UAV Antarctica

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Bachelor Thesis UAV Antarctica

David Dahl ddahl@kth.se

Fredrik Stetler fstetler@kth.se May 26th, 2014

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Abstract

One of the biggest problems of our time is the global warming. A direct result of this phenomena is the melting of ice of the glaciers on the north- and the south pole. As this continues, the melted ice will contribute to an increase of the sea level, and may cause enormous natural disasters. To be able to prevent this, it’s important to study its affects. This reports contains a concept study of a Unmanned Aerial Vehicle, a UAV, set on the coast of Antarctica by the Australian owned base Davis Station to document the changes and retracting of the glacier borderline. The purpose of the aircraft is to scout a pre-determined path whilst documenting the glaciers with photography from above.

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Foreword

A special thanks would like to be denoted to our mentor Arne Karlsson for his help and guidance throughout this project. We would also like to thank the Royal Institute of Technology for the possibility to carry out this project. Lastly we would like to thank the Royal Institute of Technology’s library for its contributions with information on the topics in hand.

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Introduction

Aerial vehicles play a huge role in todays society. They have become a great part in ev- erything from personal traveling to global warfare. Most aircrafts today are controlled by a pilot, and just like every other human being, a pilot makes errors, which may cause devastating results. For pilots in big commercial aircrafts, a mistake can lead to their own and hundreds of people’s deaths, and for military pilots the cost can be both their own lives and multi million dollar weaponry. So what if the human aspect of flight was removed? There is now a solution to this, and its called a UAV.

A UAV, or as it is commonly known, a drone, is an aerial vehicle that is controlled during flight by a computor, i.e. no human pilot aboard. The way it is controlled is either autonomously or by a remote control from the ground or another vehicle. The first UAV was simply a pack of unmanned, bomb-filled balloons, sent out by the Austrians as a way to attack Venice. The concept of a real unmanned aerial vehicle has long been sought out after. As technology got better, more advanced types of UAVs could be used during warfare, like during WWI.

Today the UAVs are most known for their participation in the US Military, where it is a great insurance for the concern of losing a valuable pilot for a dangerous mission.

For very simple missions, like scouting an area where it can be very dangerous to be due to the high enemy density in that specific area, the UAV is perfect. This way no pilot is harmed, but the mission can still executed and successful. The UAV is often used outside of the military aswell. It has a very wide range of use, as for example exploring oil-, gas-, and mineral sites, and is therefore used alot by private owned businesses.

Another example is exploring just what this report is about, glaciers.

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Background

The climate change is one of the biggest topics around the world in this day and age.

Due to global warming, caused by greenhouse gases leaking into the atmosphere, dras- tic changes are occurring on earth. One of the biggest affects global warming has on the earth is the melting of ice on the north and south pole, causing the sea level to rise.

One very important thing to do is to measure these changes, to better be able to un- derstand what is happening and the repercussions it has. This is the reason for this project. By doing a concept study of an unmanned aircraft with a mounted camera, we can plan a determined route for it to fly. Whilst flying, it will take pictures of the glacier below it. After completing its route, it will safely land on an airstrip and wait a set amount of hours until the next flight, while recharging. This will occur once a day, during a six month period. At the end of this half year, sufficient amount of data will have been collected to be able to study the glaciar changes over time.

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Chapter 1 Requirement specification

1.1 Mission profile

The aircraft will first takeoff and begin it’s ascent. Within 10 km, the aircraft must reach an altitude of 1 km. The glacier that will be observed is about 80 km long, and the aircraft will remain in steady state level flight until this entire distance has been covered. The total mission distance is approximately 100 km.

Figure 1.1: Mission profile 1.1.1 Takeoff and ascent

During takeoff the aircraft should get airborn within 300 m. It should be able to start from ice aswell as dry ground. The aircraft should lift off at a speed that is sufficient enough to satisfy the lift to weight ratio. The aircraft should climb with a fix angle and speed, and since the aircraft will operate in the Antarctic region, there are no legal restrictions to the rate of climb during ascent. However, the climb angle must be great enough for the aircraft to reach cruising altitude within a distance of 10 km.

1.1.2 Steady state level flight

During steady state level flight the aircraft should maintain a cruising speed at 100-150 km/h. To be able to clearly observe the changes of the glaciers, the aircraft should keep

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a cruising altitude of 1 km. This state of flight does not require any harsh manourvers at all. The cruising range should be at least 80 km.

1.1.3 Descent and landing

The engine will be shut off to save energy and the aircraft will glide during the decent.

The glide distance depends on the rate of decent, which is not restricted initially. Before landing, the speed will be reduced so that the aircraft can be able to touch down and brake within 300 m. Just as for the the take-off process, the aircraft should manage to land on icy ground aswell as dry ground.

1.2 Weigth and geometry

The total take-off mass of the aircraft should not exceed 300 kg. The aircraft geometry should be of conventional fashion, with a main wing attached to the body, and a tail with a horisontal stabilizer and vertical stabilizer. The wing span as well as the total lenght of the aircraft should be less or equal to 8 m. The main body must be capacious enough to carry the equipment required for the mission.

1.3 Propulsion system

The propulsion system in this case, will consist of an electric engine, and a propeller.

The engine should be a brushless permanent magnet synchronous electric engine, and it should be able to deliver the required amounts of power for the different flight phases.

The engine mass should not exceed 50 kg. It should be able to operate in sub zero temperatures without significantly losing performance. During steady state level flight, the engine should not operate above the continuous power setting.

The propeller will be attached in a pusher-configuration at the rear of the aircraft. The blade diameter of the propeller will be restricted to the start and landing phase, since the propeller must not hit the ground during rotation.

1.4 Power source

A direct current battery pack will be used as a power source. It should be powerful enough to sufficiently fuel the aircrafts trip from start to landing. The battery pack must be rechargable, and be able to discharge and recharge enough times without significantly lowering the performance. The batteries should be able to withstand sub zero temperatures, down to −20^◦C. The total mass of the battery pack must not exceed 50 kg.

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Chapter 2 Perfomance analysis

2.1 Aerodynamics and control surfaces

2.1.1 Basics

As the aircraft is moving, the entire body is exposed to aerodynamic forces, changing the motion of the aircraft. The wing is the most essential part of the aircraft. It generates next to all of the lift force, which makes it possible for the aircraft to get airborne. The body aswell as the wing also contributes to the drag force. The magnitude of the lift force and the drag force depends on multiple variables, such as wing shape, air speed, air density, wing size, body shape and more. The drag force and the lift force can be expressed as

D = C_D· ρ · S ·v²

2 (2.1)

L = C_L· ρ · S ·v²

2 (2.2)

where CD and CLare non-dimensional constants, determined by the shape of the wing and body, ρ is the density of the surrounding air, S is a reference area of the wing, and v is the velocity of the aircraft. A very common way of expressing CD is using the drag polar

C_D = C_D0+ C_Di= C_D0+ K · C_L² (2.3) where C_D0is the zero-lift drag coefficient and C_Diis the lift-dependent drag coefficient.

C_D0can be expressed as

C_D0= C_{F e}·S_wet

S (2.4)

where CF eis the equivalent skin-friction coefficient, Swetis the total wetted area of the entire aircraft. The drag-due-to-lift factor, K, is usually defined as

K = S

π · b²· e₀ (2.5)

where b is the wingspan, and e₀ is the Oswald efficiency factor.

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2.1.2 Control surfaces

Control surfaces on aircrafts controls the movement around the axes of motion, see Figure 2.1.

Figure 2.1: Control surfaces and axes of motion

The ailerons control the movement around the roll axis, the rudder controls the movement around the yaw axis and the elevators control the movement around the pitch axis.

These three surfaces are called primary control surfaces.

There are also surfaces called secondary control surfaces such as flaps, slats, and air brakes, which are mostly used on large and heavy aircrafts. The flaps and the slats are used to increase the camber and often the area of the wing, making it more effective at low speed, thus creating more lift. Air brakes are used to increase drag.

Ailerons

When initially designing the ailerons, there are some rule of thumbs which can be used [18]. The aileron-to-wing-chord ratio, Ca/Cw, is usually about 15 to 25 percent, the aileron-to-wing-span ratio, b_a/b_w is about 20-30 percent. Also, the max deflection, A_maxis usually around 20-25 degrees.

Rudders

In the design of the rudder, four parameters must be determined; rudder area, Sr, rudder chord, Cr, rudder span, br, and maximum rudder deflection, Rmax. Good inital values are Sr/Sv = 0.38, Cr/Cv = 0.42, and Rmax = 25 degrees. [18]

Elevators

In the design of the elevator, four parameters should be determined. They are elevator planform area, Se, elevator chord, Ce, elevator span, be, and maximum elevator deflection, Emax. As a general guidance [18], the typical values for these parameters are as

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follows Se/S_h= 0.4−0.15, b_e/b_h= 0.8−1, C_e/C_h = 0.2−0.4, and E_max,up= −25 degrees, Emax,down = +20 degrees.

2.2 Phases of flight

2.2.1 Takeoff and ascent

During takeoff, the propeller must produce enough thrust power, Tprop, for the aircraft to accelerate from zero to the lift-off speed, vLOF, within a certain distance. The balance equation during ground roll is

→ : T_prop− D − f_roll= m · a (2.6)

where m is the total mass of the aircraft, froll is the roll friction force and a is the aircraft’s acceleration.

Figure 2.2: Forces during takeoff The propeller thrust force can be expressed as

T_prop = P_eng· η_prop

v (2.7)

where Peng is the engine power and ηprop is the propeller efficiency. The roll friction force depends on the generated lift force, and on the roll friction coeffiecient, µ_roll.

froll = µroll· (W − L) (2.8)

where W is the total weight of the entire aircraft. The value of the roll friction coefficient depends on the contact materials. Eq. (2.6) is a non-linear ordinary differential equation. It’s possible to obtain speed data, time data, distance data and more, by solving this equation numerically in MATLAB using the function ODE45.

When the aircraft has reached its lift-off speed, it will leave ground and begin its transition to climb. The lift-off speed, is set to be 20% greater than the stall speed, v_s. The

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stall speed is the absolute minimum speed an aircraft can fly at, and can be expressed as

V_s= s

2 · W

ρ · C_L,max· S (2.9)

where C_L,maxis the maximum value of of C_L, which varies for different wing config- urations. Bacause of the short endurance of the transition to climb, no further performance analysis has been done regarding this flight phase.

Figure 2.3: Forces during ascent

The aircraft will keep a constant speed during the ascent, at a fixed angle of climb, γ_climb, which is the angle between the horisontal axis and the aircraft’s flight path. The balance equations can be expressed as

-: L − W cos(γ_climb) = 0 (2.10)

%: T − D − W sin(γ_climb) = 0 (2.11)

The rate of climb, R/C, is the vertical climb speed and it can be expressed as

R/C = v_climbsin (γ_climb) (2.12)

Since the air density is decreasing with an increase in altitude, and since the speed and angle of climb is required to be constant, the propeller thrust and the engine power will change with altitude. The times it takes for the aircraft to reach cruise altitude is calculated by

∆t_ascent = t₂− t₁ =

t2

Z

t1

dt =

h2

Z

h1

dh

R/C (2.13)

2.2.2 Steady state level flight

To maintain steady state level flight, it’s required that

↑: L − W = 0 (2.14)

→: T_prop− D = 0 (2.15)

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The thrust vector isn’t actually completely parallell to the speed vector. However, the angle between them is often insignificantly small, and it can therefore be neglected in most cases.

Figure 2.4: Forces during steady state level flight

The velocity at which the required engine power is at it’s minumum during steady state level flight is [14]

V_Pr,min= s

2 ρ

·

r K

3 · CD0

· W S

(2.16) To prevent instability issues, the cruise speed if set to be 35 % greater than this.

2.2.3 Descent and landning

Figure 2.5: Forces during descent

During the descent, the engine will be shut off and the aircraft will glide it’s way down.

As earlier mentioned, the air density will change with changes in altitude, and this will affect the angle of descent since the speed is required to be constant. The balance equations for the descent are espressed as

%: L − W cos(γ_descent) = 0 (2.17)

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-: D − W sin(γ_descent) = 0 (2.18) Because of the short endurance of the transition to landing, no further performance analysis has been done regarding this flight phase.

Figure 2.6: Forces during landing

For the landing phase, an analythical model can be applied to determine the performance. Two constants for the landing are calculated through the equations below.

A_break = g · (−µ_roll,b) (2.19)

where µ_roll,bis the friction coefficient with breaked tyres.

B_break= g

W(0.5 · ρ · S · (C_D− µ_roll,b· C_L)) (2.20) The speed at which the wheels begin to break, is usually lower than the touch down speed. An assumption is this case is that the aircraft will begin to break one second after touch down, calling that thold, and that the break speed, vbreak, is equal to the touch down speed, vT D.

The total distance it takes for the aircraft to come to a stop is given by S_land= S_roll+ S_break= t_hold· v_{T D}+ 1

2 · B_breakln

1 −B_break A_break · v_{T D}²

(2.21) Just as for the take-off process, the performance of the landing process can also be determined by numerical calculations, using the function ODE45.

2.3 Propulsion system

For a propeller propulsion system, the main system parameter will be the power. For this project, an electric engine will be used. The shaft power generated by the electric engine, P_eng, will be greater than the propeller thrust power, P_prop, due to the propeller

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efficiency, ηprop. The relation between the propeller thrust power and the engine shaft power is

Pprop= ηprop· P_eng (2.22)

The propeller thrust power at steady state flight is Pprop= Tprop· v, which implies that the engine shaft power can be written as

Peng= Tprop· v ηprop

(2.23) The key feature of the electric engine is that the voltage is constant and the current varies. Given a steady voltage across the leads of a engine, the rotational frequency will remain constant regardless of load. The higher the power load on the engine, the more current is drawn from the power source. A way of expressing the propeller efficiency is by applying the Actuator Disk Theory [1]. This theory doesn’t take the rotational frequency into consideration. The efficiency can then be expressed as

ηprop= 2

1 +q

1 +_A^T^prop

disk·v²·^ρ₂

(2.24) where A_disk is the propeller disk area. If the required thrust and propeller diameter is known at a certain airspeed and altitude, the propeller efficiency can now easily be calculated. The propeller tip Mach number should be kept below 1 to prevent high noise levels and unexpected viscous effects. It can be expressed as

Mtip = q

(π · D · n)²+ v²

a (2.25)

where n is the rotational frequency of the propeller, D is the diameter of the propeller and a is the speed of sound.

2.4 Energy demand

The aircraft will use a rechargeable battery pack as power source. It has to provide the engine and all the aircraft’s avionics with enough power to complete the mission with margin. The total amount of energy (Joule) required to power the propulsion system for the entire flight is calculated with

∆E_eng= ∆t ·

n

X

i=1

Pi,eng

η_eng

(2.26) where n is the number of flight phases. Every avionic component requires a specific voltage and current. The energy amount required to power the avionics is

∆Eavionics= ∆t ·

n

X

j=1

(Ij,req· U_j,req) (2.27)

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where n is the number of avionic components. To get the total amount of energy required for the entire trip is calculated by

∆E_tot = ∆E_eng+ ∆E_avionics (2.28)

The total energy consumption can easily be expressed in kilowatt hours kWh. One Wh is equal to 3 600 J, which implies that,

kW h_tot = ∆E_tot

3600000 (2.29)

2.5 Stability

2.5.1 Center of gravity

Regarding flight stability, it’s very important that the center of gravity, CoG, is placed correctly. The CoG is often designated to be placed within the main wing’s chord. The CoG can be expressed with the following coordinates below.

XCoG= P W (i) · x(i)

W (2.30)

Y_CoG= P W (i) · y(i)

W (2.31)

ZCoG= P W (i) · z(i)

W (2.32)

where W(i) is the weight of component i. The x-axis is placed along the roll axis, the y-axis is placed along the pitch axis, and the z-axis is placed along the yaw axis.

2.5.2 Pitch stability and trim

To begin with, the aerodynamic center, a.c, of the main wing is the point at which the pitching moment coefficientfor the main wing does not vary with lift coefficient (i.e.

angle of attack). For thin symmertical airfoils, the a.c is usually placed around 1/4 of the chord. This is not entirely true for cambered airfoils, but it’s a good approximation.

With that said, the pitching moment about the aircraft’s CoG can be expressed as C_M,CoG= xw

c · C_L,w− xt· S_h

c · S_w · C_L,h+ C_M,a.c,w (2.33) where xw is the distance from the main wing’s aerodynamic center back to the CoG, c is the chord of the wing, C_L,wis the lift coefficient for the main wing, x_tis the distance from CoG back to the a.c of the horisontal stabilizer, Shis the stabilizer reference area, Sw is the main wing reference area, and CM,a.c,w is the wing pitching moment about its a.c. Trim can now be achieved by setting the incidence of the tail surface (which adjusts its CL) to make CM,CoG= 0

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Chapter 3 Dimensioning

3.1 Weight and geometry

The aircraft has been designed to meet many different requirements. The final wing span has been set to six meters, and the fuselage’s length to four meters. The fuselage will be made out of carbon fibre to minimize the weight while keeping some strenght and flexibility to the structure. For the fuselage to be capacious enough, the front end has been designed a little bit larger, with a maximum diameter of 0.5 m at the attachment of the main wing. The landing gear will be a tricycle-configuration, with the back wheels positioned approximately 0.5 m behind the CoG, and the front wheel positioned in front of the wing. After take-off, the landing gear will retract to a slick position outside and along the fuselage. The final total weight of the aircraft is 134 kg.

3.2 Electric engine

For the propulsion system, a six phased synchronous brushless electric engine has been chosen. Unlike brushed engines, brushless engines are purely inductive. This means that the engine life is limited primarily by the bearings, which makes it very reliable.

The engine is provided by Joby Motors [10]. They primarily producine lightweight, ultra-efficient engines for use in electric aircrafts. The continuous engine power required during steady state level flight has been calculated to 7 kW. The engine chosen has a continuous shaft power of 14 kW, which is sufficiently enough.

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Figure 3.1: J2 Engine

Construction Inrunner

Nominal voltage 100-600 [V]

Poles 46

Nominal RPM 2500

Maximum RPM 3500

Diameter 200 [mm]

Mass 4000 [g]

Length 75 [mm]

Continuous Torque 53 [Nm]

Continuous Shaft Power at Nominal RPM 14 [kW]

Peak torque 80 [Nm]

Peak Shaft Power at Nominal RPM (15s) 20.9 [kW]

Table 3.1: Engine specifications

The engine will be controlled by an Electronic Speed Controller, ESC. The function of the ESC is to obtain flight status data from the flight control system, such as flight speed, angles, and accelerations. With this data, the ESC will vary the current that is applied to the engine, thereby controlling the torque for each point in time. No specific ESC has been chosen for this engine.

3.3 Propeller

The propeller is provided by Pipistrel [9], which manufactures three different kinds of propellers, the LN, the BAM, and the VARLO, with four, three, and two blades respec- tively. They are designed to be used for ultra light and experimental aircrafts. The one chosen for this project is BAM. The layer and inside parts of this propeller are made of

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composite materials, fibre reinforced plastics, which makes it very light and tough. The base of the blade is made of aluminium and stainless steel tube.

Figure 3.2: Propeller specification

3.4 Battery pack

The total energy requirements has been calculated to approximately 7 kWh for the entire trip (see Appendix/Matlab code/169). The energy contribution from the avionics have been neglected, due to its insignificance. By knowing this energy amount, a battery pack can be chosen for the assignment. To extend the life cycle, the battery’s depth of discharge should not exceed 70%. In addition, due to unknown variable changes like head wind, colder temperatures etc. an energy excess of an additional 20 % will be required. By using these numbers, the final energy requirements is calculated to approximately 12 kWh.

One of the more optimal choices for a battery pack that will operate in a cold climate are battery packs using lithium polymere cells. These are among the most powerful rechargeable batteries on the market due to their high ratio of dimensions to weight and capacity, and are especially good during colder environments compared to other sorts.

Lithium polymere cells have an volumetric energy density of 350Wh/L and gravimetric energy density of 135Wh/kg [8]. To meet the required energy need for this project, with today’s polymere cells, the battery pack would need to have a volume of at least 34 litres and a mass of 89 kg. However, with today’s rate of development, a presumption is a 50% increase of the gravimetric energy density within 20-25 years. For the final calculations, battery mass is set to 50 kg.

A battery pack for this application needs a Battery management system, BMS. The BMS works by managing the different levels of the output, making sure it doesn’t operate outside it’s safe zone, monitoring its state, controlling environment etc. No specific BMS has been choosen for this project.

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Figure 3.3: Lithium polymere cell

3.5 Charging

The aircraft will take its place in the hangar after its route where it will be plugged in and recharged. Solar panels will be used for this purpose. The solar panels will be put on the roof of the hangar, and they must be dimensioned to be able to fully recharge the batteries within its time frame. The solar panels are operating with their highest efficiency when facing the sun directly. This will be dealt with by using a small engine with two axles, spinning and tilting to autonomously track the sun. The latest solar panel on the market right now is from the Deutsch corporation Fraunhofer, which with world breaking technology has an efficiency of 44.7 % [7]. A reasonable assumption is that this number will rise to 60 % within 20 years. (6 % over an 8 year period, 12 % over 16 years, assuming linearity, most likely exponentially).

Assuming regular solar power from the sun, at least a 1000 watts per square meter [6], and an minimum time of sun exposure of 8 hours a day [5], the time, t, to recharge the battery with a solar panel of 4 square meters is

t = ∆Etot

1000 · 0.6 · 4 = 3h (3.1)

3.6 Wing design

3.6.1 Profile

The inital idea for the wing profile was to choose a cambered one with a high lift and relatively low drag. Cambered profiles are used on most gliders and low speed aircrafts, so it seemed to be a good idea. In the end it came down to two different but similar profiles; Clark-Y and the NACA 2412. To analyse the performance of these airfoils, the

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platform XFLR5 [4] was used, which is an analysis tool for wings, planes and profiles.

All calculations were made with a Reynolds nummer of 1.6 · 10⁶. The final choice of airfoil was the Clark-Y foil due to it’s high lift to drag ratio. It has a maximum thickness of 11,7% at 30,4% of the chord. The angle of incidence was set to +5, 4^◦to maximize the lift to drag ratio during steady state level flight.

Airfoil

CL

C_D

max CL CD

Clark-Y 52,1 @ 5, 4^◦ 0,98 0,019 NACA 2412 49,8 @ 6, 9^◦ 0,98 0,019 Table 3.2: Comparing airfoils for the main wing

Figure 3.4: The Clark-Y profile

3.6.2 Aileron design

The aileron-to-wing-chord ratio was set to 20%. With a wing chord of 0.8 m, the aileron chord was set to 0.16 m. The aileron-to-wing-span ratio was set to 25%. With a wing span of 6 m, the final aileron chord length is 0.75 m. The max deflection was set to +- 25 degrees.

3.6.3 Ice protection system

To keep ice from accumulating on the leading edge of the wing, a anti-ice protection system is needed. For this aircraft, a passive system could be used. Such a system employs hydrophobic surfaces that repells water. This surface can either be a coating or some sort of fabric.

3.7 Tail design

3.7.1 Horisontal stabilizer

From the pitch stability analysis, the NACA 0012 was chosen as the horisontal stabilizer.

To eliminate the pitch moment about the a.c for the tail, a symmetrical airfoil was

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chosen. To fully satisfy the pitch moment about CoG, the angle incidence was chosen to be +4^◦. The final stabilizer chord was set to 0.45 m and the stabilizer span was set to 1.8 m.

Figure 3.5: NACA 0012 airfoil

Elevators

The elevator-to-stabilizer-span ratio was set to 0.9, giving a final elevator span of 0.81 m. The elevator-to-stabilizer-chord ratio was set to 0.3, giving a final elevator chord of 0.13 m. The maximum deflection was set to +-25 degrees.

3.7.2 Vertical stabilizer

For a Cessna 177, the planform area of the vertical stabilizer is 10% of the main wing planform area. Also, the aspect ratio of the vertical stabilizer is about 1.71 [8]. Using these number as reference, gives a vertical stabilizer with a chord of 0.5 m and a span of 0.85 m.

Rudder

The rudder-to-stabilizer-chord ratio was set to 0.40, and the rudder-to-stabilizer-area ratio was set to 0.40. Using these ratios, the rudder chord was set to 0.2 m and the rudder span was set to 0.80 m.

3.8 Center of gravity

The center of gravity has been placed within the fuselage, where the chord of the wing is at it’s thickest. The ”free” masses (the avionics, battery pack, and engine) has been placed in order to achieve the required CoG. The landing gear configuration is assumed to have a horisontal CoG that coincides with the required one. Figure 3.6 and Table 3.3 below shows the placement of the masses.

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CoG x y z

Wings 0 0 0,12

Fuselage 0,45 0 0

Empennage 2,626 0 0,13

Landing gear down 0 0 -0,7

Landing gear up 0 0 -0,3

Propeller 2,66 0 0

Engine 1,66 0 0

Battery pack -0,44 0 0

Avionics -0,94 0 0

Total 0 0 0

Table 3.3: Center of gravity

Figure 3.6: Figure showing placements of masses

3.9 Optimisation of climb angle

To minimize the energy consumption during ascent and steady state level flight, it is important to study the angle of climb. It is not completely obvious how the energy usage depends on the angle of climb, since the air density is changing with altitude, affecting the performance. A way of studying this is to let the climb angle vary, while calculating the total amount of energy required for each angle. The aircraft must reach an altitude of 1 km within a horisontal range of 10 km. This gives that the minimum climb angle required is around 5.7^◦. If the climb angle increases, the horisontal climb

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distance decreases while the steady state level flight distance increases.

Figure 3.7: Optimisation of climb angle

To keep a constant climb angle and a constant speed during, the engine power is forced to change with altitude. Figure 3.8 below presents the results from this optimisation.

Figure 3.8: Climb energy optimisation

At a climb angle of 19.3 degrees, the engine power exceeds the recommended continuous power of 14 kW. From Figure 3.8 it is clear that the best climb angle for this purpose

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is tbe lowest possible. Therefore, the climb angle is set to 5.7 degrees. By using this climb angle compared to using 19.3, an energy save of 3.23% is obtained.

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Chapter 4 Avionics

The word ”Avionics” comes from the words Aviation and Electronics. It is a collective word for all the electrical components in an aircraft such as wiring, navigation systems, flight controls and so on. The aircrafts mission is to scout a predetermined route. To be able to do this, the aircraft has to be equipped with the necessary electronics for it to complete it’s mission.

4.1 Navigation

For the aircraft to be able to find its location on the earth and make sure it correctly follows its route it needs to be equipped with a GPS. GPS stands for Global Positioning Systemand it works by uses three different orbiting satellites to pinpoint its location.

The aircraft will use this technology to be able to steer its way through the predetermined route.

Figure 4.1: How a GPS works

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4.2 Flight Control

For the aircraft to be able to steer its way it needs to be equipped with some sort of flight control system. This system will make sure that the aircraft can turn left and right, and ascend/descend at the required speeds for the aircraft to be able to complete its route.

There are several different kinds of UAV navigation systems, used for different sizes and missions.

The flight control system picked for this UAV is called Proton [3]. The Proton weighs only 50 grams, and contains a combined set of gyroscopes, accelerometers and GPS, all there to control the flight on its course, whilst drawing only 0.7 W. The specifications of the Proton are listed below.

Figure 4.2: Flight control system Proton

4.3 Observation equipment

When chosing a observation system for the aircraft it is important to have one with a gyro stabilizer to make sure you get a steady focused picture. A gyro stabilizer is a device for measuring orientation, meaning it can measure its position around a certain axis. This way it knows if its wrongly positioned, and if it needs to adjust itself. The stabilizer works by using a spinning wheel in which the axis is free to assume any orientation without changing the axis of the wheel.

The equipment used is called E-sky 02 and is produced by Dat Con [2]. This camera device is used by other UAVs in a wide range of fields, and is explained as a ”day and night gyro stabilizer observation system capable of stabilizing demanding oscillations

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up to approximately 45^◦. It operates with a continous power of 3.6 W, and weighs only 213 grams. There are two different E-sky models, either with thermal imaging, E-sky 02, or without thermal imaging, E-sky 01. For this mission a thermal imaging system is of great use, hence the pick.

Figure 4.3: E-sky 02

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Summary

The main goal of the project was to be able to plan a route for an unmanned aerial vehicle, UAV, along the coast of Antarctica, which has now been successfully done.

The engine chosen for the project was a six phased synchronous brushless electric engine provided by Joby Motors, with a continuous shaft power of 14 kW, and a peak of 20.9 kW.

Along with the motor, a battery pack made of lithium polymere cells were chosen.

These are some of the more powerful batteries on the market, and they work well under low temperatures.

For the main wing, a profile named Clark-Y was chosen, due to its high lift to drag ratio.

The propeller of the aircraft has been provided by the website pipistrel.si and is called BAM. It has three blades and a diameter of 1.66 m.

To be able to navigate its own course, the aircraft had been installed with a flight control system called Proton. To be able to properly scout the area beneath, a camera called E-Sky 02had been installed, mounted right beneath the nose of the aircraft.

When combining the complete aircraft into one piece, the center of gravity has been placed where the main wing chord is at its thickest.

The project will begin during the September month, due to longer days of the win- ter half year, meaning more sunshine for the solar cells. When the aircraft is on the runway, it has a 300 m runway ahead. During take-off, a continuous thrust power of 15 kW is generated. The aircraft travels 82 m in almost 4.7 seconds, and takes off at a speed of 25.6 m/s. During the ascent, the aircraft climbs at an angle of 5.7 degrees at a speed of 30.2 m/s, until reaching an altitude of 1 km. This takes approximately 5.3 minutes.

At this point, it continues to travel at approximately the same speed for 83.5 km, which takes about 47.9 minutes. After this phase, the aircraft starts to descend. During this phase the engine is shut off, and the aircraft will glide a distance of 6.5 km, which takes about 3.7 minutes. The aircraft will touch down with a speed of 25.6 m/s, where the brakes will be applied after 1 second. The total distance from touch down to stop is 119-232 m, depending on the ground conditions. At this point it will be brought to the hangar where it will recharge with the help of solar panels, making it ready for the next days flight. The entire project will continue once a day for a six month period. At the end of this period, the data of the glacier changes will be collected and analysed.

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Bibliography

[1] http://web.mit.edu/16.unified/www/fall/thermodynamics/notes/node86.html.

[2] http://www.dat-con.com/sites/default/files/download/e-sky-2012.pdf, May 2014.

[3] http://www.uavnavigation.org/products/uav-autopilot-proton, May 2014.

[4] http://www.xflr5.com/xflr5.htm, March 2014.

[5] http://www.timeanddate.com/worldclock/astronomy.html?n=468, April 2014.

[6] http://en.wikipedia.org/wiki/sunlight, April 2014.

[7] http://www.ise.fraunhofer.de/en/press-and-media/press-

releases/presseinformationen-2013/world-record-solar-cell-with-44.7-efficiency, April 2014.

[8] http://www.powerstream.com/li-pol.htm, May 2014.

[9] http://www.pipistrel.si, April 2014.

[10] http://www.jobymotors.com, April 2014.

[11] Arne Karlsson. The lift to drag ratio. 2004.

[12] Arne Karlsson. The aeroplane - some basics. 2012.

[13] Arne Karlsson. Steady climb performance with propeller propulsion. 2013.

[14] Arne Karlsson. Steady and level flight of an airplane with propeller propulsion.

2013.

[15] Arne Karlsson. Cruise performance of airplanes with propeller propulsion. 2013.

[16] Arne Karlsson. How to estimate cd0 and k and the simple parabolic drag polar.

2013.

[17] Arne Karlsson. Cruise performance. 2004.

[18] Mohammad Sadraey. Aircraft design: A systems enginering approach. 2012.

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[19] D.P. Raymer. Aircraft design: A conceptual approach. 1992.

[20] Arne Karlsson. Aeroplane weight, balance and pitch stability. 2013.

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Appendix

CAD-pictures

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Mission data

Engine power Speed Time Distance γ

Start 15 kW 0-25.6 m/s 4.7 s 82/83 m * -

Ascent 10.9 kW 30.2 m/s 5.3 min 10 km 5.7

SSF 7 kW 29 m/s 47.9 min 83.5 km -

Descent 0 kW 29 m/s 3.7 min 6.5 km -8.8

Landing 0 kW 25.6-0 m/s 20.4/9.2 s * 232/119 m * -

* Ground condition ice/dry

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Matlab code

1 close all

2 clear all

3 clc

4

5 [m,g,Wto,b,c,S,stot,H,v_sound,rho,Swet,CL,CFe,...

6 CD0,AR,e0,K,CD,eta_prop,eta_eng,...

7 CL_max,v_stall,P_eng_1] = Initconstants();

8

9 %% Weight estimation

10

11 W_wings = 15; % Wings [kg]

12 W_lg = 6; % Landing gear [kg]

13 W_tail = 8; % Tail [kg]

14 W_body = 30; % Fuselage [kg]

15 W_prop = 6; % Propeller [kg]

16 W_avion = 15; % Avioncics [kg]

17 W_bat = 50; % Battery pack [kg]

18 W_eng = 4; % Engine [kg]

19

20 TotalW = (W_wings+W_lg+W_tail+W_body+W_prop+W_avion+W_bat+W_eng);

21

22 disp(['Total mass of aircraft: ' num2str(TotalW) ' kg'])

23 disp(' ')

24

25 % Center of gravity

26 Wi = [W_wings W_lg W_tail W_body W_prop W_avion W_eng W_bat];

27 Xi = [ 0 0 -2.7 0 -3 0.5 2 0]';

28 Zi = [0.15 -0.5 0.2 0 0 0 0 0]';

29

30 Xcg = (Wi*Xi)/TotalW; % X-coordinate

31 Zcg = (Wi*Zi)/TotalW; % Z-coordinate

32

33 %% VELOCITY FOR MINIMUM POWER DURING SSF

34

35 Vminpwr = ((2/rho(end))*((K/(3*CD0))ˆ(0.5))*(Wto/S))ˆ(0.5);

36

37 %% TAKE OFF (Ground roll)

38

39 % Solving ground roll time and distance with ODE45

40

41 my_start = [0.025 0.05];

42

43 index = 1;

44 global mu_s

45 for mu_s = my_start

46 [t_s,z] = ode45(@startfil,[0 20],[0;10e-5]);

47 s1 = z(:,1); % Distance

48 v1 = z(:,2); % Velocity

49 s1_spar(:,index)=s1;

50 v1_spar(:,index)=v1;

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51

52 index=index+1;

53 end

54 55 56

57 subplot(211)

58 plot(t_s,s1_spar)

59 title('Distance as function of time')

60 xlabel('Time [s]'), ylabel('Distance [m]'), grid on

61 subplot(212)

62 plot(t_s,v1_spar)

63 title('Velocity as function of time')

64 xlabel('Time [s]'), ylabel('Velocity [m/s]'), grid on

65 hold on

66

67 [˜,j1] = min(abs(v1_spar(:,1)-1.2*v_stall)); % Find ...

lift-off speed

68 [˜,j2] = min(abs(v1_spar(:,2)-1.2*v_stall)); % Find ...

lift-off speed

69

70 E1 = (P_eng_1*t_s(j1))/(eta_eng); % Energy req to ...

engine [J]

71

72 disp(' ')

73 disp('TAKEOFF PERFORMANCE---')

74 disp(['Takeoff speed: ' num2str(1.2*v_stall) ' m/s'])

75 disp(['Engine power: ' num2str(P_eng_1/1000) ' kW'])

76 disp(['Ground roll time: ' num2str(t_s(j1)) ' seconds'])

77 disp(['Ground roll distance: ' num2str(s1_spar(j1,1)) ' meters'])

78 disp(['Ground roll time: ' num2str(t_s(j2)) ' seconds'])

79 disp(['Ground roll distance: ' num2str(s1_spar(j2,2)) ' meters'])

80

81 %% ASCENT ...

---

82

83 s2hor = stot/10; % Horisontal distance [m]

84 s2 = sqrt(Hˆ2+s2horˆ2); % Distance speed ...

direction [m]

85 alpha = acosd(s2hor/s2); % Angle of attack ...

[degrees]

86

87 H_as = linspace(0,1000,1000); % Cruise ...

altitude SSF

88 s_as = H_as./sind(alpha); % Distance

89 v_as = 1.35*sqrt(2*Wto*cosd(alpha)./... % Speed [m/s]

90 (CL*rho(end)*S));

91 T_as = CD*rho.*S*0.5.*(v_as.ˆ2)+Wto*sind(alpha); % ...

Propeller thrust [Nm]

92 T_average = sum(T_as)/length(T_as); % Average ...

thrust

93

94 ROC = v_as*sind(alpha); % Rate of ...

climb

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95

96 D_prop = 1.66; % ...

Propeller diameter

97 Adisk = pi*(D_prop/2)ˆ2; % Disk area

98 eta_prop_as = 2./(1 + sqrt(2*T_as./... % ...

Propeller efficiency

99 (Adisk*(v_as.ˆ2).*rho) + 1));

100

101 P_as = (T_as.*v_as)./(eta_prop_as); % Engine ...

power climb [W]

102 P_average = sum(P_as)/length(P_as);

103 ds_as = diff(s_as(1:2));

104 dt_as = ds_as./v_as;

105 t2 = sum(dt_as);

106 E2 = sum(P_as.*dt_as)/eta_eng; % Total ...

energy for climb

107

108 disp(' ')

109 disp('CLIMB PERFORMANCE---')

110 disp(['Rate of climb: ' num2str(ROC) ' m/s'])

111 disp(['Angle of climb: ' num2str(alpha) ' degrees'])

112 disp(['Speed: ' num2str(v_as) ' m/s'])

113 disp(['Average engine power: ' num2str(P_average/1000) ' kW'])

114 disp(['Average propeller thrust: ' num2str(T_average) ' Nm'])

115 disp(['Time: ' num2str(t2/60) ' minutes'])

116 disp(['Distance: ' num2str(s2hor) ' meters'])

117

118 %% DESCENT ...

---

119

120 Index = 1;

121 for u = 0.1:0.01:30

122 vdes = sqrt(Wto*sind(u)/(CD*rho(2)*S*0.5));

123 Descentspeed(:,Index) = vdes;

124 if abs(1.35*Vminpwr-vdes) <= 0.01

125 k = [vdes u];

126 u = 20;

127 else

128 end

129 Index = Index+1;

130 end

131

132 s4 = H/sind(k(2)); % Distance [m]

133 s4hor = s4*cosd(k(2)); % Horisontal ...

distance [m]

134 t4 = s4/k(1); % Time [s]

135 ROD = k(1)*sind(k(2)); % Rate of descent

136

137 disp(' ')

138 disp('DESCENT PERFORMANCE---')

139 disp(['Speed: ' num2str(k(1)) ' m/s'])

140 disp(['Rate of descent: ' num2str(ROD) ' m/s'])

141 disp(['Angle of descent: ' num2str(k(2)) ' degrees'])

142 disp('Engine power: 0 kW')

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145

146 %% STEADY STATE FLIGHT ...

---

147

148 s3hor = stot - s2hor - s4hor; % Distance [m]

149 v3 = 1.35*Vminpwr; % Speed [m/s]

150 T3 = CD*rho(end)*S*0.5*(v3ˆ2); % Propeller ...

thrust [N]

151 P_eng_3 = (T3*v3)/(eta_prop); % Engine power [W]

152 t3 = s3hor/v3; % Time [s]

153 E3 = P_eng_3*t3/eta_eng;

154

155 disp(' ')

156 disp('SSF PERFORMANCE---')

157 disp(['Speed: ' num2str(v3) ' m/s'])

158 disp(['Engine power: ' num2str(P_eng_3/1000) ' kW'])

159 disp(['Propeller thrust: ' num2str(T3) ' Nm'])

162

163 disp(' ')

164 disp(['Total mission time: ' num2str((t2+t3+t4)/60) ' minutes'])

165

166 %% ENERGY ...

---

167

168 Etotal = E1+E2+E3; % Total engine energy ...

needed [J]

169 kWh = (Etotal/3600000); % Total kWh from battery

170 disp(' ')

171 disp(['Total kWh required from battery is: ' num2str(kWh)])

172 173

174 %% LANDING ...

---

175

176 my_break = [0.38 0.12]; % [Dry ground, Wet ice]

177 A_land = g*(0-my_break);

178 B_land = (g/m)*(0.5*rho(1)*S*(CD-my_break*CL));

179 S_land = abs((1./(2*B_land)).*log(1-((B_land./A_land).*...

180 ((1.2*v_stall)ˆ2))))+1*1.2*v_stall;

181 t_stop = 1 + S_land./((1.2*v_stall)/2);

182

183 index = 1;

184 global mu_l

185 for mu_l = my_break

186 % Solving ground roll time and distance with ODE45

187 [t_b,z] = ode45(@landfil,[0 30],[0;sqrt(2*Wto/(CL*rho(1)*S))]);

188 s = z(:,1); % Distance

189 v = z(:,2); % Velocity

190 sspar(:,index)=s;

191 vspar(:,index)=v;

UAV Antarctica

Contents

Foreword

Introduction

Background

Chapter 1

Requirement specification

Chapter 2

Perfomance analysis

Chapter 3

Dimensioning

Chapter 4

Avionics

Summary

Bibliography

Appendix