[Type text] Page 1
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
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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.
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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.
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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
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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
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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
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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
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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.
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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:
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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
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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.
The A-Team VP-69: “The Black Wing” Page 16
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.
The A-Team VP-69: “The Black Wing” Page 17
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
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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
The A-Team VP-69: “The Black Wing” Page 19
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.
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Figure 13 - Lift with respect to Drag
Figure 14 - Lift and Drag with respect to alpha
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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.
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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.
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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.
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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
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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
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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
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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.
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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.
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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
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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
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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
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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
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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
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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
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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.
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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
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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.
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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.
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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
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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.
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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.
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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
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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
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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
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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.
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BEC
Elevator Servo
Receiver Transmitter
ESC Lithium
Polymer
Battery
Aileron Servo
Motor
Elevator Servo
Rudder Servo
Nose Gear Servo
Aileron Servo
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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.
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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
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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
The A-Team VP-69: “The Black Wing” Page 50
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.
The A-Team VP-69: “The Black Wing” Page 51
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
The A-Team VP-69: “The Black Wing” Page 52
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
The A-Team VP-69: “The Black Wing” Page 53
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
The A-Team VP-69: “The Black Wing” Page 54
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
The A-Team VP-69: “The Black Wing” Page 55
vpbod.DAT
#========================================