. This was still

not very thin but the professor suggested that an aspect of over 10 would be too high.

The wing chord had to be increased to bring the aspect ratio close to the suggested

figure of 8-9. After a lot of calculations to meet the specified criteria, a chord length of

26cm was chosen.

c (variable) b (fixed) AR Area Weight (fixed) W/S 0.21 2.3 11.39909035 0.464072118 3.6 7.75741498 0.22 2.3 10.90212414 0.485226542 3.6 7.419214917 0.23 2.3 10.44845137 0.506295126 3.6 7.110477303 0.24 2.3 10.0326608 0.527277868 3.6 6.827519635 0.25 2.3 9.650206988 0.54817477 3.6 6.567248612 0.26 2.3 9.297243807 0.568985832 3.6 6.327046825 0.27 2.3 8.970494923 0.589711052 3.6 6.104684636 0.28 2.3 8.667152053 0.610350432 3.6 5.898250925 0.29 2.3 8.384794267 0.630903971 3.6 5.706098178

Span 2.3 m

Chord 0.26 m

Area 0.56 m2

Wing loading 6.32 kg/m2

Aspect ratio 9.29

The sizing chosen above met all criteria set by the project guidelines and suggestions by

the professor. This however, changed the look of the iSpy meaning that an exact replica

of the global flyer wouldn’t be possible.

Empennage design 3.1.6

The dimensions of the vertical and horizontal tail planes were determined using the

following equations:

Horizontal Surface:

Vertical Surface:

In order to make these calculations the following assumptions and values were

calculated:

Mean wing chord (c): 0.26

Wingspan (b): 2.3

Wing Area (c*b): 568986

Volume co-efficient (𝑉 ): 0.7

Vertical volume co-efficient (𝑉𝑉 ): 0.04

Position of centre of gravity on wing: 0.195

Centre of lift of horizontal tail: 0.054

Centre of gravity position of horizontal tail mass: 0.099

The distance between the rear edge of the wing and the leading edge of the tail (L(w-t))

was taken as a variable.

Using these values, the distance between the the CG of the plane and the centre of lift of

the tail (Lh) as well as the distance between the CG and the CG of the tail mass (Lv) could

be calculated as follows:

LH= L(w-t)+0.195+0.054

LV=L(w-t)+0.195+0.099

Thus all values were available to solve for the initial equations given above:

The area of horizontal tail:

𝑆𝐻 =𝑉 𝑆𝑐

𝑙𝐻

=0.7∗0.56898∗0.26

𝑙𝐻

The area of the vertical tail:

𝑆𝑉 =𝑉𝑣 𝑆𝑏𝑐

𝑙𝑉

=0.04∗0.568986∗2.3

𝑙𝑉

Thus, with the areas of the horizontal and vertical tails known, the dimensions of width

and height for each had to be determined.

The horizontal tail chord had previously been chosen to be 0.18. This meant that the

width of the horizontal tail would be: 𝑆𝐻

0.18

The vertical tail calculations were a little more complicated. The vertical tail chord was

chosen to be 0.24. As well as this, ratios were chosen as to the dimensions of the tail

which were sloping and which were straight. A ratio of 0.4 to 0.6 was chosen. Thus the

Vertical tail height was calculated as: 𝑆𝑉

0.24(0.6 + (0.5x0.4))

As the design involved twin vertical tails, the vertical tail height could be divided by 2 so

as to gain the individual height of each vertical tail:

𝐼𝑛𝑑𝑖𝑣𝑖𝑑𝑢𝑎𝑙 𝑡𝑎𝑖𝑙 𝑕𝑒𝑖𝑔𝑕𝑡 =𝑡𝑎𝑖𝑙 𝑕𝑒𝑖𝑔𝑕𝑡

2

The professor suggested the use of a safety factor when gaining the height of the

individual tails, as it was unsure of the differences in effect that may be encountered

with the use of two vertical tails as opposed to the single tail which the equation was

designed for. Thus a safety factor of 1.1 was decided upon:

𝐼𝑛𝑑𝑖𝑣𝑖𝑑𝑢𝑎𝑙 𝑡𝑎𝑖𝑙 𝑕𝑒𝑖𝑔𝑕𝑡 =𝑡𝑎𝑖𝑙 𝑕𝑒𝑖𝑔𝑕𝑡 ∗ 1.1

2

The above equations vary with the variable of L(w-t). A spread sheet was made to

illustrate these changes and help to find appropriate values for the dimensions. The use

of this spread sheet was highly useful as it allowed all variables to be easily changed so

that their effect was immediately obvious. This spread sheet is shown below.

Figure 3.1.601

Landing gear geometry 3.1.7

Landing gear design initiated with the decision that a quadricycle configuration would

be used for this twin fuselage design as it has been used historically with great success

for similar designs. The quadricycle design consists of 4 landing gears in total, with two

nose gears and two rear or main gears. The four wheels usually take off at the same time

with both the front gears being used to control direction of aircraft with the help of

rudders during taxi, take off and landing. Figure 4.1.501 shows the key idea of a

quadricycle configuration.

Figure 4.1.501- The quadricycle landing gear configuration

At that stage, the group thought that landing gear was a minor component and not a lot

of time was spent on it. A conventional quadricycle configuration was chosen after

discussion with the professor. This configuration was easy to design and the group

believed that it would serve its purpose.

Control surface sizing 3.1.8

The control surfaces consist of ailerons, elevator and rudder. The sizing was chosen

after suggestions from the professor that ideally all control surfaces are approximately

30% of the chord and around 30% of the span. Final control surface sizing and

calculations are shown in appendix D.

Pod Layout 3.1.9

The pod of the aircraft was made the payload for the whole mission. This was because

the pod was designed to be easily removable and assembled. The pod did nothing but to

carry a little spy camera in it to take video during the whole flight profile. The nose of

the pod was made pointy for aerodynamic purposes.

Analysis of concept 3.2

Analysis of propulsion system 3.2.1

Propulsion testing was done in the first week of November. Relationship between the current and the pulling force generated by the motor was obtained for the experiment. Motor #2006 and 14x9.5 mm propeller was used in the propulsion test. Note that the batteries provided to us were old and their performance might have been altered. In the real flight test, two 4 cell (14.8 V) batteries were connected in series. A choice of motors was not given. But from the propulsion test, we concluded that the motor given to us generated enough thrust.

Figure 3.2.101-shows the schematic of the propulsion testing rig.

Figure 3.2.102-Setup of propulsion test

The following data was obtained from the propulsion test. The maximum thrust obtained was 1.424 kg. Since we had to motors, the thrust would be doubled to 2.848 kg.

Force in grams

rpm Current (Amp)

Current drain

119 2405 37.3 0

425 4625 33.37 3.93

637 5344 28.17 9.13

890 6020 23.76 13.54

1124 6833 16.43 20.87

1382 7450 8.8 28.5

1424 7550 7.8 29.5

Figure 3.2.103 - is shows how the thrust increases with increasing rpm

Figure 3.2.104 - shows how the current drain increases with increasing rpm

0

200

400

600

800

1000

1200

1400

1600

0 1000 2000 3000 4000 5000 6000 7000 8000

Forc

e (g

)

Propeller rotation speed (rpm)

Propulsion Test (Thrust)

05

101520253035

0 2000 4000 6000 8000

Cu

rre

nt (

A)

Propeller rotation speed (rpm)

Propulsion Test (Current Drain)

Weight estimation and Centre of Gravity (cg) estimation 3.2.2

Weight Estimation:

Estimation of weights was done using the formulae provided to us by Dr. Yu during the

conceptual design phase. The formulae used were in imperial units. A conversion factor

was used to convert pounds to kilograms. The weight added by glue and wiring were

estimated from the aircrafts made by previous batches. Exact measurements of battery

and the motor weights were taken from the weighing scale. Weight of the aircraft was

then checked using CATIA during the detail design phase. Densities of materials used

were inserted into CATIA which then provided us slightly more accurate results. Note

that the weights of servos were negligible and hence were included in the wiring.

Part Formula Lb Kg

Wing 0.14676·Sw0.4852·ARw0.7082·(100 t/c)-

0.2210 0.942617 0.42756

Fuselage (x2) 0.07092·(Wm/Hm)0.04832·L1.6566 1.77084807 0.803238976

Horizontal tail 0.1570·Sh0.1939 0.16958158 0.076920509

Vertical Tail (x2) 0.1393·Sv0.6729 0.22111 0.10029

Landing gear Wto·0.07 0.5555648 0.251998638

Propellers (x2) 3.0346×10-6·D3.76468 0.11023 0.1

Glue

0.88185 0.4

Wiring

1.10231 0.5

Motors (x2)

0.66139 0.3

Batteries (x2)

1.98416 0.9 TOTAL weight: (excluding 500 g payload)

8.50689 3.85866

Table 3.2.201 shows the weight estimations during conceptual design phase

The centre of gravity positions were measured from the weights obtained during the

conceptual design phase. Moment equations were taken from which the estimated

centre of gravity was found. The estimated position of the cg was approximately 35%

from the leading edge of the wing without placement of batteries and this gave us rough

estimates of the battery movement required to keep the aircraft’s CG within the

required 25%-30% region.

Estimations were also taken from CATIA which gave a cg position at approximately 30%

from the leading edge of the wing including the battery and motors but excluding

weight added on by wiring, glue, skin and thermal shrink film.

Stability and Aerodynamic performance analysis 3.2.3

Aerodynamic performance was analysed using two main methods. One was using basic

equations provided by the supervising professors to calculate drag polar and maximum

lift coefficient. There are many levels of depth at which these calculations can be

conducted. Simplified models include just using theoretical data and empirical data

which may work to a certain extent for conventional designs of aircrafts. A slightly more

intermediate method involves using vortex lattice method which is suitable for more

unconventional designs such as ours and finally, detailed analyses of aerodynamic

performance would include panel method or CFD (computational fluid dynamics) which

is more suitable for later stages such as preliminary and detail design.

Drag Calculations:

We know that,

𝐶𝐷 = 𝐶𝐷0 + 𝐶𝐷𝑖

Where,

CDi is the induced drag (KCL2)

𝑘 =1

𝜋𝐴𝑅. 𝑒= 0.06058

e = Oswald’s efficiency factor

𝑒 = 4.61 1 − 0.045𝐴𝑅0.68 . 𝑐𝑜𝑠𝛬𝐿𝑒𝑎𝑑𝑖𝑛𝑔 𝑒𝑑𝑔𝑒 0.68

− 3.1 = 0.56508

Λ Leading edge: Sweep of leading edge = 0

∴ 𝑪𝑫𝒊 = 𝟎. 𝟎𝟓𝟎𝟓𝟖 ∗ 𝟎. 𝟑𝟏𝟐𝟐 = 𝟎.𝟎𝟎𝟓𝟗

CD0 is the zero lift drag or the parasite. The Following table can be used as a reference for calculating the parasite drag values for each component.

The following tables give additional values of parasite drag calculated from the formulae mentioned.

Adding 5% to CD0for interference drag Adding 10% to CD0for roughness and protuberances Adding 8% to CD0for fixed landing gear

Cdo skin friction 0.035028

Cdo interference 0.0017514

Cdo roughness 0.0035028

Cdo landing gear 0.00280224

𝐶𝐷0 = 𝐶𝑓𝑒 ∗𝑆𝑤𝑒𝑡

𝑆𝑟𝑒𝑓 (𝑓𝑜𝑟 𝑒𝑎𝑐𝑕 𝑝𝑎𝑟𝑡)

From historical data, Cfe = 0.0045

∴ 𝐶𝐷0 = 0.04308

∴ 𝑪𝑫 𝒕𝒐𝒕𝒂𝒍 = 𝟎.𝟎𝟒𝟑 + 𝟎. 𝟎𝟎𝟓𝟗 = 𝟎. 𝟎𝟒𝟖𝟗

Parasite drag related values

Fuselage Wing Tail

width 0.09 Sref wing 0.568985832 Sref Htail 0.138258239

height 0.072 Swet wing 1.169265884 Sref Vtail 0.065937829

variable 0.5 Swet/Sref 2.055 Swet Htail 0.280525967

front cone 0.1 Swet Vtail 0.131875658 behind wing rectangle 0.15 Swet/Sref H 2.029

behind wing taper 0.35 Swet/Sref V 2 h tail length addition 0.18

Area top 0.07875

Area side 0.063

Sref fuselage x2 0.2835

Swet fuselage x2 0.48195

Swet/Sref 1.7 Swet/Sref TOTAL 7.784

Although these values met requirements for this phase of design, soon calculations had to be in more detail and thus AVL had to be used to calculate and perform in depth analyses on aerodynamic performance and stability. AVL (Athena Vortex Lattice) version 3.26 was mainly used to calculate and analyse aerodynamic performance characteristics of the aircraft. Values obtained from AVL were also used to ensure stability in all dimensions. There were a few assumptions involved at this stage of design as AVL was only a freeware and had limited functionality. Shapes and geometry entered as part of the design had to be of minimal complexity and consideration to extrusions such as landing gear or location of motors were not taken into account either. The software itself had an easy to follow interface in command prompt but did not have a particular graphical user interface. The input files were made as accurate as possible by adding detailed coordinate points of the fuselage in order to replicate its shape and size accurately. Airfoil files were then exported from Profili into text files. These files had to be manipulated in order to be used for AVL analyses. The output files were used to extract key numerical values important to our design and stability analysis. AVL considered the airfoil shapes of the horizontal tail and the wing and since no airfoil file was linked up for the vertical tail, it was taken to be a plate by default as the program was designed to do so. Once the files required for AVL analyses were in place, they had to be run on the program to ensure the aircraft displayed accurately to a reasonable extent. The final design of our aircraft was similar to that which is shown below in figure 3.2.301.

Figure 3.2.301 – The aircraft as displayed on AVL.

After this was obtained, the program was run and the output results were saved and used to understand and demonstrate the stability and aerodynamic performance of the aircraft. The stability axis derivatives were run for over 50 different cases from angle of attack (α) ranging from -7° to 21°, plotting values for intervals of 0.5° as shown in figure 3.2.303 for zero degree angle of attack. All the values obtained for CL, CD and Xnp were recorded for all values of α. The neutral point values were used to obtain the static margin of the aircraft at all different possible values of α. Equation used is shown below

𝑆𝑡𝑎𝑡𝑖𝑐 𝑀𝑎𝑟𝑔𝑖𝑛 = 𝑋 𝑛𝑝 − 𝑋 𝐶𝐺 ∗ 100

Figure 3.2.302 – Lift coefficients and other parameters of the aircrafts at different points in two-dimension.

A 30% static margin was calculated for almost all working values of α if the centre of gravity was fixed at 25% of the wing chord (figure 3.2.304). Since the battery and the payload were movable, placing the CG at a location that provided optimum stability was not an issue. The high value of static margin calculated increased our range of possible positions for CG location. The CG position was later calculated to be safe at any value between 25% and 45% because the static margin varied from 30% to 10% respectively. Thus, during flight tests we decided to make 2 possible positions of CG; one at 25% of the wing chord and the other at 32% so that the pilot may decide which position to operate at on the day.

Figure 3.2.303 – Stability axis derivative output for zero degree angle of attack.

The stability axis derivatives for all values of α also provided us with values of CL and CD that were graphed against each other. This graph was compared to the theoretical graph of CL vs CD and they matched to a reasonable extent.

Figure 3.2.304-Stability axis derivatives

AVL was then used to run a steady level flight case with parameters set to meet an average cruise condition. The total weight was set at 3.6kg, α was set to 0°, air density was set to 1.225kg/m3, velocity was set to 18ms-1 and the acceleration due to gravity was set to 9.81ms-2. The resulting loading on the aircraft was plotted as shown in figure 3.2.305. This shows the aerodynamic loading on the exposed surface areas of the wing and empennage. The wingtips showed high loads acting on it and thus it was recommended by the stability team to add strength to this area by implementing a carbon rod or any other suitable method. The team in charge of wing decided to add a ply outline along the leading edge and also a 6mm carbon rod was placed through the ends of the wing. This ensured a stable and strong structure to the high load bearing wing.

Figure 3.2.305 – Aerodynamic forces acting on wing and empennage surfaces.

Aircraft performance analysis 3.2.4

Flight performance analysis was done using the formulae given to us by the professor.

The takeoff distance was calculated using the following formula:

𝐿𝑇.𝑂 =0.908 (

𝑤𝑠 )

𝐶𝐿 .𝑚𝑎𝑥 .𝑇 .𝑂(𝑇𝑤 − 𝑓)

Where

w/s is the wing loading in kg/m2

T/w is the average value of the thrust to weight ratio

f is the ground friction coefficient approximately equal to 0.035

CL.max.T.O is the maximum takeoff lift coefficient

The weight taken into consideration was from the conceptual design phase, i.e. from the

formulae provided to us. Refer to propulsion testing data for values of T/w ratio.

∴ 𝐿𝑇.𝑂 =0.908 ∗ 6.327

1.325 ∗ (0.749 − 0.035)= 6.594 𝑚

The stall speed was calculated as follows:

𝑉𝑠𝑡𝑎𝑙𝑙 = 2. 𝑤. 1

𝐶𝐿,𝑚𝑎𝑥 .𝑆.𝜌

𝐶𝐿,𝑚𝑎𝑥 = 0.9. 𝐶𝑙𝑚𝑎 𝑥𝑎𝑖𝑟𝑓𝑜𝑖𝑙. cos 𝛬 = 0.9 1.3731 1 = 1.23579

Substituting values of wing loading and the maximum lift coefficient in the equation,

∴ 𝑽𝒔𝒕𝒂𝒍𝒍 = 𝟗.𝟎𝟓𝟓 𝒎/𝒔

Endurance Calculations:

𝑇 =𝐶𝑎

𝐼𝑎 ∗ 1000

Ca = Capacity of the battery = 4400 (mAh)

Ia = Flow current during endurance = 15.067 (Amp)

The flow current was measured during the propulsion test. The average value of the

current drain was calculated which is 15.067 Amp. Substituting that into the above

equation, we get an endurance of 0.292 hours or 17.52 minutes.

CAD definition of the concept 3.2.5

The basic layouts of the whole aircraft were done using CATIA of a better

visualisation of the whole aircraft. No major changes were made throughout the project

in terms of the design on the aircraft so all CAD drawings were actually the initial

drawings made.

Figure 3.2.501: The outline of the fuselage

Figure 3.2.502: The modified nose section of the fuselage with the motor

Figure 3.2.503: The wing

Figure 3.2.504: The horizontal tail

Figure 3.2.505: The vertical tail

Figure 3.2.506: The pod

Figure 3.2.507: The isometric view of the whole aircraft

Figure 3.2.508: The front view of the aircraft

Preliminary and Detailed Design 4

Structure layout, initial sizing, and internal layout Design 4.1

Wing structure 4.1.1

The internal structure of the wing is one of the most important sections of the aircraft

since the wing carries most of the load as lift is produced. The location of ribs, spars,

webs, rods is important for the structural integrity of the aircraft.

It was suggested by the professor that distance between ribs should be approximately

8-10cm. It was identified that the distance has to be variable as certain areas of the wing

were to be placed under significantly more loads. These sections were on top of both

fuselages as the weight of the fuselages would cause excessive loads and hence a

reduced distance between the ribs would be ideal to transfer the load. The 6 types of

ribs used in the wing are shown in figures 4.1.101 to 4.1.106. Weight reduction was

achieved by cutting un-required parts in the middle. To maintain structural integrity of

the rib, a distance of 8mm was kept for each hole from the edge.

Figure 4.1.101 Figure 4.1.102

Figure 4.1.103 Figure 4.1.104

Figure 4.1.105 Figure 4.1.106

The ribs on top of the fuselage were extended so that they could slide in into the

fuselage where they would be screwed.

Figure 4.1.107

The two ribs in the middle of the wing were also extended so that the pod could be

screwed on to them.

Figure 4.1.108

The wing had a spar running throughout at the front of each rib. Spar caps were placed

at 30% of chord at the top and bottom. These spar caps were to be joined by webbing at

the front and back. At 70% of chord, a web-spar was added which would lock in the ribs

and hold them together at specified distance.

Figure 4.1.109 Figure 4.1.110

The curved edges would suffer enormous loads so a carbon rod was to be inserted in

the ribs to provide extra support apart from the webspar and the front curve. The two

aileron servos were to be placed into plates in front of the aileron section.

The front of the edge ribs had a curved front spar to provide the necessary shape.

Figure 4.1.111

Plates were wired onto the extended fuselage connection ribs so that fuselages can be

screwed in a different direction to transfer the load onto more components.

Figure 4.1.112

The rear of the wing had balsa sheets at the top and bottom throughout the whole

length of the wing. These were placed to provide enough contact surface for the thermal

shrink film and to provide shape after the film has been ironed.

Figure 4.1.113

Figure 4.1.114

Figure 4.1.115 Figure 4.1.116

Material type was chosen for each component. The components that were going to carry

maximum load were ply, paulownia was used where some load would act. Balsa was

used where minimal load would act and for components that were merely for providing

shape.

Most of the ribs were 2mm thickness except the ones on top the fuselage which carried

maximum load and therefore they were 4mm thick. Ribs were of all three types ply,

paulownias and balsa depending on the amount of load they were carrying. Webspars

were 3mm thick. The spar caps at the front and at 30% chord were 6mm x 6mm pine

rods. All webbing was 2mm balsa. Rear end sheets were 2mm balsa because their

purpose was to provide shape and contact surface for the film.

The following figures from 4.1.117 to figure 4.1.122 show the materials used for the

wing components. Blue colour is for balsa, green for ply, pink for paulownias, black for

carbon rod and yellow for pine.

Figure 4.1.117 Figure 4.1.118

Figure 4.1.119 Figure 4.1.120

Figure 4.1.121 Figure 4.1.122

Fuselage structure 4.1.2

Figure 4.1.201: The internal structure of the fuselage

The internal structure of the fuselage was designed to be light weighted, strong but with

minimum material used. In figure 4.1.201, we can see the complete internal structure of

the whole fuselage. The right fuselage is almost identical to this left fuselage except for

the position of the bolt holes for the horizontal tail connection of which is a mirror

image to the right fuselages.

There are thirteen ribs altogether in a single fuselage. Of which contains one nose rib, a

nose connecting rib, two basic ribs, 6 wing ribs that connects to the top plate, and three

other back ribs with a different size each at the tapered part of the fuselage. All ribs are

made of 2mm ply to ensure strength in the fuselage and the ability of load taking. The

ribs placed in the fuselage each have its certain individual task. This will be discussed in

more detail when we talk about each section of the fuselage.

Figure 4.1.202: Fuselage nose section

The nose section of the fuselage consists of a smaller nose rib connecting to a fuselage

basic rib with two longerons, a 4mm paulownia top plate and two centre plates. The

centre plate runs throughout the fuselage, only stopping at the centre of the back

tapered tail section of the fuselage. This is to ensure that there will be no geometrical

twist in the fuselage body when assembling the whole fuselage.

The connection of the nose rib to the fuselage body was simple. It was then found out

that the thrust the motor was producing and the torsion it produces was a little too

risky for the connection to be left alone. This is because other than glue, there was

nothing there to stop the nose rib from popping out from the connections. The team

then decided to drill two tiny holes near the edge of the nose rib and tied wires around

it to the other rib of the fuselage. This will then stop the nose rib from moving together

with the motor and away from the fuselage.

Figure 4.1.203: The fuselage body

Figure 4.1.204: Top view

The third rib of the fuselage is connected to the nose landing gear. Its job is solely to

take the load from the landing gear during the landing of the aircraft. This is also as

forward as the battery could move in the fuselage. The fourth rib is there just to add

structural strength to the fuselage. As you can see from figure 4.1.203, there are 6

longerons running throughout the fuselage. Like the centre plate, the longerons too stop

at the last third rib.

Since the dimensions of the material given were limited, we had to make the centre

plate and the longerons into two parts to obtain the length that we needed. The

connections of the centre plate could be clearly seen in figure4.1.204. Other than the

longerons and the centre plate, we have a bottom plate sliding through the whole

constant part of the fuselage body. This plate is for the battery to slide on to adjust the

position of the desired cg. This plate can also act as a strengthening plate for the whole

fuselage system. The orange block in fugure() is nothing but the battery.

Figure 4.1.205: The top plate

Figure 4.1.206: The side view of the part

Figure 4.1.207: The rib connecting to the top plate

This part of the fuselage is the most important part of the fuselage as it connects the

wing to the fuselage. That is why it has 4 ribs with each rib doing different jobs. The top

part of the rib has a locking mechanism to make sure the connection to the top plate is

not solely dependent on glue. That is why the top plate is made out of two 2mm ply

stuck together. The bottom plate is made smaller to slide through the smaller gap on the

top part of the rib while the top plate is made to just click down on the ribs and sit on

the bottom plate. The centre of the rib too has two holes to interlock the two centre

plate together. The same goes with the bottom battery plate.

As can be seen, 4 white blocks and 2 green blocks sit on either side of the centre plate,

each connecting to different ribs and the top plate. They are all wing connection blocks.

The first rib that connects to the 2 white blocks is to take the load from only those

blocks. The same goes to the fourth rib which is attached to the other two white blocks.

As can be seen from figure 4.1.205 we have 4 holes on the top plate. These holes connect

to the white blocks which also have holes in them through to the centre plate. There will

be a bolt going through the holes vertically up to connect the wing to the fuselage.

The second rib is connected to the 2 green blocks with a hole on each block going

horizontally through. Two slots can be seen from figure 4.1.205 on the top plates. These

slots for the wing rib extensions to slide through. Similarly, there will be two holes on

the wing rib connection. As the wing and the fuselage are connected, a 100mm bolt is

run through the four holes and tightened by a nut at the other end. This is a double

precaution that the team were taking as the wing to fuselage connection is the most

important connection for the aircraft.

The third rib is again the landing gear rib and its only job is to take the load of the

landing gear.

Figure 4.1.208: the Tail section of the fuselage

As mentioned earlier, the longerons and the centre plate stoped at the third last rib of

the fuselage. Initially, there was no carbon rod running through the back section of the

fuselage and it was only connected with the centre plate, longerons and the top plate.

The team found that the design was a bit flimsy as the back part of the fuselage was

really weak. Carbon rod was added to provide extra strength to the tail section of the

fuselage. This was when an extra rib was added to sandwich the brown block that

connects the carbon rod to the rest of the back section. Carbon rod was chosen because

it was light in weight and it was also very strong. Like the rest of the fuselage ribs that

were connected to the top plate, these back ribs also have locking gaps.

Figure 4.1.209: The back section of the fuselage that connects to the horizontal tail.

As can be seen from figure 4.1.208 and figure 4.1.209, there are two plates on the back

of the fuselage. The second plate is to let the little brown horizontal tail connection

block to sit on. There were two holes going right through the blocks and plates for to

bolts to go through. Since the whole aircraft had to be detached, nuts and bolts were

used to connect the horizontal tail to the fuselage. The second plate also has a role in

strengthening the back part of the fuselage.

There was no particular connection on the fuselage for the vertical tail as the vertical

tail was very small as the fuselage was very long, and the professor suggested that it

should be stuck on with only adhesives. Just to be on the safe side, tiny holes were made

in the fuselage to allow wires to go through to add on the connection of the vertical tail

to the fuselage.

Empennage structure 4.1.3

Preliminary and detail design (Empennage):

Horizontal tail:

Figure 4.1.301 shows the horizontal tail and the elevator assembly in CATIA

As discussed earlier, the horizontal tail was to sit on the fuselage. We had to make sure there was no angle of incidence for the horizontal. A major issue while designing the tail was its assembly with the fuselage. Details on how this problem was fixed are discussed later. The internal structure of the horizontal tail was similar to that of the wing. From the calculations done during conceptual design phase, the chord length was set to 180 mm. The entire horizontal tail was 763 mm long. Due to its size; all ribs were made of 2 mm ply to provide enough strength. The structure consisted of 11 ribs, 3 spars, 1 web spar, webbings, a servo plate and two blocks at the sides. Ribs: The 4 mm ribs on the ends were given flat bottom to avoid any angle of incidence. The ribs in the middle had slots so that the servo plate could sit on it. The distance between each rib was 100 mm except the sides where the distance was 30 mm and in the middle where the distance was 50 mm. This was done to evenly distribute the spacing between the ribs. The rear end of the ribs in the middle had a 35 degree cut which gave enough space for the elevator deflection. Holes were cut in the ribs in order to save space and to connect wires to the battery. Spars: The structure consisted of 3 spars (one in the front, the other two at 25% of the chord). The spars were made of 6x6 mm pine rods. Web spar: This part did functions of both, the spar and the webbing. This part was at the rear side of the horizontal tail structure. The web spar was made of 3 mm ply. It had slots in it so that the ribs and the end plates could easily slide in. The main purpose of these slots was to avoid as much glue as possible and eliminate errors during fabrication. Webbing: 3 mm balsa was used as webbing just to add more support to the structure.

Servo plate: The 6 mm thick servo plate was used to hold the servo. The plate was made of ply to provide strength.

Figure 4.1.302 shows the position of the servo plate assembled in CATIA

End plates: The end plates were the only places which could connect the horizontal tail to the fuselage. The plates were 4 mm thick made of ply. As seen in the figure, the plates sat in between the ribs. These plates were also slotted in. We had to make sure the plates were strong enough to hold the tail.

Figure 4.1.303 shows the position of the end plate in CATIA

Vertical tail: A decision was taken during the conceptual design phase that the vertical tail would be made of a flat plate and not airfoils because it would be easier to design and manufacture. The vertical tail had a trapezoidal shape with a thickness of 5 mm. Its dimensions were 156x188mm.Due to its small size; the tail was designed as a single part as it was within the limits of the laser cutting machine. Ply was selected as the most suitable material due to its strength. The diagonal strips acted as trusses which added on strength in all directions. The edges of the holes were given a circular shape to evenly distribute stress concentration. A hole in the centre was to hold the servo. The vertical tail was to be stuck on to the fuselage using nothing but glue. A small hole (diameter of 1.5mm) was inserted at the sides allow wires to pass through which held the vertical tail and the fuselage together.

Figure 4.1.304 shows the vertical tail design during preliminary design (without rounded edges)

An ‘L’ shaped support was also made to prevent the sideways movement of the tail. The support was stuck on to the fuselage from the bottom and the vertical tail from the other side.

Figure 4.1.305 shows the ‘L’ shaped support for the vertical tail.

Figure 4.1.306 Figure 4.1.307

Figure 4.1.308 Figure 4.1.309

Control surface structure 4.1.4

Ailerons

The aileron structure included 6 ribs. The front of the ribs was to be glued onto a

webspar and balsa sheets would provide shape at the back of the ribs. The two ribs in

the middle had to hold a plate that would have the servo connection part screwed onto

it.

Figure 4.1.401 Figure 4.1.402

All ribs were decided to 2mm paulownia, the front webspar was 2mm paulownia as well.

The servo part holder plate was 2mm ply and the rear sheet was 2mm balsa.

Figure 4.1.402

Elevator: The structure of the elevator was similar to that of the ailerons. It consisted of nine ribs made of 2 mm thickness. The total length of the elevator was 703mm. The ribs slotted into the front plate. The back sheet was made of 1 mm balsa. Its purpose was to make it easier to put the skin on. A plate was slotted in the middle for the servo connection. Due to its small size, a flat plate of similar dimensions was also made as a backup. Ply was selected as the most suitable materials for all empennage control surfaces. Rudder: The thickness of the rudder was 5 mm. Its internal structure was similar to the vertical tail with diagonal strips that acted as trusses. The structure of the rudder was the most simple compared to other control surfaces since it was made from a flat plate.

Landing gear 4.1.5

Initial stages of design included basic ideas and requirements being noted down to

ensure that all criteria can be satisfied during this process of design. This included

loading conditions, ability of front gears to provide enough turn radius, overturn and

tipback angle conditions and ground clearance from the propeller.

Landing gear geometry which needed to be established at the design phase includes

track (B), wheel base (b), tipback angle (θ), overturn angle (φ) and the length of the

landing gears (XV). Since the design required a quadricycle and the wheels were going to

be under the fuselage, it was understood that a vertical line would be suitable for this

case of landing gears and thus the track was fixed at the distances between the fuselages

of 0.768m. The loading on the landing gear was arguable as there was no real equation stating the

values of loads that the nose landing gear may be subjected to. The optimum range for

the percent of the aircraft’s weight which is carried by the nose wheel is about 6-15%,

for most-aft and most-forward CG positions for a tricycle configuration. If the nose

wheel is carrying less than 5% of the aircraft‘s weight, there will be not enough nose-

wheel traction to steer the aircraft. Since there are 2 wheels in front, we decided to use

25% load on the nose gears and 75% on the rear gears after discussing with Professor

Yu. This results in a 12.5% load on each nose wheel and 37.5% load on each rear wheel.

The nose gears were subjected to higher loads than required but this seemed the only

solution at the time to provide suitable locations of front and main landing gears. The

case was for the mot forward CG position of just 25% of the chord length of the wing.

Later during flight test, the CG position used was more toward 32% of the chord length

which meant a lower load on the front landing gears which was initially understood to

be more appropriate for this design.

The criteria that were put forward initially included the major conditions in the design

of a basic landing gear system. The vertical length of the fuselage had to be greater than

250mm in order to avoid any possible propeller and ground interference. So the vertical

height (XV) was set to be greater than 250mm in length. The next main condition was

the horizontal distance of the main landing gears from the CG position. Once the sum of

moments was calculated to be zero at a fixed front landing gear position, the range of

values appropriate for the rear landing gear was obtained. All the conditions were put

into an excel spreadsheet to give a range of possible values we could use. A screenshot

of this spreadsheet is shown in figure 4.1.502. Equations for phi (overturn angle) and

theta (tipback angle) were input into excel after calculations were initially made on

hardcopy and rechecked for trigonometry. The values highlighted in yellow were

allowable within the criteria put forward.

After the possible range of values that could be used was obtained, a discussion was

held to finalise on landing gear geometry. These values were then highlighted in green.

The fuselage team and the landing gear team then coordinated in order to design the

attachment of landing gear into the fuselages. The implementation of suspensions were

discussed slightly and then discarded till fabrication stage due to complexity in design

and lack of expertise at this particular phase.

Figure 4.1.502

Integration of propulsion and control system 4.1.6

Two motors with propellers were to be used as part of the propulsion system. These

motors were to be placed at the front of both fuselages. To make sure that enough space

was left for electrical wiring, fuselage internal structure was made big enough on the

inside. One battery was placed in each fuselage. The receiver was planned to be placed

in one of the fuselages. The choice was made later as the group was aware that after

fabrication, both fuselages won’t turn out to be exactly the same weight. The fuselage

that held the receiver would have more electrical wiring in it and therefore the lighter

fuselage would hold the receiver to balance the two fuselages. The two batteries would

be connected in parallel to make sure that if the batteries don’t discharge at the same

rate, both motors would still operate. The wiring between the two fuselages would be

run through the wing ribs.

Pod Structure 4.1.7

Figure 4.1.701-Pod detail design

The internal structure of the pod is made fairly simple with only three circular rings.

There is a plate going through the centre of the circular ribs to hold the structure

together and to let the 4 blocks sit on it. The 4 longerons helped to straighten the

structure of the pod too. The 4 brown blocks shown in figure () had holes running

through them to ensure that the pod had three different positions to move. This is

another way we can adjust the position of the cg with. Wing rib extensions will be

outside the pod and they will be held together with two 110mm bolt.

The yellow plate at the bottom with a big hole is for the camera to sit on. The lens of the

camera would be placed on top of the hole. The hole was deliberately made bigger to

avoid any obstruction of the camera’s sight.

Payload arrangement 4.1.8

The pod was the designated payload. A spy camera was placed near the front to capture

in flight video according to the mission. Additional weight was to be placed inside the

pod to make sure total payload weight is 500 grams. The additional weights were to be

placed just underneath the cg position so as to not to affect it. The weights could also be

placed in the fuselages as balancing weights if the cg position was to be altered. The

position of the pod could be changed as different holes could be used to screw the pod

onto the wing. This design idea meant that 3 different positions of the pod were possible

underneath the wing. This could also be used to alter the cg position if required. The

additional weights were to be stuck on with double sided tape. Figure 4.1.701 shows a

detailed view of the structure of the payload holder

Structural Analysis 4.2

It is common practice in aircraft projects to check the integrity of the structure. The

group was introduced to NASTRAN to check whether the internal structure created in

CATIA would be able to carry the loads in flight. Due to time limitations, this couldn’t be

done. The group had to rely on advice from experienced teaching assistants and the

professor. The structure was discussed and with minor changes, the professor was

confident that it would be able to carry the calculated loads. Later on during fabrication

and testing, it was discovered that the structure was reliable and caused no problems at

all.

Fabrication 5

Preparation for materials 5.1

The laser cutting machine available in the laboratory read .dxf files. All individual CAD

parts had to be converted to .dxf format in CATIA and then all components had to be

arranged in AutoCAD according to the type of material they were to be cut from. All

components were converted to .dxf files and then sorted according to material type.

Ply, pawlonia and balsa sheets were sorted for cutting. Carbon, ply and pine rods were

also sorted and selected. All tools were collected and placed in one section of the lab.

Superglue, AB glue and white glue were to be used for different purposes.

The group faced many problems in terms of limitation of available materials which

meant that every time an alternative had to be found. Rod, screw hole sizes had to be

continuously changed in the CAD parts due to this.

Figure 5.1.01

Refer Appendix C

Fuselage fabrication 5.2

Since there were two fuselages, it was decided that it would be wise to start with the

first one and get familiarised with it, then start with the second one. This is very useful

as there may be mistakes during manufacture the first fuselage and the mistake could

actually be corrected without wasting additional material. A couple of mistakes were

found while making the first fuselage. For example, the two top plates were made

exactly the same initially in CATIA. During the fabrication process of the second fuselage,

we found that the bottom plate did not fit as there were locking mechanisms stopping it

from clicking in. The bottom top plate was then cut out so that it could easily slide

through the ribs and not click in like the top one. Realisation occurred at the time about

the fact that processes in CATIA does not necessarily ensure fabrication success

The order of the fabrication was fairly simple. CATIA files had to be converted into .dxf

files and rearrange all the parts in AutoCAD to fit the given dimensions of different

materials that we were using. The files were then run in the laser programme and the

laser will cut the exact part out for us. The problem with this part is that some of the

planks of wood were not entirely flat so there were minor errors when the laser

machine was cutting the parts. Those errors were small enough to neglect so it did not

cause a big problem. It had to be ensured that tiny parts were required to cut, as it

might fall through the gaps under the planks. It is best to slide a thin sheet of used balsa

to stop that from happening.

The fuselage design consisted on many interlocking parts to ensure all connectivity was

well secured. Gluing was only a secondary process of connection and only in place to

avoid slipping.

The design of the fuselage had readymade slots for everything to be locked together so

figuring out the actual position of each rib was not at all a problem. Extra effort was

taken into strengthening the connections, for example, the centre plate and the

longerons that had to be split into two due to the restriction of length were to sandwich

it with two carbon strips on the joints. That would have ensured that the joining of two

would not split apart during flight or when carrying a load.

After gluing everything into place, the fuselage was to be left overnight to dry

completely. There were three different types of glue that we used (AB Glue, Superglue,

PVA Glue). Each glue needed different amount of time to dry so it all depends on which

glue was used. PVA glue was the glue used for parts that did not do heavy duty work

and superglue was the glue when you needed something to be in place instantly while

AB glue was the strongest of all and it was to be used on parts the take heavy loads.

When the skeleton of the fuselage was ready, the balsa skin was ready to be placed on it.

1.5mm balsa sheet was used as the skin of the fuselage there was no 1mm balsa sheet

available. To prevent the balsa sheet from cracking, they had to be placed or submerged

in water first. The balsa in water technique worked great on the fuselage as we had a

circular fuselage. It would be really difficult to bend the balsa sheet if it was just a dry

sheet. After bending the fuselage into shape, the balsa is left to dry up and take the

shape of the fuselage, and then we stick it on. We deliberately left two major parts of the

balsa not completely stuck onto the fuselage as we wanted to have an easy access into

the fuselage for maintenance purposes. The whole top part before the top plate was left

open with a balsa sheet acting as a door and half of the side of the fuselage under the

wing was left open as they are the most crucial part of the design and requires high

maintenance.

After all the balsa skin was done, thermal shrink films were ironed on top of the balsa

skin for finishing touch.

Wing fabrication 5.3

Wing fabrication was started by cutting all parts (ribs, spars, webs, rods etc.) and

labelling them. Once all the required parts were cut and collected, the process of

sticking everything together was started. AB glue was used for most of wing structure

however superglue was occasionally used. A total of 5 webspars were used and firstly

all the ribs were glued and locked in. Webspars ensured that the ribs couldn’t move in

any direction. The only problem was that the wood used was not perfectly straight

which meant that some ribs and webspars had a bend in them. The bending was

corrected once all rods and spars were glued on. After the webspars, spar caps at 30%

at top and bottom of the ribs were glued on. The sparcaps run through most of the wing

structure which meant that 2 rods had to be joined due to lack of availability of longer

rods. After this, front spar was glued on and it was also done by joining 2 pieces.

Figure 5.3.101

The servo holder plates in front of the aileron section were placed in between the ribs.

As these plates were locked in, no glue was used. Balsa webs were glued on after this.

Carbon rods blocks were also inserted near the curved edges to make sure the carbon

rods don’t move. Once the carbon rods were inserted into the blocks and the edge ribs,

glue wasn’t needed to stick it. The front curves of the edges were glued on next. At the

rear of the wing, balsa rear sheets were glued on onto the ribs at the top and bottom.

Balsa sheets were then glued on to the front top and bottom. These sheets are mostly

for thermal shrink film. Fuselage connection plates were then wired to the ribs and

spars. The 2 servos were then screwed onto the plates. Thermal shrink film was then

ironed onto the whole wing.

Figure 5.3.102 Figure 5.3.103

Empennage fabrication 5.4

Horizontal tail:

Figure 5.4.01 shows the position of the end plat and the servo plate during fabrication process

All parts of the horizontal tail were cut using the laser machine. In order to avoid any twist in the structure, the following order of fabrication was observed.

The 4mm side ribs and the end plates were made by sticking two 2 mm ply ribs. The 6 mm thick servo plate was made by sticking two 3mm parts together. The glue was allowed to dry overnight.

All ribs were first stuck on to the web spar and the front spar. The slots on the web spar made sure the ribs were straight and aligned perfectly.

The top and bottom spars were then added to the structure. All parts were stuck on using AB glue and super glue.

The end plates were slotted and stuck in before attaching the end ribs. The servo plate was then stuck on to the ribs in the middle. Webs were then added to the structure. Finally, 1.5 mm balsa sheets of exact measurements were stuck in the front to

make it easier to add skin. The skin was then added.

Vertical tail:

The 5 mm thick vertical tails were made by sticking one 2 mm ply and one 3 mm ply together. The glue was allowed to dry overnight.

The ‘L’ shaped support was made by sticking three 6 mm pine rods. After the skin was introduced, a balsa sheet was glued on to the vertical tail

which was then glued to the fuselage.

Landing gear fabrication 5.5

The design layout of the landing gears was very basic in detail and the manufacture of

them was taken into little consideration during the design stages. The fuselage was

designed in order to hold the 4 basic gears in place with the front gears being able to fit

into the same compartment as their corresponding servo. The landing gears by

themselves had not been thought out so during fabrication, many decisions had to be

made. The gears had to be strong in tension and compression yet light enough to stick to

weight estimations. Thus, steel rods which were easily available were opted for.

Figure 5.5.01

The bending of landing gears in order to hold the wheels were more complicated than

expected. Since no special machinery was available and due to time constraints, it was

decided that they would just be done by hand at the laboratory instead of seeking

assistance at any local metal working store. This meant that the landing gears were not

bent to an exact right angle. They also did not have a clear and defined bend, but rather

curved upwards to meet the required angle. This proved fatal later during ground tests

and the fact the metal itself was not stable on tarred road but performed better on

flatter concrete proved that the coefficient of friction between the two different surfaces

made a significant difference on the forces experienced by the gears.

The major improvement that was implemented into the design was a suspension

system for the rear gears that was holding approximately 75% of the total load. This

system is shown in figure 5.5.01 and 5.5.02. The suspension had to be limited to a 1cm

spring due to its lack of earlier addition to the design. The fuselages were not designed

to allow for such a system and the bottom plate of the fuselages restricted the allowance

for the suspension spring. The spring was placed between the gear holder placed in the

fuselage above and a stopper below. Finding screws and stoppers also seemed a

challenge at the time as hex key that was required to use some of the stoppers were not

available. We overcame this by using a Philips screw driver to force out the hex screw

on the stoppers.

Once the suspension was designed, installation was in play. This task also seemed to

turn out to be more demanding than expected. The design of the fuselages were based

more on structural integrity than accessibility in later stages of manufacture and that

mistake cost the team time and effort in a small scale. Issues were minimal and

overcome quick as we progressed. It required at least two team members to

successfully install a landing gear and time taken to do so was also not appealing

enough to have a removable landing gear system.

figure 5.5.02

The front gears at the time did not appear to have many complexities but as

manufacture began it was understood that the task will not be effortless. The

installation of the landing ears by itself was not demanding as such but the connection

metal rod to the servo was intricate. After the installation of these were don't, we

realised that the right wheel of the front gear was not as sturdy as the other three

wheels in place and the stopper used in this gear had lost its thread to an extend that

the screw would not tighten beyond a certain limit. This meant that the landing gears

neutral point had to be fixed after any major movement with a load on it. The turn

radius of the servo also did not match as we had planned with the front landing gear.

They would function well if the servos used were trimmed to a certain angle but any

jerks outside these angles would cause the landing gear rod to trip into an awkward

position at which it will no longer function as the steering. This issue was only resolved

by adjusting angle of the servo horn at neutral position and strengthening the

connection rod.

After the landing gears were fabricated, a basic run test was conducted and result

turned out positive with all functions working. Improvements to the design would have

included an appropriate implementation to fuselage design to allow for an increased

suspension system. A more active and suitable position for the servo would also have

helped in creating a more adequate and controllable front gear movement. The steel

rods used could have been bent at a metalworking shop which would have ensured a

more reliable and sturdy right angle at a point rather than a curved turn. This would

definitely have improved on strength.

Fabrication of Pod 5.6

Like the fuselage, the pod parts were also design to interlock each other. After the glue

dries, balsa skin was stuck onto the pod and the thermal shrink film was then ironed on.

Part of the balsa was not stuck onto the pod as it acts as a door for the access of the

camera.

Control system installation and test 5.7

Figure 5.8.01

Motor

(R)

ESC

(L)

ESC

(R)

Battery

(L)

Battery

(R)

Motor

(L)

Receiver

Elevator Rudder

(R)

Rudder

(L)

Ailerons

Nose

Gear (L)

Nose

Gear (R)

Channel 3

(heavy duty)

Channel 3

Channel 4

Channel 1

Channel 2

The control system installation was more tedious than expected and there were many unforseen obstacles. The basic schematic of the control system layout is shown in diagram 5.8.01. The stages that we had to undergo in order to prepare and install the control system can be described as follows.

1. Measurements were made as to required lengths of wiring for all control surfaces and landing gear, ESC (Electronic Speed Controller), batteries and motor.

2. Wires were acquired according to their task (heavy duty for batteries and motor

and normal 3 way wires for the rest) and then cut to the required lengths.

3. The wires were all stripped at ends and resistance measured to ensure high standard of conductivity and reliability.

4. The faulty wiring found was discarded and the rest were then soldered as per required for usage with appropriate connector pins.

5. The wiring was split into three categories; wiring that would stay in left fuselage, wiring that would stay in right fuselage and wiring that would stay in the wing. On connection of all three and the tail, the connections could easily be made within a matter of minutes with the connections that the electronics team had set in place.

6. All the control surface and landing gear wires were placed, connected and then tested to ensure control surface functionality.

7. The ESC connection tests partially failed as one of the ESC’s burnt up on testing. The cause of this is still unknown and is suspected to be caused by the motor as many ESC’s followed this pattern of burning up during testing. The working ESC was tested on both motors and it seemed to work completely accurately and thus a conclusive decision was never made on the cause of these damages to ESCs

8. ESC replacements finally seemed to function properly. Motor functionality however seemed to cause minor failures at start-up at times with only one motor running and the other motor cranking to a stop. This was never fixed as the pilot agreed that it was not a major issue and would be dealt with during test flight. This proved to be minor as the pilot managed two great successful flights on testing of the aircrafts with these motor issues in place.

9. The control surfaces were put through a more thorough test and with the opinion of the pilot in charge, changes were made to turn angles of all control surfaces by moving the servo horn position and the metal bar lengths to provide optimum performance.

10. During all stages, the electronics was constantly tested with a remote control and battery to ensure the wiring was accurate and notes were made to ensure proper polarity matches on connection during test flights.

Control surface fabrication 5.9

Ailerons:

The aileron parts were cut and labelled. AB glue was used to stick all ribs onto the

leading edge webspar. The rear end balsa sheet was then stuck on. It was later

discovered that the ailerons wouldn’t fit into the slot in the wing. The mistake in the

CAD model was that the ailerons were flush with the wing when placed in the slot.

However, in fabrication the sizing wasn’t perfect and this meant that the aileron had to

be shortened span wise. This was done easily by cutting 3 mm on each edge of the front

webspar and re-gluing the last rib.

Elevator: The elevator ribs were first slotted into the front plate and glued on using super

glue. The servo plate was slotted in before attaching the centre rib. The 1mm balsa sheet was the attached to the rear end of the elevator for the skin. The ‘elevator servo part’ was then attached on to the plate. The skin was added.

The elevator got damaged while adding the skin. Hence the backup flat plate was used as the elevator for the aircraft. Rudder: The fabrication process of the rudder is exactly the same as the vertical tail. After fabrication, the control surfaces were attached to the aircraft using tape.

Assembly and test 5.10

The first task which had to be undertaken on the day was the assembly of the aircraft.

This was necessary due to the fact that the aircraft had to be dismantled for

transportation. In order for the aircraft to be unassembled, several elements needed to

be added to the design of the plane. Firstly, the wing and horizontal tail plane had to be

removable from the fuselage, resulting in four separate components for transportation;

the two fuselages, the wing and the horizontal tail. This meant that these components

had to be attached with screws only; no glue could be used when joining them, as it

would have no time to dry on site. Hence, rather than using glue, a more innovative

solution had to be found to securely attach the components. This was as follows: The

wing had four ribs which protruded downwards, two of which went into each fuselage,

where a screw went through both ribs as well as a block fixed inside the fuselage,

appending the wing and the fuselage together. As well as this, 4 screws projected

upwards through the top of the fuselage and bottom of the wing and fastened inside the

wing, further strengthening the connection between the wing and the twin fuselages. It

was also necessary to connect the horizontal tail to both fuselages. This was again done

using screws. Two screws were connected through blocks which were previously

attached inside the fuselage. The horizontal tail could then be placed on top and slotted

onto these screws. The pod was connected in a similar way to the fuselages, with two

ribs extended downwards, and then two large screws going all the way through both

ribs as well as predetermined screw holes in the fuselage. It was decided that as the

vertical tails were comparatively small relative to the rest of the components, it would

not be compulsory to disconnect them for transport. Hence the vertical tails were glued

and attached to the fuselages prior to the test flight day.

Another adjustment to the design of the aircraft essential for the dismantlement was the

necessary discontinuity of the wires within the aircraft. The wires between the

fuselages, wing and horizontal tail had to be able to easily connect and disconnect at a

point so as to allow each component to be transported independently. This meant that

on assembly, the wires within the wing had to be pulled down between a hole in the

bottom of the wing and top of fuselage, and then connected with their fuselage

counterparts. Similarly, wires protruding from the fuselage were connected to their

counterparts on the horizontal tail. The task of attaching the two fuselages to the wing

and the horizontal tail plane, and connecting all the necessary wires was in theory a

relatively simple task, which had on all previous occasions been performed in less than

ten minutes. However, on the day of the flight tests, various complications arose which

hindered the speed of the assembly.

The first such complication occurred as the left fuselage was being attached to the wing.

Each fuselage was attached to the wing using five screws. So as not to lose the nuts

involved in this procedure, these had been screwed onto the screws protruding from

the fuselage while it was separated from the wing. This first problem arose due to the

fact that while trying to find an innovative solution using super glue, the team had

inadvertently glued several nuts and screws together. This led to an excruciating 10

minutes in which the team wrestled to remove the super glued nuts.

After overcoming this initial issue, assembly was continued. However, after both

fuselages had been attached to the wing, it was noticed that one of the wires in the left

fuselage had wrapped itself around a rib in the wing, and was caught between the wing

and the fuselage. This meant that the left fuselage had to be completely removed from

the wing to untangle the wire. This wasted another few minutes, and contributed to the

fact that the assembly took almost twice as long as expected. The final task performed

pre-flight was the connection of the electronics, and the subsequent checking of the

control surfaces and motors. This was performed without any glitches.

Measurement of Weight and CG location 5.11

Part Conceptual design phase using formulae (kg) CATIA (kg) After Fabrication

(kg)

Wing 0.42756 0.38311494

2 1.080

Fuselage 0.803238976 0.8372 1.230+1.210=2.44

Horizontal tail 0.076920509

0.081958742 0.215

Vertical Tail 0.10029 0.0572 Landing gear 0.251998638 0.25 Propeller 0.1 0.1 Total Weight 1.959 1.7094 3.73

Note: The weights taken after fabrication include wiring and glue for each part. There is

a difference between weights of the 2 fuselages due to different wiring methods. The

fuselage weight also includes the weight of the vertical tail as it was stuck on.

After the aircraft had been assembled, a correct position of the centre of gravity of the

whole aircraft was to be measured. This is done to calculate the exact positioning of the

battery and other movable weights. It was found that without the payload pod whilst

the batteries were placed at the forward most position in the fuselages, the centre of

gravity was at 32% of wing chord. The pod being extremely light did not affect this cg

position. To bring the cg forward, two lead pieces with a total weight of 330 grams were

placed in front of the batteries in both fuselages. This addition moved the cg to 25% of

wing chord with or without the pod. Cg positions of 25% and 32% were marked on the

wing edges for the pilot. If the pilot wanted the cg to be at 25% then payload lead

weights could be placed in front of the fuselages. However, if the pilot wanted the cg to

be at 32% then the lead weights could be placed in the pod underneath the cg. In both

cases, the required payload weight of 500 grams would be achieved at 2 different

change-able cg positions.

Tests 6

The final hurdle in any design engineers work is the test flight. This is the concluding

phase of designing an aircraft, and occurs after the design and fabrication phases of

design have been completed. The test flight serves as the ultimate test of the success or

failure of the aircraft. In this phase, any flaws from the previous phases will surface, and

areas of weakness and electronic and material malfunctions will be illuminated. This

stage will uncover weakness not only in the manufacturing techniques, but will also

highlight flaws in the basic design of the plane.

The series of tests which an aircraft is subjected to in this stage is dependent on the

authorities which the aircraft is subservient to. In the case of civilian planes, the Civil

Aviation Safety Authority sets stringent standards and guidelines along which tests

must be conducted. However, whatever level the aircraft is being tested on, the test will

involve analysis of similar sections.

Firstly, the basic structure must be strong enough to withstand the forces acting on it

during maneuvers such as taxi, take off and landing, as well as maneuvers performed

during flight. This structure should ideally prove to be strong enough not just to endure

these forces once, but to be able to endure them repetitively.

Equally importantly, the testing process is designed to check the functionality of the

electronics and wiring of the plane. This must be effective in the control of the plane,

and the powering of the engines, as well as safe enough that it poses little to no risk to

the aircraft or any payload.

Beyond the capability of an aircraft to become airborne still intact, aircraft testing is

designed to analyze the performance of the aircraft once in the air, as well as on the

ground. This involves experimenting with the maneuvers which the plane is able to

perform, as well as the overall aerodynamic performance of a flight vehicle. In the

corporate world, this offers investors a chance to see their investment in action, and to

compare it with the design parameters and requirements initially laid out.

In the case of the EPUAV designed by RMIT students, the testing was far less stringent

than that a civil or military aircraft is subjected to. The testing phase involved two

ground tests, and three flight tests. These tests were designed to analyze similar areas

to those highlighted above. The EPUAV needed to be capable of safely and securely

handling an array of maneuvers, both on the ground and in the air. Not only did this

mean that the design chosen had to be steady and aerodynamic enough to perform such

tasks, but also that the structure had to be able to withstand the forces applied due to

these maneuvers.

Rather than being subject to the authority of the Civil Aviation Safety Authority, or

private investors, the EPUAV “iSpy” was subject to the critique of NUAA professors. The

design parameters set involved the capability to carry a payload of 500g, to take off and

land within a designated amount of space, and the ability to fly at various speeds. For

‘iSpy’, and the students who created it, this was the ultimate test.

Ground tests 6.1

The aim of a ground test is to check the check that the aircraft electronics and wiring, as

well as engines and control surfaces are all in proper working order. If this is the case,

the aircraft will be able to be easily controlled using the remote and the structure and

appropriate control surfaces will be adequate to allow the display of various ground

maneuvers. Although all control surfaces are checked during a ground test, those of

utmost importance are the landing gear, as these are the primary steering device during

taxi.

Ground test 1 6.1.1

Ground test 1 was performed on Monday the forth of January, 2009, on the asphalt area

outside building A10 on the grounds of the Nanjing University of Aeronautics and

Astronautics. This area was chosen due to its close proximity to the lab where the crafts

were manufactured; however it had the drawback of a large number of potholes.

The ground test began by connecting the wires, the ESC and the battery and then testing

these connections. The control surfaces were checked while the aircraft was stationary,

by checking that the remote moved the ailerons, rudders and landing gear to the

appropriate extent in the appropriate direction. The propellers were then checked by

switching them on, but keeping the plane stationary.

It was here that the first problem was identified. The left engine would sometimes fail to

power on, resulting in the left propeller spinning slowly or not at all, while the right

worked perfectly. Although investigations were made into the causes of this

phenomenon, it was concluded that the problem lay with the motor itself, rather than

with any error made by the fabrication team, and hence the team was unable to find a

permanent solution. The solution embarked upon was simply to switch the engine on

and off until both engines began working to their full extent. It was estimated that the

engine failure would occur approximately 50% of the time. This approach worked well,

although it did raise some concerns that the engine might fail mid flight, which would

have had a disastrous result. However, the team did not deem this to be too imminent a

threat, as in all of the tests the engine only ever failed at the initial power on, never once

it was already running.

After it was satisfied that the engines and control surfaces were working effectively, the

engines were powered up to the point where the aircraft began to move. After this

began a series of ground exercises, designed to test the maneuverability of the EPUAV.

The first task undertaken was to simply taxi in a straight line.

Here the second problem was encountered. When no controls were pressed other than

the throttle, the vehicle should ideally have moved in a straight line. However, it was

discovered that iSpy instead was turning slightly right. This was initially immensely

dangerous to iSpy, as all the team members were relatively new to the use of the

controls, and would try to compensate for the movement right with a large and sudden

movement left, which on several occasions put the plane in peril, as it veered towards

obstacles such as bikes. The cause of the crafts tendency to veer right was investigated,

as was determined to be the fault of the right landing gear.

There were several reasons for which the right landing gear was malfunctioning. The

first of which was that during the manufacturing process the steel rod from which the

landing gear was created was bent into shape by hand. This meant that a perfect right

angle turn could not be achieved, and therefore the angle between the horizontal where

the wheel was attached and the vertical component connecting to the fuselage was

slightly greater than the desired 90 degrees. This was more a problem for the front

landing gear, as the steel rods from which they were constructed were provided by the

university, and were more difficult to bend than those used in the rear landing gear,

which were purchased by the team. As a solution to this problem, it was suggested that

the landing gear could be taken to a metalworkers, however, the problem was initially

not deemed of great importance, and this was never followed up on.

Another issue which attributed to the malfunction of the landing gear was the right

wheel itself. The wheels were all scavenged from the wreckage of previous planes as

new wheels were not supplied. This led to some difficulty in finding wheels of adequate

quality. As a result of this, the wheel used for the right nose landing gear was slightly

faulty, as the groove inside was uneven.

Another potential problem associated with the landing gear was the load the nose gear

was being subject to. It was stated that the nose gear should take no more than 6-15%

of the weight of the aircraft. However, this fact was for a single nose gear. As iSpy

possessed twin nose gear, it was assumed that it could withstand double the weight of a

single nose gear. Thus the nose gear was subject to 25% of the load. This was a very

high percentage of the load, and the idea that it could withstand this much was based

simply on an assumption. Thus the possibly excess load, coupled with the faulty wheel

meant that extra force was placed on a steel rod which was already slightly bent out of

shape.

Other than this landing gear issue, the first ground test was deemed to be an overall

success. The controls and wiring were all in excellent working condition, and the plane

was able, despite complications, to maneuver and taxi well.

Ground Test 2 6.1.2

This second ground test was not a part of the course plan, and was made necessary due

to the disastrous events which occurred several hours after ground test 1. When testing

the engines later on the day of the first ground test, it was discovered that both ESCs

had somehow burnt out, and needed to be replaced. This meant that new ESCs had to be

prepared and attached in the place of the old ones. Once the new ESCs had been

installed, the team deemed it necessary to perform a subsequent ground test to check

that the new ESCs were working. Hence, ground test 2 took place on Tuesday the fifth of

January, 2009.

Ground test 2 was conducted in a similar manner to ground test 1. The controls and

motors were checked first while the aircraft was stationary. Then the vehicle was

guided through an array of maneuvers. The new ESCs proved to be in excellent working

condition, and the controls and engines all worked resoundingly well, although with the

lingering problem of the left engine. The landing gear had by this stage been

disassembled and re-formed, and the aircraft was able to perform admirably.

Flight Test 6.2

The flight test aimed to examine the ability of the EPUAV to perform maneuvers such as

takeoff and landing, as well as its ability to undertake certain mid air operations. For a

successful flight test, all control surfaces must be in excellent working form, as they are

all necessary to maintain control of the vehicle during the mission. The mission profile

in this case was of relative simplicity, the aircraft needed to be able to take off, fly for at

least 11 minutes, while performing some basic flight maneuvers and then safely land.

The flight tests were conducted on Wednesday the sixth of January, 2009, at a site

approximately 30 minutes from NUAA which had an adequate runway for the task. On

this day, three flights were undertaken by the EPUAV iSpy.

Flight test 1 6.2.1

The first flight test was performed with the plane carrying no payload, although the

small spy camera was taped to the underside of the wing to document the crafts

performance. The Centre of gravity for the aircraft with no payload had been calculated

to be at 32% of the chord, this position was marked on the wing to in order for the pilot

to have a choice of CG position.

The first flight test began on a worrying note. As the plane built up speed down the

runway, it appeared that the landing gear issues had resurfaced. As the plane reached

high speeds on the rough surface, it was obvious that the wheels were shaking and

moving around a lot, indicating that the large forces exerted on them from the

roughness of the ground were having a highly negative impact on the landing gear. The

aircraft resumed its previous inclination to veer to the right. At higher speeds, the

veering was even more severe, and the pilot was forced to steer the plane left in order

for it to remain on the runway. Thus, the aircraft had to be corrected on takeoff by

ailerons being deflected at takeoff

However, once the plane was airborne, the problems with the landing gear became

obsolete. iSpy performed marvelously in flight, proving that it was able to successfully

turn, climb and dive. In fact, the pilot commented that the in-flight controls were

excellent, and that the aircraft could easily dive and regain height, although he too

acknowledged that the landing gear was problematic.

The landing was excellent and problem free. The aircraft descended gracefully to touch

down gently without an issue, and was quick to stop once hitting the tarmac, meaning

that the faulty landing gear had little effect.

Once upon the ground, the team examined the landing gear and found that the previous

problems had indeed resurfaced, and were again causing the aircraft to veer right. It

was agreed that this problem could not be solved on site; however the team concluded

that the pilot was clearly incredibly skilled, and able to successfully take off even with

problematic landing gear, and therefore the landing gear would not impinge upon the

actual flight to a great extent.

Flight test 2 6.2.2

This flight test was conducted after adding 500 grams of payload. Some of the payload

was in the form of the pod, which was attached for this stage of the test flight. As

previously mentioned, the centre of gravity without payload was at 32% of the chord.

The team had calculated that this could be moved to 25% if an additional 200g was

added in front of the battery. However, the pilot suggested that the aircraft would attain

greater stability if the centre of gravity remained at 32%. Thus rather than adding

balancing weights, the extra payload was added in the pod, directly under the centre of

gravity, so as not to change its position.

iSpy was able to successfully perform while carrying half a kilogram of payload, and

managed to take off, maneuver and land as successfully and well as in the first test flight.

However the problem with the landing gear remained and the second takeoff was

equally as shaky as the first, although the pilot managed to brilliantly overcome all

obstacles and attain lift-off.

Flight test 3 6.2.3

The third test flight was more in keeping with the planes mission as a spy plane. The

payload was removed, and the camera taped to the wing again. Then the team

instructed the pilot to fly as low as possible over the team so that they could appear on

the video being taken by the spy camera. This was an important mission, as it is a

critical attribute of a spy plane that it be able to fly low and conduct reconnaissance

work.

iSpy completed this mission stupendously well, and then spectacularly undertook a few

more maneuvers before beginning descent. On this third and final descent, the pilot shut

off the engines for landing, to test the gliding ability. iSpy rose splendidly to this final

challenge and glided magnanimously to safety.

Overall, these test flights can be viewed as an overwhelming success. On a basic level,

ISpy was able to successfully taxi, takeoff, fly and land, and is still intact enough to be

able to perform all these procedures again if necessary. All the control surfaces worked

perfectly, and the plane proved to be fantastically well designed and built, such that it

was able to withstand the forces applied to it during flight. The exception to this is the

malfunction of the landing gear, but even this was not too severe and did not drastically

affect the performance of the plane. Despite its faults, the landing gear was able to

withstand the forces applied to it during landing. The test flights also showed iSpy to

have successfully achieved the aims set out in the design parameters, both those set by

the professor, and those set by the team. Despite the shaky takeoff, the takeoff distance

was not too great, and the plane was never in danger of overrunning the allotted

distance. As well as this, iSpy was easily able to lift the 500 gram payload, and remain

airborne for the designated length of time. iSpy also fulfilled the mission set by the team,

and showed itself to be capable of undertaking reconnaissance work. Overall the test

flights were a marvelous success, and iSpy emerged triumphant.

Appendix 7

Appendix A – AVL files

iSpy 0.0 Mach 0 0 0.0 iYsym iZsym Zsym 0.568985832 0.260 2.3 Sref Cref Bref 0.065 0.0 0.0 Xref Yref Zref 0.04308444 CDo #============================================= SURFACE Wing 10 1.0 30 -2.0 YDUPLICATE 0.00 ANGLE 1.0000 SCALE 1.0 1.0 1.0 TRANSLATE 0.0 0.0 0.0 #---------------------------------- # Xle Yle Zle chord angle Nspan Sspace SECTION 0.0 0.0 0.0 0.260 0.0 10 -2.0 AFIL NACA4415.dat #---------------------------------- # Xle Yle Zle chord angle Nspan Sspace SECTION 0.0 0.89 0.0 0.260 0.0 10 -2.0 AFIL NACA4415.dat #---------------------------------- # Xle Yle Zle chord angle Nspan Sspace SECTION 0.259 1.15 0.0 0.001 0.0 10 -2.0 AFIL NACA4415.dat #============================================= SURFACE Horizontal tail 10 1.0 YDUPLICATE 0.00000 ANGLE 0.000 SCALE 1.0 1.0 1.0 TRANSLATE 0.760 0.00000 0.00 #--------------------------- SECTION

0.000 0.00000 0.00000 0.180 0.000 10 0 AFIL MM010.dat #--------------------------- SECTION 0.000 0.365 0.00000 0.180 0.000 10 0 AFIL MM010.dat #============================================= SURFACE Vertical tail 14 1.0 YDUPLICATE 0.0 SCALE 1.0 1.0 1.0 TRANSLATE 0.78400 0.00000 0.00000 #--------------------------- SECTION 0.00 0.38400 0.0000 0.24000 0.000 10 1.5 #--------------------------- SECTION 0.096 0.38400 0.188856 0.144 0.000 10 0.5 # #============================================= BODY Fuse 12 1.0 # TRANSLATE -0.450 -0.384 0.0 # BFIL fuse.dat #============================================= # BODY Pod 12 1.0 # TRANSLATE -0.150 0.0 0.0 # BFIL pod.dat #============================================= # BODY Fuse 12 1.0 #

TRANSLATE -0.450 0.384 0.00 # BFIL fuse.dat #=============================================

Appendix B – Airfoil Data

NACA4415 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 0.1515 0.0125 12.12 -0.105

-2.5 0.2064 0.0121 17.058 -0.104

-2 0.2605 0.0117 22.265 -0.104

-1.5 0.314 0.0113 27.788 -0.103

-1 0.3651 0.0107 34.122 -0.102

-0.5 0.4121 0.0102 40.402 -0.099

0 0.46 0.01 46 -0.096

1 0.633 0.0104 60.865 -0.108

1.5 0.677 0.0106 63.868 -0.106

2 0.7228 0.0109 66.312 -0.104

3 0.8172 0.0116 70.448 -0.1

3.5 0.8647 0.012 72.058 -0.098

4 0.9127 0.0123 74.203 -0.097

4.5 0.9607 0.0127 75.646 -0.095

5 1.0087 0.0132 76.417 -0.094

5.5 1.0544 0.0136 77.529 -0.092

6 1.0976 0.014 78.4 -0.09

6.5 1.1411 0.0145 78.697 -0.087

7 1.1835 0.015 78.9 -0.085

7.5 1.2223 0.0156 78.353 -0.082

8 1.2564 0.0162 77.556 -0.078

8.5 1.2856 0.0171 75.181 -0.074

9 1.3159 0.018 73.106 -0.07

9.5 1.3408 0.0191 70.199 -0.066

10 1.3594 0.0207 65.672 -0.061

10.5 1.3669 0.023 59.43 -0.055

11 1.3715 0.0257 53.366 -0.049

11.5 1.3662 0.0292 46.788 -0.044

12 1.3513 0.0338 39.979 -0.039

WORTMANN FX 77-W-153 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 -0.052 0.0192 -2.682 -0.024

-2.5 -2E-04 0.0181 -0.011 -0.023

-2 0.052 0.0175 2.9714 -0.022

-1.5 0.1044 0.0169 6.1775 -0.021

-1 0.1587 0.0166 9.5602 -0.02

-0.5 0.206 0.0154 13.377 -0.019

0 0.3244 0.0143 22.685 -0.03

0.5 0.4363 0.0147 29.68 -0.042

1 0.4978 0.0152 32.75 -0.043

1.5 0.5489 0.0155 35.413 -0.043

2 0.5999 0.0159 37.73 -0.042

2.5 0.6508 0.0163 39.926 -0.041

3 0.702 0.0167 42.036 -0.04

3.5 0.7524 0.0174 43.241 -0.04

4 0.8019 0.0178 45.051 -0.039

4.5 0.8519 0.0181 47.066 -0.038

5 0.9028 0.0185 48.8 -0.037

5.5 0.9523 0.0194 49.088 -0.037

6 0.9994 0.0199 50.221 -0.035

6.5 1.0481 0.0204 51.378 -0.035

7 1.0993 0.0207 53.106 -0.034

7.5 1.1456 0.0215 53.284 -0.033

8 1.1903 0.0221 53.86 -0.031

8.5 1.2401 0.0223 55.61 -0.031

9 1.2884 0.0228 56.509 -0.03

9.5 1.3289 0.0229 58.031 -0.028

10 1.3818 0.0225 61.413 -0.027

10.5 1.4143 0.0233 60.7 -0.024

11 1.4593 0.0232 62.901 -0.023

11.5 1.4931 0.0238 62.735 -0.02

12 1.5314 0.0236 64.89 -0.018

12.5 1.5581 0.0242 64.384 -0.015

13 1.5799 0.0247 63.964 -0.011

CLARK YM-15 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 0.0874 0.0128 6.8281 -0.087

-2.5 0.138 0.0122 11.312 -0.086

-2 0.1902 0.0118 16.119 -0.085

-1.5 0.2413 0.0114 21.167 -0.084

-1 0.2899 0.0108 26.843 -0.082

-0.5 0.3342 0.0101 33.089 -0.079

0 0.391 0.0101 38.713 -0.078

0.5 0.4694 0.0103 45.573 -0.081

1 0.5769 0.0105 54.943 -0.091

1.5 0.6577 0.0105 62.638 -0.096

2 0.7009 0.0106 66.123 -0.094

2.5 0.7418 0.0108 68.685 -0.091

3 0.7815 0.0111 70.405 -0.088

3.5 0.8187 0.0115 71.191 -0.084

4 0.8538 0.012 71.15 -0.08

4.5 0.8882 0.0126 70.492 -0.076

5 0.9263 0.0132 70.174 -0.073

5.5 0.966 0.0138 70 -0.07

6 1.0069 0.0144 69.924 -0.068

6.5 1.0471 0.0151 69.344 -0.065

7 1.0876 0.0159 68.403 -0.062

7.5 1.1296 0.0165 68.461 -0.06

8 1.17 0.0173 67.63 -0.058

8.5 1.2076 0.0181 66.718 -0.055

9 1.244 0.0188 66.17 -0.052

9.5 1.2785 0.0198 64.571 -0.049

GOE 741 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 -0.073 0.0193 -3.798 -0.015

-2.5 -0.02 0.0181 -1.105 -0.014

-2 0.0328 0.0176 1.8636 -0.013

-1.5 0.0863 0.0171 5.0468 -0.013

-1 0.1398 0.0167 8.3713 -0.012

-0.5 0.1933 0.0164 11.787 -0.011

0 0.2461 0.0163 15.098 -0.01

0.5 0.2975 0.0159 18.711 -0.009

1 0.3506 0.0158 22.19 -0.008

1.5 0.4622 0.0143 32.322 -0.019

2 0.6552 0.0146 44.877 -0.045

2.5 0.7058 0.0149 47.369 -0.044

3 0.7565 0.0151 50.099 -0.043

3.5 0.8072 0.0152 53.105 -0.042

4 0.8576 0.0155 55.329 -0.041

4.5 0.9071 0.0157 57.777 -0.04

5 0.9573 0.0157 60.975 -0.039

5.5 1.0056 0.0161 62.46 -0.038

6 1.0541 0.0162 65.068 -0.037

6.5 1.102 0.0165 66.788 -0.035

7 1.1479 0.0168 68.327 -0.034

7.5 1.1931 0.0171 69.772 -0.032

8 1.2367 0.0173 71.486 -0.031

8.5 1.2763 0.0179 71.302 -0.029

9 1.3162 0.0184 71.533 -0.026

9.5 1.3481 0.0192 70.214 -0.023

10 1.3771 0.02 68.855 -0.02

10.5 1.3997 0.0209 66.971 -0.015

11 1.4045 0.0222 63.266 -0.009

11.5 1.399 0.0236 59.28 -0.001

NACA3412 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 0.0129 0.0119 1.084 -0.079

-2.5 0.0723 0.0112 6.4554 -0.079

-2 0.1295 0.0106 12.217 -0.079

-1.5 0.1824 0.0099 18.424 -0.078

-1 0.2285 0.0091 25.11 -0.076

-0.5 0.2914 0.0086 33.884 -0.075

0 0.3993 0.0087 45.897 -0.085

0.5 0.4708 0.0088 53.5 -0.089

1 0.5184 0.0089 58.247 -0.087

1.5 0.5659 0.0091 62.187 -0.085

2 0.613 0.0093 65.914 -0.083

2.5 0.6594 0.0096 68.688 -0.081

3 0.7053 0.0099 71.242 -0.079

3.5 0.7494 0.0103 72.757 -0.076

4 0.7947 0.0107 74.271 -0.074

4.5 0.8406 0.0111 75.73 -0.072

5 0.8854 0.0116 76.328 -0.07

5.5 0.9306 0.0121 76.909 -0.068

6 0.9746 0.0125 77.968 -0.065

6.5 1.0153 0.0133 76.338 -0.063

7 1.0559 0.0141 74.887 -0.06

7.5 1.087 0.0155 70.129 -0.056

8 1.1084 0.0177 62.622 -0.051

8.5 1.1308 0.0196 57.694 -0.046

9 1.1552 0.0212 54.491 -0.041

9.5 1.1826 0.0226 52.327 -0.037

10 1.2108 0.0239 50.661 -0.034

10.5 1.2386 0.0253 48.957 -0.031

11 1.2638 0.0268 47.157 -0.028

11.5 1.2734 0.0296 43.02 -0.024

12 1.2675 0.0336 37.723 -0.019

12.5 1.2656 0.0377 33.57 -0.016

CLARK YM-15 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 0.0874 0.0128 6.8281 -0.087

-2.5 0.138 0.0122 11.312 -0.086

-2 0.1902 0.0118 16.119 -0.085

-1.5 0.2413 0.0114 21.167 -0.084

-1 0.2899 0.0108 26.843 -0.082

-0.5 0.3342 0.0101 33.089 -0.079

0 0.391 0.0101 38.713 -0.078

0.5 0.4694 0.0103 45.573 -0.081

1 0.5769 0.0105 54.943 -0.091

1.5 0.6577 0.0105 62.638 -0.096

2 0.7009 0.0106 66.123 -0.094

2.5 0.7418 0.0108 68.685 -0.091

3 0.7815 0.0111 70.405 -0.088

3.5 0.8187 0.0115 71.191 -0.084

4 0.8538 0.012 71.15 -0.08

4.5 0.8882 0.0126 70.492 -0.076

5 0.9263 0.0132 70.174 -0.073

5.5 0.966 0.0138 70 -0.07

6 1.0069 0.0144 69.924 -0.068

6.5 1.0471 0.0151 69.344 -0.065

7 1.0876 0.0159 68.403 -0.062

7.5 1.1296 0.0165 68.461 -0.06

8 1.17 0.0173 67.63 -0.058

8.5 1.2076 0.0181 66.718 -0.055

9 1.244 0.0188 66.17 -0.052

9.5 1.2785 0.0198 64.571 -0.049

SAUTER1 - Re = 250000

Alfa Cl Cd Cl/Cd Cm

-3 0.3356 0.0101 33.228 -0.078

-2.5 0.3937 0.0101 38.98 -0.077

-2 0.475 0.0104 45.673 -0.08

-1.5 0.5819 0.0105 55.419 -0.09

-1 0.6603 0.0106 62.293 -0.096

-0.5 0.7033 0.0107 65.729 -0.093

0 0.7441 0.0108 68.898 -0.091

0.5 0.7833 0.0111 70.568 -0.088

1 0.8203 0.0115 71.33 -0.084

1.5 0.8554 0.012 71.283 -0.08

2 0.8903 0.0126 70.659 -0.077

2.5 0.9281 0.0132 70.311 -0.074

3 0.9673 0.0139 69.59 -0.071

3.5 1.0086 0.0145 69.559 -0.069

4 1.05 0.0152 69.079 -0.067

4.5 1.0904 0.0159 68.579 -0.065

5 1.1318 0.0166 68.181 -0.063

5.5 1.1725 0.0173 67.775 -0.061

6 1.2098 0.0182 66.473 -0.059

6.5 1.2454 0.0189 65.894 -0.056

7 1.2794 0.0198 64.616 -0.054

7.5 1.3097 0.0207 63.271 -0.051

8 1.3378 0.0216 61.935 -0.048

8.5 1.362 0.0228 59.737 -0.044

9 1.3802 0.0244 56.566 -0.041

9.5 1.3943 0.0264 52.814 -0.038

10 1.3934 0.0296 47.074 -0.033

10.5 1.3889 0.0334 41.584 -0.029

11 1.3906 0.037 37.584 -0.027

11.5 1.3929 0.041 33.973 -0.025

Appendix C – Preparation of materials

Position Material Name thickness no. JOINT

Wing BALSA PLATE AILERON BACK 1.5 2 H- Tail: Balsa ribs in the middle 2 2 H- Tail: Balsa webbing 2 6 Wing BALSA REAR PLATE CENTRE 2 1 Wing BALSA EDGE PLATE REAR 2 2 Wing BALSA TYPE 2 2 4 Wing BALSA TYPE 1.1 2 4 Wing BALSA WEB 3 4 Wing BALSA WEB 3 4 Wing BALSA WEB 3 2 Wing BALSA WEB 3 4 Wing BALSA WEB 3 2 Wing BALSA WEB 3 1

H- Tail: PAULONIA back plate 1 1 H- Tail: PAULONIA elevator back plate 1 1

Fuselage PAULONIA Landing Gear Blocks 32x80x10 2 2 H- Tail: PAULONIA Ribs at 30 mm 2 2 H- Tail: PAULONIA Ribs connecting servo plate 2 1 H- Tail: PAULONIA ribs in the middle 2 3 H- Tail: PAULONIA elevator ribs 2 7 H- Tail: PAULONIA elevator ribs connecting servo 2 2 Wing PAULONIA EDGE MEDIUM 2 2 Wing PAULONIA EDGE BIG 2 2 Wing PAULONIA AILERON 2 8 Wing PAULONIA AILERON SERVO 2 4 Wing PAULONIA PLATE AILERON FRONT 2 2

Fuselage PAULONIA Battery Plate 3 2 Fuselage PAULONIA Nose Plate 4mm 4 2 Fuselage PAULONIA Landing Gear Blocks 32x80x10 4 4 H- Tail: Ply elevator front plate 1 1

Fuselage Ply Wing Plate Top 2mm 2 2 Fuselage Ply Wing Plate Bottom 2mm 2 2 Fuselage Ply Back Connection Plate 2mm 2 1 Fuselage Ply Wing Connection Block 10x10x32mm 2 8 Y

Fuselage Ply Carbon Rod Connection Block

10x40x56mm 2 4

Fuselage Ply Horizontal Tail Connectio Block

8x14x36mm 2 2 Y H- Tail: Ply Joint plate 2 2 Y H- Tail: Ply Joint plate 2 2 Y H- Tail: Ply Ribs connecting servo plate 2 1 Wing PLY TYPE 1.3 2 2 Wing PLY TYPE 1.1 FUSE 2 4 Y Wing PLY TYPE 1.1 FUSE 2 4 Y Wing PLY TYPE 1.1 2 2 Wing PLY TYPE 1.2 NEW 2 2

Position Material Name thickness no. JOINT

Wing PLY TYPE 2 SERVO 2 4 Wing PLY EDGE SMALL 2 2 Wing PLY SERVO BACK PART HOLDER PLATE 2 2 Wing PLY FUSE CONNECTION PLATE 2 1 Y Wing PLY FUSE CONNECTION PLATE 2 1 Y Wing PLY TYPE 1.1 POD 2 2 V- Tail ply Vertical tail 2.5 4 V- Tail ply rudder 2.5 4

Fuselage Ply Nose Rib 3mm 3 2 Fuselage Ply Fuselage Rib connecting Nose Rib 3 2 Fuselage Ply Battery Rib 3 4 Fuselage Ply Wing Rib 3 12 Fuselage Ply First Back Rib 3 2 Fuselage Ply Middle Back Rib 3 2 Fuselage Ply Tiny Rib 3 2 Fuselage Ply Centre Plate 3mm 3 4

Fuselage Ply Wing Rib Connection Block

18x40x32mm 3 24 Fuselage Ply Wing Connection Block 10x10x32mm 3 80 Y

Fuselage Ply Carbon Rod Connection Block

10x40x56mm 3 4

Fuselage Ply Horizontal Tail Connectio Block

8x14x36mm 3 4 Y H- Tail: Ply elevator servo plate 3 1 H- Tail: Ply servo plate 3 1 Y H- Tail: Ply servo plate 3 1 Y H- Tail: Ply webspar (3x6) 3 1 Wing PLY SERVO HOLDER PLATE 3 2 Y Wing PLY SERVO HOLDER PLATE 3 2 Y Wing PLY WEBSPAR CENTRE 3 1 Wing PLY WEBSPAR EDGE 3 2 Wing PLY WEBSPAR AILERON RIBS 3 2 Wing PLY FRONT CURVE PLATE 3 2

H- Tail: Ply side ribs 4 2

Longerons Fuselage Straight 5x5x760mm 6 Pine

Longerons Fuselage Back 5x5x272mm 4 Pine

Longerons Nose 5x5x52mm 2 Pine

tail connection to fuselage 6x6 v tail 6

spar caps(6x6) h tail 3 Pine

SPAR CAP WING 6x6 3 Pine

CARBON ROD wing 6mm 2

Carbon Rod fuselage D12x800 1

Wing Connection Bolt D3.5x60 4

Wing Rib Connection Bolt D4x40 2

wire

Appendix D – Control surface sizing

AILERONS

wingspan

(mm) 2300

wing chord

(mm) 260

aileron chord (% of wing

chord)

aileron chord (mm)

span (min) (% of wing

span)

span (max) (% of wing

span)

span (min) (mm)

span (max) (mm)

each aileron (min) (mm)

each aileron (max) (mm)

30 78 35 40 805 920 402.5 460

35 91 35 40 805 920 402.5 460

ELEVATOR

h tail chord

(mm) 180

h tail span

(mm) 763

chord (% of tail chord)

elevator chord (mm)

elevator span % of tail span

elevator span mm

30 54

90 686.7

35 63

95 724.85

RUDDER

v tail base

length (mm) 240

v tail height

(mm) 188

chord (% of tail chord)

rudder chord (mm)

rudder height %

rudder height mm

30 72

90 169.2

35 84

95 178.6

FINAL SIZING

chord (mm)

span/height (mm)

Aileron (each) 91 460 90 440

Rudder (each) 84 178.6 84 178

Elevator 63 724.85 63 743

Appendix E – Sketches from Team notebook

List of tasks and assignment areas

Kom

al S

Abhira

m R

Kiros L

Kate

R

Tanm

ay B

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。。。。。。

Co

ncep

tual D

esig

n

Synthesis of concept

Configuration design 3 3 3 3 3

Initial Sizing 3 1 2 3 1

Fuselage layout 0 0 3 0 0

Airfoil design 0 3 0 0 3

Wing planform design 3 0 0 0 0

Empennage design 1 0 1 3 0

Geometry for landing gear 0 3 2 0 0

Analysis of concept

Analysis of propulsion system 0 0 0 0 3

Weight and C.G. estimation 0 3 0 0 3

Aerodynamic performance analysis 0 3 0 0 0

Flight performance analysis 3 0 0 0 3

Stability analysis 0 3 1 1 0

CAD definition of the concept 3 2 3 2 2

Pre

lim

inary

D

esig

n

Structure layout , initial sizing, and internal layout

Wing structure 3 0 0 0 0

Fuselage structure 0 0 3 0 0

Empennage structure 0 0 0 2 3

Control surface structure 3 0 0 0 3

Landing gear 0 3 3 0 0

Integration of propulsion system 0 0 3 0 0

Payload arrangement 0 0 0 3 0

Structural analysis

Wing structure analysis 0 0 0 0 0

（continued）

Kom

al S

Abhira

m R

Kiros L

Kate

R

Tanm

ay B

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-

。。。。。。

Deta

il

Desig

n

Fuselage detail design 0 0 3 0 0

Wing detail design 3 0 0 0 0

Empennage detail design 0 0 0 1 3

Control surface detail design 3 0 0 0 3

Landing gear detail design 0 3 3 0 0

Control system design 3 3 1 1 0

Structure analysis for key parts 0 0 0 0 0

Fab

ricati

on

Preparation for materials 0 3 1 3 0

Fuselage fabrication 0 0 3 0 0

Wing fabrication (inner section) 3 0 0 0 0

Wing fabrication (outer section) 3 0 1 0 0

Empennage fabrication (inner section) 0 3 0 3 3

Control surface fabrication 3 3 1 0 0

Landing gear fabrication 1 3 0 0 0

Installation propulsion system in to airframe

and test 0 0 0 2 3

Control system installation and test 3 3 1 0 0

Assembly and test 3 3 2 0 0

Measurement of Weight and C.G location 3 0 3 0 0

Tests

Test plan 3 3 3 3 3

Ground tests 3 3 3 3 3

Air test (1) 3 3 3 3 3

Air test (2) 3 3 3 3 3

Note：The number is the indication of a student’s contribution to a specific task.

3 – primary contribution; 2 - secondary contribution; 1 – minor contribution; 0 – no contribution.