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Coursework Title: Final Report Coursework number (i.e. CW1 CW2) Module code: EGA 302 Module title: Aerospace engineering design Submission deadline: 15/05/15 Supervisor: Dr Ben Evans

Student number: 710820 Group Name Absolute Zero

Email: 710820@swansea.ac.uk

Degree course: Meng Aerospace engineering

Contents

1 Introduction 4

2 Concept Design Process 42.1 Team Role Agreement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.2 Mission requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 42.3 Competitor Survey . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4

3 Regulations 8

4 Preliminary Design 84.1 Aerodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

4.1.1 Aerofoil selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84.1.2 Aerodynamic characteristics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94.1.3 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

4.2 Structural Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104.2.1 Structural Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

4.3 Powertrain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.3.1 Motor selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124.3.2 Propeller selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

4.4 Material selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134.4.1 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13

4.5 Control Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134.5.1 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.6 Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 144.6.1 Aircraft Changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 14

4.7 Weight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154.8 Aircraft Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 15

4.8.1 Flight envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154.8.2 Rate of Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 154.8.3 Sink Rate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.8.4 Range . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.8.5 Take-off and Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 164.8.6 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

4.9 Preliminary design conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16

5 Detailed Design Process 175.1 Aerodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17

5.1.1 Virtual Wind Tunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175.1.2 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 185.1.3 Assessment of Analysis tools . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

5.2 Structural . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195.3 Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 225.4 Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

5.4.1 Aircraft changes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 225.5 Control Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 235.6 Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 245.7 Weight and Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

5.7.1 Gliding flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 255.7.2 Turning flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 255.7.3 Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 255.7.4 Range and Endurance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 255.7.5 Final design flight parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25

2

6 Aircraft testing 256.1 Structural . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 256.2 Aircraft scaling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 266.3 Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 266.4 Flight Simulator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

7 Aircraft costing 28

8 Conclusion 28

9 Risk Assessment 28

10 Appendix 1 (VTS) 32

3

Design Report

Absolute Zero

University of Swansea

Abstract

A detailed design report has been carried out for the purpose of creating an Unmanned Aerial Vehicle (UAV)for scientific research in the mapping of glacial retreats in Greenland. A comprehensive look into how and whythe aircraft looks and performs the way it does along with costings to determine the viability of creating a UAV fitfor this purpose.

1. Introduction

Unmanned aerial vehicles (UAV) are becoming moreprominent due to their versatility and abilities to com-plete tasks that man can not always accomplish. UAVsare no longer only used for military applications butare starting to become viable options for scientists andhobbyists alike. UAVs or drones are divided into twomain categories, remote controlled or autonomous. Forthe purpose of this design report an autonomous dronewas chosen due to the large range required.

Glacial retreats are slowly being recognised by thegeneral public as a problem due to what may happenif glaciers start to disappear from around the world.‘Glacier mass balance’ is the key to understanding glacialretreats, this balance is the yearly addition of frozenwater to the yearly melted water determining whetherthe glacier is healthy or in retreat.[1] If glaciers fromaround the world were to disappear then it would leaveregions without fresh drinking water effecting animals,wildlife and over a longer period of time sea levels.[2]

The aim of this project is to help develop a scientificresource that allows glaciologists to map and keep recordof glacial retreats more easily and relatively quickly al-lowing for preservation of these prehistoric glaciers.

2. Concept Design Process

2.1. Team Role Agreement

Based on individual’s strengths an agreement towhich roles each members would specialise in along witha secondary role for support, Table 2 shows each memberand there specified subject area.

2.2. Mission requirements

As part of this design a guideline has been providedwith the minimum values required from the aircraft.Certain values have been upgraded as it was felt thebenefits of producing an aircraft with certain capabilitieswould be favourable to the mission.

Table 1: Mission requirements

Minimum desired

Take off distance 20m 20mRange 60km 90kmGust conditions 20− 30km/h 30− 40km/hCruise speed 50km/h 50km/hService ceiling 1000m 1000mPayload 500g 1kgcost £1000 under £1000Reusability Yes Yes

Due to Greenland’s unforgiving weather and howquickly weather fronts can form, were some of the keyfactors in changing values in Table 1. Gusty conditionsare a major concern due to the mountainous regionstherefore the decision to design a more stable aircraftwas chosen. Range was increased incase areas of in-terest arise during a mission along with an increase inpayload to lift better camera equipment or measuringequipment. A challenge was set within the design tocreate an aircraft under budget however this was not avital consideration in component costing.

2.3. Competitor Survey

To fully understand market needs and define a nichein the market a competitor survey was conducted.

Table 2: Team Role Agreement

J.Jacob J.Johnson I.Milodowski D.Parish M.Rowland-jones M.Satha W.Shackley J.Tang

Specialisation Structure Materials & Propulsion Dynamics & Stability Aerodynamics Structural Weight & Performance Aerodynamics Control SystemsSecondary Role Dynamics & Stability Weight & Performance Structure Control Systems Aerodynamics Materials & Propulsion Structure Aerodynamics

4

Hirrus UAV [15]

Specifications Hirrus UAV Weight 7kg Max speed 130 km/h Flight time 180 mins Range 30km (auto pilot) Payload 0.7kg Service Ceiling 3 km

Aeromapper 300 [16]

Specifications Aeromapper 300 Wing span 3m Fuselage length 1.23m Material carbon fibre fuselage and fiberglass payload bay Take off hand launch or launcher Empty weight 3.6kg Takeoff weight 5.2kg Cruise speed 58km/h Max speed 120km/h Endurance 90mins Cost £10, 200 all included

Mugin 2600 UAV [17]

Specifications Mugin 2600 UAV Wingspan 2.6m Weight (No Engine) 6.5kg Max Take Off Weight 15kg Payload 4kg Cruise Speed 120 km/h Flight Time ~2 Hours Cost £700 airframe

DuraFly Zephyr V-70 [19]

Specifications Zephyr V-70 Wing span 1.53m Fuselage length 1m Material Expanded PolyOlefin Take off Hand launch or

launcher Motor EDF 500 watts Takeoff weight 1.15kg

Specifications Skywalker X8 Wing span 2.12m Material Expanded PolyOlefin

Take off Hand launch or launcher Motor 400-800 watts Takeoff weight 3.5kg Cost £110 empty shell

Specifications UAV 3000 Wing span 3m Fuselage length 1.5m Material Glassfiber/ply fuselage Take off Hand launch or launcher Empty weight 5.2 kg Takeoff weight Dependant on motor up to 2

kg Cost £ 180 empty shell

 

UAV  3000  [20]  

Skywalker X8 [18]

Taking initial concepts and

evaluating by performance

characteristics, the sailplane

design was evidently the

most suitable

The Skywalker X8 is the airframe in use

by Aberystwyth University currently. This

airframe meets all design parameters

except payload capacity.

Step 1 - Generate and analyse initial concepts

Step 2 - Analyse available airframes

The BlitzRCWorks Sky Surfer is a

commercially available airframe. This

airframe is large enough to meet the

payload capability but compromises in

order to produce a scale-model

appearence reduce usable interior space.

The Durafly Zephyr is an alternate

design that uses an EDF jet to

climb to high altitude then

functions as a glider in flight. This

airframe has exceptional range

but low payload capacity.

Step 3 - Using knowledge gained through

research, produce a concept design to

take forward

The concept design is a high aspect ratio aircraft with a

semi-blended wing and twin propellers. It positions the

fuselage forward with ample space for flight systems

and payload, this design will be developed in the

detailed design phase.

Figure 1: Positioning map detailing niches in the marketwhere this design hopes to sit

After initial research into each area multiple droneswere picked for further consideration. A quick tabledetailing each aircrafts properties along with a pictureof each aircraft can be found above. From here a po-sitioning map was created using a marketing tool thatallows users to find niches in the market. Fig1 showswhere each UAV fits in the market and where the aircraftdetailed in this reports aims to fit in.

3. Regulations

It is important to understand the regulations thatmay effect the design of the UAV. Although the aircraftwill be flown in Greenland where the UAV regulationsare much more relaxed, it was important the UAV is ableto conduct missions within the UK for testing purposes.

Under the Civil Aviation Authority two key concernsdictate the need for a permission of flight:

• Is the aircraft flying on a commercial basis (i.econducting ‘aerial work’)

• Camera or surveillance equipment fitted to theaircraft within congested areas.

Although the aircraft will not be working on the basisof monetary gain it will still be conducting work for anorganisation therefore a certificate will be required to flythe aircraft. The second key parameter will however notapply to this aircraft due to the surveillance or cameraequipment not operating in congested areas.[3]

UAVs are classified into 3 types based on overallweight. Class 1 under 20 kg, class 2 between 20-150kgand class 3 anything above 150kg. A certificate of Air-worthiness is required for any UAV over the weight of150kg, for the purpose of this project it will not berequired due to a very low weight under 10 kg.[4]

4. Preliminary Design

From detailed calculations and research, each area ofthe conceptual design was looked into and accessed forviability and purpose ultimately determining the finalaircraft at the preliminary stage.

4.1. Aerodynamics

4.1.1 Aerofoil selection

Reynolds numbers are a vital step in choosing anaerofoil. A low reynolds number (Re) is favourabledue to low Re values experience more laminar flow andtherefore the aircraft will produce more efficient wingsgenerating lift.

Re =ρV c

µ(1)

From Equation (1) a value of Re = 250000 was foundallowing for the comparison of multiple aerofoils.

Figure 2: Comparison of multiple Aerofoils over four keyareas at the given reynolds number

The generation of the graphs in Fig.2 were producedby Xfoil,[5] a program and analysis tool available foraerofoil selection. The program was developed at MITand is only applicable if certain criteria are not meti.e Compressible flow, Viscous Flow, etc... Based onthe low reynolds number and the values calculated byXfoils, the NACA 2412 was chosen due to its low dragproperties and relatively high Clmax. Initial studies intothe NACA 2412 using both Tornedo[6] and XFLR[7]which will be discussed later on, yielded low values oflift. Two possible options are to increase surface areas,mainly wing span, or change the aerofoil to a highercamber therefore generating more lift. The decision tochange to a NACA 6412 was chosen due to changing thewing span by the amount needed would have resulted ina difficult aircraft to launch.

8

4.1.2 Aerodynamic characteristics

To determine the aerodynamic characteristics of thewing geometry and performance, multiple methods wereused. The advantage of using multiple methods meantthat these calculations could be made more accurate.The following is a brief explanation of how each theoret-ical method works.

Prandtl Lifting Line theory assumes that there isonly one horseshoe vortex for each wing segment, thusmaking the wing finite. The theory predicts the distribu-tion of lift generated along the span of the wing throughits three dimensional geometry. The strength of thisvortex reduces along the span. To ease calculations thetheory does not take into account the following; Com-pressible flow, Viscous Flow, Swept Wings, low aspectratio wings and unsteady flows. The use of this theoryrevolved around using different aerofoils from an aerofoilgenerating software online known as airfoiltools.[8] Fromthis website estimations of Coefficient of lift propertieswere taken from graphs which showed the characteristicsof each aerofoil. This was then introduced into the ap-propriate equations to calculate the Coefficients of dragand lift.

CL =2L

ρv2S(2)

CD = Cd0 +C2l

πeAR(3)

Results from equations (2) and (3) can also be used todetermine the optimum angle of attack to fly at a givenwing geometries through the greatest CL/CD ratio.

Vortex Lattice Methods models the lifting surfacesof a wing by assuming that the wing is an infinitely thinsheet made of small vortices, this is influenced mainly bythe thickness of the sheet. It is an extension of Prandtlslifting line theory however instead of the theory assumingthat there is a single vortex per wing segment a latticeof these vortices are generated. To simplify calculationsthe software makes the following assumptions; the flowis incompressible, inviscid and non-rotational.

The lifting surfaces are assumed to be thin and theinfluence of the thickness is neglected. It is also assumedthat the angle of attack and the angle of sideslip are bothnegligible. Two different types of software were used todo these calculations both with different advantages overthe other. Tornado which allows ease of design of thewings through coordinate systems, which creates threedimensional lifting surfaces such as the wings, horizontaland vertical tails. However this software lacked the abil-ity to create a fuselage and simulate how the fuselagewould interact with the lift generating surfaces. The sec-ond was a program called XFLR5 which uses the same

interaction system as Tornado however has the ability tocreate a fuselage and shows the effect this will have onthe wings and also has some very basic computationalfluid dynamics entwined into the software to producegraphical representations of flows.

Table 3: Aerodynamics characteristics based on multiplemethods

NACA 2412 NACA 6412

Wing Span 3 2.4Mean Geometric Chord 0.3 0.25Reynolds number 290000 220000Wing inclination 4 3.7

TornadoCL 0.54 0.75CD 0.009 0.024

XFLRCL 0.44 0.8CD 0.01 0.024

Finite wing methodCL 0.55 0.692CD 0.02 0.05

Table 3 shows the two key iterations and the differ-ence between both theories.

4.1.3 Aircraft changes

First iteration: Originally a weight estimation of 6kgmeant that the coefficients of lift and drag were calcu-lated to produce enough lift for the aircraft to fly atstraight and level un-accelerated flight where lift is equalto weight. Using a constant chord is often used wherelow cost is important because of their ease to build andmanufacture, but they are less efficient in the outer sec-tions of the wing. Through structural analysis the wingspan overall was not needed and could be shortened toreduce loading factors caused by the span of the wingand amount of material used.

Second iteration: With the estimated mass of theaircraft still at 6kg, it was determined that changingthe aerofoil to a NACA 6412 meant that more lift couldbe generated because of the increase in camber of theaerofoil. This also allowed for partitions of the wing tobe tapered so to increase the aspect ratio of the wingsmaking them more structurally and aerodynamically ef-ficient by reducing wing tip vortex strength. It was alsoslightly swept back so as to increase the aerodynamicstability. The change in aerofoil also meant that therewas now a larger CLmax available, this means that thedistance required to take off would be shortened, alongwith better stall characteristics.

9

4.2. Structural Design

The design of certain features were governed by theresearch from the aerodynamics, such as the wing shapeand aerofoil, and this evolved as time went on. Themain difference from the original wing concept was theintroduction of a taper ratio of 0.5 and increasing thesweep to 10 degrees.

One of the conceptual designs for the internal struc-ture of the wing was to produce a hollow shell with aspar. This was produced to try and reduce weight whileremaining as strong as possible, however when tryingto optimise weight and structural integrity it was foundreducing the wall thickness of the aerofoil to reach anoptimal weight was problematic for structural loads.

It follows that a more classic ribs and spar configu-ration has been adopted and the material selection hasincreased stiffness and reduced weight. The tail geom-etry again changed with the larger single-vertical tail.Identical to the front wing, these were originally designedto be a plastic shell but are now ribs and spars fromthe original concept, the fuselage design has changedsubstantially: removing the joined front and tail wingconstruction as it increased weight too much and chang-ing the battery choice altered the front hub. With theshortened hub the design became structurally weakerdue to the point at which the tail of the wing cut intothe fuselage was near where the hub ended creating athin section of material, so a new design was looked intowith an additional aim of being more aerodynamicallysuited to the flight parameters.

A high lift generating (HLG) fuselage was then de-signed by using Mathematical optimisation techniquewhich then favours with more lift and low drag char-acteristics at lower angles of attack, short landing andtake-off capabilities.[9] The main challenge in the fuse-lage design was the space requirements and to get atechnical structure which withstands the load factors.The largest stresses act at the joints where the wings areconnected to the fuselage, thus this area was strength-ened with a larger wall thickness. Fillets were appliedat the sharper edges which again gives a uniform flowof the loadings. Hence, the internal structure resists thetensile and compressive loadings.

Fig 5 shows an overview of the aircraft design atthis stage along with placement of components. Theplacement of each component was derived with the helpof control and stability to ensure a stable aircraft duringflight, see the section on stability.

4.2.1 Structural Diagram

A vital part of producing an aircraft is evaluating thelimitations. A load factor, n, can be calculated based

on equations (4) and (5). This graph shows three keyareas, the first is the stall properties of the aircraft thisis important when flying slowly. The top horizontal lineshows the structural limitations in-terms of manoeuvres,the last line is the the vertical line where a maximumspeed is applied before structural loads become too high.

n =q

W

S

√CD0

k(4)

n =

√√√√√qπAe

W

S

[(T

W)max −

qCD0

W

S

] (5)

This can be depicted as a ‘V-n’ diagram found inFig. 3.

Figure 3: Structural limitations of the aircraft

10

Figure 4: Aircraft evolution

11

Figure 5: Aircraft cross section along with component placing based on CG calculations

4.3. Powertrain

4.3.1 Motor selection

There are many types of engines ranging from elec-tric propeller driven , diesel propeller driven, electricducted fan (EDF) and jet engines. Initially both jetengines and electric ducted fan were ruled out due tothere lack of efficiency. EDF systems are designed tooperate at large RPM and produce large amounts ofthrust however, they are predominantly used in the RCworld as motors installed in model jet fighters reachinglarge speeds and relatively low flight times, this wasdecided, for the purpose of this mission inadequate. Jetengines become very inefficient when scaled down tothe size we need, they also generally have high specificfuel consumption compared to a small diesel or petrolpropeller driven aircraft therefore this motor was alsoruled out.

When comparing an electric motor versus a diesel orpetrol engine the main differences are in efficiency andweight. With recent advances in brushless technologyelectric motors can reach anywhere from 75%-85% effi-ciency much higher than internal combustion engines.[10]The weight of a petrol or diesel engine is much higherincreasing the aircrafts overall weight, not a desirablefeature. Due to the small thrust requirements producedby weight and performance along with aerodynamicsresults the choice for a mid powered electric engine waschosen producing roughly 200 watts of power.

4.3.2 Propeller selection

There are two main aspects to all propellers, the di-ameter of the propeller and the pitch of the blades. Thediameter of the propeller is the distance from tip to tip,

the pitch or twist of the blade is defined as the distancethe propeller would move the airplane forward in onerotation in a perfect world. However this is impracticalas perfect conditions will almost never arise due to thefact that propellers are never 100% efficient and this isalso considering an incompressible flow.[11] Although atthe speeds the aircraft is flying it would typically notencounter compressibility effects it may be encounteredat the tips of the propeller.

The effects of the diameter of the propeller in generalwill result in a larger amount of thrust produced by theengine, whereas the pitch will increase the speed of theaircraft. For example a small diameter coupled with alarge pitch will move faster through the air however onlymove small amounts of air meaning it will be perfectfor small aircraft looking to move fast. A large diame-ter propeller with a shallow pitch angle will move largeamounts of air meaning large amounts of thrust but theshallow pitch angle means it will move through the airmore slowly.[11]

Based on Fig.6 and the more desirable shallow pitchand larger diameter it can determined that a propellersize of 10× 5 is more desirable for the Greenland appli-cation. The choice of propeller size also means that anincrease in torque benefits Take-off and Landing proper-ties of the aircraft.

12

Figure 6: Propeller sizing guild based on engine size[12]

4.4. Material selection

The material selection was based on work completedby the structural group along with the help of the Edu-pack software. The structural design allowed for a maxi-mum of 1000kg/m3 for critical components and a min of100kg/m3. From these values a list of possible materialswere chosen based of manufacturing routes, structurallimitations and overall viability for the aircraft. Duringthe detailed design phase materials will be assessed andsimulated to verify functionality. A short list of possiblematerials for key component will be carried forward are:• Nylon 6 \10• High density polystyrene• hard wood, Spar• Balsa wood

4.4.1 Aircraft changes

The conceptual design brought forward featured twoengines mounted on the underside of the wings. Thetwo engine configuration has since been dropped to asingle engine due to two main reasons, the first of whichis weight. Due to the relative lightness of the proposedaircraft, having two very small engines produce the sameamount of thrust as having one slightly larger enginewith next to no real benefits with regards to excessweight. The second reason is due to the effect it willhave on the range of the aircraft. Most brushless engineswill have a 75-85% efficiency therefore losses associatedwith having two engines is much higher than just onesingle engine. Not only do the losses in efficiency reduce

the range of the aircraft but due to each engine drawingseparate currents, the amount of energy needed for bothengines will far exceed that needed for one single engine.

Towards the end of the preliminary phase a prob-lem was found in the design and material choice for thefuselage therefore a complete overhaul of structure andmaterial choice was done which will be discussed lateron.

4.5. Control Systems

Due to the difficult nature of the mission a detailedlook into the flight controls and telemetry for the aircrafthas been conducted. The Greenland project requires anaircraft that is autonomous and able to capture imagesof the landscape it is flying through.

Multiple autopilot systems have been studied andthe AMP 2.6 board with GPS is a viable option at thiscurrent stage. It includes 3-axis gyro, accelerometer,magnetometer, barometer and other high performancerecording instruments that can be streamed live to theground station while in range or recorded while out ofrange. The system also features an open source autopi-lot systems using Invensenses 6 DoF Accelerometer andGyro MPU-6000.

Camera equipment is one of the most important as-pects of design. The mission requires scientists to analysepictures captured from the aircraft to help map glacialretreats. The initial design was to have two cameras, thefirst facing forward and the main camera facing downmapping the landscape. However due to the aircraftchanging from 2 engines to 1 engine, there is no longerroom for two cameras therefore one main camera will bepointing down mapping the landscape. A GoPro hero 4will be used as a high resolution device is needed.

Research on battery quantity and quality has beencarried out and there are clear advantages using LiPobattery packs. Calculations have been conducted:

Batterylife =mAh

mA× 0.7 (6)

Based on Equation 6, where an efficiency of 70%was used for environmental factors, and values foundby the propulsions section it was determined that twohigh capacity LiPo batteries will be required for thegiven flight time however three will be used for extrarange allowing for a safety factor and redundancies. Thebatteries under consideration at this time is the ZippyTraxxas 7600mAh 2S 1P 30C.

Table 14 shows the current selection of equipmentproposed for this mission along with dimensions andweight of each component.

13

Table 4: Control Systems

component Length (mm) width (mm) Height (mm) Weight (g)

Ardupilot 2.6 70 40 10 32Turnigy MX-353S 17g Servo x 4 38 13 27 17× 4Zippy Traxxas 7600mAh Battery x3 157 25 45 367× 3Turnigy Dual Power Unit 100 50 20 89Turnigy Plush 60A Speed Controller 80 31 14 60Turnigy D3536/8 1000KV motor 52 35 35 1023DR uBlox GPS + Compass 38 38 8.5 16.83DR Video/OSD System Kit N/A N/A N/A 100− 150

Total 1594

4.5.1 Aircraft changes

Due to the amount of components to fit within thefuselage it was required at an early stage to help re-design the fuselage to accommodate all necessary flightequipment. Although this can not be seen externally,internally new compartments where created.

4.6. Stability

For an aircraft to be stable it must, after a periodof time, return to an equilibrium point in flight follow-ing disruptive forces, such as a gust of wind or controlsurface deflection. The first task with regards to the dy-namic and static stability of the UAV was to determinethe static margin. The static margin is defined as thedistance between the centre of gravity and the neutralpoint as a percentage of the mean chord. For an aircraftto be statically stable, the centre of gravity must beforward of the neutral point, therefore the static marginmust also be positive. Generally a margin greater than5% [1] should provide sufficient stability.

An increase in angle of attack, α, should generatea nose-down pitching moment, directing the aircraftback towards equilibrium. In reverse too, a decreasein α should generate a nose-up moment, directing theaircraft once more towards stability.

dCMdαα

< 0 (7)

At straight, level and steady flight the pitching momentshould be zero, as the aircraft should be in its equilib-rium position. Therefore:

CMα(αα = 0) < 0 (8)

To calculate the static margin, this equation was used:

CMα = (h− hn)CLα (9)

From this equation, it can be seen that the centreof gravity must be located in front of the neutral point

in order to achieve static stability. The static margin isnoted as: h− hn. To make the calculation of the staticmargin easier, a MATLAB script was written, enablingit to be calculated quickly during changes to the UAVparameters.

Taking into account the configuration of the UAVin the early design stages, the static margin was calcu-lated to be 0.2359, 23.59%, demonstrating static sta-bility. With the updated design, this was then revisedto 21.15%, matching also the values achieved from Tor-nado. However, this value was deemed to be too high toachieve sufficient manoeuvrability, and the target for astatic margin of around 15% was set, as UAV are usuallyexpected to have a static margin in the region of 5%to 15%. To achieve this, the positions of the masseswithin the airframe structure were shifted closer to theneutral point. A static margin of 14.69% ended up beingcalculated for the configuration detailed above. Table5shows the change in CoG based on changes made to thestatic margin.

Table 5: Centre of Gravity variation from leading ledge ofwings

Before After

X-CoG 30.7mm 108.46mmY-CoG 0 0

4.6.1 Aircraft Changes

The tail design for this aircraft will have a majoreffect on stability of the aircraft during flight. Choosingthe correct configuration will make for a more stableaircraft resulting in clearer pictures and therefore moreaccurate aerial photography for the Greenland project.The tail configuration carried forward from the con-ceptual design was a twin tail plane, this design wasconsidered favourable at the time due to its larger surfacearea and therefore increased stability during turbulent

14

winds. This has since been changed to a more classictail plane design for several reasons.

The first reason for changing this was due to theextra weight that would be added to this design dueto structural reasons. Placing such a heavy weight atthe end of the horizontal stabiliser would mean that theinternal structure would outweigh that of the classic taildesign.

The second reason was due to the redundancy levelof having two rudders. This would increase drag, andother aerodynamic problems associated with having thesecond surface.

Weighing the advantages and disadvantages of thetwo types of designs it was more favourable to revert thedesign back to a classic tail design making for a lighter,more efficient and simpler design.

A major change due to stability reasons was thelength of the fuselage and the positioning of the wings.The increased length of the fuselage and moving thewings positioning further back allows for a better staticmargin, moving the centre of gravity closer to the neutralpoint of the aircraft. This produced a more staticallystable aircraft.

4.7. Weight

Keeping track of weight through the design is a vitalpart of producing an aircraft. A comprehensive tableknow as a Vehicle Technical Specification, VTS, was pro-duce to keep track of all incoming data from each designgroup, this can be found in Appendix 1. A small tableshowing weight estimations at the start and throughoutthe initial build can be found in table 6.

Table 6: Weight estimation

Weight (g)

Battery 367× 3 = 1101Motor 160Propeller 35Airframe 3000Camera 1000Autopilot & GPS 28Power module 17Telemetry module 33Servos 17× 4 = 68Total 5500 (estimated 6000)

The performance parameters are mainly dependenton the estimated aircraft weight, propulsion system andthe aircrafts aerodynamic characteristics.

4.8. Aircraft Performance

4.8.1 Flight envelope

The flight enveloped is developed from stall proper-ties and available power.

PR = (1

2CD0ρS)v3∞ + (

2W 2

πeARρS)

1

v∞(10)

vstall =

√2W

ρSCLmax(11)

The rearrangement of equation (10) along with equa-tion (11) allowed for Fig.7.

Figure 7: Flight envelope

Fig.7 shows the aircraft power requirements as afunction of altitude and velocity. This graph enables theoperator to determine minimum flight speed at given alti-tude either based on stall properties from wing geometryor that maximum trust used.

4.8.2 Rate of Climb

The rate of climb is determined using drag charac-teristics and energy considerations. Fig.8 shows the rateof climb against altitude. Rate at which power is be-ing used by the aircraft is due to varying the energyof the system and the rate at which drag is affectingthe power. When density, available power and velocitychanges during the flight, the climb rate changes.

15

Figure 8: Rate of Climb

4.8.3 Sink Rate

Due to the nature of the aircraft it is desirable tocreate an aircraft with an effective sink rate for efficiencypurposes.

vsink = v∞sinγ =PRW

(12)

Based on equation 12, Fig. 9 was produced.

Figure 9: Aircraft Sink rate

4.8.4 Range

With the aid of control systems and propulsions, anunderstanding of the aircraft’s range capabilities was

calculated. Based on the energy consumptions usedby onboard electronics a range of 41km was calculated.This value is to short for the purpose of this missionand has been extended as mentioned during the detaileddesign stage.

4.8.5 Take-off and Landing

The take-off distance consists of several parts, thefirst of which is the ground run. The take-off distanceis calculated for maximum weight at standard density.The ground run is currently 8.84 meters at this earlystage.

4.8.6 Aircraft changes

After initial calculation it was found that the takeoff distance was too long therefore two solutions wherefound, increasing available motor power and increasingwing surface area to generate more lift therefore reduc-ing take off distance. The NACA 6412 was thereforeadopted along with a more powerful motor reducing theoverall take off distance.

4.9. Preliminary design conclusion

At the end of the Preliminary design many prob-lems were resolved however the aircraft produced hasmoved away from the conceptual design. This is in partdue to an ease of calculations at this initial stage andalso due to problems found in the conceptual design.Moving forward into the detail design phase testing willbe carried out on all aspect of the aircraft along withslight modifications for improvement to aerodynamiccalculations along with structural design.

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5. Detailed Design Process

Moving out of the Preliminary design phase wherethe aircraft’s overall characteristics have been decidedand early calculations have been conducted to ensure afunctional plane, the aircraft now moves into the detaileddesign phase where parameters are refined to ensure ef-ficiency and more desirable characteristics. During thisstage aircraft modelling followed by aircraft testing hasbeen accomplished to ensure an aircraft that not onlyflies but handles correctly.

5.1. Aerodynamics

The aircraft wings underwent a large change in shapefor both aerodynamic and structural redundancy pur-poses. This resulted in an overall aircraft weight reduc-tion and more sleek aerodynamic geometry allowing forreductions in drag, Fig.10 shows the evolution of theaircraft wings from an aerodynamic point of view.

Figure 10: The first iteration is a simple wing, rectangularin shape, this was then tapered at the ends in aneffort to reduce wing tip strength in the seconditeration. The last and final iteration has anincreased taper ratio along with a semi blendedwing, larger root cord, for structural reasons toincrease stiffness along with increase in lift andaerodynamic properties.

During the detailed design stage of the project morerealistic software was adopted to verify lift and dragproperties. This was done to verify values that havealready been calculated and to fully understand theaircraft.

5.1.1 Virtual Wind Tunnel

Virtual Wind Tunnel by Altair[13] is a Computa-tional fluid dynamics software package designed for au-tomotive vehicles. The key benefit of this software isthat it uses viscous flow, much more similar to the realworld. Gathering data from the virtual wind tunnelallows for more accurate calculations for instance lift,drag, pressure distribution and flow separation.

The Virtual Wind Tunnel works on the utilisationof the Navier-Stoke equations. It is possible to calcu-late the forces acting in the X, Y and Z co-ordinatesystem, for the purpose of this study, lift, drag and crossforces (usually negligible) respectively. The values foundare then normalised into values of CL and CD usingequations (2) and (3).

Generating a mesh is extremely important whentesting in the virtual wind tunnel, there are two basictwo-dimensional shapes that are used for meshing. Thebenefits of using a triangle is that is it the simplest typeof mesh to create and has the ability to give a moreaccurate concave or convex shapes. This is of greatimportance when generating a mesh for the UAV as thewings are in the shape of an aerofoil which is smooth,any difficult to produce a smooth profile of the wingwhich will greatly affect results.

Element size and the maximum deviation of elementsis important to create a fine mesh. The benefit of havinga finer mesh is that it will give more accurate results.For this test the number of elements created is around30,000. Which is more than enough for simple geometryinput, to achieve more accurate results a finer meshcould be generated, however the finer the mesh the morecomputationally heavy and time consuming it will be.The mesh created for the virtual wind tunnel can befound in Fig.11

Figure 11: Meshing of the aircraft for the Virtual WindTunnel

17

After multiple results where found using the virtualwind tunnel and testing commenced in the flight sim-ulator it was noticed that the aircraft was not actingas anticipated. After some initial research it was foundthat the software uses the Spalart-Allmaras model whichwas found to have some disadvantages when performingin the boundary layer resulting in a reduced drag values.The advantage of this model however is the computa-tional time is much smaller than different models.

Creating a large enough area around the body of theaircraft is important, this will result in more accuratevalues based on better flow separation over the wingand behind the UAV. The results produced will be moreaccurate as the flow will have more time to convergeback to a laminar flow, as turbulent flows will affect thechange in pressure over and under the wing causing lossin lift.

Figure 12: Computational area around the aircraft for sim-ulations

Table 7 Shows two iterations and there correspondingaerodynamic values.

5.1.2 Aircraft changes

Third iteration: As the first structural design iter-ation was completed there was a large change in theestimated mass, this has meant that the second iterationwas producing a lot more lift than desired, to combatthis the wing span was decreased. Once this model hadbeen constructed it was placed into the virtual windtunnel made by Altair. As the results show, the air-crafts co-efficient of lift and drag are both higher thanthe other calculations, this is because VWT takes intoaccount viscous flow, whereas the other theories are allbased on ideal flow and cannot predict flow separationvery accurately. From this it was determined by thegroup that the VWT values are more accurate and sothe redesign of the wing was undertaken.

Table 7: Aerodynamics characteristics of two aircraft itera-tions using different aerofoil geometry. The tableis a comparison of methods used throughout thedesign and the difference in accuracy.

NACA 6412 NACA 2412

Wing Span 2 2Mean Geometric Chord 0.29 0.29Reynolds number 290000 250000Wing inclination 3.5 4.5

TornadoCL 0.78 0.55CD 0.027 0.014

XFLRCL 0.66 0.36CD 0.022 0.009

Finite wing methodCL 0.73 0.57CD 0.055 0.031

Virtual wind TunnelCL 0.99 0.7CD 0.12 0.1

Fourth iteration: There were multiple options thatcould have been undertaken to reduce the amount oflift being generated by the wing, these options includedecreasing the wing span, or the angle of inclination atwhich the wings are fixed, or to change the aerofoil toone that has less camber. There were advantages anddisadvantages to each available option. The easiest op-tion would be to change the angle of inclination; howeverthis would mean that the wing would be less efficientand produce more drag than desired. If the wing spanwas decreased this would further reduce the aspect ratioand so would not make the wing as efficient.

Table 8: Analysis of aerofoil selection

NACA 2412 NACA 6412

Wing Inclination (degree) 4.5 0CL 0.7 0.63CD 0.1 0.1Lift (N) 42.7 38.45Drag (N) 3.9 4

The last option was to change the aerofoil to a lesscambered design. A comparison was composed of chang-ing the wing inclination of the NACA6412 to 0 degrees,and also using a NACA2412 at 4.5 degrees. From table8 it can be determined that it would be more efficient touse a NACA2412 aerofoil as a correct lift is generatedand although marginally smaller there is a lower value

18

of drag.

The decision to change the aerofoil was passed on tothe structural devision along with updated values to thevehicle technical specification to allow each section toupdate values where necessary.

5.1.3 Assessment of Analysis tools

Through the course of this design multiple analysistools where used. During the preliminary design stagetools such as Tornado, a Matlab script, and XFLR5 wereused. Aircraft changes were based on values found usingthese packages however on further examination usingmore accurate packages such as Virtual Wind Tunnelit was noticed that initial values calculated were muchlower than needed. The software used during the prelim-inary stage ended up forcing design changes that werereverted back during the detailed stage due to inaccu-rate values found. On closer examination calculationsbased on previous published wind tunnel test to find CLvalues proved more reliable than certain software andcorrelated with the Virtual Wind Tunnel results.

5.2. Structural

Initially a hollow shell was considered for a wing usinga high strength polystyrene, this was considered throughthe sweep wing and tapered wing designs, though it wassubstantially heavier than anticipated (around 1.8kg)and deemed too heavy, so ribs and spar models havebeen chosen since.

Moving away from a purely polystyrene based wingdesigns the adoption of balsa wood for spars and nylon6/10 for ribs. The addition of a low density polystyrenewas considered as a potential material to hold aerofoilform between spars allowing for better aerodynamics.Up to the current iteration, very little has changed ex-cept that balsa has been replaced with a stronger stifferwood such as bamboo.

Figure 13: Detailed view of main wing assembly

For the final iteration, a two spar system has beenadopted with the majority of the loads being constrainedby the leading spar, and the back spar is used to con-trol the ailerons (Fig.13). There were several importantconsiderations made when designing the spars: since thebending moments that act on the wing are the greatestin the centre it was important to make sure there wassignificant material there to resist these forces.

The shape of the wing uses a NACA2412 aerofoil anda triple taper with ratios of 0.89, 0.67 and 0.33 respec-tively, this provided some difficult challenges to run aspar the length of the wing and retain structural rigidity,with the tip aerofoil being 33% of the root aerofoil. Aftertesting three different configuration of spars (Straightdrafted spars; Angled and drafted; Drafted spars thatfollow the leading edge), the drafted and angled sparsprovided the best results in simulations, whilst still rela-tively light at 187g and would be one of the easiest tomanufacture being a less complicated design.

The rear spar needed to be at a constant distancefrom the edge in each design, as this would be used torotate the ailerons. It was also important to considermanufacturability, cost and weight of each design. Theservos are placed in the centre of the wing as this isgeometrically one of the only places they will fit. Ifplaced closer to the ailerons the polystyrene holding theservo in place is simply too thin to remain structurallysounds, and therefore are placed where the polystyreneis thickest, in the centre. Notably less ribs have beenincluded, since they are made from a denser materialthey added unnecessary weight.

Figure 14: Detailed view of tail plane

Similar restraints as the front wing governed the spardesign for the tail, with the rear spars needing to bea fixed distance from the leading edges, and taperingproved the same challenges as before. The servos for theelevator are located in the centre, much in the same as

19

Figure 15: Detailed design wing evolution

the front wing, as this was the only viable position forthem to actuate the elevators due to geometric reasons.The rudder servo is not located directly onto the rearspar, unlike all other servos in this model instead itactuates the rudder via wires connected to the spar andservo, this was due to there simply not being enoughspace underneath, or above, the rudder to fit a servo.

Figure 16: Detailed view of wing fixture to chassis.

The materials that were chosen for each part were asfollows: ribs: Nylon 6/10, spars: bamboo wood, inserts:polystyrene, as mentioned in more detailed in the ma-terial section. All simulations were done on Solidworks,and bamboo and polystyrene were imported as custommaterials with appropriate values, as shown in Section(6.1) later in this report. Construction of the wing andtail is designed to be as simple as possible, gluing eachrib and polystyrene insert in turn along the spars. The

wing and tail connect to the fuselage via the aluminiumfixtures to the leading spars and chassis as shown inFig.16. The wing cannot then move horizontally as it isrestricted by the polystyrene, and vertically the chassisitself restricts the movement of the trailing edge. Thisis designed so that the forces of flight are run throughthe chassis and not the shell.

Minimising the drag during flight is important forefficiency and speed. Although speed is not vital forthe Greenland mission, range and therefore efficiencyis. Working alongside control systems and stability eachcomponent is positioned at specific points through thefuselage, keeping this in mind a sleek aerodynamic fuse-lage was created.

Mathematical optimisation techniques were imple-mented namely Equ.13, based on Fineness ratio to getan aerodynamic fuselage sketch (Fig.17). As a result ahigh lift generating fuselage, aerofoil geometry, conceptwas drawn.

The taper ratio at the tail of the aircraft fuselage wasminimised to 1.4deg, this was in effort to reduce drag andto escape flow seperation. During each iteration (Fig.17)structural problems and serviceability problems werefound and as a result new iterations where produced.After the first iteration was created, simulations wererun and it was evident that it struggled with structuralloads at the tail of the aircraft therefore reinforcementswere needed which led to the second iteration.

Finenessratio =D

L= 0.074 (13)

20

 

Fuselage  evolution  

                                                                                                                   

 

                                                         

 

 

 

                             

 

 

                                                       

 

FIRST  ITTERATION  OF  THE  FUSELAGE  AFTER  THE  CONCEPT  Figure  1

SECOND  ITTERATION  Figure  2

MATHEMATICAL  OPTIMISATION  OF  FUSELAGE  DESIGN  FROM  AN  AIRFOIL  GEOMERTY  

Figure  3

THIRD  ITTERATION  ON  AERODYNAMIC  DESIGN  Figure4  

Airframe  structure  Figure  5 Fuselage  skin  of  2mm  thickness  

Figure  6

Assembly  of  airframe  and  the  skin  Figure  7  

Figure 17: Aircraft fuselage evolution

Multiple techniques were used and as a result, thethird iteration gave a better aerodynamic fuselage, bothin terms of fineness ratio and the tail taper ratio. Withthe overall geometry of the fuselage determined, a modelwas created ready for testing using a monocoque struc-ture. Working with the material section it was noticedthat this design made it difficult to service componentsand build in conventional managers meaning more com-plex materials would be required resulting in a heavierand less stiff fuselage. The fourth iteration was createdusing a complete airframe design, where a tray of in-ternal components could be placed within the airframefor quick access and ability to swap trays for immediaterelaunch capabilities. Each iteration can be found inFig.17.

5.3. Motor

The Power-train in the aircraft has evolved multipletimes over the course of the design. With changes toaerodynamic performance along with criteria set out byweight and performance it was crucial that the correctmotor was chosen.

Based on the current drag estimates under straightand level flight where thrust must equal drag equation(14) was used to determine power required.

PR = Tv (14)

The value of 56 Watts is required for straight andlevel flight, however this value is assuming perfect condi-tions therefore equation (15) was used to include motorefficiency, shaft efficiency and propeller efficiency.

PR = Tvηshaftηmotorηpropeller (15)

Based on equations (14) and (15) a total power of171.5 Watts is needed during straight and level flight.

A certain amount of redundancy is required for takeoff, strong winds and any other eventualities that mayoccur during a flight. Table 9 shows the characteristicsof the chosen motor.

Table 9: Motor selection

Voltage 7.4\14.8VRPM 1000 KvMax Power 430 WattsWeight 102g

Using the current motor and Equ.(6) it has beenfound that the aircraft will be able to sustain flight forup to 1.7 hours.

The propeller, and engine type proposed during thepreliminary design is still fit for purpose therefore a 10×5will be used along with an electric brushless motor.

5.4. Material

Edupack[?] was continuously used to determine vi-able materials based on two key areas, weight and stiff-ness. The two parameters are vital for wing design andgiven by equations (16) and (17)

m = ALρ (16)

S =CEA2

12L3(17)

Equations (16) and (17) are substituted into eachother to provide a material index which can then beapplied to the Edupack software.

m = L

√12L3S

C× ρ

E12

(18)

The information in equation (18) provides importantvalues to achieve the stiffest and lightest material fit forpurpose, the value of ρ

E0.5 must be as small as possible.This was applied to Fig.18 and is depicted as the

line that traverses the graph. Everything above this lineis a suitable candidate for material choice however thehigher the Young’s modulus the better. A criteria wasadded so that the overall weight of the aircraft did notrise significantly and therefore a maximum density of2000 kg/m3 was chosen eliminating metals.

From Fig.18 the decision to use bamboo was adoptedfor the main spar. This is due to the high stiffness andrelatively low density.

A light plastic film will be applied to the aircraftand act as the skin of the UAV. This was chosen dueto its waterproof properties and the heat shrinking thatallows for exact moulding. Low density polystyrene willbe placed in-between each rib to allow for support of theskin and ensure a perfect aerodynamic shape is kept.

5.4.1 Aircraft changes

Although the material selection has not had a majorimpact on aircraft size or shape it has been a majordriving force to completely change the internal structureof the fuselage. During the second iteration it was no-ticed that parts were not easily accessible for repairs andservicing (Fig.15). A larger problem was caused when itwas noticed that the fuselage would have to be manufac-tured in a very specific way using certain materials thatwhere not feasible for this mission. Therefore a completeoverhaul of the internal structure of the fuselage was

22

Figure 18: Edupack material selection softare[14]

done. Working along side the structure group a newairframe system was implemented and allowed for morematerial choice and manufacturing route.

The use of 3D printing is a large topic in the world ofmanufacturing, the versatility of printing broken partsalong with upgrading components is endless. Thereforeincorporating this was felt as important especially ifscientists are in remote areas, all they would require isa 3D printer to repair components. This move to 3DABS plastics was only possible due to the adoption ofan airframe instead of the heavier monocoque design.Under testing the ABS plastic was able to withstandthe forces that would be expected from flight.

5.5. Control Systems

The process to determine the sizing of ailerons, rud-ders and elevators was undertaken by using the AerofoilFlap Modeller script in Matlab. It takes the two dimen-sional curves of the aerofoil used on the UAV, changingsome figures by percentage of the aerofoil chord lengthto simulate a virtual control surface at a set range ofdeflection angles, calculating the changes in lift coeffi-cient and centre of pressure from the simulated controlsurface.

Figure 19: CoP change during aileron deflection

The Aerofoil Flap Modeller script has been run twicein order to size the ailerons, rudders and elevators. Fig.19relate to the aileron sizing; in Fig.19, it indicates thata smaller aileron with a 20% of the chord length, hassmaller changes in centre of pressure as it deflects, whileit is still generating sufficient lift coefficient changes toroll the aircraft. The smaller change in centre of pres-sure is desirable because it does not cause huge changesin drag produced at the aileron, it will give the air-craft a benefit in terms of power consumption duringmanoeuvring.

23

The same method was used to calculate elevator sizeand yielded 50% length should be used for both elevatorand rudder configuration.

Two methods were used to calculate aileron sizebased on the aerofoil flap modeller and the second basedon hand calculated that take into account roll rates andbanking moments.

To complete the hand calculations the following val-ues where used, Wing span = 2m, Take-off weight =3.5kg, Wing area = 0.5725m2, Aspect Ratio = 6.99, Ta-per Ratio = 2, Horizontal tail planform area = 0.0983m2,Stall Speed = 8.67m/s, and Bank angle = 10 degreesturning in 2 seconds.

ClδA =2CLawτCr

Sb[y2

2+

2

3(λ− 1

b)y3]yoyi (19)

Equation (19) was used to calculate the roll momentrequired for the aircraft, from this the position of theinboard and outboard aileron should be 60% and 95%of the wing span respectively. This gave rise to a valueof yi = 0.61m and yo = 0.95m. The difference of thesevalues allow for the aileron length of 0.34m and thedepth of the aileron was found to be 0.057m, Fig.20 is adiagram of the ailerons.

Figure 20: Aileron sizing

The calculations fit with the aerofoil flap modellertherefore these values were deemed satisfactory andproved sufficient during flight testing.

The sizing for both elevator and rudder were doneon the same bases and yielded the Fig.21 and Fig.22

Figure 21: Rudder sizing

Figure 22: Elevator sizing

5.6. Stability

The XFLR5 program, as mentioned in section 2.1,allows users to determine where avionics and flight in-struments should be placed within the aircraft. Usingthe measurements of the length of the fuselage thathave been provided, a model of the UAV has been cre-ated in this software, the user is able to move differentcomponents of the aircraft, the wings and weighted in-struments within the fuselage. A 3D model has beenproduced within the program to determine positioningof component to enable the structural group to buildcompartments for housing all the electronics. This inturn enabled for a more stable aircraft, Fig.23 depictswhere components should be placed for stability reasons.

Over the course of the design multiple iterations

have been created. Changes in weight, structure andcomponents all upset the balance of the aircraft. There-calculation of the static margin along with a smallshift in Cog was needed to maintain the desired 15%.The resulting calculations produced Table 10.

Table 10: Centre of Gravity position and component posi-tioning from nose of aircraft

Distance (m)

CoG 0.48Battery and camera 0.34

24

Figure 23: Positioning of components within the fuselage for stability reasons

5.7. Weight and Performance

5.7.1 Gliding flight

For gliding flight the most important factor is theglide ratio, with a ratio of 30:1 deemed a good value.The glide ratio for our UAV currently stands at 37:1above aspected. This glide ratio correlates to 117ft/mindescent with no throttle or pilot input.

5.7.2 Turning flight

Due to the induced drag increasing with the loadfactor, the thrust required for a level turn will be morethan for straight level unaccelerated flight. Manoeuvra-bility of the aircraft was not a very high priority. Evenso the turning radius was found to be 5 meters witha turn rate of 1.3 m/s. This will be adequate for theaircraft to successfully accomplish the mission. In thecase of the pull up and pull down manoeuvre the radiuswas found to be 13.47m and 1.9m respectively.

5.7.3 Landing

The aircraft was originally designed to used aparachute system, providing a soft landing to protectequipment onboard. However due to the gusty condi-tioned found in Greenland it was decided to scrap theparachute and opt for a belly landing which negates theneed to an undercarriage.

5.7.4 Range and Endurance

After the initial value of 43 km being to low an ad-ditional battery was added resulting in an endurance of1.7 hours for cruise flight enabling the aircraft to have a

theoretical range of 85 km, close to the groups’s goal of90 km.

5.7.5 Final design flight parameters

During the detailed design phase it is importantto update all calculated values as parameters changerapidly. As each sections make adjustments to specificareas the performance of the aircraft was updated andthe completion of the VTS was done. Fig.11 is a smallpart of the VTS showing key iterations along with theirupdated values.

Table 11: Final design flight parameters

Weight 3.5 kgStall Velocity 9.62 m/sTake off velocity 12.8 m/sMaximum rate of climb 5 m/sTime to climb 95 sTake off distance 6mRange 85 km

6. Aircraft testing

6.1. Structural

To test each spar a simulation was ran using150N/m2, this value was chosen using the V-n Diagramand by finding the maximum value during a violent ma-noeuvres. Pressure was applied to the outer surfacesin an upwards direction, and was fixed in place at theribs where they will join onto the chassis. The smallestmaximum displacement were at the tips, displaced by

25

0.13m, which in comparison to the 2m wing span is arelatively small amount.

Figure 24: Aircraft wing testing using a 150 N/M2 pres-sure.

6.2. Aircraft scaling

In order to simulate the aircraft in the Merlin flightsimulator the aircraft needs to be scaled up in a processidentical to the method used for recreating large aircraftin small wind-tunnels. To model in the Merlin flight sim-ulator the minimum weight for the aircraft is the drivingfactor. The Merlin flight simulator can accurately model12.5kg as minimum, in order to have a reliable model theaircraft has been scaled up proportionally by a factorof 3. To do this a series of equations were applied togive the dimensions necessary to accurately representthe aircraft dynamic stability.

Table 12: Components

Scale Factor Non-scaled Scaled

Linear dimension n 1 3Wing span (m) n 2 6Relative density 1 1.112 1.112

Weight, mass (kg) n3

σ 3.5 102

Moment of inertia IYY n5

σ 0.35 94

Moment of inertia IZZ n5

σ 1.1 286

Moment of inertia IXX n5

σ 0.7 195

Linear velocity (m/s) n12 14 24

Time n12 1 1.7

Reynolds number n1.5 vv0

250000 2050000

6.3. Stability

Analysis using the Simulink package, part of theMatlab group, was used to demonstrate the Stability ofthe UAV in dynamic flight.

An initial phugoid test was done to show that theaircraft’s static margin was correct along with proof of

a dynamically stable aircraft. Fig.25 shows the phugoidconverging, therefore it can be seen the aircraft is stableduring periods of acceleration and certain motions.

Figure 25: The phugoid test shows the aircrafts longitudinalstability

Figure 26: A similar phugoid motion tested in the MerlinFlight simulator

The same test was repeated in the merlin flight sim-ulator, and resulted in Fig.26. A large difference canbe noted in the time to dampen out the phugoid. Thiscould be due to many reasons including the simulinkmodel is not the most accurate software as input valuesare minimal. However it is believed to be largely dif-ferent because of the initial perturbation input in themerlin flight simulator.

Following from this test it is important to understandthe characteristics of the aircraft during climb as theaircraft can become unstable when climbing to rapidlyor when the aircraft is trimmed. Fig.27 and Fig.28 wasgenerated and can be seen that both tests produced sta-ble results with no simulation stalling or doing anything

26

unexpected.

Figure 27: This shows the aircrafts ability to climb rapidlywhile maintaining stable flight

Figure 28: The same variables as before however the au-topilot is switched on and corrections are madesimilar to trimming the aircraft.

It can be noted that at the start of both tests a smallphugoid is generated this is due to the instantaneous

acceleration from 0 to cruise speed when the simulationis started.

6.4. Flight Simulator

To verify that all calculations and theory works, theaircraft has been loaded into the Merlin Flight simu-lator. A number of flight tests were carried out andrecorded Table 13. Two flight parameters were plottedon graphs to graphically see and understand how theaircraft performs, (Fig.29) and (Fig.30).

Figure 29: This shows the rate of descent of the aircraftwith no input and without trim, it was able toreach a rate of descent of 117ft/min

Figure 30: The figure shows the aircrafts speed increaseuntil take-off is achieved at 43 knots

27

Table 13: Flight simulator test

Test Design Actual Comments

T/O distance 18m 20m Design value unscaledT/O speed 43 43 Without back pressureClimb to cruise 2.7 min 7:35 min From sea levelStall characteristics 32 kt 31 kt Design to not stall aggressivelyRate of descent 117 ft/min 130ft/min At 41 knots, 0 trim and no inputApproach 225 ft/min 225 ft/min at 2nm and 500 ft

Although many characteristics did perform as de-signed a few were out, for instance climb to cruise. Acalculated value of 2.7 mins was found to climb to thedesired cruise height however when tested a value of 7.5mins was experienced. This is not fully understood whyhowever an estimation of maximum rate of climb wasfound and used for all calculations that may not have

been achievable due to the tail stalling at low angles.

This also explains why it was not possible to stall theaircraft during flight. Although it is a desirable aspectcreating an aircraft extremely difficult to stall it hasnegative impacts on manoeuvrability and the rate ofclimb.

7. Aircraft costing

Table 14: Components

component Length (mm) width (mm) Height (mm) Weight (g) Cost(£)

Airframe 162

gopro hero 59 21 41 321 110Ardupilot 2.6 70 40 10 32 160Turnigy MX-353S 17g Servo x 4 38 13 27 17× 4 21.72Zippy Traxxas 7600mAh Battery x3 157 25 45 367× 3 69.42Turnigy Dual Power Unit 100 50 20 89 10.61Turnigy Plush 60A Speed Controller 80 31 14 60 20.78Turnigy D3536/8 1000KV motor 52 35 35 102 11.833DR uBlox GPS + Compass 38 38 8.5 16.8 97.403DR Video/OSD System Kit N/A N/A N/A 100− 150 189.99

Total 1594 743.75

8. Conclusion

The aim of this report is to show the possible so-lution, to creating a greenland aerial mapping vehicleto aid scientific discovery and conservation. Althoughthere are already possibilities out on the market, theproposed UAV would cost a fraction of this putting it inthe hands of a wider audience. With the possibility ofpooling informations from many sources, more accurateresults could be found thanks to this design.

9. Risk Assessment

28

Risk assessment

Group name: Absolute zero What are the hazards?

Risk

Who/What  might  be  harmed  and  how?   Risk Level

Prevention  of  Risk  

Battery-­‐  Zippy  Traxxas  7600mAh    

• Over  heating  • Movement  during  flight  • Leaking  

• Other  Components,  Airframe,  Risk  of  human  phyiscal  injury  (burns).  

• Cause  imbalance  with  the  aircrafts  stability,  possible  break  other  components.  

• Damage  to  components,  potential  irritation  to  skin  of  handlers,  enviromental.  

• Low    • Low    • High  

• Disconnect  when  not  in  use,  don't  over  insulate,  handle  with  care,  clearly  label  warning  on  battery,  do  not  use  battery  beyond  expiry  date  

• Ensure  fixed  within  airframe  before  every  flight.  

• Handle  with  care,  check  battery  exterior  prior  to  every  flight,  keep  battery  away  from  sharp  objects,  protect  from  potential  impact  -­‐  safe  storage    

Motor    /  Propeller  10x6"  

• Over  heating  • Electrocution  • Moving  parts  

• Damage  to  other  components,  and  airframe  structure,  anyone  handling  motor  after  use    

• Other  components,  anyone  handling  component  

• Anyone  near/handling  the  aircraft,  Risk  of  human  physical  injury  from  rotor  blades  

• Low    • Low    • Medium  

• Turn  off  when  not  in  use,  Limited  time  use,  avoid  time  spent  at  max  power,  allow  for  time  to  cool  after  landing  

• Disconnect  power  before  handling  

• Disconnect  power  before  handling,  handle  with  care,  keep  hands  away  from  device  when  active  

Camera  -­‐  GoPro  Hero  4  

• Movement  during  flight  

• Low  cause  of  imbalance  with  the  aircrafts  stability,  possible  to  break  other  components  

• Low     • Ensure  fixed  within  airframe  before  every  flight.  

General  electrical  equipment,  wires,  small  components    

• Over  heating  • Movement  during  flight  • Electrocution  • Sharp  wire  edges  

• Other  Components,  Airframe,  Risk  of  human  physical  injury  (burns)  

• Slight  imbalance  with  aircrafts  stability  

• Risk  of  physical  injury,  Other  components  

• Delicate  instruments,  risk  of  human  physical  injury  

• Low    • Low    • Low    • Low    

• Disconnect  from  power  source  before  handling  and  when  not  in  use,  allow  for  time  to  cool  

• Ensure  fixed  within  airframe  before  every  flight.  

• Disconnect  from  power  source  before  handling  and  when  not  in  use  

• Ensure  no  sharp  edges  on  majority  of  components,  precaution  to  be  taken  when  handling  

Obstacles   • UAV  collides  with  an  obstacle  during  flight  

• Entire  UAV,  risk  of  human  physical  injury  upon  landing/takeoff  

• Medium  

• Plan  Accurate  flight  path  before  every  flight,  observe  the  local  enviroment  before  flight,  Stick  to  CAA  requirements,  don't  fly  within  50m  of  buuldings/groups  of  people  

Flying  Animals   • Collision  with  flying  animals  

• Entire  UAV,  Wildlife  

• Low    

• Avoid  flying  near  known  nesting/roosting  sites  (if  it  can  be  helped),  observe  the  local  enviroment  before  flight  

Ground  Animals    

• Dangerous  animals  for  user,  or  collision  upon  Landing  

• Entire  UAV,  Wildlife,  Operator  

• Low    

• Check  local  enviroment  before  any  flight,  keep  a  safe  clearance  above  ground  during  operation,  in  terms  of  dangerous  wildlife,  follow  safety  leaflets  and  advice    

Weather  Conditions    

• Ice,  Snow,  Cold  Temperatures,  Strong  winds  

 

• Entire  UAV,  Operator    

• Medium    

• Check  Weather  forecast  before  every  flight,  wear  appropriate  warm  clothing  and  high-­‐grip  footwear,  keep  time  in  cold  to  a  minimum  

 

Systems    

• Power  failure  during  flight  • Communication  failure  

during  flight  • Control  system  failure  

during  flight  

• Entire  UAV,  (rare  likelihood  of  wildlife  or  people)  

• Entire  UAV,  (rare  likelihood  of  wildlife  or  people)  

• Entire  UAV,  (rare  likelihood  of  wildlife  or  people)  

 

• Medium  • Medium  • Medium  

 

• Routinely  check  equipment  and  aircraft  before  every  flight,  Do  not  operate  near  large  groups  of  people  or  buildings  

• Routinely  check  equipment  and  aircraft  before  every  flight,  Do  not  operate  near  large  groups  of  people  or  buildings  

• Routinely  check  equipment  and  aircraft  before  every  flight,  Do  not  operate  near  large  groups  of  people  or  buildings  

 

References

[1] Nichols.edu. Alpine Glacier MassBalance [Internet]. Available from:http://www.nichols.edu/departments/glacier/mb.htm

[2] Thomas Mlg. Worldwide glacier re-treat. RealClimate. available atwww.realclimate.org/index.php?p=129

[3] Caa.co.uk. Do I need a Permission for anUnmanned Aircraft (UAS) — Aircraft —Operations and Safety [Internet]. Availablefrom: http://www.caa.co.uk/default.aspx?catid=1995&pageid=16006

[4] Unmanned Aircraft System Operations In UKAirspace Guidance, CAP 722. 6th ed. CAA, 2015.Print.

[5] Web.mit.edu. [Internet]. 2013 Available from:http://web.mit.edu/drela/Public/web/xfoil/

[6] Redhammer.se. Tornado, the Vortex lat-tice method. [Internet]. Available from:http://www.redhammer.se/tornado/

[7] Xflr5.com. XFLR5 [Internet]. 2015. Available from:http://www.xflr5.com/xflr5.htm

[8] Airfoiltools.com. Airfoil Tools [Internet]. Availablefrom: http://www.airfoiltools.com/

[9] High lift generating fuselage concept http ://www.ijetae.com/files/V olume2Issue5

[10] Quantum Devices INC. Brushless Mo-tors vs Brush Motors, what’s the differ-ence? [Internet]. 2010 Available from:https://quantumdevices.wordpress.com/2010/08/27/brushless-motors-vs-brush-motors-whats-the-difference/

[11] Brown, M. (2014). Sizing RC Airplane Pro-pellers. [online] Hooked on RC Airplanes. Availableat: http://www.hooked-on-rc-airplanes.com/sizing-rc-airplane-propellers.html

[12] Carpenter. P. RC Airplane Propeller Size Guide[Internet]. Rc-airplane-world.com. 2015. Availablefrom: http://www.rc-airplane-world.com/propeller-size.html

[13] Altairhyperworks.com. HyperWorks: Open Archi-tecture CAE solution [Internet]. 2015 Availablefrom: http://www.altairhyperworks.com/

[14] CES Edupack. (2014). United Kingdom: Granta.

[15] Hirrus mini UAV system [Internet]. 1st ed.Bucharest: TeamNet International S.A;2015 [cited 8 March 2015]. Available from:http://www.aft.ro/bro.pdf

[16] Aeromao.com. Aeromao - Aeromapper 300 [Inter-net]. 2015 [cited 2 March 2015]. Available from:http : //www.aeromao.com/aeromapper300

[17] 3. Fpvflying.com. Mugin 2600 UAV FPVplatform - FPV flying [Internet]. 2015[cited 18 March 2015]. Available from:http://www.fpvflying.com/products/Mugin-2600-UAV-FPV-platform.html

[18] HobbyKing Store. Skywalker X8 FPV / UAV FlyingWing 2120mm [Internet]. 2015 [cited 12 May 2015].Available from http://www.hobbyking.co.uk/

[19] HobbyKing Store. Durafly Zephyr V-70 High Per-formance 70mm EDF V-Tail Glider 1533mm (PNF)[Internet]. 2015 [cited 12 May 2015]. Available fromhttp://www.hobbyking.co.uk/

[20] HobbyKing Store. UAV-3000 Composite FPV/UAVAircraft 3000mm (ARF) (EU Warehouse) [Inter-net]. 2015 [cited 12 March 2015]. Available fromhttp://www.hobbyking.co.uk/

10. Appendix 1 (VTS)

32

Vehicle  Technical  Specification  

 

11/2/14   18/11/14   25/11/14   4/12/14  

 3/2/15  

 23/02/15   24/02/15   25/02/15   11/3/15  

Main  wing  geometry    

  Wing  Span,  b   2   3   2.4   2.4    

2.4  

2.4   2.3   2   2  

Cord  length,  c   0.2   0.3   0.3   0.3    

0.3  

0.3   0.3   0.3   0.3  

Root  Chord    

0.45   0.45  

Surface  area,  s   0.4   0.9   0.6154   6154    

0.6  

0.6   0.634   0.5725   0.5725  

Aerofoil   NACA2412   NACA2412  NACA6412  NACA6413    NACA6412  

NACA6412   NACA6412   NACA6412   NACA2412  

oswald  efficiency,  e   0.7   0.8   0.8   0.8    

0.8  

0.8   0.8   0.8   0.8  

Aspect  ratio   10   10   9.37   9.37    

9.6  

9.6   8.73   7   7  

Mean  Aerodynamic  Chord    

0.26  

0.26   0.30162   0.31252   0.31252  

Mean  Geometric  Chord    

0.25  

0.25   0.26359   0.28625   0.28625  

Span  partition  1    

0.4   0.4    

0.4  

0.4   0.4   0.15   0.15  

Span  partition  2    

0.8   0.8    

0.8  

0.8   0.6   0.25   0.25  

Span  partition  3    

0.6   0.6  Sweep  angle  (degrees)  partition  1  

  0   0  

 0  

0   0   0   0  

Sweep  angle  (degrees)  partition  2  

  10   10  

 10  

10   10   10   10  

Sweep  angle  (  degrees)  partition  3  

  10   10  

Taper  ratio  partition  1    

1   1    

1  

1   1   1   1  

Taper  ratio  partition  2    

2   2    

2  

2   2   1.125   1.125  

taper  ratio  partition  3    

2   2  

Wing  Inclination    

3.7   3.25   3.5   4.5  

partition  1  

root  chord  =  0.45  tip  

chord  =  0.4    

partition  2  

root  chord  =  0.4  tip  chord  

=0.3    

partition  3  

root  chord  =  0.3  tip  chord  

=  0.15    

Winglet  Span    

0.15     Winglet  Sweep  

 

20    

Winglet  Aerofoil    

NACA0018     Winglet  Dihedral  

 

90    

Tail  wing  geometry     Veritcal  tail     cvt   0.3   0.03   0.03   0.03  

 0.03  

0.03   0.03   0.03   0.03  

Vertical  tail  chord    

0.16   0.16    

0.16  

0.2   0.2   0.2   0.2  

verticle  height,  hvt   0.12   0.405   0.324   0.2592    

0.2592  

0.2767   0.22   0.22   0.22  

SVT    

4.43E-­‐02  

4.43E-­‐02   4.40E-­‐02   4.40E-­‐02   4.40E-­‐02  

Sweep  angle  (degrees)    

10   10    

10  

10   10   20   20  

Taper  ratio  of  Vertical  tail    

2   2    

2  

2   2   2   2  

Aerofoil    

NACA0012    NACA0012  

NACA0012   NACA0012   NACA0012   NACA0012  

 

Horizontal  Tail     cht   0.6   0.6   0.6   0.6  

 0.6  

0.6   0.6   0.6   0.6  

horizontal  cord,    

0.2   0.2    

0.2  

0.2   0.2   0.2   0.2  

horizontal  span,  spanht   0.24   0.54   0.864   0.864    

0.576  

0.48   0.49   0.46   0.66  

SHT    

0.8856  

0.072   0.0727   0.069   0.0983  

Sweep  angle  (degrees)    

10   10    

10  

10   10   10   10  

Taper  ratio  of  Horizontal  tail    

2   2    

2  

2   2   2   2  

Aerofoil    

NACA0018    NACA0012  

NAAC0012   NAAC0012   NACA0012   NACA0012  

  Tail  Aerodyanmics    

tail  volume,  VH    

0.2807929802  

0.2807929802  

 

0.5399865003  

0.45   0.43  

 0.52  

Cltalpha  3deg    

0.36  

Weight  and  performance    

Weight  (kg)     6    

6   6    

4.3   4.3   4.3   3.5  

Vstall  (m/s)    

9.4329   9.1334   10.0051    

8.47   8.24   8.67   9.62  

Vtakeoff  (m/s)    

11.3195   10.96   12.0061    

10.16   9.89   10.41   11.54  

Absolute  ceiling    

7700   7800    

6600   14700   14000   14600  

Service  ceiling      

13700  Maximum  rate  of  climb,  ROCmax  

 

2.010347135   2.0333  

  1.07   11.14   10.97   11.1  

Load  factor  (maximum)    

1.33  

Time  to  climb    

541.1524   5.34E+02    

1.03E+03   9.48E+01   94.79   95.22  

Sink  rate  at  13.889  m/s    

0.981   9.10E-­‐01    

0.65   0.65   0.95   0.8  

Take  off  distance    

72    

61   61   62   8.84(S_G)  

Range    

1.41E+05   1.41E+05  141392.642

2   4.15E+04  

Endurance     Turning  radius    

5.06  

Pull  up  radius    

13.47  

Pull  down  radius    

1.9  

Max  payload  (N)     Ix    

0.531   0.333   0.252   0.252  

Iy    

0.412   0.4071   0.407   0.407  

Iz    

0.904   0.6153   0.523   0.523  

 

Aerodynamics    

CD0   0.01   0.01   0.025    

0.024  

0.024   0.024   0.024   0.013  

Clmax   1.25   1.25   1.6    

1.6  

1.6   1.6   1.6   1.3  

Clalpha   0.1   0.1   0.1    

0.1  

0.1   0.1   0.1   0.1  

Stall  angle  of  attack  Degrees    

11.4176    

11.4176  

11.4176   9.915   11.9021   inclination  angle   4   4   4   4  

 4  

 3.7   3.25   3.5   4.5  

CL  STLUF  Tornado    

0.75069   0.78    CD  STLUF  Tornado  

  0.023935   0.0275  

 

 

CL  STLUF  matlab   0.5535   0.5535   0.6919      

0.826   0.826      CD  STLUF  matlab   0.0203   0.0203   0.0488  

   0.0481   0.0481  

   CL  STLUF  XFLR      

0.798  

0.7682   0.7682      CD  STLUF  XFLR  

    0.0244  

0.0225   0.0225  

   Lift  Tornado    

51.007   47.8955    Drag  Tornado  

  1.6263   1.6868  

 Lift  matlab   34.9367   47.1645   60.0968    

82.2606    

53.1547   53.1547      Drag  matlab  

   3.0953   3.0953  

   Lift  XFLR      

49.4352   49.4352      Drag  XFLR  

   1.4479   1.4479  

   WIND  TUNNEL  CL    

0.69479  

WING  TUNNEL  CD    

0.10156  

Reynolds  number   175000   290340   290340    

290340  

290340   290340    

250000  

Static  margin  (Tornado)    

0.50502    

0.2117     Neutral  Point  

  0.3235  

 0.3728  

Hand  calc.  Static  Margin    

0.1515    

0.2359  

0.2115   Centre  of  Gravity  coords  

  0.075  0  0   0.381  

  Center  of  Pressure  Coords    

0.38    

Power  available,  Pa0    

1300      

90   90   430   Static  thrust,  kg   4  

      1.8   1.8  

  Dynamic  thrust  (w/o  efficiencies)  

  6.5   6.5   36  

Engine  efficiency    

85%   85%    

85%   85%   83%   Propeller  efficiency  

    45%   45%   45%  

Engine  weight      

160g   160g   102g   propeller  weight  

    35g   35g   35g  

propeller  size    

15x8  

10x5   10x5   10x5  

Control  Systems    

APM  2.6  Autopilot   28g  

7  cm  x  4.5  cm  x  1.5  

cm  

TTL  3DR  Radio  3DRobotics  Telemetry  433Mhz  module  

 

6.9  cm  x  1.7  cm  x  0.5  cm  

MAVLink-­‐OSD   6g  

19mmx39mm  (no  include  

connector)  *UART  

4Pin  cable  150mm  

APM  Power  Module   17g  

25mm  x  21mm  x  9mm  

GoPro  Hero4  Silver  Edition  

83g  /  147g(with  housing)  

    ZIPPY  Traxxas  7600mAh  2S1P  30C  Lipo  Pack  x3  

 

367g  x3  =  1101g  

157x45x25mm