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VP-69: “The Black Wing” The A-Team

EPUAV Design and Build Project at NUAA, Nanjing

2012

The A-Team:

Joshua Richardson

Yasin Gulec

Matthew Moore

Edwin Cheah

Weibin Lin

Angus Chang

Nicholas O’Leary

Figure 1 - The A-Team

Contents Table of Figures ....................................................................................................................................... 5

Executive Summary ................................................................................................................................. 6

Statement of Purpose ............................................................................................................................. 7

Design Requirements .............................................................................................................................. 7

Performance ....................................................................................................................................... 7

Operation ............................................................................................................................................ 7

Cost ..................................................................................................................................................... 7

Conceptual Design .................................................................................................................................. 7

Design Considerations ........................................................................................................................ 7

Endurance ....................................................................................................................................... 8

Stability ........................................................................................................................................... 8

Complexity of Manufacture ............................................................................................................ 8

Aesthetics ........................................................................................................................................ 8

Flow Seduction Principle ................................................................................................................. 8

Key Design Objectives ......................................................................................................................... 9

Initial Sketches .................................................................................................................................... 9

Initial Sizing ........................................................................................................................................... 10

Wing Loading and Wing Size ............................................................................................................. 10

VTVC .................................................................................................................................................. 11

HTVC .................................................................................................................................................. 11

Accurate Sizing .................................................................................................................................. 12

Control Surface Sizes ......................................................................................................................... 12

Aileron - ......................................................................................................................................... 12

Elevator and Rudder - ................................................................................................................... 13

Calculated Values .............................................................................................................................. 14

Preliminary Design ................................................................................................................................ 15

Structure Layout ............................................................................................................................... 15

Fuselage ........................................................................................................................................ 15

Wing .............................................................................................................................................. 15

Empennage ................................................................................................................................... 16

Electronic Components ................................................................................................................. 18

Testing ............................................................................................................................................... 18

Propeller Testing ........................................................................................................................... 18

Endurance Testing ......................................................................................................................... 19

Calculations ....................................................................................................................................... 19

Profili ............................................................................................................................................. 19

AVL ................................................................................................................................................ 21

Detail Design ......................................................................................................................................... 24

Fuselage ............................................................................................................................................ 24

Wing .................................................................................................................................................. 28

Empennage ....................................................................................................................................... 28

Vertical Stabiliser .......................................................................................................................... 28

Rudder Control Surface ................................................................................................................. 31

Horizontal Stabiliser ...................................................................................................................... 32

Elevator Control Surface ............................................................................................................... 35

Full Empennag ............................................................................................................................... 36

Horizontal and Vertical Stabilizer Fins .......................................................................................... 37

Electronics Layout ............................................................................................................................. 37

Fabrication ............................................................................................................................................ 37

Fuselage ............................................................................................................................................ 37

Wing .................................................................................................................................................. 37

Empennage ....................................................................................................................................... 39

Vertical Stabilizer .......................................................................................................................... 40

Rudder Control Surface ................................................................................................................. 41

Horizontal Stabilizer ...................................................................................................................... 41

Elevator Control Surface ............................................................................................................... 41

Servo and Hinge Installation ......................................................................................................... 41

Horizontal and Vertical Fins .......................................................................................................... 42

Covering Film ................................................................................................................................ 42

Electronics ......................................................................................................................................... 43

Electric Motor ............................................................................................................................... 43

Electronic Speed Controller (ESC) ................................................................................................. 43

Propeller ........................................................................................................................................ 44

Electric Ducted Fan (EDF) .............................................................................................................. 44

Battery Eliminator Circuit (BEC) .................................................................................................... 44

Servos ............................................................................................................................................ 44

Lithium Polymer Battery ............................................................................................................... 45

Transmitter (TX) ............................................................................................................................ 45

Receiver (RX) ................................................................................................................................. 45

Electronic Schematic ..................................................................................................................... 45

Ground Testing...................................................................................................................................... 47

Ground Test 1.................................................................................................................................... 47

Ground Test 2.................................................................................................................................... 47

Ground Test 3.................................................................................................................................... 48

Flight Test .............................................................................................................................................. 50

Damage ............................................................................................................................................. 50

Causes ............................................................................................................................................... 50

Appendices ............................................................................................................................................ 51

A.1 Calculation spreadsheet ............................................................................................................. 51

A.2 VP-69 AVL File ............................................................................................................................. 52

The A-Team VP-69: “The Black Wing” Page 5

Table of Figures FIGURE 1 - THE A-TEAM ............................................................................................................................................... 2

FIGURE 2 - FLOW SEDUCTION ........................................................................................................................................ 8

FIGURE 3 - TOP VIEW OF INITIAL CONCEPT ....................................................................................................................... 9

FIGURE 4 - SIDE VIEW OF INITIAL CONCEPT ....................................................................................................................... 9

FIGURE 5 - ESTIMATION OF WING CHORD BASED ON WING LOADING ................................................................................... 10

FIGURE 6 - CONTROL SURFACE SIZING ........................................................................................................................... 13

FIGURE 7 - CALCULATED SURFACES ............................................................................................................................... 14

FIGURE 8 - CALCULATED VALUES .................................................................................................................................. 14

FIGURE 9 - FUSELAGE AND WING JOIN ........................................................................................................................... 15

FIGURE 10 - WING PRELIMINARY DESIGN ....................................................................................................................... 16

FIGURE 11 - EMPENNAGE PRELIMINARY ......................................................................................................................... 17

FIGURE 12 - PROPELLER TESTING .................................................................................................................................. 18

FIGURE 13 - LIFT WITH RESPECT TO DRAG ...................................................................................................................... 20

FIGURE 14 - LIFT AND DRAG WITH RESPECT TO ALPHA ...................................................................................................... 20

FIGURE 15 - LIFT/DRAG AND MOMENT WITH RESPECT TO ALPHA........................................................................................ 21

FIGURE 16 - AVL GEOMETRY ....................................................................................................................................... 22

FIGURE 17 - VALUES GIVEN THROUGH AVL .................................................................................................................... 23

FIGURE 18 - LIFT VISUALISATION .................................................................................................................................. 23

FIGURE 19 - REAR SECTION OF THE FUSELAGE ................................................................................................................. 24

FIGURE 20 - SIDE VIEW OF THE REAR OF FUSELAGE ........................................................................................................... 25

FIGURE 21 - FUSELAGE FRAME ..................................................................................................................................... 25

FIGURE 22 - FUSELAGE SPAR ....................................................................................................................................... 26

FIGURE 23 - FUSELAGE WING BOX SECTION ................................................................................................................... 26

FIGURE 24 - JOINING FRAME ....................................................................................................................................... 27

FIGURE 25 - VERTICAL STABILISER ................................................................................................................................. 29

FIGURE 26 - SPAR FOR VS ........................................................................................................................................... 29

FIGURE 27 - VS LAYOUT ............................................................................................................................................. 30

FIGURE 28 - JOIN BOX FOR EMPENNAGE ......................................................................................................................... 30

FIGURE 29 - RUDDER RIB ............................................................................................................................................ 31

FIGURE 30 - JOIN SPAR FOR RUDDER ............................................................................................................................. 31

FIGURE 31 - HS LAYOUT ............................................................................................................................................. 32

FIGURE 32 - HS RIB ................................................................................................................................................... 33

FIGURE 33 - HS SPAR ................................................................................................................................................. 33

FIGURE 34 - HS DESIGN .............................................................................................................................................. 34

FIGURE 35 - FULL EMPENNAGE .................................................................................................................................... 36

FIGURE 36 - WING BUILD ............................................................................................................................................ 38

FIGURE 37 - WINGLET BUILD ....................................................................................................................................... 39

FIGURE 38 - EMPENNAGE LASER CUT ............................................................................................................................. 39

FIGURE 39 - EMPENNAGE BUILD ................................................................................................................................... 41

FIGURE 40 - COMPLETE EMPENNAGE ............................................................................................................................ 42

FIGURE 41 - FIRST GROUND TEST .................................................................................................................................. 47

FIGURE 42 - SECOND GROUND TEST .............................................................................................................................. 48

FIGURE 43 - THIRD GROUND TEST ................................................................................................................................. 48

FIGURE 44 - GROUND TEST CLOSE TO TAKEOFF ................................................................................................................ 49


Executive Summary Students from RMIT were sent to NUAA on exchange to perform a design and build of an EPUAV

(electric powered unmanned aerial vehicle). The UAV built and described herein was labelled the VP-

69: The Black Wing.

The Black Wing is a single engine tractor driven UAV, with a major focus on aesthetics. The design

team were given requirements to fulfil in terms of size and weight of the aircraft. Within these

limitations the group went through all major stages of aircraft design.

The design of this aircraft was done using Catia software. Taking the 3D designs, the group used a

laser printing device at NUAA to produce the necessary components for the UAV manufacture.

Issues were found and solved during the design and manufacture process, as can be expected with

any engineering project.

The final aim of this process was to produce a UAV that flew within the limits of the design

requirements. The Black Wing was built and ready by test day and performed relatively well during

the flight. Unforeseen issues arose during flight, causing loss of stability and a stall. Following the

stall, the UAV impacted with the ground. The group took time to analyse the causes and have learnt

from this experience.


Statement of Purpose

The following report describes the process and systems taken to design and manufacture an electric

powered unmanned aerial vehicle (EPUAV); being the major task assigned to students in the “Project

Design of Aircraft” course at Nanjing University of Aeronautics and Astronautics (NUAA). The team

involved in the design of the EPUAV described here consisted of seven exchange students, and

collectively will be named A-Team. The task was set by the university to introduce the problems

faced by professional engineers in a way that the students would learn from these problems and

also apply the theoretical knowledge they had gained in study thus far.

Design Requirements

There were a number of requirements for this project by Professor Yu. These requirements were to

be adhered to in the same manner as an engineering team would do to the requirements of its

clientele. These requirements were given as follows;

Performance Endurance – Minimum flight time of 10 minutes.

Vmax – Maximum level flight velocity must not exceed 18 m/s.

Vmin – Stall speed must be below 10 m/s.

TOFL – Takeoff distance below 20 m.

Weight – Gross weight must not exceed 2.8 kg.

Payload – During flight a payload of at least 0.5 kg must be carried.

Operation Wing span – Span must not exceed 2 m.

Fuselage – Must not exceed 1.8 m.

Cost Material costs must not exceed 2000RMB.

Conceptual Design

Design Considerations In addition to the parameters mentioned under ‘Design Requirements’ as defined by the NUAA

course instructors, A-Team decided to begin conceptual design by setting considerations that were

important to the group itself. These considerations were to guide decision making and to help define

the major task in a way that would streamline the early design process. The following considerations

were first labelled as important, and used as a basis for configuration design; these considerations

overlapped the requirements of the project at times, however this did make clear to A-Team the

characteristics that were most important.


Endurance

A-Team highlighted this as a major consideration when designing the aircraft. It was hoped

that this would be a key area that the EPUAV could exceed the basic performance requirement set

through the course. This factor did not play a large part in early design, at least not alone; the aim of

exceeding the endurance requirement did lead to discussion on reducing drag in early conceptual

design.

Stability

This project is the first major design and build that the students have undertaken, as such it

was seen that stability should be a major focal point. Designing an aircraft that would have natural

stability in flight would help overcome some issues concerning the flight test and the initial

calculations. There was a lack of experience in piloting aircraft of this nature, so stability would have

also helped with the final performance flight.

Complexity of Manufacture

A-Team considered the manufacturing process very early on in design. Prior to completing

the conceptual layout, most group members had read through the previous design submissions for

students completing a similar task. A major stumbling block that many groups had faced was the

difficulty in fabrication. It became very clear that any design discussion, not only during the initial

stages but throughout the completion of the project, would always take in to consideration the ease

with which parts could be manufactured.

Aesthetics

The belief that the EPUAV should be visually appealing was shared amongst all members

from the beginning. This helped to form a few concepts during the conceptual stage and also helped

at times to discern between equally effective methods of construction. Further to this, the

discussion on aesthetics and the application of aerodynamic principles led the group to create a

concept that has come to be known as the Flow Seduction Principle.

Flow Seduction Principle

The modern climate of aircraft design requires an aircraft that is visually pleasing as well as

meeting the expectations of performance. In order to achieve these requirements, the design

process needs a balance of ideas and concepts. The Flow Seduction Principle is the integration of

aesthetics in to the known and working principles of aerodynamic studies. In the application of this

project, it was seen to include the shape of the fuselage and the extremities of the wing and

empennage. To this end, the conceptual design was to focus on the smooth shape changes in the

body and the flared winglets and tail pieces. The below image gives an indication of the curvature

desired when applying the Flow Seduction Principle to aircraft design; as is seen, the fuselage tapers

in a smooth dual-asymptotal curve.

Figure 2 - Flow Seduction


Key Design Objectives Following discussion on the major considerations the group saw as important, the configuration was

discussed. The major components of configurations needed to be decided; this would include

positioning of the wing, the shape and layout of the empennage, sweep and dihedral of the wing,

landing gear positioning and layout and positioning of the propulsion system. The following were

chosen as the key configuration objectives for the EPUAV:

High Wing Placement – This would give natural roll stability.

Possible Dihedral Wing – Further adding to roll stability. This would later be

discarded as the high wing placement was decided to be enough.

Conventional or T-Tail – The empennage was not decided upon until later in

the conceptual design process. The final concept took a conventional layout;

this was to avoid overloading the tail with moments from the horizontal

surface, and also easing construction.

Tractor Driven – Placing the propulsion in the tractor position gives a cleaner

flow, and also keeps the fuselage construction to something manageable in

the time limit the team had.

Initial Sketches After deciding on these objectives, the following sketches were initially made. These show all major

considerations in the layout. Sizing for these sketches was based upon appearance, with only the

wingspan being maximised as early as possible.

Figure 3 - Top View of Initial Concept

Figure 4 - Side View of Initial Concept


Initial Sizing The following calculations were initial estimations of sizing for the UAV. These were performed using

rough values for loading, gross weight and volume coefficients for both stabilisers.

Wing Loading and Wing Size

m0 – normal take-off weight (kg)

g – Gravity (m/s2)

p0 – wing loading (10 N/m2)

S – Wing area (m2)

The parameters for this project state that 3kg is the maximum available for weight. In our

calculations we added another 1 kg to account for payload and a marginal safety factor. The

wingspan in our conceptual design was 2m, so this is the value we have used. Professor Yu had

suggested that 4.6 kg/m2 is the approximate wing loading based on historical data. After group

discussion and research on previous students’ UAV projects it was decided that the wing loading we

would work with for initial calculations would be between 6 and 9 kg/m2. Using these estimations

the below table of values was created. With these values we have decided to pursue a wing chord of

0.25m, giving us a maximum wing loading of around 8kg/m2 and an aspect ratio of 8. This fit with our

conceptual design which was to have an AR of between 7 and 9.

weight (kg) 4

wing span (m) 2

wing loading (kg/m2) 5 6 7 8 9 10

wing area (m2) 0.8 0.666667 0.571429 0.5 0.444444 0.4

wing chord (m) 0.4 0.333333 0.285714 0.25 0.222222 0.2

AR 5 6 7 8 9 10 Figure 5 - Estimation of Wing Chord based on wing loading


VTVC

Sv – The reference area of the vertical tail

lv – The distance from cg to the tail cg

S – The wing area

b – Wing span

The conceptual design for this project was based on something similar to a single engine general

aviation aircraft; according to Raymer (1992) the vertical tail volume coefficient for this type of

aircraft is 0.05, however after discussion with the team and on the advice of Professor Yu it was

decided to use 0.06 for initial calculations.

Using the results from the table above, and the estimation of the volume coefficient for the vertical

tail; the sizing of the vertical stabiliser was calculated as below.

Raymer (1992) suggested an AR for the vertical stabiliser of 1.3-2, after discussion the team decided

to go with an AR of 1.7. The chord of the fin was found to be 0.206m, and the span of the vertical

surface was 0.351m.

HTVC

SH – the reference are of the horizontal surface

lH – the distance from the tail cg to the cg

S – The wing area

c – The mean wing chord


As with the vertical tail, we based our horizontal tail volume coefficient on the single engine general

aviation aircraft. Raymer (1992) suggests 0.65 for this coefficient, though we have increased this to

0.75 on the advice of the professor to scale down to an RC aircraft.

With the above calculations for the sizing of the wing as well the estimation of the volume

coefficient we calculated, following the same procedure as the vertical tail, the reference area of the

horizontal tail as below.

Raymer (1992) suggested an AR for the horizontal stabiliser of 3-5, after discussion the team decided

to go with an AR of 4. The chord of the tail was found to be 0.172, and the span of the horizontal

surface was 0.689m.

Accurate Sizing After discussion with the teaching assistants and a lengthy discussion with the group, we decided to

recalculate the size values. For the wing, we used 1.9 as our wingspan to account for our wingtip

design. The volume coefficients for all surfaces were altered under the new sizing. The calculated

values are given here, with the associated table provided in the appendices.

Control Surface Sizes

Aileron -

The size of the ailerons was found using the following equation:

Where:

Va – Aileron volume coefficient

Sa – Area of aileron

la – Distance between midpoints of two ailerons

S – Wing Area

b – Wing span

Distance between the ailerons is found via a constant, given by the below table, and the equation

following.


Figure 6 - Control Surface Sizing

This gives an area of 297.3 cm2 for each aileron. To cover this area a chord of 10 cm was selected,

this gives a ratio of 0.43, which is slightly higher than Raymer (1992) has suggested; our advice from

the Professor and the teaching assistants has been to scale all surfaces up compared to full size

aircraft. The ailerons were then 10 cm2 x 30cm2.

Elevator and Rudder -

No set equation was given for the sizing of the elevator and rudder. Based on historical data, A-Team

was informed to use a chord length of approximately 0.45 of the main surface, and a span of 0.9 of

the main surface. This produced the following values:


Rudder

c 0.0675

b 0.27

Elevator

c 0.1

b 0.36 Figure 7 - Calculated Surfaces

Calculated Values The following table presents the final calculated values obtained during the initial sizing of the

aircraft.

Value Value

W/S 9 λ 0.904

b 1.9 cr-vert 0.316

S 0.444 ct-vert 0.158

c-bar 0.234 bvert 0.402

Sv 0.095 cr-hor 0.2

Sh 0.09 ct-hor 0.1

Sa 0.0594 2bhor 0.6

ca 0.1 cgx 0.4232

ba 0.3

Figure 8 - Calculated Values


Preliminary Design

Structure Layout

Fuselage

The earliest design for the structure of the fuselage contained two major spars supporting a set of

frames. This design was seen as conventional and best represents the reproduction of a full-sized

engineering project. This was used as the basis for all structural design decisions.

During conceptual design it was decided that the layout would contain a high wing position, allowing

the control of stability. In order to achieve this, in terms of manufacture, A-Team decided to place

the wing within the fuselage as opposed to resting on top.

Figure 9 - Fuselage and wing join

A wing box was to be used as the mounting point for the wing, with major frames and a flat, solid

mounting to screw the wing to. This wing box would need to resist the load given through lift as well

as the torsion created during rolling.

The desired curvature of the fuselage was to be attained using the frames, the shape and size of

which would follow the asymptotal curve defined by A-Team.

Wing

As with the other major components, a major consideration for the layout of the wing was the

overall appearance. To this end, ribs were to be placed close enough together to retain the smooth

shape of the outside skin. It was decided that two major spars were all that would be needed to give

span-wise strength to the structure. Under the advice of Professor Yu, the ribs were placed no more

than 50mm apart. In sections where increased strength would be needed, they would be placed

slightly closer, with a thicker and stronger material applied in these sections.


Figure 10 - Wing preliminary design

The above image shows the preliminary layout for the wing. Exact rib shape and internal structure

was not defined at this stage, however rib placement and the spar designation was close to

finalisation. After analysing this design, a number of changes were to be made with respect to

function and fabrication.

1. The main spars were to extend vertically from the lower surface of the wing to the upper

surface. The purpose of this change was to help the skin adhere to the wing structure

and retain the airfoil shape.

2. The rear spar was to be shifted towards the leading edge. This would leave the spar at

the same position as the beginning of the aileron, preventing the need for the spar to be

broken at this point, increasing strength and continuity throughout the wing.

3. A carbon rod was to be placed on the trailing edge. The trailing edge as it is shown above

has a drastically sharp point; this sharpness would easily pierce the film to be used as

the skin for the UAV. A rounded carbon rod would easily take the filming, and keep the

airfoil shape.

Empennage

The preliminary design of the empennage section of the aircraft was focused on the stylising of the tail section of the aircraft and its integration into the main fuselage. Preliminary design also focused on the type of airfoil required for the tail control surfaces. Conventional control surfaces utilize symmetrical airfoils that are designed only for control, not lifting airfoils. Typical choices include the symmetric NACA 00XX series. Both the vertical and horizontal stabilisers were chosen to use the NACA 0015 airfoil for a large amount of control from the empennage. The tail volume coefficients were used to determine the size scaling of the empennage. The volume coefficients were scaled up as that the UAV is considerably smaller than a full sized aircraft yet the airflow remains the same. The sizes of the control surfaces were also scaled up to increase the amount of control of the aircraft at lower cruising velocities. Initial empennage concepts focused on an anhedral horizontal stabiliser to counter any possible Dutch roll induced by the high main wing and winglets. Early concepts had a ‘shark fin’ vertical surface and rectangular horizontals. Later concepts included the shark fin design on the horizontals.


Preliminary design saw the removal of the anhedral on the horizontal stabilisers, a conventional taper added to both sets of stabilisers and winglet-type fins added to the edges to replicate the shark fin design in its earlier stages. The method of integrating the horizontal surfaces onto the vertical stabiliser, and then attaching the entire empennage onto the tail of the fuselage was still in discussion until the beginning of the detail design phase where the wing box structure and the carbon rod connections were introduced into the design.

Figure 11 - Empennage preliminary


Electronic Components

Testing Two sets of experimental testing were performed; the first was to determine the most efficient

propeller to use with the given motor and battery combination, and the second was to test for the

endurance of that propeller and motor set.

Propeller Testing

A-Team set out to perform testing on various sizes of propellers, with the purpose of determining

the most efficient propeller throughout a range of current draws; efficiency being defined by A-

Team as the most thrust for the least current drawn from the battery.

Propellers are defined by two numbers, their diameter and pitch respectively. Diameter is simply

the total length of the propeller; important in ground clearance considerations. Pitch is the angle of

attack of the propeller; the number associated with the pitch is the ideal travel distance of the

propeller, in inches, given one full rotation of the propeller.

Over a range of current draws, set with a servo-tester and a multimeter, a static thrust force was

measured using handheld electronic scales and is measured in the equivalent number of kilograms.

Figure 12 - Propeller testing

The above graph shows the static thrust generated over current drawn by the motor for four

propellers. As can be seen, the 12x6e propeller provides the best performance over the full range of

the test for all propellers tested. Theoretically, the 12x8e should be capable of producing greater

performance, however, the 12x8e places unnecessary strain on the motor which diminished

performance. This strain on the motor would more than likely lower the endurance of the system;

0.00

0.20

0.40

0.60

0.80

1.00

1.20

1.40

1.60

1.80

0.00 5.00 10.00 15.00 20.00 25.00 30.00 35.00 40.00 45.00

Me

asu

red

Sta

tic

Thru

st (

Kg)

Current Draw (Amps)

Propeller Testing

12x6e

11x7e

10x7e

12x8e


although that was not the concern of this test, the 12x6e outperformed the 12x8e regardless of

theoretical understanding.

Endurance Testing

After selecting the 12x6e propeller for use, A-Team undertook a test on the endurance of the

propeller with the given motor and battery set. Under the advice of Professor Yu, a thrust to weight

ratio of 0.4 was used to estimate the required thrust for this test. The experimental set up was the

same as for propeller testing. With the assumed weight of approximately 4kg for the total weight of

the UAV, a thrust of 1.6 was used during testing. When this thrust was reached on the electronic

scale, the motor was left running under the relative current draw. After a period of 11 minutes at

this level, the thrust had dropped to only 1.5 kg, and a total time of 14 minutes was reached before

any significant amount of thrust had been lost; by this indication the propeller chosen would

perform the required endurance with ease.

Calculations Further to the testing performed by A-Team, a series of calculations were used on two pieces of

software. The first was Profili, an airfoil database that analysis the shape and characteristics of an

airfoil. Profili can also analyse the structure of a wing, however this feature was not used during this

design. The second software was AVL, a vortex-lattice program used to analyse airflow over a body

and a wing structure. The purpose of this software was to test for stability based on the current

design, and identifying the best way forward.

Profili

Airfoil selection was needed at this stage, as the wing was seen as the most crucial component of the

UAV. It has been stated the aesthetics was a major factor in the design of The Black Wing, the

functionality of the wing and the selection of the airfoil was seen by A-Team to be important to the

success of the final flight. To this end, various airfoils were researched using a number of databases

found online.

The key characteristics identified by A-Team in the selection of an airfoil were:

1. Stall angle

2. Maximum CL

3. Attack angle at cruise CL

After 4 airfoils had been identified as satisfying these factors, they were analysed using the software

Profili. This software would give lift, drag and moment values for an airfoil based on attack angle and

Reynolds number. Re for the UAV was calculated to be close to 300000, this value was used for the

Profili analysis. Shown below are the results of these calculations.


Figure 13 - Lift with respect to Drag

Figure 14 - Lift and Drag with respect to alpha


Required CL at cruise condition was identified to be around 0.5. All 4 of these airfoils achieved this

value at an alpha of 0, leading to further analysis. The CL/CD curves shown previously and the

CD/alpha curve identify the pair of GOE airfoils as most appropriate. The CM curve in particular shows

that they are laterally more stable than the others. In terms of deciding between the two GOE

airfoils, it was seen as negligible the difference between them. This was at least the view of the

group for the purpose in this project. A-Team eventually selected the GOE 527 as the airfoil for the

main wing.

Figure 15 - Lift/Drag and Moment with respect to alpha

AVL

The stability of the UAV was estimated using the program AVL, a vortex-lattice software for stress

and flow analysis. Applying the conceptual design to this program, with the selected airfoils and

rough fuselage shape, the required placement of CG for stable flight and a guide to CL at various

angles of attack could be attained. Shown below is the geometry presented after entering the coding

for the VP-69 in to AVL. The winglets and taper of the wing can be seen, as well as the curvature of

the fuselage.


Figure 16 - AVL geometry

Stability coefficients can be taken from AVL by entering variables such as attack angles, flow velocity,

mass and CG positioning. These values were run, and a series of results shows an approximate

position of CG to be close to the mid-chord position. Exact value was calculated using the equation

given for static margin:

The AC was found through the AVL calculations, as the CG was shifted until a moment of 0 was

achieved on the aircraft; for stability however, the moment should be slightly forward (nose down)

to aid in recovery and to prevent unexpected wing stall. The static margin suggested for use by

Professor Yu was 12-15%. A static margin of 15% was used in this approximation, giving a position

for CG at 120mm from the leading edge of the wing.

AVL also helped with the visualisation of lift, lift distribution and drag across the UAV. The following

images present these visualisation steps. The changing of attack angles on the entire aircraft allow

for calculation of CL during various stages of flight. When this was combined with the results from

Profili V2 the angle of incidence of the wing with respect to the fuselage could be chosen.


Figure 17 - Values given through AVL

The overall lift distribution of the VP-69 can be seen in the following image. The results shown above

give an Oswald’s efficiency of approximately 0.95, although the shape shown below is not quite as

elliptical as would be preferred. The shape of the lift distribution, regardless of the given result, was

seen as satisfactory for the purpose of this design.

Figure 18 - Lift Visualisation

AVL gave A-Team the ability to analyse the conceptual and preliminary design with practicality in

mind. With the results attained during this stage, the decision to continue to detail design was

made, with a view to begin fabrication as soon as was viable.


Detail Design

Fuselage The fuselage was designed with two main spars or “side plates” running through as much of the

fuselage as possible, these spars take axial and bending loads, as well as torsion when coupled with

the interior torsion boxes.

Figure 19 - Rear section of the fuselage

Fuselage main spar dimensions

Height 63(mm)

Length 1152(mm)

Thickness 4(mm)

Distance Apart 40(mm)

The spar is broken roughly every 100mm with a frame, these frames give the fuselage its shape as

well as supporting some of the loads that the fuselage experiences.

The large height (63mm) of the spar allows the spar to support all of the back of the fuselage, as the

diameter of the fuselage grows forward; a second underside spar becomes necessary to stop the

bending of the fuselage. Due to the size of the spar, much of the risk of breaking is impeded; the

underside spar can then be simply used for contouring the shape and to maintain rigidity.

The main fuselage spar contains several elliptic weight-saving holes. These holes exist between the

frames for most of the length of the fuselage, except for the last few frames, where the empennage

would attach; it was deemed necessary for the spar to have the extra strength.

In the translation from the drafting of the spar to the .dxf file as read by the laser-cutter, an error

occurred which resulted in the loss of two weight holes.

The spars were designed as a pair, 40mm apart, in order to maximize the bending strength in the

lateral direction. A small gap exists in the fore-most spar for the same reason, but this also had the


added benefit of creating a small compartment for the battery or payload, at the very front of the

plane.

The frames that form the shape of the fuselage vary in size along the length of the rear of the

fuselage, giving the plane a distinctive sweep.

The join between the frames and the spar is a simple half-lap joint, so that the top of the spar lies as

flush as possible against the top of the frame. As the frames grow in size, the spar rests higher up the

frame, giving more free space on the frame for weight holes, other spars and stringers.

Figure 20 - Side view of the rear of fuselage

Figure 21 - Fuselage Frame

As all the frames along the fuselage were different, the shape of each one had to be individually

catered. The holes themselves were designed conservatively, favouring possible strength over the


desire for weight savings. Each of the corners in the weight holes is chamfered to some extent;

distributing the load and minimizing the risk of stress cracks.

The Lower Portion of the fuselage is supported by two lower spars, designed to stop bending on the

fuselage.

Figure 22 - Fuselage Spar

The lower spar is 829mm long and extends from under the wing to the beginnings of the contouring

tail. The lower spar stops well before the end of the wing because of the large size of the upper

fuselage spar; it was deemed unnecessary to include two spars to prohibit bending when the large

upper spar would suffice.

In order to allow this, the main spars were to run stop and a second, wing-box spar was added as

well as two, much thicker frames to accommodate the loads provided by the wing.

Figure 23 - Fuselage Wing Box Section

The thicker frames were implemented because of the large forces exerted by the wing as well as the

landing gear, which was positioned inside the second large frame. It was also anticipated that either


a battery or the payload may be positioned here, but due to concerns over the position of the centre

of gravity, no payload was supported under the wing.

In the positions where two types of spars came together, a special joining frame was used which

could accommodate the two spars.

Figure 24 - Joining Frame

The shape of the fuselage was designed with a rounded front, which would allow the air pushed

behind by the motor to flow over it without causing excess drag and prevent the fuselage from

blocking off potential thrust. After this rounded off section the fuselage would retain a cylindrical

shape with a diameter of 150 mm. This would allow sufficient storage space for the various

components of the aircraft.

The front of the fuselage was designed bearing in mind that it would hold a large amount of weight.

It acts as a storage compartment for electronics including；the ESC，the battery，the receiver and

payload.

In order to allow room and strength to hold these components the frames were designed with a

large amount of space taken out from them. They were also made using double thickness ply wood,

4 mm, so that they would be able to be strong enough to act as a container for the various

electronics and payload. The frames were designed to be approximately 10 cm apart from one

another to ensure an optimal strength to weight ratio, these frames were connected by two spars

running throughout the front of the fuselage. These spars were of a 4mm thickness and were used

to withstand the bending, tension and compression loads.

Two shelves were placed as a placed as a system to hold the electronics and strengthen the nose of

the aircraft, however in the end they were not used for the electronics and acted purely as structural

support.

Also located in the nose of the fuselage is the front landing gear and nose gear servo.


The nose gear and landing gear were placed directly above one another so that the plane would

have greater manoeuvrability on the ground, as any degree of turning performed by the nose servo

would result in an identical turn for the wheel. This also helped with ease of manufacture as the

push rod was a simple rod vertically connecting the servo and landing gear. The nose gear was

mounted on a sheet of 2mm ply wood that ran across the top of the front two frames, between the

spars. The landing gear was mounted to a plate of 6mm thickness placed on the front frame. This

meant the landing gear would have more strength if it came under a rough landing.

Placed at the very front of the fuselage is a firewall to hold the motor. The motor needed to be

mounted to a firewall to ensure that the motor wouldn’t rip out when it produced a large amount of

thrust. The firewall was made out of 6 mm thickness ply that would ensure that the motor would be

tightly secured. In the firewall a total of 7 holes were placed; four for screws, one for the shaft that

sticks out behind the motor, another for wires and lastly a hole that would allow air to flow through

the fuselage acting as a cooling system for all the electronics.

Finally to keep the aesthetic geometry of the aircraft, stringers were input into the design process.

These were not influential to the structure of the plane and were expected to take minimal to no

loads. Ensuring that they were light was the top priority. The ideal size and material was determined

to be stringers composed of a 5x10mm balsa.

Wing

Empennage Much of the design of the empennage revolved around the integration of the horizontal stabilizers into the vertical stabilizer. Carbon fibre rods were used in order to increase the strength of the vertical stabilizer and link the horizontal stabilizers together through the central wing box.

Vertical Stabiliser

The vertical stabiliser was designed so as that the ribs easily sit in slots on the trailing edge plate, which in turn automatically sets the height so the fabrication can become relatively simple.


Figure 25 - Vertical Stabiliser

This is the training edge stabiliser spar. This structure is designed as the backbone of the entire vertical stabiliser. Note the notches along the entire surface – they act as the calibration to achieve the correct height on each of the ribs in the vertical to obtain the desired degree of taper. The two closely spaced notches at the lower portion of the spar are spaced so as that the main wing carry through box can be sandwiched between the ribs. The middle two ribs spaced closely together are so as the servo can easily slot between the two.

Figure 26 - Spar for VS


The wing box in the centre of the vertical stabiliser was designed in order to secure the horizontal stabilisers onto the vertical stabiliser and efficiently transfer the loading from one to the other. Its construction will involve gluing layers of laser cut ply material together whilst the angled hole in the top will need to be hand milled in order to fit the carbon rod in the leading edge of the vertical stabiliser.

The entire vertical stabiliser goes together as follows: the ribs slot into their respective positions on the trailing edge spar piece, the wing box is installed, the leading edge carbon rod is put in place and the stringers are cut to size and glued in. The trailing edge carbon rod is left to be installed once the entire of the empennage is integrated into the rear of the fuselage.

Figure 28 - Join box for empennage

Figure 27 - VS layout


Rudder Control Surface

The construction of the rudder is designed to be extremely simple and facilitate the installation of metal hinges in order to allow the control surface to move freely during ground taxiing and flight testing.

The design of the rudder only calls for one main structural member, and is meant to be light. This main spar structure allows the rudder ribs to simply slot on and lock into place. The ribs are spaced evenly along the length of the spar to allow for the even distribution of aerodynamic loading when the rudder control surface is under deflection.

The simple design of the rudder ribs allows them to slot on the main rudder spar relatively easy. Note the circular notch placed in the trailing edge of the rib – this is designed to facilitate the integration of a carbon rod acting as the trailing edge stringer to prevent the ribs from moving during flight and making the application of the thermal contact film much easier.

Figure 30 - Join spar for rudder

Figure 29 - Rudder rib


Horizontal Stabiliser

The entire of the horizontal stabiliser has a similar design to that of the vertical stabiliser: several carbon rods are designed to bear most of the forces induced by flight along with a trailing edge spar and stringers. The exact same hinging mechanism is also used.

Figure 31 - HS layout


This structure exactly replicates the trailing edge spar of the vertical stabiliser – the notched automatically set the height of the ribs. The two closely spaced ribs also allow the facilitation of the servo to control the elevator surface.

The ribs of the horizontal stabilisers are also similar to that of the vertical stabilisers – note the circular holes designated for the carry-through carbon rods and the ovular holes designed for weight reduction within the stabiliser. The notching for the stringers and the trailing spar piece are also present, as also seen in the ribs of the vertical stabiliser.

Figure 33 - HS spar

Figure 32 - HS rib


Both of the horizontal stabilisers are fabricated in the following manner: the ribs slot into their respective positions on the trailing edge spar piece, the stringers are cut to size and glued in then the two carbon rods are cut to size once the entire empennage has been mocked up. The carbon rods can be installed at the final phase of fabrication when the entire empennage needs to be finalised.

Figure 34 - HS design


Elevator Control Surface

As with the design of the rudder control surfaces, the elevators are designed to be extremely simple and facilitate the installation of metal hinges in order to allow the control surface to move freely during ground taxiing and flight testing.

The design of the elevator only requires one main structural member, and is meant to be light. This main spar structure allows the elevator ribs to simply slot on and lock into place. The ribs are spaced evenly along the length of the spar to allow for the even distribution of aerodynamic loading when the elevator control surface is under deflection.

The simple design of the elevator ribs allows them to slot on the main rudder spar relatively easy. As with the rudder control surface ribs, note the circular notch at the trailing edge to allow for a carbon rod. Note also that the root elevator rib is marginally shorter that all the others – this is designed so that the rudder will not impact the elevator when it is deflecting. The wedge cut-out of the elevator allows for the maximum deflection of the rudder without compromising the elevator.


Full Empennag

The full empennage is designed to be put together by stringing the two carbon rods through the central wing box, passing fully through and reaching their respective ribs. At this stage it may be possible to glue the carbon rods in place as well as install the servos and hinges. Note that the trailing edge vertical stabiliser carbon rod is designed to be installed once the empennage has been installed onto the fuselage in order to provide a solid and efficient connection between the two.

Figure 35 - Full empennage


Horizontal and Vertical Stabilizer Fins

The horizontal and vertical stabiliser fins are designed to be purely aesthetic and are in no way

designed to take any aerodynamic loading. The vertical fin was designed to sit on the trailing edge

carbon rod of the vertical stabiliser and hence needs to be installed at the end of fabrication. The

horizontal fins are designed to be simply glued onto the end of the horizontal stabiliser.

Electronics Layout

Fabrication

Fuselage During the fabrication phase of the project, the fuselage underwent many changes compared to the

original design. The frames near the wing, in the section called the wing box, were quadrupled for

added strength. The two spars that ran the length of the body were also doubled for strength, as this

was seen as an important feature. During all fabrication stages, strength was seen as the most

important factor by the build team.

After the fitting of the frames to the spars, the fuselage appeared longer than the original concept.

At some point during design, 300mm was added to the total length of the fuselage. It was believed

that this occurred due to the designation of two designers of the fuselage. The resulting construction

led to a removal of the last two sections, the two frames and the spar piece attached. This

shortening not only created something more like what was expected, but also shifted the CG

towards the front of the UAV.

There was another shortening of the fuselage, via the same process, due to the realisation that the

CG was further towards the rear than was anticipated. This removal of sections was needed and

finally produced the expected fuselage length; however the curvature had changed due to the loss

of frames and attached spars.

Wing The wing was the first component constructed, and as such was the first to encounter trouble. The

main issue during construction for all components was the size of the material. Ply was used for all

major components, and the thickness of this ply was between 2 and 3 mm. During design it was

expected that the group would have material of 2 and 4 mm in thickness, and this was the pieces cut

via the laser cutter. When the group realised that the thickness was either two small or two large for

the designated holes, a process of sanding and boring of all slots and holes was undertaken. This

process was applied to all components during fabrication.

Shown below is the setup of the wing during fabrication. The two main spars can be seen, as well as

the ribs throughout. The stage shown is prior to the addition of the carbon rods, and also prior to

the connecting of the entire wing span.


Figure 36 - Wing build

After the main wing had been constructed, the group focussed on the build of the winglets. Below is

shown the original construction of these winglets; however this was changed during filming. The

change between the wing and the winglet was too aggressive for the plastic film to cover it

adequately, and as such the winglets were placed closer to the wing tip. This choice did not impact

on the effect of the winglets, as they were seen as an aesthetic component.


Figure 37 - Winglet build

Empennage The fabrication of the empennage section of the UAV was relatively simple, yet hindered slightly by the change in thickness of design material and the inaccuracies of the laser cutting machine. The fabrication of the empennage was performed in stages, as that several components are dependent on each other and cannot be completed before any others. The entire empennage was able to be cut out of a single sheet of 3mm ply material with the laser cutting machine. Cutting duration took 30 minutes are resulted in most of the parts being cleanly cut out from the ply material. Some parts, namely the vertical stabiliser ribs, had to be either cut free

Figure 38 - Empennage laser cut


from the ply sheet or recut entirely as the laser’s power varied during cutting. The stringers for the empennage were manually cut from separate strips of material. Most of the parts that had to be slotted together had to be milled and cleaned before they could be put together due to the variation of the thickness of the material used for fabrication.

Vertical Stabilizer

The design of the vertical stabiliser allowed the ribs to slot into the trailing edge plate piece and automatically have their height set. The central wing box designed to join the two horizontals was then cut, glued and milled before being installed into place. The leading edge carbon rod was then strung through this wing box to securely lock all of the vertical ribs into place. The stringers were then cut by hand to ensure they were the correct length and installed to complete the exterior of the vertical stabilizer. The balsa sheet around the leading edge of the stabilizer was cut and glued into place. The trailing edge carbon rod was installed after the empennage was placed on the fuselage in order to secure them together.


Rudder Control Surface

The rudder control surface was easily fabricated to the simplicity of design laid out during the detail design phase of the project. The rudder ribs slot onto the rudder spar piece and glued into place. The design of the ribs placed a notch in the trailing edge so as that the 2mm diameter rod of carbon could be placed for reinforcement and to make applying the covering film easier. Balsa sheets were glued over the trailing edge to also help with the covering film. A solid plate of ply material was glued between two of the rudder ribs to allow placement of the servo horn. Lengths of light balsa material were added to allow the installation of the metal hinges.

Horizontal Stabilizer

The horizontal stabilizers were constructed in a similar manner to that of the vertical stabilizer. The main difference between the two designs was that the horizontals were designed without a main spar but with two carbon rods running all the length between the horizontals and the vertical stabilizers. Both of the horizontal stabilizers were constructed in the same method as the vertical stabilizer before they were connected to the vertical via the two carbon rods strung through the central wing box.

Elevator Control Surface

The construction of the two elevator control surfaces was remarkably similar to that of the rudder control surface. Elevator ribs fit into the slots on the elevator spars and were glued into place. Again, notches allowed the installation of a trailing edge carbon rod for strength and ease of applying the covering film. Balsa sheets were also glued to the trailing edge for the covering film.

Figure 39 - Empennage build

Servo and Hinge Installation

The initial design of the empennage required the use of metal hinges for both elevator and rudder control surfaces. After the construction of the empennage it was found that it was too heavy, and after some deliberation it was decided that some of this additional weight was attributed to the use of the metal hinges for hinging the control surfaces. Modifications were made to both the rudder and elevator control surfaces in order to be able to use tape as the hinging mechanism.


The empennage was designed so as that the servos were to be placed in between closely spaced ribs, allowing easy access if necessary as well as a solid mounting point. The servos were installed and set to neutral before the arms and pushrods were installed.

Horizontal and Vertical Fins

The horizontal and vertical fins were added as part of stylizing the aircraft and making the overall design of the plane more aesthetically pleasing. These fins also complimented the winglets that were added to the ends of the wings. The horizontal fins were attached to the horizontal stabilizer after its fabrication was completed. The vertical fin was installed onto the vertical stabilizer after the entire empennage was fixed to the fuselage. This was done as that the mounting system for the empennage to the fuselage prevented the fin installation until after completion.

Covering Film

Covering the entire of the empennage was relatively simple process. The thermal film was ironed onto the balsa sheeting and stretched over the gaps. Most of the empennage was filmed once it was completed, but some elements had to be filmed after it was installed onto the fuselage. Elements such as the join between the empennage and the fuselage, as well as the vertical fin had to be filmed at a later stage.

Figure 40 - Complete empennage


Electronics Most conventional remote control aircraft utilize a variety of electronic components for power, control and sustained flight. These components are all critical to flight, and if any one fails the aircraft could fail to fly. Below is a brief overview of the main radio control components used in the design, build and flight of UAV aircraft.

Electric Motor

The majority of modern day RC aircraft use outrunning electric motors due to their lower weight and higher power output than their brushed counterparts. Brushless motors differ to conventional brushed motors by lacking the iron brushes and commutator inside the motor in order to reverse the direction of the induced magnetic force every rotation. A typical brushless motor with utilize 3 motors leading to the ESC, not just the two usually used in brushed motors. The third wire acts as a sensor that electronically reverses the direction of the current during operation as opposed to mechanically, thus drastically reducing the internal friction and hence energy losses across the motor. Hobby grade electric RC motors are usually rated in terms of maximum voltage, current draw and its KV rating. The KV rating is an indication of how fast a motor will rotate when a voltage has been placed across it. The KV is literally the number of revolutions per minute (rpm) the motor will make when 1 volt of electricity is placed across it. For example, a 1000KV motor with a 4S (14.8V) lithium cell would rotate at 14,800 rpm. Of the brushless-type motors there are two main variants – inrunner and outrunner. The main difference between the two is the part of the motor that rotates – inrunners behave exactly like conventional brushed motors with the shaft rotating and the copper coils inside the body of the motor remaining stationary. Outrunners on the other hand have the entire outer shell of the motor and the driving shaft rotating, the copper coils around the core of the motor remain completely stationary. Both outrunner and inrunner types are used widely across the RC community, and both have their advantages and disadvantages. Inrunner motors are more conventional and provide more power and efficiency at higher rpms. Another characteristic of inrunner motors is the fact that they are usually much smaller than an equivalent outrunner and hence are more often found in the likes of RC cars. Outrunner motors are comparatively larger, but have the additional bonus of active air cooling through the motor. Due to these two factors, they are the most common motor utilized in hobby-scale electric remote control aircraft.

Electronic Speed Controller (ESC)

The ESC is what provides power to both the motor and the receiver within an RC circuit. The ESC links the flight battery to the motor and the receiver. There are two main types of ESCs - brushed and brushless to match the type of electric motor being used. Most modern ESCs should have a type of BEC inbuilt to provide power to the radio receiver. ESCs are rated on the current draw they are capable of holding, known as a constant current. This current is the maximum current that the motor can constantly draw through the ESC. Most ESCs also have what is called a burst current. The burst is usually significantly larger than the constant current


and can only be held by the ESC for a limited amount of time, usually on the order of 10 to 20 seconds before irreversible damage is done to the ESC.

Propeller

The propeller of an aircraft is the most conventional way to provide a driving force pulling the aircraft through the air. Propellers work by deflecting air backwards through the air resulting in a driving force forwards. In the case of remote control aircraft, there are numerous types of propeller arrangements: tractor, pusher, and folding, counter-rotating and contra-rotating variants. The most common by far is the tractor configuration. Remote Control scale propellers are usually denoted by a set of numbers: 12x6 or 14x7 for example. The first number denotes the diameter of the propeller in inches, and the second denotes its travel. Under ideal circumstances, the travel indicates the amount of distance the propeller would be drawn forwards after undergoing one complete revolution. For example, a 12x6 prop would ideally travel 6 inches forward after one complete rotation.

Electric Ducted Fan (EDF)

Electric Ducted Fans are another method of providing thrust to an aircraft. They are more representative of modern turbines than of conventional propellers. Numerous small fan blades are joined to a central shaft, and are spun at very high speeds. EDFs are designed for high end speed and hence are usually found in replica jet aircraft. The nature of EDF units makes them not as efficient as propellers at lower speeds, hence providing lower acceleration and longer take off distances. They also require more power than any propeller equivalent. Another drawback of an EDF system is that the entire unit must be perfectly balanced in order to operate both efficiently and safely. Any imbalance at very high rpm could cause the shaft and fans to destabilize and possibly strike the inside shrouding of the fan unit.

Battery Eliminator Circuit (BEC)

BECs are required to provide power to the servos on a model aircraft allowing directional control during flight. BECs are commonly integrated into ESCs, but on occasion with larger model aircraft, a completely separate BEC is required to provide power to the servos. The BEC is designed to divert some of the power from the flight battery into the receiver so that it can power each of the servos linked to each of the control surfaces. On the occasion that a large number of servos are required for an aircraft, an additional external BEC is required to pick up the deficit power required by the receiver to power them.

Servos

Servos are small electronic components containing an electric motor and gearbox system. They are designed to rotate a lever arm one way or the other according to the input signal sent by the receiver unit. Most servos utilize 3 wires for normal operation: positive 5 volts, negative and a signal wire. Servos draw power from the receiver; hence a typical standalone receiver can only support a minimal number of servos simultaneously. BECs however can be added into an RC circuit in order to provide additional power to the receiver so as it can power additional servos. Wires such as


extension leads and y harnesses can also be used to extend the distance of a servo from the receiver unit and allow two servos to operate simultaneously with each other.

Lithium Polymer Battery

Lithium Polymer is widely used within the RC community as a more powerful alternative to heavier Nickel Cadmium (Ni-Cd) batteries. Lithium Polymer batteries are rated by the number of cells they have, the milliamp hour rating (MAh) and the discharge rate (C). The number of cells in a lithium battery is directly proportional to the voltage that the battery can provide. Each cell can provide a maximum of 4 volts, thus a 4 cell battery would provide a maximum of 16 volts. The milliamp hour rating is simply expressing the capacity of the battery – a larger milliamp hour rating yields longer flight times, likewise for smaller ratings. The discharge rate (the C value) gives an indication of how much current can be drawn through the battery before permanently damaging it. For example, a 1000mah battery with a 25C discharge rate gives a maximum discharge current of 25 Amps. Similarly, all lithium batteries have a maximum charge rating of 1C, meaning that the same 1000mah battery can be safely charged at a rate up to 1 Amp before damage to the battery occurs. Lithium Polymer batteries can become very volatile under the right circumstances. Hence, the operating voltage of each cell must remain between 3 and 4 volts – lithium cells cannot be discharged to 0 volts like Ni-Cd cells otherwise it will cease to function. Comparatively, over charging the battery or exceeding the maximum charge or discharge current can cause individual cells to ignite and become very unstable. Caution is recommended at all times when using these batteries.

Transmitter (TX)

The radio transmitter is the control unit from which signals are sent to the receiver in order to control the aircraft. Previous model transmitters transmit their signals over the kilohertz frequency range; however with the increased risk of interference over this wavelength current model transmitters are designed to operate over the frequency of 2.4 GHz. Radio transmitters transmit several channels of PWM (pulse width modulation) or PPM (pulse pitch modulation) in order to control the servos attached to different channel outputs of the receiver unit.

Receiver (RX)

The radio receiver is designed to wirelessly receive the signals sent from the transmitter and convert them into the motion of servos or the motor. The signals sent to the servos are either in PWM or PPM depending on the branding of transmitter and receiver set. The receiver also provides power to the servos; however a BEC is required to provide additional power to the receiver in the event there are too many servos connected to the receiver for it to handle alone.

Electronic Schematic

Shown below is the set up of the electronic system.


BEC

Elevator Servo

Receiver Transmitter

ESC Lithium

Polymer

Battery

Aileron Servo

Motor

Elevator Servo

Rudder Servo

Nose Gear Servo

Aileron Servo


Ground Testing A series of ground tests were performed. The objective of these tests was to determine the

functionality of the electronics and the integration of these components in to the structural design

of the UAV.

Ground Test 1 The first ground test was performed indoors, within the hallways of the Aerospace building at NUAA.

This test was performed when the fuselage had finished initial construction, prior to the application

of the wing and the empennage. The landing gear had been attached and the relevant electronics

were placed in position, including the servo for the nose gear and the battery and esc for the motor

and prop system.

Figure 41 - First ground test

The UAV taxied along the length of the hallway a number of times, with the velocity increasing as it

proved safe. The response of the UAV to control given by the pilot was satisfactory, within the limit

of the width of the hallway.

This first test was seen as successful, as the UAV taxied at a relatively high velocity and performed an

adequate turning circle. The control given by the nose gear was seen as adequate by the group.

Ground Test 2 The second ground test was performed outdoors, on the grounds of NUAA. The surface was a

tarmac similar to what was expected to be used during the flight test. During this test the wing was

attached using the bolt system applied. The empennage had not yet been connected as it was not to

be detachable and alterations still needed to be made.


Figure 42 - Second ground test

The UAV taxied on the surface of the tarmac fairly well. It reached speeds close to 10m/s. The

control of the UAV on this surface was obviously less stable, the height of the roughness with

respect to the size of the wheels led to some slipping while taxiing. Although it was not expected

that sharp turns were to be made at higher speeds, the loss of control was identified as an issue.

This second test was also seen as a success, with the only issue the slipping during turning. The cause

of this was seen as the roughness of the surface, and also the lack of strength in the nose gear. This

led to changes being made to the material used in the nose gear, through the addition of steel rods

to the set up.

Ground Test 3 The third ground test was again performed outside the Aerospace building at NUAA. This test was

done after the application of the film to the UAV, and the empennage and wing were both attached.

At this stage the group intended to push the UAV to velocity close to lift off. This velocity was

determined visually.

Figure 43 - Third ground test


The UAV was again taxied around the tarmac at NUAA, however the addition of all major

components, as well as the film left the UAV with a lot more weight. After the addition to the nose

gear, the UAV responded much better to the input controls from the pilot. The UAV taxied for

approximately 5 minutes before the group decided to accelerate it until it appeared to lift off. The

UAV begun to lift, and the throttle was released.

This ground test was seen as a complete success, with the overall aim of near lift off velocity

attained. The UAV handled well under the conditions, and there was lift generated at approximately

50% of input thrust from the pilot.

Figure 44 - Ground test close to takeoff


Flight Test On the 26th of December, the final flight test was performed. The lift off was taken at a tar road in

Nanjing. Prior to the flight test, the teaching assistants assigned to the group advised that the CG

was further back than was needed for stable flight. Weight was added to the UAV before takeoff to

shift the CG forward. This added weight was not accounted for in the design of the aircraft.

After pre-flight checks and a ground taxi test, the pilot deemed that the aircraft was ready to take off. The aircraft accelerated down the runway, and only slowly rose into the air after rotation at a long ground roll distance. The aircraft slowly lifted into the sky and briefly circled around. During this banking manoeuvre, the ailerons were seen to be fluttering very severely. After a second circle, the pilot tried to bank in the opposite direction, into the wind, to turn back the other way. The aircraft shuddered before tipping over and stalling. This subsequent stall caused the aircraft to tip over and plummet into the ground.

Damage The damage to the nose and wing of the aircraft was quite severe; however the remainder of the fuselage did not suffer as much damage due to the nature of the impact. The nose dive of the aircraft caused the entire nose section to sustain heavy damage, shattering many of the nose frame sections and extensions to the main spar. Upon impact, the firewall sheared off the spars, the propeller had shattered into pieces and the nose landing gear had also broken clean off. Part of the main wing also struck the ground on impact, causing the carbon to snap and breaking off the fuselage. Surprisingly the plate designed to join the fuselage and wing together did not shatter on impact, but the remainder of the wing did.

Causes From analysing the crash, there are two main causes that can be attributed to causing the aircraft to fail. The first cause being that the aircraft was severely overloaded even before it had taken off. The adequate centre of gravity as set by the designers was deemed too far after and hence additional nose weight had to be added to the aircraft to compensate. The nose weight only slightly increased the aircraft centre of gravity further forward, but more importantly it drastically increased the overall weight of the aircraft. This had the effect of making it much more difficult to take off, increase the take-off distance, and as a result the power system integrated into the aircraft had become obsolete and had to be operated at 100% capacity in order to remain airborne even for the small flight duration. The second factor is the phenomena of aileron flutter. Before take-off it was noted that the aileron control surfaces were slightly loose, but was dismissed as that the servos still had the capacity to fully deflect the surfaces both up and down. However, during flight severe aileron flutter was noticed with any banking manoeuvre. This instability prevented the additional control required to recover from the stall.


Appendices

A.1 Calculation spreadsheet

weight 4 kg

wing span 1.9 m

wing loading kg/m^2 5 6 7 8 9 10

wing area m^2 0.8 0.666667 0.571429 0.5 0.444444 0.4

wing chord m 0.421053 0.350877 0.300752 0.263158 0.233918 0.210526

AR 4.5125 5.415 6.3175 7.22 8.1225 9.025

Vertical Tail

VTVC 0.07 0.07 0.07 0.07 0.07 0.07

S-Wing Area 0.8 0.666667 0.571429 0.5 0.444444 0.4

b-Sing Span 1.9 1.9 1.9 1.9 1.9 1.9

l-cg to tail cg 1.15 1.15 1.15 1.15 1.15 1.15

Sv-vertical area 0.052961 0.063553 0.074145 0.084737 0.095329 0.105921

AR 1.7 1.7 1.7 1.7 1.7 1.7

b 0.300055 0.328694 0.35503 0.379543 0.402566 0.424342

c 0.176503 0.193349 0.208841 0.22326 0.236803 0.249613

taper 0.5 0.5 0.5 0.5 0.5 0.5

root c 0.235337 0.257799 0.278455 0.297681 0.315738 0.332817

tip c 0.117669 0.128899 0.139227 0.14884 0.157869 0.166408

Horizontal Tail

HTVC 1 1 1 1 1 1

S-Wing area 0.8 0.666667 0.571429 0.5 0.444444 0.4

c-Wing chord 0.421053 0.350877 0.300752 0.263158 0.233918 0.210526

l-cg to tail cg 1.15 1.15 1.15 1.15 1.15 1.15

Sh- Horizontal Area 0.292906 0.203407 0.149442 0.114416 0.090403 0.073227

AR 4 4 4 4 4 4

b 1.082416 0.902013 0.773154 0.67651 0.601342 0.541208

c 0.270604 0.225503 0.193289 0.169128 0.150336 0.135302

taper 0.5 0.5 0.5 0.5 0.5 0.5

root c 0.360805 0.300671 0.257718 0.225503 0.200447 0.180403

tip c 0.180403 0.150336 0.128859 0.112752 0.100224 0.090201


A.2 VP-69 AVL File

VP69

#MACH

0.0

#IYsym IZsym Zsym

0 0 0.0

#Sref Cref Bref

0.444 .26 1.9

#Xref Yref Zref

0.3 0 0

#CDo

0.023

#====================================

SURFACE

MAIN WING

# Nchord Cspace Nspan Sspace

10 1.0 50 -1.0

YDUPLICATE

0.00000

ANGLE

0.00000

SCALE

1.0 1.0 1.0

TRANSLATE

0.3 0.0 0.055

SECTION

#Xle Yle Zle chord angle

0.0 0.0 0.0 0.26 0.0

AFIL


527.DAT

SECTION

0.0 0.53 0.0 0.26 0.0

AFIL

527.DAT

SECTION

0.06 0.95 0.0 .20 0.0

AFIL

527.DAT

SECTION

0.09 0.95 0.05 0.22 0.0

AFIL

0012.DAT

#===========================================

SURFACE

HORIZONTAL TAIL

#Nchord Cspace Nspan Sspace

10 1.0 20 -1.0

YDUPLICATE

0.00

ANGLE

0

TRANSLATE

1.274 0.0 0.175

SECTION

0 0 0 0.226 0

NACA

0015

SECTION

0.113 0.34 0.0 0.113 0.0


NACA

0015

#=========================================

SURFACE

VERTICAL TAIL

# NCHORD CSPACE NSPAN SSPACE

10 1.0 20 -1.0

SCALE

1.0 1.0 1.0

ANGLE

0

TRANSLATE

1.2 0.00 0.075

SECTION

0.00 0.00 0.00 0.3 0.00

NACA

0015

SECTION

0.135 0 0.38 0.15 0.00

NACA

0015

#========================================

BODY

VP Fuselage

20 1.0

SCALE

1.0 1.0 1.0

TRANSLATE

0.0 0.0 0.0

BFIL


vpbod.DAT

#========================================

vp-69: “the black wing” - nuaa.edu.cnaircraftdesign.nuaa.edu.cn/pd-2007/report/2011-the black...

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